MODULE 11 COMPLETE

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MODULE 11.01 Theory of Flight

Contents 1 MODULE 11 (AEROPLANE AERODYNAMICS, STRUCTURES AND SYSTEMS) .......................................................................................... 1-2 1.1

1.2

AEROPLANE AERODYNAMICS AND FLIGHT CONTROLS ..................... 1-2 1.1.1 Fixed Aerofoils .............................................................. 1-2 1.1.2 Moveable Control Surfaces ........................................... 1-6 1.1.3 High Lift Devices ........................................................... 1-13 1.1.4 Drag Inducing Devices .................................................. 1-14 1.1.5 Airflow Control Devices – Wing Fences......................... 1-17 1.1.6 Boundary Layer Control ................................................ 1-18 1.1.7 Trim Tabs ...................................................................... 1-21 1.1.8 Mass Balance ............................................................... 1-24 1.1.9 Control Surface Bias ..................................................... 1-26 1.1.10 Aerodynamic Balance – Horn Balance .......................... 1-26 1.1.11 Aerodynamic Balance – Inset Hinge.............................. 1-27 HIGH SPEED FLIGHT ..................................................................... 1-28 1.2.1 Speed of Sound ............................................................ 1-28 1.2.2 Subsonic Flight ............................................................. 1-29 1.2.3 Transonic Flight ............................................................ 1-30 1.2.4 Supersonic Flight .......................................................... 1-32 1.2.5 Aerodynamic Heating .................................................... 1-39 1.2.6 Area Rule ...................................................................... 1-40 1.2.7 Factors Affecting Airflow in Engine Intakes of High Speed Aircraft 1-41 1.2.8 Effects of Sweepback on Critical Mach Number ............ 1-43

Issue 1 – 04 Sept 2001

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MODULE 11 (AEROPLANE AERODYNAMICS, STRUCTURES AND SYSTEMS)

The principles of Aircraft Theory of Flight are covered in JAR 66 Module 8.

1.1 AEROPLANE AERODYNAMICS AND FLIGHT CONTROLS

An aircraft is equipped with fixed and moveable surfaces, or aerofoils, which provide stability and control. Each item is designed for a specific function during the operation of the aircraft.

Typical Aircraft Flight Controls Figure 1 1.1.1 FIXED AEROFOILS

The fixed aerofoils are the wings or mainplanes, the horizontal stabiliser or tailplane and vertical stabiliser or fin. The function of the wings is to provide enough lift to support the complete aircraft. The tail section of a conventional aircraft, including the stabilisers, elevators and rudder, is occasionally known as the empennage.

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Horizontal Stabiliser

The horizontal stabiliser is used to provide longitudinal pitch stability and is usually attached to the aft portion of the fuselage. It may be mounted either on top of the vertical stabiliser, at some mid-point, or below it. Conventional horizontal stabilisers are placed aft of the wing and normally set at a slightly smaller or negative angle of incidence with respect to the wing chord line. This configuration gives a small downward force on the tail with a value dependent on the size of the stabiliser and its distance from the Centre of Gravity (CG).

Horizontal Stabiliser Figure 2


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T-Tail Arrangement

The T-Tail Arrangement places the complete stabiliser/tailplane and elevator assembly on top of the vertical stabiliser. This ensures that pitch control is not affected by turbulent air from the wing. It also makes the vertical stabiliser and rudder control more effective, due to the so-called „end plate effect‟. However a T-Tail (and rear engine) configuration, would be dangerous if the aircraft entered what is termed a „deep stall’. At a very high angle of attack (i.e.: stalling angle), airflow could make pitch control non-effective (and may cause the engines to flame out). To prevent this, T-Tailed aircraft will have a „stick push‟ system, in order to automatically recover them safely from excessive angles of attack. The T-Tail has another disadvantage in that the empennage structure will be heavier than normal, due to the strengthening required to combat greater bending loads. However since the pitch moment arm is increased, the stabiliser and elevators can be made smaller and therefore lighter than conventional designs. Often, the complete stabiliser can be moved to provide longitudinal trim, negating the use of trim tabs (later in Module 11.09).

T–Tail Arrangement Figure 3 Page 1-4


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Vertical Stabiliser

The vertical stabiliser for an aircraft is the aerofoils forward of the rudder and is used to provide directional stability. A problem encountered on single-engined propeller driven aircraft is that the propeller causes the airflow to rotate as it travels rearward. This strikes one side of the vertical stabiliser more than the other, resulting in a yawing moment. These aircraft may have the leading edge of the stabiliser offset slightly, thereby causing the airflow to pass around it in such a manner to counter the yaw.

Off-Set Vertical Stabiliser Figure 4


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1.1.2 MOVEABLE CONTROL SURFACES

Moveable control surfaces are normally divided into Primary and Secondary controls. The primary control surfaces include the elevators, rudder, ailerons and roll spoilers. The secondary control surfaces consist of trim controls (tabs), high lift devices (flaps and slats), speed brakes and lift dumpers (additional spoilers). Note: Traditionally, spoilers have not been included as primary controls, but those which operate in conjunction with the ailerons during roll, are considered to be primary in the JAR 66 syllabus, so this is how these notes will define them. The primary control surfaces are used to make the aircraft follow the correct flight path and to execute certain manoeuvres. The secondary controls are used to change the lift and drag characteristics of the aircraft or to provide assistance to the primary controls.

Moveable Control Surfaces Figure 5

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Roll Control - Ailerons

These primary controls provide lateral (roll) control of the aircraft, that is, movement about the longitudinal axis. They are normally attached to hinges at the trailing edge of the wing, near the wing tip. They move in opposite directions, so that the up-going aileron reduces lift on that side, causing the wing to go down, whilst the down-going surface increases the lift on the opposite side, raising the wing. Large aircraft often use two sets of aileron surfaces on each wing, one in the conventional position near the wing tip and the other set at mid-span or outboard of the flaps. The inboard set is referred to as „high speed ailerons‟. The outboard surfaces, or sometimes both sets, work at low speeds to give maximum control during take off and landing, for example when large movements may be required. At high cruising speed the outer ailerons are isolated and only the inboard set operate. If the outer ailerons were permitted to operate at high speed, the stress produced at the wing tips may twist the wing and produce „aileron reversal‟. This is particularly likely with modern highly flexible thin wings, where the possibility of structural damage may result if the outboard surfaces were too powerful. The ailerons are operated by a control wheel, a control column or a side-stick. Movement of any of these inputs away from neutral towards one side, will result in the aircraft rolling to that side. Returning the control to neutral at this stage will leave the aircraft in a banked condition and a similar but opposite movement will be required to bring the aircraft level once more.

Aileron Controls Figure 6


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The ailerons are usually operated in conjunction with the rudder and/or elevator during a turn and are rarely used on their own. A co-ordinated turn is one that occurs without slip or skid. Too little bank will cause the aircraft to skid outwards, too much bank will cause the aircraft to slip downwards.

1.1.2.2

Roll Control - Spoilers

The use of spoilers as a primary control, will be to operate asymmetrically in conjunction with aileron movement and are normally referred to as Roll Spoilers. Roll spoilers are mounted on the top of the wing just inboard of the outboard set of ailerons.

Roll Spoiler Controls Figure 7 Movement of the aileron control wheel on the flight deck will deploy each spoiler progressively upwards with the up-going aileron, whilst on the side of the downgoing aileron, the spoiler will remain flush with the upper wing camber. .This is achieved by the control system being routed via a spoiler/aileron mixer unit. The up-going spoiler will effectively spoil the lift on the down-going wing and augment the similar effect of the up-going aileron. Alternatively, on some aircraft the spoilers will replace the ailerons completely to provide the sole means of roll control. Note: Other spoiler functions are covered later under Secondary Controls.

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Pitch Control - Elevators

The elevators are the control surfaces which govern the movement of the aircraft in pitch about its lateral axis. They are normally attached to the hinges on the rear spar of the horizontal stabiliser. When the control column of the aircraft is pushed forward, the elevators move down.. The resultant force of the „airflow generated lift', acting upwards, raises the tail and lowers the nose of the aircraft. The reverse action takes place when the control is pulled back.

1.1.2.4

Pitch Control – Stabilators

A special type of pitch control surface that combines the functions of the elevator and the horizontal stabiliser is the stabilator, often referred to as a slab or allflying tailplane . The stabilator is a complete all-moving horizontal stabiliser which can change its angle of attack when the control column is moved and thereby alter the total amount of lift generated by the tail.

Elevator Controls Figure 8

Stabilator Controls Figure 9


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Pitch Control – Variable Incidence Stabilisers

Incorporating a conventional elevator control system, the variable incidence horizontal stabiliser is often used for pitch trim. Normally a powerful electric motor is used to vary its angle of attack when trim switches on the flight deck are operated.

Variable Incidence stabiliser Figure 10 1.1.2.6

Canards

Some earliest powered aircraft, such as the Wright Flyer, had horizontal surfaces located ahead of the wings. This configuration, with the forward surface usually referred to as a canard or foreplane, has been used on occasions, up to the present day. Conventional aircraft have the tailplane located at the rear of the fuselage which provides a small, stabilising down force. This means that the wing has to produce slightly more lift to balance this down force. As we have seen, in order for a wing to produce lift it must also generate drag. With the tailplane located at the front of the aircraft, the stabilising force is directed upwards. This contributes to the total lift of the aircraft, thereby reducing drag from the lift producing wing.

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A fundamental feature of a canard design is that the angle of attack of the foreplane, (in front of the CG of the aircraft) is set at a greater angle than the main wing. This feature will ensure that the foreplane reaches the stalling angle first, resulting in a predictable dropping of the nose and a certain recovery. Additionally, stall sensing systems (later), can be triggered just before the foreplane reaches its critical angle of attack, leaving the main wing safely below the stalling angle and still producing adequate lift.

Canard Design – Beech Starship Figure 11 1.1.2.7

Yaw Control - Rudder

The rudder is a vertical control surface that is hinged at the rear of the fin and is designed to apply yawing moments. The rudder rotates the aircraft about its vertical axis and is controlled by rudder pedals that are operated by the pilots‟ feet. Pushing on one pedal, the right for example, causes the rudder to move to the right also. This causes the rudder to generate a 'lifting' force sideways to the left which turns the nose of the aircraft to the right. Because of the power of some rudder systems, particularly assisted systems, they may have their range reduced at high speed by means of a speed-sensitive range limiting system.(later). The rudder is normally a single structural unit but on large transport aircraft it may comprise two or more operational segments, moved by different operating systems to provide a level of redundancy. Issue 1 – 04 Sept 2001

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Rudder controls Figure 12

1.1.2.8

Combined-Function Controls – Elevons and Ruddervators

An example of combined-function controls is found on delta-wing aircraft, where control surfaces for pitch and roll must be fitted on the trailing edge of the wing. Controls with a dual-function (elevators and ailerons) called elevons, provide both pitch and roll, by moving symmetrically in pitch or asymmetrically in roll via a mixer unit, when the control column or control wheel are operated on the flight deck..

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Another example are ruddervators normally used on aircraft fitted with a 'V' or Butterfly tail. These surfaces serve the purposes of both rudder and elevator.

Ruddervator Controls Figure 13 1.1.3 HIGH LIFT DEVICES

Aerodynamic lift is determined by the shape and size of the main lifting surfaces of the aircraft. In order to produce the outstanding performance achieved by a large modern, swept wing, passenger jet such as the Boeing 777, the wing is designed to give optimum lift to support the aircraft whilst in cruise (typically Mach 0.87). This has meant, that to be able to control and land the aircraft weighing around 200-tonne on runways of reasonable length, the landing speed needs to be slower than the „clean‟ stalling speed of the aircraft. In order to achieve this, more lift is required and this is obtained from so-called high lift devices. These are divided generally into leading edge devices, namely slots, slats and Krueger flaps and trailing edge devices including plain, slotted and fowler flaps. They will increase lift and as a result, reduce the stalling speed. Consequently the landing speed, (about 1.3 times the stalling speed), will also be reduced, since drag is also increased with large angles of trailing edge flap deployment.


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Flaps and Slats Figure 14 Additionally, some aircraft incorporate ailerons, both of which are designed to move downwards together whenever the trailing edge flaps are extended to the landing position. These will act as additional plain flaps and provide extra drag (and lift), but will still provide roll control if required. These surfaces are referred to as „Droop Ailerons‟ or „Flaperons‟.

Droop Aileron Figure 15 1.1.4 DRAG INDUCING DEVICES

There are several situations where the aircraft must slow down fairly quickly. With slower, high drag, light aircraft, simply closing the throttle allows the high drag of the airframe and the idling propeller to slow the aircraft down, to gliding speed prior to landing approach, for example.

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As previously stated, a modern airliner is an extremely smooth, low drag design which, if only the throttles are retarded, will continue in level flight for many miles before slowing down. Furthermore, if the nose were lowered more than a degree or so, the aircraft will begin to accelerate again. In order to overcome the problems of low drag on large aircraft with high momentum, the designers have introduced a variety of drag inducing devices. These include spoilers, lift dumpers, speed brakes and in unusual circumstances, lowering the landing gear and operating in-flight thrust reversers.

1.1.4.1

Spoilers and Lift Dumpers.

Spoilers and Lift Dumpers are usually hinged panels located about mid-chord position on the upper surface of the wing. Hydraulically operated, they produce a large amount of turbulence and drag when deployed, resulting in a reduction of lift.

Lift Dump Spoilers Figure 16 Spoilers, have a variety of uses, all of which involve spoiling the lift of the wing. Some of the following facilities can be combined, so that one set of panels can have more than one job. Firstly, they can be the primary roll control of the aircraft as described previously. Secondly, the spoilers can be used in a symmetrical, part-deployed position, allowing the aircraft to slow down quickly in the cruise, or descend at a much steeper rate without accelerating. On some aircraft, the deployment angle of the spoiler panels can be varied by changing the position of the control lever in the flight compartment.


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Lift dumpers are, as their name describes, are spoiler panels incorporated solely to dump lift. They are normally deployed after landing, destroying the lift of the wing and producing high drag, to assist in stopping the aircraft efficiently and thereby allowing the wheel brakes to be operated more effectively.

1.1.4.2

Speed Brakes

Whilst it is true that the in-flight use of spoilers may be referred to as selecting the 'speed brakes', the term more accurately describes devices which are solely for the production of drag without any change of trim. The rear fuselage mounted 'clamshell-type‟ doors which open up on the BAe 146 and Fokker 70/100 aircraft are true speed brakes (or air brakes) and have the following major advantage over the use of spoilers for producing drag. When the wing mounted spoilers are deployed, vibration or rumble is often felt in the passenger cabin, which some people may find disturbing. The aft mounted speed brakes not only produce high drag at any airspeed, but their selection is virtually vibration free. Also, lift will be completely unaffected, thus permitting their deployment on approach and making a go-around much safer. (This will be covered later in powerplants).

Speed Brake Installation Figure 17

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1.1.5 AIRFLOW CONTROL DEVICES – WING FENCES

These devices are usually fitted to aircraft with swept wings. Total airflow over a swept wing, splits into two components, one moving across the wing chord parallel to the airflow and the other flowing spanwise towards the wing tip. The fences are fitted about mid-span, on the leading edge of the wing and extending rearwards. They are designed to control the spanwise flow of the boundary layer air over the top of the wing. Also they will straighten the airflow over the ailerons, improving their effectiveness and straighten the air nearer the wing tip, resulting in less 'spillage' of air from beneath the wing to the top, thereby producing less drag. (See Winglets later).

Wing Fences Figure 18 1.1.5.1

Airflow Control Devices – Saw Tooth Leading Edges

This form of airflow control is more common on military aircraft than modern commercial airliners. The saw tooth or notch is simply a small increase in wing chord on the outer portion of the wing. The step where the change occurs, tends to form an invisible 'wall' of high velocity air, which flows over the wing and straightens the spanwise flow. It functions in much the same way as the wing fence but removes the extra drag and weight penalty.

Leading Edge Notch Figure 19 Issue 1 – 04 Sept 2001

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Airflow Control - Winglets

These can be seen on a variety of the later generation airliners and business jets. The outboard part of the wing are upswept to an extreme dihedral angle. These winglets work best at higher speeds and, by clever aerodynamic design, will give better airflow control and reduce the drag produced by the wing. It does this by using the up-flow from below the wing to produce a forward thrust from the winglet, rather like a yacht sail. The winglets add weight to the aircraft as well as increasing parasitic drag, but the large reduction in induced drag at the wingtip, results in a significant fuel saving.

Winglets Figure 20 1.1.6 BOUNDARY LAYER CONTROL

The boundary layer is that layer of air adjacent to the aerofoil surface (the boundary between „metal‟ and „air‟). If measured, the air velocity in the layer will vary from zero directly on the surface, to the relevant velocity of the free stream at the outer extremity of the boundary layer. Normally, at the leading edge of the wing the boundary layer will be laminar, (in smooth thin sheets close to the surface), but as the air moves over the wing towards the trailing edge, the boundary layer becomes thicker and turbulent. The region where the flow changes from laminar to turbulent is called the transition point. .As airspeed increases, the transition point tends to move forward, so the designer tries to prevent this thus maintaining laminar flow, over the top of the wing for as far back as possible. Methods of boundary layer control are as follows:

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Boundary Layer Control - Vortex Generators

One way of stimulating the boundary layer and stopping the airflow becoming increasingly sluggish towards the trailing edge is the use of vortex generators. Vortex generators are small plates or wedges projecting up from the surface of an aerofoil about 25mm.(about 3 times the typical boundary layer thickness), into the free stream air. Their purpose is to shed small but lively vortices from their tip, which act as scavengers to direct and mix the high energy free stream air into the sluggish boundary layer air and invigorate it. This action pushes the transition point backwards towards the trailing edge . In this way,the small amount of drag created by the vortices is far more than compensated by the considerable boundary layer drag which they save. They also weaken the shock wave at high speed and reduce shock drag also. (later).

Vortex Generators Figure 21

1.1.6.2

Boundary Layer Control - Stall Wedges

We have seen previously that washout on a wing permits the root of the wing to stall first, allowing the pilot to retain roll control during the stall. Even with a degree of washout, the aircraft will „drop a wing‟ on occasions due to adverse boundary layer air causing the outer part of the wing to stall first. This can be overcome with the use of stall wedges, or stall strips, as they are sometimes known.


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Stall Wedges are small, wedge-shaped strips mounted on the leading edge of the wings at about one third span. The are designed to disrupt the boundary layer airflow, at large angles of attack approaching the stall, thus ensuring the airflow breaks away,(stalls), at the root end of the wing first. Additionally they produce a similar effect to a wing fence at smaller angles of attack resulting in a smoother airflow over the ailerons, thus retaining optimum roll control.

Stall Wedges Figure 22 1.1.6.3

Boundary Layer Control - Leading edge Devices

Other devices to prevent laminar separation at the low speed end of the range and thus control boundary layer air are leading edge droop flaps and Kreuger flaps. They can be a droop snoot or permanent droop type, or can be adjusted during flight.

Krueger (left) and Drooped (right) Leading Edge Flaps Figure 23

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1.1.7 TRIM TABS

During a flight an aircraft will develop a tendency to deviate from a straight and level „hands-off‟ attitude. This may be due to changes in fuel state, speed, load position or flap/landing gear selection and could be countered by applying a continuous correcting force to the primary controls. This would be fatiguing for the crew and difficult to maintain for long periods, so trim tabs are used for this purpose instead. Trim tabs move the primary control surface aerodynamically in the opposite direction to the movement of the tab. To correct an aircraft „nose down’ out of trim condition, the elevator tab is moved down, resulting in the elevator moving up, the tail of the aircraft moving down, so that the nose comes up, correcting the fault.

1.1.7.1

Fixed Trim Tabs

A fixed trim tab may be a simple section of sheet metal attached to the trailing edge of a control surface. It is adjusted on the ground by simply bending it up or down, to a position resulting in zero control forces during cruise. Alternatively, the tab is connected to the primary control by a ground-adjustable connecting rod. Finding the correct position for both types is by trial and error.

Fixed Trim Tab Figure 24

1.1.7.2

Controllable Trim Tabs

A controllable trim tab is adjusted from the flight deck, with its position being transmitted back to a flight deck indicator showing trim units, left and right of neutral.


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Flight deck controls are trim-wheel, lever, switch, etc., with the actuation of the tab by mechanical, electrical or hydraulic means. Trim facilities are normally provided on all three axes.

Controllable Trim Tab Figure 25 Note: Aircraft with hydraulic fully powered controls do not have trim tabs. Since fully powered controls are termed irreversible, trim tabs if fitted, would be aerodynamically ineffective. With these systems, trimming is achieved by moving the primary control surface to a new neutral datum.(later). 1.1.7.3

Servo Tabs

Sometimes referred to as the flight tabs, servo tabs are positioned on the trailing edge of the primary control surface and connected directly to the flight deck control inputs. They act as a form of „power booster‟, since pilot effort is only required to deflect the relatively small area of the servo tab into the air stream. Movement of the flight deck control input moves the tab up or down and the aerodynamic force created on the tab, moves the primary control, until the aerodynamic load on the control surface balances that on the tab. Moving the tab down will cause the primary control to move up and vice-versa.

Servo Tab Figure 26

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Balance Tabs

Balance tabs assist the pilot in moving the primary control surface. The flight deck controls are connected to the primary control surface whereas the balance tab, hinged to the trailing edge of the primary surface, is connected to the fixed aerofoil. For example, the elevator balance tab, will be connected by an adjustable rod to the horizontal stabiliser and is so arranged, that it tends to maintain the tab at the same relative angle to the stabiliser when the pilot moves the elevator. Aerodynamically, therefore, the tab is moving in the opposite direction to the control surface and assists its movement. Adjusting the length of the connecting rod will alter the displacement of the effective range of the tab about the mid-point datum. Some types of balance tab have more than one point of attachment and it is possible with these so called „geared balance tabs’, to alter the range of tab deflection. The function of a balance tab can also be combined with that of a trim tab, by adjusting the length of the balance tab connecting rod from the flight deck. This is usually achieved by installing a form of linear actuator in the rod and is termed a trim/balance tab (Geared balance and trim/balance tabs will be covered later in the notes).

Balance Tab Figure 27 1.1.7.5

Anti-Balance Tabs

Anti-balance tabs operate in a similar way aerodynamically as balance tabs but with a reverse effect. The difference is in the way it is connected to the fixed aerofoil. It is routed so that the tab moves, relative to and in the same direction as, the primary control surface. The effect is to add a loading to the pilot effort, making it slightly heavier and thus providing „feel‟, to prevent the possibility of over-stressing the airframe structure.

Anti-Balance Tab Figure 28 Issue 1 – 04 Sept 2001

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Spring Tabs

At high speed, control surfaces operated directly from the flight deck, become increasingly difficult to deflect from neutral, due to the force of the aerodynamic loads caused by the airstream around them. The spring tab is progressive in its operation and provides increasing aerodynamic assistance in moving the control surface, with an increase in aircraft forward speed. The flight deck controls are connected to the spring tab in a similar manner to the servo tab previously described, except the linkage is routed via a torque rod assembly (or spring box) attached to the primary control surface. When the aircraft is stationary or flying at low airspeed the airloads are nonexistent or very small. If the flight deck controls are deflected from neutral, the rigidity of the torque tube (or spring force) causes the primary control to be deflected together with the spring tab. The tab will remain in the same relative position with the primary control and consequently provides no additional aerodynamic assistance. As the aircraft flies faster, the increased force produced by the airflow, opposes the movement of the primary control surface from its neutral position. Deflection of the flight deck controls in this case causes the torque tube to twist (or the spring to compress), resulting in a deflection of the spring tab. The tab deflection provides an added aerodynamic load which assists the flight deck effort. The faster the aircraft flies, the greater the airflow force and therefore the greater the spring tab deflection, resulting in a progressively increasing assistance in moving the primary control.

Spring Tab Figure 29 1.1.8 MASS BALANCE

All aircraft structures are distorted when loads are applied. If the structure is elastic, as all good structures are, it will tend to spring back when the load is removed, or its point of application is changed. Since a control surface is hinged near its leading edge, the centre of gravity (C of G) will be behind the hinge and as a consequence, there will be more weight aft of the hinge line than in front of it . Page 1-24


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In the case of an aileron for example, should the air load distort the wing upwards, it is likely that the aileron will „lag‟ behind and distort downwards. This effectively produces an extra upward aerodynamic force which pushes the wing up even further. Due to its elasticity, the wing will spring back and the aileron will lag again but this time upwards, aerodynamically forcing the wing down further than it would normally go due to elastic recoil alone. Now the cycle is repeated and a high speed oscillation will result. This unwanted phenomenon is referred to as flutter. Flutter can be prevented if the C of G of the control surface is moved in line with, or slightly in front of, the hinge line. The normal way of achieving this is to add a number of high density weights, either within the leading edge of the surface itself or externally, ahead of the hinge line. The addition of these weights, normally made from lead or depleted uranium, is closely controlled and calculated to ensure that the exact balance is obtained. This procedure of adding weights is referred to as mass balancing of the controls.

External Mass Weights Figure 30

Integral Mass Weights Figure 31 Issue 1 – 04 Sept 2001

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1.1.9 CONTROL SURFACE BIAS

When a control surface is set so it is not in the true neutral position it is referred to as having a bias. There are many reasons for not having the controls in a true central position, including compensating for design features. As an example, a single propeller aircraft may have a tendency to roll in the opposite direction to the engines torque, to counteract this moment the ailerons could be offset with one slightly up and the other down. Once the aircraft is flying level with the bias set the trim gauge in the cabin would then be set to read zero. 1.1.10 AERODYNAMIC BALANCE – HORN BALANCE

In order to overcome the high stick forces on larger aircraft at higher speeds, the surfaces themselves are used to lighten the forces. This is referred to as Aerodynamic Balancing and the three principal ways of achieving it are: horn balance, inset hinge and pressure balancing. This method, a small part of the primary control surface ahead of the hinge will project into the airflow when the control is deflected from neutral. The airflow on this side assists the movement of the control in the desired direction and will attempt to move the control further away from the neutral position. Air loads on the control side, aft of the hinge, try to push the surface back towards neutral. (This is the force that would normally make the controls heavy). If the proportion of balance area forward of the hinge and control area aft of the hinge is correct, the pilot will feel that his control loads are more manageable, making the aircraft easier to fly.

Horn Balance Figure 32 Page 1-26


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1.1.11 AERODYNAMIC BALANCE – INSET HINGE

This method is similar to and has the same effect as the horn balance. Instead of having a forward projection at one or both ends of the control surface, the hinges are set back so that the area forward of the hinge line, which projects into the air flow when the control surface is moved from neutral, is spread evenly along its whole length.

Inset Hinge Balance Figure 33 1.1.11.1

Aerodynamic Balance – Balance Panels

A device fitted to a few aircraft is the aerodynamic balance panel. Often used in the aileron system, the panel is fitted between the leading edge of the aileron, ahead of the hinge and the rear face of the wing. When the aileron is deflected upwards (downwards) from neutral, the high velocity, low pressure air passing over the lower (upper) gap decreases the air pressure under (above) the balance panel and pulls it down (up). The force on the balance panel is proportional to airspeed and control surface deflection and assists the pilot in moving the controls accordingly.

Aerodynamic Balance Panel Figure 34


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1.2 HIGH SPEED FLIGHT Advancement in modern aircraft and engine design has produced very large airliners capable of cruising at 87% of the speed of sound. Typically at an altitude of 11,000 metres (approximately 36,000feet), this will amount to an airspeed of about 575 miles per hour. Earlier in the course the effects of subsonic air were considered. As airspeed increases, the aerodynamic effects of airflow passing over an aircraft, go through a series of changes, which will now be considered.

1.2.1 SPEED OF SOUND

One of the most important measurements in high speed aerodynamics is based on the speed of sound and so called mach number. Mach number is named after the Austrian physicist Ernst Mach (1838-1916) and is the ratio of true airspeed of an aircraft to the local speed of sound at that altitude. (This will be covered in more detail later). Sound waves, like those produced by a stationary object vibrating at certain frequencies, will cause a continuous series of pulses or pressure waves, to radiate outwards equally in all directions from the point of origin and travel in exactly the same manner as the ripples on a pond.

Pressure Waves – Stationary Object Figure 35 The actual speed at which the waves radiate, depends on the type and density of the material in which they are travelling. Air and Water are both fluids but water is more dense than air, so sound waves will travel faster (about 4 times) in water than in air.

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Additionally, in any one of the fluids, speed will vary with a change in temperature. As temperature increases, the speed of sound will increase and vice-versa, so that in Air on a standard day at sea level (15oC approx), the waves will travel at 761mph (661.7 knots), whereas at 11,000 metres altitude, the speed will fall to 661mph, since the temperature has dropped to -56oC at this altitude. Note: At altitudes above 11,000 metres and up to about 27,000 metres, the temperature and hence the speed of sound, will remain constant.

1.2.2 SUBSONIC FLIGHT

The propagation of the pressure waves from a stationary object has been discussed above. When an aircraft begins to move through the air at subsonic speeds, (a speed less than pressure wave propagation speed) the waves still travel forward and it is as if a message is sent ahead of the aircraft to warn of its approach. On receipt of this message, the air streams begin to divide to make way for the aircraft but there is very little, if any change in the density of the air as it flows over the aircraft. This warning message can be detected perhaps 100metres in front of the aircraft. Consequently, anyone standing ahead of the aircraft, would hear it coming and be able to detect the change in the nature of the pressure waves as the aircraft passed by. It would be similar to the change in the pitch of the siren of a passing emergency road vehicle. This is often referred to as Doppler shift or Doppler effect.

Pressure waves – Subsonic Flight Figure 36


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1.2.3 TRANSONIC FLIGHT

At subsonic speeds, the study of aerodynamics is simplified by the fact that air passing over a wing experiences only very small changes in pressure and density. The airflow is termed incompressible as, when it passes through a venturi, the pressure changes without the density changing At higher speeds, the change in air pressure and density becomes significant and is called the compressibility effect. When air enters a venturi at supersonic speeds, the airflow slows down and must compress in order to pass through its throat. Once a fluid compresses, its pressure and density will both increase.

Subsonic Airflow Figure 37

Supersonic Airflow Figure 38

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The transonic flight range encompasses sound wave velocity and consequently is the most difficult realm of flight since some of the air flowing over the aircraft, particularly the wings, is subsonic and some is supersonic. As the aircraft approaches the speed of sound, the pressure waves ahead of it will be travelling at the same speed as the aircraft and are therefore relatively stationary. They accumulate to form a continuous pressure wave and consequently will result in the removal of any advance warning of the approach of the aircraft.

Transonic Flight Pressure Waves Figure 39 At these speeds other pressure waves, or shock waves form wherever the airflow reaches the speed of sound. These waves will upset the aerodynamic balance of the wing and this phenomenon will be covered later in the notes.


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1.2.4 SUPERSONIC FLIGHT

Once the aircraft is supersonic, all parts of it are considered to be above the speed of sound and therefore travelling faster than the rate of propagation of the pressure waves. An infinite number of pressure waves are produced and form a cone, the inclination of which will change as the aircraft speed changes.

Mach Cone Figure 40 1.2.4.1

Mach Number

As previously mentioned, Mach number is the ratio of the true airspeed of the aircraft and the local speed of sound at that altitude. An aircraft travelling at exactly the speed of sound is said to be travelling at Mach 1. It follows therefore that an aircraft travelling at twice the speed of sound would be travelling at Mach 2 and at half the speed of sound, Mach 0.5, etc,. The following definitions regarding airflow and mach number apply: Subsonic Flow Mach Numbers below Mach 0.75 Transonic Flow Mach Numbers between Mach 0.75 and Mach 1.2 Supersonic Flow

Mach Numbers between Mach 1.2 and 5.0

Hypersonic Flow

Mach Numbers above

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Critical Mach Number

At any constant aircraft forward speed, the speed of the airflow will vary over the curves and cambers on the different areas of the airframe. The behaviour of the airflow over the wing will be particularly significant, since this is the major lift provider for the aircraft. As air flows over the camber on the upper surface of the wing, its speed will increase as it flows rearwards from the leading edge, reaching a maximum at the thickest part of the wing chord. This means that although the aircraft itself may be travelling at an airspeed well below Mach 1, the airflow over the thickest part of the wing chord, may have already reached Mach 1 As will be discussed later, many unwanted effects occur when the wing approaches and reaches Mach 1. Therefore, the designers may either incorporate features that will lessen the unwanted effects, or limit the aircraft to a predetermined maximum airspeed, that will ensure the wing speed remains below Mach 1 and thus avoids the unwanted effects altogether. For each aircraft type therefore, a unique maximum aircraft forward speed will be calculated, corresponding to a wing speed of Mach 1. This aircraft speed (always be less than Mach 1) is called the Critical Mach Number or M.crit and nonsupersonic aircraft flying in the transonic flight range, will normally be limited to a maximum speed set below the Critical Mach number.

Critical Mach Number Figure 41 A thick wing will cause the airflow to speed up over the camber and reach Mach 1 more quickly than a thin wing of similar chord length. Consequently, the Critical Mach number for the thinner wing will be a higher value than the thicker wing. This in turn will mean that the aircraft with a thin wing, will be able to fly faster in the transonic flight range than the one with the thicker wing, before the unwanted effects caused by the wing reaching Mach 1 ensue. Conversely, less lift will be produced by a thin wing, than a thick wing of similar chord length, but this can be overcome by the so called Supercritical wing chord.


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In this design, the total amount of lift lost by the shallower camber of the thin wing is restored by making the chord longer. This is perfect for transonic cruise conditions, but at low airspeeds, lift on a clean wing will be insufficient and so extensive use of high lift devices (slots, slats and flaps) is necessary

Supercritical Wing Figure 42 1.2.4.3

Adverse Transonic Effects

Even though the onset of compressibility is gradual, it begins to have a significant effect as the Critical Mach number is approached. Unwanted adverse effects including, buffeting, shock waves, increase in drag, decrease in lift and movement of the centre of pressure occur. If uncontrolled, these effects could result in the aircraft becoming difficult to fly and to behave in a similar manner to a low speed high incidence stall, even though the aircraft is at high speed and low angle of incidence.

1.2.4.4

Compressibility Buffet

Previously discussed has been the build up of the pressure wave in front of the aircraft as it approaches Mach 1, including the fact that other parts of the airframe, in particular the wing, are likely to reach Mach 1 well before the complete aircraft does. When this occurs the smoothness of the airflow over the wing is severely affected. This region, as well as those on the flying control aerofoils, experience violent vibration and so-called compressibility buffeting of the airframe. If allowed to continue, control loss or possible structural damage can occur. 1.2.4.5

Shock Wave

Previously in the notes, the build up of pressure waves and the change from incompressible to compressible flow as the aircraft or an aerofoil surface approaches the speed of sound, has been discussed. Transonic flight presents major design problems for the aerofoil in particular, because only a portion of the airflow passing over the wing becomes supersonic.

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When an aerofoil moves through the air at a speed below its critical Mach number, all of the airflow is subsonic and the pressure distribution is predictable.The first indication of a change in the nature of the flow will be a breakaway of the airflow from the aerofoil surface as described previously in boundary layer control. Any turbulence resulting from the separation will cause an increase in drag and a corresponding reduction in the amount of lift. As speed begins to increase, the point of separation moves forward, extending the turbulent wake.

Subsonic Flow Over all the Surface Figure 43 However, as flight speed reaches and exceeds the critical Mach number, the airflow over the top of the wing speeds up to supersonic velocity and a shock wave starts to form.

The First Sonic Flow is encountered Figure 44

A Normal Shock Wave Begins to Form Figure 45


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Note: If the aerofoil is symmetrical and set at zero degrees angle of attack, the incipient shock wave as it is called, would form equally on the upper and lower surfaces. However, because the wing is usually set to an angle of incidence of about 3 degrees, even a symmetrical aerofoil section would produce the incipient wave on the top surface first. The wave extends outwards more or less at right angles to the aerofoil surface and is referred to as a normal (perpendicular) shock wave This normal shock wave forms a boundary between supersonic and subsonic airflow. As we have seen the high velocity airflow over the top of a wing creates an area of low pressure. The shock wave causes it to decelerate to subsonic speed, resulting in a rapid rise in pressure. The separation point and turbulent wake will now start from this point, resulting in a sudden and considerable increase in drag (about 10 times) and therefore a large loss of lift. Severe buffeting is likely, which could even lead to a shock stall and the centre of pressure will be altered, affecting the pitching moment. This „extra‟ drag, so called Shock Drag, will be made up of two components, namely Wave Drag, resistance caused by the wave itself and Boundary Layer Drag, due to the increased turbulent region over the surface of the wing. Furthermore, this shock-induced separation is likely to reduce flying control effectiveness The velocity of the air leaving the shock wave remains supersonic, so both the static pressure and the density of the air increase adding to the high drag/ low lift condition. Additionally, some of the energy in the airstream will be dissipated in the form of heat. As the aircraft speed continues to increase, the wave will extend outwards and begin to move aft towards the trailing edge of the wing. A second wave begins to form on the lower surface, as the airflow here also speeds up to supersonic velocity

Shock Induced Separation Occurs Figure 46

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As the airspeed reaches the upper end of the transonic range, both shock waves move aft, become stronger and will eventually attach to the wing's trailing edge.

Almost all Flow is Supersonic, Some Shock Induced Separation Figure 47 Further increases in forward speed will now result in the characteristic normal shock wave forming ahead of the aerofoil. This continuous wave, known as a Bow wave, will move towards and subsequently attach itself, to the leading edge of the wing. Once attached, all airflow over the wing will be supersonic and many of the unwanted transonic effects are eliminated.

The Bow Wave is starting to Form Figure 48


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As can be seen in figure 49, the transonic region has a great affect on the lift and drag. Both values rise until Mach 0.81, when shock induced separation drastically reduces the coefficient of lift. As speed approaches Mach 0.99, a bow wave is forming and airflow over the wing is slowed to subsonic speeds, resulting in an increase in lift coefficient and a reduction of drag.

Lift / Drag Comparison at 2º Angle of Attack Figure 49

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1.2.5 AERODYNAMIC HEATING

One of the biggest problems of sustained supersonic flight is aerodynamic heating of the aircraft structure. An extreme example of aerodynamic heating might be a „shooting star‟, when its material overheats to the point of destruction, from the heat generated by friction heating with the earth's atmosphere. In the commercial world, Concorde was probably the only airliner where aerodynamic heating presents a significant problem. When the aircraft was flown at Mach 2, the friction of the air passing around the aircraft heats the skin considerably even at altitudes in excess of 17,000 metres. The point of maximum heating is on the nose where the rise in temperature could reach 175 0C. As a precaution, a probe on the nose of the aircraft monitors the temperature during flight. When a reading of 1270C is reached, the flight deck is directed to reduce the speed to about Mach 1.8, to bring the temperature back within limits. Concorde used conventional aluminium alloys in its construction. If future aircraft were required to travel within the atmosphere at even higher Mach numbers, other materials such as titanium alloy or stainless steel would need to be considered.

Concord Skin Temperature Figure 50


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1.2.6 AREA RULE

Area rule is an aerodynamic technique used in the design of high-speed aircraft. If drag is to be kept to a minimum at transonic speeds, aircraft must be slim, smooth and streamlined. In general terms it means that the wings, fuselage, empennage and other appendages have to be considered together when working out the total streamlining. This is necessary so that the cross-sectional area of successive „slices‟ of the aircraft from nose to tail, conform to those of a simple body of streamline shape. Area rule is defined as: “For the minimum drag at the connections, (wing/fuselage), the variation of the aircraft‟s total cross-sectional area along its length, should approximate that of an ideal shape having minimum wave drag”. Without area rule, the greatest frontal cross-sectional area of the fuselage would occur where the wings are attached to the fuselage. Therefore, one method of achieving area rule in this situation is to reduce the cross-sectional area of the fuselage, thereby cancelling out the increase caused by the wings. Alternatively, the fuselage cross-section could be increased with the use of enlarged sections behind and in front of the wings to eliminate sudden changes in the cross-sectional area and achieve the same result.

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1.2.7 FACTORS AFFECTING AIRFLOW IN ENGINE INTAKES OF HIGH SPEED AIRCRAFT

Engine intakes on aircraft that operate in the subsonic flight range only can be of almost any form. The main criteria are that the airflow reaching the compressor stage of the engine during cruise ideally does not exceed Mach 0.5. This is normally achieved by the careful design of the intake ducts. Obviously, if the aircraft never exceeds Mach 0.5, a parallel intake duct could be employed, but if the aircraft is to cruise at airspeeds in excess of this, yet below Mach 1, a divergent duct must be utilised to slow the airflow at the compressor down to Mach 0.5. If the aircraft is designed to cruise above Mach 1, the air entering the intakes will be supersonic and will behave in accordance with the rules of supersonic flow. In this case a convergent duct would be necessary to slow down the airflow to the compressor. However the aircraft must fly through the transonic range in order to reach supersonic speed so both types of duct will be necessary. One way to overcome the problem is to have moveable doors that change the intake duct shape from divergent to convergent cross-section as the aircraft passes through Mach 1. See figure 52. This technique can be found on the intakes of Concorde. Other methods to control airflow reaching the compressor is to make use of the fact that air passing through a shock wave slows down to a lower speed. This type of intake design is usually characterised by the „bullet fairing‟, which on some aircraft can translate in and out of the intake to reposition the shock wave during low or high supersonic flight speeds. See Figure 53


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Intake Moveable doors Figure 52

„Bullet Fairing‟ Intake Figure 53

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1.2.8 EFFECTS OF SWEEPBACK ON CRITICAL MACH NUMBER

In order to fly at high speed in the transonic range without encountering the problems caused by the production of shock waves, the Critical Mach number needs to be as high as possible. As has already been shown, one way is to have as thin a wing as possible. This of course is an acceptable solution in theory, but in practice there will be structural integrity problems, such as wing loading, strength and flexibility. Another way of raising the Critical Mach number without the structural limitations is by the use of swept wings. Sweepback not only delays the production of the shock wave, but reduces the severity of the shock stall should it occur. The theory behind this is that it is only the component of velocity over the wing chord that is responsible for the pressure distribution and so for causing the shock wave to develop. The other velocity component that travels spanwise causes only frictional drag and has no effect on shock wave production. This theory is borne out by the fact that when it does appear, the shock wave lies parallel to the span of the wing. Therefore only that part of the velocity perpendicular to the shock wave, i.e. across the chord, is reduced by the shock wave to subsonic speeds. The greater the sweepback, the smaller will be the component of velocity affected, resulting in a higher Critical Mach number and a reduction in drag at all transonic speeds. Additionally sweepback results in a thinner mean aerodynamic chord, which raises the Critical Mach number even more.

Effects of Sweepback Figure 54 Issue 1 – 04 Sept 2001

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MODULE 11.02 AIRFRAME STRUCTURES CONTENTS

2

AIRFRAME STRUCTURES – GENERAL CONCEPTS................ 2-1 2.1 AIRWORTHINESS REQUIREMENTS FOR STRUCTURAL STRENGTH ..... 2-1 2.1.1 STRUCTURAL CLASSIFICATION ...................................................... 2-1 2.1.2 Primary structure ........................................................... 2-2 2.1.3 Secondary Structure ..................................................... 2-4 2.1.4 Tertiary Structure .......................................................... 2-4 2.2 FAIL SAFE, SAFE LIFE AND DAMAGE TOLERANT CONCEPTS ............ 2-4 2.2.1 Fail Safe........................................................................ 2-4 2.2.2 Safe Life........................................................................ 2-4 2.2.3 Damage Tolerance........................................................ 2-5 2.3 ZONAL AND STATION IDENTIFICATION SYSTEM................................ 2-7 2.3.1 Zonal System ................................................................ 2-7 2.3.2 Station Identification System ......................................... 2-8 2.4 LOADS FOUND W ITHIN THE STRUCTURE – STRESS AND STRAIN ...... 2-9 2.4.1 Compression ................................................................. 2-10 2.4.2 Tension ......................................................................... 2-10 2.4.3 Bending......................................................................... 2-11 2.4.4 Torsion .......................................................................... 2-12 2.4.5 Shear ............................................................................ 2-12 2.4.6 Hoop Stress .................................................................. 2-13 2.4.7 Metal Fatigue ................................................................ 2-13 2.5 DRAINAGE AND VENTILATION PROVISIONS ..................................... 2-16 2.5.1 External Drains ............................................................. 2-16 2.5.2 Internal Drains............................................................... 2-18 2.5.3 Ventilation ..................................................................... 2-18 2.6 LIGHTNING STRIKE PROVISION ...................................................... 2-19 2.7 CONSTRUCTION METHODS ............................................................ 2-20 2.7.1 Stressed Skin Fuselage ................................................ 2-20 2.6.1 Frames and Formers..................................................... 2-21 2.6.2 Bulkheads ..................................................................... 2-21 2.6.3 Longerons and Stringers ............................................... 2-22 2.6.4 Doublers and Reinforcement ......................................... 2-23 2.6.5 Struts and Ties .............................................................. 2-23 2.6.6 Beams and Floor Structures .......................................... 2-24 2.6.7 Methods of Skinning...................................................... 2-24 2.6.8 Anti-Corrosive Protection .............................................. 2-26 2.6.9 Construction Methods – Wing ....................................... 2-27 2.6.10 Construction Methods – Empennage ............................ 2-28 2.6.11 Construction Methods – Engine Attachments ................ 2-29 2.6.12 Structural Assembly Techniques ................................... 2-31 2.6.13 Solid Shank Rivets ........................................................ 2-31 2.6.14 Special and Blind Fasteners. ......................................... 2-33 2.6.15 Bolts and Nuts............................................................... 2-38 2.6.16 Adhesive Bonded Structures ......................................... 2-43 2.6.17 Methods of Surface Protection ...................................... 2-45 2.6.18 Exterior Finish Maintenance .......................................... 2-47

Issue 1 – Module 11.02 21 Dec 2001

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MODULE 11.02 AIRFRAME STRUCTURES

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AIRFRAME STRUCTURES – GENERAL CONCEPTS

2.1 AIRWORTHINESS REQUIREMENTS FOR STRUCTURAL STRENGTH Airworthiness requirements are necessary with respect to aircraft structures, because established standards of strength, control, maintainability, etc. will ensure that all aircraft will be constructed to the safest possible standard. Requirements for aircraft above 5700kg MTWA (maximum total weight authorised) are listed in Joint Airworthiness Requirement 25 (EASA-25) and for aircraft below 5700kg MTWA, in EASA-23. These publications cover not only the basic requirements, like maximum and minimum 'g' loading, but a vast range of other requirements with respect to the structure such as: 

Control Loads



Door Operation



Effect of Tabs



Factor of Safety



Fatigue



High Lift Devices



Stability & Stalling



Ventilation



Weights

The list is all-embracing and provides a useful means of searching for specific structural details.

2.1.1 STRUCTURAL CLASSIFICATION

For the purpose of assessing damage and the type of repairs to be carried out, the structure of all aircraft is divided into three significant categories:

Primary structure



Secondary structure



Tertiary structure

Diagrams are prepared by each manufacturer to denote how the various structural members fall into these three categories.


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STRUCTURES

In the manuals of older aircraft the use of colour may be found to identify the three categories. Primary Structure is shown in Red, Secondary in Yellow and Tertiary in Green. Note: This system has been discontinued for many years, but with some aircraft having a life of 30 or more years and still being operated, it may still be possible to find the old system in use.

2.1.2 PRIMARY STRUCTURE

This structure includes all portions of aircraft, the failure of which in flight or on the ground, would be likely to cause: 

Catastrophic structural collapse



Inability to operate a service



Injury to occupants



Loss of control



Unintentional operation of a service



Power unit failure



Examples of some types of primary structure are as follows: 

Engine Mountings



Fuselage Frames



Main Floor members



Main Spars

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Primary Structure :Wing Spars Figure 2


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2.1.3 SECONDARY STRUCTURE

This structure includes all portions of the aircraft which would normally be regarded as primary structure, but which unavoidably have such a reserve of strength over design requirements that appreciable weakening may be permitted, without risk of failure. It also includes structure which, if damaged, would not impair the safety of the aircraft as described earlier. Examples of secondary structure include: 

Ribs and parts of skin in the wings.



Skin and stringers in the fuselage

2.1.4 TERTIARY STRUCTURE

This type of structure includes all portions of the structure in which the stresses are low, but which, for various reasons, cannot be omitted from the aircraft. Typical examples include fairings, fillets and brackets which support items in the fuselage and adjacent areas.

2.2 FAIL SAFE, SAFE LIFE AND DAMAGE TOLERANT CONCEPTS

2.2.1 FAIL SAFE

A fail safe structure is one which retains, after initiation of a fracture or crack, sufficient strength for the operation of the aircraft with an acceptable standard of safety, until such failure is detected on a normal scheduled inspection. This is achieved by part and full scale airframe testing and fatigue analysis by usually by the aircraft manufacturer and by subsequent in-service experience.

2.2.2 SAFE LIFE

Safe life structure and components are granted a period of time during which it is considered, that failure is extremely unlikely. When deciding its duration, the effects of wear, fatigue and corrosion must be considered. For example, if tests show that fatigue will cause a failure in 12,000 flying hours, then one sixth of this might be quoted as the safe life.(2000 hours then scrapped) If wear or corrosion prove to be the likely cause of failure before 12,000 hours, then one of these will be the deciding factor. The safe life time period may be expressed in flying hours, elapsed time, number of flights or number of applications of load, ie; pressurisation cycles.

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2.2.3 DAMAGE TOLERANCE

The fail safe method has proven to be somewhat unreliable following some accidents that proved that the concept was not 100% guaranteed. It was also a severe limitation that the addition of extra structural members to protect the integrity of the structure considerably increased the weight of the aircraft.. The damage tolerant concept, has eliminated much of the extra weight, by distributing the loads on a particular structure over a larger area. This requires an evaluation of the structure, to provide multiple load paths to carry the loading. The main advantage is that even with a crack present, the structure will retain its integrity and that during scheduled maintenance programmes, the crack will be found before it can become critical. For example, a wing attachment to the fuselage, which in the past would have been designed with one or two large pintle bolts, will now have a larger number of smaller bolts in the fitting. The single or dual bolt attachment had to be heavily reinforced to take the wing loading, adding more weight, whereas the multiple load path can be constructed in a lighter manner, whilst still maintaining its strength.

Single Pin Attachment Figure 3


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Multiple Pin Attachment Figure 4

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2.3 ZONAL AND STATION IDENTIFICATION SYSTEM

2.3.1 ZONAL SYSTEM

During many different maintenance operations including component changes, structural repairs and trouble shooting, it is necessary to indicate to the engineer where, within the structure, the correct location is to be found for the work to be carried out. When attempting to establish a specific location or identifying components, some manufacturers make use of two systems, a zonal system and a frame/station method. The zonal system divides the airframe into a number of zones, (usually less than 10), to give engineers and others a rough idea of where they need to look. The zonal system may also be used in component labelling and work card area identification. In the illustration below, an engineer might have for example a work card numbered 500376, indicating it was Job 376 located on the left wing (Zone 500).

Zonal Identification Figure 5 Issue 1 – Module 11.02 21 Dec 2001

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2.3.2 STATION IDENTIFICATION SYSTEM

Most manufacturers use a system of station marking where, for example, the aircraft nose is designated Station 0 and other station designations are located at measured distances aft of this point. Component and other locations within the wings, tailplane, fin and nacelles are established from separate dedicated station’s zero. Fuselage Locations A particular fuselage station (or frame) would be identified, for example, as Station 5050. This means that if the metric system of measurement is employed, the frame is located at 5.05 metres (5050mm) aft of station zero.

Frame Stations Figure 6 Lateral Locations To locate structures to the right or left of the aircraft, many manufacturers consider the fuselage centre line as a station zero. With such a system, the wing or tailplane ribs could be identified as being a particular number of millimetres (or inches) to the right or the left of the centre line. Vertical Locations These are usually measured above or below a ‘water line’, which is a predetermined reference line passing along the side of the fuselage, usually, somewhere between the floor level and the window line.

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2.4 LOADS FOUND WITHIN THE STRUCTURE – STRESS AND STRAIN Aircraft structural members are designed to carry a load or to resist stress and a single member may be subjected to a combination of stresses during flight. When an external force acts on a body, it is opposed by a force within the body. This force is called Stress. If the body is distorted by the stress, it is said to be subject to Strain. Stress and strain can be defined as follows: Stress is load or force per unit area acting on a body. Stress = Load or Force Cross Sectional Area Strain is the distortion per unit length of a body.

Strain = Distortion Original Length

There are five major stresses and all will be found somewhere within an aircraft structure. In the design stage, the stresses will have been assessed by the designer and the structure made strong enough to carry them adequately. Furthermore, a reserve of strength will also have been included for safety. The five types of stress are: 1. Compression 2. Tension 3. Bending (a combination of compression and tension) 4. Twisting/Torsion 5. Shear


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2.4.1 COMPRESSION

Compression is regarded as a primary stress and is the resistance to any external force which tends to push the body together. Compressive stresses applied to rivets for example, expand the shank as they are driven in, completely filling the hole and forming the head to hold sheet metal skins together.

Compression Figure 7 2.4.2 TENSION

Tension is the primary stress that tends to pull an object apart. A flexible steel cable used in flying control systems is an excellent example of a component designed to withstand tension loads only. It is easily bent, has little opposition to compression, torsion or shear loads, but has an exceptional strength/weight ratio when subjected to a purely tension load.

Tension Figure 8 Page 2-10


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2.4.3 BENDING

Bending, when applied to a beam, tends to try to pull one side apart while at the same time squeezing the other side together. When a person stands on a diving board, the top of the board is under tension while the bottom is under compression. Wing spars of cantilever wings are subject to bending stresses. In flight, the top of the spar is being compressed and the bottom is under tension while on the ground, the reverse occurs, the top is in tension and the bottom is under compression. If the wing is supported, the strut will be in tension in flight and in compression on the ground.

Bending Figure 9


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2.4.4 TORSION

A torsional stress is one that is put into a material when it is twisted. When we twist a structural member, a tensile stress acts diagonally across the member and a compressive stress acts at right angles to the tension. A good example is a crankshaft of an aircraft piston engine which is under a torsional load when the engine is driving the propeller.

Torsion Figure 10 2.4.5 SHEAR

A shear stress is one that resists the tendency to slice a body apart. For example a clevis bolt in a flying control system is designed to take shear loads only. It is normally a high strength steel bolt with a thin head and a fat shank. These bolts secure the flexible steel cables to the control surfaces and allow the cable to move with the control surface without bending. The airload on the control surface attempts to slice the bolt apart or shear it.

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2.4.6 HOOP STRESS

An aircraft which has its fuselage pressurised inside to allow the carriage of passengers at altitude, will have other stresses acting on the fuselage skin. The circumferential load about the fuselage is known as hoop stress and resisted by the fuselage frames and tension in the so called stressed skin. The longitudinal (axial) load along the fuselage is also resisted by tension in the skin and by the longerons and stringers.

Hoop stress Figure 12 2.4.7 METAL FATIGUE

The phenomenon of metal fatigue has long been known, but has become of greater concern in recent years with aircraft which remain in service long after their original expected fatigue life has expired. It is relatively easy to design a structure to withstand a steady load, but aircraft are subjected to widely varying loads in flight and many components experience load reversals, an example being the wings, where the aerodynamic forces during flight manoeuvres cause tension and compression loads to alternate continually. Unfortunately, any metal part subjected to a wide variation or reversal of even a relatively small load is gradually and progressively weakened. The subject was vividly highlighted in 1954, with another type of load reversal, that of pressurisation cycles of the passenger cabin. which resulted in a number of disastrous accidents with the De-Havilland Comet airliner. Small fatigue cracks in the fuselage skin accumulated around the corners of the square shaped windows and hatches and led to a fatal explosive decompression of the cabin. Following the incidents the most extensive research to this hitherto unwarranted menace was undertaken, and led to fatigue loading being included into future design considerations. Metal fatigue refers to the loss of strength, or resistance to load, experienced by a component or structure as the number of load cycles or load reversals increases. Issue 1 – Module 11.02 21 Dec 2001

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Load reversals refer to a material being continually loaded and unloaded and as long as the elastic limit is not exceeded, the material should be unaffected and return to its original state. In reality, however, the load application may result in minute, seemingly inconsequential cracks, which, as the cycles continue, get larger and join up with other, newer cracks. Eventually, after many cycles, the cumulative effect will be such that the strength of the metal will be compromised and could result in catastrophic failure. The fatigue strength of a metal can be found by experimentation on full scale fatigue rigs, which can be subjected to a programme of load reversals, 24 hours a day, 365 days a year, to accumulate information and a fatigue life, years ahead of the oldest aircraft of the particular type in the fleet. How the in-service aircraft subsequently consumes this fatigue index, depends on its operating theatre. For example, the number of times the pressurisation cycles are applied to aircraft on long or short haul flights, steep or conventional take off and landing etc., are taken into account to calculate fatigue life consumed. Stress amplitude can be plotted against endurance for one particular value of mean stress, the so-called ‘S/N Curve. Using a chart such as this, it can be determined at what point, in cycles, the metal has reached its minimum acceptable strength. This will be the ultimate fatigue life and is normally allotted a fatigue index of 100.

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Even when the fatigue index of 100 is eventually reached on each individual aircraft, the designers can extend it beyond 100, by examining, as previously mentioned, how the fatigue was consumed and recommending specific structural inspection and possibly strengthening or replacement of fittings and components. Fatigue is a natural phenomenon and cannot be prevented. The ability to correctly predict its effects and take the necessary action is the problem faced by the aircraft design and maintenance personnel. Different metals have different fatigue characteristics and the way parts are designed, also affects their fatigue life. Fastener holes, sharp changes in thickness and small seemingly insignificant cracks for example, can directly affect the fatigue life of a part. Fatigue cracking can also accelerate the onset of corrosion, by exposing unprotected metal to the elements. The crack growth and the consequential increase in corrosion, can cause serious structural problems over a relatively short period. With the ageing of the airliner fleet, a number of extra inspections, including non-destructive testing and structural sampling techniques have been introduced. The maintenance technician must carefully monitor the aircraft structure, paying particular attention to the integrity of surface finish and general corrosion.


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2.5 DRAINAGE AND VENTILATION PROVISIONS

Drainage The aircraft structure requires many different types of drain holes and paths to prevent water and other fluids such as fuel, hydraulic oil etc., from collecting within the structure. These could become both a corrosion and fire hazard. The forms of drainage can be divided into two areas. 1. External drains 2. Internal drains

2.5.1 EXTERNAL DRAINS

These ports are located on exterior surfaces of the fuselage, wing and empennage to ensure fluids are dumped overboard. In small unpressurised aircraft and unpressurised areas of larger airliners, these drains may be permanently open. However, in pressurised aircraft, the cabin air would leak uncontrollably through the drains and so it is necessary to use drain valves to prevent loss of cabin pressure. There are a number basic types of drain valve used for this purpose. Two similar types rely upon pressurised air in the cabin to keep the valve closed. One valve has a rubber flapper seal and the other a spring loaded valve seal. Normally located on the keel of the fuselage, both are open when the aircraft is unpressurised on the ground, allowing the fluids to drain overboard. During flight, the increased air pressure in the cabin closes the valves, thus preventing any pressurisation losses. These valves are shown below, where it can also be seen that a levelling compound has been used in areas which might become fluid traps. This compound is usually a rubberised sealant which fills the cavity, bringing the level up to the lip of the drain hole.

Fuselage Drains Figure 14

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Another similar type of drain valve also uses the cabin air pressure to close off the drain path, this time by moving the plunger down to seal the drain. This valve will also be open when cabin pressure is removed.

Fuselage Drains Figure 15 Fluids from some places, such as galleys and wash basins, require more than simple drain holes. The temperature at cruising altitude can fall to -60°C and water draining overboard could freeze and cause blockage problems. The method used in these cases are drain masts, which are like small aerofoils projecting from the bottom of the aircraft skin, on the centre line, through which the water is discharged. The drain masts are heated to prevent icing and also discharge the liquids well away from the aircraft's skin.

Boeing 747 Drain Masts Figure 16 Issue 1 – Module 11.02 21 Dec 2001

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2.5.2 INTERNAL DRAINS

To enable the external drains to function as designed, means must be provided within the various locations of the airframe and powerplant installation, to ensure that all fluids are directed towards the site of the external drain points. This is achieved by using internal drain paths and drain holes. The internal structure is provided with tubes, channels, dams and drain holes, to direct the flow of fluid towards the external drain points. All structural members are designed so that they do not trap fluids by ensuring, for example, that all lightening holes and ribs face downwards, allowing fluids to run off them.

2.5.3 VENTILATION

It is essential that the internal cavities within the structure are properly vented to prevent the build up of flammable vapour from the drain lines and to allow any other moisture residue to properly evaporate. Consequently sumps, tanks and cavities will all be provided with vent pipes and in some cases, such as engine cowlings, ram air inlets and outlets are utilised to ensure all zones where fluids are contained are adequately ventilated. System Installation Provisions The installation of various systems within the airframe, require adaptations from the perfect ‘drawing-board’ design. When systems like the air conditioning and pressurisation, hydraulic, pneumatic, electrical, avionics and others are designed, there must be facilities incorporated in the plans, to provide a location for all the system components, their associated lines and cables. It must also be borne in mind that many components have to be either serviced ‘in-situ’, or will be a line replaceable unit (LRU), both of which requires easy access for the maintenance engineers. To this end, on modern aircraft, there are normally compartments allocated to each of the major systems where the majority of components will be installed. Thus, it can be possible to find dedicated Avionics bays, Hydraulic bays, Air conditioning bays, etc., all of which allow access for the easier replacement of 'black boxes' (LRU’s) and mechanical components like control units, valves, filters etc,. Older aircraft will still have components scattered throughout the airframe, with difficult access in some places through small panels, all of which will obviously make maintenance on these systems much more difficult.

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2.6 LIGHTNING STRIKE PROVISION When aircraft are flying in cloud or in close proximity to storms, there is always the risk of the aircraft being struck by lightning. Whilst this is a rare occurrence, there are many protection devices installed in the aircraft to ensure that a strike does as little damage as possible when it does happen. A lightning strike on an aircraft can have a peak current of up to 100,000 amperes, so precautions must be taken to ensure that the least damage is done to the aircraft, its systems and components as the charge passes through. Most important is the electrical bonding of all the major components of the airframe. Bonding is achieved by electrically connecting all the components of an aircraft structure together. These precautions will ensure all components are at the same electrical potential by providing a return path through the airframe, since modern aircraft utilise an earth return system. This means that current from the lightning strike cannot build up on one part of the structure and create a voltage high enough to allow it to jump to another part, that might be electrically separated, such as flying control surfaces. Note: Electrical bonding also protects equipment from the build up of static electricity, which is produced as the aircraft collects ions from the atmosphere as it passes through. Bonding cables are referred to as secondary conductors. As well as electrical bonding, dedicated lightning protection systems are employed to cater for the high current and these are usually known as primary conductors. They can be found, connecting system earth returns, as mentioned earlier, connecting power-plants to the airframe and ensuring that all major structural items, (which are often manufactured in different factories in different countries), are properly connected together after final assembly. Occupants of the aircraft are also protected from electrical shock in this way by the surrounding aircraft structure with what is referred to as a ‘Faraday Cage’.

Electrical Bonding Figure 17 Issue 1 – Module 11.02 21 Dec 2001

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CONSTRUCTION METHODS

2.7.1 STRESSED SKIN FUSELAGE

As previously described, a variety of loads act on the airframe during flight. If a proportion of these loads can be carried by the skin covering, the underlying framework can be made lighter without loss of overall strength. In early aircraft, all loads were taken by the framework and the covering of fabric, doped to pull it taught or of thin sheets of wood achieved streamlining, but contributed little or nothing to the strength of the airframe. As aircraft design evolved, the fabric and wood was replaced with aluminium alloy sheet. Because of its extra strength, a large part of the load can be borne by this skin, reducing the weight of underlying structure. This is called Stressed Skin construction and this method also provides a very smooth surface, because the skin is stiff enough not to be distorted by the airflow. With the advent of pressurised cabins the usefulness of a strong skin is evident when considering pressurisation loads. A method of construction where the skin carries all the loads without supporting structure is called pure monocoque construction. A good example of a pure monocoque construction is a chicken’s egg, since it has no internal support, only the egg shell carries the load. In practice, this construction is difficult to achieve, as the skin would have to be so thick, that the extra weight penalty incurred, would severely impair the ability to fly. However, the principle is sometimes used in the construction of composite material external fuel tanks, mainly for military aircraft and even here some internal strengthening is necessary.

Monocoque Construction Figure 18 Page 2-20



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STRUCTURES

In a stressed skin fuselage construction, about half the loads are carried by the skin and half by the supporting structure. This type of construction is called semi monocoque and its advantage is that the space within the structure is unobstructed and is used for passengers and freight.

Semi-Monocoque Construction Figure 19 2.6.1 FRAMES AND FORMERS

Frames and formers provide the basic fuselage shape, with the frames, being of more robust construction, providing strong points for attachment of other fittings such as the wings and tailplane.

2.6.2 BULKHEADS

Where extra support is required within a fuselage for mounting of components such as wings and landing gear, bulkheads are to transfer the loads to the fuselage structure without producing stress raising points. Bulkheads can be either a complete or a partial circular frame, which usually reinforces a fuselage frame. Other examples are solid pressurisation bulkheads which are normally found at the front of the fuselage ahead of the flight deck and at the rear of the pressure cabin, or an engine firewall on the nacelles. Issue 1 – Module 11.02 21 Dec 2001

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2.6.3 LONGERONS AND STRINGERS

Longerons are used in fuselage construction, where either an aperture such as a door or window requires greater support, or where a number of structural high load points such as floors, landing gear attachments, etc. need to be interconnected. They are usually of much heavier construction than stringers and can be solid extrusions or fabricated multiple part construction. Stringers provide longitudinal shape and support to the fuselage skin. They are also the spanwise members of the mainplanes, vertical and horizontal stabilisers and flying control surfaces. Often stringers are attached to frames with fillets or gussets.

Longerons and Stringers Figure 20

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2.6.4 DOUBLERS AND REINFORCEMENT

Where the skin requires extra strengthening, at the junction of plates or around small apertures, a second layer of skin is attached over the original to reinforce it. This extra plate is known as a doubler or a doubler plate. Where loads are concentrated within the structure, it can be strengthened at these places by either making the material thicker, or by the addition of a number of layers of similar material. The actual amount of reinforcement being dictated by the amount of stress carried in each area.

Doubler Plate Figure 21 2.6.5 STRUTS AND TIES

Any structural item that is designed solely to take a compressive load is called a strut. Whereas an item that only takes a tensile load is called a tie. They can be found throughout a modern aircraft structure, although an ideal example would be a high performance biplane. In this type of aircraft often used for aerobatics, the struts which separate the pairs of wings, in compression and the interconnecting flying wires, in tension, take all the loads produced by the wing.

Struts and Wires Figure 22 Issue 1 – Module 11.02 21 Dec 2001

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2.6.6 BEAMS AND FLOOR STRUCTURES

Beams are often used laterally and longitudinally along the fuselage to support the flight deck and passenger cabin floors. Additionally they provide strong point attachments for the crew and passenger seats and as such, constitute primary structure. Modern cabin flooring is usually made up from a number of removable composite honeycomb core panels, examples of which are shown below, whereas the flight deck is often made from metal panels supported on beams.

Floor Structures Figure 23 2.6.7 METHODS OF SKINNING

Skins for light aircraft are usually simple, thin sheets of aluminium alloy, wrapped around and riveted to the internal structure. Larger aircraft, developed since the 1950’s have their skins manufactured from heavier material with the additional use of even thicker sections in certain places where more strength is required. As the aircraft designs became more complex, the excess weight of thicker skins in places where they are not necessarily required, became too big a penalty. To overcome this problem, the skins were rolled individually to produce a variety of differing thickness across each sheet, to cater for variations in stress.

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The latest methods are to machine or mill each skin panel individually from a solid billet, to include all stringers and risers and to provide a varying thickness all over the sheet. In this way, the skin panel is exactly the right thickness at each location, with no excess material and hence no extra weight. This method results in what is termed milled skin or machined skin. Milled wing skins give maximum strength and rigidity with minimum weight. Panels containing areas of different thickness can also be produced from a chemical etching process where areas which have been treated, will be removed to about half their thickness by the chemical etch. The nature of the etching process ensures that no ‘stress raisers’ are introduced into the material. So called ‘waffle plates’ can be produced in this way and are shown in Fig 24.

Skinning Methods Figure 24


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2.6.8 ANTI-CORROSIVE PROTECTION

Materials used in aircraft construction are selected primarily for their strength and tenacity. Unfortunately, many may readily suffer serious damage from corrosion unless effectively protected and the rate of corrosion attack can be extremely rapid in certain environments. One of the main considerations in the design of aircraft structure therefore, are measures for the control and prevention of corrosion. During manufacture and assembly, a range of surface treatments are applied. Materials are heat treated to refine grain structure, sacrificial coatings in the form of plating and cladding are employed, to retard the onset of corrosion. Epoxy primers, special paint finishes, wet-assembly techniques and the use of barrier sealants to prevent the ingress of dirt and moisture between component parts, all help to reduce the risk of corrosion. Additionally, drain holes, drainage paths and attention to good corrosion resistant design techniques for each component part, ensure that aircraft newly off the production line are protected as much as possible, before entering airline service. Aircraft are required to operate in widely varying, often highly corrosive environments throughout the world and despite the high standard of protective treatments applied during manufacture, corrosion will still occur. Corrosive attack may extend over an entire metal surface, may penetrate locally to form deep pits or may follow the grain boundaries within the metal. The weakening effect of corrosive attack may be aggravated by stresses in the metal and result in premature failure of the component. These stresses may be due to externally applied loads or may be internal stresses locked into the metal structure during manufacturing processes, despite the care taken to keep the risk to a minimum. Whatever the cause and type of corrosive attack, unless preventative maintenance is carried out, damage may become so severe, it could present a serious hazard to the airworthiness of the aircraft. Rectification of advanced corrosion damage is time consuming and much of the corrosion during service can be prevented or contained by simple corrosion prevention measures Corrosion seldom occurs on a clean dry aircraft especially if the protective coatings are completely in tact. Since aircraft have to operate outside throughout their lives, they are difficult to keep dry, but keeping the protective coatings free from scratches, dents and scores, ensuring drains which might allow water to accumulate are kept clear and keeping the aircraft clean and free of dirt are all within the scope of a good maintenance engineer. In addition, the engineer should clear up spills from the galleys and toilets and remove deposits from engine exhausts as these are also very corrosive if left on the skin for too long.

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2.6.9 CONSTRUCTION METHODS – WING

The basic requirement for wing construction, particularly with cantilever types is for a spanwise member of great strength, usually in the form of a spar. Conventionally, there are three general designs, monospar, two-spar or multispar. Most modern commercial airliners, have a wing comprising top and bottom skins complete with spanwise stringers, front and rear spars and a set of wing ribs running chordwise across the wing between the spars. This forms a box-like shape which is very robust and the addition of nose ribs and trailing edge fittings produce the characteristic aerofoil shape. Wing structures carry some of the heaviest loads found in aircraft structure. Fittings and joints must be carefully proportioned so they can pick up loads in a gradual and progressive manner and redistribute them to other parts of the structure in a similar manner. Special attention must be paid to minimising stress concentrations, by avoiding too rapid a change in cross section and to provide ample material to handle any concentration in stress or shock loading that cannot be avoided, such as landing loads.

Typical Wing Construction Figure 25


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2.6.10 CONSTRUCTION METHODS – EMPENNAGE

The vertical and horizontal stabilisers, elevators and rudder are constructed in a manner similar to the wings but on a smaller scale. The main structural members are the spars, with the stringers, ribs and stressed skin completing the basic design.

Typical Stabilizer Construction Figure 26

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2.6.11 CONSTRUCTION METHODS – ENGINE ATTACHMENTS

Engine mountings consist of the structure that transmits the thrust provided by either the propeller or turbojet, to the airframe. The mounts can be constructed from welded alloy steel tubing, formed sheet metal, forged alloy fittings or a combination of all three. Some typical examples are shown in Figures 27 to 29. All engine mounts are required to absorb not only the forward thrust during normal flight, but the reduced force of reverse thrust and the vibrations produced by the particular engine/propeller combination..

Fabricated Piston Engine Mounting Figure 27

Tubular Turbopropeller Mounting Figure 28 Issue 1 – Module 11.02 21 Dec 2001

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Machined Turbojet Side Mounting Figure 29

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2.6.12 STRUCTURAL ASSEMBLY TECHNIQUES

The integrity of an aircraft joint depends on the way the parts are attached together. The most common method of attachment is by the use of rivets or more sophisticated types of rivets, known as fasteners. However, where high strength is required, nuts and bolts are used whilst other structural assembly is achieved by the use of adhesive bonding techniques. Although aluminium alloy is the most common material for aircraft construction, more and more structural components and in some cases, complete aircraft, are being manufactured from composite materials like glass or carbon fibre. Riveting is generally divided into two types: (1) solid shank rivets and (2) special fasteners. The special fastener category being sub-divided further into special and blind fasteners. 2.6.13 SOLID SHANK RIVETS

The vast majority of aircraft structure is held together with solid rivets. As will be explained later, many of the more modern designs use special fasteners and some bonded construction, but the majority are still solid rivets. Head Shapes In the past there have been a large number of rivet head shapes used in aircraft, but in recent years these have been reduced and standardised to four main types: The Universal Head, sometimes known as AN70 or MS20470, is most popular and may be used to replace any protruding-head rivet. It is streamlined on top but thick enough to provide strength without protruding too much into the airflow. A Round Head rivet, AN430, is used on internal structure where the thicker head is more suitable for automatic riveting equipment. In internal locations where a flat head rivet can be driven more easily than either a round or universal head rivet, the AN442 Flat Head rivet may be used. Where a smooth skin is important, flush rivets such as AN426 or MS20426, with a 100 countersink head are used. Additionally, rivets with a different countersink angle, such as 90 and 120 degrees can be found.

Rivet Head Types Figure 30 Issue 1 – Module 11.02 21 Dec 2001

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Types of Alloy used for Solid Shank Rivets The identification marks on rivet heads serve two important functions. Firstly, the marks are used to identify the rivet alloy required for a special installation area and, secondly, the head markings are necessary when trying to identify which kind of rivets are being removed from an aircraft during disassembly or repair. The alloy identifying marks are made on rivet heads at the time they are being stamped out during manufacture. Generally, solid rivets are manufactured in five different materials:

Solid Rivet Identification Figure 31 For non-structural applications, rivets made from pure aluminium, sometimes known as 'A' rivets, may be used. A very popular rivet is the 'AD' rivet, which has copper and magnesium added to the aluminium base metal. This rivet is heat treated during manufacture to make it strong, whilst still being soft enough to be formed easily. When much more strength than the 'AD' rivets is required, there are two stronger rivets available. These are 'D' and 'DD' rivets but they must be heat treated to make them softer before they can be formed. The 'D' types are of 2017 alloy and the 'DD' types are manufactured from 2024 alloy. Both of these rivet types, after heat treatment, must be formed within a specific period of time (one hour for 'D' and ten minutes for 'DD' types) or they may be put into a refrigerator to maintain the softening effect. Once refrigerated they will remain useable for about 10 days.

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When riveting magnesium alloy sheets, there must be no copper in the rivet alloy, or dissimilar metal corrosion will set in. Therefore, a 'B' rivet, manufactured from 5056 alloy is used. This contains a large amount of magnesium with a little manganese and chromium but no copper. Dimensions Aircraft rivet dimensions are categorised by the diameter of the shank, ‘D’, and the length, ‘L’, measured from the end of the shank to the portion of the head that will be flush with the surface of the metal. This means that a countersink rivet is measured from the top of its head, whilst the remainder are measured from under the head.

Rivet Dimensioning Figure 32 Identification The complete identification of a rivet includes its head style, its material, its diameter and its length. The identification code shows the diameter as a number of 1/32ths of an inch and the length as a number of 1/16ths of an inch. For example, An MS20470AD4-4 has a universal head (MS20470), is made from alloy 2117 (AD), is 1/8" diameter (4 x 1/32”) and 1/4" long (4 x 1/16”).

2.6.14 SPECIAL AND BLIND FASTENERS.

When solid shank rivets become impractical to use, then special fasteners are used. These, you will remember, are of two types; special and blind fasteners. The term ‘Special Fasteners’ refers first to their job requirement and second to the tooling needed for the installation. In certain locations, aircraft require strength that cannot be produced by a solid shank rivet, so a special high strength fastener is used. For example, if high shear strength is required, then special High Shear rivets are used. These are usually installed with special tools and will be discussed later in this chapter.


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Blind Fasteners There are several different types of blind fasteners which can be hollow or selfsealing. They include the following types, all of which can be installed from one side of the work. 

Chobert



Avdel



Tucker/Pop



Cherry

Note: It is most important that the correct tools are always used with the types of rivets mentioned above. Chobert Rivets These are available with a snap (round) head or a countersink head and are closed by forcibly pulling a mandrel through the bore of the rivet. This closes the 'tail' and expands the rivet tightly into the hole. To seal Chobert rivets, a separate sealing pin is driven into the hollow bore of the rivet.

Chobert Rivet Figure 33

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Tucker or 'Pop' Rivets Tucker/'Pop' rivets are manufactured with either domed or countersunk heads and are supplied on individual mandrels. The rivets can be either ‘break head’ or ‘break stem’ and when closed, can be sealed or open depending upon their application. Break head rivets are rarely used due to the 'foreign object' risk from the broken off heads lying within the internal aircraft structure. Break stem rivets are be divided into two groups, short and long break mandrels. Long break types leaves the stem in place, greatly increasing the shear strength of the rivet.

Tucker ‘Pop’ Rivet Figure 34 Cherry Rivets These rivets, of American manufacture, are similar to Avdel rivets, except that the stem is positively locked in the rivet bore. During final forming, a locking collar is forced into a groove in the stem, preventing further movement. After the closing operation, the remainder of the stem is milled flush with the skin. There are many different types of Cherry rivets, two of the most popular being the Cherry Lock and the Cherry Max. The Cherry Lock, however, requires a range of closing tools for different sized rivets, whilst the Cherry Max series can all be closed with a single tool. Cherry Lock rivets are manufactured from 2017 or 5056 alloys, Monel metal or Stainless Steel, whereas Cherry Max are made from 5056 alloy, Monel or Inconel 750. They are all available with either universal or countersink heads and due to their positive locking method, can be installed in place of solid shank rivets.


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Cherry Lock Rivet Figure 35

Cherry Max Rivet Figure 36 Page 2-36


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Avdel Rivets These are similar to Chobert rivets, but each is fitted with its own stem (mandrel). The stem is pulled through the rivet body to close the rivet and at a predetermined load, breaks off proud of the manufactured head. This leaves part of the stem inside the body which seals the rivet. The excess stem is then removed by nipping it off and carefully milling it until flush with the surface of the aircraft skin. The shear strength of an Avdel rivet is greater than a Chobert rivet of equivalent material and size and similar to a solid rivet.

Avdel Rivet Figure 37 Special Fasteners These can include Hi-Shear, Avdelock, Jo-Bolts, and Rivnuts. The first three are all formed by means of a collar which is swaged into the grooves in fastener shank or expanded over the shank to form a blind head. Rivnuts are formed using a similar method to cherry locks, but with a threaded mandrel screwed into the Rivnut. The advantage of Rivnuts, (see Fig 38), is that after closing, a fixed nut is left behind which may be used for the attachment of de-icing boots, floor coverings and other non-structural parts.

Rivnuts After Installation Figure 38


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2.6.15 BOLTS AND NUTS

Bolts A bolt is designed to hold two or more parts together. It may be loaded in shear, in tension, or both. Bolts are designed to be used with nuts and have a portion of the shank that is not threaded, called the grip, whereas Machine screws and Cap screws have the entire length of the shank threaded. The dimensions required to identify a bolt are expressed in terms of the diameter of the shank and the length from the bottom of the head to the end of the bolt. The grip length should be the same as the thickness of the material being held together. This measurement can be found by reference to the applicable charts. Bolt heads are made in a variety of shapes, with hexagonal being the most common.

Bolt Terminology Figure 39 General Purpose Bolts All-purpose structural bolts used for both tension and shear loading is made under 'AN' standards from 3 to 20, the bolt diameter is specified by the AN number in 1/16"; for example: AN3 = 3/16" diameter AN11 = 11/16" diameter The range is from AN3 to AN20 which have hexagon heads, are made from alloy steel and have UNF (fine) threads. The length of the bolt is expressed as a dash number. Bolts increase in length by 1/8" and the dash number(s) will show the length.

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For example: AN3-7 = 7/8" long AN3-15 = 1 5/8" long Other markings will identify whether the bolt has a drilled shank, a drilled head for locking and indicate what material the bolt is made from. Clevis Bolts These bolts (AN21 to 36) are designed for pure shear load applications such as control cables. The slotted, domed head results in this bolt often being mistaken for a machine screw. A clevis bolt has only a short portion of the shank threaded with a small notch between the threads and the plain portion of the shank, which allows the bolt to rotate more freely in its hole. Because the length of this bolt is more critical than normal bolts, its length is given in 1/16" increments.

Clevis Bolt Identification Figure 40

Nuts All nuts used on aircraft must have some sort of locking device to prevent them from loosening and falling off. Many nuts are held in place on a bolt, by passing a split pin through a hole in the bolt shank and through slots, or castellations, in the nut. Others have some form of locking insert that grips the bolt's thread, whilst others rely on the tension of a spring-type lock-washer to hold the nut tight enough against the threads to prevent them from vibrating loose. Sometimes, nuts that are plain with no locking devices are used and prevented from coming undone, once they have been tightened, by the use of locking wire attached to an adjacent nut or to the aircraft structure.


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There are two basic types of nuts, self-locking and non self-locking. As the name implies, a self-locking nut locks onto a bolt with no external help, whilst a non selflocking nut relies on either a split pin, lock-nut, locking washer or locking wire, to stop it from undoing.

Standard Nuts Figure 41 Another type of nut in general use is the Anchor nut. These are permanently mounted on nut plates that enable inspection panels and access doors to be easily removed and installed, without access being required on the reverse side of the work. To make fitment of the panel easier when there is a large number of screws, the nuts are often mounted 'floating' on their mounts, which allows for small differences in the position of the attaching screws. Although rarely used on large commercial airliners, Tinnerman nuts are manufactured from sheet steel and are used mainly on light aircraft, for the fitting of instruments into the flight deck panels, the attachment of inspection panels, etc. Some light aircraft engine cowlings have U-type tinnerman nuts fitted over the inner edge of the cowling frame. When the retaining screws are tightened, spring action holds them tightly and safely in place.

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Examples of self-locking nuts, anchor nuts and U-type tinnerman nuts are shown in figures 42 and 43 below.

Self Locking and Anchor Nuts Figure 42

U-Type Tinnerman Nut Figure 43


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2.6.16 ADHESIVE BONDED STRUCTURES

Adhesive Bonding is the technique of joining materials using special adhesives. In the past a common type of adhesive widely used in metal to metal joints was the ‘Redux’ epoxy resin system. ‘Redux’ is the trade name for a range of adhesives produced by the Ciba-Geigy company and the epoxy bonding procedure in general, refers to a hot-melt, hot-cure adhesive, which is available in partly cured strips or sheets. Note: This type of epoxy resin is also used to provide the reinforcement for fibre composite construction and has already been covered as a separate topic in Module 6. In metal to metal bonding, the sheets of partly cured adhesive, which at this stage resemble strips of chewing gum, are cut to exact size. With the backing paper peeled away, they are carefully placed between each of the components being joined together and the joint securely clamped. The complete assembly, which for example might consist of a wing skin with all its stringers and ribs in place, is then loaded into an autoclave (pressure cooker) to complete the curing process. The adhesive melts and flows evenly into the narrow gaps between the component parts and cures to produce a very strong bond. In the autoclave the temperature limits are strictly controlled, (typically not above 100-150C, depending on type of adhesive used), and subjected to a constant clamping force (usually by a vacuum process), resulting in perfect bonded joints which are as strong as, or stronger than, equivalent riveted joints. For composite repairs, figure 45, a portable Autoclave process is employed. There are a number of aircraft, in which the majority of the primary metal structure is joined together entirely with adhesive bonding, with very few rivets being used. The Fokker 50/70/100 and BAe 146/RJ are good examples of aircraft employing this technique extensively. In fact British Aerospace claims that by using adhesive bonding techniques on the BAe 146/RJ airframe, over 10,000 rivets are not required. This means the weight of the rivets, the work that would be expended in closing them and the risk of subsequent in-service cracks (see Figure 44) emanating from rivet holes, are all saved on each airframe. A further important advantage of using adhesively bonded structures, is improved sealing of integral fuel tanks, eliminating the leakage problems that are typical of riveted assemblies.


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Comparison between Machined and Bonded Structure Failure Rates Figure 44

Autoclave Curing Process During Composite Repair Figure 45 Page 2-44


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2.6.17 METHODS OF SURFACE PROTECTION

As mentioned in an earlier chapter, there are many different types of surface protection added to the basic structural materials and hardware. Anodising A method of protecting aluminium based alloys from corrosion, especially when cladding is impractical, is by a process called Anodising. This is an electrolytic treatment which coats the host metal with a film of oxide. This film is hard, waterproof, air-tight and to aid in identification of some parts, will permanently accept a coloured dye. The film also acts as an insulator, so when bonding leads are to be attached to an anodised part, the surface treatment must be carefully removed before the bonding lead is attached. Finally, anodising a part also provides an excellent base for the addition of an organic finish and bonding adhesives. There are a number of different organic finishes applied to aircraft to protect the surfaces: Synthetic Enamel.- An older finish which cures by the process of oxidation It has a good surface finish, but is poor when it comes to its resistance to chemicals or wear. Acrylic Lacquer.- A popular finish in the mass production market, easy to apply and has a fairly good resistance to chemical attack and weather. Polyurethane.- One of the most durable finishes which has high resistance to wear, fading and chemicals. It also has a 'wet look'. Chromating Chromate coatings are used to protect Magnesium-based alloys, as well as zinc and its alloys. Components are immersed in a bath containing potassium bichromate and results in a yellowish coating on magnesium alloys. The coating can be restored locally with Alocrom 1200 treatment. Cladding There are two metals most commonly alloyed with aluminium, to produce high strength skin and component parts for aircraft manufacture. These are, Copper and Zinc. These alloys suffer extensively from the effects of corrosion, so a cladding technique is used as a form of corrosion protection. ‘Alclad’ as it is termed is a soft, highly corrosion-resistant, pure aluminium skin, rolled onto the face of each base alloy sheet, effectively sandwiching the alloy.


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Surface Cleaning Most aircraft will be cleaned before starting on large inspections, but it is common sense to keep an aircraft clean all of the time. Dirt can cover up cracked or damaged components as well as trap moisture and solvents which can lead to corrosion. Note: Materials mentioned in this chapter are only used as an example, each aircraft type will have a list of suitable and prohibited materials in its maintenance manuals (AMM). Exterior Cleaning Exterior cleaning is an important facet of corrosion control, but there are a number of points which must first be protected from cleaning materials and high pressure water sprays. The pitot tubes and static vents must be properly blanked off to prevent water ingress and the wheels, tyres and brake assemblies need to be covered to keep them free of aggressive cleaning agents. Only cleaning agents and chemicals recommended by the manufacturer are to used. for the job in hand or the risk of serious contamination may result. One of the unseen effects of using non-approved cleaning agents is hydrogen embrittlement. This is caused by hydrogen from the agent being absorbed into the metal, causing minute cracks and will lead to stress corrosion failure. Aircraft should ideally be washed on a proper platform with suitable drains. It is better if the outside air temperature is not too high, so the cleaning agent does not evaporate. Typically, a mix of water and an emulsion-type cleaner, to a ratio of between 3:1 and 5:1 is applied, allowed to soak for a few minutes and then rinsed off with a high pressure stream of water. Engine cowlings and wheel well areas usually have grease, oil or brake dust deposits that require special treatment. These require stronger mixtures ratios and scrubbing with a soft bristle brush to loosen the dirt before rinsing off with a high pressure water jet. It must be borne in mind however, that oil and grease could be accidentally removed from places where they are meant to be, for example in wheel bearings etc. These will often require re-lubrication after washing has been completed. Exhaust residue from both piston and jet engines is very corrosive and must be removed on a regular basis. These deposits usually require a special proprietary solvent to mix with the water. Sometimes a simple emulsified mix of kerosene and water may be approved. Dry-cleaning solvent or naptha is sometimes used for oil and grease removal. Some naptha compounds are harmless to rubber or acrylic items, whilst others will attack these same materials, so only approved specifications are to be used.

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2.6.18 EXTERIOR FINISH MAINTENANCE

All materials used on the exterior of an aircraft must be approved by the manufacturer of that aircraft to ensure no abrasives or solvents are applied where they can do damage. Non-Metallic Cleaning Non-metallic components sometimes require different cleaning techniques from metal parts. For example, the slightest amount of dust on plastic or acrylic panels will scratch and severely reduce the optical quality if rubbed with a dry cloth. This can also build up a static charge and attract more dust so the correct procedure in this situation is to wash down, rinse with water without rubbing with a cloth. Oil and hydraulic fluid also attack rubber components such as tyres, so any spillages must be cleaned up immediately. Neoprene rubber leading-edge de-icer boots and composite structures are other examples of parts that need special cleaning procedures, all of which will be detailed in the AMM. Engine Cleaning Apart from external cleaning carried out on the engine cowlings, with the associated protection of electrical components; gas turbine engines are regularly washed internally to remove the deposits of dust, sand and salt, that tend to accumulate on internal parts of the engine. This coating if not removed, can have a serious effect on the engine's performance. Indeed, the output of the engine could fall below the manufacturers minimum figures, resulting in an unscheduled and expensive engine change Alignment and Symmetry Aircraft can have abnormal occurrences during their life, when for example, a very heavy landing could occur, some accidental external damage or the need to replace a major component, etc. All of these instances will require special checks to be carried out to guarantee that the aircraft is perfectly symmetrical and aligned before its next flight. The checks consist of measuring very accurately from a number of datum points on the airframe, such as from wing tips, the nose, the horizontal stabiliser and the top of the vertical stabiliser. The checks vary, depending on the aircraft manufacturers requirements, but all ensure that measurements taken on the lefthand side of the aircraft are within a minimum tolerance of the measurements from the right-hand side. These checks are usually taken with the aircraft on jacks and in the rigging position, ie: a nominally level ‘in flight’ attitude. Issue 1 – Module 11.02 21 Dec 2001

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On light aircraft, these measurements are usually taken using a surveyors tape measure. (It is a check of comparison, not of outright measurement). As the aircraft get larger, optical theodolite style methods are used. These can be a microscopic level with the use of sighting rods or even a laser ranging alignment device. Deeper checks that are carried out after any of the above mentioned situations, as well as on a routine basis, include checks on the wing, tail and control surfaces to ensure that they are set at the correct angles. These checks are usually known as 'rigging checks' and are carried out using purpose built levelling boards and an accurate measuring device known as a Clinometer.

Rigging Checks - Older Aircraft Figure 46

Symmetry Checks – Modern Aircraft Figure 47

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MODULE 11.03 AEROPLANE STRUCTURES

Contents 3

AIRFRAME STRUCTURES - AEROPLANES .............................. 3-3 3.1

3.2

3.3 3.4 3.5

FUSELAGE ................................................................................... 3-3 3.1.1 Truss Fuselage Construction ........................................ 3-3 3.1.2 Truss Fuselage - Warren Truss ..................................... 3-3 3.1.3 Stressed Skin Structure................................................. 3-4 3.1.4 Pressurised Structure .................................................... 3-5 3.1.5 Attachments .................................................................. 3-6 3.1.6 Passengers and Cargo ................................................. 3-9 3.1.7 Doors ............................................................................ 3-10 3.1.8 Windows and Windscreens ........................................... 3-12 WINGS ......................................................................................... 3-14 3.2.1 Construction .................................................................. 3-14 3.2.2 Fuel Storage ................................................................. 3-16 3.2.3 Landing Gear ................................................................ 3-18 3.2.4 Pylons ........................................................................... 3-19 3.2.5 Control Surface and High Lift/Drag Attachments ........... 3-20 STABILISERS ................................................................................ 3-21 FLIGHT CONTROL SURFACES ........................................................ 3-22 NACELLES AND PYLONS ................................................................ 3-23


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AIRFRAME STRUCTURES - AEROPLANES

3.1 FUSELAGE The fuselage of a light aircraft is the body of the aircraft, to which the wings, tail, landing gear and engines may be attached. Larger aircraft can have their main landing gear attached to the wings and, on multiple engined aircraft, a number of the power-plants can be wing mounted also. The loads produced either on the ground or in flight, will at some time, have to pass through the fuselage. In order to absorb these tremendous loads imposed upon the structure, the fuselage must have maximum strength, but this must be combined with the other constraint, that of minimum weight. There are two types of construction found in the majority of modern aircraft fuselage design, the truss and the stressed skin type.

3.1.1 TRUSS FUSELAGE CONSTRUCTION

By definition, a truss is a form of construction in which a number of members (or struts), are joined to form a rigid structure normally covered with non-load carrying material such as cloth, fabric or thin sheets of wood. Very early aircraft used a method of construction referred to as a Pratt Truss, where struts were held in compression, and wires, which ran diagonally between the struts, were in tension.

Truss Fuselage – The Pratt Truss Figure 1 3.1.2 TRUSS FUSELAGE - WARREN TRUSS

When fuselages were subsequently made from welded tubes, the Warren Truss became popular. In this arrangement, shown overleaf, the longerons are separated by diagonal members which carry both compressive and tensile loads.


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Warren Truss Figure 2 3.1.3 STRESSED SKIN STRUCTURE

The neccessity of having to build a non-load-carrying covering over a structural truss led to designers to develop the stressed skin form of construction. In this method, a proportion of the load is carried by the outside skin, which can be also be formed into a much smoother and more efficient shape. The commonest form of a stressed skin structure is a chicken egg (pure monocoque). The seemingly fragile shell can resist high loads, as long as they are applied in a proper direction. Pure-Monocoque Structure This form of stressed skin construction is rarely seen in its purest form, because it is normal to add some form of light internal structure to help support the skin. However, there are some aircraft (normally gliders and sailplanes) made from glass reinforced plastic (GRP), which are constructed as a pure monocoque structure. In this design, the GRP skin is quite thick, often with a core of some other lightweight material such as balsa wood or composite honeycomb, so there is no need for any internal, supporting structure. Semi-Monocoque Structure This form of construction has a skin carrying a large amount of the loads, but with an internal structure of frames and stringers to keep the skin to its correct shape, where it can best carry the loads. Some have longerons which are more substantial than stringers and carry most of the longitudinal structural loads, with the frames carrying the radial loads.

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3.1.4 PRESSURISED STRUCTURE

High altitude flight places the occupants in a hostile environment in which life cannot be sustained without oxygen. To avoid the need to wear oxygen masks, the pressure in the cabin is raised higher than it is outside, which provides sufficient oxygen in the air for the passengers to breathe normally. In the 1950’s, piston-engined aircraft, had a pressure differential across the cabin wall about two pounds per square inch (psi) maximum. Modern aircraft cabins can sustain a pressure differential between 8 and 10 psi, so there must not be any part of the structure containing 'stress raisers' which would concentrate stress to an unacceptable level. Much of the structure of modern aircraft has been built to the 'fail safe' philosophy, in which the structure is built with multiple load paths for the major stresses to pass through, to cater for the unlikely failure of a single structural item. Pressurisation Sealing All joints in the structure, as well as openings such as doors, panels, emergency exits, etc. must be completely airtight during flight, to prevent the cabin pressure leaking below its required level. Joints are constructed with an interface of sealing compound, whereas windows and doors employ pre-formed rubber seals around their edges. The points where control tubes and cables pass in and out of the pressure hull, utilise some form of flexible bellows which are leak proof but move with the controls.

Pressurisation Sealing Figure 3


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3.1.5 ATTACHMENTS

The fuselage can, as mentioned earlier, carry most of the major loads, both on the ground and in flight. To this end, most of the other airframe components such as the wing, stabilisers, pylon and undercarriage, can be fitted to the fuselage. The wings can be mounted above or below the passenger compartment. As already mentioned, wings are usually attached to the fuselage with multiple attachments, although light aircraft may still have wings attached with as few as two bolts.

Early High Stress Attachment Figure 4

Multiple Fastener Wing Attachment Figure 5 Page 3-6


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The horizontal and vertical stabilisers can be fitted to the fuselage in numerous different ways. When the horizontal stabiliser is fitted part-way up or on the top of the vertical stabiliser, there will be only one strong attachment point. Otherwise, there will be separate attachments for the fin and for the left and right tailplane sections. Where a moving horizontal stabiliser is employed, the attachment will consist of left and right rear pivot fittings and a single forward attachment to a trim actuator. On rare occasions, the rear fuselage is manufactured, together with the stabilisers, as one integral unit. Because the loads generated by the empennage, it is usual to find that the rear fuselage structure has stronger frames around the stabiliser attachment points. These frames transmit the loads along the fuselage and away from the tail. The same technique is used when the engines are attached to wing or to rear fuselage mounted pylons The Fokker 70/100, for example, has oblique frames to connect the vertical stabiliser to the top mounted tailplane and to the fuselage, plus two heavy frames to transmit all the engine thrust loads into the fuselage.

Strengthened Frames Figure 6


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As previously mentioned, the landing gear can be attached either to the fuselage, the wings, or within wing mounted engine nacelles. Because of the need for cabin space, fuselage mounted landing gear on passenger and freight-carrying aircraft, often have the main landing gears mounted in fairings or nacelles beneath the fuselage as in the ATR-72, detailed below.

Faired ATR 72 landing Gear Figure 7 The landing gear, as for the other attachments, is mounted on to strong fuselage frames which in this case, are also used to mount the wings, attached above the fuselage. The loads that these frames carry, both in flight and on the ground, are transmitted into the fuselage by means of longitudinal stringers and longerons.

Fuselage Strong Points Figure 8 Page 3-8


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3.1.6 PASSENGERS AND CARGO

Aircraft that carry passengers as well as crew, all have to have seats that comply with crashworthiness regulations. These regulations dictate that the seats with a person correctly strapped in place, must be able to survive a sudden stop of over 20 times the force of gravity, (20g), without the floor mountings (to which the seat is attached) failing, or the seat itself collapsing. Although aircraft seats appear to resemble normal domestic seats, the tubular framework and floor attachment 'feet' are very strong, yet are light in weight and can be disconnected from the floor if necessary, by releasing a few quick-release fasteners. Passenger compartment floors of modern aircraft are often panels of the composite material ‘Fibrelam’, which are strong enough to carry most of the general loads created by passengers and galley equipment. The panels are themselves supported by lateral and longitudinal beams, which are primary structure, into which the panels fit. Lateral beams are attached to the lower portion of the (usually) circular fuselage frames and longitudinal beams supported by the lateral beams, are those upon which the seats are fitted.

Seat Track Fittings Figure 9 The top of each longitudinal beam is fitted with location holes which are a standard size and into which all seats are slotted. Additionally, the galleys and bulkhead partitions can also be attached to them. The frequent and equal spacing of the seat track attachment holes, allows the seats to be fitted at a variable increment, or pitch, to cater for different classes of cabin (economy or first class). On some aircraft, such as the Fokker 100, there are five longitudinal seat tracks in the cabin floor which allow a five abreast seating to be installed (3+2 or 2+3), with the off-set aisle on whichever side the customer wishes.


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Cargo Loading Systems Aircraft which are used for carrying all or part freight loads have to have the floor modified to allow the movement of pallets or containers. Usually this will consist of substantial reinforcement of the flooring with tracks, guides and rollers fitted, to allow safe and easy motorised movement up and down the freight bay. In the entrance door area, a ‘ball-mat’ is installed to allow the freight to be easily loaded, rotated and man-handled on to the rollers.

3.1.7 DOORS

This topic covers most methods of entry and exit from the fuselage, including those for passengers, crew, refreshments and meals, baggage and major maintenance access. In addition, some doors are dedicated to emergencies only and will therefore remain unused during normal operations. If the aircraft has a cabin pressurisation system, the doors have to be more substantial than for a non-pressurised type and be fitted with safety devices to prevent accidental opening. One method to prevent this happening is allow the door to open inwards so that the door 'plugs' the aperture when closed and is held in place by the cabin pressure in addition to the door frame locating bolts. Any door on pressurised aircraft that does open outwards, must have additional devices and protection mechanisms fitted to prevent accidental opening and a flight deck warning system to inform the crew if it is not properly closed and secured. Non-pressurised aircraft doors still have to be safe, with a system of handles and latches that operate in a specific order or after the application of a certain force. Doors on most aircraft are constructed in a similar way to the fuselage with an inner and an outer skin and vertical and horizontal members. The sometimes complex locking and latching mechanisms, plus the indicating and warning electrical wiring systems are all contained within this structure. Most fuselage doors are operated manually, but much larger freight/cargo doors are either electrically or hydraulically operated. Another requirement on all cabin doors, (normal exit/entry and emergency type) is the need for efficient emergency egress in the event of a mishap on the ground. They must be operable by a single handle whose operation shall be ‘rapid and obvious’. Most doors have decals and large red arrows, to clearly indicate the way in which the handles are to be rotated or moved to open the door. Dedicated emergency exits are almost always 'plug' type and, therefore, cannot be opened in flight due to the cabin pressure acting on door opening mechanism (usually an over-centre type a cam arrangement) thus preventing handle rotation.

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Door Mechanism Figure 10

Door Structure and Sealing Figure 11


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To prevent leakage of the cabin pressure, all doors have to have a substantial seal around their edges to keep the aperture between door and surrounding fuselage frame airtight. Some seals just compress and fill the space when the door is closed, others use cabin air to inflate and therefore expand the seal to achieve the same result. Fig 11 shows a typical door seal arrangement. 3.1.8 WINDOWS AND WINDSCREENS

All the transparencies on non-pressurised aircraft are normally made from acrylic or some other clear plastic material. On pressurised aircraft, flight deck windscreens have to comply with very strict bird-strike regulations and are made from a toughened sandwich of glass/plastic/glass The passenger cabin windows are manufactured from acrylic, mylar or other plastics. It must be considered that an aircraft travelling at 400 knots which collides with a bird weighing 3kg, could suffer severe structural damage, engine failure and more importantly, if the bird struck a windscreen and broke through, it could cause serious injury. Furthermore, rapid decompression of the pressure cabin would result. The regulations state that during testing, when a dead bird is fired at it from a large air gun, the screen must be able to survive the impact. Consequently, the glass/plastic/glass sandwich is fitted with a heating element between the interface of the front glass panel and the plastic core. Not only does the heater provide anti-icing protection, but helps absorb impact since it makes the plastic core more pliable and shock absorbent. The section through a typical windscreen below shows how the lamination of glass and plastic layers is arranged.

Windscreen Construction Figure 12

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Passenger cabin windows are almost always made from acrylic plastic. This saves quite a lot of weight as well as cost. For added safety, the acrylic cabin windows are actually two layers with a space in between, so that if one fails the other will carry the pressurisation loads, a typical case of fail safe. In addition, some cabin window assemblies have a third, pane of acrylic fitted to help reduce the engine noise in the cabin from the power-plants outside.

Passenger Cabin Window Figure 13 Most aircraft require one or more flight deck windows that can be opened for signalling to the ground-crew, for fresh air ventilation if the air conditioning is 'off' on the ground and to be able to see out in emergency situations, for example, the windscreen becoming obliterated. To achieve this, aircraft are usually fitted with a pair of opening front corner or side windows, sometimes called Direct Vision windows. If the cabin is pressurised, they will be unable to be opened due to the provision of a similar ‘pressure on’ safety lock system as the cabin doors.


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3.2 WINGS 3.2.1 CONSTRUCTION

The methods by which the wings produce lift were covered in Module 8, so this module will concentrate on wing construction and their attachments. To classify the many types of wing it is best to break them down into different groups. The first sub-division is either those that are externally braced or those that are of cantilever construction. (no external bracing). In the early days the majority of aircraft were constructed with the whole aircraft, including the wings, being braced by wires and struts. These produced very high drag, although the overall structural weight could be kept down. As materials and the wing construction became stronger, the number of wires were progressively reduced, until in the mid-1930's the first genuine fully cantilever wings with no external bracing, were put into production. This does not mean the bracing has been eliminated, it just means that all ‘bracing’ is included within the wing structure and made much stronger. Fig 14 below, shows how the external bracing of a biplane has been replaced with more efficient internal bracing on a cantilever wing.

Biplane and Cantilever Wing Bracing Figure 14 To illustrate how complex the inside of even a small aircraft wing can be, the following two pictures show the internal structure of both a wood and a metal wing.

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The heart of a wing is the spar (or spars), to which are attached the ribs stringers and other structural items. The number of spars is decided by the designer or design team, but modern airliners normally have two. It is usual to attach landing gears, primary flying controls, leading and trailing edge devices, to one or other of the spars within the wing on larger aircraft. Simpler wings on, for example, a light aircraft, will have only one main spar but some aircraft can have up to five, which has a measure of 'fail safe' philosophy. If military aircraft are considered, some modern fighters can have more than 15 spars as part of the ‘damage tolerant’ design application. Wing planforms can show an infinite number of different shapes, that are purpose built and satisfactory for providing lift. These could be generally grouped into straight, swept, delta and combination wings. Straight wings include those with a slightly swept leading edge, trailing edge or both. Swept wings are usually categorised as those with both leading and trailing edges swept back, at a variety of different angles, whilst the delta-winged shape (from the Greek for triangle) is self-explanatory. Under the cover-all title of 'Combination', the selection of silhouettes below should give an idea of the wide range of wings that can be found on modern day aircraft, in addition to the more conventional planforms mentioned above.

Wing Planforms Figure16


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3.2.2 FUEL STORAGE

Rigid Tanks Because of their shape, wings are often designed to be used for fuel storage. They can either contain separate fuel tanks within the wing structure, or use the wing structure itself, suitably sealed, to make integral tanks. Separate internal tanks are usually manufactured from either light alloy or from flexible, rubberised fabric. Rigid light alloy tanks are first riveted, then welded to make them fuel tight and are securely clamped into the wing structure by straps or tie bars. They will often have baffles inside, to prevent fuel surge from one end of the tank to the other.

Rigid Fuel tank Figure 17

Flexible Fuel tank Figure 18

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Flexible Tanks Flexible tanks, (Fig 18), also referred to as 'bladder' tanks, have to be located snugly into the tank bay within the wing, because the sides of the bay provide support to the relatively weak tank skin. Older types of flexible tanks were made from rubber- covered fabric. These days the fabric is replaced by man-made fibres, impregnated with neoprene or some similar fuel tight material. Integral Tanks Integral fuel tanks are found on most, if not all, modern commercial aircraft. During manufacture, practically the entire wing structure becomes a box, comprising front and rear spars, top and bottom wing skins, inboard and outboard sealed ribs, into which are installed pumps, drains, filler caps and vents. The main advantage of the integral tank, is that it provides maximum fuel capacity for the minimum amount of weight and the only sealing required, is that applied to the seams after construction is completed.

Boeing 737 Integral Fuel Tank capacities Figure19


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3.2.3 LANDING GEAR

As mentioned earlier, the attachments for major components can often be strong points on the wing spars, or even a separate spar built specifically for that purpose.. One such component that falls into this category is the main landing gear, otherwise known as the undercarriage. On some very large aircraft, like the Boeing 747 or Airbus A340, additional body gears, as well as conventional wing gears are to be found. These have to have reinforcements built into the lower fuselage structure to absorb the extreme loads at touch down.

Landing Gear Attachments Figure 20 Page 3-18


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3.2.4 PYLONS

Many aircraft have engines mounted on pylons attached to the wing. With this so called ‘podded engine’ configuration, the pylons have to take very large thrust forces from the engines and transfer it to the airframe. This is normally achieved by attaching the engine to strong points on the pylon and attaching the pylon to the wing spars. Thrust links are then fixed to the engine frame and the wing spars to transfer the engine thrust efficiently. Pylons must be positioned low enough so that the engine exhaust doesn’t strike the wing structure, but not too close to the ground to risk a runway scrape. The Boeing 737-600 is a fine example of this compromise.

Pylon Engine mounting Figure 21

Turbo-Propeller Mounting Figure 22 Issue 1 – 04 Sept 2001

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Boeing 737-600 Engine Pylon Mountings Figure 23

3.2.5 CONTROL SURFACE AND HIGH LIFT/DRAG ATTACHMENTS

SR 99

All of the flying controls on the wing will be attached to strong points on either the front or rear spars. This includes high and low speed ailerons, leading and trailing edge flaps, slats, roll spoilers, speed brakes and lift dumpers. The wing structure must therefore be made strong enough not only to carry the lift forces in flight but the additional loads of pilot control inputs, additional drag devices, etc. Consequently, the spars, are always the strongest part of the wing structure.

Control Surface mountings - Wings Figure24

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3.3 STABILISERS The vertical stabiliser (fin) produces directional or lateral stability, whilst the horizontal stabiliser (tailplane) produces longitudinal stability. As was mentioned in the aerodynamics section, these surfaces are of similar construction to the wings with spars, ribs, stringers etc,. They have to resist the twisting forces from the control surfaces mounted on the trailing edges. In many cases, the fin is similar to one half of the tailplane and on a number of light aircraft, it is actually constructed in this way, thereby simplifying production and component parts. Light aircraft have stabilisers manufactured from welded tube or fabricated from thin aluminium sheet of simple construction. As the aircraft size and weight increases, the surfaces will be made from stronger milled or machined skins and forged spars. Below can be seen examples of the empennage of light aircraft, Piper Cub and Cherokee and Cessna 150, showing their simple construction.

Empennage Construction Figure 25


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AIRFRAME STRUCTURES

3.4 FLIGHT CONTROL SURFACES The construction of most flight control surfaces is critical, since the designer wants to make them as light as possible. The control surfaces in the early years of aviation were a light, tubular frame covered with fabric and in later years when light alloy was adopted the quest for lightness continued. Today, metallic structures with honeycomb cores or epoxy reinforced composite construction are utilised for most control surfaces. The control surfaces are attached to the wing, fin or stabiliser by hinges, the spars being reinforced where these attachments are located. The cutaway below shows an elevator from a Fokker 100 and it can be seen that the construction is very similar to other main surfaces. The only difference is that the rear half of the surface has no internal framework but instead, a core of shaped aluminium honeycomb with the skin adhesively-bonded to it.

Elevator Structure Figure 26 To prevent the risk of flutter, as previously described, the ailerons, elevator and rudder, are all constructed so that the part of the surface behind the hinge line, is as light as possible and a number of calibrated weights are added to the leading edge of the surface. These weights are known as mass balance weights, (see cutaway above) and the procedure is known as mass balancing. In addition to mass balancing, surfaces that do not have the benefit of hydraulic power assistance, (see later) and are difficult to move when the aircraft is at high speed, have the benefit of aerodynamic balancing. To achieve this simply and as previously discussed, the hinge of the control is inset, so that part of the surface in front of the hinge line projects into the airstream, when the control is deflected from neutral. Page 3-22


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3.5 NACELLES AND PYLONS It has been mentioned previously, how the nacelles and pylons are attached to the wings, generally and to other parts of the airframe on selected aircraft. The main purpose of all these engine fairings is to keep the engines outside of the airframe itself. There are several reasons for this, but the major reasons are that it is safer, in the event of a fire or explosion, if it isolated from the fuselage or the wings by firewalls. Also, it is much easier for routine maintenance and engine changes, if the engine is externally mounted. Most nacelles are simply fairings which cover the power-plant in a streamlined manner, although, they usually also serve as the intake for jet and turbo-propeller engines. Most are covered by large, easy-to-open doors and panels, which allow quick and easy access. On some designs there can be smaller, quick release panels fitted into the larger ones, which allow access for maintenance, such as oil level quantity indicators, which need to be checked every time the engines are shut down. On light aircraft, engine nacelles are usually fairly simple GRP fairings which are split into two parts and removed by releasing a few screws or quick release fasteners. These also contain a small intake for the air to reach the carburettor of the piston engine. On many larger aircraft, particularly those with fan bypass engines, are fitted with thrust reversers as part of the cowlings. These are usually doors which translate rearwards and open up panels containing cascade vanes, which re-direct the exhaust thrust in a forward direction, when reverse thrust is selected after landing. These will be covered later in the power-plants chapter. Although they are much more efficient that the older designs, modern jet engines produce harmful high frequency noise. One way that the noise may be kept below the safe and legal minimum, is by making the cowlings out of honeycomb sandwich, which as well as being very light in weight is excellent at absorbing sound. The honeycomb can be manufactured from glass or carbon fibre and covered with composite or light alloy skin facing panels. The pylons which support the engines fitted on to the wings or the rear fuselage all have one main purpose, which is to transmit the full thrust of the engines into the airframe. They must be extremely strong and yet flexible, as the wing mounts especially have to move with the flexing of the wings. On many large aircraft, the space within the pylons is utilised to fit such components as heat exchangers, (radiators); air valves; fuel valves; pipes containing air, oil and fuel and electric cabling. All engines must be isolated from the rest of the aircraft, so that a fire can be completely contained within the nacelle and extinguished if the aircraft is equipped a fire extinguishing system. To this end, there will be a sealed bulkhead or divider between the engine and the airframe made of a fire resistant material such as titanium or stainless steel. Issue 1 – 04 Sept 2001

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AIRFRAME STRUCTURES

All engines are subject to vibration that can be sensed inside the aircraft. To reduce this, the engine mounts are designed not only to hold the engine securely and to transmit the thrust, but the mounts themselves are fabricated with a shock absorbing material. This is usually an elastomeric or metallic woven block and will absorb a large proportion of the vibration providing the passengers and crew with a smooth flight.

Typical Fan Engine Cowlings Figure 27

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Cowling and Pylon Fairing Installation Figure 28


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MODULE 11.04 AIR CONDITIONING AND CABIN PRESSURISATION

CONTENTS 4

AIR CONDITIONING AND CABIN PRESSURISATION ............... 4-1 4.1 4.2

4.3

4.4

4.5 4.6

4.7

4.8 4.9

4.10

4.11 4.12

INTRODUCTION ............................................................................. 4-1 AIR SUPPLY ................................................................................. 4-1 4.2.1 Engine Bleed Air (compression) .................................... 4-1 4.2.2 Air Compressors or Blowers .......................................... 4-2 4.2.3 Auxiliary Power Unit (APU) ........................................... 4-2 4.2.4 Ram Air ......................................................................... 4-3 4.2.5 Ground Cart .................................................................. 4-3 COOLING ..................................................................................... 4-4 4.3.1 Air Cycle Cooling........................................................... 4-4 4.3.2 Vapour Cycle Cooling ................................................... 4-9 HEATING ...................................................................................... 4-11 4.4.1 Exhaust Heating Systems ............................................. 4-11 4.4.2 Combustion Heating Systems ....................................... 4-12 TEMPERATURE CONTROL ............................................................. 4-12 HUMIDITY CONTROL ..................................................................... 4-14 4.6.1 Water Separation – Water Extractor .............................. 4-14 4.6.2 Water Infiltration ............................................................ 4-17 MASS FLOW CONTROL ................................................................. 4-18 4.7.1 Mass Flow Controller .................................................... 4-18 4.7.2 Spill Valve Flow Controller ............................................ 4-19 DISTRIBUTION SYSTEMS ............................................................... 4-20 4.8.1 Re-circulation Air System .............................................. 4-23 PRESSURISATION SYSTEMS .......................................................... 4-23 4.9.1 Control And Indication ................................................... 4-26 4.9.2 The Un-Pressurised Mode ............................................ 4-26 4.9.3 The Isobaric Mode ........................................................ 4-27 4.9.4 The Constant-Differential Pressure Mode ..................... 4-27 4.9.5 Cabin Air Pressure Regulator ........................................ 4-27 4.9.6 Isobaric Control System ................................................ 4-28 4.9.7 Differential Control System ............................................ 4-29 4.9.8 Safety Valves ................................................................ 4-31 ELECTRONIC PRESSURISATION CONTROL ...................................... 4-31 4.10.1 Flight Deck Control Panel.............................................. 4-32 4.10.2 Automatic Pressure Controller....................................... 4-33 4.10.3 Outflow Valve ................................................................ 4-33 4.10.4 Inward and Outward Safety Relief Valves ..................... 4-34 CABIN PRESSURE INDICATION ....................................................... 4-35 SAFETY AND WARNING DEVICES .................................................. 4-36 4.12.1 Overheating .................................................................. 4-36 4.12.2 Duct Hot Air Leakage .................................................... 4-37 4.12.3 Excess Cabin Altitude ................................................... 4-37 4.12.4 Smoke Detection ........................................................... 4-37

Issue 0 - 30 Jan 2003

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4 AIR CONDITIONING AND CABIN PRESSURISATION 4.1 INTRODUCTION The atmosphere above10,000ft is too thin and cold for normal breathing. Passenger carrying aircraft, operating above this height need an air conditioning and pressurisation system. The temperature of the air passing through the passenger cabin, flight deck and other compartments must be strictly controlled, as well as flow rate and level of humidity. Cabin temperature will normally be maintained between 15 and 30 degrees Celsius. Additionally, a controlled amount of pressurisation is necessary, so that the air pressure in the passenger cabin and adjacent areas does not exceed the equivalent of the ambient air pressure at 8000ft. Air conditioning is also essential for un-pressurised aircraft types. A typical air conditioning and pressurisation system comprises eight principle sub-systems:        

Air Supplies (Pneumatics ATA 36) Cooling Heating Temperature Control Humidity Control Mass Flow Control Distribution Pressurisation

4.2 AIR SUPPLY The source of fresh air supply and arrangement of essential components will vary between aircraft type and each air conditioning system, but in general one of the following methods described in the following paragraphs will be adopted: 4.2.1 Engine Bleed Air (compression) This method is the most common and is installed on the majority of modern aircraft types. Very hot air is tapped from the main engine compressor stages and supplied to the cabin, flight deck and other areas. Before the air enters the cabin, it is passed through a temperature control system, which reduces its temperature and pressure. Additionally, a means of flow control is utilised and in some aircraft, humidity control forms part of the system. (See Fig 1) In pressurised aircraft, the discharge of the conditioned air is regulated to maintain the cabin pressure at the selected pressure altitude.

Issue 1 - 20 March 2001

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AUXILIARY POWER UNIT

NON RETURN VALVE

SHUT OFF VALVES FLOW CONTROLLER

ECU TEMPERATURE CONTROL VALVE

NRV SECONDARY HEAT EXCHANGER

RAM AIR

TO CABIN

MIXER UNIT PRIMARY HEAT EXCHANGER

NRV

WATER SEPARATOR COUPLED COMPRESSOR TURBINE

Typical (Compression) Bleed Air System Figure 1

4.2.2 Air Compressors or Blowers This method is used on turbo-prop, piston engine or even turbo-jet aircraft where main engine compressor bleed is unavailable or unsuitable. Normally the compressor or blower will be mechanically driven from the accessory gearbox of the main engine and its air supply routed via a temperature control system, in a similar manner to the engine bleed method. 4.2.3 Auxiliary Power Unit (APU) The APU is a small gas turbine engine, which can be connected into the main air supply system and provide an independent means of air conditioning and pressurisation, either on the ground or in flight, when the main engines cannot supply. It will utilise the engine bleed air principle outlined above.

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This method is normally found as the primary ventilation system on unpressurised aircraft. A ram air scoop placed directly into the airflow, will provide the means of air supply as the aircraft moves forward. Since the air at altitude will be cold, the temperature control system through which it passes before entering the cabin, will normally be a form of heater. A self-contained combustion type heater will be employed, or the some form of exhaust gas heater. The air conditioning ducting will be routed around the combustion heater casing or around engine exhaust duct to obtain convection heating. On pressurised aircraft, a ram air system can be used as a means of emergency ventilation, following a complete loss of the main system. RAM AIR

COLD AIR OUTLETS

DEMISTER

WARM AIR OUTLETS

EXHAUST

COMBUSTION CHAMBER FLOW CONTROL VALVE FUEL SOLENOID VALVE

ENGINE DRIVEN AIR BLOWER AIR SUPPLY

FUEL SUPPLY OFF

ON

Typical Combustion Heater System Figure 2

COMBUSTION HEATING AIR CONDITIONING SYSTEM 4.2.5 Ground Cart This will be an independent means of heating or cooling the passenger cabin on the ground. It can be used on aircraft that do not have an APU. The trolley will be connected externally to the aircraft, via a purpose built inlet into the air conditioning system and normally employs a combustion type heater and the means to control the output of the air temperature from a control panel the cart.


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4.3 COOLING When bleed air is used as the air supply, the air tapped off the engine compressor can reach a temperature in excess of 300 degrees Celsius. This is obviously far too hot to be fed directly into the air-conditioned areas, so it must first be cooled down to around 20 degrees Celsius. There are two main methods of cooling; Air Cycle and Vapour Cycle cooling systems. 4.3.1 Air Cycle Cooling Air cycle cooling relies on three basic principles; surface heat exchange, expansion and energy conversion. Surface heat exchange, provides cooling by passing the air tapped from the engine compressor (charge air) across some form of heat exchanger. The charge air is subjected to the effect of a colder cross flow, normally ambient air, scooped by an intake and passed across the heat exchanger as the aircraft moves forward (ram air). Although 90% of heat is given up in this way, the charge air temperature can never be reduced below the ram air temperature by this method alone. Expansion, provides cooling when the pressure of the charge air is reduced by increasing its velocity and expanding it across the turbine of a so-called Air Cycle Machine (ACM) or Cold Air Unit (CAU). In this way, the temperature of the charge air can be rapidly lowered to zero degrees Celsius, irrespective of the ram air temperature Energy Conversion, cools by making the hot air do work. This is achieved by using the charge air to drive a turbine, which is connected by a shaft to the compressor or fan within the cold air unit, thus converting heat energy into kinetic energy. This method will also help to reduce the charge air to zero degrees Celsius.

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engineering HOT AIR INLET PRIMARY HEAT EXCHANGER

SECONDARY HEAT EXCHANGER

RAM AIR

TEMPERATURE CONTROL VALVE WATER SEPARATOR TO CABIN MIXER UNIT COMPRESSOR

TURBINE

Turbo Compressor Figure 3 4.3.1.1

HEAT EXCHANGERS

These are components within the air conditioning system that transfer heat from one gas stream to another. Ram air is used as the cooling medium to cool the very hot charge air ducted from the engine compressor or the gearbox mounted air compressor or blower. Depending on where they are placed within the air conditioning system, heat exchangers are often described as;  A ‘Pre-cooler’ or ‘Primary Heat Exchanger’  An ‘Inter-cooler’ or ‘Secondary Heat Exchanger’ The basic construction is a sealed unit containing a series of cooling passages; through which the charge air flows and over which the ram air is directed. Between these passages are thin corrugated strips, that also serve to dissipate heat as the ram air passes over them.


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engineering 4.3.1.2

AIR CYCLE MACHINE (ACM) OR COLD AIR UNIT (CAU)

The ACM/CAU is the primary component in an air cycle cooling system. A number of different types can be found including; The turbo-compressor, the brake turbine and the turbo-fan. All three use the charge air to drive the turbine and the major differences between each type, relates to the overall weight for a given mass flow, the size and method of dissipating the power output of the turbine. TO DISTRIBUTION SYSTEM

DIFFUSER

FROM INTERCOOLER

NOZZLE BLADES

BLEED AIR

COMPRESSOR

TO INTERCOOLER

TURBO COMPRESSOR Turbo Compressor Cold Air Unit Figure 4

The turbo-compressor type consists of a turbine driving a centrifugal compressor and operating in conjunction with an inter-cooler connected between the compressor and turbine stages. Its basic construction consists of two main casings, the turbine volute and compressor volute casings. The two casings are connected together and enclose a bearing housing with two bearing assemblies, supporting a shaft upon which the turbine and compressor wheels are mounted. Page 4-6


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The turbine wheel revolves within a nozzle ring and the compressor wheel rotates within a diffuser ring. The very hot charge air from the engine compressor bleed and routed via the pre-cooler, enters the eye of the ACM/CAU compressor. It becomes compressed on passing through the diffuser ring, increasing its temperature and energy. From the compressor, the hot air is directed across the inter-cooler matrix over which ram air passes and is then directed into the turbine volute nozzle ring, where it drives the turbine. The resultant expansion and energy conversion, rapidly lowers the air pressure and temperature. It is then directed towards the passenger cabin. (See Fig 3) The ACM/CAU compressor and turbine wheels rotate at extremely high speeds, often in excess of 80,000 rpm, so efficient bearing lubrication is essential to ensure smooth and trouble-free running. Two lubrication methods are used; Integral wet sump arrangements, or pressurised air bearings that need no oil lubrication. The wet sump type normally has a sump containing oil and a means of metering it to the bearings usually by the use of integral ‘wicks’ or with an ‘oil slinger’ that pumps an optimum oil/air mix to the bearings. This ensures the correct amount of oil at the bearings at all times. Oil replenishment is critical however, as too much oil will lead to the charge air being oil contaminated and too little oil, may result in a premature seizure of the rotating shaft. The air bearing type uses a pressurised air supply to support the shaft in a similar manner to the ‘hovercraft principal’. As the rotor ‘floats’ on a thin layer of air, it is essential that this type is kept clean and dry and completely free from oil and grease. AMBIENT AIR OUTLET

TURBINE

COMPRESSOR

TO CABIN

AMBIENT AIR INLET

HEAT EXCHANGER

MIXER UNIT

BLEED AIR

CONTROL VALVE RAM AIR

Brake Turbine Cold Air Unit Figure 5 Issue 1 - 20 March 2001

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The brake-turbine type of ACM/CAU, has its charge air routed directly from the pre-cooler to drive the turbine. The air expands across the turbine as before, resulting in a large temperature and pressure drop. Since this layout dispenses with the need for an inter-cooler, it results in a greater efficiency due to weight saving. To safeguard against the turbine rotating too fast, it is coupled with a compressor, which rotates in ambient air and consequently acts as a braking medium. Additionally, the slower rotation of the shaft further improves turbine output efficiency. (See Fig 5) BLEED AIR

RAM AIR OUTLET

TURBINE

RAM AIR

TO CABIN MIXER UNIT LARGE FAN

HEAT EXCHANGER

CONTROL VALVE

Turbo Fan Cold Air Unit Figure 6 TURBO FAN COLD AIR

UNIT

The turbo-fan type is mechanically similar to the brake-turbine arrangement. In this case however, the turbine drives a large centrifugal fan instead of a normal compressor. The fan is draws a large quantity of ambient air over the pre-cooler, which cools the incoming charge air. The major advantage of this type over the other two, is that with the fan-induced airflow over the pre-cooler, it can be used with the aircraft stationary on the ground with the aircraft engines running. It does not need to rely solely on ram air as the cooling medium for the pre-cooler.

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4.3.2 Vapour Cycle Cooling The vapour cycle cooling system can be used as an alternative to the air cycle cooling system. Although not commonly used these days for air conditioning systems, the system may be used as the means to remove heat from electrical and electronic equipment. The system relies on the principle of the ability of a refrigerant to absorb heat when changing from a liquid to a gas, through the process of vaporisation or expansion. For example, if you were to put a drop of a highly volatile liquid such as methylated spirits or petrol on the back of you hand, it will feel cold. This is because the liquid starts to evaporate and draws the heat necessary for evaporation from your hand. Liquids with a low boiling point have a stronger tendency to evaporate at normal temperatures than those with a high boiling point. Furthermore, the amount of pressure acting on a liquid substance will affect its state. A sufficient reduction in pressure will cause any liquid to change state into a vapour or a gas. Conversely, a corresponding increase in pressure will reverse the process.

CONDENSER RAM AIR RECIEVER DRYER

THERMOSTATIC EXPANSION VALVE

AIR SUPPLY

CAPILLARY TUBE TURBO COMPRESSOR

EVAPORATOR

TEMPERATURE SENSOR

TEMPERATURE CONTROL VALVES

AIR DISTRIBUTION

Schematic Vapour Cycle System Figure 7 Issue 1 - 20 March 2001

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The major components of a typical system are a liquid receiver, a thermostatic expansion valve, an evaporator, a turbo-compressor, a condenser and a condenser fan. Often these components are mounted close together to form a line-replaceable refrigeration pack or vapour cycle cooling pack. The liquid receiver acts as a reservoir and provides storage for the refrigerant, normally a highly volatile chemical such as Freon. The refrigerant will pass from the liquid receiver to a thermostatic expansion valve where it is metered and released into the evaporator. The very hot charge air from the main engine bleed flows across the evaporator, releases heat that vaporises the liquid refrigerant and passes into the passenger cabin at a much lower temperature. Meanwhile, the now vaporised refrigerant gas is directed towards the turbocompressor. It is drawn into the compressor wheel, the coupled turbine of which is driven by the main engine bleed air. (Note: In some cases, an independent means instead of a turbo-compressor may be used to compress the refrigerant gas, such as an electric motor, as in a domestic refrigerator). The refrigerant gas leaves the compressor at a high pressure and temperature and passes across the matrix of the condenser. The gas is cooled by the ram air, flowing across the matrix and so condenses back into a liquid once again. It then returns to the liquid receiver to repeat the refrigeration cycle once again. The condenser fan is used to induce air across the condenser matrix when the aircraft is stationary on the ground and no ram air is available.

Typical Vapour Cycle System Figure 8 Page 4-10


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4.4 HEATING Un-pressurised aircraft use a ram-air system for ventilation. At altitude, the ramair passing through the cabin would be very cold, so a heating system is required. Heating systems can be generally divided into two types:  Exhaust heating systems  Combustion heating systems 4.4.1 Exhaust Heating Systems In its simplest form, this type of heating system employs a heater muff that surrounds the exhaust pipes coming from a piston engine, or the jet pipe of a turbo-jet. A ram air scoop at the forward end of the heater muff allows some of the cold air to go to directly to a mixing valve. The remainder, enters the muff and surrounds the exhaust/jet pipes. Heat from the pipes is transferred into the ram air and carried to the mixing valve. The heated air joins the cold air at the mixing valve and the combined flow is directed into the passenger cabin. Some form of control lever, operated from within the aircraft and connected to the mixing valve, allows the proportion of hot and cold air to be modulated in order to suit the cabin heating requirements. To cater for the possibility of the ventilation air becoming contaminated from the exhaust pipes, some aircraft will be fitted with carbon monoxide detectors within the cabin area. These are indicators filled with brightly coloured crystals, which turn black if exposed to dangerous levels of carbon monoxide.

Exhaust System Heater Figure 9 Issue 1 - 20 March 2001

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4.4.2 Combustion Heating Systems This system uses a purpose built combustion chamber heater assembly to provide the heat source, rather than the previously described exhaust heating method. Fuel is directed from the aircraft fuel system, through a pressure regulating and shut off valve that ensures the fuel is at the correct pressure for atomisation. Other components include a fuel filter, a fuel pump and spray nozzle, where it is atomised and ignited with an igniter plug. The combustion chamber assembly heats up the ram air that passes around it.

Typical Combustion Heater System Figure 10 4.5 TEMPERATURE CONTROL In order to operate the aircraft in an infinite number of climatic and operating conditions, the temperature in the passenger cabin, flight compartment and other areas needs to be regulated for comfort. Temperature regulation for the majority of aircraft that employ the engine bleed air method is usually accomplished by controlling the proportion of hot and cold air coming from the air supply system. An electric motor driving a double butterfly type air mixing valve, regulates the cabin temperature, by allowing a controlled amount of hot air to by-pass the air cycle system. This air is then recombined in proper proportions with the cold air that has been directed through the air cycle system at a down stream mix chamber. The position of the air-mixing valve is determined by signals from the temperature control system. The temperature control system is normally operated automatically or as a manual system, if the automatic temperature controller should fail. Page 4-12


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During automatic operation, the temperature controller continually monitors cabin temperatures and repositions the air mixing valve if necessary to keep the temperature at the selected level. In order to achieve this, the controller receives signals from temperature selector on the flight deck (the temperature requested) and from temperature sensors in the passenger cabin, flight compartment and supply ducts (the actual temperature). If a difference between the requested and actual temperatures occurs, the controller will send an output signal, to re-position the air mixing valve until parity exists once more. During manual operation, the temperature control circuit bypasses the controller and connects the temperature selector on the flight deck, directly to the air-mixing valve. Other sensors in the system transmit compartment temperatures to indicators on the flight deck overhead panel, so that the actual temperatures and the position of the air-mixing valve can be monitored.

Temperature Control Figure 11


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4.6 HUMIDITY CONTROL Humidity control is the means to ensure that the correct amount of water moisture content is in the air conditioning air within the aircraft cabin. This is necessary to ensure occupants do not suffer from low humidity levels that are experienced with high altitude flight. Humidity control can be achieved two ways;  Water Separation  Water Infiltration Water separation is the removal of excessive moisture from the charge air, normally by a water extractor or separator. Water infiltration is the addition of moisture into the conditioned air as it enters the cabin using a water pump and spray nozzle. 4.6.1 Water Separation – Water Extractor Water can be introduced into the air conditioning system due to the compression and expansion of the air in the ACM/CAU and other areas of the air cycle process. There are three types of water separator in general use; the coalescer/diffuser type, the coalescer/bag type and the swirl vane type. 4.6.1.1

COALESCER/DIFFUSER TYPE

This type consists of a coalescer constructed from layers of monel metal gauze and glass fibre cloth sandwiched between layers of stainless steel gauze. It is supported by the diffuser cone and held in place by a relief valve housing. As the air leaves the diffuser and passes over the coalescer, moisture in the air is converted into water droplets. The droplets enter the collector shell and are deposited into collector tubes where they drain down to a collector box from where the water is ejected overboard. COALESCER

COLLECTOR SHELL

DIFFUSER

PRESSURE RELIEF VALVE

CONDENSER TUBES

DRAIN

COALESCER WATER EXTRACTOR FIGURE 12 Page 4-14


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COALESCER/BAG TYPE

A porous bag, supported by a shell is fitted within the extractor to convert moisture into water droplets. A swirl is imparted into the conditioned air and the centrifugal effect forces the droplets to the outlet shell where it collects and drains from the component. A bag visual indicator operated by back pressure, will show when the coalescer bag becomes dirty or blocked. In this case, a relief valve will open to ensure flow is still available.

BLOCKAGE INDICATOR OUTLET SHELL BAG

PRESSURE RELIEF VALVE WATER DRAIN

Bag Type Water Extractor Figure 13


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engineering 4.6.1.3

SWIRL VANE TYPE

This type uses centrifugal force to spin the moisture-laden air outwards against the exit shell. The swirl vane, either fixed or rotating imparts the swirl by rotating the airflow at high speed. The action, separates the heavier water droplets in the moisture and collects them in a sump, to be drained away.

SEPARATOR SHELL

SWIRL VANE

WATER SUMP DRAIN Swirl Vane Type Water Separator Figure 14

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engineering 4.6.2 Water Infiltration

Humidity control can also include the addition of water into the air conditioning system. As an aircraft climbs to high altitude, the moisture level in the air reduces to a much lower amount than at lower levels of altitude. The reduction in moisture may cause discomfort to the aircraft occupants. To counteract this, moisture is added into the conditioned air, by pumping water from a tank to a spray nozzle positioned at the cabin air inlet. Humidity sensors will detect low humidity conditions and automatically turn on the controller water pump to restore the humidity to acceptable levels.

WATER SEPARATOR DRAIN

COLLECTOR TANK CABIN HUMIDITY SENSOR

SPRAY NOZZLE

OVERFILL DRAIN

WATER PUMP AND CONTROLLER TO CABIN

Typical Humidity Control System Figure 15


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4.7 MASS FLOW CONTROL Legislation requires that a minimum amount of fresh air be supplied to passengers and crew. In addition stale air must be removed and odours eliminated. Most pressurisation systems rely on the fact that air is delivered at a constant rate under all conditions of flight in order to function correctly. Mass flow control systems constantly monitor the velocity and density of the air supply by either increasing or decreasing the demand upon the source of supply, or by spilling excess supply air overboard. The mass of air must be controlled at a constant value regardless of aircraft altitude or cabin pressure. It must also adjust for changes in main engine compressor speed in bleed air systems, or changes in rotor speed when a separate air supply from an accessory gearbox driven blower is incorporated. 4.7.1 Mass Flow Controller This type automatically caters for changes in air density, cabin back pressure and engine compressor supply pressure. At ground level and during take off and the early stages of flight, the pressure available from the main engine compressor outlet is high. As altitude increases or when the engines are set to cruising speeds, the supply pressure drops. The amount of pressure from the engine compressor bleed acting on an altitudecompensated piston valve, determines the position the valve will adopt when opposed by a spring and back pressure from the cabin. The pressure drop across the valve, will vary the size of outlet ports and will thus determine the valve’s degree of opening and closing. This will result in a constant mass flow downstream of the valve at all times.

Mass Flow Control Figure 16 Page 4-18


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4.7.2 Spill Valve Flow Controller This type receives the charge air supply through a metering duct, which senses variations in the velocity and density of the air. The metering duct on sensing these variations, transmits the information to a mass flow controller, which converts the air pressure signals into electrical signals. The electrical signals in turn control the position of spill valves. They will move towards a more open or closed position, to vary the amount of air spilled overboard, thereby ensuring a constant flow rate into the cabin. At sea level, with the engines at low power, the absolute capsule D will be compressed by atmospheric pressure. The contacts A, B and C will be in the position shown and the spill valve will be towards closed. With the main engines at take off power, the air velocity through the venture increases, causing a pressure differential across the controller diaphragm. This will cause contact B to move towards contact C and when they touch, the spill valves will be driven towards the open As the aircraft climbs, the static pressure in the metering duct and controller will decrease. The absolute capsule will now expand and the position of contacts A and C, will be adjusted in relation to contact B. When contact B is touched, the spill valves will move towards closed once more and once again the mass flow to the cabin will remain constant.

Mass Flow Controller Operations Figure 17


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engineering 4.8 DISTRIBUTION SYSTEMS

The air distribution system on most aircraft takes cold air from the air conditioning packs and hot air bleed from the engines and mixes the 2 in a mixer unit to the required temperature. The air is then distributed to side wall and overhead cabin vents. On some aircraft the cabin air is then drawn back into the mixing unit by recirculating fans where it is mixed with new air and then re-distributed. All major components are usually located together in a designated bay for ease of maintenance. ( Figure 14). A gasper fan provides cold air to the individual overhead air outlets for the aircrew and passengers. This air can be drawn direct from outside or from the cooling packs. Each passenger or crew can control the amount of air received by controlling the position of the air outlet. This outlet could be a rotary nozzle or a louvre. TO SIDEWALL DUCTS TO GASPER OUTLETS

GASPER FAN MIXER VALVES MANIFOLD RELIEF VALVE

TO COCKPIT TO SIDEWALL DUCTS

CONTROL VALVES WATER SEPARATOR

TO OVERHEAD DUCTS

CONTROL VALVE SELECTOR LINKAGE

Air Conditioning Distribution Manifold Figure 18 Page 4-20


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Conditioned air systems dispense temperature controlled air evenly throughout the cabin and crew areas. One duct system supplies the cockpit (Figure 17) while another supplies the cabin. The cabin ducting is then divided into 2 systems, the overhead (Figure 15) and the sidewall systems (Figure 16). The overhead system releases air into the cabin from outlets in ducting running fore and aft in the cabin ceiling. The sidewall duct system takes air through ducting between the sidewall and cabin interior linings and releases it through cove light grills and louvres. A cockpit controlled selector valve located on the main distribution manifold allows all overhead, side wall or any combination of the two systems to be used and varies the flow between the two.

DUCTING

FLOOR EXHAUST DUCT

ADJUSTABLE AIR OUTLETS

GASPER FAN

Overhead Panel Figure 19 Duct sections throughout both the cabin and cockpit are joined together with clamps or clips. Means of equalising the duct pressures and balancing the air flows are designed into each system. The systems are protected from excess pressures by use of a spring loaded pressure relief valve usually located in the main distribution manifold. The main manifold is located immediately downstream from the mixing units in the air conditioning bay. On large aircraft a cockpit controlled dual selector valves divides the air between cockpit and cabin areas. These butterfly valves are interlinked. When one is fully open the other is fully closed and vice versa.


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Air is exhausted from the passenger cabin through grills and outflow valves in the sidewalls above the floor. This air can then be directed around the cargo compartment walls where it assists in compartment temperature control. Some air then flows to the cargo heat distribution duct under the compartment floor and is then discharged overboard through the outflow valves.

DISTRIBUTION BOXES

WALL FEEDER DUCTS

WINDOW DEMISTER

FLOOR EXHAUST VENTS

DISTRIBUTION DUCT

Sidewall Ducting Figure 20 Below each floor air exhaust outlet is a flotation check valve. This valve is a plastic ball held in a cage. If the cargo compartments become flooded the balls float up the cage and seals off the floor to help prevent water from entering the cabin. CABIN TEMPERATURE SENSOR

AIR VENT

FLIGHT DECK TEMPERATURE SENSOR

SILENCER

FAN ASSY COOLING FANS FAN ASSY PRESSURE SWITCH

Cockpit Air Distribution Figure 21 Page 4-22


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Aircraft may be separated into zones each with its own air conditioning system and controls for that zone located in a distribution bay. Some areas may have a remote heat exchanger and fan assembly in the vapour cycle system, to allow cooling to specific areas such as avionics bays, fed from one of the zone packs. 4.8.1 Re-circulation Air System To improve cabin ventilation and supplement airflow the cabin air is recirculated back to the main distribution manifold where it is mixed with conditioned air form the cooling packs. The use of re-circulated air improves airflow and offloads the air supply system. This off loading of the air conditioning packs is converted into a fuel saving. The re-circulation fan will draw air from the cabin area, through a check valve and filter assembly to remove any smoke and noxious odours before passing it to the mixer unit for re-distribution. The check valve prevents any reverse flow through the fan and ducting when the fan is not in use. 4.9 PRESSURISATION SYSTEMS As aircraft became capable of obtaining altitudes above that at which flight crews could operate efficiently, a need developed for complete environmental systems to allow these aircraft to carry passengers. Air conditioning could provide the proper temperature and supplemental oxygen could provide sufficient breathable air. The problem was that not enough atmospheric pressure exists at high altitude to aid breathing in and even at lower altitudes the body must work harder to absorb sufficient oxygen, through the lungs, to operate at the same level of efficiency as at sea level. This problem is overcome by pressurising the cockpit/ cabin area. Cabin pressurisation is a means of adding pressure to the cabin of an aircraft to create an artificial atmosphere that when flying at high altitudes it provides gives an environment equivalent to that below 10000 feet. The minimum quantity of fresh air supplied to each person on board must be at least 0.5lb/ minute. Aircraft are pressurised by sealing off a strengthened portion of the fuselage. This is usually called the pressure vessel and will normally include cabin, cockpit and possibly cargo areas. Air is pumped into this pressure vessel and is controlled by an outflow valve located at the rear of the vessel. Sealing of the pressure vessel is accomplished by the use of seals around tubing, ducting, bolts, rivets, and other hardware that pass through or pierce the pressure tight area. All panels and large structural components are assembled with sealing compounds. Access and removable doors and hatches have integral seals. Some have inflatable seals.


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Pressurisation systems do not have to move large volume of air. Their function is to raise the pressure inside the vessel. Small reciprocating engine powered aircraft receive their pressurisation air from the compressor of a coupled turbocharger. Larger reciprocating engine powered aircraft receive air from engine driven compressors and turbine powered aircraft use compressor bleed air Small Reciprocating Engine Powered Aircraft Turbochargers are driven by the engine exhaust gases flowing through a turbine. A centrifugal compressor is coupled to the turbine. The compressors output is fed to the engine inlet manifold to increase manifold pressure which allows the engine to develop its power at altitude. Part of this compressed air is tapped off after the compressor and is used to pressurise the cabin. The air passes through a flow limiter (or sonic venturi) and then through an inter-cooler before being fed into the cabin. A typical system is shown at Figure 22. 

Sonic Venturi

A sonic venturi is fitted in line between the engine and the pressurisation system. When the air flowing across the venturi reaches the speed of sound a shock wave is formed which limits the flow of air to the pressurisation system RAM AIR HEATING AIR PRESSURISED AIR

RAM AIR SHUT OFF VALVE

EXHAUST GASES

COUPLED TURBO COMPRESSOR

COMBUSTION HEATER

SONIC VENTURI INTERCOOLER

OUTFLOW VALVE

SAFETY VALVE

Small Reciprocating Engine Aircraft Pressurisation System Figure 22 Page 4-24


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Large Reciprocating Engine Powered Aircraft These aircraft use engine driven compressors driven through an accessory drive or by an electric or hydraulic motor. Multi engine aircraft have more than one air compressor. These are interconnected through ducting but each have a check valve or isolation valve to prevent pressure loss when one system is out of action. Turbine Powered Aircraft The air supplied from a gas turbine engine compressor is contamination free and can be suitably used for cabin pressurisation (Figure 23). Some aircraft use an independent compressor driven by the engine bleed air. The bleed air drives the coupled compressor which pressurises the air and feeds it into the cabin

FLUSH AIR INTAKE

TURBO COMPRESSOR

PRESSURE VESSEL (CABIN/COCKPIT)

BLEED AIR

OUTFLOW VALVE

ENGINE

Turbo Compressor Figure 23 Some aircraft use a jet pump to increase the amount of air taken into the cabin (Figure 24). The jet pump is a venturi nozzle located in the flush air intake ducting. High velocity air from the engine flows through this nozzle. This produces a low pressure area around the venturi which sucks in outside air. This outside air is mixed with the high velocity air and is then passed into the cabin


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FLUSH AIR INTAKE

PRESSURE VESSEL (CABIN/COCKPIT) JET PUMP BLEED AIR OUTFLOW VALVE ENGINE

Jet Pump Figure 24 4.9.1 Control And Indication There are 3 modes of pressurisation, un-pressurised, the isobaric mode and the constant–differential pressure mode. In the un-pressurised mode the cabin altitude remains the same as the flight altitude. In the isobaric mode the cabin altitude remains constant as the flight altitude changes and in the constantdifferential pressure mode, the cabin pressure is maintained at a constant amount above the outside ambient air pressure. The amount of differential pressure is determined by the structural strength of the aircraft. The stronger the aircraft structure the higher the differential pressure and the higher is the aircrafts operating ceiling. 4.9.2 The Un-Pressurised Mode In this mode the outflow valve remains open and the cabin pressure is the same as the outside ambient air pressure. This mode is usually from sea level up to 5000` but does vary from aircraft to aircraft.

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engineering 4.9.3 The Isobaric Mode

In this mode the cabin pressure is maintained at a specific cabin altitude as flight altitude changes. The cabin pressure controller begins to close the outflow valve as the aircraft climbs to a chosen cabin altitude. The outflow valve then opens or closes (modulates) to maintain the selected cabin altitude as the flight altitude changes up or down. The controller will then maintain the selected cabin altitude up to the flight altitude that produces the maximum differential pressure for which the aircraft structure is rated. At this point the constant differential mode takes control. 4.9.4 The Constant-Differential Pressure Mode Cabin pressurisation puts the aircraft structure under a tensile stress as the cabin pressure expands the pressure vessel. The cabin differential pressure is the ratio between the internal and external air pressures. At maximum constant-differential pressure as the aircraft increases in altitude the cabin altitude will increase but the internal/external pressure ratio will be maintained. There will be a maximum cabin altitude allowed and this will determine the ceiling at which the aircraft can operate. 4.9.5 Cabin Air Pressure Regulator The pressure regulator maintains cabin altitude at a selected level in the isobaric range and limits cabin pressure to a pre-set pressure differential in the differential range by regulating the position of the outflow valve. Normal operation of the regulator requires only the selection of the desired cabin altitude and cabin rate of climb the adjustment of the barometric control. STATIC ATMOSHERE CONNECTION ADJUSTER CONTROL

ISOBARIC METERING VALVE ADJUSTER CONTROL

DIAPHRAGM

BAROMETRIC CAPSULE

RESTRICTOR

DIFFERENTIAL METERING VALVE HEAD

SOLENOID DUMP VALVE

REFERENCE CHAMBER PILOT

BASE

ACTUATOR DIAPHRAGM OUTFLOW VALVE BAFFLE PLATE

Cabin Pressure Regulator Figure 25 Issue 1 - 20 March 2001

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The regulator shown in Figure 25 is a typical differential pressure type regulator that is built into the normally closed air operated outflow valve. It uses cabin altitude for its isobaric control and barometric pressure for the differential control. A cabin rate of climb controller controls the pressure change inside the cabin. There are 2 main sections to the regulator, the head and reference chamber and the base with the outflow valve and diaphragm. The balance diaphragm extends outward from the baffle plate to the outflow valve creating an air chamber between the baffle plate and the outer face of the outflow valve. Cabin air flowing into this chamber through holes in the side of the outflow valve exerts a force against the outer face of the valve which tries to open it. This force is opposed by the force of the spring around the valve pilot which tries to hold the valve closed. The actuator diaphragm extends outward from the outflow valve to the head assembly creating an air chamber between the head and the inner face of the outflow valve. Air from the head and reference chamber exert a force against the inner face of the outflow valve helping the spring to hold the valve closed. The position of the outflow valve controls the amount of cabin air that is allowed to flow from the pressure vessel and this controls the cabin pressure. The position of the outflow valve is determined by the amount of reference chamber air pressure that presses on the inner face of the outflow valve. 4.9.6 Isobaric Control System The isobaric control system of the pressure regulator shown in Figure 26 incorporates an evacuated capsule, a rocker arm, valve spring and a ball type metering valve. One end of the rocker arm is connected to the valve head by the evacuated capsule and the other end of the arm holds the metering valve in a closed position. A valve spring located on the metering valve body tries to move the metering valve away from its seat as far as the rocker arm allows. When the cabin air pressure increases enough for the reference chamber air pressure to compress the evacuated capsule the rocker arm pivots around its fulcrum and allows the metering valve to move away from its seat an amount proportional to the compression of the capsule. When the metering valve opens reference pressure air flows form the regulator to atmosphere through the atmospheric chamber.

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ISOBARIC METERING VALVE EVACUATED BELLOWS

OUTFLOW VALVE

Isobaric Control Operation Figure 26 When the regulator is operating in the isobaric range, cabin pressure is held constant by reducing the flow of reference chamber air through the metering valve. This prevents a further decrease in reference pressure. The isobaric control responds to slight changes in reference pressure by modulating to maintain a constant pressure in the chamber throughout the isobaric range of operation. Whenever there is an increase in cabin pressure the isobaric metering valve opens which decreases the reference pressure and causes the outflow valve to open which then decreases the cabin pressure. 4.9.7 Differential Control System The differential control system of the pressure regulator (Figure 27) incorporates a diaphragm a rocker arm, a valve spring and a ball type metering valve. One end of the rocker arm is attached to the head by the diaphragm which forma a pressure sensitive face between the reference chamber and the atmospheric chamber.


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ATMOSPHERIC CHAMBER

METERING VALVE DIAPHRAGM

OUTFLOW VALVE

Differential Pressure Mode Figure 27 Atmospheric pressure acts on one side of the diaphragm and reference chamber pressure acts on the other. The opposite end of the rocker arm holds the metering valve in a closed position. A valve spring located on the metering valve body tries to move the metering valve away from its seat as far as the rocker arm allows. When reference chamber pressure increases to the system differential pressure limit set above the decreasing atmospheric pressure it collapses the diaphragm which is set at differential pressure and opens the metering valve. Air flows from the reference chamber to atmosphere through the atmospheric chamber, which causes a reduction in the reference pressure. This reduction in reference pressure causes the outflow valve to open to reduce the cabin pressure to maintain the system pressure differential.

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4.9.8 Safety Valves Cabin Air Pressure Safety Valve The pressure relief valve prevents cabin pressure from exceeding the predetermined cabin to ambient pressure differential. A negative pressure relief valve and pressure dump valve may also be incorporated into this valve assembly. Negative Pressure Relief Valve A pressurised aircraft is designed to operate with the cabin pressure higher than the outside air pressure. If the cabin pressure were to become lower than the outside air pressure the cabin structure could fail. Outside air is allowed to enter the cabin to ensure that this does not happen. It is basically an inward pressure relief valve. Dump Valve This valve is normally solenoid actuated by a cockpit switch. When the solenoid is energised the valve opens dumping cabin air to atmosphere. Cabin pressure will decrease rapidly until it is the same as the outside air pressure and cabin altitude will increase until it is the same as the flight altitude. Ditching valve If any of the cabin control valves were situated below the water level and the aircraft ditch in the water, the cabin would quickly flood. To prevent this happening, either a mechanical or electrical ditching selection, can be made by the crew to seal off all pressurisation valves and inlets. 4.10

ELECTRONIC PRESSURISATION CONTROL

Most modern airliners have the means to electronically control the cabin pressure automatically for the entire flight, from settings made by the flight crew before take off. The pressure control system consists:    

a flight deck control panel an automatic pressure controller with pressure sensing inputs and outputs to monitoring indicators an electrically-driven gate-type outflow valve inward and outward safety relief valves


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4.10.1 Flight Deck Control Panel This provides a means for the flight crew to control the cabin pressure by positioning the outflow valve. There are three mode selections available; ‘Auto’, ‘Standby’ or ‘Manual’.

Figure 28 The desired mode will normally be ‘Auto’, where all settings such as intended cruise (flight) altitude and destination airfield (landing) altitude are made before flight. This will allow automatic control of cabin pressure for the whole of that flight. This is called the fully automatic mode. Alternatively, ‘Standby’ or back up mode can be selected, where a cabin altitude setting must be made for each desired cabin pressure change. The input setting is then controlled automatically as before. This is called the semi-automatic mode. If neither the fully or semi-automatic modes are available, (i.e.: the pressure controller fails), the outflow valve can be positioned directly from the flight deck by operating the electric torque motors to drive the valve. This is called the manual mode and a choice of an ac or dc electrical supply is available.

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4.10.2 Automatic Pressure Controller The pressure controller provides output control signals to the outflow valve’s ac or dc torque motors. The motors position and modulate the valve to establish and control actual cabin pressure in accordance with the controller’s pre-programmed climb, cruise or descent schedules. This will ensure that for every aircraft altitude there will be a particular cabin altitude. Input signals to the controller are from the flight deck control panel, cabin and ambient pressure sensors, barometric correction and air/ground sensing.

Auto Mode Flight Profile Figure 29 4.10.3 Outflow Valve The valve has a moving gate designed to cover or uncover an aperture in the fuselage skin. An increase in the aperture size will cause cabin pressure to fall (cabin altitude to ascend), whereas a decrease in the aperture size results in an increase in cabin pressure (cabin altitude to descend). The gate is driven by one of two electrically driven motors, the choice of ac or dc motor being determined by flight crew input. Motor input signals come from the controller when in the auto or standby modes, or directly from a control panel switch when in the manual mode.


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Outflow Valve Figure 30 4.10.4 Inward and Outward Safety Relief Valves Fuselage frames are designed to accept tensile loads associated with and outward force from within the pressure cell. Their ability to withstand compression loads that would occur if the pressure outside the aircraft were higher than within the pressure cell is poor. Therefore an inward relief valve will open and equalise the pressure if the inward or negative differential exceeds about 0.5 psid. Two outward relief valves are fitted to prevent the maximum outward differential pressure from exceeding the structural limit. This will typically be around 8.5psid. Even though the main pressure control is electronic, the safety relief valves are mechanical operated and are completely independent of any automatic control system.

Pressurisation System Valves Figure 31 Page 4-34


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CABIN PRESSURE INDICATION

Most pressurisation systems have three basic cockpit indicators cabin altitude, cabin rate of climb and the pressure differential indicator. The cabin altitude gauge measures the actual cabin altitude.

Cabin Altitude Gauge Figure 32

The cabin rate of climb indicator tells the pilot the rate that the cabin is either climbing or descending. (I.e. the rate at which the cabin loses or gains pressure) A typical maximum climb rate is 500ft per minute and the maximum descent rate is 300ft per minute. The control can be automatic or manual depending on aircraft type.

Cabin Rate of Climb Figure 33


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The differential pressure gauge (Figure 34) reads the difference between the cabin and the outside air pressures. This differential pressure is normally controlled and maintained to a structural limitation around 7psid. This depends on the aircraft type and the operating ceiling of the aircraft. The differential pressure gauge may be combined with the cabin altitude (Figure 35).

0 10

1 2

DIFF PX PSI

9

3

8

4 7

6

5

Differential Pressure Gauge Figure 34 4.12

Dual Gauge Figure 35

SAFETY AND WARNING DEVICES

To ground test the pressurisation system with the engines running, at least three men are required inside the aircraft for safety reasons. Both air conditioning and pressurisation systems use safety and warning devices to protect the aircraft from possible catastrophic failures. Some of the protection devices may be inhibited in certain stages of flight; landing or take off where the extra distractions caused by such warnings may be too much for the crews to deal with safely. With the air conditioning system the main concerns are with overheating of the air conditioning packs and extraction and ventilation fans, as well as hot air leaks from ducting which could damage surrounding structure or components. 4.12.1 Overheating Most packs systems are protected from overheating by a thermal switch downstream of the pack outlet. If the outlet temperature reaches a pre determined figure the switch will operate causing the pack valves to shut, preventing air from getting to the packs, as well as sending a warning signal to the cockpit central warning panel with associated caution/warning lights and aural chimes and to illuminate a fault light on the pack selector switch. Page 4-36


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Once the system has cooled down sufficiently the crew may have an option to reselect the overheated system. The overheat may have been caused by a fault in the automatic temperature control system in which case the pilot may be able to control the system manually via a manual selector switch on the cockpit controller. Extraction or ventilation fans will be protected in much the same way. An overheat will signal the central warning panel with associated caution/warning lights and aural chimes. The fan may be isolated automatically or manually. Once the fan has cooled down it may be possible to re-select if required. Fans may also be protected from over or under speeding, which will also have an effect on the system temperatures. Speed sensors on the fan will indicate a fault when over or under speed limits are reached and a warning signal is sent to the cockpit central warning panel with associated caution/warning lights and aural chimes. 4.12.2 Duct Hot Air Leakage Any ducting that includes joints is liable to leak under abnormal conditions. A duct protection system will include fire-wire elements around the hot zones such as engine air bleeds, air conditioning packs and auxiliary power units if fitted. The sensing elements will be the thermistor type. As the temperature around the wire increases the resistance decreases until an electrical circuit is made. When the circuit is made a warning signal is sent to the cockpit central warning panel with associated caution/warning lights and aural chimes. The leaking duct may be isolated automatically or may require the pilot to take action to close off the air valves. The faulty system will then remain out of use. 4.12.3 Excess Cabin Altitude If the cabin altitude was allowed to increase unchecked the crew and passengers could unknowingly suffer the effects of hypoxia. This dangerous condition is obviously undesirable especially for the aircrew. Most aircraft give a warning on the CWP with associated audio and visual warnings when the cabin altitude reaches 10000`. 4.12.4 Smoke Detection Smoke detectors may be fitted within the cabin; avionics bay and cargo areas to monitor systems, which if become faulty may generate smoke on overheating, or are may be liable to catch fire. These detectors will send a signal to the CWP with associated lights and audio warnings. They may also automatically switch on extractor fans, which will remove the smoke overboard and away form the cabin and cockpit areas. In this event, the pilot may have a switch or control lever to operate a valve to isolate the cockpit air conditioning ducting from the rest of the aircraft to prevent any smoke from getting to the cockpit. Issue 1 - 20 March 2001

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MODULE 11.07 EQUIPMENT AND FURNISHINGS

CONTENTS 7

EQUIPMENT AND FURNISHINGS .............................................. 7-3 7.1 7.2 7.3 7.4 7.5 7.6 7.7 7.8 7.9

EMERGENCY EQUIPMENT REQUIREMENTS ..................................... 7-3 SEAT, HARNESSES AND BELTS ..................................................... 7-4 CABIN LAYOUTS ........................................................................... 7-6 CABIN FURNISHINGS..................................................................... 7-7 CABIN ENTERTAINMENT ................................................................ 7-8 GALLEY INSTALLATIONS ............................................................... 7-8 CARGO HANDLING AND RETENTION EQUIPMENT ............................ 7-9 CARGO RETENTION EQUIPMENT .................................................... 7-14 AIRSTAIRS ................................................................................... 7-15

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EQUIPMENT AND FURNISHINGS

7.1 EMERGENCY EQUIPMENT REQUIREMENTS On every aircraft, there can be found some form of emergency equipment. This can vary from a simple seat belt and a fire extinguisher on a micro-light aircraft, to a large list of equipment fitted to a commercial airliner. For example, a medium sized aircraft like the Fokker 50 carries over thirty different types of safety equipment. The list of equipment fitted in a 450+ seater Boeing 747-400, will include items such as seat belts, lifejackets, first-aid kits, fire extinguishers, oxygen sets, torches etc. The types of safety equipment that must be carried on any specific flight, are laid down in the Air Navigation Order, (ANO), schedule No.4. This list covers a wide range of safety equipment, from mooring equipment for seaplanes to cookers and snow shovels for arctic operation. JAR 25 - Large Aeroplanes, details amongst others, the requirements for the design and performance of safety and other equipment, ranging from size of access doors and emergency exits and the numbers required for each size of aircraft, width of cabin aisles, number of seats abreast. The list is endless, but the JAR 25 regulations are an excellent source of information. Some of the items of equipment carried may seem to be of little use, but each has a specific purpose in some emergency or other. For example the large axe carried on passenger aircraft is to enable any trapped passengers and crew to cut their own way out of the cabin. Smoke hoods are to permit the cabin staff to help passengers leave the aircraft, even if the cabin is full of smoke. Portable oxygen is used in the cases of passengers feeling ill, in addition to the 'drop-out' masks, which activated if the cabin pressurisation has failed. Life jackets use a co2 cylinder to give rapid inflation once the passenger is outside the aircraft. The buoyancy of the jacket is then controlled by use of a mouthpiece to further inflate the jacket. The jackets are inspected at 6 monthly intervals for condition and inadvertent operation. The water-activated light is checked for insulation resistance by measuring across the terminals, which should be of at least 1 Megohm. Inadvertent operation of the light is checked by signs of chemical reaction and the co2 bottle is checked by weighing it on a laboratory scales.

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7.2 SEAT, HARNESSES AND BELTS All seat belts have to restrain the passenger (or crew) in their seat, even during a crash landing. The seat to which the belt is attached, has to hold securely in the seat rails, even during the high 'g' loadings experienced in an emergency landing. The seat rails are a continuous extrusion with circular cut-outs, which allow the seats to be attached and locked at different seat spacing, (pitch). The pitch is usually in, one inch or 25mm increments.

Seat Tracks Figure 1 Aircraft seats can be divided into three main groups; passenger seats, flight attendant seats and flight deck crew seats. Passenger seats are usually part of multiple units, although in first class and executive seating, some individual seat units can be found. Most passenger seats are manufactured from aluminium alloy tube, which is riveted and welded to form the frame with supporting legs and braces, individual reclining seat backs and integral tables. The seats and rails are all classed as primary structure. Flight Attendant seats are usually more utilitarian than passenger seats and can be mounted on seat tracks, the aircraft wall structure or, as in the ATR-72, to a sliding assembly that stows away without taking up passenger space, as shown below. They will all normally be fitted with a full harness seat belt, compared with the 'lap strap' assemblies for the passengers. The harnesses should only be cleaned with acid free soap and water. Inertia reel system will lock the harness if a rapid de-acceleration of the aircraft occurs. In the locked position, backward motion is still possible but forward motion is prevented.

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Attendant Seat Figure 2 The seats in the flight deck have to be the most comfortable on the aircraft, because it is laid down in many airline regulations that there must be a full crew in the cockpit, at all times. The crew must be as 'sharp' and attentive during the landing as they were at take-off many hours ago. Flight deck seats will have many different axes of movement such as height, reach, backrest tilt, lumbar support, arm rest height, etc. Most of the larger seats will have some of these movements powered by electrical actuators. These seats will also have at least a four point harness assembly and, in many cases these days, five point harnesses, with a lower crotch strap

Crew Seat Figure 3

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7.3 CABIN LAYOUTS The layout of the cabin is a compromise between the builder/designer, who would like it to contain as many paying passengers as possible, and the airworthiness authorities, who wish to limit the maximum number of passengers. This maximum has to be the number that can be evacuated from inside the cabin, through 50% of available exits, in 90 seconds. This ruling dictates the number and size of the exits, the width of the aisles and, most importantly, the number of seats. As can be seen from the diagrams below, the position of the exits varies with the design of the aircraft.

Seating And Emergency Exits Figure 4 The majority of passenger aircraft have seats in pairs or triple units with one or two aisles. The wide body Boeing 747 usually has two aisles with triple units outboard and a pair of double units between the two aisles, giving 10 abreast seating, the normal maximum.

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The remainder of passenger cabins are fairly standard with overhead stowages. Passenger service units (PSU) are located on the bottom of the overhead stowage lockers and normally contain reading lights, call buttons, seat belt and NO SMOKING warnings and, on aircraft that are equipped with them, drop-out oxygen masks. Galleys can be found in a variety of places in the cabin, at the front the rear, and occasionally, centrally, where they can be used to divide the different classes of passenger. They have their own power supply for heating, lighting and ventilation. For maintenance the galley units are removable, as are all other dividing partitions as well as the overhead units and PSUs. Galleys are also supplied with their own water supplies to permit the making of hot drinks, washing-up etc. This means they require connections to both fresh (potable) water and ‘grey’ (waste) water from the aircraft’s own systems. Some galleys are fitted in the under floor areas of larger aircraft, which necessitates the installation of lifts between floors. 7.4 CABIN FURNISHINGS As with galleys, all furnishings have to be easily removable, not only to allow the engineers access during deep maintenance, but also to permit various items to be changed at irregular intervals due to "fair wear and tear". This can include worn carpets, torn seat covers, cracked plastic cabin wall skins, ceiling panels and damaged overhead bin doors. All of the previous items are attached by 'quick release' fittings of varying types. Shown below are examples of an overhead bin, a wall panel and a ceiling panel.

Cabin Furnishings Figure 5

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7.5 CABIN ENTERTAINMENT Cabin entertainment varies greatly depending upon the aircraft type, (and age), as well as the airline operating the aircraft. It can vary from little more than 'music' played over the cabin P.A. system on smaller aircraft, through to the most common installations of films, navigation information and cabin safety briefings displayed on multiple television monitors located throughout the cabin. To reduce weight entertainment systems can utilise multiplexing to allow the different media to be transmitted down a single cable. Some modern aircraft have, fitted to their higher class seats, a complete 'entertainment experience', which can consist of individual viewing screens either attached to the seat back of the unit in front, or individually seat arm located. These screens can offer a multiple and individual video selection; computer games; musical videos with stereo sound on headphones and, in business class, access to a satellite telephone and other business tools. 7.6 GALLEY INSTALLATIONS Galleys, as has been mentioned earlier, have to be modular units so that they can be removed for maintenance. In the case of technical problems, it mayl also be necessary, some times to remove the units. Most galley units will have a supply of electricity and potable water and facilities for the disposal of 'grey' water overboard. As most modern catering operations use pre-prepared food, the standard sized food trolleys and containers are given stowage space in the galley units, which can then keep warm, heat up and chill both food and drinks as required. The illustrations show two typical galleys, with a selection of full and half sized trolley stowages, coffee makers and most of the facilities to provide a cabin meal and refreshment service.

Galley Installations Figure 6 Page 7-8

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7.7 CARGO HANDLING AND RETENTION EQUIPMENT In the majority of commercial aircraft, cargo is carried below the cabin floor, in dedicated fire resistant compartments that can be air conditioned if animals are being carried. There are a number of different variations to the above, dependent on the size of the aircraft, the type of passenger, the routes flown, etc. In the Fokker 100, for example, most of the underfloor space is for baggage, excluding the extreme front, which is for avionic equipment.

Under Floor Baggage Holds Figure 7 Smaller aircraft such as the Dornier 227 and the Fokker 50 have their cargo carried within the cabin space, the underfloor space being limited. Aircraft at the other end of the size spectrum, known as 'wide body' aircraft, can be produced as dedicated freighters such as certain Boeing 747 models. A more popular layout these days is the 'Combi freighter' which can carry both extra freight and passengers in the cabin, whilst still carrying cargo in the underfloor space. This type of aircraft is much more flexible on routes where the cargo/passenger ratios can vary through the week, the month or year. At times, there might be only 50 - 100 passengers on board whilst the remainder of the aircraft is carrying cargo. Containers are typically boxes shaped to the contour of the aircraft fuselage to maximise the capacity of the freight bay. They can be made from, alloy honeycomb of fibreglass. LD2

LD3

LD8

Baggage Containers Figure 8 B1 Mod 11.07 Issue No 0 30 - Jan 2003

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Most containers are sized by a code established by an international agreement enabling aircraft manufacturers to build the freight bays to a common size. The width of the cargo bay determines the size of the containers that can be loaded, but by setting the position of the various guides on the cargo bay floor, more than one size of container can be carried. Typical container sizes commonly used are shown above in various configurations. The identifying code letters indicate where the container has been designed to be loaded. Thus LD2 is a container of a standard size, designed for the lower deck (LD). Other sizes range from LD3 to LD8. Automatic Cargo Loading Systems Automatic cargo loading systems represent a major advance in the speed and efficiency that airfreight can turned round. Aircraft are designed to carry farepaying passengers above the cabin floor and a vast amount of cargo underneath. With the use of an automatic loading system, one man can load many tons of equipment, usually stored in purpose built containers, in the time it takes to refuel the aircraft and board the passengers. The Electro-mechanical loading system mechanism is normally built into the aircraft during manufacture. It consists of driver and sorter devices to load, store and manoeuvre the containers into the freight bay. The method of moving the containers is with the use of rubber-tyred rollers. These are in contact with the base of the containers and are motor-powered from the aircraft electrical system. Various guide rollers are used to steer the containers to the required location in the bay. Loading and unloading The containers are normally raised to the cargo bay floor level by a hydraulically operated deck, whose load area is covered with free running rollers or balls. The containers are then manually pushed into the door area, where they enter the freight bay of the aircraft and are supported by a ball mat or ball transfer panel. The ball mats/panels are low friction devices, which permit easy container movement. Each individual ball unit consists of a self-lubricated spring-loaded steel ball which itself rides on a series of smaller ball bearings in a cup-shaped housing. A wiper ring surrounding the ball prevents the ingress of dirt into the mechanism.

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Roller Ball Guide Figure 9 Once the load begins to enter the bay, the lateral rubber-covered drive rollers can be rotated to drive the containers fully into the bay. Lateral guides keep the container square and prevent the container from running off the edge of the transfer panel, as it is driven into the bay. Guide rollers around the doorframe ensure centralisation of the container and prevent doorframe damage. Sill rollers are mounted to the lower doorframe (sill) to provide a rolling surface into and out of the door area. With the container fully into the bay, longitudinal rollers can be rotated to propel the container down the length of the bay, once the relevant lateral retaining guides have been lowered. Roller trays are mounted down the entire length of the cargo bay, to continue on the similar low friction surface as the ball mats, for moving the containers. Centre or auxiliary guides ensure the container travels longitudinally and squarely, down to the extremities of the bay.

Freight Floors Figure 10 B1 Mod 11.07 Issue No 0 30 - Jan 2003

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During unloading, the roll-out-stops are locked down to permit free passage of the container out of the bay. This is done electrically or by pressing a foot pedal on each roll out stop if electrical power is not available. The Power Drive Units (PDU) consist of an electric motor driving a rubber-tyred roller. When commanded to rotate from the control panel, the roller is raised approximately 12mm from the floor level by a cam. Only when the roller is raised, will it begin to rotate and apply a moving friction force to the base of the container to propel it over the balls and rollers. Control Panel Each cargo bay will have a control panel with switches for; system power on/off, cargo bay lights, raising and lowering of the various lateral and longitudinal guides. A joystick control with eight positions and centre-off is provided for power drive unit operation.

A Typical Control Panel Figure 11

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Cargo Loading Operations The sequence for loading LD2 containers into the cargo compartment is as follows: 1.

Door fully open.

2.

Power switch ON.

3.

Select LD2 on panel.

4.

Set roll out stops to load position.

5.

Set lateral guides to NORMAL.

6.

PDU switch to A-B ON.

7.

PDU switch on the on the other panel to FWD ON.

8.

Switch the lights on if required.

9.

Place the container level with the doorframe of the compartment.

The joystick is then used to control the PDU’s and move the container into the desired bay. As the container clears each roller they spring up to prevent it rolling back out of the bay. At the end of the containers’ travel it contacts the fixed loading stops and with all the containers loaded the roll out stops are positioned into the locked up position, holding the load firmly. Unloading. Essentially unloading is the reverse of the loading procedure, with the exception of locking the roll out stops to their retracted position and positioning all of the centre/auxiliary guides to the down position. If any guide has been manually locked down, ensure they are unlocked before electric power is applied to prevent damage to the motors. At the first sign of any container stopping or failing to move release the joy stick to the centre/off position and investigate the cause of the jam. Only approved personnel, who have received proper training in the particular installation or layout, should carry out the operation of automatic loading systems.

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Dangerous Goods Dangerous goods are those which possess potentially hazardous characteristics. However, as long as suitable precautions are taken these goods are not necessarily prohibited from air travel. They include obvious items such as; acids, explosives and radio-active materials and also some unlikely items such as magnets, breathing apparatus and other gas cylinders and instruments that contain mercury. 7.8 CARGO RETENTION EQUIPMENT Once cargo is loaded into the aircraft, it must be restrained to prevent movement, during take- off, in turbulent flight and landing, (especially hard braking). The LD containers have positive latches, which attach the containers directly to the aircraft structure. 'Loose' baggage in cargo holds are usually restrained by nets, which can be locked into the floor or the walls of the bay. This system can also be used on pallets, where cases and bags are, again, preloaded and then covered by waterproof sheet and restraint netting. Once loaded, the pallets are clamped down on to the cargo bay floor.

Baggage Hold-Down Figure 12

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7.9 AIRSTAIRS The term airstairs is usually used to describe passenger steps that are integral to the aircraft structure, meaning that it is independent of normal passenger steps and of jetways at large airports. They are often fitted to aircraft that will be operated into poorly equipped airports on a normal, day-to-day operation. Airstairs can be manually or power operated and can be as simple as a set of stairs set into the back of the entrance door or on larger aircraft, a fully powered, folding set of steps that are extended and retracted by the operation of push buttons.

Airstairs Figure 13

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The first example shown is from the ATR-72 turbo-propeller aircraft. This unit is mechanically operated and counterbalanced by a pair of large springs. As can be seen from the drawing, there are handrails, one of which can be folded, if required. The second example, (lower left), is an electrically powered airstair fitted to the new Boeing 717-200. This aircraft can also be fitted with a second airstair at the rear of the cabin, (lower right), which will allow the passengers to embark and disembark through two doors simultaneously. This will speed up the turn around maintenance.

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MODULE 11.08 FIRE PROTECTION

CONTENTS 8

FIRE PROTECTION...................................................................... 8-1 8.1

8.2

8.3

8.4

8.5

FIRE/OVERHEAT DETECTION AND WARNING .................................. 8-1 8.1.1 Unit (Spot) Type ............................................................ 8-2 8.1.2 Continuous Loop (Fire Wire) Detectors ......................... 8-3 8.1.3 Dual Loop System ......................................................... 8-4 8.1.4 Pressure-Type Sensor .................................................. 8-5 FIRE ZONES ................................................................................. 8-5 8.2.1 Hot And Cool Zones ...................................................... 8-5 8.2.2 Fireproof Bulkheads ...................................................... 8-6 8.2.3 Engine Fire Prevention .................................................. 8-6 8.2.4 Cockpit and Cabin Interiors. .......................................... 8-7 SMOKE DETECTION ...................................................................... 8-7 8.3.1 Carbon Monoxide Detectors .......................................... 8-7 8.3.2 Photoelectric Smoke Detectors. .................................... 8-7 8.3.3 Ionisation Type Smoke Detector. .................................. 8-8 FIRE EXTINGUISHING .................................................................... 8-9 8.4.1 Extinguishing System .................................................... 8-10 8.4.2 Directional Flow Control Valves (2 Way Valves) ............ 8-12 8.4.3 Fire Extinguishant Container ......................................... 8-12 8.4.4 Toilet Compartment Systems ........................................ 8-13 8.4.5 Warnings And Indications.............................................. 8-13 8.4.6 Hand Held (Portable) Fire Extinguishers ....................... 8-14 SYSTEM TESTS ............................................................................ 8-14 8.5.1 Fire System Test Switch................................................ 8-15 8.5.2 Fire Wire Loop Test....................................................... 8-15 8.5.3 Squib-Test..................................................................... 8-15

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JAR 66 CATEGORY B1 ) MODULE 11.08 FIRE PROTECTION


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8

FIRE PROTECTION

Fire is the most dangerous threat to the safety of an aircraft and is associated with external areas near the main engines and the APU. Other external hot spots are landing gear bays, where heat from brake units could affect the surrounding equipment and wiring, when the gears are retracted. Overheating of the structure, equipment and wiring from very hot air, leaking engine compressor bleed air pipes, must also be catered for. Fire from internal areas such as the passenger, flight deck and toilet compartments as well as cargo, air-conditioning and electrical/electronic equipment bays require protection too. Indeed any source on an aircraft that the manufacturer or operator considers a likely hazard will be protected. Ideally, a fire protection system will include as many as possible of the following features:  Rapid warning of fire/overheat and its accurate location  Must not cause false warnings  Continuous warning for duration of fire/overheat  Confirmation that the fire has been extinguishing  Indication that the fire has re-ignited  A means of testing the system from the flight deck  Detectors that are proof against oil, water, vibration and high temperatures  Detectors that are easily accessible throughout the aircraft  Detectors and extinguishers hot wired electrically or powered from emergency electrical buses  Adequate visual and aural indication on the flight deck and vital areas on the aircraft  Separate warnings for each engine and specific areas as determined by the aircraft manufacturers Therefore, the Fire (and Overheat) Protection system will normally be split into two main subsystems:  Fire/Overheat Detection and Warning  Fire Extinguishing 8.1 FIRE/OVERHEAT DETECTION AND WARNING Fire/Overheat detectors can be divided into two main groups:  

Unit or Spot Type Continuous Loop (Firewire) Type


Page 8-1


8.1.1

Unit (Spot) Type

This type is fitted at various strategic points within the fire/overheat zone and takes the form of a thermally activated switch. They are electrically connected in parallel with each other and in series with the audio/visual warning system. This arrangement allows any switch to operate the warning, even if other switches have failed in the remainder of the system. Some Unit detectors may have a pair of BI-metallic contacts, that close when heated and open when they are cooled down, to make or break the electrical warning circuit. However, the majority has a thin casing that surrounds two conventional electrical contacts that are normally set apart from each other. When subjected to heat, the casing expands and pulls the two contacts together, completing the warning circuit in a similar manner to the BI-metallic type. The main advantage of this so-called „High Speed Resetting Switch‟ (HSRS), is its sensitivity and fast reaction time, to initiate the warning and cancel it once the heat is removed. Spot Detectors are used mainly to detect high temperature leaks from bleed air ducts and are normally positioned at pipe to pipe connections.

Thermoswitch Type Fire Detection System Figure 1

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Thermoswitch Spot Detector Figure 2

8.1.2 Continuous Loop (Fire Wire) Detectors This method permits more complete coverage of a fire hazard area than any type of spot-type of temperature detectors. The continuous loop uses the principle of capacitance and resistance to indicate a rise in temperature at any point along the length of the detector loop. The commonest type has a stainless steel or Inconel outer tube, an inner pure nickel wire surrounded by ceramic beads wetted by eutectic salt. The effect of this design is that a rise in temperature causes a sharp fall in electrical resistance, as well as a rise in capacitance. Once the detection unit senses this effect, anywhere along the wire, it will cause an overheat warning to be generated. This continuous loop system is often referred to as a 'firewire' system. The advantage of a firewire system is that a loop can cover the complete powerplant, (Figure 3) within its cowling so that an overheat or fire will be detected quickly no matter where it starts. The firewire will also re-set the control box to remove the warning when the temperature falls below the limit temperature.


Page 8-3


Fire Wire Layout Figure 3 Firewire elements are attached to the airframe structure with quick release clips approximately 6” apart and 4” from the end fittings. The element is supported in clips with a rubber grommet to prevent rubbing and to help damp out vibrations. (Figure 4). Care is taken to eliminate strain on the element as excessive bending could result in work hardening of the capillary.

Fire Wire Clips and Connections Figure 4 8.1.3 Dual Loop System Most aircraft use the dual loop system of indication. Each sensing circuit has dual sensing loops. Each Loop A and Loop B is independent of each other. When the loop selector switch is set to BOTH, both loops must detect a fire condition before the warning system is activated. If only one loop detects a fire, the associated loop fault light will illuminate.

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If the selector is switched to a single loop (A or B) full fire warnings will activate if the selected loop senses a fire condition. Pressing the loop test button simulates a fire condition on the respective loop. This is done by earthing the inner electrode of the loop that functionally checks the system and checks the continuity of the loop. 8.1.4 Pressure-Type Sensor The pressure type detection system uses a continuous loop for the detection element. This loop is made from sealed stainless steel tube that contains an element that absorbs gas when it is cold but releases the gas when it is heated. This tube is connected to a pressure switch that will close when the pressure reaches a pre-determined level. The commonest make of this type of system is the Systron-Donner system which uses a centre titanium centre wire and the expansion of both helium and hydrogen gas to give the two-stage warnings. Whilst the firewire system actuates when any part of the loop reaches the limit temperature, the pressure type system will actuate in two different ways. If a localised fire occurs, the hydrogen gas is released and its pressure closes the pressure switch which will set off the warning system, however, if the temperature over a larger area rises to a lower level than a fire warning the helium expands and closes the pressure switch to activate the system warning. 8.2 FIRE ZONES On light aircraft, the only protection against fire is a stainless steel or titanium bulkhead (firewall), dividing the engine bay from the cabin and the rest of the aircraft. Larger aircraft have the complete engine cowlings isolated from the airframe/wing assemblies and, in addition, aircraft cowlings can be divided into a number of 'fire zones', each one usually having its own warning and extinguishing system. The types of zone dictate what type of protection that they receive, for example, light aircraft have piston engines and hence, due to the high flow of air through the bay, have no fire protection and depend on isolating the engine of fuel to put out any fire. The example has four zones around the engine that only two have firewires and extinguishing. 8.2.1 Hot And Cool Zones Engines are usually split into hot and cool zones (Figure 5). The hot zone comprises the combustion chamber turbines and exhaust areas, the cool zone comprises the intake, compressors and accessory drives.


Page 8-5


Engine Fire Zones Figure 5 8.2.2 Fireproof Bulkheads These prevent fire from spreading to other areas. Auxiliary power units and tail mounted engines are normally contained within such bulkhead compartments separating them from the rest of the airframe. The engine pylons also contain a firewall to separate the engine from the wing. These are made from titanium or stainless steel and all joints are sealed with fireproof sealants 8.2.3 Engine Fire Prevention There are a number of techniques used to help prevent a fire occurring around engines. These are, the use of flameproof or flame resistant materials, use of bonding strips to prevent arcing, drainage of spilt fuel/oil and efficient cooling. All pipes which carry fuel, oil or hydraulic fluids are made fire resistant and all electrical components and connections are made flame proof. It is essential that a fire staring in any zone is contained within that zone and is not allowed to spread to any other part of the aircraft. The engine cowlings form a natural container but they are usually made from light alloy and would not contain a ground fire for long. In flight however cooling airflow‟s through the cowlings, provide sufficient cooling to render the cowlings fireproof. The fireproof bulkheads and any cowling that has no cooling airflow are usually made from titanium or stainless steel.

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8.2.4

Cockpit and Cabin Interiors.

All wool, cotton and synthetic fabrics used in interior trim are treated to render them flame resistant. Tests conducted have shown that whilst the foam used in seat cushions is flammable, if covered with a flame-resistant fabric, there is little danger of fire from accidental contact with a cigarette, for example. Fire protection for the aircraft interior is usually provided by hand-held extinguishers. Various types are available including, Water, CO 2 and Dry Powder. Each type is best used on one kind of fire but may be used on other kinds. It is best to be sure which is safe to use on which type of fire. 8.3 SMOKE DETECTION A smoke detection system monitors certain areas of the aircraft for the presence of smoke, which is could be indicative of a fire condition. These may include cargo and baggage compartments and the toilets of transport category aircraft. A smoke detection system is used where the type of fire anticipated is expected to generate a substantial amount of smoke before temperature changes are sufficient to actuate a heat/fire detection system. 8.3.1 Carbon Monoxide Detectors The presence of Carbon Monoxide (CO), or Nitrous Oxides (N2O), is dangerous to flight crew and passengers alike and may indicate a fire condition as it is a byproduct of combustion. Detection of the presence of either or both of these gases could be the earliest warning of a possible dangerous situation. Carbon Monoxide is very dangerous, firstly due to the minute amount required to cause loss of attention and headaches; (this is approximately 2 parts in 10,000). It is colourless, odourless, tasteless and a non-irritant. Carbon Monoxide detectors are usually used in cabin and cockpit areas. The detector is usually a small card with a transparent pocket containing silica gel crystals that have been treated with a chemical, which changes colour to green or black when they are exposed to carbon monoxide.

8.3.2 Photoelectric Smoke Detectors. Air from the monitored compartment is drawn through the detector chamber and a light beam is shone on it. A photoelectric cell installed in the chamber senses the light that is refracted by the smoke particles. The photocell is installed in a bridge circuit that measures any changes, in the amount of current that it conducts. Figure 6 shows a typical photoelectric smoke detector.


Page 8-7


Air inlet Light beam

Light source

Photoelectric cell Light reflected from smoke into photocell Air outlet

Photo Electric Smoke Detector Figure 6 When there is no smoke in the chamber air, no light is refracted and the photocell produces a reference current. When smoke is in the chamber air, some of the light is refracted and sensed by the photocell. Its conductivity changes, changing the amount of current. These changes in current are amplified and used to initiate a smoke warning signal. 8.3.3 Ionisation Type Smoke Detector. A small amount of radioactive material is mounted on the side of the detector chamber. This material bombards the oxygen and nitrogen molecules in the air flowing through the chamber and ionises it to the extent that a reference current can flow across the chamber through the ionised gas to an external circuit.

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_

Ionzing beam

Radioactive material

Air inlet

Air outlet

+ Target Ionisation Type Smoke Detector Figure 7 Smoke flowing through the chamber changes the level of ionisation and decreases the current. When the current reduces to a specific level the external circuit initiates a smoke warning signal. Figure 7 shows a typical ionisation smoke detector. 

Flame Detectors

This system uses a photoelectric cell to detect a sharp rise in light, such as that from a flame in a closed bay. 8.4 FIRE EXTINGUISHING There are a variety of aircraft and ramp extinguishing agents. Their use depends upon several variables such as location, proximity to personnel, environment, possible sources of fire, etc. There are integral extinguishing systems on board the aircraft as well as hand held extinguishers


Page 8-9


8.4.1 Extinguishing System Aircraft that have an integral fire extinguisher system have a system similar to the arrangement shown in Figure 8. There are a number of pressurised bottles with extinguishant inside and each bottle has two explosive cartridges, (squibs), which can be fired from the flight deck. Each bottle can feed either the port or starboard engines through a crossfeed. The extinguishant is fed through a series of pipelines and valves to the outlet nozzles and tubes. In some aircraft, fixed systems may also be provided for the protection of landing gear wheel bays and baggage compartments. These systems may be independent of each other. They may be fully automatic or require the aircrew to initiate them when a fire is indicated.

Basic Aircraft Extinguishing System Figure 8 On multi-engine aircraft there may be one extinguisher bottle provided for each engine or one bottle may feed 2 engines (Figure 9). There is always usually a facility for cross feeding to another engine should the need arise.

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Dual Container System Figure 9 Two bottles giving either two 'shots', to a single engine or, one 'shot' each to either engine (Figure 10). The bottle condition is indicated either through a pressure gauge on each bottle, or a red/green sectioned gauge showing red when the bottle is empty or its pressure is low as well as a discharge indication on the associated fire control panel I the cockpit.

Typical 2 Shot System Figure 10


Page 8-11


There may also be pop up indicators to indicate that the squib has been fired. A pressure switch may also be fitted which gives an electrical indication to the cockpit control panel when the pressure drops to a pre-determined level. Each bottle will have protection against overpressure using a 'rupture disc', which fails if the bottle pressure becomes excessive due to overheating. 8.4.2 Directional Flow Control Valves (2 Way Valves) These valves are non-return valves designed for use in a crossfeed system to allow the contents of one or several extinguishers to be directed into any one engine (or compartment). The valves prevent the reverse flow of the extinguishant into the other bottle or engine. 8.4.3 Fire Extinguishant Container Figure 11 shows a typical extinguishant container. The cartridge is electrically ignited which drives the cartridge cutter into the disc that on rupture releases the extinguishant. The strainer prevents any of the broken disc from entering the distribution system.

Fire Extinguisher Bottle Figure 11

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The safety plug is connected by a pipeline to a red indicator disc on the outside of the compartment. If the gas pressure increases due to an increase in the compartment temperature that the bottle is located in, the fusible safety plug melts at a pre-determined temperature and the bottle contents are discharged overboard. As the bottle discharges overboard, it blows out the red indicator. The gauge shows the pressure of the extinguishant in the container. 8.4.4 Toilet Compartment Systems Small, automatic units will often be found in the toilet waste bins, where they will discharge themselves when a heat source is sensed in the region of 75 degrees centigrade. A fusible type plug will melt allowing the contents to discharge. Most aircraft with this system fitted do not generate any indications to the cockpit or attendants panel if the system was activated. Some systems have a visible temperature strip that can be checked before each flight, or by the cabin crew in flight. 8.4.5 Warnings And Indications Once a fire has been detected in the engine bay (or compartment being sensed), a signal is generated by the firewire element and this signal is sent to a control unit. The control unit processes the signal and sends a signal to the cockpit CWP, associated power lever handle, and the fire control panel. The CWP red Fire warning caption light illuminates for the affected engine (or compartment) as well as the master warning lights and audio warnings. The Affected power lever handle and fire extinguisher handle on the overhead console also illuminate red. To activate the extinguishant, the red fire handle is pulled to arm the system and then the squib button is pressed to fire the bottle. If after the bottle contents have exhausted and the fire indication remains, the second squib button is pressed to fire the contents of the other bottle into the same affected engine (or compartment). Some aircraft activate the extinguishers differently. The bottle may be fired by pressing the affected fire button on the fire panel. If the fire remains a cross feed switch is activated which opens a crossfeed valve and the same fire button is repressed to fire the other bottles contents into the same affected system. Once discharged an amber DISCH caption on the fire control panel will indicate when the corresponding bottle is empty. These captions are usually electrically activated Whatever the method of operation of the extinguisher system, the same basic principle applies. The contents of each bottle can be cross fed into the affected area that is on fire.


Page 8-13


8.4.6 Hand Held (Portable) Fire Extinguishers Each aircraft must carry portable fire extinguishers for use by the cabin crew in case of a fire. These are positioned in various places within the cabin with easy access to the crew. The amount and location depends on the type of aircraft and its size. Halon extinguishers contain a gas that interrupts the chemical reaction that takes place when fuels burn. These types of extinguishers are often used to protect valuable electrical equipment since they leave no residue to clean up. Halon extinguishers have a limited range, usually 4 to 6 feet. The initial application of Halon should be made at the base of the fire, even after the flames have been extinguished Carbon Dioxide fire extinguishers disperses the gas quickly, these extinguishers are only effective from 3 to 8 feet. The carbon dioxide is stored as a compressed liquid in the extinguisher; as it expands, it cools the surrounding air. The cooling will often cause ice to form around the “horn” where the gas is expelled from the extinguisher. They are primarily used to extinguish electrical fires in the cabin and cockpit. The CO2 can be aimed at the fire and discharged using a trigger. A dry powder fire extinguisher use compressed nitrogen to expel a dry powder such as sodium bicarbonate or potassium bicarbonate. They can be used on most fires but should never be used on the flight deck, due to lack of visibility and interference with some electrical equipment caused by the powder. Water extinguishers are also fitted to some aircraft and should be used to put out fires in ordinary combustibles, such as wood and paper. The hand held extinguishers are subject to periodic maintenance. The extinguisher is checked for its weight. This is stamped on the neck of the bottle and indicates its charged weight. If the weight is below the set limits, it is to be replaced. 8.5 SYSTEM TESTS All extinguishing systems have a method of testing their serviceability. This can vary from weighing the complete cylinder off-aircraft, (which will have the correct 'full' weight marked on it), through to the bottle having a gauge with safe and lowpressure sectors marked on it. Figure 12 shows an engine extinguisher with a fitted gauge. Other more sophisticated systems have internal pressure switches fitted to the bottle, which will notify the flight deck of the loss of bottle pressure, (or discharge), via a warning light, magnetic indicator etc.

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Regardless of the system, all bottles and squibs have a life, after which they have to be removed and returned to the manufacturer for maintenance.

Fire Bottle With Pressure Gauge Figure 12 8.5.1

Fire System Test Switch

A test switch is available for each system. When pressed all warning lights and audio warnings are checked. If a light fails to illuminate it will normally indicate a bulb filament failure. 8.5.2

Fire Wire Loop Test

A test switch on the cockpit fire panel is available to test each sensing element loop. When selected the continuity of each circuit is checked. If the system is serviceable, the Loop caption(s) will illuminate. If the caption(s) do not illuminate there is a fault in the system. 8.5.3

Squib-Test.

A squib test button is available to check the continuity of the discharge heads for each of the fire extinguisher bottles. When pressed a squib warning light or magnetic indicator will illuminate if the system is serviceable. No illumination means that there is a fault in the system. The current used during the squib test is at a much lower value than that required to fire the squib.


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MODULE 11.09 AERODYNAMICS, STRUCTURES AND SYSTEMS

Contents 9

FLYING CONTROLS .................................................................... 9-1 9.1

PRIMARY FLIGHT CONTROLS ................................................... 9-1 9.1.1 Ailerons ......................................................................... 9-3 9.1.2 Elevators ....................................................................... 9-3 9.1.3 Rudders ........................................................................ 9-4 9.1.4 Spoilers ......................................................................... 9-5 TRIM CONTROLS ........................................................................... 9-7 9.2.1 Fixed and Adjustable Trim Tabs .................................... 9-7 9.2.1.2 Controllable Trim Tabs .................................................. 9-7 9.2.1.3 Servo Tabs.................................................................... 9-8 9.2.2 Balance Tabs ................................................................ 9-8 9.2.3 Anti-Balance Tabs ......................................................... 9-9 9.2.4 Spring Tabs................................................................... 9-9 FULLY POWERED FLYING CONTROL TRIM SYSTEM ........................... 9-10 9.3.1 Typical Trim System ...................................................... 9-10 9.3.2 Rudder Trim System ..................................................... 9-10 9.3.3 Aileron Trim System ...................................................... 9-10 9.3.4 Tailplane Trim System .................................................. 9-11 ACTIVE LOAD CONTROLS ............................................................... 9-15 9.4.1 Active Load Control ....................................................... 9-15 9.4.2 Active Control Technology ............................................ 9-15 9.4.3 Advantages of Active Control Technology ..................... 9-17 9.4.4 Direct Lift Force ............................................................. 9-17 9.4.5 Direct Side Force .......................................................... 9-17 HIGH LIFT DEVICES ....................................................................... 9-18 9.5.1 Flaps ............................................................................. 9-18 9.5.2 Slats .............................................................................. 9-20 9.5.3 Drooped Leading Edges................................................ 9-21 9.5.4 Krueger Flaps ............................................................... 9-21 LIFT DUMP AND SPEED BRAKES .................................................... 9-22 9.6.1 Lift Dumpers.................................................................. 9-22 9.6.2 Speed Brakes ............................................................... 9-22 SYSTEM OPERATION ..................................................................... 9-24 9.7.1 Manual Operation.......................................................... 9-24 9.7.2 Powered Flight Controls (P.F.C.U‟s).............................. 9-24 9.7.3 Proportionality ............................................................... 9-25 9.7.4 Redundancy of hydraulic Supplies ................................ 9-25 9.7.5 Tandem PFCU .............................................................. 9-25 9.7.6 Dual Assembly PFCU‟s ................................................. 9-27 9.7.7 Duplicate/Triplicate PFCU's........................................... 9-28 9.7.8 Duplicated Control Surfaces ......................................... 9-29 9.7.8 Self Contained PFCU .................................................... 9-30 9.7.9 Input Systems ............................................................... 9-30

9.2

9.3

9.4

9.5

9.6

9.7

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AERODYNAMICS, STRUCTURES AND SYSTEMS

9.7.10 High Speed Primary Controls ........................................ 9-31 TRAILING EDGE FLAP CONTROLS .................................................. 9-33 9.8.1 Flap Control Utilising Linear Hydraulic Actuators ........... 9-33 9.8.2 General ......................................................................... 9-34 9.8.3 Hydraulic Power ............................................................ 9-35 9.8.4 Control Input Circuit ...................................................... 9-35 9.8.5 System Operation ......................................................... 9-36 9.8.6 Safety Aspects .............................................................. 9-37 9.8.7 Position Indication ......................................................... 9-38 9.8.8 Flap System - Hydraulic Motors and Torque Tube Drive 9-38 9.8.9 Maintenance of Flap Systems ....................................... 9-41 LEADING EDGE FLAP CONTROLS ................................................... 9-41 9.9.1 Leading Edge Flap Pneumatic drive Unit ...................... 9-45 9.9.2 Krueger Flap Drive Components ................................... 9-48 SPEED BRAKE/GROUND SPOILER CONTROL ................................... 9-49 9.10.1 Operation ...................................................................... 9-50 MECHANICAL & ELECTRICAL FLIGHT CONTROL SYSTEM .................. 9-53 9.11.1 Mechanical Controls...................................................... 9-53 9.11.2 Electrical Flight Controls................................................ 9-54 „Q‟ FEEL, YAW DAMPER, MACH TRIM, RUDDER LIMITER, GUST LOCKS 9-58 9.12.1 Artificial Feel ................................................................. 9-58 9.12.2 Operation ...................................................................... 9-58 9.12.3 Hydraulic „Q‟ Feel System ............................................. 9-60 9.12.4 Mach Number Correction .............................................. 9-60 9.12.5 Operation ...................................................................... 9-60 YAW DAMPING ............................................................................. 9-62 9.13.1 Yaw Control .................................................................. 9-62 MACH TRIM .................................................................................. 9-63 9.14.1 Typical System .............................................................. 9-65 9.14.2 Controller ...................................................................... 9-65 9.14.3 Mach Trim Actuator ....................................................... 9-65 9.14.4 Operation ...................................................................... 9-65 RUDDER LIMITING ......................................................................... 9-67 9.15.1 „Q‟ Limiter ...................................................................... 9-67 GUST LOCKS ................................................................................ 9-67 9.16.1 Description .................................................................... 9-67 9.16.2 Controls locking mechanism (aileron and elevator) ....... 9-68 9.16.3 Controls locking mechanism (rudder) ............................ 9-69 9.16.4 Power Supplies ............................................................. 9-71 9.16.5 Operation ...................................................................... 9-71 RIGGING AND BALANCING CONTROLS .......................................... 9-72 9.17.1 Rigging - Introduction .................................................... 9-72 9.17.2 Checks Before Rigging ................................................. 9-72 9.17.3 Rigging Procedure ........................................................ 9-73 9.17.4 Control Surface Setting Gauges .................................... 9-75 9.17.5 Checking for Sense of Movement ................................. 9-75 9.17.6 Checking for Static and Running Friction ...................... 9-77

9.8

9.9

9.10 9.11

9.12

9.13 9.14

9.15 9.16

9.17

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9.17.7 Checks After Rigging .................................................... 9-77 9.17.8 Duplicate Checks .......................................................... 9-78 9.17.9 Primary Control Systems - Example of Rigging ............. 9-78 9.17.10 Rigging a Tube-operated Control System ..................... 9-79 9.17.11 Rigging a Powered Flying Control System .................... 9-80 9.17.12 Rigging of Trimming Tab System .................................. 9-82 STALL W ARNING AND PROTECTION ................................................ 9-83 9.18.1 Stall Warning Systems .................................................. 9-83 9.18.1.1 Pneumatic Stall Warning System .................................. 9-83 9.18.2 Stall Protection System ................................................. 9-85 9.18.3 Typical System Components ......................................... 9-85 9.18.4 Actual Stall Protection System ...................................... 9-86 9.18.5 Incidence Probes .......................................................... 9-86 9.18.6 Nitrogen System ........................................................... 9-87 9.18.7 Automatic Ignition.......................................................... 9-88 9.18.8 Stall Warning................................................................. 9-88 9.18.9 Stall Identification .......................................................... 9-89 FLY BY W IRE................................................................................ 9-92 9.19.1 Introduction ................................................................... 9-92 9.19.2 Principles of FBW.......................................................... 9-92 9.19.3 Principles of FBOW ....................................................... 9-92 9.19.4 Advantages of FBOW over FBW ................................... 9-92 9.19.5 Other Inputs to Powered Flying Control Unit ................. 9-93 9.19.6 777 Flight Controls - Introduction .................................. 9-93 9.19.7 General ......................................................................... 9-93 9.19.8 777 Primary Flight Control System ................................ 9-93 9.19.9 High Lift Control System................................................ 9-94 9.19.10 Benefits of the Fly-By-Wire System ............................... 9-94 9.19.11 Abbreviations and Acronyms ......................................... 9-94 9.19.12 Primary Flight Control System - Introduction ................. 9-96 9.19.13 PFCS – General Description ......................................... 9-97 9.19.14 Manual Operation.......................................................... 9-97 9.19.15 Autopilot Operation ....................................................... 9-98 9.19.16 PFCS Modes of Operation ............................................ 9-98 9.19.17 Flight Deck Controls ...................................................... 9-98 9.19.18 Main Equipment Centre................................................. 9-99 9.19.19 PFCS – Flight Controls ARINC 629 BUS Interfaces ...... 9-99

9.18

9.19

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FLYING CONTROLS

9.1 PRIMARY FLIGHT CONTROLS Aircraft theory of flight has already been discussed in Module 11.1. We shall now look at how the Aircraft are equipped with moveable aerofoil surfaces that provide control in flight. Controls are normally divided into Primary and Secondary controls. The primary flight controls are:  Ailerons  Elevators  Rudders  Spoilers Because of the need of aircraft to operate over extremely wide speed ranges and weights, it is necessary to have other secondary or auxiliary controls. These consist of:  Trim controls  High Lift Devices  Speed Brakes and Lift Dump

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Note: There is some variation of opinion as to whether spoilers are considered to be primary controls. The EASA 66 syllabus includes them as primary controls, so that is how these notes will define them. Both types of controls are illustrated in the following diagram.

Typical Aircraft Flight Controls Figure 1

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9.1.1 AILERONS

Ailerons are primary flight controls that provide lateral roll control of the aircraft. They control aircraft movement about the longitudinal axis. Ailerons are normally mounted on the trailing edge of the wing near to the wing tip.

Inboard and Outboard Ailerons Figure 2 Some large turbine aircraft employ two sets of ailerons. One set are in the conventional position near the wingtip, the other set is in the mid-wing position or outboard of the flaps. At low speeds both sets of ailerons operate to give maximum control. At higher speeds hydraulic isolate valves will cut power to the outer ailerons so that only the inboard ailerons operate. If the outer ailerons are operated at high speeds, the stress on the wing tips may twist the leading edge of the wing downwards and produce “aileron reversal”. 9.1.2 ELEVATORS

Elevators are primary flight controls that control the movement of the aircraft about the lateral axis (pitch). Elevators are normally attached to hinges on the rear spar of the horizontal stabiliser. Fig 11.1 shows the typical location for elevators.

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9.1.3 RUDDERS

The rudder is the flight control surface that controls aircraft movement about the vertical or normal axis. Rudders for small aircraft are normally single structural units operated by a single control system. Rudders for larger transport aircraft vary in basic structural and operational design. They may comprise two or more operational segments; each controlled by different operating systems to provide a level of redundancy.

Rudder Figure 3

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9.1.4 SPOILERS

Spoilers are secondary control surfaces used to reduce or spoil the lift on a wing. They normally consist of multiple flat panels located on the upper surface of the wings. The diagram below shows the more common configuration.

Operation of Spoilers on a Typical Aircraft Figure 4 The spoilers lay flush with the upper surface of the wing and are hinged at the forward edge. When the spoilers are operated, the surface raises and reduces the lift. The spoilers may be used for different purposes.

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9.1.4.1 Flight Spoilers

Flight Spoilers are used in flight to reduce the amount of lift. If the pilot operates the controls left or right to roll the aircraft, the spoilers on the down-going wing move upward to aid rolling the aircraft. The movement of the spoilers is in proportion to the rate of roll required. On some aircraft, the spoilers are the primary flight control for rolling. If operating only as flight spoilers, only the surfaces on one wing will be raised at any one time. The flight spoilers are normally positioned outboard of the ground spoilers. 9.1.4.2 Ground Spoilers

Ground Spoilers are only used when the aircraft is on the ground. They operate with the flight spoilers to greatly reduce the lift on landing. The also reduce the drag after landing to slow down the aircraft. Ground spoilers will normally be deflected to their maximum position to give maximum drag on landing.

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TRIM CONTROLS

The majority of aircraft at some time during a flight develop a tendency to deviate from a straight and level attitude. This may be caused by a fuel state change, a speed change, a change in position of the aircraft's load, or in flap and undercarriage positions. The pilot can counter this tendency by continuously applying a correcting force to the controls - an operation, which, if maintained for any length of time, would be both fatiguing and difficult to maintain. The tendency to deviate is therefore corrected by making minor trim adjustments to the control surfaces. Once an aircraft has been trimmed back to a 'balanced' flight condition, no further effort is required by the pilot until further deviation develops. 9.2.1 FIXED AND ADJUSTABLE TRIM TABS 9.2.1.1

Fixed Trim Tabs

A fixed trim tab is normally a piece of sheet metal attached to the trailing edge of a control surface. It is adjusted on the ground by bending to an appropriate position that give zero control forces when in the cruise. Finding the correct position is by trial and error. 9.2.1.2

CONTROLLABLE TRIM TABS

Controllable Trim Tab Figure 5 A controllable trim tab is adjusted by mechanical means from the flight deck, usually with an indication of its position being displayed to the pilot. Most aircraft have trim on the pitch control and more advanced aircraft have trim on all three axes. Whilst the controls in the cockpit are by lever, switch etc., the actuation can be by mechanical, electrical or hydraulic means.

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SERVO TABS

Servo Tab Figure 6 Sometimes referred to as the flight tabs, the servo tabs are used primarily on large control surfaces, often found on larger, older aircraft. This tab is operated directly by the primary controls of the aircraft. In response to the pilot's input, only the tab moves. The force of the airflow on the servo tab then moves the primary control surface. This tab is used to reduce the effort required to move the controls on a large aircraft.

9.2.2 BALANCE TABS

Balance Tab Figure 7

A balance tab is linked to the aircraft in such a manner that a movement of the main control surface will give an opposite movement to the tab. Thus the balance tab will help in moving the main surface, therefore reducing the effort required. This type of tab will normally be found fitted to aircraft where the controls are found to be rather heavy during initial flight-testing.

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9.2.3 ANTI-BALANCE TABS

These anti-balance tabs operate in the same way, mechanically, as balance tabs. The tab itself is connected to the operating mechanism so that it operates in the reverse way to the balance tab. The effect this has is to add a loading to the pilot‟s pitch control, making it appear heavier. These tabs can often be found fitted to „stabilators‟, which are very powerful and need extra „feel‟ to prevent the pilot over-stressing the airframe. 9.2.4 SPRING TABS

The spring tabs, like some servo tabs, are usually found on large aircraft that require considerable force to move a control surface. The purpose of the spring tab is to provide a boost, thereby aiding the movement of a control surface. Although similar to servo tabs, spring tabs are progressive in their operation so that there is little assistance at slow speeds but much assistance at high speeds.

Spring Tab Figure 8

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9.3 FULLY POWERED FLYING CONTROL TRIM SYSTEM As fully powered flying controls are irreversible, i.e. all loads (reactions) are fed via mountings to structure; trim tabs would be ineffective. To overcome this, electric trim struts or actuators are used within the input system. These actuators commonly reposition the "null" position of a selfcentring spring device to hold the control-input system in a new neutral position. Thus the main control surface will be held deflected and the aircraft trimmed.

9.3.1 TYPICAL TRIM SYSTEM

The following is a typical trim system as used on a fully powered flight control system.

9.3.2 RUDDER TRIM SYSTEM

In a typical rudder trim system for a powered system, trim commands from the trim switch causes an actuator to extend or retract, which rotates the feel and centring mechanism. This provides a new zero force pedal position corresponding to the trimmed rudder position. The trim switch is spring loaded to return to neutral. Both positive and negative elements of the circuit are switched to prevent a trim runaway should one set of switch contacts become shortcircuited. The trim indicator is driven electrically by a transmitter in the rudder trim actuator. The indicator shows up to 17 units of left or right trim. Each unit represents approximately one degree of rudder trim.

9.3.3 AILERON TRIM SYSTEM

In a typical aileron trim system for a powered system, trim commands from the trim switches causes the actuator to extend or retract, which repositions the feel and centring mechanism null detent. The trim switches must be operated simultaneously to provide an electrical input to the actuator, as both positive and negative elements of the circuit are switched to prevent a trim runaway should one set of switch contacts become short circuited. The available aileron trim provides 15 degrees aileron travel in both directions from neutral.

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9.3.4 TAILPLANE TRIM SYSTEM

For trimming the aircraft longitudinally (about the lateral axis) the elevators are not trimmed. Instead the angle of incidence of the whole tailplane is altered. Raising the leading edge of the tailplane will increase lift over the tailplane, which imparts a nose-down attitude to the aircraft or vice versa. This is done by mounting the forward end of the tailplane on a screw jack. Depending on the system the screw jack is rotated by two hydraulic or electric motors via a gearbox. Movement is induced by a lever in the flight deck, which operates solenoid selector valves or an electric control circuit to operate the motors. Over-travel is prevented by micro-switch. Reasons for fitting to transport aircraft: 1. All aircraft benefit from having as large a range of useable centre of gravity as possible. This gives flexibility in cargo loading and allows for fuel usage in a swept wing. 2. Aircraft benefit from a wide speed range. Very simply, when an aircraft is trimmed at a particular speed, a reduction in speed calls for "up" elevator and an increase in speed calls for "down" elevator. This would cause extra drag. 3. The need to compensate for centre of pressure changes due to slat/flap extension, gear extension. 4. To reduce trim drag to a minimum to give the optimum performance in cruise.

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Variable Incidence Tailplane Trim System The tail-plane is pivoted at the rear of the centre section torsion box and attached to an actuator forward of the centre section. Operation of the actuator raises or lowers the leading edge of the tail-plane, altering the incidence angle.

Variable Incidence Tailplane Figure 9 Page 1

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The actuator comprises a re-circulating ball screw jack and nut assembly driven by two hydraulic motors with separate spur gear reduction trains.

Figure 10

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Friction brakes ensure that air loads cannot back-drive the actuator when the system is de-pressurised. The actuator is signalled from one of three sources: i)

Auto-pilot servo

ii)

Mach trim servo

iii)

Trim hand-wheel operation.

A cable loop runs from the pedestal in the cockpit, under the cabin floor, and ends at a cable reduction-gearing unit at the tailplane incidence actuator. Hydraulic Power Supply Each hydraulic motor is powered from a separate system. In the event of a single hydraulic system failure, a bypass valve permits that motor to "freewheel" when the system is de-pressurised. Position Indication Systems Geared indicator scales inboard of the cockpit hand-wheels present the demanded position of the tail-plane. This will be the actual tail-plane incidence with the hydraulic system(s) pressurised. Actual tail-plane position is continuously displayed on the pilot's instrument panel, signalled by a position transmitter operated by the tail-plane. External markings on the structure adjacent to the tail-plane give the approximate position of the tail-plane. Tail-plane in Motion Warning Some aircraft types have a tail-plane in motion warning system to alert the pilots of continuous motion of the tail-plane beyond a certain time period.

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ACTIVE LOAD CONTROLS

9.4.1 ACTIVE LOAD CONTROL

This system is a relatively new approach to civil aviation, although it has been in use for some time in military aircraft. It is a complex system that senses disturbances in the air that may cause both discomfort to passengers and crew, whilst causing extra unnecessary loading on the airframe. The gusts that are about to hit the aircraft are sensed either by a tiny pair of vanes on either side of the nose or by accelerometers mounted inside the nose of the aircraft. These instantly send a signal, 'bump coming', to the flight control computers, which instantly send a correcting signal to the elevators that counter the bump and give a smoother ride. The whole system requires the quick reactions of both the computers and the hydraulic jacks to be successful. If the aircraft senses a downdraft, the computers instantly signal just the correct amount of 'up elevator' to counteract the disturbance and leave the aircraft to fly smoothly on.

9.4.2 ACTIVE CONTROL TECHNOLOGY

Active Control Technology (ACT) can be defined as “the use of a multivariable automatic flight control system to improve the manoeuvrability, dynamic flight characteristics and the structural dynamic properties of an aircraft by simultaneously driving an appropriate number of control surfaces and auxiliary force or moment generators in such a fashion that either the loads which the aircraft would have experienced as a result of motion without an ACT system are much reduced or the aircraft produces a degree of manoeuvrability beyond the capability of a conventional aircraft.” In essence ACT is the use of technology to make an aircraft and its control surfaces operate in an unconventional manner to effect high manoeuvrability or to reduce airframe stress.

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ACT is nothing new, it has been used on aircraft for many decades but it has increased in usage with the advent of flight control computers and fly-by-wire systems. The Tristar aircraft has a system installed that reduces the flight loads on the wings by partially deploying the spoilers. This changes the lift profile over the wing, bringing the lift closer to the wing root, which is much stronger (see next fig). This means that the wing can be lighter and the wing stresses will be reduced.

Figure 11 Numerous control surfaces, auxiliary force and moment generators can be added to make the aircraft operate unconventionally. Fighter aircraft and some executive jets may have a number of such devices fitted to make them more agile. These include: 

Foreplanes which can only move together to give pitch control.



Canards, these differ from foreplanes as they can also move independently giving more response in roll.



Flaperons which are control surfaces that act as flaps and/or ailerons depending on the pilots selection. They have the ability to move both up and down independently for roll control, but can also move simultaneously for take off and landing.



Thrust vectoring, mainly used on combat aircraft, but the advantages gained with short take off and landing will mean that some form of vectoring system will be developed for commercial aircraft in the future.

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9.4.3 ADVANTAGES OF ACTIVE CONTROL TECHNOLOGY

The employment of Active Control Technology presents numerous advantages both for civil and military aircraft, namely: 

The aircraft is more stable in flight



The aircraft are highly agile (military only)



A more comfortable flight for passengers



Reduced fatigue on the aircraft, therefore lighter construction can be utilised



Lighter construction gives better fuel consumption



Varying lift profiles means wings can be more streamlined (less drag)



It is impossible for the aircraft to be flown beyond its design limitations – under normal conditions!

Conventional aircraft have four forces providing control and movement 

Rolling moment



Pitching moment



Yawing moment



Thrust (Drag modulation)

The use of ACT can provide two more additional forces of control and movement: 

Direct lift force



Direct side force

9.4.4 DIRECT LIFT FORCE

In order to change altitude a pilot must pitch the nose of the aircraft up, which may cause him to lose sight of his destination (the runway). Using ACT, the pilot can change altitude by causing the foreplanes and flaperons to operate together increasing the lift on the front and rear of the aircraft simultaneously. This is known as the direct lift force

9.4.5 DIRECT SIDE FORCE

The pilot, conventionally, must roll the aircraft to change its flight path in a sideways plane. ACT allows the aircraft to side step during normal flight by deploying the rudder and the canards together to pull the nose and tail of the aircraft across in the same direction. This is known as the direct side force. Page 1

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9.5 HIGH LIFT DEVICES 9.5.1 FLAPS

These devices have two primary aims, to provide extra lift during take-off and to provide greater lift as well as high drag during landing. The types of flap used on different aircraft depends on the type of aircraft, the method of aircraft operation and other variables. For example, a single engined light aircraft might only have some form of simple trailing edge flap, whilst a large airliner like the Boeing 777 has complex, triple slotted flaps.

Types of Flap System Figure 12 Flaps are fitted to most aircraft and are usually one of the types shown, together with the maximum increases of lift over the 'clean' configuration. As the complexity increases to improve performance, there is a proportional increase in weight, maintenance and cost.

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Whilst the term 'flaps' is used, it is taken as meaning trailing edge flaps, and the term 'leading edge flaps 'refers to those fitted to the leading edges of the wings of most large aircraft. The methods of operation of flaps, are numerous. They can vary from simple, mechanical push rods or cables actuated, via a lever in the cockpit, by the pilot, to complicated, multiple flaps that are electrically selected on the flight deck and hydraulically or electrically powered. Most flap systems have a number of positions, which can be selected at various times. As an example, five positions could be as follows; 00 - flaps up 80 - take-off, first position

250 - landing, first position 400 - landing, second position

150 - take-off, second position These would all be selected by movement of a lever in the cockpit, which will have 'detents' at the various positions. This movement will, as can be seen in the illustration, be transferred to the control valve and on to the motor, which moves the actuators.

Flap Mechanism Figure 13

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Other high lift devices can be found on the leading edges of the wings and include slats, drooped leading edges and Krueger flaps. All of these devices are aimed at smoothing the airflow over the leading edges of the wings when they are at a high angle of attack, thereby maintaining, or increasing lift when the wing would normally be stalled. 9.5.2 SLATS

Slats are separate small aerofoils, which can be fixed or retractable. Their purpose is to control the air passing over the top of the wing at slow speeds. On larger aircraft, the retractable slats have their extension interconnected with the trailing edge flaps.

This can be seen in the illustration, which not only shows the operation of the slats through three different positions, 'stowed', 'active' and 'open', but their association with the four positions of the trailing edge flaps. Fixed slats are usually found on light aircraft, where the complications of weight, cost etc, can be balanced by the limitation of slightly higher drag than a 'clean' wing.

Leading and Trailing Edge Flap Settings Figure 14

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9.5.3 DROOPED LEADING EDGES

Drooped leading edges are a different design, but are aiming at the same effect, that of smoothing the air over the top of the wing. They operate in much the same way as most high lift devices, by screw jack operation with the motive power for the jacks coming from the hydraulic system. 9.5.4 KRUEGER FLAPS

Krueger flaps are, again, a different design for the same effect. These are usually found fitted to the leading edges of the wing at the inboard sections where the effect of 'slats' or 'drooped leading edges' are not as efficient.

Figure 15 Krueger (left) and Drooped (right) Leading Edge Flaps

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9.6 LIFT DUMP AND SPEED BRAKES 9.6.1 LIFT DUMPERS

These devices are used to spoil lift from the wing after touchdown. This ensures that the aircraft's weight is fully on its landing gear, which enables the brakes to work at 100% for the full landing run. If this did not happen, the aircraft would tend to 'float' or „bounce‟ at touchdown, making the brakes inefficient and the risk of skidding much greater. Lift dumpers are nearly always flat, rectangular panels, hinged at their leading edge and powered by hydraulics. They can usually be found on the top of the wing, and located about the maximum thickness, where their deployment would destroy the maximum lift from the wing. To ensure that they deploy at the correct time and also without the need for the pilot to select them, at a very busy time, there is a simple system to deploy them automatically. A set of switches are fitted to the landing gear which 'make' and indicate weight-on-wheels to several systems, once the aircraft is completely on the ground. By giving the pilot a "lift dumper arming" button, he can arm the system, in flight, and know that it will deploy the lift dumpers at the correct time. 9.6.2 SPEED BRAKES

The use of speed brakes is similar regardless of the aircraft type. If the aircraft is a sailplane it is so streamlined that it requires high drag when descending and landing in unprepared fields. A large 400 seat airliner needs to be able to follow Air Traffic Control instructions to descend and maintain certain speeds and a military jet fighter needs to have very high drag on approach, permitting the engines to accelerate quickly if the landing is aborted. All types of speed brake use a variation of the same principle, to put panels of varying shapes into the airflow, to increase the drag. Some are able to modulate, (vary the amount of drag to suit the situation), whilst others are just 'IN' or 'OUT'. Some airliners use the same surfaces on the top of the wing to carry out more than one operation, such as speed brakes when in flight and needing drag; roll control to augment (or replace) ailerons; or as lift dumpers to be used after landing.

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Light aircraft rarely need speed brakes because of their generally high drag designs. A reduction in power will produce a satisfactory slowing down of the aircraft. Streamlined sailplanes, however, usually have vertical panels that project from the wing, top and bottom, which produce large amounts of drag, enabling steep, slow and safe approaches when landing. Military jets have a different need for drag, not only as mentioned during the approach to landing, but during combat and other operations where fast application of drag with a quick reduction in speed can have a life saving effect.

Speed Brake Installation Figure 16

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SYSTEM OPERATION

9.7.1 MANUAL OPERATION 9.7.2 POWERED FLIGHT CONTROLS (P.F.C.U’S)

In large modern aircraft that fly at high speeds, the air loads on the flying control surfaces far exceed the ability of the pilot to move them manually. To overcome this problem hydraulic pressure is used to move the control surfaces, a POWERED FLYING CONTROL UNIT or BOOSTER being used to convert hydraulic pressure into a force exerted on the control surface. In its simplest form, a P.F.C.U. consists of a hydraulic jack, the body of which is fixed to the aircraft structure and the ram, via a linkage to the control surface. To control the P.F.C.U. a servo valve (control valve) is mounted on the jack. The servo valve, which is connected to the pilot's controls by a system of cables and/or pushrods, called the input system, directs fluid to either side of the jack piston and directs the fluid from the other side to return. This flow of fluid will displace the jack ram and as this is connected to the control surface via an output system of pushrods or cables, the control surface is moved.

Figure. 17

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9.7.3 PROPORTIONALITY

To make the controls "proportional" (i.e. the degree of movement of the jack-ram and hence the control surface, should be proportional to the degree of movement of the pilot's controls), a "follow-up linkage" is used. This linkage connects the input system, through a series of levers to the output system in such a way that the movement of the output system (jack ram) tends to cancel the input once the desired position is reached and so output movement ceases. In effect the movement of the jack ram is always trying to re-centre the servo valve and stop fluid flow in the jack.

9.7.4 REDUNDANCY OF HYDRAULIC SUPPLIES

Hydraulically powered flight control units usually derive their hydraulic power from the aircraft hydraulic system. If a PFCU obtained hydraulic power from only one hydraulic supply, a failure of that hydraulic supply due to an engine shut down, loss of fluid due to a leak, or failure of a hydraulic pump. The result would be loss of powered control of the aircraft. The probability of hydraulic failure is too great to allow a system to rely on one hydraulic supply, so redundancy must be introduced into the flight control system. As in the previous notes on hydraulic systems, modern large multi-engine aircraft are arranged such that the engine driven pumps (and the other types of pumps) supply two or more independent hydraulic power supply systems. The following are methods that use that arrangement of hydraulic redundancy to allow failure of one hydraulic supply and still maintain control of the aircraft.

9.7.5 TANDEM PFCU

These are similar to the arrangement shown. They consist of a single jack ram but with two pistons. These pistons are housed in two co-axial cylinders each of which receives pressure fluid from separate power supply circuits via their own duplicated servo valves. The servo valves, which are controlled by the same input system, are carefully set up in the overhaul workshop to ensure they work in unison. This prevents the two hydraulic pistons working against each other. With this arrangement a loss of one hydraulic supply will allow the relevant piston to "free stroke whilst the other piston operates the control surface.

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TANDEM ACTUATOR Figure 18

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9.7.6 DUAL ASSEMBLY PFCU’S

These are similar to the tandem arrangement but two piston rams are located in cylinders mounted side by side with the piston rams connected to a common output lever that transmits the movement to the control surface. The arrangement for the input system, the duplicated servo valves and hydraulic fluid supplies are the same.

Dual Assembly PFCU FIGURE 19 Page 1

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9.7.7 DUPLICATE/TRIPLICATE PFCU'S

In this arrangement each control surface is operated by two or three separate PFCU'S. For hydraulic redundancy, each PFCU is powered from separate hydraulic supply circuits. If one supply system should fail, or if one PFCU should malfunction the effected PFCU can be switched off. In this event a bypass valve within the PFCU will open interconnecting both sides of the jack ram. Therefore, as the pilot moves the input and operates the serviceable PFCU'S, the control surface will move and, "drag" the unserviceable PFCU ram with it. The open bypass valve will allow fluid to transfer from one side of the ram to the other as the PFCU "free strokes". Thus control will be maintained by the serviceable PFCU's driving the control surface, and a hydraulic lock in the unserviceable PFCU is prevented. In this arrangement each control surface (rudder is shown in the diagram) is split into two or three independent sections. Each section is operated by its own PFCU. For hydraulic redundancy, each PFCU is powered from separate hydraulic supply circuits. If one supply system should fail, or if one PFCU should malfunction the effected PFCU can be switched of. In this event the PFCU and its control surface segment will be "blown back" to the neutral position by aerodynamic loads and held by a lock. Thus control will be maintained by the serviceable PFCU's driving their respective segments of control surface. All PFCU's are controlled via a single input system to a common input lever connected to all PFCU servo valves. Therefore if one PFCU malfunctioned it could prevent the operation of the remaining serviceable PFCU'S. To prevent this the input to the servo valves from the common input lever is via compressible spring struts or spring boxes. In normal operation these spring struts/boxes resist compression and allow full control of all PFCU'S. If a PFCU is unserviceable, pilots input will compress the spring strut to that PFCU but the remaining spring struts/boxes will resist compression and operate the PFCU servo valves normally.

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DUPLICATED CONTROL SURFACES

9.1.4

Duplicated Control Surfaces Page 1

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MODULE 11.09 AERODYNAMICS, STRUCTURES AND SYSTEMS Figure 20

9.7.8 SELF CONTAINED PFCU

A self contained PFCU consists of a jack-ram powered by its own dedicated integrally mounted hydraulic “generator" and hydraulic reservoir. The generator is a radial piston pump arrangement within a slip ring assembly. The slip ring position is control ' led by a servo valve piston arrangement. With the slip ring held concentric with the piston bank no movement of the pistons within the rotating piston bank is allowed and no fluid flow will result. If an input moves the slip ring the rotating bank of pistons will be allowed to "stroke" and a flow to the PFCU piston will occur and the PFCU ram will move. Movement of the slip ring in the opposite direction will cause fluid flow to the other side of the piston and the ram will move in the other direction. The piston bank is rotated by a drive from a 3-phase electric motor, which derives its supply from the aircraft electrical system. To maintain redundancy this type of PFCU will be duplicated and each may drive a duplicate and independent (split) control surface as above. As its source of power is electrical, it is independent of the aircraft‟s hydraulic system, therefore even with total hydraulic failure, control can still be maintained. On malfunction of a PFCU, or loss of electric power to that PFCU, it will lose hydraulic pressure and "blow back" to a neutral position where an integral lock will hold it. In this event further inputs to the servo valves are absorbed by spring-strut that allows unhindered operation of the remaining PFCU'S. To give redundancy of electrical power supply, each PFCU in a "set" (i.e. rudder) gets its power supply from a different bus bar.

9.7.9 INPUT SYSTEMS

Generally the input system of the powered flying control system is mainly a cable system with the related quadrants, pulleys and fairleads with the connections to the control column and the PFCU input lever by push rods. To guard against loss of control due to cable breaks the cable system is duplicated. All duplicated runs are routed separately through the aircraft to avoid one incident damaging both control runs. The cable systems meet at a common input lever to the PFCU'S.

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Input Systems Figure 21 9.7.10 HIGH SPEED PRIMARY CONTROLS

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Primary controls are designed to give adequate control in all flight phases. The flight phase at which the control surfaces are least effective is during low speeds (landing). This is because of the reduced aerodynamic effect with low speed. This means that the size and range of movement of each control surface must be sufficient to maintain sufficient control authority. With the control system designed to give efficient control at low speed, there may be a problem at high speed. This is that at high speeds the increased air-loads on the control surfaces will cause them to be too sensitive producing over control and possible loss of control or over-stressing of the airframe. To prevent this two systems may possibly be used.

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Geared Controls In this system a single acting hydraulic jack may be fitted to an idler lever. The control rod is attached to this jack so that the radius of operation can be altered. Thus for a given angular movement of the idler lever, if the length of the jack is shortened, the linear movement of the control rod is reduced. This will maintain a constant range of movement at the pilots‟ controls but reduce the range of movement of the control surface. Pressure at the jack is usually controlled by a pressure-modulating valve sensitive to a pressure transducer in the pitot system. High Speed Control Surfaces (ailerons) Normal, "low speed" ailerons are situated at the usual wing tips position to gain maximum authority due to the moment arm produced. But again at high speed their authority may be too great. In this system an additional set of "high speed" ailerons is also fitted at the wing root. Hydraulic isolate valves are incorporated in the control system such that at low speed the outer ailerons are functional, but at high speed, their hydraulic power is cut off and the high speed ailerons are powered to maintain roll control. The isolate valves are again controlled by pressure switches in the pitot system. 9.8 TRAILING EDGE FLAP CONTROLS On small aircraft the flaps are operated using hydraulic jacks to operate a single flap on each mainplane. This arrangement is not suitable for use on larger aircraft due to the size of the airframe that requires that the flaps are manufactured and mounted in "segments" along the trailing edge.

9.8.1 FLAP CONTROL UTILISING LINEAR HYDRAULIC ACTUATORS

The following system that may be regarded as a simple system, similarly uses linear hydraulic actuators for an aircraft that has three flap segments on each mainplane each positioned by a separate hydraulic actuator. Movement of each actuator is controlled by a servo valve (simiIar to that in a primary flight control unit). Control is by flap lever/quadrant on the centre console. This is connected to the actuator servo valves by a duplicated system of control cables and pushrods.

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9.8.2 GENERAL

The flap surfaces are operated through linkages by hydraulic actuators. The actuators respond simultaneously to the control-cable-relayed demands of a selector lever mounted on the flight compartment centre console. The piston rod end of each actuator is structurally anchored; movement being confined to the unit body. A position control element (servo-valve) incorporated in the body is controlled by an attached operating lever that has limited travel on each side of the neutral position. The lever is moved towards or away from the anchored piston rod end to retract or extend the actuator. Each actuator incorporates internal restrictors that control the rate of response and an internal mechanical lock that engages when the flaps are fully up. The lock is hydraulically released when a down selection is made. The control system consists of a duplicated input circuit, which through the medium of a spring strut, signals all six actuators. Beyond the spring strut the signal to the inner flap actuators is conveyed by a rod and lever system and to the mid and outer flap actuators by interconnected signalling cables. The purpose of the spring strut is to "store" control lever movement due to the actuators' restricted rate of travel. The adjacent ends of the mid and outer flap surfaces are connected by a link that allows sufficient free movement to accommodate normal variations of relative positions without the links being loaded. The links are incorporated as a safety feature and take effect to prevent an asymmetric flap condition. The flap selector lever is afforded the following gated positions - 0º, 5º, 15º and 30º.

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9.8.3 HYDRAULIC POWER

Hydraulic power for operation of the actuators is provided by main system pressure backed up by flap accumulator pressure, when the flight compartment selector lever is at any position other than fully up (0º). The accumulator-stored pressure is released to the flap system when a solenoid valve is energised open via a micro-switch operated by the selector lever. The 'back-up' pressure is introduced downstream of a non-return valve in the main system pressure line; thus maintenance of a selected down position is assumed, for a limited period in the event of a main system failure.

Flap System Hydraulics Figure 22 9.8.4 CONTROL INPUT CIRCUIT

From the flap selector lever on the centre console, the duplicated input cables are routed aft through the roof structure to a position immediately aft of the rear spar. At this point, the cables are directed through the roof skin terminating with a double quadrant assembly. A double acting spring strut is connected between an output lever on the quadrant and a series of levers and control rods. These:  Operate the position control elements (servo valves) on the inner flap actuators and transmit actuator movement to the inner flap surfaces.  Provide an input to the left and right mid and out flap signalling circuits. Page 1

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 The spring strut is incorporated to allow a selection to be made in one quick movement - the total input motion being absorbed by the spring strut and progressively released as all six actuators respond at their controlled rate of travel. Each of the left and right mid and outer flap signalling circuits consists of a pulley drum from which cables are routed outboard to quadrant assemblies at the mid and outer flap positions. Output levers on these quadrants are linked by control rods to the position control element (servo-valve) operating levers in the appropriate actuator package assemblies. The left and right pulley drums are interconnected by two tie-rods to ensure symmetrical operation of the left and right wing flaps.

Flap Control Input Circuit Figure 23 9.8.5 SYSTEM OPERATION

Immediately a selection is made the total input motion is absorbed by the spring strut and progressively released as all six actuators respond at their limited rate of travel. When the spring strut returns to its pre-selection settled length - the rod that connects to the position control element-operating lever on each actuator arrests. The actuators will then marginally run on until their now restrained element operating levers reach neutral positions. This simultaneously creates a hydraulic lock at all six actuators and hence arrests the surfaces in alignment at the selected position. Page 1

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9.8.6 SAFETY ASPECTS

Two main safety requirements must be met. 1. One is that a control cable break will not mean loss of control of the flaps.

System integrity is such that duplication of the input cables which allows for functioning in the event of loss of either circuit) will maintain control. 2. The other is that an asymmetric deployment of the flaps is prevented. An

asymmetric condition could happen in several ways and the following mechanisms are designed to prevent these. A. Controls jamming between an actuator and surface (input systems intact):Should this occur during a programmed selection, the input system of the relevant actuator will arrest and in consequence will stop signalling of the remaining actuators which will then run on marginally until their now restrained servo valve operating levers reach neutral positions - thus arresting all six surfaces in approximate alignment. B.

Mechanical failure between an actuator and surface (which will not impede surface movement):-

Should this occur at either of the inner flaps - the system will remain functional (full asymmetry between inner flaps can be adequately countered by aileron action). Should this occur at a mid or outer flap - the link which interconnects the adjacent ends of these surfaces will take effect to allow full functioning of both surfaces from one actuator. Thus preventing an asymmetric condition that would be beyond the ailerons ability to counter. C. Loss of signalling (cable break) to a mid or outer flap actuator. Should loss of signalling to a mid or outer flap actuator occur and the 'free' actuator become hydraulically locked at any stage during a programmed operation - the interconnecting link will arrest the adjacent functional actuator and thus its intact signalling system. This will have the effect of simultaneously arresting the interconnected input circuits of the remaining actuators that then run on marginally, until their now restrained servo valve operating levers reach neutral positions - thus arresting all six surfaces in approximate alignment. The actuator arrested by the link will remain programmed to achieve intended travel in opposition to the locked adjacent surface. For this reason and to prevent excessive structural overloading - the actuators incorporate internal relief valves. D. Loss of main system pressure

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Main system pressure is augmented by flap accumulator stored pressure via a solenoid valve when the FLAPS selector lever on the flight compartment centre console is at any position other than fully up (0'). The 'back up' pressure is introduced down-stream of a non return valve in the main system pressure line; thus maintenance of a selected down position is assured for a limited period in the event of a main system failure.

9.8.7 POSITION INDICATION

Flap position is indicated on a twin pointer scale calibrated to 0', 5', 15' and 30' settings. The flap position is signalled by two transmitters that are driven from the flap hinge arms via control rods.

9.8.8 FLAP SYSTEM - HYDRAULIC MOTORS AND TORQUE TUBE DRIVE

On large aircraft it is more common for the flaps to be driven by twin hydraulic motors, each motor deriving its hydraulic supply from a different hydraulic system. Each motor is mounted on the same gearbox, such that drive from either or both motors will drive the gearbox. The gearbox is commonly located in the main gear bay. The drive is transmitted to the flap surfaces by a system of torque tubes, gearboxes and screw jacks. The screw jacks drive trolley assemblies along flap tracks mounted to the wing structure via support units. The flap segments are mounted onto the trolleys. System Description The flap system of each side of the aircraft comprises of flap sections supported and moved by six support/operating units. (Flap Tracks) The flaps are manually controlled by a lever on the central console to UP (0º), take off (20º), approach (35º) and landing (45º) positions. This manual control operates independent Electro/hydraulic systems A and B, employed simultaneously to power the drive unit (gearbox) and their supplies are drawn from the aircraft electric and hydraulic systems bearing the same suffix letter. Both systems normally operate together, but should a hydraulic system fail, or a fault develop which necessitates selection of ISOLATE on one system, the flaps travel only at half rate due to the design of the drive unit.

a. Drive Unit The drive unit, comprises a gearbox and selector drum assembly, powered by two hydraulic motors. It rotates a torque shaft system that operates screw jack and trolley mechanisms at each support/operating unit.

The drive unit is mounted to the rear of the wing rear spar member in the left main landing gear bay. It is powered by two hydraulic motor/lock valve assemblies; one supplied from hydraulic system A and the other from system B. Page 1

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The motors drive a main shaft through a differential gear and a spur wheel reduction gearing. A gear driven selector drum operates micro-switches to arrest the flaps when they reach the selected position. b. Flap Transmission System Torque shafts extend outboard in each wing from either end of the drive unit main shaft. The sections of torque shaft couple via universal joints and serrated sleeve joints to bevel gearboxes and to intermediate bevel gearboxes. The bevel gearboxes and intermediate gearboxes are connected by serrated sleeve joints and universal joints to screw shaft assemblies located at each support unit. Flap trolleys fitted to each screw shaft engage via their rollers with trolley tracks fitted to the support units. These trolleys support the flap sections. The flaps are hinged by pins to lugs on the flap trolleys. A torque link pivoted to each flap section carries a forward flap trolley, the rollers of which engage with the cam track on the support unit. c. Hydraulic System For redundancy the flaps are supplied by two independent hydraulic systems, which are identical. The following therefore describes one system only. Hydraulic pressure is supplied to the flap selector valve via a flow control valve and isolating valve. Movement of the flap selector lever energises the appropriate solenoid selector valve to allow pressurised fluid to pass to the hydraulic motor through the lock valve. Return fluid from the hydraulic motor passes through the lock valve and flap selector valve back to the main system. The flow control valve controls the rate at which the flaps move. A throttle valve slows down the flaps at all selected positions. When the flaps reach the selected position the selector valve solenoid is deenergised, through the operation of the selector drum micro switches. The pressurised fluid is held at the selector valve and the two service lines from the lock valve are connected together and into return. The lock valve prevents the hydraulic motor from rotating. d. Flap Control Each separate flap operating hydraulic circuit is controlled by a separate 28 volt D.C. electrical system. Each supply is derived from a separate D.C. Bus Bar. Each system is controlled by three micro-switches operated by control lever movement, these provide a circuit to the selector valve solenoids via six microswitches operated by the drive unit selector drum. Cams on the outer periphery of the selector drum operate one switch at both the normal, up and down limit positions of the flaps and two switches at the take-off (20') and approach (35') positions. Page 1

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Flap control unit Figure 24 e. Over-Run Protection If the micro-switches in the drive unit selector drum should malfunction there is a probability that structural damage may occur as the flap trolleys reach the end of travel on their screw jacks. If a malfunction should happen, a set of over-run micro-switches mounted on the flap support units, will be operated to interrupt the supply to the selector valve solenoid and prevent the trolleys bottoming on their screw jacks. These microswitches are part of the complete control circuit and are operated by strikers on the flap support trolleys. f. Asymmetry Protection If a malfunction should occur in the flap transmission system causing one part to seize, great damage could occur as the drive system attempted to drive the flaps to their selected position. To prevent this, weak "fail safe" joints are incorporated in various torque tubes that are designed to fail under a certain load.

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However, this will allow the damaged portion of the system to stop and the remainder to continue travelling, so producing an asymmetric flap condition. To prevent this an asymmetry protection circuit is incorporated in the control system. This system uses an A.C. electrical supply and is controlled by four synchro's which are small devices mounted on and driven by screw shafts in board and out board of the flap systems. These are paired, and as they rotate send an alternating signal to an asymmetry control box. If the signals become out of phase with each other the over-travel/ asymmetry isolate relay will be energised to lockout the system. g. Position Indication Flap position indication is provided by a D.C. ratio meter indicating system comprising two transmitters, driven from the outboard end of the left and right torque shaft systems and dual indicators positioned on the centre in the flight deck.

9.8.9 MAINTENANCE OF FLAP SYSTEMS

Because of the exposed position of most flap system components regular lubrication of hinge bolts, screw jacks, trolleys etc is required. When carrying out this task all excess grease must be removed to prevent the accumulation of dirt or grit that may enter bearings etc. Rigging The flap operating system is a large complex system which will only work if all parts are in their correct relative positions at all times. To ensure this, whenever the system is disturbed by a maintenance task it must be checked or re-rigged. Provision is built into the system for this. 9.9 LEADING EDGE FLAP CONTROLS The following notes describe a typical leading edge flap control system (Boeing 747). There are two modes of operation, primary and alternate. The normal method of operation is by use of the primary mode and is initiated by use of the flap lever to a selected detent. On the Boeing 747 there are 28 leading edge flaps, they in turn are divided in two categories, variable camber (22) off, Kruger (6) off. Four pneumatic power units in each wing move the flaps up or down, (extend) or (retract). Air is supplied to the power units from ducts in the leading edge of the wing. The ducts also supply air to jets that spray the outboard drive units with hot air for anti icing. Each drive unit assembly has two motors, one pneumatic and one electrically powered. Torque developed by the drive units is supplied to rotary actuators. The actuators move the flaps.

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Normal operation is achieved by operation of the flap control lever. Three rotary variable differential transducers (RVDT‟s) sense movement and signal the Flap Control Unit (FCU), which control the direction control motors. If pneumatic power is not available the FCU will switch to electric drive motor operation.

Figure 25

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L/E Flap System Components Figure 26

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Figure 27

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9.9.1 LEADING EDGE FLAP PNEUMATIC DRIVE UNIT

Purpose Eight pneumatic drive units (PDU) power the LE flap system. Each power drive unit has both a pneumatic and an electric drive motor. The pneumatic motor is the primary drive source and is powered by the leading edge pneumatic manifold. The electric motor is an additional drive source for use when the pneumatic system is not available. Leading Edge Flap Drive Unit- operation a. Pneumatic Drive The flap lever is used to command FCU operation. The flap control unit signal is passed to the directional control motor and the shutoff valve. Pneumatic pressure flows from the inlet duct through the alternate valve (normally open) to the shutoff valve. The shutoff valve (normally closed) opens to pressurise the regulator and the air Motor brake. Pneumatic pressure at the regulator opens the butterfly valve and regulates the pressure to the control valve. Pneumatic pressure at the air motor brake releases the brake. The direction and speed difference between the direction control motor and the output shaft follow-up gear is sensed by the differential. The differential uses the speed differences to position the control valve and maintain PDU speed. Travel limits are governed by the primary position controller. This translates the amount of distance that the nut travels. When the translating nut reaches its travel limit it stops the direction control motor rotation that, in turn, stops PDU operation. b. Electric Drive The signal to activate the electric drive motor closes the alternate solenoid valve. The electric motor brake then releases the electric motor drive. The pneumatic brake holds the sun gear of the planetary gearbox at the air motor output shaft. The electric motor drives the output shaft through the ring gear of the planetary gear reduction. When the translating nut in the alternate position controller reaches the end of its travel it opens the electric motor limit switches. The alternate controller position switches control the electric motor shutdown in both primary and alternate control modes.

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Operating Times The approximate leading edge flap extension or retraction times are: 

Pneumatic operation: 9 seconds



Electric operation: 90 seconds

Electric Drive. Motor Control Primary Mode: in the primary mode the FCU controls the LE flap operation. If the pneumatic drive motor is not available the FCU will select the electric drive motor. The alternate controller provides signals to the FCUs for control, monitoring, and indication functions. Alternate Mode: in alternate mode the electric drive motor is the only method of moving the flaps. The alternate arming switch arms the system. Flap operation is commanded by using the rotary alternate control switch located on P-2 in the flight deck.

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Figure 28

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9.9.2 KRUEGER FLAP DRIVE COMPONENTS

Purpose/Location The Krueger flaps modify the configuration of the inboard portion of the wing leading edge to increase low speed lift. There are three Krueger type LE flaps installed on each wing inboard of the inboard engines (flaps 11 through 16).

Figure. 29 Page 1

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9.10 SPEED BRAKE/GROUND SPOILER CONTROL Spoilers will normally be controlled by the pilot through the normal roll controls or by the automatic flight control system (auto-pilot). They may also be operated automatically as part of an automatic landing system. On a typical aircraft (Boeing 757) the spoilers are electrically controlled and hydraulically powered.

Figure 30 Page 1

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9.10.1 OPERATION

Rotary variable differential transducers (RVDT‟s) convert control wheel inputs into electrical signals. Spoiler control modules receive the signals and command the Power Control Actuators (PCA‟s) to raise the spoilers. Placing the speed-brake lever in the UP position will raise all flight spoilers.

Figure 31

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Three RVDT‟s are grouped together in a can-like unit, mounted on the bottom of each control unit assembly. The RVDT‟s convert aileron control wheel rotation into a signal voltage proportional to the control wheel movement. The Spoiler Control Modules mix RVDT inputs with other inputs according to a programmed logic. Six SCM‟s control the 12 spoiler surfaces. Power Control Actuators operate the spoilers. Each spoiler has one PCA, powered by one of three hydraulic systems. Each PCA consists of a hydraulic actuator, an electro-hydraulic servo valve (EHSV) and a Rotary variable differential transformer (RVDT). The PCA extends or retracts as commanded to raise or lower the spoiler. The RVDT sends a feedback signal to the SCM proportional to the amount of surface deflection. Electro-hydraulic Servo valves controls the flow of hydraulic fluid in the PVA in response to the SCM commands. The command operates a jet pipe that supplies hydraulic fluid to the EHSV control bobbin. The EHSV is spring loaded to the retract position, so the spoiler panel will retract if there is no command signal.

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Spoiler Electro Hydraulic Servo Valve Figure 32

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9.11 MECHANICAL & ELECTRICAL FLIGHT CONTROL SYSTEM 9.11.1 MECHANICAL CONTROLS

Most aircraft use conventional mechanical controls to move the flight controls. These will normally consist of cables, chains and control tubes. Many examples of this type of system have been described and illustrated previously. The ailerons and elevators on this type of system would normally be operated by a conventional control column and control wheel. Operation of this is instinctive to the pilot, the control wheel being rotated to the left to bank left and right to bank right. Pushing the control column forwards causes the aircraft to dive and pulling back causes the aircraft to climb. A typical control wheel and other cockpit controls is illustrated.

Figure 33 Page 1

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9.11.2 ELECTRICAL FLIGHT CONTROLS

Many modern aircraft use electrical inputs to the powered control units. This eliminates the need for mechanical controls and all of the chains, pulleys, fairleads and linkages associated with this type of system. This topic is covered in more detail in the Fly By Wire section, but the following paragraphs illustrate a typical Airbus system. The electrical flight control computers are designed to ensure a high degree of safety. This is accomplished by using a high level of redundancy which consists of five EFCS computers installed in the aircraft, the use of dissimilar redundancy which consists of two types of computers with each being capable of achieving pitch and roll control along with other redundant features assuring aircraft control. Each computer is also composed of one control unit and one monitoring unit. Control and monitoring software are different and the control and monitoring units are physically separated. Monitoring In each computer, one monitoring channel is associated to a control channel by use of self- monitored channels. Each computer is able to detect its own failures (microprocessor test, electrical power monitoring, input and output test). Input monitoring by comparison of signals of the same type, but sent by different sources, and checking of the signal coherence along with permanent cross talk between associated control and monitoring channels, consolidate and validate information received. This allows permanent monitoring of each channel by its associated one. Automatic test sequences can be performed on the ground when electric and hydraulic power is applied (no surface deflection during test). Side-stick Controller The side-stick controllers are used for pitch and roll manual control and are shown below. The side-stick controllers are installed on the captains and first officer's forward lateral consoles. An adjustable arm-rest is fitted on each seat to facilitate the side-stick control. The side-stick controllers are electrically coupled.

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Figure 34 In the case of one pilot wanting to take control of the aircraft (priority), the autopilot instinctive disconnect button is used to signal the priority system. A visual indication is given to the pilots to indicate left or right side-stick priority. In autopilot operation the side-stick controllers remain In neutral position. The autopilot function can be overridden by the pilots and the autopilot then disengages. Control Laws Normal control laws selected for A320 pitch and lateral control are manoeuvre command laws with normal acceleration and roll rates used as basic parameters. Inside the normal flight envelope, the main features are a neutral static stability, short term attitude stability, along with automatic longitudinal trimming.

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The flight characteristics that can be controlled are: 

the automatic elevator in a turn



lateral attitude hold in a turn



dutch roll damping



turn co-ordination



engine failure compensation.

In addition, protections are provided against extreme attitudes (pitch and roll) excessive load factors, over-speed, and stall. The load alleviation function (LAF) is accomplished by the electrical flight control system (EFCS). The LAF is implemented in the elevator and aileron computer (ELAC) and the spoiler elevator computer (SEC). The control surfaces used are both ailerons as well as spoilers 4 and 5 (i.e. the outboard pair on both sides) for up gusts. There are four specific accelerometers that are installed in the forward fuselage station to provide the electrical flight control computers with vertical acceleration values. These sense the up gust and deploy the spoilers to smooth out the normal result of an up gust of wind as described in the before mentioned example. Four hydraulic accumulators are installed to provide the extra hydraulic flow needed to achieve the surface rates and duration of movement required for load alleviation as illustrated below.

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‘Q’ FEEL, YAW DAMPER, MACH TRIM, RUDDER LIMITER, GUST LOCKS

9.12.1 ARTIFICIAL FEEL

9.12.1.1

‘Q’ Feel System Principles

'Q' feel is an artificial force felt at the control column, which increases as the aerodynamic pressure (Q) at the control surface increases. Aerodynamic pressure conforms to the relationship Q = ½pv2,. Where “p” is air density, and v is the velocity of air flow. Thus, a 'Q' feel system has to simulate the actual control surface loading lost with the use of powered controls; preventing the pilot damaging the aircraft by pulling excessive 'g' loads. Artificial 'Q' feel" units have to increase the control column centralizing force, in proportion to the square of the airspeed. In general, 'Q' feel systems can be either mechanically or hydraulically operated. Typical systems are explained below.

9.12.1.2

Mechanical ‘Q’ Feel System

Spring feel has the disadvantage of being constant throughout the airspeed range. However, with this system the effective force provided by the spring cartridge is adjusted for given airspeeds. This is achieved by moving the fulcrum point of its bell crank lever. Rather like the study of lever mechanisms, where the given forces by distances are equal on either side. Thus, we can attain a mechanical advantage over the spring, increasing or reducing the effective feel force.

9.12.2 OPERATION

Refer to diagram overleaf: The slotted bell crank lever has the control rods attached at one end, and the spring cartridge at the other. As a control surface demand is made, this lever pivots about the roller, which is attached to the fulcrum arm. Relative positions of the fulcrum arm determine the amount of feel felt back at the stick. The fulcrum arm can be repositioned by means of an electrical linear actuator. Page 1

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Should the actuator be extended, the fulcrum arm would be lowered. This gives a short distance from the roller to the spring, relative to the control rods. Hence, there is a good mechanical advantage in the mechanism, making it easy to move the spring cartridge. This would be the configuration for low airspeeds. As the airspeed of the aircraft increases, the fulcrum arm would move up, progressively giving more feel to the system. The linear actuator operates from a closed loop positional servo system. Input is by means of an airspeed sensor, which converts the pitot/static pressure differential into an electric signal. Feedback is achieved by means of a follow-up potentiometer attached to the fulcrum arm.

PITOT/ STATIC SLOTTED BELLCRANK LEVER AIRSPEED

FOLLOW-UP POT

SENSOR SERVO AMP

ROLLER

+ _ FULCRUM ARM

LINEAR ELECT ACTUATOR

SPRING CARTRIDGE

Mechanical „Q „ Feel System Figure 36

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9.12.3 HYDRAULIC ‘Q’ FEEL SYSTEM

NORMAL OPERATION A hydraulic jack is attached to the control rods adjacent to the control column. The principle of operation is that the pressure of hydraulic fluid within "this 'Q' feel simulator jack, will be proportional to the amount of force necessary at the stick, to overcome it. Low pressure produces light feel. High pressure produces heavy feel. To provide this pressure differential relative to airspeed a special 'Q' feel unit is used. Pitot and static pressure are transmitted to the unit, but are isolated from one another by a flexible diaphragm. As the airspeed increases the pitot pressure acts to push the diaphragm down. This action is the resistive force acting against the upward tendency of the servo valve piston. Signal pressure is supplied to the 'Q‟ feel jack at differing magnitudes, by the servo valve. This signal pressure is proportional to the airspeed. Different pressures are achieved by the action of the servo valve piston acting against the force created at the diaphragm. At zero airspeed (static) the piston will be fully up, as there is no pitot pressure resisting it. This will close the valve pressure inlet and open the signal pressure lines to exhaust (return line). Hence, no feel simulated. With an increase in airspeed, there will be a greater force felt on the diaphragm side of the piston. Therefore, a greater pressure will be required in the signal pressure lines to close off the servo valve pressure inlet port. Hence, feel is simulated at the control column, and this builds up in proportion to the square of the airspeed.

9.12.4 MACH NUMBER CORRECTION

As increased Mach numbers are reached there is a reduction in the effectiveness of the control surfaces, for a given amount of deflection. This effect is due to the compressibility of air at supersonic speeds. Therefore, at such Mach numbers, the feel force has to be reduced accordingly; regardless of the aircraft speed. 9.12.5 OPERATION

On the Mach number correction side of the unit the diaphragm has differential areas, upon which pitot and static pressure may act. This is due to the underside of the capsule reducing the area on the pitot pressure side, but the static pressure can affect the whole of the diaphragm under-surface. Page 1

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Hence, an increased force is felt on the underside of the diaphragm, for relative pressures on either side. At low and moderate airspeeds this retains the capsule in the position as shown in the diagram. As higher Mach numbers are approached, an increase in pitot relative to static pressure is experienced. This has the effect of pushing the diaphragm down, in proportion to the Mach number reached. In turn, a linkage has the effect of pushing up on the servo valve piston against the normal diaphragm. Signal pressure is subsequently reduced, and there is less centering force at the stick. The pilot has less feel. Mach numbers are not always constant for a given airspeed. They change with the aircraft altitude. To compensate for this effect the capsule is evacuated, and operates on an aneroid principle. MACH CORRECTION CAPSULE PITOT

DIAPHRAGM STATIC EXHAUST

CONTROL COLUMN

PRESSURE INLET

INPUT TO PFCU

Q FEEL JACK

Hydraulic „Q‟ Feel System Figure 37

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9.13 YAW DAMPING Yaw damping is provided on some aircraft to improve the directional stability and turn co-ordination. When the aircraft yaws due to side air-loads, a hydraulic yaw damper actuator automatically compensates by generating rudder control inputs. The following notes describe a typical yaw damper system. 9.13.1 YAW CONTROL

Yaw control is provided by a single piece rudder actuated by three independently supplied hydraulic servo-jacks. They are signalled via interconnected pedals by a single cable run up to a spring loaded artificial feel unit connected to the trim screw-jacks. The commands are transmitted by a single load path linkage fitted with a centring spring device. This holds the servo-jack inputs in the neutral position should a disconnect occur. Rudder travel is limited as a function of air speed (CAS). Orders are delivered by the flight augmentation computer (FAC) controlling electric motors coupled to a variable mechanical stop, as illustrated in the picture below. Yaw dampening is operative throughout the whole flight envelope. Yaw damper commands are transmitted via a differential unit. Yaw stability augmentation orders are delivered by the FACS.

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Artificial feel is provided by a spring rod, the zero force position of which is controlled by an. electrical trim actuator. An automatic reset function initiated by pressing the RESET pushbutton allows the rudder trim position to be nulled through the FACS. Rudder trim position is displayed on an indicator adjacent to the trim switch.

Figure 38

9.14 MACH TRIM Modern transport aircraft are designed to cruise at high mach numbers, close to, or at the speed where shock waves may form on the wing. This is their "critical mach number". At this aircraft speed the formation of the shock waves causes shock induced separation and a movement of the centre of pressure forward. This produces a pitch up which must be countered. The Mach Trim System is provided to automatically maintain the correct aircraft pitch trim angle in relation to speed by varying the tail-plane trim. In achieving this function, the system maintains the same degree of longitudinal stability throughout the operational speed range of the aircraft.

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Tailplane Operating Mechanism Figure 39

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9.14.1 TYPICAL SYSTEM

The mach trim system operates within the range from 0.68 IMN (Indicated Mach Number) to 0.84 IMN when the aircraft is above 9000 ft. The system operates in passive mode when the aircraft is flown with the autopilot engaged, but becomes active if the autopilot is disengaged. A mach trim activity light on the pilot's instrument panel flashes intermittently to indicate that a trimming demand exists. Illumination of the light for a sustained period indicates a runaway or seized actuator. A mach trim ON/OFF switch located in the cockpit permits a faulty system to be isolated.

9.14.2 CONTROLLER

The controller is supplied with height and speed inputs from the aircraft pitot static system. The inputs are used to generate control signals that determine the direction and rate of rotation of the mach trim actuator. The controller also provides the 28 volts DC output to energize the clutch and connect the mach trim actuator to the tail-plane trim system.

9.14.3 MACH TRIM ACTUATOR

The actuator is located in the centre pedestal in the cockpit, and is connected by a chain drive to the manual tail-plane trim hand-wheels cross-shaft. A solenoid operated clutch connects the mach trim actuator to the drive system. The tailplane auto-trim actuator operated by the auto-pilot system is also attached to the mach trim actuator. The control system ensures that only one actuator can be engaged at a time.

9.14.4 OPERATION

With the system selected ON and the auto-pilot disengaged, the mach trim actuator is clutched to the tail trim mechanism as soon as the aircraft power supplies are switched on. The system becomes active as soon as the aircraft flies above 9000 ft and its speed is within the Mach number range 0.68 IMN to 0.84 IMN. If the manual tail-plane trim hand-wheels are operated, the mach trim actuator is declutched to permit the tail-plane incidence to be changed and the clutch reengaged when the trim hand-wheels are released. Page 1

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Typical Mach Trim System Figure 40 Page 1

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9.15 RUDDER LIMITING At slow speeds the pilot is able to utilise the full movement of the rudder to enable maximum control of the aircraft during landing and take-off. As airspeed increases, the same full movement of the control surfaces would have a much more dramatic aerodynamic effect. Structural damage could occur if the controls were moved the same amount as at low speed. The artificial feel systems previously discussed how feel is incorporated into the controls. Rudder limiting restricts the maximum movement of the rudder as airspeed increases. Two typical systems are described.

9.15.1 ‘Q’ LIMITER

The rudder 'Q pot restricts movement as airspeed increases by extending the stepped stop, which restricts movement of the clawed stop. The clawed stop is connected by rod to one end of the inner level in the trim unit that restricts the movement of the input lever. The stepped stop is extended by operation of the rudder 'Q' pot. The 'Q' pot assembly comprises a cylinder assembly, a sealed piston bolted to a springloaded piston rod. The sealed piston divides the 'Q' pot into two sealed chambers No.1 pitot static and No.2 pitot pressure. These chambers are supplied from the 'Q' pot pitot head located on the lower left-hand nose fuselage. As pressure from the pitot head rises in chamber No.2, the piston moves, compressing the spring and extending the piston rod and consequently the stepped stop into the clawed stop. A microswitch is mounted on the 'Q' pot and is operated by a cam on the stop in the extended position. 9.16 GUST LOCKS The following notes describe a typical aircraft system and refer in some cases to specific references associated with that system.

9.16.1 DESCRIPTION

The gust lock system is employed to lock the primary control surfaces in the neutral position for taxiing, parking or mooring the aircraft. The system consists of forward and aft installations that are electrically connected for operation by a single lever on the flight compartment centre console. The forward installation caters for the locking of the aileron and elevator surfaces; the aft installation, which includes an electrical actuator, locks the rudder.

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Each installation incorporates a weight switch controlled solenoid lock (snib); that in the forward installation renders it impossible to select controls LOCKED in flight. The aft snib prevents locking of the rudder should its associated gust lock actuator be subjected to spurious electrical signals. When ground selected LOCKED, the control lever, in addition to locking the primary control surfaces via the mechanisms described in the following paragraphs, also operates an interlock mechanism; this baulks engine power lever movement to restrict engine power during taxiing. Warning of the locked condition is provided by a warning light on panel 1P.

9.16.2 CONTROLS LOCKING MECHANISM (AILERON AND ELEVATOR)

The mechanism, shown in the following diagram, basically consists of two pivoted locking arms, each of which is provided with an open-ended slot. The aileron and elevator arms are connected by input springpots to levers fixed to the controls locking lever shaft. The controls locking lever handle has two positions: a. UNLOCKED - forward and b. LOCKED - aft. The handle incorporates a spring-loaded push rod that protrudes from the upper end of the handle as a push-button. The rod is provided with a collar at the lower end that can engage in either of two locking holes in a structurally anchored gated bracket on the levers shaft. The locking holes are joined by a slot that allows the collar to be push-button displaced. The handle is then ground selected from UNLOCKED to LOCKED and vice-versa. When selected LOCKED the input springpots load the locking arms against pins fitted to the aileron and elevator primary bellcranks at the positions shown; this 'arms' the mechanism such that when the associated primary circuit is brought to its lock position the slot in the related locking arm will 'snap' engage with the lock pin on the bellcrank. When UNLOCKED is selected, the input springpots will pull the locking arms clear of the bellcrank - thus freeing the controls. NOTE: THE MECHANISM ADDITIONALLY INCORPORATES A 'FAILSAFE' SPRING ASSEMBLY AT EACH LOCKING ARM; THESE WILL PREVENT THE ARMS FROM ENGAGING WITH THE BELLCRANKS SHOULD A LINKAGE FAILURE OCCUR WHEN THE CONTROLS ARE SELECTED

UNLOCKED.

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9.16.3 CONTROLS LOCKING MECHANISM (RUDDER)

The rudder gust lock mechanism shown in the following diagram is located in the upper left side of the rear fuselage. Basically the mechanism comprises a lock strut that pivots in a limited lateral arc about its bracket-attached forward end. The aft end of the strut is equipped with a bayonet fitting, encompassed by a nylon guide block assembly. The bayonet incorporates an open-ended slot to operationally engage a lock pin in the rudder control lever assembly. The strut is positioned for lock engagement/disengagement by a springpot interposed between the bayonet fitting and an electrical actuator structurally anchored to the fuselage. The linear actuator, circuit identification WM5, is a split field series wound unit driven by a bi-directional motor and equipped with internal extend and retract limit switches. The unit operationally retracts to 'arm' the strut for lock pin engagement; engagement occurs when the rudder pedals are centralized. Conversely, the actuator extends to disengage the lock. Disengagement is assisted by a tension spring.

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Gust Lock Installation Figure 41 Page 1

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9.16.4 POWER SUPPLIES

The system power supply is derived from the 28v dc right essential services busbar. C/B No. 192 (3A) on panel 2D caters for: rudder gust lock actuator operation, on ground retraction of the solenoid lock snibs and controls locked warning indication. A paralleled supply for locked warning indication is also taken from C/B No. 178 (3A) on panel 2D; this is operationally taken via the normally open contacts of microswitch WM4 when the rudder lock strut is engaged. The dual supply assures warning integrity should one or other of the two circuit breakers trip out during a locked condition.

9.16.5 OPERATION

The operation of the gust lock system is essentially as detailed in the preceding paragraphs, however, some discussion of the electrical aspects of control is necessary; this follows: Solenoid locks The two snib-type solenoid locks are flight de-energized (i.e. snibs extended) by a weight switch in the right equipment bay of the nose fuselage. In this condition, the forward unit will baulk a „toe' on the control lever; thus preventing lever movement from the UNLOCKED to the LOCKED position. The rear unit snib will prevent engagement of the rudder gust lock strut. When on the ground, 28v dc is made available through the weight switch relay to energize both solenoids; this withdraws the snibs, thus allowing unimpeded ground operation of the control lever from UNLOCKED to LOCKED and engagement of the rudder lock strut. With the aircraft on-ground and electrical power available (solenoid snibs retracted) selection of the control lever to LOCKED will cause an adjustable cam on the levers shaft to connect to pole A of a two-pole microswitch WM2 to the retract field winding of the actuator. The actuator then retracts to arm the system for rudder lock engagement; this occurs when the rudder pedals are centralized. When fully retracted the actuator limit switches changeover in readiness for a subsequent extend command. An UNLOCK selection similarly causes actuator extension to disengage in the rudder lock strut. Controls locked warning A CONTROL LOCKS warning light (red) on Panel IP will illuminate: a. if the controls locking lever is out of its UNLOCKED detent or if the rudder gust lock strut is not in a fully disengaged position. Page 1

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9.17 RIGGING AND BALANCING CONTROLS 9.17.1 RIGGING - INTRODUCTION

When applied to control systems the term 'rigging' is used to describe the practice of truing and checking the system to ensure that the flying controls operate correctly. The objective of rigging is to have the cockpit control in neutral at the same time as the control surface is in neutral. Rigging a control system ensures that: 

The pilot's control is in the correct relationship to the relevant control surface.



The control surface moves in the correct sense and to its designed maximum travel position in either direction.



Friction in the system is within acceptable limits.

The rigging and adjustment of the system is carried out: a. At specified intervals as laid down in the relevant aircraft servicing publication. b. After disturbing any part of the control system, including the control system. 9.17.2 CHECKS BEFORE RIGGING

a. Before operating any flying control system in an aircraft, first check that there are no obstructions that could damage the control surface when it is moved. It is also important to display warning notices informing personnel of the possibility of movement of the control surface. Inform personnel working in the vicinity of a control system when you are about to operate it. b. In rigging an aircraft control system it is sometimes necessary to level the aircraft both laterally and longitudinally – to put it into the rigging position, as described. The appropriate aircraft maintenance manual will state on what occasions, if any, this is necessary. c. Before starting to rig a flying control system it is advisable to ensure that all parts of the system and the control surfaces are serviceable. There is little merit in rigging a control system only to discover, subsequently, that some parts have to be replaced. Thus cables and tubes should automatically be examined for wear and corrosion, and other components for freedom of movement, security of attachment and so on. Replace components as necessary before continuing.

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9.17.3 RIGGING PROCEDURE 9.17.3.1

Establishing the Neutral Setting

The first action is to set the cockpit control to neutral and to lock it in this position, using the equipment provided for the particular system. The rest of the control run is then adjusted to the neutral setting and locked in that position, often by using rigging pins. Generally speaking, control surfaces are in neutral when they are in line with the main surface to which they are attached. An exception to this is where the trailing edge of the aileron is set a specified amount below the mainplane trailing edge. This setting is known as aileron droop.

9.17.3.2

Rigging Pins

Rigging pins are issued in sets, the type and number depending upon the aircraft and also upon the specific control run being rigged. The type, number and positions of rigging pins in the aircraft's system are shown in diagrams of appropriate aircraft maintenance manual. The first pins, called the No. I or master pin, is fitted at the cockpit end of the control run and, in conjunction with the cockpit control neutral setting bar, secures that end of the system in neutral. Between these two items, there may be an adjustable link that has to be set at the correct length. By adjusting the control cable and tubes, holes in idler gears or levers can be made to align with corresponding holes in the airframe structure; rigging pins are then used to join these two holes, thereby positively locating and locking the control system in neutral. When all the rigging pins have been fitted in this way, that particular control run has been adjusted to, and locked in, neutral. This setting may be checked by using setting gauges.

Figure 42 Page 1

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The next stage is to remove and then refit each rigging pin in turn to ensure that this can be done without strain. This indicates that the system has been set up satisfactorily, and that there is no backlash in the system; this is particularly important where the system is cable operated. Finally, it is vital to check that the complete set of rigging pins are removed from the aircraft on completion of the work. Note: There have been many accidents or near accidents attributed to failure to remove rigging pins, or the use of incorrect items to lock controls in neutral. In one particular incident, a new aircraft was taxiing out from the manufacturer for delivery to the customer. Whilst carrying out the “full and free” control test prior to take off the pilot felt a restriction in the aileron controls. When the aircraft taxied back to the hangar, a bolt was found inserted in the captain‟s control rigging pin hole. Obviously someone had used this in preference to the correct rigging pin. The correct checks had obviously not been carried out and the “rigging pin/bolt” not removed. (see diagrams following).

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9.17.4 CONTROL SURFACE SETTING GAUGES

Where control surface setting gauges are provided, they are used to check the neutral and maximum travel position of controllable aerofoil surfaces. Each gauge is manufactured for use with one specific surface. The gauge is firmly attached to a fixed part of the aircraft, next to the movable surface with which it is associated. With the controls set at neutral, the trailing edge of the control surface should coincide with the neutral mark on the gauge. Now move the control surface to the maximum travel position, in either direction, and see if the trailing edge of the control surface coincides with the appropriate mark on the gauge. The control surface movement can be quickly and easily adjusted with the gauge in position by restricting the mechanical stops. 9.17.5 CHECKING FOR SENSE OF MOVEMENT

Having established the neutral position of the control system, the next stage is to ensure that the control run being rigged operates the control surface in the correct sense. This is clearly vital; inadvertent cross-over of connections would reverse the control surface movement with possible disastrous results. The sense of operation can be readily checked by two tradesmen - one at the control in the cockpit and the other at the control surface, if you are not sure of the relationship between control movement and the corresponding control surface movement.

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Mechanical StopsThe next check is to ensure that the control surface moves to its designed maximum travel position, in both directions, when moved by the cockpit control. The maximum travel of a primary control surface is limited in either direction by mechanical (limit) stops. These stops are fitted to limit the control surface movement due to excessive travel. In a manual system, the limit stops are usually located near the control surface, and a second pair of stops, known as 'override stops' are fitted to limit the pilot's control movement should the main stop fail. Override stops are adjusted to a specified clearance under normal operating conditions. In powered control systems, the mechanical stops are-located on the input (PFCU); usually they are located next to the pilot's control in 'the cockpit, thus limiting the control system movement from that position. During the rigging procedure, the main mechanical (limit) stops may need to be re-set to ensure that the control surface reaches, but does not exceed, its maximum travel position. The maximum travel position of a control surface can be checked in a variety of ways using the instruments detailed later. Figure 44 However, most modern aircraft use control surface setting gauges for this purpose. We have now rigged the control system and also checked that it operates in the correct sense; and we have set the limit stops to give the maximum required travel in both directions. The next stage is to check the system for resistance to movement from rest and also the force required to maintain the speed of movement when the control system is operated.

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9.17.6 CHECKING FOR STATIC AND RUNNING FRICTION

The resistance to movement of a control system may be due to lack of lubrication, misalignment, or slight faults in bearing surfaces. This resistance can be measured using a spring balance attached to the cockpit control; an example of this is shown. The pull on the spring balance is in the direction that the control would normally move. Note the reading on the spring balance when the control starts to move from rest. This force is known as the breakout force, and represents the amount of static friction in the system. Once the control system is moving, the force required to keep it moving is less than the breakout force. The spring balance indicates this reduced force, which represents the running friction. The amount of static and running friction permissible in any given aircraft control run must not exceed the limits laid down in the appropriate aircraft maintenance manual. Insufficient lubrication will, of course, increase the friction of any parts that rub together.

Figure 45 9.17.7

CHECKS AFTER RIGGING

After any adjustment to a flying control system, it is necessary to carry out a functional test of the system and to carry out a visual check of the complete system; start from the cockpit and finish at the control surface. The following are typical of the checks to be carried out. a. Carry out a functional test, ensuring that no part of the system fouls the airframe structure when operated over the full range of movement.

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b. Check that turnbuckles, adjustable end fittings and limit stops are in safety and locked. c. Examine all parts of the system and supporting structure for security of attachment and check that shackle pins and nuts are correctly split-pinned. d. Check cable alignment around pulleys. e. Lubricate the system as necessary in accordance with servicing instructions for the system. f.

Examine the control surface itself to ensure that it has not been damaged in any way.

g. Check to ensure that no tools or other 'foreign objects' have been left within the system to become a FOD hazard. h. The final check is always a duplicate check - by a suitable qualified, engineer. 9.17.8 DUPLICATE CHECKS

In the interest of safety, all work on, and the functioning of, aircraft control systems must be checked twice, each time by a suitably authorised qualified person. Duplicate checks are divided into parts: a. b. c. d.

check for correct assembly and locking, and function range of movement sense check.

The term 'control systems' applies to all engine, undercarriage, flying and associated control systems and equipment directly affecting the safety of the aircraft. Full regulations concerning duplicates are described in module 10. 9.17.9 PRIMARY CONTROL SYSTEMS - EXAMPLE OF RIGGING

The method of rigging and the procedures to be adopted when adjusting a specified control run on a given aircraft will be detailed in the appropriate maintenance manual. This must always be consulted. However, to help you, and to give guidance on the sort of things you can expect to find, an example of a typical rigging procedure is given in the following paragraphs. To rig an aileron system, the procedure may be along the following lines: a. Carry out the checks before rigging described in paragraph 8.2 b. Slacken the control cables throughout the system. c. Lock the pilot's column in neutral and ensure that the control chains shown are equally disposed around the top and bottom sprockets. d. Set the aileron operating sprockets to neutral using rigging pins as necessary.

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e. Ensure that the chains are correctly positioned and then adjust the turnbuckles evenly to tension the control cables; a tension-meter may or may not be required to check cable tension, depending on whether or not the system is a regulated one. f.

Adjust the operating rods until the ailerons are in line with the trailing edge of the main plane or as specified (or the neutral setting on the setting gauge, if one is provided).

g. Remove the control column locking device and also the rigging pins, if fitted. h. Operate the ailerons, checking for freedom of movement and that they move in the correct sense relative to the control column movement. i.

Measure the range of movement of the ailerons and adjust the limit stops until the range is as specified in the maintenance manual. If the limit stops are not adjustable, and the range of movement is incorrect, replace the stops.

j.

Then, using a spring balance, check the control system for static and running friction. k. Carry out the necessary checks after rigging. l.

Arrange for a duplicate check to be carried out.

To rig other primary control systems (i.e. elevators and rudders) a procedure similar to that outlined above is carried out. Remember, however, that each system is peculiar to the aircraft in which it is installed and the need to consult the aircraft maintenance manual should be obvious. 9.17.10

RIGGING A TUBE-OPERATED CONTROL SYSTEM

To rig a primary control system, in which light alloy tubes are used, the procedure may be similar to that described below: a. Carry out the checks before rigging as in paragraph 8.2. b. Set and lock the pilot's control in the neutral position and disconnect the control tubes. c. Commence by attaching the forward control tube to the pilot's control. d. Fit the appropriate (master) rigging pin. e. Connect the tubes in sequence from the cockpit to the control surface, fitting the rigging pins in the appropriate housings. f.

Connect the control surface and adjust to the neutral setting.

g. Remove the cockpit control locking device and the rigging pins. h. Operate the control, checking for freedom of movement and that it moves in the correct sense in relation to the control surface. i.

Measure the range of movement and, as necessary, adjust the limit stops.

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Check the system for friction.

k. Carry out the necessary checks after rigging. l. Arrange for a duplicate check to be carried out. 9.17.11

RIGGING A POWERED FLYING CONTROL SYSTEM

As you would expect in this type of system, a power unit of some sort acts to move, or to assist the movement of, the control surface in response to movement of the cockpit control. A powered flying control system has other units, such as a trim actuator, artificial feel unit and yaw damper, fitted to it. But we are not concerned with any of these units at this time. Our concern is with rigging the manual part of the system, and here the same principles apply as in the other examples. In a typical rudder powered control system. Control tubes are used from the cockpit to the bell crank lever, to which the artificial feel unit is attached. The remainder of the input system to the PFCU is cable-operated, apart from the yaw damper lever group. It is convenient in rigging a powered flying control system to split the operation up into three stages as described below. Stage 1. This stage describes the operations required to rig the rudder controls in the cockpit. This part of the system uses tubes that, as we have seen, form a rigid link. a. Disconnect all control tubes and set the pilot's control - i.e. the rudder pedals to the specified initial setting (neutral). Lock the pedals in neutral with the tools provided. b. Connect the forward control tube to the rudder lever, make any necessary adjustments and insert the master rigging pin in the appropriate housings. Continue to build up the system in the cockpit; connect the various control tubes to their corresponding lever, adjusting as necessary; fit the subsidiary rigging pins in their correct positions, thereby locking the levers in neutral. Stage 2. This stage describes the operations required in that part of the control run situated in the centre of the fuselage. a. Set the control lever on the cable tension regulator vertical by inserting a subsidiary rigging pin. b. Set the artificial feel units as described in the aircraft maintenance manual. c. Set the trim actuator to neutral and connect to the artificial feel unit. d. Fit the control cables to the cable tension regulator, and run the cables around their pulleys. e. Connect the cables to the tie-rods and feed the cables around their pulleys to the rear part of the fuselage.

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Stage 3. This final stage describes the operations required to rig the control system up to the control surface (rudder) itself. a. Connect the cables to the cable quadrant and to their respective tie-rods. b. Set the cable quadrant to neutral by inserting a subsidiary rigging pin. c. The yaw damper can now be connected into the system, control tubes connecting this unit to the cable quadrant and to the PFCU. d. Adjust the system, as necessary, as detailed in the aircraft maintenance manual. e. Connect the link rod to the rudder operating lever and adjust as necessary. f.

After tensioning the cables as described earlier, remove and then re-insert each rigging pin in turn to ensure that the system is correctly adjusted to the neutral position.

g. Unlock the control circuit by removing all rigging pins and the rudder pedal locking device. h. Check the system for friction. i.

Carry out the necessary checks after rigging.

j.

Arrange for a duplicate check to be carried out.

Rigging of the PFCU will be considered later.

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RIGGING OF TRIMMING TAB SYSTEM

As we saw, manually operated primary control surfaces can be trimmed so that the aircraft flies without any load on the control column or rudder bar. Power assisted and power operated control systems, which can be manually controlled may also, have controllable trimming tabs fitted to the primary control surfaces. The trimming tab may be moved by a screw jack mechanism attached to the trimming tab-operating arm and operated from the cockpit by turning the handwheel. An indicator plate adjacent to the hand-wheel indicates the degree of trim. Trimming tab systems may be cable, control tube or electrically operated. These systems are trued-up in a manner similar to that of a primary control system. To rig a cable-operated system, slacken the cables, check the handwheel over its full travel, and then set it to the mid-travel position. Check that the graduated plate shows the correct degree of trim and then proceed as follows: a. Set the screw jack in mid-way (neutral) position and ensure that the operating chain is equally disposed around the sprocket. b. Tighten the turnbuckles evenly, and obtain the correct tension. c. With the elevator in line with the tailplane, adjust the operating rod until the tab is in line with the elevator. d. Operate the trimming tab, checking for freedom of operation and that it moves in the correct sense in relation to the hand-wheel. e. Measure the travel of the trimming tab, which must be as specified. f.

Carry out the necessary checks after rigging.

g. Arrange for a duplicate inspection to be carried out.

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9.18 STALL WARNING AND PROTECTION If an aircraft is flown at a high angle of attack, lift will be increased. However, if the angle of attack is increased to too great an angle, the airflow over the wings will separate and become turbulent. This will cause the lift to instantly fall to a very low value and the wing (or aircraft) is said to have stalled. The design of some aircraft will give an inherent indication of an approaching stall condition. The airflow or wake leaving the wings will become progressively more turbulent as the stall is approached. This turbulent wake will strike the airframe structure or tail-plane causing a condition known as "buffet". The pilot will normally recognise this as an indication of an impending stall and take appropriate action to prevent it, i.e. push the control column forward to reduce the angle of attack. Many aircraft do not have this inherent warning characteristic of buffet; therefore these aircraft require a system to warn the pilot of an impending stall. There are several stall-warning systems in use.

9.18.1 STALL WARNING SYSTEMS 9.18.1.1

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This system is common on light aircraft. In this system a plenum chamber is mounted in the wing leading edge. This is covered and sealed by an adjustable plate that acts as part of the leading edge. The plate is adjusted so that in normal flight attitude a slot in the plate coincides with the stagnation point of the wing. The plenum chamber is connected by tube to a horn/reed assembly in the cabin. As the angle of attack is increased the slot in the adjustable plate effectively moves up from the stagnation point into an area of progressively lower air pressure. The slot is so positioned that it reaches a low-pressure area sufficient to draw air through the horn/reed assembly, which will emit a noise and alert the pilot to an impending stall. 9.18.1.2

Electric Stall Warning System

This is typical of a system fitted to larger aircraft. This is a simple system that employs a micro-switch (transducer), operated by a vane. The transducer is mounted in the wing leading edge such that the operating vane is at the stagnation point during normal flight. Therefore no air-loads are imposed on the vane and it is not deflected from its null position. Page 1

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As the aircraft angle of attack increases the transducer-operating vane effectively moves up and away from the stagnation point. The air-loads on the vane will increase until at a set angle of attack they overcome a spring pressure to deflect the vane and close the micro-switch contacts. This completes a circuit to illuminate a warning light and sound a warning horn. This should occur just prior to reaching the stall. These systems are found on relatively simple or small aircraft. Larger and more complex aircraft generally require a more sophisticated system that will do more than just warn of impending stall. This is termed a stall protection system. 9.18.2 STALL PROTECTION SYSTEM

9.18.2.1

System Functions



Stall Warning - As with the previous system this tells the pilot that he is approaching a stall condition.



Stall Identification - This detects an imminent stall and automatically takes action to prevent the stall occurring, i.e. the stick is automatically pushed forward by the system. This may be achieved by a hydraulic or pneumatic jack acting on the elevator control system.



Auto Ignition - In some aircraft, particularly rear engine aircraft, disturbed airflow entering the intakes may cause the engines to flame out near or at the stall. To prevent this an auto ignition circuit may be initiated on a stall warning/identification condition to prevent this. Flap/Slat/Krueger Flap Modulation - As flap, slat and Krueger flap position affect the stall angle the stall protection system may include the monitoring of their position and delay the initiation of stall warning.

 9.18.3 TYPICAL SYSTEM COMPONENTS



Stall Warning Sensors - There are several designs in use. They may be mounted on the main-planes or side of the fuselage. They are normally duplicated, each providing a signal to a duplicated system.



Stall Warning Computer - Receives signals from the sensors and initiates warnings or control movements.



Stick Shaker - The Main stall warning device. An electrically driven, out of balance rotor, which shakes the control column when a stall warning condition, is detected.

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

Stick Pusher - A hydraulic or pneumatic ram which pushes the control column forward when a stall identification condition is sensed. It may usually be overridden by higher than normal pilot force.



Ground/Flight Sensing - To prevent unwanted operation of the system on the ground a circuit through the landing gear weight switches disarms the stall protection system on the ground.



Test - A pre-flight test facility is built into the system.



Mach Sensing - Speeds over the aircraft critical Mach number may cause high-speed stall or flame out. To prevent this an input to the computer from mach switches or the air data computer may be included to give a stall warning at high mach numbers.

9.18.4 ACTUAL STALL PROTECTION SYSTEM

The following is the description of an actual system used on a large passenger aircraft. The stall protection system provides the following during the various phases of approach to the stall: 1. Automatic ignition on all four engines. 2. Stall warning by the operation of a stick shaker on each control column. 3. Stall identification by the sounding of a klaxon for each system, allowed by operation of a ram to move the control columns forward.

9.18.5 INCIDENCE PROBES

Four slotted conical probes, are mounted, two on either side of the forward fuselage, and project into the air stream. Each probe can rotate about its own axis through 50º in pitch, 4º of which are above fuselage datum. The probe detects the direction of airflow and transmits to the computer unit a voltage, picked off from potentiometers, proportional to the angle between the airflow and the fuselage datum. When the aircraft angle of incidence is steady, pressure acts equally on the two probe slots, but as the angle of incidence changes, differential pressures are set up which, applied to the opposite sides of a paddle wheel, cause the wheel to rotate the probe until the pressures are again equal, i.e. the direction of flow bisects the angle between the slots. Ice protection for the probes is provided by heaters supplied from the No 1 and No 2 essential 28-volt dc supply. The left probe heaters are controlled by the first pilot's pressure head heater switch and the right probe heaters by the No 2 autopilot pressure head heater switch.

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Separate ammeters on the engineer‟s engine panel monitor the heater supplies. When the aircraft is on the ground the current is limited by a resistor in series with the power supply.

Ferranti Probe Stall Warning System Figure 48

9.18.6 NITROGEN SYSTEM

Nitrogen is stored at 1,500 psi in a reservoir. Nitrogen is piped via a stop valve to a pressure reducing valve and non-return valve to a low-pressure reservoir. Gauges monitoring the high and low pressure are on the right sill panel and forward roof panel respectively. A relief valve in the low-pressure line vents at 52 p.s.i. to prevent too great a pressure build-up in the system. Low-pressure nitrogen is fed to solenoid valve A and from there through solenoid B to a control ram, which operates on the control column linkage. A dump valve operated from a STALL DUMP VALVE lever on the centre console is coupled to this part of the circuit, and when the lever is set to DUMP, pressure in the line is released and prevents further operation of the stick pusher until the lever is reset.

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9.18.7 AUTOMATIC IGNITION

Automatic ignition is signalled from the lower two of four Ferranti-type probes located on each side of the forward fuselage. It is switched on at a predetermined incidence, which is modified by slat position and Mach number and remains on as long as the incidence is at or above this value. Indication of igniter operator is shown on the engine start panel. The system is brought into operation earlier whenever the slats are in or whenever 0.74M is exceeded. The system, which is physically shared with, but electrically isolated from the stall identification system, consists of two computer units, two mach switches and two angle of incidence probes. One of the two igniters on each engine is coupled to its associated computer, thus providing a completely duplicated and independent system. 9.18.8 STALL WARNING

The stall warning function, is provided by two duplicated systems, No 1 and No 2, each containing a computer unit, a lift rate modifier, an angle of incidence probe, and a stick shaker motor. Stall warning is signalled by the upper two of the four fuselage mounted probes. One probe is dedicated to shaker system. The warning is signalled at a predetermined incidence, which is modified by a combination of flap position, slat position and rate of change of incidence. It remains in operation as long as the incidence is at or above this value. One stick shaker is mounted on each control column and is connected to respective computer unit and lift rate modifier, thus providing duplicated and independent indication of stall warning.

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Stick Shaker Figure 49 9.18.9 STALL IDENTIFICATION

Stall identification is provided by two duplicated systems, No 1 and No 2, each containing a lift rate modifier a solenoid operated valve, interlock relay and delay unit, a warning horn and an angle of incidence probe which is shared with, but electrically isolated from, the auto-ignition system. Identification of a stall is signalled by the two fuselage mounted probes, which signal auto-ignition. The signal occurs at a predetermined incidence set at a level that is always above the stall warning value. This predetermined incidence is modified by, a combination of flap positions, slat position and rate of change of incidence. The stall, identification system operates only if armed by a prior stall warning signal, and remains in operation as long as the incidence is at or above the modified level. Page 1

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The system reverts to normal operation once the stall-warning signal is cancelled by resuming normal flight. When the stick shakers operate, a priority circuit receives signals from each of the computer units of the stall warning system. The first signal received is passed to the stall identification interlock relays to arm the solenoid valves circuit. The signal from the stall identification probes is fed to the appropriate computer unit, and when the signal reaches a particular value, the unit supplies a 28 volt dc output. The value of the signal can be changed by combinations of the flap and slat position compensation. The signal is passed through the lift rate modifier so that a quick rate of change of the probe angle causes an advanced signal, provided that has been preceded for 0.7 seconds by a stick shaker signal. The computer unit output is passed through a priority circuit to the stall identification relay in the interlock circuit. Providing the sequence is correct, this completes the circuit to the solenoid valves that open to allow nitrogen to the rams that extend to move the control column forward. The warning horn in each system sounds when the respective stall warning and stall identification computer units both signal, which is simultaneous with control column movement. Both solenoid operated selector valves are opened by a stall identification signal. The opening of each valve is indicated by the associated red light on the overhead panel, and the subsequent movement of the ram is indicated by the STALL IDENT amber light adjacent to the airspeed indicators on each pilot's panel also coming on. The system is pneumatically powered from a HP nitrogen bottle that feeds the stick pusher ram through a reducing valve, an LP reservoir and the two solenoidoperated selector valves. A gauge on the forward roof panel indicates the pressure in the low-pressure reservoir and another on the right sill panel indicates the pressure in the HP bottle. Minimum HP pressure for flight is 500 p.s.i. When pressure falls to 32 PSI, the LP red light on the forward roof panel comes on.

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Stall Protection System Figure 50

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9.19 FLY BY WIRE

9.19.1 INTRODUCTION

Fly by wire is used on some aircraft to operate the controls. Instead of a conventional mechanical link between the pilot‟s controls to the control surfaces or powered control servo valves, the link is by an electrical or fibre-optic cable. The abbreviations Fly by wire (FBW) or Fly by Optical Wire (FBOW) are used.

9.19.2 PRINCIPLES OF FBW

FBW is a control system that receives inputs directly by electrical signals. The flying control actuators are electro-hydraulic design converting electrical signals into movement of a hydraulic ram. Many systems on the aircraft use electrical signals to automatically control the flight path. It is a natural development to integrate the pilot‟s input with these automatic controls. Correcting signals can be sent directly to the control actuator as well as those sent by the pilot.

9.19.3 PRINCIPLES OF FBOW

An optical fibre cable consists of multiple glass fibres, each about as thick as a human hair. The cable can carry pulses of light without amplification and without electromagnetic interference. One fibre can carry over 9,000 simultaneous signals. Fibre optics transmits information using: 

A light source modulated with information



A fibre optic transmission medium (cable)



An optical receiver to de-modulate the information

9.19.4 ADVANTAGES OF FBOW OVER FBW



Increased amount of information can be passed



Increased speed of transmission



Lighter in weight

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9.19.5 OTHER INPUTS TO POWERED FLYING CONTROL UNIT

The pilot is the main controller of the aircraft controls. There are, however, other inputs as follows: 

Auto-stabilisation (variable incidence tailplane)



Datum shift caused by operation of landing gear. The system can automatically make an input to the PFCU when the gear is lowered or raised.



Mach trim will deflect the tailplane or elevators to compensate for changes in aircraft attitude at high Mach Numbers due to rearward movement of the centre of pressure



Autopilot will be interfaced directly with the PFCU‟s.



Terrain Following Radar (TFR) – The system can process information on radar or radio height to the PFCU.



Inertial Navigation System (INS)



Instrument Landing System (ILS) – programmed automatic landing sequences can be fed directly into the control system.



Airspeed – The aircraft engines can also be controlled to give fully automatic programmable airspeed.



Position of other controls, including secondary controls such as flaps and LE flaps and slats.

9.19.6 777 FLIGHT CONTROLS - INTRODUCTION 9.19.7 GENERAL

The flight controls keep the aeroplane at the desired attitude during flight. They consist of movable surfaces on the wing and the empennage. The flight controls change the lift of the wing and the empennage. There are two types of flight controls: the primary flight control system and the high lift control system. 9.19.8 777 PRIMARY FLIGHT CONTROL SYSTEM

The primary flight control system (PFCS) uses a fly-by-wire control system with digital and analogue electronic equipment. It receives commands from the flight crew and the autopilot and causes the control surfaces to move. The PFCS controls the attitude of the airplane during flight. The control surfaces operated by the PFCS are: 

One aileron on each wing



One flaperon on each wing

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

Seven spoilers on each wing



One horizontal stabiliser



One elevator on each side of the horizontal stabiliser



One tabbed rudder.

9.19.9 HIGH LIFT CONTROL SYSTEM

The high lift control system (HLCS) uses a fly-by-wire control system with digital electronic equipment. It receives commands from the flight crew and causes the flaps and slats to move. Operation of the HLCS increases the wing lift so the aeroplane can takeoff and land at lower speed and higher weight. The high lift devices operated by the HLCS are: 

Seven leading edge slats on each wing



One Krueger flap on each wing



One single slotted outboard flap on each wing



One double slotted inboard flap on each wing.

Operation of the HLCS also causes the ailerons and the flaperons to move. They droop on both wings when the high lift devices extend. 9.19.10

BENEFITS OF THE FLY-BY-WIRE SYSTEM

The fly-by-wire design of the flight controls permits: 

A more efficient structure design



Increased fuel economy



A smaller vertical fin



A smaller horizontal stabiliser



Reduced weight



Improved controls and protections.

9.19.11

ABBREVIATIONS AND ACRONYMS

ACE· actuator control electronics ACMS· aeroplane condition monitoring system ADIRS· air data inertial reference system ADIRU· air data inertial reference unit ADM· air data module Page 1

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AFDC· autopilot flight director computer AFDS· autopilot flight director system AIMS· aeroplane information management system ARINC· Aeronautical Radio, Inc. BAP· bank angle protection B/D· backdrive CMCS· central maintenance computing system CPU· central processing unit EDIU· engine data interface unit EHS· electro-hydraulic servo valve EICAS· engine indication and crew alerting system FCDC· flight controls direct current FMCS· flight management computer system FSEU· flap/slat electronics unit HLCS· high lift control system LIB· left inboard LOB· left outboard LVDT· linear variable differential transformer MCP· mode control panel MFD· multi functional display PCU· power control unit PDU· power drive unit PFC· primary flight computer PFCS· primary flight control system PMG· permanent magnet generator PSA· power supply assembly PSEU· proximity sensor electronic unit RIB· right inboard ROB· right outboard RVDT· rotary variable differential transformer SAARU· secondary attitude air data reference unit SOL· solenoid Page 1

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SOV· shutoff valve STCM· stabiliser trim control module TAC· thrust asymmetry compensation WEU· warning electronic unit WOW· weight on wheels 9.19.12

PRIMARY FLIGHT CONTROL SYSTEM - INTRODUCTION

Purpose The primary flight control system (PFCS) controls the aeroplane flight attitude in relation to the three basic axes: 

Longitudinal



Lateral



Vertical.

777 Primary Flight Controls Figure 51

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Roll Control The roll control uses the ailerons, flaperons, and spoilers to control the aeroplane attitude about the longitudinal axis. During a bank of the aeroplane, the aileron and flaperon on one wing move in an opposite direction from the aileron and flaperon on the other wing. The spoilers move up only on the down wing and do not move on the up wing. Pitch Control The pitch control uses the horizontal stabiliser and the elevator to control the aeroplane attitude about the lateral axis. The stabiliser controls long term pitch changes. The elevator supplies short term pitch control. Yaw Control The yaw control uses the rudder to control the aeroplane attitude about the vertical axis. The rudder has a tab, which moves to increase the effectiveness of the rudder. Speedbrakes The PFCS also includes the speedbrakes. In addition to roll control, the spoilers also act as speedbrakes in the air and on the ground. They deploy on both wings to increase drag and to decrease the amount of lift the wings supply. 9.19.13

PFCS – GENERAL DESCRIPTION

The pilots or the autopilot commands control the PFCS. The pilots can override the autopilot.

9.19.14

MANUAL OPERATION

Position transducers change the pilots' manual commands of the control wheel, the control columns, the rudder pedals, and the speedbrake lever to analogue electrical signals. These signals go to the four actuator control electronics (ACEs). The ACEs change the signals to digital format and send them to the three primary flight computers (PFCs). The PFCs have interfaces with the aeroplane systems through the three flight controls ARINC 629 buses. In addition to command signals from the ACEs, the PFCs also receive data from: 

The airplane information management system (AIMS)



The air data inertial reference unit (ADIRU)



The secondary attitude air data reference unit (SAARU).

The PFCs calculate the flight control commands based on control laws and flight envelope protection functions. The control laws supply stability augmentation in the pitch and yaw axes and flight envelope protections in all three axes. The digital command signals from the PFCs go to the ACEs. Page 1

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The ACEs change these command signals to analogue format and send them to the power control units (PCUs) and the stabiliser trim control modules (STCMs). The ACEs and the PCUs form control loops, which control the surfaces based on the PFCs commands. One, two or three PCUs operate each control surface. One PCU controls each spoiler, two PCUs control each aileron, flaperon, and elevator, and three PCUs control the rudder. The PCUs contain a hydraulic actuator, an electrohydraulic servo valve, and a position feedback transducer. When commanded, the servo valve causes the hydraulic actuator to move the control surface. The position transducer sends a position feedback signal to the ACEs. The ACEs then stop the PCU command when the position feedback signal equals the commanded position. Two STCMs control hydraulic power to the motors and brakes of the horizontal stabilizer. 9.19.15

AUTOPILOT OPERATION

The PFCs receive autopilot commands from all three autopilot flight director computers (AFDCs). The PFCs use the autopilot commands in the same manner as the pilots' manual commands. In addition, the PFCs supply the backdrive signals to the backdrive actuators through the AFDCs. The backdrive actuators move the control wheels, control columns, and rudder pedals in synchronisation with the autopilot commands. The movement of the flight deck controls supplies visual indications to the flight crew. 9.19.16

PFCS MODES OF OPERATION

The PFCS has three modes of operation: normal, secondary, and direct. 

Normal mode operates when the necessary data are available for the PFCs and the ACEs. All the control laws, protection functions, and the AFDCs operate.



When the PFCS detects the loss of important air and attitude data, the PFCS operation changes to secondary mode. The PFCs and the ACEs operate but the PFC control laws and protection functions downgrade. The autopilot cannot operate in secondary mode.



In direct mode, the PFCs are not used. The ACEs set the position of the control surfaces in direct response to analog pilot control inputs.

9.19.17

FLIGHT DECK CONTROLS

The control wheel, control column, and rudder pedal position transducers are below the flight deck floor in the forward equipment centre. The speedbrake lever position transducers are in the control stand. The location of these transducers is shown in other sections. Page 1

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MAIN EQUIPMENT CENTRE

The E1, E2, and E3 racks contain most of the electronic equipment of the PFCS. The E1 rack contains: 

The left PFC



The L1 ACE



The L2 ACE.

The E2 rack contains: 

The centre PFC



The centre ACE



The SAARU.

The E3 rack contains the ADIRU. Forward Cargo Compartment The E16 rack, forward of the forward cargo door, contains the right PFC. The E5 rack, aft of the forward cargo door, contains the right ACE. Control Surfaces Each PCU connects directly to its related control surface on the wing and the empennage. The ballscrew actuator of the horizontal stabiliser is in the stabiliser compartment. The location of the PCUs and the ballscrew actuator is shown in their specified sections. 9.19.19

PFCS – FLIGHT CONTROLS ARINC 629 BUS INTERFACES

a. General ARINC 629 digital data buses supply the principal means of communication among aeroplane systems. Three dedicated flight controls ARINC 629 buses connect the PFCS to: 

The three autopilot flight director computers (AFDC)



The two aeroplane information management system (AIMS) cabinets



The air data inertial reference unit (ADIRU)



The secondary attitude air data reference unit (SAARU).

Physical separation of the buses, and redundant LRUs, protects against multiple failures due to one event. b. PFCS Interface

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The three primary flight computers (PFCs) and the four actuator control electronics (ACEs) have interfaces with the flight controls data buses. The L PFC, C PFC and R PFC receive data from all three flight controls data buses but transmit data only on their on-side data bus. (On-side means that the relationship is with equipment of the same side. For example, the left bus is the on-side bus for the left PFC.) Each ACE receives data from all three PFCs through the three flight controls data buses. Each ACE processes control data from its on-side PFC. If this data is not valid, the ACE processes data from an alternate PFC. The ACEs process some data from the other PFCs at all times. For example, this occurs during data validation and voted commands. The ACEs transmit only on their on-side bus. c. AFDS Interface The autopilot flight director system (AFDS) interfaces with the PFCS through the autopilot flight director computers (AFDCs). Each AFDC transmits to its on-side flight controls data bus, but receives data from all three buses. The AFDCs receive backdrive commands, engagement status, and other data from the PFCs. The AFDCs transmit pitch, roll and yaw commands and engage requests to the PFCs. d. AIMS Interface The two AIMS cabinets receive data from all three-flight controls data buses, but normally transmit only to their on-side bus. During tests on the ground, the AIMS cabinets transmit also to the centre bus. The PFCS supplies information to the AIMS for: · The primary display system (PDS), (flight, synoptic, and EICAS displays) · The central maintenance computing system (CMCS) · The aeroplane condition monitoring system (ACMS) · The flight management computing system (FMCS) e. ADIRS Interface The air data inertial reference system (ADIRS) consists of the air data inertial reference unit (ADIRU) and the secondary attitude air data reference unit (SAARU). The ADIRU and SAARU supply air data variables and inertial data to the PFCs. The ADIRU receives data from all three buses and transmits data to the left and right flight controls data buses. The SAARU also receives data on all three buses, but transmits data only to the centre flight controls data bus. f. Air Data Modules

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Six air data modules (ADMs) supply pitot and static air data to the ADIRU and the SAARU. The ADMs transmit these data through the flight controls ARINC 629 buses. Figure: PFCS - FLIGHT CONTROLS ARINC 629 BUS INTERFACES PFCS - SYSTEMS ARINC 629 BUS INTERFACES General Many components of the PFCS transmit to and receive information from other aeroplane systems. Data on the systems ARINC 629 buses goes through the AIMS data conversion gateway function and then to the flight control ARINC 629 buses. The PFCS uses information from: · The flap/slat electronics units (FSEUs) · The proximity sensor electronics units (PSEUs) · The left and right systems card files. The PFCS transmits information to the left and right warning electronic units (WEUs). The PFCS also uses radio altimeter data supplied through AIMS. FSEU Interface The FSEUs supply flaps and slats signal to the PFCs for the gain functions of the control laws. These signals show the retracted or not retracted condition of the flaps and the slats. The FSEUs also supply a signal to the ACEs. PSEU Interface The PSEUs supply truck tilt signals and associated fault messages to the PFCS. The truck tilt signals are used together with the weight on wheels and radio altimeter functions to operate the auto speedbrake. Systems Card Files Interface The two systems card files supply signals from the hydraulic interface module (HYDIM) cards and the weight on wheels (WOW) cards to the PFCS. The HYDIM cards supply hydraulic systems condition signals to the PFCS. They also supply data about the truck tilt pressure sensors. The WOW cards supply air/ground signals to the PFCS. These signals supply air/ground information to the PFCS. WEU Interface The PFCS sends stabiliser and rudder trim position signals to the WEUs. The WEUs supply a takeoff warning if the stabiliser is out of green band or the rudder trim is out of normal limits. The PFCs receive stall data from the WEUs for stall protection. Page 9-101 Issue 1

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Engine Data Interface Units The left and right engine data interface units (EDIUs) receive the N1 speed and calculate the thrust for each engine. The EDIUs supply these data to the AIMS and to the PFCS for the thrust asymmetry compensation (TAC) function. Radio Altimeter Interface The three radio altimeters supply information to AIMS for use by the PFCS for the flare function during manual landing. The PFCS inhibits the radio altimeter test when ground speed is more than 40 knots and during flight. Indications The status message PFCS INTERFACE shows because of one of these faults: · There is a disagreement with the flap discrete signals in two or more ACEs · Only two of the four FSEU channels are available · Only two of the four digital WES channels are available · Only one of the two truck tilt pressure data sources is available · Thrust data from one of the left or right EDIU channels does not agree with the others · Data from the two WOW cards does not agree. These conditions cause the message to show in normal, secondary, and direct modes. Figure: PFCS - SYSTEMS ARINC 629 BUS INTERFACES PFCS - ANALOG INTERFACES General All analogue interfaces with the PFCS go to the ACE. The primary inputs/outputs are: · Rudder trim selector · Manual trim cancel switch · Pitch trim switches · Flight control position transducers · Flight control force transducers · FSEUs · Primary flight computers DISC/AUTO switch · Thrust asymmetry compensation switch · AIMS cabinets Page 9-102 Issue 1

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· PCUs. Rudder Trim Selector and Manual Trim Cancel Switch The rudder trim selector and the manual trim cancel switch supply signals to the ACEs. These signals show the pilot commands for rudder trim. Pitch Trim Switches The pitch trim switches supply signals to the ACEs to show the pilot pitch trim commands. Flight Control Position Transducers The flight control position transducers supply electrical inputs to the ACEs. They show the position of the: · Control wheel · Control column · Rudder pedals · Speedbrake lever. Flight Control Force Transducers The pitch and roll force transducers supply signals to the ACEs. The signals show when the pilot applies a force to the control wheel or control column. FSEUs

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INTENTIONALLY BLANK

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JAR 66 CATEGORY B1 MODULE 11.10 FUEL SYSTEMS

CONTENTS 10 SYSTEM LAY-OUT....................................................................... 10-1 10.1 10.2

10.3

10.4 10.5

10.6

10.7 10.8 10.9 10.10

10.11

10.12 10.13

10.14

RIGID TANKS ................................................................................ 10-1 Rigid Metal Tanks ....................................................................... 10-2 FLEXIBLE FUEL TANKS ................................................................. 10-3 10.2.1 Tank Coverings ............................................................. 10-3 10.2.2 Self-Sealing Coverings .................................................. 10-4 10.2.3 Attachments and Fittings ............................................... 10-4 INTEGRAL FUEL TANKS ................................................................ 10-5 10.3.1 Tank Numbering ........................................................... 10-5 10.3.2 Water Draining .............................................................. 10-9 10.3.3 Water Scavenge System ............................................... 10-10 SUPPLY SYSTEMS ......................................................................... 10-11 ENGINE FUEL FEED ...................................................................... 10-11 10.5.1 Design Requirements of an Aircraft Fuel Feed System . 10-11 10.5.2 Engine Fuel Feed (Multi Tank and Booster Pumps) ...... 10-12 10.5.3 Engine Fuel Feed (Collector Tanks) .............................. 10-12 10.5.4 Engine Fuel Feed (Fuel Cells) ....................................... 10-13 FUEL FEED COMPONENTS .............................................................. 10-14 10.6.1 Fuel Pumps (Booster Pumps) ....................................... 10-14 10.6.2 Jet Pumps ..................................................................... 10-16 10.6.3 Sequence Valves .......................................................... 10-18 10.6.4 Transfer Valves ............................................................. 10-18 L.P. Valve ................................................................................... 10-19 Cross Feed Valve ....................................................................... 10-19 APU FUEL FEED .......................................................................... 10-20 DUMPING, VENTING AND DRAINING................................................ 10-22 DUMPING (JETTISON).................................................................... 10-22 THE VENT SUB-SYSTEM ............................................................... 10-25 10.10.1 General ......................................................................... 10-25 10.10.2 Venting Due to Heat ...................................................... 10-25 10.10.3 Unpressurised System Venting ..................................... 10-25 10.10.4 Pressurised Fuel Tanks................................................. 10-25 10.10.5 Float Valves .................................................................. 10-28 10.10.6 Vent Pipe Drains ........................................................... 10-29 CROSS-FEED AND TRANSFER ....................................................... 10-31 10.11.1 Auto – transfer .............................................................. 10-31 10.11.2 Manual – transfer .......................................................... 10-32 INDICATIONS AND WARNINGS ............................................... 10-33 FUEL LEVEL SENSING................................................................... 10-36 10.13.1 High Level Sensing ....................................................... 10-36 10.13.2 Overflow Sensing .......................................................... 10-37 10.13.3 Low Level Sensing ........................................................ 10-37 10.13.4 Calibration Sensing (Fuel Trim only) ............................. 10-37 10.13.5 Under Full Level Sensing .............................................. 10-37 FUEL QUANTITY SYSTEM MEASUREMENT AND INDICATION ............. 10-37

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10.15 PRINCIPLE OF CAPACITANCE GAUGING ......................................... 10-38 10.16 FUEL QUANTITY INDICATING SYSTEM ............................................ 10-38 10.16.1 Capacitance Index Compensator .................................. 10-39 10.16.2 Measurement................................................................ 10-42 10.17 REFUELLING AND DE-FUELLING .................................................... 10-42 10.18 REFUELLING ................................................................................ 10-42 10.18.1 Pressure Refuel – Functional Description ..................... 10-45 10.19 DEFUELLING ................................................................................ 10-48 10.20 LONGITUDINAL BALANCE FUEL SYSTEMS ..................................... 10-48 10.21 SUPERSONIC FLIGHT FUEL TRANSFER .......................................... 10-49

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10 SYSTEM LAY-OUT The purpose of the fuel system is to store and deliver fuel to the engines and the apu. An aircraft must be able to carry sufficient fuel to enable the engines to operate over long periods. to meet this requirement there must be some way of storing this fuel safely and supplying it to the engines in a suitable condition and at a controlled rate. A typical fuel system therefore will consist of a number of tanks, fuel lines, connections and fittings, which are compatible with all types of fuel meeting engine and apu specifications. Often, the fuel system is subdivided into storage, refuelling, distribution, transfer, venting and indicating subsystems. The following example of a system layout is for a typical large commercial twin aircraft. The number of tanks and system complexity will vary from aircraft to aircraft. Clearly a four-engine aircraft will have more components than a twin. The figure shows a typical fuel cell layout

Typical Fuel Cell Layout Figure 1 NOTE: For additional range, some operators will install centre tanks, these are offered as optional on most single isle and wide bodied aircraft.Fuel Tanks Fuel tanks normally fall into three categories of construction:  Rigid  Flexible  Integral 10.1 RIGID TANKS These are normally made from metal or plastic material, they are fitted internally where space permits. Flexible fuel tanks have an advantage over rigid tanks, because they can be shaped and fitted into odd shaped spaces where rigid tanks cannot be fitted. In general, flexible tanks are lighter and easier to handle and store than rigid tanks. Integral fuel tanks are of rigid construction because they are part of the airframe structure. They are not independent items like the other tanks. B1 Mod 11.10

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Whatever the construction method, fuel tanks should be shaped so that almost all the fuel is available to the engine. Awkward pockets which prevent fuel from leaving the tank are undesirable and are avoided if possible. 10.1.1 Rigid Metal Tanks

Typical Rigid Internal Fuel Tank Figure 2 Fuel tanks are made in shapes and sizes to fit the spaces available in each particular airframe and therefore the size and shape of the fuel tanks will not be the same for all aircraft. Metal fuel tanks are constructed from aluminium alloy, stainless steel or tinned steel and they are riveted, welded, or soldered together. The tank is a light structure which is strengthened by the use of internal stiffeners, angle pieces and by incorporating baffles to give strength and which are necessary, in large tanks, to reduce the effects of fuel surge caused when the aircraft manoeuvres. Secure attachment of a rigid tank within the airframe may be achieved by built-in padded cradles and padded metal straps. The cradle is shaped to match the contours of the tank and the straps secure the tank to its cradle. Each tank will have the brackets, strap guides and fittings to match the aircraft structure into which the tank is to be fitted. It must be stressed that very few aircraft over 5,700 kg would utilise metal rigid tanks, except when long range tanks are fitted in the cargo hold, i.e. commercial IATA LD6 containers, etc.

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10.2 FLEXIBLE FUEL TANKS Flexible fuel tanks may be constructed with thin and very flexible walls (called bag tanks) or they may be made of thicker less flexible material. These tanks are made in shapes to fit particular spaces in the aircraft structure and their flexibility enables the tanks to be folded and inserted through a small aperture, which would not allow a rigid tank of similar capacity to be fitted. Because flexible tanks can be made in shapes to suit most of the space available, a greater fuel capacity is made available to a particular aircraft when flexible tanks are used. Some aircraft fuel systems are designed to include rigid, flexible and external fuel tanks so that the greatest possible fuel load is carried. The compartment for a flexible fuel tank is made as smooth as possible on the inside and projecting joints are covered to prevent chafing the tank material. Before a tank can be fitted, the compartment must be properly cleaned out and all swarf and loose items removed. After a flexible tank has been inserted into the tank compartment, the tank is carefully unfolded and the various external fittings are aligned. Usually the walls of the more flexible tanks are attached to the compartment walls by a type of press-stud fitting. When filled with fuel, the tank expands to contact the walls of the tank compartment so that the weight of the fuel is carried by the aircraft structure and not by the tank. Because the load is not carried by the tank, flexing of the aircraft structure does not impose harmful loads upon the tank material. Flexible fuel tanks are resilient, like an inner tube and because they are resilient, the tanks can withstand a considerable amount of distortion or shock loading. If a flexible tank is not completely full it is unlikely to burst on a crash impact. 10.2.1 Tank Coverings 10.2.1.1 PROTECTIVE COVERING

A protective covering may be fixed to the outside of a flexible fuel tank. The covering is not special to type and similar covering materials are used to protect different types of tank. The protective covering usually consists of several layers of fabric, or fabric and rubber, which are cemented to the material of the tank with adhesives. When a tank is fitted with a protective cover it, in general, becomes stiff enough to support its own weight and retain its shape. However, when the various metal fittings are added, the tank will sag and it needs support when fitted. Some tanks, which do not have protective covers, are reinforced by nylon fabric or net. This type of reinforcement does not stiffen the tank, which remains very flexible and limp. This type of tank cannot support its own weight and is the type which is sometimes called a „bag tank‟.

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10.2.2 Self-Sealing Coverings

These coverings have been developed to reduce the magnitude of a fuel leak if, for any reason, the fuel tank is pierced or ruptured. The self-sealing covering is usually made from layers of cellular rubber with an overall protective cover of glass fabric or nylon fabric on the outside. This type of rubber is a material that is immediately affected by contact with fuel. If a tank leaks, the cellular rubber swells on contact with the fuel and forces its way into the puncture to block the hole and reduce or stop the leak. Unfortunately, minor leaks may remain undiscovered for some time until the self-sealing cover begins to swell and bulge on the outside. 10.2.3 Attachments and Fittings

To complete a flexible fuel tank, provision must be made for attaching fuel system components and for joining each tank into the fuel system. The fuel tank is constructed with moulded connectors and apertures of an appropriate size and position but because of the flexible nature of the material, each aperture needs to be reinforced before a system component can be fitted. Each aperture is strengthened and stiffened by fitting a metal attachment ring. The attachment rings are sometimes called „stud rings‟ or „bolt rings‟.

Attachment Rings and Moulded Connections Figure 3

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10.3 INTEGRAL FUEL TANKS Primary wing structure is used for aircraft integral tanks. They are normally located between the front and rear wing spars and between the upper and lower wing skin. Solid „tank end‟ ribs close the ends of each tank, while all the other ribs act as fuel baffles to minimise fuel slosh. Often a centre tank traverses the fuselage between the two inner wing root ribs. All fuel tanks are fuel tight. Close metal-to-metal fit of all parts forms the basic seal, with sealing compounds and sealing fasteners on all joints to complete the fluid tight seal. The centre tank will have a secondary external barrier coating to prevent fuel vapour entering the pressurised section of the fuselage. Some of the wing ribs contain a series of free-swinging, fuel-actuated baffle check valves, to prevent fuel flow away from the electric boost pumps. Access panels, usually on the underside of the wing, provide access to each tank. The outer portion of the wing provides fuel overflow by means of a surge tank, which also affords venting into the system. The fuel tanks hold all the necessary equipment for refuelling/ de-fuelling and engine fuel feed. Equipment used for fuel quantity indicating is also contained within the fuel tank structure.

Integral Fuel Tanks Figure 4 10.3.1 Tank Numbering

The majority of aircraft carry fuel only in wing and centre tanks. However, a few „extended range‟ aircraft will have an additional integral tank in the vertical stabiliser. Aircraft manufacturers number fuel tanks, in which case the philosophy will be from left to right, nose to tail. Generally, fuel tanks are numbered as shown in Fig 4. B1 Mod 11.10

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Tank Access Panels Figure 5

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Baffle Check Valves Figure 6

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Before assembly, all the structural parts that become integral fuel tanks are cleaned to a particular specification; the clean parts are immediately coated with a special sealant and assembled wet. It is important that the joints are finished (rivets closed or bolts tightened) before the sealant sets. This first coating of sealant is called the „interfay‟ and it should bond with all parts of the joint. After the joint is tightened it is necessary to remove the surplus sealant that has been squeezed out as the joint closed. After cleaning the work, a neat coating of sealant is applied at the edges of the joint; this coating is called the fillet (see the figure) and it should be strong enough to cope with any flexing between the parts. A final brush-on coat of sealant is applied to overlap the joint and fillet. Interfay, fillet and the brush-on coat are part of the standard treatment for sealing integral fuel tank structures and all use a similar sealant. As an aid to quick production, the joint can be covered by a barrier coating of a quicker drying substance. The barrier-coating material is not the same as the sealant used for jointing and it will not prevent or cure leaks. The barrier-coat becomes tack-free in a relatively short time and it is applied over partially cured sealants to reduce the possibility of contamination from swarf, when work must continue in the area of an uncured joint. To extend the leak-free life of the integral fuel tank, take great care when handling or working on the skin area which covers the integral fuel tank.

Integral Tank Sealing Figure 7

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Generally speaking, large commercial aircraft have three tanks in each wing, inner fuel tank, outer fuel tank and a surge tank. On some aircraft the fuel tanks are referred to as fuel cells. A centre tank is sometimes available as a standard option. Each fuel tank has additional space for 2% expansion of the fuel without spillage into the surge tank. Removable access panels are provided in the lower wing surface. The centre tank, if fitted, is accessible through manholes in the rear spar. 10.3.2 Water Draining

Water drain valves are provided at low points of each tank. All valves may be opened with standard tools and the outer seal of the valve is replaceable without emptying the tanks.

Water Drain Valve Figure 8

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Fuel Tank Drain Points Figure 9 10.3.3 Water Scavenge System

A typical system has a water scavenge system fitted in the optional centre tank. Two jet pumps using tappings on the tank pumps for motive power, collect water from low points and discharge it towards the fuel pump inlet. Removable access panels are provided in the lower wing surface. The optional centre tank is accessible through two manholes in the rear spar.

Water Scavenge System Figure 10 Page 10-10



10.4

SUPPLY SYSTEMS

10.5 ENGINE FUEL FEED 10.5.1 Design Requirements of an Aircraft Fuel Feed System

On an aircraft, a fuel system should be designed to comply with many requirements as laid down in Joint Airworthiness Requirements. An example of these requirements is as follows: 1. Each fuel system should be constructed and arranged to ensure a flow of fuel at a rate and pressure to ensure proper functioning of the engine for each likely operating condition. 2. The fuel system must allow the supply of fuel to each engine through a system independent of the system supplying fuel to any other engine. 3. The system design should be such that it is not possible for any pump to draw fuel from two or more tank simultaneously unless means are provided to prevent the introduction of air into the system. 4. If fuel can be pumped from one tank to another in flight, the fuel tank vents and transfer system must be designed so that no structural failure can occur because of over-filling. 5. Integral tanks must have facilities for interior inspection and repair. 6. Fuel tanks must be designed, located and installed so that no fuel is released in or near the engines in sufficient quantities to start a fire in otherwise survivable crash conditions. 7. Pressure cross-feed lines passing through crew, passenger or cargo compartments shall either be enclosed in a fuel and vapour proof enclosure, ventilated and drained to the outside, OR consist of a pipe without fittings and routed or protected against accidental damage. 8. The system shall incorporate means to prevent the collection of water and dirt or the deposition of ice or other substances from satisfactory functioning of the system. 9. Lines, which can be isolated from the system by means of valves or fuel cocks, shall incorporate provision for the relief of excess pressure due to expansion of the fuel. 10. Each fuel tank filler connection must be marked with type of fuel and be provided with a bonding point and drain discharging excess fuel. 11. There must be a fuel strainer at each fuel tank outlet or for the booster pump(s). 12. Each fuel line must be designed, installed and supported to prevent excessive vibration and allow a reasonable degree of deformation and stretching without leakage.

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10.5.2 Engine Fuel Feed (Multi Tank and Booster Pumps)

Multi tank fuel systems can use a low-pressure fuel booster pump in each tank as shown. Location of Pump Canister Assemblies Figure 11 The pumps are located in collector tanks which are equipped with check valves which provide a one way fuel flow. 10.5.3 Engine Fuel Feed (Collector Tanks)

Rather than use booster pumps in each tank, some aircraft fuel systems use groups of tanks that feed collector tanks as shown in the diagram.

Engine Fuel Feed Collector Tanks Figure 12 Page 10-12



10.5.4 Engine Fuel Feed (Fuel Cells)

Another multi tank system is the use of fire cells. In normal conditions, each engine is supplied from one pump in the optional centre tank or both pumps in the tank of its own wing. Any one pump can supply the maximum demand of one engine. A cross-feed pipe, controlled by a double motor actuated spherical plug valve, allows both engines to be fed from one side or all the fuel to be used by one engine. The valve is mounted on the rear spar in the centre section. Two plug-in a.c. driven booster pumps supplied from different busbars are fitted in each tank. Each pump has a suction inlet. On each side, the two pumps in the wing tank and one pump in the centre tank (when fitted) deliver fuel via a built in non-return valve into a single pipe. The pumps in the wing tanks are fitted with pressure relief sequence valves that ensure that when all pumps are running, the centre tank pumps will deliver fuel preferentially. No sequence valves are provided on a two tank version aircraft. In each wing tank the pumps are located in a collector box. The box is fed by gravity through flap non-return valves. This ensures that the system can continue to supply fuel under negative „g‟ or transient manoeuvres. A bypass is provided at the pumps to permit gravity feed. Air release valves are fitted to the feed lines. The supply of fuel to each engine can be shut off by an engine LP valve mounted on the front spar. This is a spherical plug valve driven by a double motor actuator. To provide the maximum integrity, the two actuators are supplied from different busbars and the cables are routed separately. Controls and indications for pumps and crossfeed valves of the feed system are located on the overhead panel. In normal operation, all wing pumps will remain on throughout the flight. If a centre tank is fitted, switching of pumps is automatic. If there are no malfunctions, no action is required during flight. The engine LP valves are controlled by operating the engine fire handles.

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Engine Fuel Feed Figure 13 10.6

FUEL FEED COMPONENTS

10.6.1 Fuel Pumps (Booster Pumps)

Pumps employed in aircraft fuel systems differ in size, shape, output, etc. However regardless of type and any special features they may have, they all operate on the same principle and consist of very similar components. Each tank is normally provided with two fuel pumps. They are all identical and interchangeable. These pumps are installed in the canister assemblies to enable replacement without de-fuelling the tank. The fuel pumps are centrifugal pumps driven by 115 volts, three phase motors. The output of each pump is about 250-300 litres per minute. Maximum fuel pressure at zero flow is about 38 p.s.i. Each pump includes a non-return and a by-pass valve. The by-pass valve is to reduce the pressure drop allowing an engine to be operated on suction feed up to about 6000 feet.

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They are protected by a thermal fuse, which is activated at approximately 175 degrees centigrade.

Fig 14 Some pumps have special features that are dictated by the aircraft role and any design requirements namely: a. Pressure relief valve. b. Non-return valve. c. AC DC motor. d. Thermal trip devices. e. Cannister shut off valve to facilitate pump replacement with fuel in the tanks.

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Cannister Assembly Figure 15 10.6.2 Jet Pumps

These are another method of transferring fuel around an aircraft fuel system. They use fuel bled from the booster pump which is continually fed through the central nozzle into a venturi. The depression created in the venturi draws fuel from the surrounding tank, in through the filter then up through the venturi tube and either into the next fuel tank or straight to the collector box.

JET PUMP FIGURE 16

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Figure 17

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10.6.3 Sequence Valves

Sequence valves are fitted to give an automatic transfer from one tank to another, the following example is for an aircraft with pumps in the centre tank, inner tank and outer tank. The valve limits the fuel pressure of the outer tank pumps from 38 psi to 17.5 psi. This is to give priority to the inner tank fuel pumps for structural reasons. When the inner tanks are empty, the engines will be automatically supplied from the outer tanks So the outer fuel pumps run continuously.

Sequence Valves Figure 18 10.6.4 Transfer Valves

The example at figure 13 shows the fuel tank split into two cells at rib 15. To enable transfer to take place, two transfer valves are fitted in this instance at rib 15. Operation of these valves is actuated by a signal from low level sensors shown just inboard of rib 2.

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10.6.5 L.P. Valve

LP Valve

Figure 19 The L.P. shut off valve enables isolation of the fuel system in the event of fire and engine maintenance, i.e. engine removal. Located at the top of the pylon on the outside of the front wing spar it will be controlled normally be operation of the fire handles and activated by either a pair of electric motors or mechanically as shown above. 10.6.6 Cross Feed Valve

Cross-feed Valve

Figure 20 B1 Mod 11.10

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The cross feed valve enables fuel to be fed to any engine from any tank. Normally of a spherical type construction with two 28 VDC electric motors mounted on a differential gearbox. One motor only will drive the valve at any time, the other motor is a back up. The cross feed valve would normally be fitted on the rear spar as shown in the figure. 10.7 APU FUEL FEED The feed to the APU is taken from the left engine feed but may be taken from the right engine feed when the cross feed valve is open. The tank booster pumps can supply fuel to the APU at the required pressure. For starting the APU without electrical power available for the tank pumps, a separate pump is provided that can be operated from the aircraft batteries and is mounted in the feed line on the rear spar of the centre section. The supply of fuel to the APU can be shut off by a valve mounted on the rear spar of the centre section. It is a spherical plug valve driven by a double motor actuator. To provide the maximum integrity, the two actuators are supplied from different busbars and the cables are located in separated routes. The feed pipe emerges from the top of the tank and passes through the pressurised fuselage in a drained and vented shroud that extends to the APU fire wall.

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APU Fuel feed Figure 21

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10.8 DUMPING, VENTING AND DRAINING 10.9 DUMPING (JETTISON) Fuel jettison systems are fitted to a number of large commercial aircraft to allow the jettisoning of fuel in an emergency thus reducing weight so as to prevent structural damage when landing. Fuel jettison systems are often fitted after the installation of a centre tank, because of the extra fuel weight. The system illustrated is from a wide-bodied twin fitted with multi tanks and booster pumps. The jettison pipe is branched off the feed pipe between the inner tank fuel pump and the inner tank shut off valve. A check valve is installed to separate the outer tanks during jettisoning. The function of this check valve is to prevent the dumping of the outer tanks fuel. The jettison pipe runs inside the wing tanks through the ribs into the outer tanks, where the jettison valves are installed. These valves are fitted to the bottom of the tank.

Jettison System Figure 22 Because of electrical emergency situations, the valve will be driven by two 28 VDC electric motors. The motors are mounted from the outside and are attached to the bottom of the tank through a gearbox and in many instances are interchangeable with the cross feed valves.

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Shut Off Valve Figure 23 The outlet of the jettison pipe is normally at the end of the flap track fairing and fitted with an anti corona device to avoid vaporisation of the fuel. A normal transfer rate will be in the region of 30-350 litres per minute.

Fuel Jettison Figure 24

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The jettison operation is controlled from a jettison panel located either on a flight engineers station or from an overhead panel on a two crew configuration. Normally the panel is protected by a quick release cover. In the following example, two switches are provided to operate the jettison valve. i.

A primary switch for motor number one.

ii.

A guarded secondary switch for motor number two.

Fuel Jettison Control Figure 25 The position of the right and left-hand jettison valve is monitored by two magnetic indicators, showing green cross-line when the valve is closed and in-line when the valve is open. As is common with this type of indicator, it will show amber cross-line to indicate transit or malfunction.

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10.10

THE VENT SUB-SYSTEM

10.10.1 General

An air vent is fitted to the top of each tank to allow free flow of air in and out of the tank as the fuel level rises and falls. This is known as inwards and outwards venting and is required to prevent over pressurisation of the tanks as the fuel level rises and depressurisation as it falls. 10.10.2 Venting Due to Heat

Another important aspect of an aircraft vent sub-system is that it must be able to cope automatically with any expansion and/or contraction of the fuel. As the fuel expands, due to heat, the vent must allow air and sometimes fuel, to escape to atmosphere via vent pipes. Conversely the sub-system must allow air into the tanks during contraction of the fuel when the outside air temperature (OAT) is decreasing. 10.10.3 Unpressurised System Venting

This is a very simple method of venting tanks which requires only that fuel tank vent orifices be connected to a vent pipe gallery, which leads to atmosphere directly. Venting of this type is found mainly in small aircraft; some helicopters and aircraft with low flight ceilings. The disadvantages of “open orifice” or “open vented” tanks are that they are subject to fuel venting during manoeuvres, they limit the maximum ceiling of the aircraft due to the fact that fuel boils at the low ambient atmospheric pressure found at altitude; danger of cavitation in fuel supply lines if fuel should boil; increased rate of evaporation (REID VAPOUR PRESSURE) leading to a greater fire risk. REID VAPOUR PRESSURE (RVP) – the rate at which fuel gives off vapour. Obviously there are many inherent problems with the open vented system. It is for many reasons that most aircraft fuel systems are pressurised. 10.10.4 Pressurised Fuel Tanks

On most large aircraft, the fuel tanks are vented through a pipe connected to the surge vent tank. The vent pipes are sized to prevent tank overpressure in the event of a refuel cut off failure. In the example shown, the centre tank vent pipe is connected to the left-hand surge vent tank. The inner and outer tank pipes are connected to the relevant side surge vent tank. The centre tank vent pipe ends inside the surge vent tank at the top. The inner and outer vent pipes end about 3 centimetres above the bottom of the surge vent tank. These ends are arranged so that any fuel overflowing into the surge vent tank is drawn back into the wing tanks by suction, as long as one or more fuel pumps are running. On some aircraft fuel pumps are fitted to pump the fuel back to the tanks from the surge tank and will be activated by a float switch. Each vent tank is vented to atmosphere via the NACA valve. This valve ensures tank pressurisation during flight and allows the fuel to flow out in the event of a high level cut-off failure during refuelling.

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On some aircraft a frangeable disc is fitted in the surge tank to prevent structural damage caused by over pressure. A flame arrester is also fitted in the NACA intake in case of ground fires.

Venting System Figure 26

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Figure 27

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10.10.5 Float Valves

Each wing tank is provided with an additional vent opening. This opening is connected to the corresponding venting line and controlled by a vent float valve situated at the highest point of the tank.

Vent Float Valve Figure 28

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10.10.6 Vent Pipe Drains

At the lowest points of each vent pipe, a self-draining non-return valve is connected. The type shown is made of synthetic rubber.

Vent Drain Valve Figure 29

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The use of the centre wing box as a tank on some aircraft has made it necessary to protect this area against leaking fuel. A vapour seal is installed around the forward and lower part of the tank. The space between the tank sink and the vapour seal is ventilated with air coming from the air conditioning system. The air is directed to the outside through several small outlets. If the tank has a fuel leak, the vent air line will collect this fuel and drain it through these outlets.

External Ventilation of Centre Tank Figure 30

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10.11

CROSS-FEED AND TRANSFER

Cross-feed Valves permit the transfer of fuel from any tank to any engine, whereas Transfer Valves enable fuel to be transferred from tank to tank. 10.11.1 Auto – transfer

When an aircraft has a wing with lateral dihedral the fuel pumps will normally be inboard and the fuel flow towards the wing root. When the wing contains more than one tank, the outboard tank will automatically transfer into the inboard tank and so be the first to empty. Since the inboard tanks will be feeding the engines, a transfer valve between the inboard and outboard tanks will be opened automatically, whenever a high level float switch in the inboard tank detects it being not full.

Cross Feed Control Panel Figure 31

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Figure 32 10.11.2 Manual – transfer

No in-flight transfer of fuel between left and right mainplanes is possible for reasons of trim. However fuel can be fed from any tank to any engine by means of boost pump selection and the opening of a crossfeed valve from the flight deck.

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10.12

INDICATIONS AND WARNINGS

Provision is made to display fuel tank quantity, boost pump low pressure, crossfeed valve and fuel/fire shut off valve position, on the flight deck overhead panel. Though the layout will vary from aircraft type to type, generally it will be similar to the examples shown below.

Overhead panel - Push switch type Figure 33

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Overhead panel - Toggle switch type Figure 34

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Additionally, aural and visual warnings on the glare shield will result if the fuel system develops a fault.

Glare Shield and Fire Warning Panel Figure35

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ECAM System Display Figure 36

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10.13 FUEL LEVEL SENSING A modern aircraft will use thermistors to send signals through amplifiers to actuate warnings, sequencing, etc. Older aircraft may use float switches as shown in the following diagram.

Low Level Sensing Figure 37 Float operated switches are of a magnetic type, similar to the one shown above and are designed to isolate the electrical mechanism from the fuel tank for safety reasons. Upward movement of the float brings the armature closer to the magnet and at a predetermined fuel level, it has sufficient influence to attract the magnet, which results in operation of the micro switch. As the fuel level and the float fall, the attraction of the armature is eventually overcome by the combined forces of the counterweight and the micro switch spring and the counterweight falls, changing the micro switch circuit. Whether they are float switches or thermistors, their functions are as follows. 1 2 3 4 5

High level sensing. Overflow sensing. Low level sensing. Under full level sensing. Level sensing for calibration (Fuel Trim only).

10.13.1 High Level Sensing

High level sensing is installed to prevent an overfilling of the fuel tanks. When the fuel washes around the respective sensor, the: Page 10-36



 associate refuel/defuel valve closes.  blue FULL light on the fuelling panel comes on. The high level signal from the inner and outer tanks could be used for computation purposes in the fuel quantity computer, when refuelling in AUTO MODE. 10.13.2 Overflow Sensing

If during refuelling the high level shut off system fails, fuel enters the adjacent vent tank and washes around the overflow sensor. This is indicated by the amber FULL light on the refuel panel. 10.13.3 Low Level Sensing

Low level sensing is divided into:  outer tank low level and  inner/centre tank low level sensing. If the outer tank LO LVL sensor is exposed to air, the associated amber LO LVL light comes on. The inner/centre tank low level sensing have only in the AUTO MODE a function (ref. fuel pump control). 10.13.4 Calibration Sensing (Fuel Trim only)

Calibration sensors are installed in centre tanks, inner tanks and trim tank. They give a signal at a predetermined filling level in the trim tank for accuracy test of the fuel quantity indication during refuelling. For the trim tank the calibration sensor switching level is corrected by the stabiliser position. 10.13.5 Under Full Level Sensing

When the fuel quantity drops in either outer tank below a certain level, the maximum flight speed (VMO) becomes reduced in order to protect the wing structure. The sensor signals are sent to the ADC (Air Data Computer). 10.14 FUEL QUANTITY SYSTEM MEASUREMENT AND INDICATION The system has the following tasks: 1 Measuring of the fuel quantity in the tanks. 2 Indicating of the fuel quantity on:  The fuel quantity indicator.  The pre-selector.  The ECAM system fuel page.  ECAM/EFIS. 3 Controlling of automatic refuelling. 4 Fuel quantity messaging to the flight management computer. The system comprises: 1 2 3 4 5 6 7

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fuel quantity computer. capacitance probes. capacitance index compensator. cadensicon sensor. attitude sensor. THS position detector. associated indicator in the flight compartment.

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10.15 PRINCIPLE OF CAPACITANCE GAUGING A capacitor is an electrical device which stores electrical charge. The amount of charge it can hold depends upon three physical properties of the capacitor itself, namely: a. The surface area of the plates. b. The size of the gap between the plates. c. The insulating material (dielectric) between the plates. In a fuel tank “capacitor stack” two of the above are fixed, ie. the area of the plates and the gap between them. The only variable is the dielectric which, in a fuel tank, is either fuel or air or both. The amount of charge held in the capacitor, when the tank is full, will be of a preset value. As the fuel level falls, the dielectric will gradually change to air and the amount of charge stored will reduce. This change in capacitance is sensed by a signal conditioner and the change in fuel level is thus sensed. 10.16 FUEL QUANTITY INDICATING SYSTEM Each tank has installed a group of probes arranged so that a minimum of one probe is immersed at all times, the number of probes will vary from aircraft to aircraft. The following example is from a wide-bodied twin fitted with a fuel trim system. The number of probes is:  6 in each outer tank.  6 in each inner tank.  4 in the centre tank.

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The probes of each group are wired in parallel and connected to a summing adapter, located on the wing rear spar. The probe level signals are sent to the fuel quantity computer.

Wing Capacitance Probe Installation Figure 38 10.16.1 Capacitance Index Compensator

One compensator is installed in each tank to the lowest located capacitance probe. Separated wiring for these units is routed to the fuel quantity computer.

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The purpose of the index compensator is to sense the different types of fuels, additives, etc. and make correction signals for accurate fuel readings.

Capacitance Index Compensator - Installation Figure 39

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Probe Installation – Trim Tank Figure 40

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10.16.2 Measurement

The signals from the capacitance probes in each tank are sent via adapters to the fuel quantity computer. The computer calculates the fuel quantity. To increase the measuring accuracy, further signals enter the computation:  

capacitance index compensator, balances different fuel types. Condensicon sensor: Senses while refuelling the: Density Dielectric constant of running fuel.  Attitude Sensor: Senses on ground and in flight the attitude of the aircraft to the: Roll axis (longitudinal) Pitch axis (lateral) The attitude signal computation depends on the AIR/GRND signal (wing bending direction).  THS Position Detector Senses the THS position steady and gives its signal to the fuel quantity computer for correction of trim tank fuel measurement. Fuel Quantity Indicator The fuel quantity of the tanks is normally displayed in 10 kg steps. Power supply and the indication signals are delivered by the fuel quantity computer. To avoid transmission errors, the indicator sends feedback signals to the computer. The indicator is also used for test purposes. In the test mode, the indicator displays different number codes. The examples shown are from an aircraft with a two-man crew. The refuelling system will be looked at later. The aircraft is a twin with a centre tank, an inner and an outer tank. Note: The LO LVL lights in the indicator receive their signals from the outer tank LO LVL sensing circuit. 10.17

REFUELLING AND DE-FUELLING

10.18 REFUELLING As you will be aware, as any liquid flows through a pipeline, it will produce Static Electricity. If this static electricity were allowed to discharge in the presence of aviation fuel vapour, an explosion would result, with possible catastrophic results. To therefore minimise the explosion risks, the following guidelines must be followed. Safety Precautions:  Use correct grade of fuel (Av-gas, Av-tur, Av-tag).  No smoking within 15m.

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          

No metal studded or tipped footwear. Correct bonding of Aircraft and Bowser. Correct positioning of Bowser. No vehicles or Ground Equipment under the aircraft. Maintenance activity kept to a minimum. No replenishment of LOX. No transmitting of Radar Aircraft & Bowser not to be left unattended. Check and remedy fuel spillage or leakage. Appropriate Fire Appliance readily available. The electrical state of the Aircraft must not change while connected to the Bowser.

Refuelling a small aircraft is no more complex than filling the family car. One limitation is that on some aircraft it is not possible to fly the aircraft with all the seats occupied with full baggage allowance, when the tanks are full. This means that if the aircraft is to be flown fully loaded, it may be necessary to re-fuel to less than full, to keep the aircraft within its weight limits. As the aircraft become more complex, the refuelling exercise has to be carried out with more care. If the aircraft is small but has say, two tanks in each wing, and the fuel load is to be three quarters full; then it may be the rule for that aircraft that the inner tanks have to be filled to the top first and the remainder put into the outer tanks. This puts less bending load on to the wing spars. When we get to larger aircraft, there are several further problems to consider. Not only must the aircraft be filled laterally in the correct order but, if the aircraft has the fin, tailplane and rear fuselage tanks mentioned earlier, it must be refuelled in the correct order longitudinally as well to ensure the aircraft stability is maintained.

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Modern large aircraft utilise pressure refuelling, which has replaced open line refuelling on most aircraft with high fuel capacities. The time taken to fill a Boeing 747 through a normal hose and nozzle system would take hours. With pressure refuelling, a large diameter hose is rigidly connected to a coupling in the aircraft and fuel under pressure of about 40 psi is pumped into the aircraft tanks. To assist this operation, most aircraft can have the total fuel load pre-set at the point of connection so that the aircraft stops the refuelling at the correct time. The illustrations show the location and layout of a typical, Boeing 777, refuelling panel.

Boeing Refuelling Panel Figure 41

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10.18.1 Pressure Refuel – Functional Description

Fuel flows from the refuel adapters into the refuel/jettison manifold. When the refuel valves open, fuel flows from the manifold into the fuel tanks. A flow tube at the end of each refuel valve decreases the exit force of the fuel. The flow tube also puts the fuel in different parts of the tank.

Refuel System Layout Figure 42

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As each tank reaches full, the high level sensor signals the refuel valve to close to stop fuel flow. When all refuel flow ceases, fuel that is left in the refuel/jettison manifold goes through the manifold drain valves and into the main tanks. The manifold has two vacuum relief valves. These valves permit air into the manifold when the fuel leaves via the manifold drain valves. Refuel Manifold Drain Valve Fig 43

Refuel Manifold Drain Valve Figure 43

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If a refuel system failure prevents the refuel valves from closing, fuel goes into the surge tanks. If the fuel gets to the level of the surge tank float switches, the switch closes, and all refuel valves are closed.

Surge Tank Float Switch Fig 44

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10.19

DEFUELLING

Defuelling a pressure type fuel system is almost the reverse of the refuelling procedure. A de-fuel bowser would be connected to the single fuel point coupling, and using a combination of both the bowser‟s suction pump and the aircraft‟s own fuel supply booster pumps, selected tanks can have their contents returned to the bowser.

Defuel System Layout Fig 45

10.20 LONGITUDINAL BALANCE FUEL SYSTEMS The weight of the fuel is a large percentage of an aircraft‟s total weight, and the balance of the aircraft in flight changes as the fuel is used. These conditions add to the complexity of the design of an aircraft fuel system. In small aircraft the fuel tank or tanks are located near the centre of gravity so the balance changes very little as the fuel is used. In large aircraft, fuel tanks are installed in every available location and fuel valves allow the flight engineer to keep the aircraft balanced by scheduling the use of the fuel from the various tanks. High performance military jets and more modern civil aircraft will use a fully automatic fuel scheduling system to reduce the workload on the flight crew.

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10.21

SUPERSONIC FLIGHT FUEL TRANSFER

Longitudinal Fuel Transfer Figure 46

In supersonic flight the aerodynamic centre of pressure moves aft, thus changing the longitudinal stability. This was compensated in aircraft like Concorde, by moving the centre of gravity, shifting fuel as necessary between the fuel tanks in the rear fuselage and the wings as shown in the previous diagram.

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MODULE 11.11 HYDRAULIC POWER

engineering

CONTENTS 11 HYDRAULIC POWER INTRODUCTION ...................................... 11-3 11.1

11.2 11.3 11.4 11.5 11.6 11.7 11.8 11.9 11.10 11.11 11.12 11.13 11.14 11.15 11.16

11.17

11.18 11.19 11.20 11.21

11.22 11.23

COMPARISON W ITH OTHER POWER TRANSFER SYSTEMS ............... 11-3 11.1.1 Mechanical Systems ..................................................... 11-3 11.1.2 Electrical Systems ......................................................... 11-4 11.1.3 Pneumatic Systems ...................................................... 11-4 BASIC HYDRAULIC PRINCIPLES ...................................................... 11-5 COMPRESSIBILITY ......................................................................... 11-5 PASCALS LAW OF FLUID COMPRESSIBILITY .................................... 11-5 FORCE DUE TO FLUID PRESSURE .................................................. 11-6 DIFFERENTIAL AREA ..................................................................... 11-7 HYDRAULIC FLUIDS ....................................................................... 11-8 EFFICIENCY .................................................................................. 11-8 PROPERTIES OF AN IDEAL HYDRAULIC FLUID .................................. 11-8 TYPES OF HYDRAULIC FLUID ......................................................... 11-8 SEALS ......................................................................................... 11-9 11.11.1 Types of Seals .............................................................. 11-10 HYDRAULIC POWER SYSTEMS ............................................... 11-13 SIMPLE HYDRAULIC SYSTEM ......................................................... 11-15 11.13.1 Operation (Fig. 15) ........................................................ 11-16 SYSTEM COMPONENTS ................................................................. 11-20 INTRODUCTION ............................................................................. 11-20 RESERVOIR‟S ............................................................................... 11-20 11.16.1 Vented Reservoir .......................................................... 11-21 Pressurised Reservoir................................................................. 11-22 11.16.3 Remote Servicing Point ................................................. 11-25 11.16.4 Filters ............................................................................ 11-26 ACCUMULATORS ........................................................................... 11-28 11.17.1 Purpose ........................................................................ 11-28 11.17.2 Construction .................................................................. 11-28 11.17.3 Charging Operation ....................................................... 11-29 11.17.4 Bladder & Diaphragm Type Accumulators ..................... 11-29 PRESSURE GENERATION (HYDRAULIC PUMPS) ............................... 11-32 HAND PUMPS ............................................................................... 11-33 SUCTION BOOST PUMPS............................................................... 11-34 POWERED PUMPS ........................................................................ 11-35 11.21.1 Constant Volume Fixed Displacement Pumps............... 11-35 11.21.2 Piston Pumps ................................................................ 11-37 11.21.3 Unloading (cut-out) Valve .............................................. 11-41 11.21.4 Constant Pressure/Variable Displacement Pump .......... 11-42 11.21.5 Stratopower Pumps....................................................... 11-44 11.21.6 Operation ...................................................................... 11-45 EMERGENCY PRESSURE GENERATION ........................................... 11-47 HAND PUMPS ............................................................................... 11-47

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engineering 11.24 11.25 11.26 11.27 11.28 11.29 11.30

11.31

11.32 11.33 11.34

11.35 11.36 11.37 11.38

DUPLICATION OF SUPPLY .............................................................. 11-47 ELECTRIC MOTOR DRIVEN PUMPS (EMDP‟S) 115V AC .................. 11-51 AIR TURBINE MOTOR DRIVEN PUMPS (ATM‟S OR ATDP‟S) ............. 11-52 POWER TRANSFER UNITS (PTU‟S) ................................................ 11-53 HYDRAULIC RAM AIR TURBINES (HYRAT‟S) ................................... 11-54 HYDRAULIC VALVES ................................................................ 11-55 PRESSURE CONTROL VALVES........................................................ 11-55 11.30.1 Pressure Relief Valve .................................................... 11-55 11.30.2 Pressure Regulators ..................................................... 11-56 11.30.3 Thermal Relief Valve ..................................................... 11-57 11.30.4 Pressure Reducing Valve .............................................. 11-58 FLOW CONTROL VALVES ............................................................... 11-59 11.31.1 Non-Return (Check) Valve ............................................ 11-59 11.31.2 Selector Valves ............................................................. 11-61 11.31.3 Priority Valves ............................................................... 11-65 11.31.4 Sequence Valves .......................................................... 11-66 11.31.5 Hydraulic Fuses ............................................................ 11-68 POWER DISTRIBUTION .................................................................. 11-70 POWER CIRCUITS ......................................................................... 11-71 COMPONENT CIRCUITS ................................................................. 11-73 11.34.1 Flaps ............................................................................. 11-73 11.34.2 Landing Gear ................................................................ 11-76 HYDRAULIC POWER - INDICATION AND W ARNING SYSTEMS ............. 11-76 HYDRAULIC PRESSURE ................................................................. 11-77 HYDRAULIC QUANTITY .................................................................. 11-80 INTERFACES WITH OTHER SYSTEMS ............................................... 11-83

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11 HYDRAULIC POWER INTRODUCTION This section explains the basic principles, advantages, operation and layouts of aircraft hydraulic power systems. It also describes the various materials used and the function of the associated components that make up, operate and control different types of hydraulic systems and Fluid power: The the interface of hydraulic power with other systems. transmission of force by Fluid power systems are mechanical systems in which a moving fluid performs work. This fluid may either be a compressible gas or an incompressible liquid. Systems that use compressible fluids (gasses) are called pneumatic systems, and those that use incompressible fluids are called hydraulic systems. Hydraulic power is often used to operate aircraft landing gear, flight controls, flaps and slats, air brakes, wheel brakes, nose-wheel steering, freight doors etc. in conjunction with other systems. This method of operation is termed; Hydraulic Actuation. 11.1 COMPARISON WITH OTHER POWER TRANSFER SYSTEMS

the movement of a fluid. i.e. Hydraulic and pneumatic systems.

Fluid: A substance, either a gas or a liquid, which flows and conforms to the shape of its container.

Hydraulic actuation has the following advantages over mechanical, electrical and pneumatic forms of remote control: 11.1.1 MECHANICAL SYSTEMS

a Hydraulics provides smoother and steadier movement.

Hydraulics: A fluid power system, which transmits force through an incompressible fluid.

b Hydraulic power is confined to pipelines and components, which avoids the extra strengthening of airframe structure required for mechanical operations. c Hydraulics systems have a higher Power/weight ratio than mechanical systems, particularly on large transport aircraft. d Installation of hydraulic equipment is simpler. Pipelines between components for example, can be routed around obstructions and structure, whereas to solve this problem mechanically requires the use of levers, guides, bell-cranks and pulleys to change direction of mechanical pushrods and cables. e Variation in speed of operation can be achieved without the use of complex gearing. f

Finally, hydraulic actuation normally obtains its power from the aircraft engines, which relieves the pilot of unnecessary fatigue when operating a service.

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Pneumatics: A fluid power system, which transmits force through a compressible fluid.

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11.1.2 ELECTRICAL SYSTEMS

The obvious advantage of electrical systems is that cables can be routed around obstructions even easier than pipelines. They are also generally lighter in weight, however, the power required to actuate landing gear and flight controls of large aircraft, would require large electric motors powered by equally large (and heavy) electrical generators, requiring high current cables connecting the system components. Therefore, electrically operated systems are normally limited to light aircraft. 11.1.3 PNEUMATIC SYSTEMS

Some older type aircraft used pneumatics to operate brakes systems and emergency landing gear extension systems. Modern, large transport aircraft use high-pressure pneumatics to actuate systems in high temperature, fire hazard areas such as; jet-engine thrust reversing systems and engine starting operations, also cabin pressurisation and air-conditioning systems. However, the main disadvantages over hydraulic actuation is its compressibility when actuating highly loaded systems such as landing gears and flight control operations. Also, difficulty in detecting leaks in the system, and problems with moisture and corrosion contamination have limited the use of pneumatic power as a remote control system. Pneumatic power has some advantages such as; lightness and return lines are unnecessary.

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11.2 BASIC HYDRAULIC PRINCIPLES 11.3 COMPRESSIBILITY All liquids have a high resistance to compression. The example in figure 1 shows two cylinders of equal volume, each fitted with pistons, one containing liquid, the other air. If a force of 20,000 N (Newton‟s) is applied to the pistons, the decrease in volume of the air is large compared to that of the liquid, which is negligible.

Compressibility of Fluids Figure. 1 11.4 PASCALS LAW OF FLUID COMPRESSIBILITY Power transmission in a closed hydraulic (or pneumatic) system, is best explained by PASCAL‟S LAW, which states: “Pressure in an enclosed container is transmitted equally and undiminished to all parts of the container and acts at right angles to the enclosing walls.” See figure 2 Container (a), shows that pressure produced by a fluid in an open container is caused by the height of fluid above the point at which the pressure is measured. The higher the fluid above the gauge, the greater the pressure.

Container (a) Container (b) Pascal‟s Law Figure. 2 Issue 1 Mod 11.11

Container (b), shows that, when pressure is applies to a liquid in a closed container, the pressure rises to the same amount in all parts of the container

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11.5 FORCE DUE TO FLUID PRESSURE It has been stated that fluid pressure is transmitted equally in all directions, but in hydraulic actuation it is more important to know the total effect of the pressure upon a particular surface. In figure 3, a pressure of 10 N/mm² is applied to one side of a piston in a cylinder actuator. The piston diameter is 40mm. It‟s area is multiplied by the piston radius squared (r2) i.e. 3.142 x 20mm² = 1,256.8mm². Therefore the force (load) that the piston can push is: 10N x 1,256.8mm² = 12,568Nf

RETURN

PRESSURE = 10Nf/mm2

FORCE A = 12,568Nf

40mm

Force by an Actuator, due to Hydraulic Pressure Figure. 3. When the same value of hydraulic pressure is applied at the opposite side of the piston (figure 4), the force will be smaller. This is due the ram, reducing the effective piston area upon which the hydraulic pressure is acting. In this case the effective area will be The area of the piston, minus the area of the ram i.e. (3.142 x 20mm x20mm) - (3.142 x 5mm x 5mm) = 1256.8mm2 - 78.55mm2 = 1178.25mm2 The force is now reduced to 10Nf/mm2 x 1178.25mm2 = 11782.5Nf.‟ RETURN

PRESSURE = 10Nf/mm2 10mm

FORCE B = 11782.5Nf

40mm

Reduced Force due to smaller piston area Figure 4 Issue 1 Mod 11.11

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11.6 DIFFERENTIAL AREA Another aspect of force produced by a fluid is the effect of differential area. When the two fluid ports are connected together, as in actuator in figure 5, the pressure is the same on both sides of the piston. The piston will move to the right. This is caused by the area of the piston being reduced on one side by an amount equal to the cross sectional area of the piston rod. Since the force is 12568Nf on the larger area of the piston and 11782.5Nf on the smaller area of piston, the resultant force will be 785.5Nf and the piston will extend. = 10Nf/mm2 C = 785.5Nf

12568Nf

11782.5Nf

Resultant Force when equal pressure applied to both sides of piston Figure 5

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11.7 HYDRAULIC FLUIDS 11.8 EFFICIENCY The efficiency of a hydraulic system is governed by the resistance to motion, which is encountered by the fluid. In practice, a certain amount of force is necessary to overcome friction between pistons and cylinders, piston rods against bearings and seals, etc. Friction between the fluid and the walls of pipelines and hoses depends upon the: a

Velocity of the fluid in the pipelines.

b

Bore, length and internal finish of the pipelines.

c

Number of bends in the pipelines and the radius of the bends.

d

Viscosity of the fluid.

11.9 PROPERTIES OF AN IDEAL HYDRAULIC FLUID Fluids used in an aircraft hydraulics system must have the following properties: a

Be as incompressible as possible.

b

Have a very low viscosity rate.

c

Be free flowing over a wide temperature range.

d

Be chemically stable.

e

Not affect, or be affected by the materials in the system components.

f

Must not foam during operation when subject to sudden pressure increases or decreases.

g

Have good lubrication properties.

h

Have a high flash point.

i

must not deteriorate or form sludge.

Not all fluids have these properties, therefore, the only type of fluid allowed in a specific hydraulic system is that recommended by the manufacturer of the hydraulic components (specified in the Maintenance Manual). Technical bulletins issued by the fluid manufacturer provide information about the compatibility of the hydraulic fluids with various aircraft materials. 11.10

TYPES OF HYDRAULIC FLUID

There are three basic types of hydraulic fluids used in aircraft hydraulic systems: vegetable base, mineral base, and synthetic base. 1. Vegetable (Castor oil) base,

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DTD 900/4081 (MIL- H- 7644) - Golden yellow (or Blue) in colour, used with natural rubber seals. It is inflammable, strips paint and attacks synthetic rubber. It is toxic in a fine spray mist. These systems can be flushed with alcohol. (Only found on very old aircraft types) 2. Mineral base,

DTD 585 (MIL- H- 5606) - Red in colour, used with synthetic rubber seals. It is a kerosene-type petroleum product with good lubricating properties, but it is inflammable and attacks natural rubber. It can be flushed with naphtha, varsol, or Stoddard solvent. Neoprene seals and hoses may be used with this fluid. It‟s density and lubricating properties vary with temperature. 3. Synthetic ester base,

SKYDROL 500B - Purple in colour, used with Butyl, Ethylene Propylene, or Teflon seals. It is fire resistant, strips paint and attacks natural and synthetic rubbers. It can operate in a very wide temperature range: -20ºC ( -68ºF) to 107ºC (225ºF). Skydrol systems can be flushed with trichlorethylene. Components can be cleaned with methyl ethyl ketone (MEK), or isopropyl alcohol. Skydrol will cause irritation of the skin and burning of the eyes, therefore protective equipment and clothing should be worn when handling this fluid. CAUTION: These fluids are not compatible with each other and must never be mixed, or used to replace each other. Note: If a system has been inadvertently serviced with the wrong fluid, the complete system must be drained and flushed with an approved solvent, and all the seals in the system must be replaced. Seals can only be identified by Part number, obtained from the appropriate Illustrated Parts Catalogue. 11.11

SEALS

Seals are used throughout hydraulic and pneumatic systems to minimise internal leakage and the loss of system pressure. The two main types of seals used in aviation are: a) Gaskets. These are used where there is no relative movement between the surfaces. (Covers, inspection panels and end-plate sealing etc.) b) Packings. Used where relative movement does exist. (Piston and actuator sealing, rotating shaft sealing etc.) All rubber seals have a Shelf life starting from the Cure date (Date of manufacture) This shelf life is dependant on the type of material, it‟s use and the conditions of storage.

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Note: All rubber items should be stored in a constant, dry and relatively cool environment, away from any form of Ultra-Violet (UV) light, (Sunlight or strong artificial light) and ionised atmospheres. (Storage batteries and strong magnetic fields). Such varying conditions and harsh atmospheres can cause rapid deterioration and reduced self-life of all rubber components. Rubber seals are supplied individually in hermetically-sealed packaging, the Cure date being clearly marked on each package, together with the manufacturers part number, Batch number and Mil Spec. The seals should be stored in their original, unopened packaging until required for use. The issue of seals from the Bonded Store should be as they are received. “First in – First Out” 11.11.1

TYPES OF SEALS

There are many different types of seals available for a variety of applications. Most can be broken down into six general designs: Chevron/V-ring, U-section, Square section, O-ring, Bonded Seal, Wiper ring, Duplex,

One-way seals: Both Chevron (V-ring) and U-section seals derive their name from their shape. (See fig. 6a) These seals will prevent fluid flow in one direction only. To prevent flow in both directions, two sets of seals must be installed placed back-to-back. (See fig. 6b) Both seal types are used in very high-pressure situations, normally with two or more seals placed together as in fig. 6b.

Figure 6a Chevron/V-ring & U-section Seals

The apex or point of the seal rests in the groove of a back-up ring. A spreader ring is installed in front of the seal and compressed by an adjusting nut, expanding the seals and holding them tight against the actuator cylinder wall. U-section seals are used in the same manner but with different shaped back-up and seal retaining methods. Figure 6b Correct placement of Chevron seals on a double-acting hydraulic piston

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Double-acting (Two-way) seals: (Fig.7) are suitable for applications where a positive seal and long life are essential. The “T” section profile provides a stable base thus preventing rolling and spiral failure. The PTFE backing rings positioned either side of the seal prevent extrusion (distortion) of the seal under high pressure and piston speeds.

Figure. 7 Double Acting seal

Note: Extrusion is when the seal is forced to distort and wedge between the piston and cylinder wall due to high pressures and speeds. (See „O‟ ring illustration fig. ?) Figure 8 Duplex Seal

Duplex seals: (Fig.8) are often installed in accumulators, floating pistons and emergency air circuit components. They consist of an inner layer of soft rubber bonded to a harder outer layer, allowing it to seal against varying oil and air pressures. Square section seals: (Fig. 9) Often used on piston heads and Landing gear Oleo‟s. It can withstand high pressures and sudden, high speed piston deflections. Soft metal or Tufnol back-up rings are sometimes installed to provide additional seal compression for good sealing and prevent extrusion.

Figure 9 Square Section Seal

Wiper Ring Seal: (Fig. 10) This type does not act as a pressure seal, but as a scraper, by removing dirt, oil and water from the piston shaft, preventing damage to the pressure seal, thereby prolonging the pressure seal life.

Fig. 10 Wiper Ring Seal

Note: It is extremely important to ensure the Wiper ring is installed the correct way! otherwise it will allow FOD to pack up against the pressure seal, causing rapid seal failure and piston shaft wear. Bonded Seal: (Fig. 11) These seals are fitted to banjo unions, adaptor plugs, flush-mounted components etc. The rubber seal is hermetically bonded to the metal washer and is fitted between the two components thereby compressing the seal to the extent of the metal washer thickness when the components are tightened together. Issue 1 Mod 11.11

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O-ring Seal: This is the most commonly used double-acting (Two-way) seal used in fluid and pneumatic systems. It can be used either as a gasket or a packing seal in both static and reciprocating applications. The seal fits into a groove in one of the surfaces to be sealed, the depth of which should be 10% less than the seal diameter. (See fig.12). This provides the compression of the seal against the mating component to provide a seal under zero pressure conditions. Fig. 13 (A) shows the correct sealing condition. Fluid pressure forces the seal against the side of the grove and wedging it tightly against the piston and cylinder wall. With less than 10% “pinch”, fluid will leak past the seal under low pressure conditions. (See fig. 13 (B).

Figure. 12 The groove in which an O-ring seal fits should be wider than the O-ring, but the depth should be 10% less than the O-ring diameter.

In some high pressure applications a back-up ring is installed on the nonpressurised side of the O-ring on oneway operations, but both sides of the Oring should have back-up rings installed on two-way operations to prevent extrusion of the seal between the piston and cylinder wall. (Fig. 13 (C)). The mouth of a cylinder in which an Oring equipped piston fits must be chamfered to avoid cutting or pinching of the O-ring during installation. (Fig. 13 (D)).

Chamfer:-

Correct

Incorrect

Sealing action of an O-ring Figure. 13. Issue 1 Mod 11.11

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HYDRAULIC POWER SYSTEMS

As aircraft have become more complex, the demand for hydraulically operated equipment has increased. Retractable landing gear, wing flaps, brakes, engine cowl flaps, passenger doors and stairs, hydraulically powered flight controls, i.e. elevators, rudders, ailerons, air brakes and lift dump systems, leading edge flaps and slats. On modern aircraft, this demand has warranted the design of a complete and independent, „Hydraulic Power Supply System‟ Figure 14 shows a block diagram of a large, jet transport aircraft. To aid in understanding the development of the systems, we will start with a very basic hydraulic system and build on it as we discuss the various components.

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engineering

Figure 14

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Large Aircraft Hydraulic System

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SIMPLE HYDRAULIC SYSTEM

Aircraft hydraulic systems consist of a varying number of components, depending on the complexity of the system; i.e. fluid to transmit the force, pipelines and hoses to carry the fluid to the components, a reservoir to store the fluid, a pump to move the fluid, actuators to change the flow of fluid into mechanical work, and valves to control the flow, direction and pressure of the fluid. We will start with a simple system and add components to it, thereby developing to a more complex system resembling that which you are likely to encounter in the „Aircraft Maintenance work-place‟. Simple hydraulic system using a Reservoir, hand-pump, non-return valves, double-acting, linear actuator and a three position, selector valve.

Simple Hydraulic System Figure 15

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OPERATION (FIG. 15)

Hydraulic fluid, stored in the reservoir, is drawn into the hand pump via a pipeline attached to the bottom of the reservoir, through a non-return valve (NRV) and into the hand pump. The pump pushes the fluid through another NRV, via the pressure pipeline, to a 3-position selector valve. Depending on the position selected, it will either direct the fluid through a port, to one side of the doubleacting, linear actuator piston, or the other. Or it can be selected to the “Off” position, which locks the fluid in the actuator and prevents any movement of the piston in either direction. Fluid from the “non-pressure” side of the actuator piston, is diverted back to the reservoir by another port in the selector valve via a return pipeline. By installing an Engine driven pump (EDP) (See figure 16) the pilot is relieved from the physical task of hand pumping, which allows him to concentrate fully on flying the aircraft. The hand pump is still retained however, and is used as an Emergency back up, in case of an EDP failure. The hand pump is also used for testing the hydraulic system when the aircraft is on the ground during servicing operations and to build up the pressure in the system to operate the brakes before the engines are started.

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The use of an EDP creates a problem in that the pump is still maintaining pressure in the system when it is not needed during cruise flight, thereby wasting valuable engine power. The pump absorbs very little power when it is not moving fluid against an opposition. This problem is overcome by the installation of a pump, unloading valve. (Also called an; Automatic Cut-out valve). This valve relieves the pressure off the pump by diverting the fluid back to the reservoir. The fluid circulates freely from the pump, to the reservoir and back to the pump again with no opposition, thereby using very little engine power. The selector valve holds fluid trapped in the actuator, preventing any movement, or creep of the piston rod. (The actual operation of the unloading valve (cut-out valve) will be discussed in detail in a later section.)

Non Self-idling Hydraulic System Figure 16 When the piston has reached the end of it‟s stroke, pressure will build up in the system. This is relieved by the system pressure relief valve, which dumps the excess pressure fluid back to the reservoir. To maintain a positive pressure in the system when it is not operating, a nonreturn valve is installed in the pressure line from the pump, just after the unloading valve. This prevents the back-pressure being sensed by the pump and allows the unloading valve to divert the fluid back to the reservoir.. Issue 1 Mod 11.11

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An accumulator is installed to maintain a pressurised supply of fluid to absorb the initial pressure drop in the system when a selector valve is opened,. It also acts as a “shock absorber” to cushion the pressure surges of the fluid when the actuator pistons reach the end of their travel, thus preventing damage to the components. The accumulator has two compartments separated by a movable piston or diaphragm. One compartment is connected to the “pressure manifold” (pressure supply line) The other compartment is charged with air or nitrogen through a charging valve. (Nitrogen is used because all water vapour is removed during the processing of the gas at manufacture and the fact that Nitrogen is an inert gas). This nitrogen pressure is felt across the piston or diaphragm by the system fluid. To actuate any hydraulic system with the engines running, the pilot places the selector lever in the desired position. (Let us use a “Flap selection as an example) The system senses the “pressure-drop” and pressurised fluid flows from the accumulator, through the selector valve to the desired side of the actuator. The pressure-drop is also sensed by the unloading valve, which stops dumping pressure back to the reservoir via the return manifold and allows full pump pressure to feed the pressure manifold again during the operation of the actuator. This action also charges up the accumulator again until the system pressure relief valve senses the maximum system pressure, above which the relief valve dumps the excessive pressure back to the reservoir via the return manifold. Also at this time, the unloading valve once again senses the highpressure build-up and diverts the pump pressure back to the reservoir. The system continues to recycle in this manner whenever there is a demand for hydraulic power. As we continue to evolve the hydraulic system, you will notice that the reservoir has been altered to include a supply line to the EDP which is set higher in the reservoir than the emergency hand pump supply line. This extension is called a “standpipe” or “stackpipe”. It‟s function is to ensure that sufficient fluid is retained to operate the essential services such as brakes and landing gear extension, in the event of loss of fluid due to an excessive leak, down-stream of the brake and landing gear fluid pressure supply line. If the broken line, or leaking component can be isolated, there will still be enough fluid remaining in the reservoir to allow the emergency hand pump to lower the landing gear and operate the brakes. We can now add a few other items to the system to make it more usable. (Figure 17) To keep the fluid in the system clean, we need a filter through which all the fluid will pass. A typical location for the filter is in the return line just before the fluid enters the reservoir. This is called the Scavenge or Return filter. Here, it will catch all of the fluid, both that which is used to operate the actuators and that which circulates through the pump via the unloading valve. A second filter is installed immediately after the EDP to protect the rest of the hydraulic system from contamination in case of EDP failure.

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Typical Constant Delivery (non-self idling) Hydraulic Power Circuit Figure. 17

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SYSTEM COMPONENTS

11.15

INTRODUCTION

The following paragraphs describe various hydraulic components, including those used in the circuits. Some components are similar in construction and operation, but vary in the function they perform. Therefore, it is usual for the name of the component to indicate its purpose. Unfortunately, due to a difference in the terms used by the various manufacturers, some components with different names serve similar functions, such as a selector valve and a control box act fundamentally as a control valve. However, where different terms are used for similar components, it will be mentioned in the appropriate paragraph. 11.16

RESERVOIR‟S

The reservoir stores the hydraulic fluid. It supplies fluid to the system through a pump and receives the return fluid from the system. It accommodates the extra fluid caused by thermal expansion and compensates for slight leaks, which may occur throughout the system. Through it‟s design, it provides a reserve supply of fluid for emergency operation of systems which are essential for flight control and landing. This is done by the installation of a standpipe (stackpipe). It should also be observed that when the actuator piston rod is moved inwards, less fluid is required as the piston rod occupies space within the cylinder. With the actuator in this position, the surplus fluid is stored temporarily in the reservoir until the piston travels in the opposite direction.

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VENTED RESERVOIR

“Non-Pressurised” (Vented) Reservoir Figure. 18 CONSTRUCTION 1. Welded, Aluminium Alloy. 2. Vented Filler Cap. 3. Metal, gauze strainer, To prevent FOD (Foreign Object Damage) and contamination, during the filling operation. 4. Sight glass, Indicating Maximum, Minimum and Normal Operating fluid level. 5. Remote level indicator, (To gauge on pilot‟s instrument panel) 6. Inlet connection. (From system Return manifold) 7. Outlet connections, to Engine driven pump, (EDP) and Emergency handpump, (EHP) A Vented reservoir is the type normally fitted to a Piston-engine, un-pressurised, aircraft, which would normally operate below 20,000 feet altitude. The reservoir is located at a higher level than the EDP‟s to ensure a positive “head of pressure” supply of fluid throughout all normal flight manoeuvres. However. when flying through turbulent air, negative “g” forces or high roll angles, could cause a temporary loss of supply to the EDPs‟ allowing them to “run dry”, resulting in pump inlet cavitation. This could seriously damage the pump and cause it to fail. To compensate for this, a low-pressure pump is sometimes installed between the reservoir and the EDP‟s to ensure a positive head of pressure during such conditions.

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engineering 11.16.2

PRESSURISED RESERVOIR

Typical “Pressurised” Reservoir Figure.19 Jet and Turbo-prop aircraft that fly at altitudes higher than 20,000 feet require the hydraulic reservoir to be pressurises to prevent foaming of the fluid due to the low ambient air pressure at high altitudes, and to prevent pump cavitation in it‟s inlet. There are several ways in which pressurisation can be achieved: a

A nitrogen charged cylinder.

b

Cabin pressurisation air.

c

Engine Compressor/ Bleed air. (P3)

d

Hydraulic system pressure

Construction 1. Welded Aluminium Alloy. 2. Pressurised via a Pressure Reducing Valve (PRV) from Engine Compressor/ Bleed air, Cabin pressure, or from a Nitrogen storage cylinder. 3. Fluid quantity sight glass. (Indicating Max, Min, and Normal Operating fluid levels) 4. Max, pressure relief/ depressurising valve.

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Remote fluid level and temperature indicators (To gauges on pilots instrument panel) Return fluid de-aerator (Separates any air bubbles (foaming) absorbed into the fluid during pressure changes, allowing de-aerated fluid to fall back into the reservoir

“Pressurised” Reservoir using an Aspirator Regulator. Figure 20 Figure 20 shows a typical method of pressurising a reservoir using Engine bleedair (P3) or Pressurised Cabin air. Pressurisation can vary between 30 to 45psi depending on system design. Figure 21 shows a typical reservoir pressurised by hydraulic system pressure. Operation System pressure act‟s on one side of a small piston attached to the bottom of the main piston shaft, which exerts pressure on the fluid through the main piston. Pressure ratio‟s of about 50:1 are common for this type of reservoir. This means that a 3,000 p.s.i. system pressure can pressurise the reservoir fluid to 60psi. The fluid level in this type of reservoir is indicated by the amount the piston sticks out of the body at the bottom of the reservoir. Low fluid level is sensed by the “Level sensing switch”, which illuminates a light on the pilots instrument panel. In this pressurised condition, both the return line from the system, and the EDP supply line will be pressurised. Issue 1 Mod 11.11

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Pressurised Reservoir Figure 21

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REMOTE SERVICING POINT

On modern Jet and Turbo-prop aircraft it is common practice to install a Remote Servicing Point (Fig. 22.) in a convenient place, with easy access from ground level for maintenance personnel to carry out replenishment of the hydraulic fluid level. The Service point usually consists of; a

Self-sealing, quick release, filler point

b

Hand pump.

c

Reservoir de-pressurisation valve.

d

Level indicator.

e

Selector Valve SELECTOR VALVE SHOWN “CLOSED”

FWD

PRESSURE FILL CONNECTION TO SYSTEM A RESERVOIR

HAND PUMP

FILTER FILL SUPPLY TO STANDBY AND „B‟ SYSTEM RESERVOIR

SUCTION HOSE. (STAY‟S WITH AIRCRAFT)

Hydraulic Reservoir, Remote Servicing Point. Figure. 22 The servicing point allows fast and efficient servicing of the complete system contents at all reservoirs. Before connecting to the system, the Maintenance Manual procedures must be followed and all hydraulic systems must be in the prescribed position to ensure the correct fluid level is being indicated. The reservoir de-pressurisation valve must be operated to relieve the reservoir pressure.

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FILTERS

The extremely small operating clearances in modern hydraulic pumps, valves and components, require very effective filtration of the fluid. Therefore, filters are rated by the size of particles, which they can arrest. The size of these particles is measured in Microns. One micron is equal to one millionth of a meter or 0.000039 inch. An indication of just how small these particles are can be seen by the information in Fig. 23. (e.g. Particles as small as 40 microns are just visible with the naked eye)

Filters, which will remove particles less than 10 microns will maintain a very clean fluid

Relative Size of Particles Arrested by a Hydraulic Filter Figure 23.

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There are several types of filtration design‟s, two of the most common types used are shown in Fig 24. The paper element type, is made of specially treated paper folded into pleats to increase its surface area. The micronic element is wrapped around a spring wire coil to prevent it from collapsing under hydraulic pressure. Such filters normally have a bypass valve across the filtering element in case the filter becomes blocked with contamination, in which case the fluid bypasses the filter allowing “unfiltered” fluid into the system rather than “starving the system completely of fluid. Aircraft hydraulic filters are fitted at strategic locations throughout the system. The main locations being: L.P. (Low pressure) filter. H.P. (High Pressure) filter. By- pass filter.

Filters Figure 24

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ACCUMULATORS

11.17.1

PURPOSE

a.

To absorb fluctuations in pressure.

b.

To ensure immediate response and delivery of pressurised fluid on demand.

c.

To allow limited operation of systems when the EDP is not running.

Hydraulic fluid is non-compressible, and pressure can only be stored with compressible fluids. The compressibility effect can be gained by the using an accumulator. 11.17.2

CONSTRUCTION

Accumulators are constructed from high-strength materials such as cast, or machined, Aluminium alloys, or stainless steels. They consist of a container divided into two compartments by some form of movable, sealing partition, There are three types commonly used in aircraft hydraulic systems: Piston type, Bladder type, and Diaphragm type. The Piston type, (Figure 25.) is in the form of a cylinder with a floating piston. One compartment is connected to the system pressure manifold, the other is charged with compressed dry-air, or nitrogen, through a high pressure charging valve. The charging pressure is normally around, 1,500psi. (Approximately half system operating pressure).

Sliding Piston Accumulator Figure 25

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CHARGING OPERATION

As the accumulator is charged, (With “zero” system hydraulic pressure) the piston moves to the top of the cylinder until it reaches it‟s full stroke. The nitrogen pressure is then allowed to build up to approximately 1,500psi. The accumulator is now charged. A special, High-pressure (HP) valve, (See Figure 26) is then checked for leaks, and the dust cap installed. NOTE: HP valve cores are identified by a letter “H”, embossed on the end of the stem, and are NOT interchangeable with inner-tube and tubeless tyre cores.

AN812, High Pressure (HP) Air Valve for Accumulators and Air-Oil Shock Struts. Figure 26.

11.17.4

BLADDER & DIAPHRAGM TYPE ACCUMULATORS

CONSTRUCTION: Figure. 27 (A) & (B). These accumulators are spherical in shape, usually made of cast, or moulded aluminium, sometimes steel wire-wrapped. Others are of stainless steel. Both form two compartments as in the “piston” type. One to accept the dry-air, or nitrogen charge, the other connected to the fluid system pressure manifold. OPERATION: The operation is similar the “piston” type in that, the lower compartment is charged with dry-air, or nitrogen to a specified pressure, (usually between: 1,200 / 1,500psi). As pressure builds up in the hydraulic pressure manifold above the nitrogen pressure, hydraulic fluid is forced into the fluid compartment of the accumulator and deflects the bladder, or diaphragm, compressing the nitrogen until maximum system pressure is reached. (Usually around, 2,500 / 3,000 psi), Thereby providing a flexible cushion of In-compressible fluid via the medium of a compressible gas, transferred through a flexible bladder, or diaphragm. Some systems have a pressure gauge connected to the nitrogen side for quick monitoring during servicing, without disturbing the charge valve.

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Bladder and Diaphragm Type Accumulators Figure 27 Charging Valves On the previous page, Figure 26. illustrates a simple high-pressure valve, which seals through the valve core. Figure 28 shows two types of metal-to-metal sealing vales which are more commonly used. The AN6287-1 valve does not depend on the valve core to provide the seal, but seals through metal-to-metal contact between the stem and the valve body. To release air, loosen the swivel nut one turn and de-press the valve core. To charge air, connect the special, high pressure hose fitting and apply pressure through a regulator valve with the swivel nut open at leased one full turn. CAUTION: Use great care and protect eyes and skin while charging, or releasing high pressure air, or nitrogen. The MS28889-1 valve is also used in many high pressure systems and is similar to the AN6287-1, but with different features. a) The swivel nut is the same size as the hexagon valve body, whereas the swivel nut on the AN valve is smaller. b) The stem is retained in the valve body by a roll pin to prevent the stem from being unscrewed fully. c) There is no valve core in this type, just the metal-to-metal sealing surface. CAUTION: ALWAYS install the special, high pressure valve cap after you have checked for leaks, and on completion of the work. Issue 1 Mod 11.11

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Charging Valves Figure 28

Deflation Cap Figure 29 Figure. 29. Shows a special cap for safely deflating an accumulator, or air-oleo strut under controlled conditions. Screwing on the cap progressively, pushes the valve core off its seat slowly, allowing gradual de-pressurisation to take place.

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PRESSURE GENERATION (HYDRAULIC PUMPS)

Hydraulic power is transmitted by the movement of fluid by a pump. The pump does not create the pressure, but the pressure is produced when the flow of fluid is restricted. We often use a hydraulic analogy for studying electricity, Therefore, we will use our knowledge of electricity to help us understand hydraulic power. The flow of fluid in a line is equivalent to the flow of electrons in a wire, the current (I). The pressure that causes the flow is the same as the voltage (E), and the opposition to the flow of fluid is the same as the resistance (R). If there is very little friction in the line, very little pressure is needed to cause the fluid to flow. In Fig. 30. we have a very simple electrical system, consisting of a battery, an ammeter, a voltmeter, and a resistor. The ammeter measures the flow of electrons in the circuit, and the voltmeter measures the voltage (pressure) drop across the resistor. The hydraulic system in Fig. 31, is very similar in its operation. The pump moves the fluid through the system and may be compared to the battery, which forces electrons through the circuit. The flowmeter measures the amount of flow, the valve acts as a variable opposition to the flow, and the pressure gauge measures the pressure drop across the valve. When the variable resistor is set to its minimum resistance, the current will be maximum and there will be a minimum voltage drop across the resistor. In the same way, when the valve is fully open, there will be a maximum flow of fluid and a minimum pressure drop across the valve. When the resistance in the electrical circuit is increased, the voltage across the resistor will increase and the current will decrease. In the hydraulic system, as the valve is closed, the flow will decrease and the pressure will increase. When the valve is fully closed, there will be no flow and the pressure will increase to a value as high as the pump can produce. If the pump is of the constant displacement type, there must be some provision in the system to relieve the high pressure; otherwise the pump will be damaged, or components in the system damaged.

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HAND PUMPS

Single-action, piston type pumps, move fluid on one stroke only, while doubleaction pumps move fluid on both strokes. Most modern aircraft hydraulic systems use the double-action type because of their greater efficiency. Figure 32 illustrates the operating principle of a typical double-action hand pump. This type is called a “Piston rod displacement pump” because the pumping action is caused by the difference in area between the two sides of the piston, due to the piston rod area displacement. In view (A), the handle is pulling the piston to the left. Fluid is drawn in through the inlet check valve, When the piston reaches the end of its stroke, chamber “1” is full of fluid and the inlet check valve closes by the action of its spring. As the handle is moved to the right, as in view (B), the piston is pushed to the right, forcing fluid through the outlet check valve and into chamber “2”. The volume of chamber “2” is smaller than chamber “1” because of the piston rod area, therefore, the excess fluid is displaced through the outlet port. On the return stroke, (To the left again) the remainder of the fluid in chamber “2” is also displaced through the outlet port. At the same time, a new charge of fluid is being drawn into chamber “1”, from the inlet port, through the inlet check valve.

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Hand Pump Operation Figure 32 (B)

11.20

SUCTION BOOST PUMPS

This is a low-pressure pump, (Approx. 100 psi) whose prime function is to provide a positive pressure to the inlet side of the main system pressure pump, to prevent cavitation. It is located between the reservoir fluid supply and the Engine-driven pump (EDP) inlet. The pump can be mounted independently, or attached to the reservoir. It is normally powered by a 3-phase electric motor, and in some cases, by a hydraulic motor driven by system pressure. Many modern hydraulic pumps have a “Spur-gear” type pump built into the body of the main pressure pumps. (This will be discussed in more detail under Variable displacement, piston type pumps). In the event of a boost pump failure, The EDP (Main pressure pump) and system will still operate, but at a possible reduced efficiency with a risk of cavitation of the EDP. in severe cases.

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POWERED PUMPS

The only function of a pump is to move fluid through the system. There are a number of ways powered pumps can do this. The two basic types are: 1. Constant Volume/Fixed displacement (Non-self idling). Figures. 33 & 34. 2. Constant Pressure/Variable displacement (Self-idling). Figures. 43 & 44. A Constant Volume/Fixed displacement, (Non-self idling) pump moves a specific volume of fluid for each revolution of the drive-shaft. It requires some form of Regulator, or Relief valve (Sometimes called a; Cut-out, or Unloading valve) in the system to relieve the pressure which builds up when the pump delivers more fluid than the system requires. (See Figs. 16, 17, and 33.).

Constant Volume/Fixed displacement (Non-self idling) Pump System Figure 33 11.21.1

CONSTANT VOLUME FIXED DISPLACEMENT PUMPS

(Non Self- Idling) The most common type of Constant volume (CV) pump for medium-pressure systems, is the Gear pump type. (See Fig. 34.) These pumps are very rugged and dependable, with few moving parts, relatively easy and in-expensive to manufacture, compared with other types.

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The left-hand gear is driven by the engine through a splined shaft. This gear rotates in a close fitting housing and drives the right-hand gear housed in the same manner. As the gears rotate in the direction shown, fluid is transported between the teeth around the outside of the gears, from the inlet side of the pump. When the teeth mesh with each other, in the outlet chamber, fluid is displaced into the outlet side of the pump. A very small amount of fluid is allowed to leak past the gears and around the shaft for lubrication, cooling, and sealing. This fluid drains into the hollow shafts of the gears where it is picked up by the low pressure on the inlet side of the pump. A relief valve holds the oil in the shafts until it builds up to about 15 psi. This is called; case pressure. This is maintained so that, in the event of the shaft, or seal, becoming scored, fluid will be forced out of the pump rather than air being drawn in.

Gear Type Hydraulic Pump Figure.34 Spur gear pumps provide a good, non-pulsating, high flow rate, but are limited to pressures up to about 800psi. Because of this, they are more commonly used on smaller aircraft, but also as pressure back-up pumps for the more powerful, piston-type pumps on larger aircraft, who‟s hydraulic systems operating pressures are between: 1,200 to 3,000psi.

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engineering 11.21.2

PISTON PUMPS

Aircraft hydraulic systems that require a relatively small volume of fluid under a pressure of 2,500 psi or more, often use fixed-angle, Multi-piston pumps as shown in fig. 28. a

Axial Piston Pump, (Figure 35)

This type of pump consists of a bronze cylinder block, rotated by a splined drive shaft, driven by the engine, through a universal link. The cylinder housing is mounted at a fixed angle to the drive shaft and bearing housing. The cylinder block usually has seven, or nine axially-drilled holes, which accommodate, “High precision, close fitting pistons”. These in turn are attached by a ball-jointed rod to a pump drive plate which is rotated by the engine. As the piston and cylinder block assembly are rotated by the drive-shaft, the pistons on one side (upper pistons) are at the bottom of their stroke, and open to the Inlet port. due to the angle of the housing. The pistons on the opposite side (Bottom pistons) are then at the top on their stroke, open to the Outlet port. (See fig. 35)

Fixed Angle, Axial, Piston Type Hydraulic Pump Figure 35 The stroke (Displacement) of the piston is dependent on the angle of the cylinder housing to that of the bearing housing. As the whole assembly is rotated, fluid is drawn in by the piston moving down in the one side of the cylinder block, while fluid is being pushed out by the piston moving up in the opposite side of the cylinder block. Issue 1 Mod 11.11

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A valve plate with two crescent-shaped openings cover the end of the cylinders. One above the pistons moving up, thereby pushing fluid through the Outlet port. The other, above the pistons moving down, drawing fluid into the cylinder, through the Inlet port. b

Radial Piston Pumps

In this type of fixed volume pump, the cylinders are arranged radially around an eccentric crankshaft. (See Fig.29A & B). When the crankshaft is rotated, the pistons move outwards in each cylinder, forcing pressurised fluid into the annular outlet port through each cylinder delivery valve. When each piston is at the bottom of it‟s stroke, the pistons uncover the inlet port, allowing a fresh charge of fluid to enter each cylinder. The fresh charge of fluid is then compressed as the piston moves outwards again forcing fluid once more through the delivery valve. This process is repeated with each revolution of the eccentric crankshaft Typical Radial, Piston-type, Hydraulic Pump - Side View Figure 36

Radial Piston Hydraulic Pump – End View Figure 37 Issue 1 Mod 11.11

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Typical Radial, Piston-type, hydraulic pump Constant Volume/Fixed Displacement Figure 38

Operation of radial, piston-type pump. (constant volume/ fixed displacement) Figure 39 Issue 1 Mod 11.11

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engineering Vane Pumps

These pumps are used in systems, which required moving a large volume of fluid, but at relatively low pressures. The vanes are allowed to float freely in slots machined in the rotor, and are held in place by a spacer. This rotating assembly is attached to a drive shaft and is driven by the engine, or, an electric motor. The rotating assembly is mounted “concentrically” in a ported, steel sleeve which is pressed into a cast, aluminium housing.

OPERATION: As the rotor turns in the direction of the arrow, (Fig. 29.) the volume between the vanes on the inlet side increases, while the volume between the vanes on the outlet side decreases. This change in volume draws fluid into the pump through the inlet port, and discharges it through the outlet port and into the system

Vane-type Hydraulic Pump (Constant Volume Fixed Displacement Fig. 40 This type of pump is normally used on light aircraft, particularly in “POWERPACK” type hydraulic systems, but is more generally used in fuel and pneumatic systems than hydraulic systems.

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engineering 11.21.3

UNLOADING (CUT-OUT) VALVE

An Unloading (Cut-out) Valve of some sort is needed when a Constant volume/Fixed displacement pump is used to relieve the engine of the pump loading when there is no demand on the hydraulic system.

Fig. 41 Unloading (Cut-out) Valve during system demand

FIG. 42 UNLOADING (CUT-OUT) VALVE – PUMP IDLING POSITION Issue 1 Mod 11.11

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engineering 11.21.4

CONSTANT PRESSURE/VARIABLE DISPLACEMENT PUMP

A Constant pressure/Variable displacement, (Self-idling) pump, only moves an amount of fluid, which the system requires, hence the term: Variable displacement. As the pressure in the system builds up due to no actuation (no fluid movement), the pump delivery displacement is automatically reduced to noflow. By varying the pump output, the system pressure can be maintained at a constant, within the desired range without the use of Regulators (Cutout/Unloading valves). It allows the pumps to turn without delivering fluid to the system. However, this can cause overheating of the pump. To prevent this, fluid is by-passed back to the reservoir, by the LP spur-gear back-up pump, ensuring a continuous flow of fluid through the HP piston pump at all times, even when there is no fluid delivery to the system. Thus providing cooling of the pump.

Fig. 43 Constant Pressure/Variable displacement, (Self-idling) hydraulic Pump This type of pump is similar in construction to the fixed volume, axial-piston type, (Figure 35) It is normally a 2 stage pump. The first stage usually consists of a low pressure (LP), high volume, spur gear pump, (similar to the Radial pump shown in Figure. 37). This ensures a positive supply of fluid to the second stage, high pressure (HP), axial, Multi -piston pump, the cylinder block of which is driven by a common drive shaft.

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The piston stroke is varied by a Yoke mechanism, sometimes called a Swashplate, or Cam. (See Figures. 43. & 44.) The pistons are attached to shoes that rotate against the stationary Yoke. The angle between the Yoke and cylinder block is varied, to increase, or decrease the piston stroke. This action is carried out by a Servo Control Piston, which senses “system pressure”. This pressure pushes the Servo Control Piston against the return spring pressure, and reduces the Yoke angle, thereby, reducing the HP piston strokes. When the Yoke is at 90º to the drive shaft, (Perpendicular to the pistons) the piston stroke is zero and there is no flow of fluid, therefore, no load on the drive-shaft.

Fig. 44 Schematic of Constant Press./Variable displacement pump

Fig. 45 Constant Press/Variable displacement (Self-idling) hydraulic pump Issue 1 Mod 11.11

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engineering 11.21.5

STRATOPOWER PUMPS

As previously discussed, some kind of “unloading valve” is required when using a constant displacement pump. But the same force, (system operating pressure) which controls this valve can be used to control the output of the variable displacement pump. Figures 43, 44 and 45 shows variable displacement pumps, which are controlled by a spring-loaded piston, which moves a pivoted yoke, or swash-plate to adjust the stroke of the delivery pistons, thereby regulating the fluid flow. Another commonly used variable displacement pump for high pressure aircraft hydraulic systems is the Stratopower demand-type pump illustrated in Figure. 38.

Fig. 46 Constant Pressure/Variable displacement, (Selfidling) hydraulic Pump. (Stratopower Pump, demand-type) This pump uses nine axially-orientated pistons and cylinders. The pistons are driven up and down in the cylinders by a fixed-stroke cam. The stroke of the pistons is the same regardless of system demand. In this type, the effective length of the piston stroke controls the amount of fluid delivered to the system. This type of pump usually has a delivery capacity of between 22-37gpm. (gallons per minute) and maintains a nominal supply pressure of 3,000psi.

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engineering 11.21.6

OPERATION

The forces which control the pump output and system pressure is between the compensator spring and the compensator stem piston. Pump out-put pressure is ported around the compensator stem which acts as a piston and opposes the compensator spring. As the pressure increases, the stem piston compresses the compensator spring.

Fig. 47 Stratopower pump, flow and pressure controlling mechanism. Issue 1 Mod 11.11

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The spider, which is connected to the compensator stem, moves the sleeves up and down the delivery pistons. When the pressure is high, the stem piston moves the spider, compressing the compensator spring and uncovers the relief holes near the bottom of the delivery pistons during the full stroke. This allows the fluid to be dumped during the compression stroke to the inlet side of the pump, preventing fluid flow through the check-valves and into the system. The pump is allowed to deliver a small amount of fluid even at it‟s minimum stroke to ensure adequate lubrication and cooling of the pump at all times during operation. When system pressure drops, the compensator spring forces the stem and spider assembly down the piston, covering the relief holes at the bottom of the delivery piston stroke. This prevents bleed-off of fluid during the compression stroke. The compressed fluid is then forced out through the check valves and into the system to meet the fluid demand. During any intermediate pressure condition the spider sleeves cover the relief holes at some point along the discharge piston‟s stroke, thereby maintaining system pressure and fluid flow to the required value. The value of the compensator valve is set by the pressure adjusting screw, which varies the tension of the compensator spring.

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engineering 11.22

EMERGENCY PRESSURE GENERATION

A failure of the hydraulic supply circuit may have a disastrous effect on the operation of the aircraft. If such an emergency arises, provision must be made to supply the services which are hydraulically operated by some alternative source of power. There are several ways in which this can be achieved; a

Hand-pump operated by the pilot,

b

Duplication of supplies,

c

Electrically operated AC or DC pumps,

d

Compressed air, Air turbine motor driven pumps, (A.T.M.)

e

Ram-air turbine pumps, (R.A.T.)

11.23

HAND PUMPS

The Hand-pump operation has been explained in Chapter 6.1 Almost all aircraft with a hydraulic power system installed have an Emergency hand-pump mounted in the cockpit or flight deck. It is usually mounted and stowed under the floor, between the pilot‟s seats, thereby allowing either pilot or co-pilot to operate it with relative ease while still flying the aircraft. A quick-release access cover is usually marked in Red or Yellow and Black stripes, indicating; “Emergency operation”. The hand-pump is connected in parallel with the Engine driven pump (EDP) but has an independent fluid supply line from the Reservoir which draws hydraulic fluid from a lower level in the reservoir than the EDP supply, This ensures a positive supply of fluid if the level is low. (See Figure 18) In some systems the hand-pump is also used to initially pressurise the system to ensure adequate system pressure to operate the emergency or park brake system prior to towing, parking and engine start-up of the aircraft. 11.24

DUPLICATION OF SUPPLY

On Multi-engined aircraft, where hydraulic power is used extensively, and also as a safety factor, it is often necessary to have a power circuit using two or more pumps to meet the demand when most systems are being operated at the same time. i.e. (Landing and Take-off) The circuit illustrated in Figure 48 is fitted with two self-idling pumps which, should one pump fail during flight, the remaining pump will still provide fluid flow but at half the normal rate. The primary purpose of the Accumulators in this circuit is to dampen out the pulsation‟s of the pumps, also to give speedier operation of components when initially selected, and to provide a source of hydraulic power when the engine-driven pumps (EDP‟s) are not working. Issue 1 Mod 11.11

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Figure 48

Fig. 49 Typical Engine-Driven Hydraulic Pump (EDP) as fitted on Boeing 737 Issue 1 Mod 11.11

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Multi-engined aircraft normally have one EDP mounted on each engine similar to the one in Figure. 49. However, some aircraft like the Lockheed L1011 Tri-Star, have one EDP driven by each wing mounted engine, (No‟s. 1 & 3 engines.) and two EDP‟s driven from the rear fuselage mounted engine. (No. 2 engine.) This is to ensure adequate flow and pressure supply to a large and complex hydraulic system and to cater for redundancy and continued safety in the case of an engine or pump failure. Modern Jet transport aircraft now have at least two hydraulic systems completely independent of each other with duplicated actuation of all primary hydraulically powered flight control systems. Figure 50 shows a schematic diagram of the Boeing 737 hydraulic system. This consists of two Main systems (Systems A & B) with EDP‟s drawing fluid from separate reservoirs and a Standby system as an additional back-up in case of failure of one or both main systems.

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Fig. 50 Dual Hydraulic System Schematic Diagram (Boeing 737

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engineering 11.25

ELECTRIC MOTOR DRIVEN PUMPS (EMDP‟S) 115V AC

It is common practice to install Electric Motor Driven Pumps (EMDP) primarily as a back up to the EDP when system demand is high, but also to provide hydraulic power in case of EDP or engine failure.

Fig.51 Typical 115v. AC Motor-driven Pump as fitted to Boeing 737 aircraft Issue 1 Mod 11.11

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A three phase 115v AC EMDP is connected to the main hydraulic circuit in parallel with each EDP. It draws its fluid from the same reservoir, but its fluid supply line is mounted lower in the reservoir to ensure a continued supply to the EMDP when the fluid level is low. These pumps are very similar in operation to the EDP‟s but with a lower capacity, usually about 6-10gpm (gallons per minute) and maintain a pressure of about 2,700 p.s.i. Hydraulic fluid enters the pump by way of the electric motor housing to provide cooling of the pump and motor assembly during operation. On some aircraft a Low capacity (3 g.p.m. at 2,700 p.s.i.) 28v DC motor driven pump is installed as an Emergency hydraulic power source which is also used to provide initial hydraulic pressure to charge up the system for brake operation, prior to towing the aircraft or engine starting. 11.26

AIR TURBINE MOTOR DRIVEN PUMPS (ATM‟S OR ATDP‟S)

Some aircraft such as the Airbus 300 series and B767 use hydraulic pumps operated by air turbines, which are driven by bleed air from the engines. These Air-turbine driven pumps (ATDP) receives pressurised air from the aircraft‟s main bleed air system. The flow of air is controlled and modulated by a solenoid operated pressure regulator and shut-off valve to maintain the turbine speed within set parameters. The turbine is connected by a shaft to the pump. (See figure 52)

Fig. 52 Typical Air Turbine Motor driven hydraulic pump. (ATDP) Issue 1 Mod 11.11

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engineering 11.27

POWER TRANSFER UNITS (PTU‟S)

PTU‟s consist of a hydraulic motor, which is supplied fluid under pressure by one hydraulic system. This motor turns a drive shaft, which powers a hydraulic pump, which is connected to a second hydraulic system in the aircraft. The PTU is an integrated unit housed in one casing. (See Fig. 53) The purpose of the PTU is to use pressure from one system to power the motor which drives the pump to provide pressure in the other system. The PTU motor may be isolated from pressure when system operation is normal but may be selected manually or automatically (by a pressure switch) in the event of a pressure drop or failure of the other system pumps. The B737 incorporates a PTU to supply pressure to the slat system automatically in the case of reduced pressure.

Fig. 53 Typical Power Transfer Unit (PTU) Schematic.

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engineering 11.28

HYDRAULIC RAM AIR TURBINES (HYRAT‟S)

HYRATS may be used as an emergency source of hydraulic power in the case of major failure within the normal system. The HYRAT consists of a turbine (similar in appearance to a small propeller) which is normally stowed in a compartment in the fuselage as in the Lockheed L1011 trustier and Boeing 767 aircraft. (See Fig. 54.)

Fig. 54 Hydraulic Ram Air Turbine (HYRAT) Pump Unit. It is only deployed in the case of a major hydraulic failure to provide minimum hydraulic supply for the safe recovery of the aircraft. The HYRAT may be deployed automatically or by manual selection. Pressure output is governed by varying the blade angle in response to aircraft speed and pressure demand. Issue 1 Mod 11.11

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engineering 11.29

HYDRAULIC VALVES

The valves used in hydraulic systems may be divided into pressure control and flow control valves. a

A pressure control valve adjusts, regulates and/or limits the amount of pressure in the power supply system or any component circuit.

b

A flow control valve selects and directs the flow of fluid through the system or circuit in a particular direction and is not normally concerned with the pressure.

11.30

PRESSURE CONTROL VALVES

11.30.1


The flaps are comparatively fragile and if they are lowered when the aircraft is flying at high speeds, are liable to be damaged by the airflow. The flaps are designed to be used only when the aircraft is landing or taking-off. To prevent such damage occurring, a pressure relief valve is provided in the circuit. This valve, which acts as a “blow-back” valve, bypasses pressure fluid in the Down line to the return line. In effect, the valve enables the flaps to “blow-back” if they are left down and the aircraft speed is increased. It also prevents the pilot from lowering the flaps at high air speeds.

Fig. 55 Operation of “Pressure Relief Valve” (PRV)

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engineering 11.30.2

PRESSURE REGULATORS

In Chapter. 6.3 We discussed the two basic types of pumps used, The Constant Volume/Fixed Displacement (Non-self idling) type. The Constant Pressure/Variable Displacement (Self-idling) type. It was stated that; the Non-self idling type required an Unloading, or Cut-out valve to relieve the pressure which builds up in the system when the out-put from the pump is greater than the system demand. It also regulates the system pressure within a normal operating range. A complex Unloading valve was discussed previously. A simpler pressure regulator (the Balanced-type), is illustrated in Figure 56.

Fig. 56 “Balanced-type” pressure regulator valve. OPERATION The pump delivers a fluid flow through the NRV into the system and charges the Accumulator with fluid and pressure builds up in the system. This pressure is sensed on the under-side of the regulator piston. The same pressure is sensed on the upper surface of the ball, forcing it onto its seat as the pressure increases. The spring is acting downwards against the piston and a balance of forces is Issue 1 Mod 11.11

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reached between the fluid pressure on the ball, the spring pressure on top of the piston, and the system pressure acting upwards under the piston. At the condition of balance, when the pressure is 1,500psi, there will be a force of 1,500 pounds (lbs) pushing up on the piston. The total downward force of 1,000 lbs applied by the spring and a 1/3 of 1,500 lbs (500 lbs) of fluid force pushing down on the ball. If the system pressure rises above this balanced pressure, the spring pressure is constant and not effected by hydraulic pressure, therefore the piston will move up and lift the ball off its seat. This allows the pump delivery (flow) to return to the reservoir with very little resistance and therefore virtually zero pressure. The NRV holds the pressure trapped in the system and the accumulator. This condition will continue until the pressure in the system drops to 1,000 psi, at which point the spring will force the piston down, allowing the ball to re-seat and the pressure will rise again to the unloaded pressure of 1,500 psi. This gives a system cycling pressure of: 1,000 – 1,500 psi. 11.30.3

THERMAL RELIEF VALVE

This valve is designed to relieve excessive pressure caused by expansion of the hydraulic fluid due to increase in temperature. It is situated in a pipeline between components where the fluid is in a closed circuit, such as between an NRV and an actuator, where there is a hydraulic lock. The excessive pressure is relieved back to the reservoir via the return line. The restrictor pack ensures that only the slow pressure changes from thermal expansion effects the operation of the valve.

Fig 57 Thermal Relief Valve Issue 1 Mod 11.11

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engineering 11.30.4

PRESSURE REDUCING VALVE

Some hydraulically operated components require a much lower pressure than system pressure to operate them. In such cases a Pressure Reducer Valve similar to the one in figure 58 is used. This valve reduces system pressure by the action of a balance between hydraulic and spring forces.

Fig. 58 Pressure Reducer Valve

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engineering OPERATION

Assume that the piston in figure 58 has an area of one square inch (1in²) and is held on its seat by the large spring with 100 pounds force (lbsf). The piston has a shoulder area of ½ square inch, which is acted on by the full 1,500psi. system pressure. The reducer valve seat area is ½ square inch (Same as piston shoulder) and is acted on by the 200 psi reduced pressure. A small hole in the piston bleeds fluid into the chamber behind the piston and the relief valve maintains this pressure at 750 psi. This relief action is determined by the pressure inside the piston cavity, acting on one side of the relief ball and the spring, and reduced pressure (200 psi) acting on the opposite side. When the reduced pressure drops, the hydraulic force on the ball drops, allowing it to unseat. This decreases the hydraulic force on the piston and allows it to move up. Fluid now flows into the reduced pressure line and restores the 200-psi. This increased pressure closes the relief valve so that the pressure behind the piston can again increase up to 750 psi and seat the valve. The small bleed hole also prevents the piston from chattering by giving the piston a relatively smooth action. The piston remains off its seat just enough to maintain the reduced pressure as it is used.

11.31

FLOW CONTROL VALVES

Flow control valves in hydraulic systems control fluid flow and the direction of flow. They may control manually (direct operation by flight or ground crew) or automatically (by flow, pressure or remote sensing devices) Flow Control valves can be mechanically, electrically or hydraulically operated. The valves may be of the ball, sleeve, poppet, rotary, piston or sliding- spool type.

11.31.1

NON-RETURN (CHECK) VALVE

This valve is the simplest of all flow control valves and is used in most systems. Its basic function is to allow fluid flow in one direction only. The different types are shown in fig. 59. An NRV or Check valve, is always fitted just down-stream of the pump to ensure there is no reverse-flow through the pump which could cause damage to it when stationary or not in use. Some applications require full flow in one direction and a restricted flow in the other. This valve is known as a Restricted or Orifice Check valve (fig. 60)

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HYDRAULIC POWER

engineering

NON-RETURN (CHECK) VALVES

Fig 59

Fig 60

A

Ball Check Valve

A

Orifice Check Valve

B

Cone Check Valve

B

Orifice type, installed in

C

Swing Check Valve

a landing gear system

(Flapper Valve)

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engineering 11.31.2

SELECTOR VALVES

Selector valves may be considered to be the first valve in the Services System and not part of the Power System. The purpose of the Selector Valve is to direct fluid to the appropriate side of an actuator, and to provide a return path for the fluid displaced from the opposite side of the actuator, back to the reservoir. Many flow control valves are simple four-way valves, connecting the pressure and return lines to alternate sides of the actuator, without a neutral position, however, control valves in open-centre systems often lock fluid in the actuator while providing an idling circuit for the pump.

Fig. 61 Manual Selector Valves Issue 1 Mod 11.11

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Figures. 61. Illustrates the Ball, Rotary and Sliding-spool type valves, which are normally used in relatively low-pressure actuation. Higher-pressure systems require a more positive shut-off of fluid flow and Poppet-type selector valves are often used. OPEN CENTRE, POPPET TYPE, SELECTOR VALVE OPERATION When the Control Selector handle is in the Neutral position, Poppet valve 3 is off its seat. Fluid flows straight through the valve from the pump to the next selector valve and on to the reservoir. All the other poppet valves are closed. When Gear Down is selected, movement of the cams causes valve 3 to close, and valves 1 and 4 open, redirecting pump pressure to the Gear Down side of the gear actuator, through valve 4 Fluid on the other side of the actuator piston is then redirected back to the reservoir through valve 1 via the return line. When the actuator reaches the end of it‟s travel, the pressure increases to a specific value and operates a mechanism, which returns the selector handle to the neutral position, thereby closing valves 1 and 4 and reopening poppet valve 3 When Gear Up is selected, valve 3 once again closes and valves 2 and 5 open. This directs pump pressure to the Gear Up side of the actuator through valve 2 Fluid on the other side of the actuator is redirected back to the reservoir through valve 5 via the return line. When the actuator again reaches the end of it‟s stroke, the pressure increase is again sensed by the return mechanism and the selector handle is returned to the Neutral position, thereby closing valves 2 and 5 and reopening poppet valve 3 again. Note: This type of selector requires a pressuresensing device which moves the selector handle back to the Neutral position when the actuator is fully extended or retracted. Fig 62 Open Centre, Poppet Type, Selector Valve

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engineering

ELECTRICALLY OPERATED SELECTOR VALVES. These valves use electrically operated solenoids to control the position of a spool valve which, in turn, controls the direction of fluid flow to the system actuator. Switches located on the flight deck, or remote sensors, operate these valves. The advantage of this type over mechanical valves is the elimination of bulky lever mechanisms, torque tubes, bell-cranks, levers and pulley‟s, which add extra weight to the aircraft. On large Transport aircraft, this is especially important when considering the large distances from the controlling point to the actuation point. “Fly-by-Wire” systems are modern examples of this method of Power Control. SINGLE SOLENOID TYPE, SELECTOR VALVE The selector illustrated in Figure. 63. is a Single solenoid, two-way valve. Typically used for emergency operation of the Flaps or Landing Gear

Fig. 63 Electrically operated, Slide-valve, Selector (Single Solenoid) OPERATION With the solenoid de-energised, the pilot valve is spring-loaded against the return seat, and fluid from the emergency power system passes to both sides of the slide valve. Since the right-hand end of the valve is a larger diameter than the left, the valve is moved to the left by the greater force, and system pressure fluid passes to the actuator to extend its ram. Fluid from the opposite side of the actuator passes through the slide valve, to the reservoir, via the return line.

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With the solenoid energised, the pilot valve is held against the pressure seat, and supply pressure acts on the left-hand side of the slide valve only. The right – hand side being open to return, thereby forcing the slide valve to the right. This directs system pressure fluid to the actuator to retract its ram. Fluid from the opposite side of the actuator, being open to return to the reservoir, via the return line. DOUBLE SOLENOID TYPE This valve is similar to the single solenoid valve but is used where the service has intermediate positions. With both solenoids de-energised, the valve will hold the service in any rigid position (Hydraulically locked) but with the supply pressure isolated from the utility system.

Fig. 64 Double Solenoid type Selector valve OPERATION With both solenoids de-energised the ball valves are held against their respective return seats, (same as the Single Solenoid valve). In this position, system pressure is directed to identical area pistons on both sides of the Spring-loaded shuttle valve. With no hydraulic power in the system, the springs, which are also of equal tension, hold the shuttle valve in the centred position, thereby shutting off, both lines to the actuator and creating a hydraulic lock in the actuator. When the left-hand Solenoid is energised, the ball valve is held against the pressure seat. This allows the pressure on the left-hand side of the shuttle valve piston to be vented to the return line through the left-hand shuttle valve chamber, thereby, causing a pressure imbalance which forces the shuttle valve to the left. This allows pressure to one side of the actuator, and directs the other side of the actuator to the return line through the right-hand chamber of the shuttle valve. Issue 1 Mod 11.11

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engineering 11.31.3

PRIORITY VALVES

These valves are similar to Sequence valves except they are opened by hydraulic pressure rather than by mechanical means. They are called priority valves because such devices as “Wheel-well doors”, which must operate first, require a lower pressure than the Main Landing Gear. The valve will shut off all the flow to the Main Gear until the doors have actuated to the fully Open position and the pressure builds up at the end of the actuator stroke. The priority valve senses the pressure build-up and opens, allowing fluid to flow to the Main Gear actuators.

Fig. 65 Typical Priority Valve Operation

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engineering 11.31.4

SEQUENCE VALVES

Most modern aircraft with retractable landing gear have landing gear doors, which close in flight to cover the wheel well to ensure a streamlined airflow. When the gear is selected “UP” or “Down”, by the pilot, the gear doors must open first before the gear starts to retract. For this reason, a Sequence Valve is sometimes installed. These are usually similar in construction to Check valves (NRV‟s) which allow a flow of fluid in one direction, but may be opened manually to allow fluid to flow in both directions. These valves are similar to Priority Valves regarding their function, by allowing one hydraulic component to function before another is allowed to function. The difference between them is that Priority valves are controlled by fluid pressure, whereas Sequence valves are controlled by mechanical displacement of a plunger, which moves a Ball valve off its seat to redirect fluid to another component.

Fig 66 Section through a typical Sequence Valve

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The illustration in Fig. 67 (below) Shows the location and typical use of Sequence Valves in a simple Landing Gear system. It explains the basic sequence of operation as the gear is selected “UP”.

(a)

i Ii iii

Gear selected “UP” Wheel-well door “Fully OPEN” Gear retracting

(b)

i ii iii

Gear fully “UP” Door Sequence valve “OPEN” Wheel-well door “Fully CLOSED”

Fig. 67 Sequence Valve location and operation, in landing gear system

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HYDRAULIC FUSES

PURPOSE These special valves are used to block off fluid flow if a serious leak should develop. There are two types of hydraulic fuse. The first type shuts off the fluid flow if the pressure drop across the fuse falls below a specified limit. The second type shuts off the fluid flow after a specific amount of fluid has flowed through it

Pressure Sensing Fuse Valve This fuse senses the pressure drop across the valve.

Fig 68 Pressure Sensing Fuse Valve During normal flow through the valve, the spring keeps the piston against its seat. If a serious leak or pipe failure occurs downstream of the outlet (B) the pressure drop is sensed across the piston, which generates a force greater than that of the spring. This allows fluid pressure upstream at the inlet (A) to move the piston to the right, thereby shutting off the fluid flow. This condition will be maintained until the system pressure as inlet (A) is relieved, i.e. The system is shut down, allowing the spring to return the piston to its normal operating position.

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Flow Sensing Fuse Valve Operation: This fuse does not operate on the pressure sensing principle. It will shut off the flow after a given amount of fluid has passed through it. In the static condition, all the ports are closed off. When fluid begins to flow in the normal direction of operation, system pressure on the sleeve valve moves it to the right against the spring pressure, thereby opening the ports and allowing fluid to flow through the valve. During this time some fluid passes through the metering orifice and progressively moves the piston to the right until it shuts off the primary delivery ports which stops fluid flow. When fluid flows in the reverse direction, the sleeve valve and the piston are both moved to the left which keeps all the ports open and allows fluid to flow through the fuse unrestricted, in the opposite direction. Normal operation of the unit protected by this type of fuse doesn‟t require enough flow to allow the piston to drift completely to the right and seal the primary delivery ports. Only when there is a serious leak will there be sufficient fluid flow to move the piston to the right and close off the primary delivery ports. Fig. 69 Flow Sensing Hydraulic Fuse

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POWER DISTRIBUTION

As previously stated, the hydraulic system in an aircraft may be used for operating various services, such as landing gear, flaps, airbrakes, wheel brakes, control surfaces, nose-wheel steering, etc. As it will be necessary to operate these services independently, provision must be made to ensure adequate fluid flow and pressure at all times, not only during operation of all the primary circuits at the same time, but also in the case of a complete failure of one power supply system.

Fig. 70 Block diagram of hydraulic power distribution to component circuits Issue 1 Mod 11.11

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Figure 70 shows a block diagram of a typical hydraulic power distribution system and the supply to the various component circuits. The example shown, has two main independent systems, System “A” and “B” with a “Standby” System to cater for redundancy. The complete hydraulic system consists of: a.A power circuit, b.Various component circuits. c. An emergency circuit in the event of hydraulic power failure. 11.33

POWER CIRCUITS

The power circuit supplies fluid to the component circuits and accommodates the fluid returned from these circuits. The system varies with the type of aircraft and may contain more than one Engine-driven pump (EDP). The circuit may be selfidling or non-self-idling. The self-idling circuit is designed to “idle” when the working pressure has been achieved, while in the non-self-idling circuit the pump supplies fluid continuously to the circuit and necessitates the installation of an automatic unloading (cut-out) valve.

Fig. 71 Power Circuit “Constant Pressure/Variable displacement”, (Self idling) Pump System Issue 1 Mod 11.11

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Power Circuit Constant Volume/Fixed displacement”, (Non self-idling) Pump System Figure 72 Figures 71 & 72 show the differences in the hydraulic power system design when a Constant Pressure/Variable displacement Pump is installed, (Fig. 43 and 44.) Compared to a system that has a Constant Volume/Fixed displacement Pump installed. (Figs. 34, 35, 36 & 37)

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COMPONENT CIRCUITS

Each component (system), has its own hydraulic circuit within the hydraulic system. These circuits are usually connected to a common pressure line and a common return line of a power circuit. Fluid expelled from each component circuit is conveyed back to the reservoir by the return line. 11.34.1

FLAPS

The flap circuit illustrated in Fig 73 consists of port and starboard flap jacks, synchronising jacks, and various components interconnected by pipelines. As with the landing gear circuit, fluid is supplied by the main system power circuit, to a control (selector) valve, via a Shut-off valve, which directs the fluid to the desired end of the jacks, at the same time connecting the other end of the jacks to the reservoir. A non-return valve before the control valve prevents operation of other services, such as alighting gear, from interfering with the flaps.

Typical Wing Flaps Hydraulic Circuit Figure 73 Issue 1 Mod 11.11

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THROTTLING VALVE During flight, it is essential that the Wing flaps are lowered and raised slowly to prevent sudden change in the trim of the aircraft, therefore, a throttling valve is provided in the circuit. This valve reduces the rate of flow of fluid to and from the flap actuators and is normally situated in the DOWN line.

Typical Balanced spring Throttling Valve Figure 74 This valve, which is a form of two-way restrictor valve, maintains the flow of fluid to and from a service, but at a constant rate. It automatically sets the flow rate in proportion to the supply pressure and is used to slow down the operation of the flaps.

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With fluid pressure normal, fluid flows through the piston ports, but an increase in the Inlet pressure or an increase in flow to the valve would increase the pressure on one side of the piston which in turn will move the piston in the direction of the flow. The leading needle valve approaches and restricts the outlet port, thereby restricting the flow of fluid out of the valve. The design of the valve and strength of the springs ensures that the needles will not seat and completely shut off the flow. SYNCHRONISATION Owing to slight variations in jack volume or piston friction, or to unequal air loading such as will occur when landing the aircraft in a crosswind, the rate at which the port and starboard flaps move may differ. To minimise this possibility the movement of the flaps is synchronised. The methods of synchronisation vary and may consist of a single jack, mechanical linkage (cable), hydraulic synchronisation valves or jacks. The method of synchronisation varies with the type of aircraft, but the method illustrated in Fig. 69 employs two additional jacks interconnected by transfer pipes, which are not connected to the power circuit. The operation is as follows. When the flaps travel in alignment, the fluid in the synchronising jacks is merely transferred from one side of the piston in one jack, to the opposite side of the piston in the other jack. There is normally a tendency for the travel of one flap to be slower than the travel of the other flap. When this occurs, the synchronising jack on the slow flap will provide an assisting force to the slow flap. Example: Assume that Flaps Down has been selected and that the Port flap tends to move down faster than the starboard flap. The piston in the port synchronising jack would expel more fluid from D into A, therefore, pressure is generated in A which, acting on the piston of the starboard synchronising jack, produces an assisting force helping to keep the starboard flap in alignment with the port flap. The fluid expelled from B is accommodated in C. (Fig. 74) NOTE : Only the basic flap synchronising circuit is described and illustrated. In the aircraft, provision must be made for priming, thermal expansion and contraction; in some aircraft, the flap synchronising circuit is connected to the power circuit. A hydraulic lock will be formed between a non-return valve (or the Control/Selector valve when selected in the “OFF” position) and the jacks. A hydraulic lock or a closed circuit can be designed into a system as a Landing gear protection device. NOTE : In some aircraft the synchronising circuit is independent of the main hydraulic circuit, the fluid in the circuit having no pressure except that built-up by transfer.

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THERMAL RELIEF VALVE When flying the aircraft to high altitude, the low temperature of the atmosphere will cause the fluid in the pipelines to contract. From high to low altitude, due to the increase in temperature of the atmosphere, the fluid will expand (thermal expansion). Whereas contraction of the fluid will be compensated by the pump supplying more fluid to the pipeline, thermal expansion of the fluid, especially in a closed circuit such as previously described may burst the pipelines. To prevent this, two thermal relief valves are fitted in the circuit, one in the up line and the other in the down line. The valves relieve fluid pressure from the pressure line to the return line. In the flap circuit illustrated in Fig. 74, the pressure relief valve will act as a thermal relief valve. 11.34.2

LANDING GEAR

The Landing gear circuit illustrated consists of two main undercarriage jacks, a nose wheel jack, fairing door jacks and various components interconnected by pipelines. Normally, fluid is supplied by the power circuit to the control valve, which may be manually or electrically operated and directs the fluid to the desired end of the jacks, at the same time connecting the other end of the jacks to the reservoir. A non-return valve positioned before the control valve provides a hydraulic lock in both UP and DOWN position, which ensures that the alighting gear remains in its selected position, when any other service is operated.

11.35

HYDRAULIC POWER - INDICATION AND WARNING SYSTEMS

Information regarding the condition of the hydraulic system must be relayed to the flight deck, for normal and abnormal situations. Generally the information comprises, actual hydraulic pressure, temperature and quantity (normal indications) and warnings of low pressure and low quantity or high oil temperature (abnormal indications). Since most hydraulic reservoirs are pressurised by engine bleed air, to prevent the oil from foaming due to low ambient pressure at altitude, a warning of low air pressure is also included. From a servicing viewpoint, direct reading quantity gauges are often to be found on the side of the hydraulic reservoirs and to show the gas pressure in hydraulic accumulators.

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System Control and Monitoring Figure 75 11.36

HYDRAULIC PRESSURE

Since the hydraulic bay is often some distance from the flight deck and to avoid the inherent risk of hydraulic oil leaking onto electronic equipment, no oil pipes run directly to the flight deck instruments. On all modern aircraft, electro-hydraulic transducers fitted in the hydraulic bay, relay pressure information for each system to the flight deck. In this way all hydraulic lines stay out of the pressure cell. Instead, electrical cables are routed from each transducer to some form of ratiometer or solid state liquid crystal display (LCD), calibrated to read hydraulic pressure.

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Pressure Indication Circuit Figure 76 Additionally, a pressure switch set to minimum pump output pressure, is routed to the aircraft alerting and warning system, to some form of visual warning (warning lamp/ flashing glareshield lights) and an aural warning (chimes).

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Low Pressure Warning Figure 77

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HYDRAULIC QUANTITY

The quantity of oil in the reservoir of each system will be relayed to a quantity gauge on the flight deck by means of a float switch in the tank. Alternatively, a capacitance type detector can be employed. Both types cause a voltage change at the gauge corresponding to a change in oil level in the tank. The gauges can be calibrated to show the actual quantity (litres) or displayed as a percentage of full.

Hydraulic Fluid Quantity Figure 78

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Additionally, a low level switch can be fitted in the reservoir or within the gauge/indicator which will trigger visual and aural warnings, when the level reaches a pre-calibrated minimum value.

Standby Hydraulic System – Low quality Light Figure 79

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Hydraulic Temperature The system may be fitted with a temperature transducer relaying system temperature to a gauge, but normally this is unnecessary. Usually, all that is required is a temperature switch, usually in the return line as it enters the reservoir, to trigger the visual/aural warning if the temperature should exceed a pre-determined maximum value. Such temperature sensors are often associated with electric motor driven hydraulic pumps and may monitor the temperature of the motor windings as well as actual oil temperature.

System Overheat Indication Figure 80

Reservoir Low Air Pressure A low pressure switch fitted in the Bleed Air line downstream of the pressure regulator just before it enters the reservoir, will trigger the visual/aural warning system if the pressure drops below a pre-determined minimum value. Accumulator Gas Pressure Gauges are fitted to accumulators, to indicate the pre-charge gas (nitrogen) pressure when all hydraulic pressure has been dissipated. These gauges are usually direct reading and will show system pressure when the hydraulic pumps are running. Issue 1 Mod 11.11

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INTERFACES WITH OTHER SYSTEMS

Hydraulic Power is used for the operation of a large number of aircraft systems. These include: 

Powered Flying Controls – Primary Controls



Leading edge and Trailing Edge Flaps



Spoilers



Speed Brakes and Air Brakes



Wheel Brakes and Anti-skid



Nosewheel Steering



Landing Gear Retraction & Lowering



Windscreen Wipers

Hydraulic pumps can be driven: Mechanically from the main engine accessory gearbox or from the APU Electrically from the main electrical buses By means of a Ram Air Turbine deployed into the airflow (emergency) Air-driven from the aircraft bleed air system (emergency) Hydraulic Accumulators can be used for parking brake pressure storage.

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MODULE 11.12 ICE AND RAIN PROTECTION

Contents 12 ICE FORMATION, ........................................................................ 12-3 12.1 12.2 12.3

12.4

12.5 12.6

12.7 12.8 12.9

12.10 12.11

12.12

12.13

12.14

12.15 12.16

CLASSIFICATION AND DETECTION INTRODUCTION ...................... 12-3 FACTORS AFFECTING ICE FORMATION ................................. 12-3 TYPES OF ICE FORMATION ..................................................... 12-3 12.3.1 Hoar Frost ..................................................................... 12-3 12.3.2 Rime Ice........................................................................ 12-4 12.3.3 Glaze Ice ....................................................................... 12-4 12.3.4 Pack Snow .................................................................... 12-5 12.3.5 Hail ............................................................................... 12-5 AREAS TO BE PROTECTED...................................................... 12-5 12.4.1 Effects On Aircraft ......................................................... 12-6 12.4.2 Effects of Icing on The Ground ...................................... 12-7 ICE DETECTION ........................................................................ 12-7 METHODS OF ICE DETECTION ................................................ 12-7 12.6.1 Visual (Hot Rod) Ice Detector)....................................... 12-8 Pressure Operated Ice Detector Heads ...................................... 12-9 Serrated Rotor Ice Detector Head ............................................... 12-10 12.6.4 Vibrating Rod Ice Detector ............................................ 12-11 ICE FORMATION SPOT LIGHT ......................................................... 12-12 ANTI-ICING AND DE-ICING SYSTEMS .............................................. 12-12 INTRODUCTION ............................................................................. 12-12 12.9.1 De-icing ......................................................................... 12-12 12.9.2 Anti-icing System .......................................................... 12-13 DE-ICING/ANTI-ICING SYSTEMS - GENERAL ................................. 12-13 FLUID SYSTEMS ........................................................................... 12-13 12.11.1 Windscreen Protection .................................................. 12-13 12.11.2 Aerofoil Systems ........................................................... 12-16 12.11.3 Propeller Systems ......................................................... 12-18 PNEUMATIC SYSTEMS ............................................................. 12-19 12.12.1 Air Supplies ................................................................... 12-20 12.12.2 Distribution .................................................................... 12-20 12.12.3 Controls and Indication ................................................. 12-20 12.12.4 Operation ...................................................................... 12-21 THERMAL (HOT AIR) SYSTEM .................................................. 12-22 12.13.1 Exhaust Gas Heating System ....................................... 12-23 12.13.2 Hot Air Bleed System .................................................... 12-25 ELECTRICAL ICE PROTECTION SYSTEN ................................ 12-27 12.14.1 Heater Mat .................................................................... 12-27 12.14.2 Spray Mat ..................................................................... 12-28 Windscreen Anti-icing ................................................................. 12-31 WINDSCREEN CABIN W INDOW DE-MISTING SYSTEMS ..................... 12-33 RAIN REPELLANT AND RAIN REMOVAL ........................................... 12-35

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12.17 WINDSCREEN CLEARING SYSTEMS ....................................... 12-35 12.18 WINDSCREEN WIPER SYSTEMS ..................................................... 12-36 12.18.1 Electrical System........................................................... 12-36 12.18.2 Electro-Hydraulic System .............................................. 12-38 12.18.3 Windscreen Wiper Servicing ......................................... 12-39 12.19 PNEUMATIC RAIN REMOVAL SYSTEMS.................................. 12-41 12.20 WINDSCREEN WASHING SYSTEM .......................................... 12-41 12.21 RAIN REPELLANT ..................................................................... 12-43 12.22 DRAIN MAST HEATING .................................................................. 12-46 12.23 WATER SUPPLY AND DRAIN LINES ................................................. 12-46 12.24 DRAIN MASTS .............................................................................. 12-46

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12 ICE FORMATION, 12.1 CLASSIFICATION AND DETECTION INTRODUCTION The operation of aircraft in the present day necessitates flying in all weather conditions and it is essential that the aircraft is protected against the build up of ice which may affect the safety and performance of the aircraft. Aircraft designed for public transport and some military aircraft must be provided with certain detection and protection equipment for flights in which there is a probability of encountering icing (or rain) conditions. In addition to the requirements outlined above, certain basic standards have to be met by all aircraft whether or not they are required to be protected by the requirements. These basic requirements are intended to provide a reasonable protection if the aircraft is flown intentionally for short periods in icing conditions. The requirements cover such considerations as the stability and control balance characteristics, jamming of controls and the ability of the engine to continue to function. 12.2 FACTORS AFFECTING ICE FORMATION Ice formation on aircraft in flight is the same as that on the ground; it can be classified under four main headings, i.e. Hoar Frost, Rime, Glaze Ice and Pack Snow. Dependent on the circumstances, variations of these forms of icing can occur and two different types of icing may appear simultaneously on parts of the aircraft. Ice in the atmosphere is caused by coldness acting on moisture in the air. Water occurs in the atmosphere in three forms, i.e. invisible vapour, liquid water and ice. The smallest drops of liquid water constitute clouds and fog, the largest drops occur only in rain and in between these are the drops making drizzle. Icing consists of crystals, their size and density being dependent on the temperature and the type of water in the atmosphere from which they form. Snowflakes are produced when a number of these crystals stick together or, in very cold regions, by small individual crystals. 12.3 TYPES OF ICE FORMATION 12.3.1 HOAR FROST

Hoar frost occurs on a surface which is at a temperature below the frost point of the adjacent air and of course, below freezing point. It is formed in clear air when water vapour condenses on the cold airframe surface and is converted directly to ice and builds up into a white semi-crystalline coating; normally hoar frost is feathery.

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When hoar frost occurs on aircraft on the ground, the weight of the deposit is unlikely to be serious, but the deposit, if not removed from the airframe, may interfere with the airflow and attainment of flying speed during take-off, the windscreen may be obscured and the free working control surfaces may be affected. Hoar frost on aircraft in flight usually commences with a thin layer of glaze ice on the leading edge, followed by the formation of frost which gradually spreads over the whole surface. Again the effects are not usually serious, though some change in the landing characteristics of the aircraft can be expected. 12.3.2 RIME ICE

This ice formation, which is less dense than glaze ice, is an opaque, rough deposit. At ground level it forms in freezing fog and consists of a deposit of ice on the windward side of exposed objects. Rime is light and porous and results from the small water drops freezing as individual particles, with little or no spreading, a large amount of air is trapped between the particles. Aircraft in flight may experience rime icing when flying through a cloud of small water drops with the air temperature and the temperature of the airframe below freezing point. The icing builds up on the leading edge, but does not extend far back along the chord. Ice of this type usually has no great weight, but the danger of rime is that it will interfere with the airflow over the wings. If the super-cooled droplets are small enough and the temperature is low, each droplet freezes instantly on impact as an individual particle and being a nonadhesive dry powder in the slipstream the accumulation on the aircraft is not serious. This is called "opaque rime". 12.3.3 GLAZE ICE

Glaze ice is the glassy deposit that forms over the village pond in the depth of winter. On aircraft in flight, glaze ice forms when the aircraft encounters large water drops in clouds or in freezing rain (or super-cooled rain) with the air temperature and the temperature of the airframe below freezing point. It consists of a transparent or opaque coating of ice with a glassy surface and results from the liquid water flowing over the airframe before freezing. Glaze ice may be mixed with sleet or snow. IT WILL FORM IN GREATEST THICKNESS ON THE LEADING EDGES OF AEROFOILS AND IN REDUCED THICKNESS AS FAR AFT AS ONE HALF OF THE CHORD. Ice formed in this way is dense, tough and sticks closely to the surface of the aircraft, it cannot easily be shaken off and if it breaks off at all, it comes away in lumps of an appreciable and sometimes dangerous size.

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The main danger of glaze ice is still aerodynamic but also the weight of the ice produces unequal loading and propeller blade vibrations. Glaze ice is the MOST SEVERE and most dangerous form of ice formation on aircraft because of its high RATE OF CATCH. Super-cooled rain is rare in the British Isles but is more common on the Continent and East coast of North America. 12.3.4 PACK SNOW

Normally, snow falling on an aircraft in flight does not settle, but if the temperature of the airframe is below freezing point, glaze ice may form from the moisture in the snow. The icing of the aircraft in such conditions, however, is primarily due to water drops, though snow may subsequently be embedded in the ice so formed. 12.3.5 HAIL

Hail is formed when water droplets, falling as rain, pass through icing levels and freeze. Air currents in some storm clouds (Cumulo-nimbus) may carry the hail vertically through the cloud a number of times, increasing the size of the hailstone at each pass until it is heavy enough to break out of the base of the cloud and fall towards earth. Aircraft encountering this type of ice formation may suffer severe damage in the form of dented skin, cracked windscreens, blocked intakes and serious damage to gas turbine engines. 12.4 AREAS TO BE PROTECTED The following areas are critical areas on the aircraft where ice forms and where protection is essential. a. all aerofoil leading edges b. engine air intakes (including carburettor intakes) c. windscreens d. propellers e. pitot static pressure heads

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ICE AND RAIN

engineering

PROTECTION

Icing – Areas to be Protected Figure 1 12.4.1 EFFECTS ON AIRCRAFT

The build up of ice on the aircraft is known as 'ice accretion' and, from the foregoing, it is evident that if ice continues to be deposited on the aircraft one, or more, of the following effects may occur. a. Decrease in Lift This may occur due to changes in wing section resulting in loss of streamlined flow around the leading edge and top surfaces. b. Increase in Drag Drag will increase due to the rough surface, especially if the formation is rime. This condition results in greatly increased surface friction. c. Increased Weight and Wing Loading The weight of the ice may prevent the aircraft from maintaining height. d. Decrease in Thrust With turbo-prop and piston engines, the efficiency of the propeller will decrease due to alteration of the blade profile and increased blade thickness. Vibration may also occur due to uneven distribution of ice along the blades. Gas Turbine engines may also be affected by ice on the engine intake, causing disturbance of the airflow to the compressor. Furthermore, ice breaking away from the intake, may be ingested by the engine causing severe damage to the compressor blades and other regions within the engine. Mod 11.12 issue 1

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e. Inaccuracy of Pitot Static Instruments Ice on the pitot static pressure head causes blockage in the sensing lines and produces false readings on the instruments. f.

Loss of Inherent Stability

This may occur due to displacement of the centre of gravity caused by the weight of the ice. g. Radio antennae Reduced efficiency h. Loss of Control Loss of control may occur due to ice preventing movement of control surfaces. (This is not usually a problem in flight but may occur on the ground). 12.4.2 EFFECTS OF ICING ON THE GROUND

The effects of ice accretion on the ground are similar to those occurring in flight but the following additional effects may be caused. a. Restriction of the controls may occur if ice is not removed from hinges and gaps in the controls. b. The take off run may be increased because of the increase in weight and drag. c. The rate of climb may be reduced because the weight and drag are increased. 12.5 ICE DETECTION The ANO Schedule 4 states that: In the case of an aircraft of MTWA exceeding 5700 kg (12500 lb), means of observing the existence and build up of ice on the aircraft must be provided. The equipment will be carried on flights when the weather reports or forecasts available at the aerodrome at the time of departure indicate that conditions favouring ice formation are likely to be met. 12.6 METHODS OF ICE DETECTION Ice detection systems use one of the following methods of detecting and assessing the formation of ice.

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12.6.1 VISUAL (HOT ROD) ICE DETECTOR)

This consists of an aluminium alloy oblong base (called the plinth) on which ismounted a steel tube detector mast of aerofoil section, angled back to approximately 300 from the vertical, mounted on the side of the fuselage, so that it can be seen from the flight compartment windows. The mast houses a heating element, and in the plinth there is a built-in floodlight.

Hot Rod Ice Detector Figure 2 The heating element is normally off and when icing conditions are met ice accretes on the leading edge of the detector mat. This can then be observed by the flight crew. During night operations the built-in floodlight may be switched on to illuminate the mast. By manual selection of a switch to the heating element the formed ice is dispersed for further observance.

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12.6.2 PRESSURE OPERATED ICE DETECTOR HEADS

MAST

OUTLET HOLES

INLET HOLES

ELECTRICAL CONNECTORS

Pressure Operated Ice Detector Figure 3 These consist of a short stainless steel or chromium plated brass tube, which is closed at its outer end and mounted so that it projects vertically from a portion of the aircraft known to be susceptible to icing. Four small holes are drilled in the leading edge of this tube and in the trailing edge are two holes of less total area than those of the leading edge. A heater element is fitted to allow the detector head to be cleared of ice. In some units of this type a further restriction to the air flow is provided by means of a baffle mounted through the centre of the tube. Each system comprises an ice detector head, a detector relay and a warning lamp. When in normal flight, pressure is built up inside the tube by the airstream, this pressure is then communicated by tubing, to the capsule of an electro-pneumatic relay tending to expand it and separate a pair of electrical contacts. When icing conditions are met, ice will form on the leading edge and close off the holes. As the holes in the trailing edge will not be covered by ice the air-stream will now tend to exhaust the system, collapsing the relay capsule and so closing the relay contacts. Generally these contacts operate in conjunction with a thermal device, to illuminate a warning indicator in the flight compartment and to switch on the heater in the detector head; the latter clears the head of ice and is then switched off allowing continued detection of icing conditions. A heater energised by the detector relays, automatically clears the ice from the head, but a cam holds the lamp on for a further 4 minutes and the heater for a further 30 seconds. Mod 11.12 issue 1

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Should icing conditions persist and the detector heads again ice up, the cam is automatically re-set and the time cycle repeated.The pilot will switch on the deicing system when the warning lights indicate icing conditions. In some systems the warning phase is connected to automatically switch on the de-icing system. This cycling will continue until such time that the icing conditions no longer exist. 12.6.3 SERRATED ROTOR ICE DETECTOR HEAD

Serrated Rotor Ice Detector Figure 4 This consists of a serrated rotor, incorporating an integral drive shaft coupled to a small ac motor via a reduction gearbox, being rotated adjacent to a fixed knifeedge cutter. The motor casing is connected via a spring-tensioned toggle bar to a micro-switch assembly. The motor and gearbox assembly is mounted on a static spigot attached to the motor housing and, together with the micro-switch assembly, is enclosed by a cylindrical housing. The detector is mounted through the fuselage side so that the inner housing is subjected to the ambient conditions with the outer being sealed from the aircraft cabin pressure.The serrated rotor on the detector head is continuously driven by the electrical motor so that its periphery rotates within 0.050 mm (0.002 in) of the leading edge of the knife-edge cutter. The torque therefore required to drive the rotor under non-icing conditions will be slight, since bearing friction only has to be overcome. Under icing conditions, however, ice will accrete on the rotor until the gap between the rotor and knife-edge is filled, whereupon a cutting action by the knife edge will produce a substantial increase in the required torque causing the toggle bar to move against its spring mounting and so operate the microswitch, to initiate a warning signal. Once icing conditions cease, the knife edge cutter will no longer shave ice, torque loading will reduce and allow the motor to return to its normal position and the micro-switch will open-circuit the ice warning indicator. Mod 11.12 issue 1

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12.6.4 VIBRATING ROD ICE DETECTOR

This ice detector senses the presence of icing conditions and provides an indication in the flight compartment that such conditions exist. The system consists of' a solid state ice detector and advisory warning light. The ice detector is attached to the fuselage with its probe protruding through the skin. The ice detector probe (exposed to the airstream) is an ice-sensing element that ultrasonically vibrates in an axial mode of its own resonant frequency of approximately 40 kHz.

SENSOR UNIT

DETECTOR PROBE

VIBRATING ROD

Vibrating Rod Ice detector Figure 5 When ice forms on the sensing element, the probe frequency decreases. The ice detector circuit detects the change in probe frequency by comparing it with a reference oscillator. At a predetermined frequency change (proportional to ice build-up), the ice detector circuit is activated. Once activated, the ice warning light in the flight compartment is illuminated and a timer circuit is triggered. The operation of the time circuit switches a probe heater on for a set period of time to remove the ice warning indicator and returns the system to a detector mode, providing that icing conditions no longer exist. If, however, a further ice warning signal is received during the timer period, the timer will be re-triggered, the warning light will remain on and the heater will again be selected on. This cycle will be repeated for as long as the icing conditions prevail.

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12.7 ICE FORMATION SPOT LIGHT Many aircraft have two ice formation spot lights mounted one each side of the fuselage, in such a position as to light up the leading edges of the mainplanes, when required, to allow visual examination for ice formation. Note:

In some aircraft this may be the only method of ice detection.

Spotlight Ice detectors Figure 6 12.8 ANTI-ICING AND DE-ICING SYSTEMS 12.9 INTRODUCTION There are various methods of ice protection which can be fitted to an aircraft but they can be considered under one of two main categories, de-icing and anti-icing. 12.9.1 DE-ICING

In this method of ice protection, ice is allowed to form on the surfaces and is then removed by operating the particular system in the specified sequence. Mod 11.12 issue 1

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12.9.2 ANTI-ICING SYSTEM

Ice is prevented from forming by ensuring that the ice protection system is operating whenever icing conditions are encountered or forecast. 12.10

DE-ICING/ANTI-ICING SYSTEMS - GENERAL

There are four primary systems used for ice protection. These are: 1. Fluid 2. Pneumatic 3. Thermal 4. Electrical 12.11

FLUID SYSTEMS

These may be used either as an anti-icing or de-icing system. When used as an anti-icing system it works on the principle that the freezing point of water can be lowered if a fluid of low freezing point is applied to the areas to be protected before icing occurs. When used as a de-icing system the fluid is applied to the interface of the aircraft surface and the ice. The adhesion of the ice is broken and the ice is carried away by the airflow. The system is normally used on windscreens and aerofoils and has also been used successfully on propellers. It is not used on engine air intakes - which are usually anti-iced. 12.11.1

WINDSCREEN PROTECTION

The method employed in this system is to spray the windscreen panel with an ALCOHOL based fluid. The principal components of the system are: 

Fluid storage tank



Hand operated or electrically driven pump



Supply pipelines



Spray tubes

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The diagram illustrates a typical aircraft system in which the fluid is supplied to the spray tubes by two electrically driven pumps.

Typical Fluid De-icing System Figure 7 This design enables the system to be operated using either of the two pumps, or both pumps, according to the severity of the icing.

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The next diagram shows a hand pump installation on the HS 125 aircraft where it is used as an auxiliary system.

Windscreen Auxiliary De-icing System Figure 8

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ICE AND RAIN PROTECTION

AEROFOIL SYSTEMS

The fluids used for aerofoil ice protection are all GLYCOL based and have properties of low freezing point, non-corrosive, low toxicity and low volatility. They have a detrimental effect on some windscreen sealing compounds and cause crazing of perspex panels. The components in the system are the tank, pump, filter, pipelines, distributors, controls and indicators normally consisting of a switch, pump power failure warning light and tank contents indicator. When icing conditions are encountered, the system may be switched on automatically by the ice detector or manually by the pilot. Fluid is supplied to the pump by gravity feed from the tank and is then directed under pressure to the distributors on the aerofoil leading edges. After an initial 'flood' period, during which the pump runs continuously to prime the pipelines and wet the leading edge, the system is then controlled by a cyclic timer which turns the pump ON and OFF for predetermined periods. The leading edge distributors appears in one of two forms, i.e. strip and panel. Strip Distributor The distributor consists of a 'U' channel divided into two channels, called the primary and secondary channels, by a central web. The outer part of the channel is closed by a porous metal spreader through which the de-icing fluid seeps to wet the outer surface. The primary and secondary feed channels are interconnected by flow control tubes to ensure an even spread of fluid over the outer surface. The strips are let into the leading edge so that the porous element is flush with the surface of the leading edge curvature. This type of distributor is rarely used and would only be found on very old aircraft. Panel Distributors This type of distributor consists of a micro porous stainless steel outer panel, a micro-porous plastic sheet and metering tube. The fluid passes through the metering tube that calibrates the flow rate into a cavity between the plastic sheet and a back-plate. This cavity remains filled when the system is operating and the fluid seeps through the porous stainless steel outer panel. The airflow then directs the fluid over the aerofoil.

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PROTECTION

The outer panel is usually made of stainless steel mesh although a new technique of laser drilling of stainless steel sheet is appearing on some new aircraft. DISTRIBUTOR PANELS FILTER

VENT MAIN FEED PIPES

GALLEY PIPES

PUMP

TANK

DISTRIBUTOR PANELS

Fluid De-icing System with Distribution Panels Figure 9 Mod 11.12 issue 1

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When a system is to be out of service, or unused for an extended period of time, it should be functioned periodically to prevent the fluid from crystallising and causing blockage of the metering tubes, porous surfaces and pipelines. Distributors should be cleaned periodically by washing with a jet of water sprayed on to the distributor at an angle.

Section of a TKS Distribution Panel Figure 10 12.11.3

PROPELLER SYSTEMS

It is necessary to de-ice the propeller blade root and a section of the propeller blade to prevent the build up which could change the blade profile and upset the aerodynamic characteristics of the propeller. Uneven ice build up will also introduce imbalance of the propeller and cause vibration. The leading edge of the propeller blade is therefore de-iced and the ice is shed by centrifugal force. The blade root has a rubber cuff into which the de-icing fluid is fed by a pipeline from a slinger ring on the spinner back plate. From the cuff the fluid is spread along the leading edge of the blade by centrifugal force.

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Fluid is fed into the ‘slinger ring’ from a fixed pipe on the front of the engine.

Propeller ‘Slinger Ring’ De-Icing Figure 11 12.12

PNEUMATIC SYSTEMS

Pneumatic (or mechanical) systems are used for de-icing only, It is not possible to prevent ice formation and works on the principle of cyclic inflation and deflation of rubber tubes on aerofoil leading edges. The system is employed in certain types of piston engine and twin turbo-propeller aircraft. The number of components comprising a system and the method of applying the operating principle will vary but a typical arrangement is shown. The de-icer boots (or overshoes) consist of layers of natural rubber and rubberised fabric between which are disposed flat inflatable tubes closed at the ends. They are fitted in sections along the leading edges of wing, vertical stabilisers and horizontal stabilisers. The tubes may be laid spanwise, chordwise or a combination of each method. The tubes are made of rubberised fabric vulcanised inside the rubber layers and are connected to the air supply by short lengths of flexible hose secured by hose clips. Mod 11.12 issue 1

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Depending on the type specified, a boot may be attached to the leading edge either by screw fasteners or by cementing them directly to the leading edge skin. The external surfaces of the boots are coated with a film of conductive material to bleed off accumulations of static electricity.

Pneumatic De-Icing Boots Figure 12 12.12.1

AIR SUPPLIES

The tubes in the overshoes are inflated by air from the pressure side of an engine driver vacuum pump or, in some types of turbo-propeller aircraft, from a tapping on the engine compressor. At the end of the inflated stage of the operating sequence, and whenever the system is switched off, the boots are deflated by vacuum derived from the vacuum pump or from the venturi section of an ejector nozzle in systems using the engine compressor tapping. 12.12.2

DISTRIBUTION

The method of distributing air supplies to the boots depends on the system required for a particular type of aircraft. In general three methods are in use: 

shuttle valves controlled by a separate solenoid valve



individual solenoid valves direct air to each boot



motor driven valves

12.12.3

CONTROLS AND INDICATION

The controls and indication required for the operation of a system will depend on the type of aircraft and on the particular arrangement of the system. In a typical system a main ON-OFF switch, pressure and vacuum gauges or indicating lights form part of the controlling section. Pressure and vacuum is applied to the boots in an alternating, timed sequence and the methods adopted usually vary with the methods of air distribution. In most installations, however, timing control is affected by an electronic device. Mod 11.12 issue 1

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Pneumatic De-icing System Layout Figure 13 12.12.4

OPERATION

When the system is switched on, pressure is admitted to the boot sections to inflate groups of tubes in sequence. The inflator weakens the bond between ice and the boot surfaces and cracks the ice that is carried away by the airflow. At the end of the inflation stage of the operating sequence, the air in the tubes is vented to atmosphere through the distributor and the tubes are fully deflated by the vacuum source. The inflation and deflation cycle is repeated whilst the system is switched on. When the system is switched off, vacuum is supplied continually to all tubes of the overshoes to hold the tubes flat against the leading edges thus minimising aerodynamic drag.

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Pneumatic De-Icing Boots - Operation Figure 14 12.13

THERMAL (HOT AIR) SYSTEM

The thermal (hot air) system fitted to aerofoils for the purpose of preventing the formation of ice employs heated air ducted span-wise along the inside of the leading edge of the aerofoil and distributed between double thickness skins. Entry to the leading edge is made at the stagnation point where maximum temperature is required. The hot air then flows back chord-wise through a series of corrugations into the main aerofoil section to suitable exhaust points.

Thermal (Hot Air) de-Icing System Figure 15 In anti-icing systems a continuous supply of heated air is fed to the leading edges, but in de-icing systems it is usual to supply more intensely heated air for shorter periods on a cyclic basis. Hot gas may be derived from heat exchangers around exhausts, independent combustion heaters or direct tappings from turbine engine compressors. Mod 11.12 issue 1

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EXHAUST GAS HEATING SYSTEM

The following diagram illustrates the principle of a thermal system using exhaust gases to heat ambient air. Ambient air enters an intake formed on one side of the engine nacelle and is ducted to pass through tubes of a heat exchanger. The exhaust gases from the jet pipe are partially diverted by electrically actuated flaps to flow between the tubes of the heat exchanger before discharging to atmosphere. The heated air from the heat exchanger passes to a duct containing an electrically operated hot air valve before passing to the leading edges. In the event of failure of the gas flap in the open position, an emergency manual override facility is provided to close the hot air valve and open an actuator operated spill valve to direct the hot air overboard. The gas flap actuator and the hot air valve actuator are electrically interlocked in such a way that the hot air valve must be fully open before the gas flap opens. Conversely, the gas flap must be fully closed before the hot air valve closes. This arrangement, controlled by the limit switches in the actuators, prevents overheating of the heat exchanger. Temperature control is automatic with a standby 'manual' facility. A control unit, in conjunction with 'normal' control and 'overheat' thermistors, provides automatic control and overheat protection. An overheat control unit, in conjunction with an 'override' thermistor and flame-stat provides a final overheat protection system.

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Exhaust Gas Heating system Figure 16

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HOT AIR BLEED SYSTEM

In this system, air is bled from a late stage of the gas turbine engine compressor before being distributed to aerofoil leading edges in the same manner as the exhaust system. The system may be used for anti-icing or de-icing purposes on wing and tail leading edges. It may also be used for ice protection of engine intakes In principle, the system works by either maintaining the temperature of the skin above that at which ice occurs or by raising the skin temperature to melt the ice after it has formed. On aircraft with engines mounted on the rear fuselage, distribution of air along the wing leading edges may be graded to give a higher intensity of heating for the inboard section. This is to prevent the shedding of ice accretions into the engine intakes of a size that could result in hazards to the engine. The following diagram illustrates, in schematic form, a thermal system for a four engine aircraft. In operation, anti-icing shut off valves on each engine open to supply air to the leading edge ducting at temperatures of about 200ºC. Wing and fuselage cross over ducts ensure a supply to all surfaces in the event of an engine shut down in flight.

Hot Bleed Air Anti-icing System Figure17 Mod 11.12 issue 1

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On some installations, air temperature in the ducting may be controlled by mixing compressor bleed air with ram air admitted to the system by a cold air control valve.

Hot and Ram Air Mixing Figure 18 When initially switched on, hot air is fed undiluted into the cold leading edge ducting. Temperature sensors in the leading edge monitor the temperature rise and progressively open and close the cold air valve via an inching unit to control the skin temperature. In the event of failure of the: 

temperature sensor to control the temperature of the leading edge



cold air valve

or blockage of the ram air inlet, the overhead sensor will control the temperature by regulation of the hot air valve. Note: Temperature regulation may also be achieved by controlling the position of the hot air valve.

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ELECTRICAL ICE PROTECTION SYSTEN

Electrical heater elements are attached to the outer surface of the area to be protected. There are two methods; these being the heater mat and spray mat. 12.14.1

HEATER MAT

This type of element consists of two thin layers of rubber or PTFE sandwiching a heater element. Each mat is moulded to fit snugly over the section to be protected. Heater elements differ in design, construction and materials according to their purpose and environment. The latest mats have elements made from a range of alloys woven in continuous filament glass yarn. The diagram below shows the application of a heater element to the air intake of a turbo-prop engine.

Electrical Anti-Icing Heater Mat Figure 19

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SPRAY MAT

This type of element is so called because it is sprayed directly on to the surface to be protected. The technique was developed by the Napier Company to provide a lightweight system for use on aerofoils and is ideally suited for application to compound curves. A base insulator is brushed directly on to the airframe and is composed basically of synthetic resin. The insulator is normally about 0.03 inches thick although in some cases this may vary. The heater element, made of either aluminium or Kumanol (copper manganese alloy) is sprayed on to the base insulation using a flame spraying technique. The insulation is of the same material as the base insulation and about 0.01 inches thick. Finally, a protective coating is used where the heater requires extra protection from mechanical damage, eg on leading edges. This protective coating known as 'stoneguard' consists of stainless alloy particles bonded with synthetic resin.

Napier Type Anti-Icing Spraymat Figure 20 Mod 11.12 issue 1

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The system layout shows the distribution and heating elements on the leading edges of an aircraft tailplane and fin.

Distribution of Heating Elements Figure 21 Some of the elements are supplied continuously with electrical power (anti-icing) whilst others are supplied intermittently on a cyclic basis (de-icing). Areas provided with continuous anti-icing heating are situated immediately in front of areas on which limited ice formation is tolerable but which require de-icing by the cyclic application of heat. Heating of these areas is rapid in order to break adhesion as quickly as possible, allowing the detached ice to be blown away by the airflow. To ensure a clean breakaway of the ice, the cyclically heated areas are separated by continuously heated 'breaker' strips. A system requiring different intensifies of anti-icing and cyclic de-icing would require one or more cyclic switches, temperature sensing elements and temperature control units. In general, control methods may be classified as antiicing and de-icing.

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Anti-icing Anti-iced areas have their heat supplied continuously, the heating intensity being graded such that under operating conditions no ice formation occurs. The heat is regulated by means of either a sensing element embedded in the mat and an associated thermal controller or a surface mounted thermostatic switch which is pre-set to give cut-in and cut-out temperature levels. Cyclic De-icing Cyclic de-icing areas are usually arranged in groups being connected to a cyclic switch. The detailed design of the cycling switch depends upon the loading and type of power supply, e.g. dc or 3-phase ac. Its operation is controlled either by timed impulses from a pulse generator or by an electronic device built into the switch. The timed impulses are set to the appropriate rate for the range of ambient temperatures likely to be encountered. At a relatively high ambient temperature the atmospheric water content, and consequently the rate of icing, is likely to be high but only a comparatively short heating period will be required to shed the ice. At very low temperatures the atmospheric water content and rate of icing are lower and longer heating periods are required. The ratio of time ON to time OFF, however, remains unchanged. The typical ratio is 1:10. Setting of the pulse generator may be manual, as estimated from indications of ambient air temperature, or by an automatic control system in which the ON:OFF periods are varied by signals derived from an ambient air temperature probe, working in conjunction with either an ice detector or a rate of icing indicator. The source of power may be dc, single phase ac or 3-phase ac. In a 3-phase system the heated areas are arranged so as to obtain balanced loading of phases for both anti-icing or de-icing circuits, if possible. De-icing heaters are connected in such a manner that, as far as practicable, current requirements are constant. To achieve this the OFF period for certain areas is made to coincide with the ON period for others.

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WINDSCREEN ANTI-ICING

The windscreens and other critical windows in the cockpit (e.g. direct vision windows, sliding side windows) of high performance pressurised aircraft are complicated and expensive items of the airframe structure as they are designed to withstand varying air pressure loads, possible shock loads due to impact of birds and hailstorms, and thermal stresses due to ambient temperature changes. In all cases, a laminated form of construction is used, similar to that shown

Typical Laminated Glass Windscreen Figure 22 Laminated glass panels were conceived in order to impart shatter proof characteristics to the glass. Such panels are produced by interposing sheets of clear vinyl plastic (polyvinyl Butyral) between layers of preformed and pretempered glass plies. The vinyl and glass plies are then bonded by the application of pressure and heat. Since the desired bird-proof characteristics of a windscreen depend to a large degree, on the plasticity of the vinyl, it therefore follows that it also depends upon its temperature. The optimum temperature range for maximum energy absorption by the vinyl is between 27ºC and 49ºC and the electrically heated windscreen panel assemblies are normally maintained within these limits. Below this range the bird-proof characteristics decline rapidly and depending upon the actual configuration, a panel's impact resistance can be reduced by 30% to 50% when still at quite a moderate temperature of 16ºC. Electric heating of a windscreen therefore is an important factor in maintaining the optimum bird-proof characteristics. Mod 11.12 issue 1

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The heating element is an extremely thin transparent conductive coating which is 'floated' on to the inside surface of the outer glass ply; this being normally thinner in section allows a more rapid heat conduction. The coating may be a tin oxide or a gold film depending on a particular manufacturer's design. The conductive coating is heated by alternating current supplied to busbars at the edges of the windscreen panel. The power required for heating varies according to the size of the panel and the heat required to suit the operating conditions.

Windscreen Temperature Control Figure 23 The circuit of a typical windscreen de-icing system embodies a controlling device, the function of which is to maintain a constant temperature at the windscreen and also to prevent over-heating of the vinyl inter-layer(s). The controlling device is connected to temperature-sensing elements embedded in the windscreen. There are two methods of temperature sensing commonly in use. One of these utilises a grid in which the resistance of the grid varies directly and linearly with temperature. The other uses a thermistor, in which the resistance of the thermistor varies inversely and exponentially with temperature. The number of sensing elements employed depends on the system and circuit design requirements. A system of warning lights and/or indicators also forms part of the control circuit and provides visual indications of circuit operating conditions, e.g. 'normal', 'off' or 'overheat'.

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When the electrical power is applied, the conductive coating heats the glass. When it attains a temperature predetermined for normal operation the change in resistance of the appropriate sensing element causes the controlling device to isolate the heating power supply. When the glass has cooled through a certain range of temperature, power is again applied and the cycle is repeated. In the event of a failure of the controller, the glass temperature will rise until the setting of the overheat system sensing element is attained. At this setting an overheat control circuit cuts off the heating power supply and illuminates a warning light. The power is restored again and the warning light extinguished when the glass has cooled through a specific temperature range. 12.15 WINDSCREEN CABIN WINDOW DE-MISTING SYSTEMS

Glass is a very poor conductor of heat and at altitude the low atmospheric temperature will maintain the inside of the windscreens and cabin windows at low temperature resulting in condensation on the inner surface and obscured vision. Windscreens are normally kept mist free by blowing hot air, from the air conditioning system, across the inner surface of the glass. In addition, demisting of some windscreens and, usually, all cabin windows is achieved by using windows of "dry air sandwich" construction. This is rather like double-glazing with outer and inner layers of glass sandwiching a layer of dry air between them. The outer layer of glass is of thick laminate construction (glass and vinyl) to give the necessary impact and shatterproof qualities. The inner layer of glass is much thinner allowing it to be warmed by the cabin air temperature, thus preventing condensation. The air sandwich is kept dry to prevent internal condensation of the outer glass, by one of two methods: During manufacture the two layers of glass are hermetically sealed with dry air between them. The space between glass layers is vented to the cabin to allow the pressure in the air space to equalise with cabin pressure. Venting takes place through a desiccant unit that absorbs moisture from the air during the venting process to maintain the dry air sandwich.

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On some larger aircraft the fixed cabin windows are interconnected to a common desiccant unit whilst escape windows have their own integral unit. The diagram shows typical fixed window and escape hatch desiccant systems. The desiccant used is Silica Gel crystals which are blue in colour but gradually change to pink or white as they absorb moisture. Frequent checks must be made on the state of the desiccant which must be replaced when it begins to turn pink. Failure to take this action may result in condensation within the dry air sandwich which may involve lengthy rectification to dry out the sandwich or may require the windscreen/window to be replaced.

Cabin Window Desiccant System Figure 24

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12.16

RAIN REPELLANT AND RAIN REMOVAL

12.17

WINDSCREEN CLEARING SYSTEMS

Vision through windscreens may become obscured by factors other than ice and misting. For example, rain, dust, dirt and flies can impair vision to an extent where methods of clearing the screens must be provided to enable safe ground manoeuvring, take off and landing. Windscreen clearing systems may be considered under the following headings: a. Rain clearing systems which can be further broken down into a.

windscreen wipers

b.

pneumatic rain removal

c.

rain repellent

d.

windscreen washing

Windscreen washing systems.

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12.18 WINDSCREEN WIPER SYSTEMS 12.18.1

ELECTRICAL SYSTEM

In this type of system the wiper blades are driven by an electric motor(s) taking their power from the aircraft electrical system. Sometimes the pilot's and copilot's wipers are operated by separate motors to ensure that clear vision is maintained through one of the screens in case one system should fail. The following diagram shows a typical electrical wiper and installation. An electrically operated wiper is installed on each windscreen panel. Each wiper is driven by a motor-converter assembly that converts the rotary motion of the motor to reciprocating motion to operate the wiper arm. A shaft protruding from the assembly provides an attachment for the wiper arms.

Electric Windshield Wiper System Figure 25 Mod 11.12 issue 1

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The wiper is controlled by setting the wiper control switch to the desired wiper speed. When the "high" position is selected, relays 1 and 2 are energised. With both relays energised, fields 1 and 2 are energised in parallel. The circuit is completed and the motors operate at an approximate speed of 250 strokes/minute. When the "low" position is selected, relay 1 is energised. This causes fields 1 and 2 to be energised in series. The motor then operates at approximately 160 strokes/minute. Setting the switch to the OFF position allows the relay contacts to return to their normal positions. However, the wiper motor will continue to run until the wiper arm reaches the "park" position. When both relays are open and the park switch is closed, the excitation of the motor is reversed. This causes the motor to move off the lower edge of the windscreen, opening the cam operated park switch. This de-energises the motor and releases the brake solenoid applying the brake. This ensures that the motor will not coast and re-close the park switch.

Windshield Wiper Circuit Diagram Figure 26

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The path swept by the wiper blade may clear an arc as shown in the diagram on the left, or in a parallel motion as shown on the right. The parallel motion is preferred as it provides a greater swept surface, but the operating mechanism is more complex.

Windshield Wiper Swept Areas Figure 27 12.18.2

ELECTRO-HYDRAULIC SYSTEM

Older aircraft employed hydraulic motors instead of electric motors to drive the wiper blades. A typical example is shown in the figure below. It consists of two independently operated motors powered from each hydraulic system with control valves operated from a selector on the flight deck

Figure 28 Mod 11.12 issue 1

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WINDSCREEN WIPER SERVICING

Servicing of the windscreen wiper systems consists of inspection, operational checks, adjustments and fault finding. Inspection a. Examine the system for cleanliness, security, damage, connections and locking b. Examine blades for security, damage and contamination. Blades should be replaced at regular intervals. c. Check level of fluid in pump reservoir (electro-pneumatic system) d. Examine hydraulic pipes for leakage and electrical cables for deterioration and chafing

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Operational Check Before carrying out an operational check, the following precautions must be taken: a. Ensure that the windscreen is free of foreign matter b. Ensure that the blade is secure and undamaged During the check ensure that the windscreen is kept wet with water. NEVER operate the windscreen wipers on a dry screen. It may cause scratches. Adjustments The following adjustments may be made: a. Blade tension should be adjusted to the value stated in the Maintenance Manual. This is carried out by attaching a spring balance to the wiper arm at its point of attachment to the wiper blade and lifting at an angle of 90º. If the tension is not within the required limits, the spring may be adjusted by the appropriate pressure adjusting screw. b. Blade angle should be adjusted to ensure that the blade does not strike the windscreen frame. This would cause rapid blade damage. This may involve re-positioning the operating arm on the drive spindle. Where a parallel motion bar is used, the length of the tie rod may be altered to vary the angle of sweep. c. Proper parking of the wipers are essential to ensure that they do not obscure vision. If the wipers do not park as they should, they should be adjusted by the method laid down in the Maintenance Manual. Trouble shooting may be carried out using charts in the Maintenance Manual (Chapter 30-42-0 in the ATA100 Scheme).

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PNEUMATIC RAIN REMOVAL SYSTEMS

Windscreen wipers suffer from two basic problems. One is that at speed the aerodynamic forces tend to reduce the blade pressure on the screen and cause ineffective wiping. The other problem is to achieve blade oscillation rates that are high enough to clear the screen during heavy rain.

Pneumatic Rain Removal System Figure 29 Pneumatic rain clearance systems overcome these problems by using high pressure bleed air from the gas turbine engine and blowing it over the face of the windscreen from ducts mounted at the base of the screen. The air blast forms a barrier that prevents the rain spots from striking the screen. 12.20

WINDSCREEN WASHING SYSTEM

A windscreen washing system allows a spray of fluid (usually de-icing fluid, e.g. Kilfrost), to be directed on to the windscreens to enable the windscreen wider to clear dust and dirt from dry windscreens in flight or on the ground.

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PROTECTION

The fluid is contained in a reservoir and sprayed on to the screen through nozzles. The fluid may be directed to the nozzles by an electrically driven pump or by pressurising the top of the reservoir with compressor bleed air via a pressure reducing valve. An example of an electrically driven system is shown.

Electrically Driven Windscreen Wash System Figure 30 Servicing of the system involves functionally testing the system, replenishment of the reservoir and checks for security, leaks and damage. The system may be used in flight and on the ground. Mod 11.12 issue 1

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RAIN REPELLANT

When water is poured onto clear glass it spreads evenly to form a thin film. Even when the glass is tilted at an angle and subjected to an air stream, the glass will remain wetted and reduce vision. However, when the glass is treated with certain chemicals (typically silicone based), the water film will break up and form beads of water, leaving the glass dry between the beads. The water can now be readily removed. This principle is used on some aircraft for removing rain from windscreens. The chemical is stored in pressurised, disposable cans and is discharged on to the windscreen through propelling nozzles. Examples of rain repellent systems are shown. The following system shows a combined rain repellent and windscreen washing system.

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Combined Windscreen Wash And Rain Repellent System Figure 31

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The system shown below is a rain repellent only system and uses a disposable pressurised canister.

Rain repellent System Figure 32 The system is operated by a push button which causes the relevant solenoid valve to open. Fluid from the container is discharged onto the windscreen for a period of about 5 seconds under the control of a time delay unit. About 5cc of fluid is used with each discharge from the container which holds approximately 50 cc. The solenoid will be de-energised and the button must be re-selected for a further application. The fluid is spread over the screen by the rain which acts as a carrier. The system may be used with, or without wipers, depending on the aircraft speed, but it is normally used to supplement the wipers in heavy rain at low altitude where airspeeds are low. It is essential that the system is not operated on dry windscreens because: 

heavy undiluted repellent will cause smearing



the repellent may form globules and distort vision

If the system is inadvertently operated, the windscreen wipers must not be used as this will increase the smearing. The screen should be washed with clean water immediately. The windscreen wash system, if fitted, may be used. Rain repellent residues can cause staining or minor corrosion of the aircraft skin. Mod 11.12 issue 1

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DRAIN MAST HEATING

On many large aircraft, the water supply and water drain lines are electrically heated to prevent ice formation. Power is normally supplied via the AC bus line and is available both on the ground and in flight. 12.23

WATER SUPPLY AND DRAIN LINES

Heater tapes and blankets are wrapped around some water supply and drain lines, the temperature being controlled by thermostats. In a typical aircraft (Boeing 757), the thermostats control the heating, to open when the temperature exceeds 15.5ºC and closes when the temperature drops to 7.2ºC. Heating gaskets may be installed on the ends of toilet drain pipes. 12.24

DRAIN MASTS

Drain masts are heated to allow in-flight drainage without freezing. Drain mast heating is controlled by an air/ground relay. Low heat is supplied on the ground and high heat in flight. Figure 37 overleaf illustrates some of the heating methods used.

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Waste Water Heater Components Figure 33 Mod 11.12 issue 1

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INTENTIONALLY BLANK

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MODULE 11.13 LANDING GEAR

engineering

CONTENTS 13 LANDING GEAR .......................................................................... 13-1 13.1 13.2 13.3 13.4 13.5

13.6 13.7

13.8 13.9 13.10 13.11 13.12

13.13 13.14 13.15 13.16

13.17

13.18

INTRODUCTION ............................................................................. 13-1 GENERAL ..................................................................................... 13-2 CONSTRUCTION............................................................................ 13-3 MULTIPLE AXLES AND WHEELS .................................................... 13-5 SHOCK ABSORBING ..................................................................... 13-6 13.5.1 Oleo-pneumatic without separator ................................. 13-6 13.5.2 Oleo-pneumatic with separator...................................... 13-8 13.5.3 Liquid Spring ................................................................. 13-8 SERVICING – FILLING AND CHARGING ............................................ 13-8 EXTENSION AND RETRACTION SYSTEMS ....................................... 13-10 13.7.1 Extension System ......................................................... 13-11 13.7.2 Retraction System ......................................................... 13-12 SELECTOR VALVE ........................................................................ 13-12 UPLOCK MECHANISM .................................................................... 13-13 DOWNLOCK MECHANISM ............................................................... 13-14 EMERGENCY LANDING GEAR OPERATION ...................................... 13-16 LANDING GEAR DOORS SEQUENCING............................................ 13-17 13.12.1 Door Operated Sequencing System .............................. 13-18 13.12.2 Gear Operated Sequencing System .............................. 13-19 SAFETY BARS .............................................................................. 13-19 INDICATIONS AND WARNING INDICATIONS AND WARNING ............... 13-19 SAFETY SWITCHES ....................................................................... 13-25 WHEELS, BRAKES, ANTISKID AND AUTOBRAKING ........................... 13-26 13.16.1 wheels........................................................................... 13-26 13.16.2 Types of Wheels ........................................................... 13-27 TYRES ......................................................................................... 13-29 13.17.1 Tyre inflation and deflation ............................................ 13-29 13.17.2 Tyre Construction .......................................................... 13-29 13.17.3 Tyre Wear Assessment ................................................. 13-30 13.17.4 Tyre Damage ................................................................ 13-33 13.17.5 Leak Holes (Awl Holes) ................................................. 13-33 13.17.6 Vent Holes .................................................................... 13-33 13.17.7 Balance Marks .............................................................. 13-33 13.17.8 Electrically Conducting Tyres ........................................ 13-33 13.17.9 Aquaplaning .................................................................. 13-34 BRAKES....................................................................................... 13-34 13.18.1 Energising Brakes ......................................................... 13-34 13.18.2 None Energising Brakes................................................ 13-34 13.18.3 Expander Tube Brakes ................................................. 13-35 13.18.4 Single Disc Brakes ........................................................ 13-35 13.18.5 Multi Disc Brakes .......................................................... 13-35 13.18.6 Brake systems .............................................................. 13-36

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MODULE NO 11.13 LANDING GEAR

13.18.7 Brake control valve ....................................................... 13-38 13.19 ANTI SKID SYSTEMS .................................................................... 13-39 13.19.1 Introduction ................................................................... 13-39 13.19.2 Electronic Anti Skid System .......................................... 13-40 13.19.3 Mechanical Anti Skid System ........................................ 13-46 13.20 AUTOBRAKING ............................................................................. 13-48 13.20.1 Selector Panel .............................................................. 13-48 13.20.2 Auto-Brake Control Unit ................................................ 13-48 13.20.3 Auto Brake Solenoid Valve ........................................... 13-49 13.20.4 System Operation ......................................................... 13-49 13.20.5 Auto Brake Termination ................................................ 13-49 13.21 STEERING .................................................................................... 13-50 13.21.1 Steering Mechanisms ................................................... 13-52 13.21.2 Nose Wheel Self Centring............................................. 13-53

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engineering 13 LANDING GEAR 13.1 INTRODUCTION

Landing gears have two main functions:  Supporting the weight of the stationary aircraft on the ground  Absorbing the loads during touchdown, the landing run and taxiing. They are divided into two main categories, fixed (non-retractable) or fully retractable. Early aircraft had fixed landing gear, which unfortunately produced a large amount of parasitic drag in flight. Since drag increases at the square of forward speed, as aircraft began to fly faster, the resulting amount of drag became too prohibitive. In the short term, this problem was resolved by simply installing streamlined fairings over the wheels. However it soon became clear that this drag could be almost completely eliminated, if the landing gear were retracted after take off and stowed out of the air-stream.

Tail Wheel Type Undercarriage FIGURE 1


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JAR 66 CATEGORY B1 MODULE NO 11.13 LANDING GEAR

13.2 GENERAL Early landing gear designs consisted of two main legs set just in front of the centre of gravity (C of G) of the aircraft and a small tailwheel at the rear end of the fuselage. Putting the C of G just aft of the main gear, ensured the aircraft very quickly attained flying attitude during take off. All aircraft at that time, were propeller-driven types and the inclined fuselage gave ample clearance between the propeller and the ground during taxiing, take-off and landing. However the main disadvantage of this configuration was the risk that the aircraft was likely to „nose over‟ when heavy braking was applied and poor vision for the crew during taxiing and the initial part of the take off run. This problem was overcome by the development of the Tricycle configuration, which is now used almost exclusively. This places the main landing gear aft of the C of G and a supporting nose gear at the forward end of the fuselage. As aircraft became larger and heavier, landing gear design included multi-leg and multiwheel configurations.

Nose Wheel Type Undercarriage Figure 2

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All landing gears have to be attached to strong points on either the fuselage or the wing structure, so that the landing loads can be absorbed and transferred safely to the aircraft structure. Smaller light aircraft use a steel leaf or tubular steel spring to act as an undercarriage (figure 3). One end is attached to a strong point on the airframe while located on the other end is the wheel and axle. The deflection of the spring tube on landing absorbs the landing loads and transmit them to the airframe. A properly conducted landing will not cause any undercarriage rebound.

Spring Tube Type Figure 3 Another simple method was to use elastic bungee cord encased in a loose weave cotton braid (Figure 4). The bungee cord is located on a series of support struts which support the wheel and axle. The bungee cord stretches on landing and transfers the landing forces into the airframe.

BUNGEE SHOCK CORD Bungee Cord Type Landing Gear Figure 4 B1 Mod 11.13 Issue No

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Larger more modern aircraft, require more complex and heavier retractable systems (Figure 5). The larger the aircraft the larger the system. The components remain similar just the size and quantities change (Figure 6). Each landing gear unit is basically a wheeled shock absorber (oleo). A forged cylinder body is attached to the airframe on trunnions to allow it to pivot when lowered and raised. Articulated side stays are located between the cylinder body and airframe strong points to give the landing gear strength and rigidity and allow the landing gearleg to fold. Drag or bracing struts may also be fitted. These absorb the high acceleration loads during take off and deceleration loads during braking. MAIN SUPPORT FRAMES

TRUNNION

MAIN ACTUATOR

DOWNLOCK ACTUATOR

BRACING STRUT

DOWNLOCK LINKAGE (TOGGLE LEVERS) SIDE STAY

MAIN OLEO

PISTON

Landing Gear Leg With Bracing Struts Figure 5 The wheel and axle assembly (bogey) is attached to the piston end. A hinged torque (scissor) link is located between the axle yoke and the cylinder body. This allows the piston to move freely in and out of the cylinder but prevents the piston and wheel assembly from swivelling. Two actuators are usually fitted. A main actuator attached to the cylinder body to raise and lower the gear and a downlock actuator located on the bracing strut which operates to cause a mechanical lock when lowered. It also unlocks the gear mechanism before raising. A Hop Damper is often used with multi wheel units to align the bogie at the correct angle for landing and absorbs minor shock loads during taxiing. It is connected between the main landing gear body and the bogie. Page 13-4

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MAIN ACTUATOR

DOWNLOCK ACTUATOR CYLINDER BRACING STRUT

PISTON SCISSOR (TORQUE) LINK

WHEEL

Oleo Type Landing Gear Figure 6 13.4 MULTIPLE AXLES AND WHEELS To allow for maximum utilisation of aircraft when operating from different runways multi wheel landing gear is used. Typical configurations are shown in Figure 7. SINGLE

DOUBLE

TANDEM

BOGIE

Wheel Axle Configurations Figure 7


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LANDING GEAR

The advantages of using multi-wheel configurations are: 

They spread the landing loads over a larger area (footprint).



They are easier to stow as the wheel volume is reduced.



They provide greater safety. As the loads are spread over several wheels a burst tyre is not so critical as the remaining wheels accept the extra loads.

The main disadvantages are: 

There are more moving parts so they need more maintenance.



They are expensive to produce



Due to the large footprint the turning circle is increased to prevent the tyres from crabbing and increasing wear.

13.5 SHOCK ABSORBING In order to absorb and dissipate the tremendous shock loads of landing, the kinetic energy of the impact must be converted into other forms of energy. This is achieved on most landing gear legs by using self-contained hydraulic shock absorbing struts. There are three main types of strut commonly used in commercial aircraft:   

Oleo-pneumatic without separator Oleo-pneumatic with separator Liquid Spring

13.5.1 Oleo-pneumatic without separator

The strut uses a compressed gas (normally nitrogen) combined with a specific quantity of hydraulic oil to absorb and dissipate the shock loads. It is essentially an outer cylinder into which an inner hollow piston is inserted. When the aircraft is airborne, the landing gear is no longer supporting the aircraft weight, consequently the piston fully extends under the influence of the nitrogen pressure. The nitrogen gas being lighter than oil, will settle in the upper portion of the cylinder with the heavier oil at the bottom. Since in this particular type of strut there is no separator between the oil and gas, there will be some aeration („froth‟) as the oil and gas mix together at the demarcation line.

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On landing, the inner piston is forced up into the outer cylinder, reducing the internal volume. A tapered metering pin and snubber knob which are an integral part of the piston, are forced into a snubber tube carried by the outer cylinder. (See Figure 8). OIL BLEED VALVE

NITROGEN/OIL CHARGING VALVE

FLAPPER VALVE (OPEN)

CYLINDER

FLAPPER VALVE (CLOSED )

INNER CYLINDER

SNUBBER KNOB CYLINDER

SNUBBER TUBE

METERING PIN

SNUBBER KNOB

PISTON

(Strut Compressed) Figure 8 PISTON

Oleo - Pneumatic without separator

(Strut Extended) Figure 9

Oil is forced into the upper chamber through a series of holes in the snubber tube and through the open flapper valve. The tapered shape of the metering pin steadily reduces the available orifice area as it compresses. The landing energy is therefore absorbed by the oil, as it is forced through the ever-decreasing sized orifice and by the compression of the nitrogen gas, as the oil is forced into the reduced volume of the upper chamber. The problem now is to absorb the recoil, to prevent the aircraft from bouncing back up from the runway. As the piston starts to extend, the oil is now forced downwards into the hollow piston. The rate at which this transfer takes place is greatly restricted by the flapper valve slamming shut, leaving only a reduced number of holes in the snubber tube to permit transfer the oil. This restriction in flow and the associated increase in internal volume, prevents rapid strut extension and thus dampens the recoil energy. (See Figure 9).


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LANDING GEAR

13.5.2 Oleo-pneumatic with separator

In this design, the principle is exactly the same as the oleo-pneumatic without separator type previously described. The main difference is the inclusion of a floating piston, to separate the oil chamber from the nitrogen chamber and therefore prevent oil and gas mixing together. It also means that the nitrogen chamber does not have to be positioned at the top of the leg, or indeed be limited to one chamber. This makes shock absorbing more efficient, less severe jolting during taxiing and will simplify servicing (see later). 13.5.3 Liquid Spring

This type does not have a gas compartment. Instead, it relies on the fact that if a piston is forced into a cylinder completely filled with oil under a static pressure, energy absorption will take place due to oil compression. Oil is generally considered to be incompressible, however it is a fluid and will obey the same rules as for a gas. At normal hydraulic system pressures (typically 3000 psi), the amount of compression is negligible. However, in liquid spring shock absorbers, pressures in excess of 60,000 psi will often be generated and in this case the oil will be compressed. During touchdown, the inner piston is forced up into the upper cylinder as before, compressing the oil as the volume progressively reduces by what is known as, „jack ram displacement‟. A restrictor valve inserted as before, will absorb the recoil in a similar manner to the previous two types. 13.6 SERVICING – FILLING AND CHARGING To guarantee the correct operation of the shock absorber, the strut must be serviced in order to fill the leg with the proper quantity of oil. Additionally, the oil must be completely free of air. The nitrogen chamber must also be charged to the correct value in order to maintain the correct oil/gas ratio. When correctly filled and charged, the strut will adopt the correct extension when supporting the aircraft on the ground and the risk of the inner piston coming into contact with the outer cylinder („bottoming‟) during touchdown will be eliminated. Filling and charging procedures will vary between aircraft type, will be detailed in the Aircraft Maintenance Manual (AMM) and must be strictly adhered to. A general sequence of events to fill and charge a typical oleo-pneumatic without separator type of strut (conforming to relevant health and safety regulations), is detailed as follows:

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engineering 

Normally the aircraft will be positioned on jacks with the wheels clear of the ground.  Using an approved adapter, completely release the nitrogen pressure via the charging valve and ensure the valve remains open after all pressure has been dissipated.  Place a bottle jack under the strut and carefully compress the leg, pushing the inner piston into the outer cylinder until it bottoms and the leg is fully compressed.  Open the hydraulic bleed valve and pump oil into the oil filling connection until fresh clean oil, completely free of air bubbles, emerges from the bleed valve. The leg is now completely filled with oil to the correct quantity.  Close and tighten the oil charging valve and oil bleed valve.  Remove the bottle jack, connect a nitrogen rig to the nitrogen charging valve. Slowly and carefully inflate the leg with nitrogen until the leg is fully extended and the inflation adapter gauge shows the correct gas pressure obtained from the AMM.  Close and tighten the nitrogen charging valve and remove the charging rig.  Repeat if required on the other main leg.  Lower the aircraft off jacks. The legs are now properly filled and charged. OIL BLEED POINT

OIL CHARGING VALVE OIL OIL BLEED

SEPARATOR

SEPARATOR

CHARGING VALVE

GAS OLEO - PNEUMATIC WITH SEPARATOR Figure 10 LEG EXTENDED Figure 11


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Note 1: If the leg is an oleo-pneumatic with separator type, there will be an additional procedure before deflating the nitrogen pressure to ensure the separator is in its correct position. Note 2: The procedure is similar with a liquid spring type regarding the oil filling and bleeding, there will be no nitrogen charging procedure. Note 3: In-service, the serviceability of the shock struts can be monitored with the use of a pressure/extension graph and adjustments may be made to the nitrogen pressure as required.

GAS PRESSURE (PSI) (GAUGE PSI)

OLEO PRESSURE/EXTENSION GRAPH Figure 12 13.7 EXTENSION AND RETRACTION SYSTEMS As the speed of the aircraft becomes high enough that the parasite drag of the landing gear is greater than the induced drag caused by the added weight of the retracting system it becomes economically practical to retract the landing gear into the aircraft structure. Raising or lowering of the undercarriage is carried out either hydraulically or pneumatically via a selector lever in the cockpit which is mechanically or electrically linked to a selector valve. When the selector valve is operated it directs the fluid to one side or the other of the piston.

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The landing gear is uplocked and downlocked mechanically or hydraulically through the uplock boxes and the downlock toggle levers. Landing gear positions are sensed by proximity switches or microswitches and transmit these positions to the cockpit instrumentation via a control unit. In the case of fluid or electrical failure, a mechanical emergency lowering system is available. An emergency handle located in the cockpit is operated and by a system of push-pull cables and gearboxes, the uplocks are released. The landing gear selector valve or a freefall valve is also operated, which opens all extension and retraction lines to return. The landing gear is allowed to fall under gravity and aerodynamic forces but may be assisted by a spring or gas operated free fall assister. Smaller light aircraft may use differing methods for operating the landing gear. Electric motors may drive actuators, a winding cable system, a simple operating lever with safety locks or a manual hydraulic jacking system may be used to raise or lower the landing gear. Most modern light aircraft use a hydraulic power pack. This is a self-contained system and was designed to be lightweight and easy to maintain. The pack contains the fluid reservoir, sight glass, pressure pump, filter, thermal relief valve, pressure relief valve, ground service and replenishment connections. 13.7.1 Extension System

When the selector lever is selected to GEAR DOWN a micro-switch on the lever is made which powers up the hydraulic pump, the hydraulic pressure is then fed to the uplock actuator valves to unlock the uplocks. Once operated, the uplock hooks remain mechanically open under spring pressure. Movement of the undercarriage legs break the uplock limit switches which indicates on the instrumentation panel that the landing gears are in transit.(red triangles) and that the undercarriage is unlocked. The landing gear selector valve operates, and the down lines to the actuators and the return lines to the reservoir are opened. The fluid pressure flows through the selector valve to the actuators and extends the actuators. Once the main actuators are fully extended and the undercarriage legs have mechanically locked, excess pressure is bled back through the low pressure control valve to the reservoir. When all 3 wheels are down and locked, proximity switches send signals to a control unit which turns the hydraulic pump off, closes the selector valve lines and sends signals to the instrument panel indicating that the undercarriage is locked down, (green triangles).


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13.7.2 Retraction System The retraction procedure is basically the opposite of the extension procedure. When the selector lever is selected GEAR UP a micro-switch on the lever is made which powers up the hydraulic pump, the hydraulic pressure is then fed to the downlock actuators to unlock the mechanical locks on the bracing struts. Its is also fed to the selector valve and opens the uplines to the main actuators and the return lines to the reservoir. Movement of the undercarriage legs breaks the downlock proximity switches which send signals to the control unit which indicates on the instrumentation panel that the landing gears are in transit, (red triangles) and that the undercarriage is unlocked. The fluid pressure flows through the selector valve to the main actuators and retracts the landing gear. The undercarriage legs on full retraction mechanically lock the uplocks. Once the main actuators are fully retracted and the undercarriage legs are locked up, excess pressure is bled back through the low pressure control valve to the reservoir. When all 3 wheels are up and locked, uplock limit switches send signals to a control unit which turns the hydraulic pump off, closes the selector valve lines and change the red triangles to black on the indicating panel. If a red triangle remains on when the undercarriage is fully extended or retracted there is a fault in the system. A squat switch system and an electro-mechanical stop on the selector lever, will prevent the landing gear from being retracted when the aircraft is on the ground. The landing gear will not be able to be retracted until certain parameters are met. This is normally when all landing gear legs have fully extended after take off. This is sensed by proximity switches on each leg. 13.8 SELECTOR VALVE The selector valve on modern large aircraft will be normally operated by electrical solenoids signalled from micro-switches in the landing gear selector lever, but on some aircraft they may be mechanically operated. A spool valve in the selector valve is moved from a neutral position one way or the other allowing hydraulic pressure to one side of the main actuator piston, depending whether the landing gear is to be raised or lowered . Normal operation of the selector valve can be overridden in case the landing gear has to be lowered in an emergency, if the landing gear fails to extend due to a system fault. The spool valve is moved mechanically by a system of rods, cables and levers to allow all lines to be opened to allow the free flow of hydraulic fluid around the system. This operation is normally inter-linked with the emergency mechanical opening of the uplocks. A typical selector valve is shown in Figure 13

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SPOOL VALVE

SOLENOID

SOLENOID

MECHANICAL OVER-RIDE LINKAGE

Selector Valve Figure 13 13.9 UPLOCK MECHANISM On large modern aircraft when the landing gear is being retracted the uplocks will operate mechanically. A roller on the landing gear leg will locate and engage into the uplock hook. Limit switches will sense when the landing gear leg has engaged in the lock hook and will turn off the hydraulic pressure. The gear will then be held retracted in place purely mechanically. (Figure 14) LOCK LEVER ASSEMBLY

LIMIT SWITCH

UNLOCK ACTUATOR VALVE LANDING GEAR LEG ROLLER UPLOCK HOOK

Locked Uplock Figure 14 B1 Mod 11.13 Issue No

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Normal release of the uplock is by a hydraulically actuated valve. The supplied hydraulic pressure pushes a plunger against the lock lever which rotates about its pivot. This action allows the uplock hook to disengage under its own spring tension. The landing gear will then be extended hydraulically by the main actuator. (Figure 15) LOCK LEVER ASSEMBLY LIMIT SWITCH

PLUNGER UNLOCK ACTUATOR VALVE UPLOCK HOOK

LANDING GEAR LEG ROLLER

Unlocked Uplock Figure 15 13.10

DOWNLOCK MECHANISM

The downlock actuator can have either a single or double direction operation depending on the aircraft. A single direction operation would unlock the downlock mechanism (upper and lower toggles) prior to retraction, the leg relying on its own extension to provide the over centre lock. The double direction actuator will lock the downlock mechanism on extension and unlock it prior to retraction. Once the landing gear has been fully extended and is sensed by a limit switch hydraulic pressure is directed to the downlock actuator which extends the actuator piston. The piston acts against a toggle lever which move both toggle levers to an over centre position. This over centreing of the toggle levers forms a mechanical lock which prevents the landing gear leg from collapsing. (Figure 16)

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MAIN LEG

DOWNLOCK ACTUATOR

PROXIMITY SWITCH

SIDE BRACE

UPPER TOGGLE LEVER LOWER TOGGLE LEVER

PROXIMITY SWITCH CENTRE LINE

OVER CENTRE POSITION

Linkage Downlocked Figure 16 Once the aircraft has landed and parked up, a red flagged safety pin is inserted through alignment holes in the toggle levers to prevent inadvertent collapse or retraction of the landing gear on the ground. This safety pin is removed before flight.


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On selecting the landing gear up, the hydraulic pressure is directed initially to the downlock actuator and retracts the piston. As the piston retracts it moves the lower toggle overcoming the mechanical lock, moving both toggle levers from the over centre position to an under centre position, so that the landing gear can now fold. (Figure 17) MAIN LEG DOWNLOCK ACTUATOR

PROXIMITY SWITCH

SIDE STAY UPPER TOGGLE LEVER UNDER CENTRE POSIITION LOWER TOGGLE LEVER PROXIMITY SWITCH

CENTRE LINE

Linkage Unlocked Figure 17 13.11

EMERGENCY LANDING GEAR OPERATION

The uplocks can be released manually if the actuator or hydraulic system fails. An emergency landing gear lever, operated from the cockpit will act on and rotate the hook locks, releasing the landing gear legs from the uplock hooks. The emergency mechanism lever will also operate a lever on the landing gear selector valve which will open all hydraulic lines to return. This allows the hydraulic fluid to free flow through the system, to allow the landing gear to extend. Once the uplocks are released the landing gear legs will extend under gravity and aerodynamic forces. Spring or gas operated free fall assistors may be used to help the gear extend. The proximity and limit switches will operate as normal giving a cockpit indication of the gear in transit and down locked.

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HOOK LINK ASSEMBLY LEVER

LIMIT SWITCH

UPLOCK HOOK UNLOCK ACTUATOR VALVE CABLE

EMERGENCY OPERATING HANDLE

Emergency Release Mechanism Figure 18 On aircraft fitted with hydraulically sequenced doors if the hydraulic system fails, the door jack is mechanically unlocked. This will also be carried out by a mechanical linkage connected to the cockpits emergency release mechanism (Figure 18) 13.12

LANDING GEAR DOORS SEQUENCING

To keep the aircraft as streamlined as possible and to reduce drag, the landing gear is normally retracted into bays within the aircraft structure. However some aircrafts landing gear do not fully retract into the structure and some access doors do not fully enclose the landing gear. The bays have access doors which open and close in relation to the movement of the landing gear. Some doors are mechanically linked to the landing gear, by a system of connecting rods, bellcranks and links, whilst other doors open and close under operation from a hydraulic sequencing valve, signalled by microswitches or proximity switches via a control unit. B1 Mod 11.13 Issue No

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To further reduce the drag some doors will close when the landing gear has been extended. The landing gear doors may have a manual unlocking mechanism to allow the door to be opened on the ground for access in carrying out maintenance tasks and inspections. Anything that jeopardises the sequence can cause considerable damage to the aircraft structure and could lead to an unsafe landing condition. Door sequencing relies on the movement of valves operated by the doors and the movement of the legs. The sequencing valve can be therefore be either door operated or gear operated. 13.12.1 Door Operated Sequencing System Only when the door is fully open is pressure allowed to flow to the main actuator. If the door is not fully open the main actuator remains isolated. Hydraulic pressure is initially fed to the landing gear door actuator which operates to open the door. When the door reaches its maximum travel it abuts against. and depresses a plunger. (Figure 19) The movement of the plunger unseats a valve in the sequence valve, which opens a gallery to allow fluid pressure to the main actuator and extends the landing gear down. TO DOOR ACTUATOR

PRESSURE IN VALVE SEAT

PLUNGER

TO MAIN ACTUATOR

Sequence Valve – Door Shut Figure 19

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Retraction of the landing gear is reversed. Pressure is fed to the main actuator which retracts the landing gear leg. When the landing gear leg is fully retracted it abuts against and depresses a sequence valve plunger. The movement of the plunger unseats a valve in the sequence valve, which opens a gallery to allow fluid pressure to the door actuator which closes the door. (Figure 20) TO DOOR ACTUATOR

PRESSURE IN VALVE SEAT

PLUNGER

TO MAIN ACTUATOR

Sequence Valve – Door Open Figure 20 13.12.2 Gear Operated Sequencing System The principle of operation is very similar to the door operated mechanism. The difference being that the plunger (or slide) is operated via a cam and linkage mechanism directly attached to the landing gear leg. This ensures that when the gear starts to move the door starts to, or is in the process of opening. 13.13

SAFETY BARS

On some aircraft with hydraulically sequenced doors if the hydraulics system was to fail, to allow the landing gear to lower, the wheels will forcibly open the doors. This is done by the landing gear legs pushing against safety bars which are fitted to the doors. The doors will open without being damaged and once operated the doors will remain open. 13.14

INDICATIONS AND WARNING INDICATIONS AND WARNING

All modern aircraft fitted with retractable landing gear will have a means of indicating on the flight deck whether the legs are locked down, in transit or correctly locked up. Additionally, a separate warning system may be included to show faults, or to indicate that the legs are not in the position selected („nips‟). B1 Mod 11.13 Issue No

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Normally leg position is shown, by a dedicated set of coloured indicators on the front panel, near to the landing gear selector lever. Each leg will have its own set of indicator lights. On some aircraft a „nips‟ light is included in the selector lever itself. The actual sequence of indication often varies from aircraft to aircraft, but the modern „dark cockpit‟ philosophy during flight, usually means that all indicator lights are extinguished (no lights), when the legs are properly locked up. Red lights are often used when the legs are in transit (i.e.: not locked up and not locked down) and green lights illuminate when each leg is down and locked.

RED TRANSIT GREEN LOCKED DOWN

A RED LIGHT COMES ON WHENEVER: 1. THE LEVER IS NOT DOWN AND GEAR NOT UP. 2. THE LEVER IS DOWN AND GEAR NOT DOWN AND LOCKED. 3. ENGINE No. 1 OR 2 THROTTLE IS IN IDLE RANGE AND ANY GEAR NOT DOWN AND LOCKED. A GREEN LIGHT COMES ON WHENEVER: 1. THE GEAR IS DOWN AND LOCKED.

Gear Position Indicator Figure 21 On other aircraft, the red transit lights are replaced by the „nips‟ light in the selector lever, and separate amber warning lights on the front panel will show a fault. (I.e.: if any leg fails to reach its selected position, either locked up or locked down, within a certain time limit.) Also, where for example, visual confirmation from the cabin windows is not possible, usually for nose gear, the locked down indicator may be duplicated, as an additional „confidence light‟, in case a bulb failure occurs. Page 13-20

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ADDITIONAL NOSE LOCKED DOWN GREEN GREEN LOCKED DOWN LIGHTS

AMBER „FAULT LIGHTS‟

RED „NIPS‟ LIGHT INSIDE HANDLE

Landing Gear Selector and Indications Figure 22 Micro switches or proximity sensors are fitted to each leg to relay information the flight deck indicators. A change the output voltage whenever the uplock or downlock mechanisms are made or broken during the retraction or lowering sequences, determine indicator output.


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MAIN LANDING GEAR


MAIN LANDING GEAR DOWNLOCK SENSOR Gear Downlock Sensors Figure 23

Other methods can be mechanical indicators outside the aircraft, visible from the cockpit. There may be painted indicator lines on the landing gear legs toggle levers which align when the gear is down and locked. (Figure 24)

UNLOCKED LOCKED

Landing Gear Down Locked Visual Indicator FIGURE 24 Page 13-22

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Some aircraft have pop up indicators which stand proud on the upper wing surface when the gear is down and locked (Figure 25). These are plunger operated through a cable linkage attached to the toggle levers. When the landing gear extends and is locked down a plate attached to the toggle lever operates a spring loaded plunger which by cable connection moves the indicator from its housing, proud of the airframe skin. The indicator returns under spring pressure into its housing when the landing gear is retracted

POP UP INDICATOR

AIRFRAME SKIN

UNLOCK ACTUATOR PLUNGER

TOGGLE LEVERS SIDE STRUT

POP UP INDICATOR

AIRFRAME SKIN

UNLOCK ACTUATOR PLUNGER

TOGGLE LEVERS

SIDE STRUT

Landing Gear pop Up Indicator Figure 25 B1 Mod 11.13 Issue No

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To prevent the pilot from landing with his under carriage retracted there may be a warning system connected to the centralised warning panel with associated warning lights and audio warnings. The warning system may be activated when the aircraft descends to a certain height above the ground detected by the radio altimeter, or when the landing configuration is incorrect ie, when the engine power levers or flaps are set incorrectly. SAFETY LATCH PIN UP

LANDING GEAR SELECTOR LEVER

DOWN SAFETY LANDING

SOLENOID

GEAR LEG

DE-ENERGISED

EXTENSION CONTROL UNIT

LIMIT SWITCHES

Landing Gear Selector Lever Safety Interlock Figure 26 SAFETY LATCH PIN UP

LANDING GEAR SELECTOR LEVER

DOWN SAFETY SOLENOID

LANDING

ENERGISED

GEAR LEG EXTENSION CONTROL UNIT

LIMIT SWITCHES

Landing Gear Selector Lever Safety Interlock Figure 27 Page 13-24

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The landing gear may have an electro-mechanical safety device, which prevents operation of the selector lever on the ground. When all the landing gear legs are compressed a safety solenoid is de-energised which moves a latch pin under the landing gear selector lever. So long as the solenoid remains de-energised the latch pin prevents the selector lever from operating. As soon as each landing gear leg is fully extended the limit switch is made which sends a signal to the control unit. When the control unit receives signals from all the landing gear legs an earth is made and the safety solenoid is energised. The latch pin is withdrawn from beneath the selector lever allowing gear up when selected. (Figures 26 and 27) 13.15

SAFETY SWITCHES

Proximity switches on each landing gear leg will indicate that the landing gear leg is either downlocked or is in transit. The switch will be made when the target on the landing gear leg comes into alignment with the switch probe indicating that the landing gear is downlocked. The gap between the probe and target is set in accordance with the maintenance manual for the aircraft. When the proximity switche probes are out of alignment with their targets, the switches are broken and it is sensed that the landing gear leg is in transit. The signals will be sent to an electronic control unit or computer where they are processed and will illuminate an associated green triangle on the landing gear panel when locked down and a red triangle when the landing gear is in transit. Limit micro-switches on the uplocks will sense when the landing gear is locked up and limit switches on the oleos will sense when the oleo leg is fully extended. The signals will be sent to an electronic control unit or a computer where they are processed. When the landing gear is locked up the limit switch will change the red triangles to black. When the oleos are fully extended the limit switches will allow the landing gear to be retracted. The proximity switches and limit switches form part of the weight on wheels, weight off wheels squat switch system and will prevent inadvertent retraction of landing gear on the ground and will only allow retraction when certain circumstances are met. This mainly being that all 3 landing gear legs are weight off wheels and are fully extended, and the downlocks have been unlocked.


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WHEELS, BRAKES, ANTISKID AND AUTOBRAKING

13.16.1 wheels

The wheels on the landing gear leg provides some form of suspension and adhesion between the aircraft and the ground. Early wheels and tyres were of the bicycle type with spoke rims and with the tyres fitted using tyre levers. Most light aircraft have fixed flange one piece forged or cast wheels (Figure 28).

Fixed Flange Wheel Figure 28 Modern tyres are much more rigid, due to the load-bearing requirements, which results in the wheels having to be of two piece construction (Figure 29). The two piece wheel construction, are of 2 types, removable rim or split wheel. The removable rim wheel has an inner tube where as the split wheel is tubeless and requires a perfect seal between the halves. An O ring is located between the mating surfaces. To be as light and strong as possible they are usually constructed from alluminium or magnesium alloys and may be cast or forged. The inboard wheel section is fitted with key ways that allows the brake discs to slot into. These key ways drive the brake discs with the wheels. Larger aircraft wheels have one or more fusible plugs fitted. These plugs have a centre hole which is filled with a low melting point alloy. In the event of the tyre overheating, when a temperature limit is reached the low melting point alloy melts and allows the tyre to safely deflate.

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engineering 13.16.2 Types of Wheels

There are three basic types of wheel used for aircraft:  Well-based  Divided (or Split)  Loose and Detachable Flange 13.16.2.1 WELL-BASED

This type is limited to smaller light aircraft and is similar to those found on a typical family car. 13.16.2.2 DIVIDED (OR SPLIT)

This type is used on most modern commercial airliners. It consists two half assemblies matched up and bolted together to form the complete wheel. Each half is more or less identical and has its own tapered bearing assembly. A sealing ring is incorporated between the two halves, to provide an airtight joint when the wheel is used with a tubeless tyre. Additionally, the inner half will carry the brake rotor drive blocks and the outer half may be fitted with fusible plugs. Half Hub Assembly

Outer Bearing Inner Bearing Sealing Ring Drive Block Mounting Divided (Split) Wheel Figure 29


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13.16.2.3 LOOSE AND DETACHABLE FLANGE

This type of wheel has a main hub, which carries both bearings, brake rotor drive blocks and fusible plugs. To facilitate tyre replacement, one of the two wheel flanges can be removed. The flange when refitted to the wheel hub is retained by a locking ring (loose flange) or by means of a series of nuts and bolts (detachable flange). As with the divided wheel a sealing ring is incorporated in the flange recess to provide the airtight joint when used with tubeless tyres.

Locking Ring

Loose Flanged sealing Figure 30

Loose Flange

Spigot Joint

Three Piece Loose Flanged Wheel Figure 31

Inner Bearing

Outer Bearing Drive Blocks

Detachable Flanged Wheel Figure 32 Page 13-28

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engineering 13.17

TYRES

Tyres with patterned tread became important when aircraft got effective brakes that could be used for slowing the aircraft during landing. At first the treads were a diamond pattern that provided good braking on wet grass but the ribbed tread proved to be more suitable for operation on hard surface runways. Today almost all aircraft tyres have a ribbed tread that consists of straight grooves, which run around the tyres‟ circumference. 13.17.1 Tyre inflation and deflation

The tyres are inflated with nitrogen from a ground cart. The required pressure will be laid down in the AMM and a tyre inflation box is used to regulate the charge rate and pressure. A deflation tool is used to release the pressure and any ice that forms must be allowed to thaw before the valve core is removed. 13.17.2 Tyre Construction

. 

The Bead

The bead gives the tyre its strength and stiffness to assure a firm mounting on the wheel. The bead is made up of bundles of high strength carbon steel wire with two or three bead bundles on each side of the tyre. Rubber strips streamline the round bead bundles to allow the fabric to fit smoothly around them without any gaps. The bead bundles are enclosed in layers of rubberised fabric, to insulate the carcass plies from the heat absorbed in the bead wires. 

The Carcass

The carcass (or chord body) is the body of the tyre that is made up of layers of rubberised fabric cut in strips with the threads running at an angle of about 45 degrees to the length of the strip. These strips extend completely across the tyre around the bead and partially up the side. Each ply is put on in such a way that the threads cross each other at about 90 degrees to that of the adjacent ply. This type of construction is known as bias ply. The cords of the ply fabric were originally cotton, then nylon and now aramid fibres (kevlar) are used. This is stronger than nylon, polyester or fibreglass and even strong pound for pound than steel. Chafing strips are rubberised strips of fabric that wrap around the edges of the carcass plies and enclose the bead area. The chafing strips provide a smooth chafe resistant surface between the tyre and the bead seat of the wheel.


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The undertread is a layer of compound rubber between the plies and the tread rubber that provides good adhesion between the tread and the carcass. On top of the undertread are more plies of strong fabric that strengthen the tread and oppose centrifugal forces that try to pull the tread from the carcass during high speed rotation. The inner liner is a thin coating of rubber over the inside plies. For tubeless tyres it is made from a compound which is less permeable than other rubbers used. It seals the tyre and reduces the amount of leakage. On tyres with inner tubes the liner is very smooth to help prevent chafing. TREAD PLIES

SIDEWALL CHAFING STRIPS

CARCASS

BEAD BUNDLE

BEAD WIRES

Aircraft Tyre Construction Figure 33 

The Tread

The tread is the thick layered rubber around the outer circumference of the tyre that serves as a wearing surface. The tread has a series of moulded grooves moulded into its surface to give optimum traction with the runway surface. 13.17.3 Tyre Wear Assessment

The manner in which tread wear of a tyre is established, is dependent upon which of a number of methods of indicating wear has been incorporated into the tyre by the manufacturer. Page 13-30

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Tyres used on modern aircraft have a series of circumferential grooves in the tread, primarily to displace water on the runway and so help to prevent the tyre from aquaplaning. These grooves can be also used as a means of establishing tyre wear. If this method is adopted, then wear which results in any groove being less than 2mm in depth, for more than 25% of the tread circumference, requires the tyre to be replaced. Other ways of establishing wear assessment are by the use of:  Tie Bars  Wear indicator Grooves  Sipes 13.17.3.1 TIE BARS

These are small transverse bars of rubber, moulded at intervals in the circumferential grooves around the tyre as described above. They are set at a depth of 2mm, or as required by the particular manufacture and thus provide an easy visual means of establishing wear limits. Limits – tyre worn to the top of the tie bar.

Tie Bars

Tie Bars Figure 34 B1 Mod 11.13 Issue No

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13.17.3.2 WEAR INDICATOR GROOVES

These are dedicated grooves set in the tread pattern and have a depth graduated by the manufacturer, but typically 2mm shallower than the water-displacing grooves. Limits – tyre worn to the bottom of the indicator groove anywhere on the circumference of the tyre.

Wear Indicator Grooves Figure 35 13.17.3.3 SIPES

Certain tyres, normally those having a zigzag tread pattern have an axial slit in the tread rubber at some of the zigzag corners. The slit does not extend into the depth of the tread and is called a sipe. Limits – Tyre worn to the bottom of the sipe.

Sipes

Sipes Figure 36 Page 13-32

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engineering 13.17.4 Tyre Damage

The amount of tyre damage a tyre can suffer without becoming unserviceable is very small. Damage in the vicinity of the bead is rarely tolerated, while cuts in the casing plies must be assessed very carefully in accordance with the manufacturer‟s requirements before deciding on the degree of serviceability. Normally if the chords are exposed due any form of damage, including splits or crazing, then the tyre will be classed as unserviceable. NOTE: Always consult the Aircraft/Component Maintenance Manual. 13.17.5 Leak Holes (Awl Holes)

During inflation of a tyre/tube assembly, air may become trapped between the tube and the inside surface of the tyre, giving an incorrectly inflated assembly. The risk is reduced by allowing the air to escape through Leak Holes, pierced completely through the sidewall of the tyre, during manufacture. The holes are often made with a pointed tool called an Awl. Because of this, the holes are sometimes referred to as Awl Holes. The position of these holes is indicated by a series of 6mm diameter spots of grey or green litho ink, usually grey. 13.17.6 Vent Holes

During the manufacture of tubeless tyres, air that gets trapped between layers in the casing is permitted to escape to atmosphere through vent holes pierced in the sidewall. The vent holes do not penetrate right through the sidewall in this case and are identified, as with leak holes, by 6mm diameter spots of grey or green litho ink, usually green. 13.17.7 Balance Marks

A red spot (sometimes triangular) on either side of the tyre indicates its lightest point around the circumference as ascertained during the manufacturer‟s balancing procedure. During assembly with the wheel the red spot should be aligned with the inflation valve on a tubeless assembly. On a tubed assembly, the spot should be aligned with a red line (heavy point) on the tube. If it has no red line, align with the inflation valve of the tube. 13.17.8 Electrically Conducting Tyres

Some wheel assemblies are fitted with tyres that are designed to conduct electrical charges to earth as the aircraft touches down. Such tyres are identified with the word CONDUCTIVE or the letters ECTA (electrically conducting tyre assembly) on the sidewall.


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13.17.9 Aquaplaning

Aquaplaning is a condition that occurs on wet runways when a wave of water builds up in front of a spinning wheel. This could result in the tyre being lifted from the runway surface and to float on the thin layer of water. This is dangerous, as a complete loss of braking efficiency will occur. Although it appears only to be an aircrew problem, there is a significant factor that affects the maintenance engineer. Mathematically there is a formula for Aquaplaning speed Aquaplaning Speed (Kt.) = 9 (approx.) x Square Root of the Tyre Pressure. This speed will be placarded for the crew, so that in wet conditions they will quickly traverse through it on landing. However, if the tyre pressures are incorrect, the placarded speed will be useless and aquaplaning will occur at a different speed. Take care, therefore, to maintain tyre pressures at their correct value at all times.

13.18

BRAKES

Aircraft brake systems convert kinetic energy from the motion of the aircraft into heat energy, which is generated by the fiction between the brake linings and the brake drum or disc. There are two types of brakes in use energising (servo) and none energising. Energising brakes use the friction developed between the rotating and stationary parts to produce a wedging action that uses the momentum of the aircraft to increase the braking force which reduces the pilots effort needed in producing the required braking action. None energising brakes do not use this wedging action. 13.18.1 Energising Brakes

Energising brakes used on some smaller light aircraft have a single servo action and only operate with forward motion. Energising brakes have their shoes and linings mounted on a torque plate in such a way that they are free to move out against the rotating drum. When the brakes are applied two pistons move out and push the linings against the drum that rotates with the wheel. Rotation of the brake drum wedges the linings against it. When the hydraulic pressure is released, a retracting spring pulls the linings form the drum and releases the brakes. 13.18.2 None Energising Brakes

This is the most common type of brake used on aircraft. These brakes are actuated by hydraulic pressure and the amount of braking action depends on the pressure applied. Expander tube, single disc and multiple disc brakes are the main types of none energising brakes used.

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engineering 13.18.3 Expander Tube Brakes

Expander brakes utilise a heavy neoprene tube, but are rarely use on modern aircraft. Hydraulic fluid from the master cylinder is directed into the expander tube which is located on the circumference of a torque flange. When this tube is expanded it pushes the brake block linings out against the brake drum and the friction between the linings and the drum slows the aircraft. The heat generated in the linings is kept from damaging the expander tube by stainless steel heat shields placed between each of the lining blocks. As soon as the brake pedal is released, the return springs between the brake lining blocks collapse the expander tube and force the fluid back into the cylinder reservoir. 13.18.4 Single Disc Brakes

This is most common on light aircraft. The brakes are actuated by hydraulic pressure from a master cylinder and friction is produced when the rotating disc is squeezed between the brake linings in the brake caliper. There are two types of single disc brakes, one has the disc keyed into the wheel and it is free to move in and out as the brake is applied. This type is called floating disc fixed caliper. The second type of brake disc is rigidly attached to the wheel and the caliper moves in and out on anchor bolts. This type is called fixed disc floating caliper. Some single disc brakes have automatic adjusters and wear indicators. The automatic adjusting pin is pulled through the grip when brakes are applied. When the brakes are released the piston and the linings move back only under pressure of the return spring. The protrusion on the adjuster pin indicates lining wear. In general, when the pin is flush with the housing the linings are replaced. 13.18.5 Multi Disc Brakes

The gross weight of the aircraft and the speed at the time of brake application determines what size brakes are required. As the aircraft‟s size, weight and landing speed increases there is a need for greater braking surfaces and heat dissipation. Segmented rotor, multiple disc brakes are standard on most modern high performance aircraft. The segmented disc brake has three rotating discs keyed on to the wheel. The rotors are segmented to allow for cooling and for expansion caused by the high temperatures generated during braking. Between each disc is a stator plate or brake-lining disc, keyed on to the axle shaft. Riveted on to each side of the stator plates are the brake linings. A pressure plate is located on the inboard side of the axle shaft and a backing plate is located on the outboard side. B1 Mod 11.13 Issue No

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Automatic adjusting pins are pulled through the grip when brakes are applied. When the brakes are released the pressure plate moves back under pressure of their return springs. The protrusion on the adjuster pins indicates lining wear. In general, when the pin is flush with the housing the linings are replaced. PRESSURE PLATE PISTON BRAKE LINER SEGMENTED ROTOR CYLINDERS PADS

STATOR PLATE

BACKING PLATE

WEAR PINS

Multi Disc Brake Unit Figure 37 The brakes used on most large jet aircraft use a number of brake cylinders instead of a single annular cylinder. (Figure 37) Each cylinder has a piston which presses against the pressure plate when hydraulic pressure is applied. Each cylinder will be supplied from separate hydraulic systems so if one fails full braking can be applied from the other system. Some aircraft may have their brake discs made from carbon fibre. These are lighter in weight and they can function at higher temperatures. They are expensive to use and generally only used on transport aircraft where the weight saving makes them more cost effective. 13.18.6 Brake systems

Light aircraft will generally use hydraulic pressure generated by the pilots‟ feet. When the pilot depresses the rudder pedals, pressurised fluid is moved from the master cylinder, to a slave cylinder operating the brakes. Larger aircraft will use the aircraft‟s main hydraulic systems to provide the pressure to operate the brakes. Page 13-36

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The pressure applied to the brakes must be proportional to the force exerted on the brake pedals; the pilot must be able to hold the brakes partially applied without a build up in pressure. The hydraulic pressure to the brakes is much higher but remains proportional to the input. This is achieved with a brake control valve also known as a metering valve. The rudder pedals are connected to the brake control valve by various methods including hydraulically by use of a master cylinder (also known as foot motors), rods or cables. The hydraulic systems will operate simultaneously and usually a different system will feed the inboard wheels to the outboard. In the event of a system hydraulic failure, braking is still maintained to at least one set of wheels. PILOTS FOOT MOTORS

1ST

PILOTS FOOT MOTORS

1ST

2ND SYSTEM PRESSURE

1

2ND

RETURN

2

BRAKE CONTROL VALVE 1 2

1

2

Brake control system Figure 38


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engineering 13.18.7 Brake control valve

The schematic drawings of the brake control valve (figures 39, 40 and 41) shows a simplified version of how the proportional application is achieved. The centre slide moves to the left as the pilot applies the brakes, opening the pressure line and closing the return line. This allows pressure to the brakes and they are applied. At the same time pressure is directed to the metering chamber were pressure builds up until it equals the pedal input pressure. When the pressures are equal the slide moves to the right, until it is in the central position, with both the pressure and return lines blocked. This holds the brake pressure constant until the pressure is either increased or decreased by a change in the pilots‟ input. If the pedals are released the slider will move to the right opening a line from the brakes to return, dissipating the pressure. RETURN

PRESSURE

BRAKE

S

Figure 39 Brakes released -The return line is open for the pressure to dissipate.

RETURN

PRESSURE

BRAKE

S

Applied brake pressure

Monitoring chamber pressure Figure 40

As brake pressure is applied the slider moves to the left blocking the return line and opening the brake line to the pressure. Pressure is fed to the monitoring chamber were it starts to move the slide to the right as it equals the input force.

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engineering PRESSURE

RETURN

BRAKE

S

Applied brake pressure

Monitoring chamber pressure

Figure 41 With the pressure equal to the input force the slide moves to the central position with both the pressure and return lines blocked off. In this position a constant brake pressure is held to the brakes. 13.19

ANTI SKID SYSTEMS

13.19.1 Introduction

The anti skid system is designed to provide maximum effective braking for any runway condition without skidding and is often used in conjunction with an autobrake system. It operates by automatically overriding or modifying the metered input brake pressure from the flight deck, or braking commands from the autobrake system. Hydraulic pressure is automatically controlled at each brake unit, maintaining the optimum wheel braking requirement, regardless of prevailing weather conditions (ie: ice/heavy rain/crosswind etc). Aircraft stopping distances are minimised and directional control is maintained. Maximum braking efficiency occurs when all main wheels are at the maximum rate of deceleration just before an impending wheel skid. The system continuously modulates the hydraulic pressure at each individual brake unit in response to actual wheel speed, thus preventing blown tyres, flat spots or the risk of aquaplaning caused by a locked wheel. On a normal landing sequence, there is no need for a corrective signal as long as the rate of wheel deceleration is within limits. However, if the rate is above these limits, this is sensed as an approaching skid. A corrective signal is applied to momentarily reduce the applied brake pressure at the relevant wheel. The corrective signal is removed when the wheel speed increases again and the process repeated as required, until the deceleration rate remains within limits once more. The anti skid system can be either electronically or mechanically controlled. Most modern systems are electronic, since mechanically controlled systems are only fitted to older aircraft types. B1 Mod 11.13 Issue No

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engineering 13.19.2 Electronic Anti Skid System

The system consists of the following components:  A wheel speed transducer, located in each main landing gear axle and driven by the wheel rotation.  An electronic antiskid control unit, normally located in the electronic/electrical equipment bay, with BITE facility to provide continuous self test and fault warning.  An antiskid control valve for each mainwheel, normally located in the hydraulic equipment bay.  A control switch and failure warning indicator, on the flight deck panel.

Electronic Anti-Skid System With Auto-Brake Facility Figure 42

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13.19.2.1 WHEELSPEED TRANSDUCER

The transducer is a speed sensing generator-type device, which sends an output voltage directly proportional to wheel rotation to the electronic antiskid control unit. The control unit compares the transducer output voltage, with a reference voltage scheduled to the maximum deceleration rate for the aircraft. If the transducer output voltage exceeds the reference voltage, the error signal will be sent to the relevant antiskid control valve. This results in the hydraulic pressure at the corresponding brake unit momentarily reducing, until the voltages agree once more.

WHEELSPEED SENSOR RING NUT GENERATOR CARRIER

DRIVE CAP

V-CLAMP Wheel speed Transducers Figure 43 B1 Mod 11.13 Issue No

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13.19.2.2 ANTI SKID CONTROL VALVE

This valve is a two-stage electro-hydraulic servo valve, which meters pressure applied to the brake unit in accordance with signals from the anti skid control unit. The first stage is a torque motor-operated flapper valve set between two hydraulic ports (return and pressure). The second stage is a spool valve, spring biased to „brakes on‟ position and hydraulically controlled, by directing oil pressure into a drilled passage way at either end of the spool. When there is no control signal to the torque motor, the flapper valve is biased towards the return nozzle and maximum braking is possible. However, a signal (increase in current), will be sent to the torque motor windings from the Control Unit, if it in turn receives a signal from a wheel speed transducer that a wheel is slowing down too quickly and may skid. This causes the flapper valve to move towards the pressure nozzle, restricting fluid into the chamber and allowing more to escape to return. As a result pressure reduces in the first stage chamber and the reduction is felt on the bias spring side of the second stage spool valve. Pressure on the opposite end of the spool forces the valve to move, closing off the pressure line and connecting the brake line to return. The amount of second stage valve movement is directly proportional to torque motor current in the first stage, which in turn depends on the amount of brake pressure reduction required to achieve wheel spin up. As the mainwheel spins up again to its correct speed, the current at torque motor windings reduces. This allows the flapper to move back to the return nozzle and moves the spool valve back, closing off the return line and causing brake pressure to be re-applied to the wheel brake. If necessary, the complete cycle can be repeated, often with a rapid „Brakes off/ Brakes on‟ modulation rate of up to 50 cycles/second.

Anti-skid Control Valve Figure 44 Page 13-42

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Anti-skid Control Valves Figure 45 B1 Mod 11.13 Issue No

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13.19.2.3 ANTI SKID CONTROL UNIT

This contains all the electrical circuits necessary for full anti skid control and circuits for BITE and monitoring of control valves and transducers. Circuits for a typical aircraft having four mainwheels ( Boeing 737) are normally arranged into two separate channels, for inboard and outboard pairs of wheels. As we have seen, skid control for each individual wheel requires a self-generated signal from its wheelspeed transducer.

Two types of anti-skid control unit Figure 46 The system has three modes of operation:  Touch-down protection, which prevents landing with brakes on.  Skid-control, which makes sure there is maximum braking efficiency.  Locked-wheel protection, which prevents any wheel from locking up due to the runway condition. When the aircraft is airborne, an „in-flight‟ signal is sent to the Control Unit via the ground/flight switch relay. The signal is sent to the „Touchdown protection mode‟ circuit, causing a full brake release signal (full dump) to be sent to the skid control valves. This prevents any pressure from going to the brakes and ensures that all brake units are always connected to return, even if the brake pedals are fully depressed. The full dump signal will be removed on touchdown when the „inflight‟ signal is replaced by a „on-ground‟ signal. The „Skid control mode‟ will not commence until the wheels have spun up to predetermined speed for the particular aircraft concerned. (Examples are; 30kts and 70kts for Fokker 50 and Boeing 737 respectively). Brake pressure is now controlled by modulation of the antiskid control valves as previously described. In addition to the „Skid Control mode‟ which ensures maximum braking efficiency for the level metered from the flight deck, the „Locked-wheel protection mode‟ circuitry looks at the inboard and outboard pairs of wheels and compares their speed. Should one of the pair slow down to a pre-determined difference, a full dump signal will be momentarily sent to the slower wheel in order to restore equilibrium. (Examples are; 30% and 40% difference for Fokker 50 and Boeing 737 respectively). Below about 15kts this mode is switched off (drops out) but „Skid control mode‟ remains. Page 13-44

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13.19.2.4 CONTROL SWITCH AND WARNING SYSTEM

This is normally located on the front panel of the flight deck and is often combined with the auto brake selector, if applicable to the particular aircraft type. It usually consists of a simple on/off switch to power up the anti skid circuitry. It also contains a warning light to give a warning of system malfunction. Following illumination of this warning light, it is possible to interrogate the Anti Skid Control Unit and pinpoint the cause – for example a particular transducer, valve or the control unit itself.

Flight deck control panel for anti-skid and autobrake Figure 47


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engineering 13.19.3 Mechanical Anti Skid System

This older type system modulates brake pressure as with the electronic type, but the modulation is achieved mechanically by a single self-contained device, one for each wheel. The device, often referred to as a „maxaret‟ (maximum arresting) unit, detects a rapid deceleration of the wheel and momentarily releases the brake pressure as before. It will normally be mounted externally on the brake unit torque plate and driven by a small rubber tyred wheel in contact with the aircraft mainwheel. Alternatively, it can be mounted inside the axle and driven by the aircraft mainwheel via a splined drive shaft in the hub cap. Both types, wheel or axle mounted, incorporate an internally mounted heavy flywheel, sensitive to the angular deceleration that occurs when braking. When the braking is severe or just before the wheel is about to lock up, the flywheel is permitted to continue rotating at the higher speed due to its inertia. It will advance through an arc until it contacts a set of limit stops. The flywheel, is connected mechanically to two hydraulic system metering valves within the maxaret unit. Using a pair of thrust balls and push rods the valves change their position from the normal „pressure to brakes‟ position, to the „no pressure in and brakes to return‟ position. With brake pressure removed, the wheel regains speed and the flywheel returns to its original position assisted by a return spring. The brakes are re-applied and the „brakes on/ brakes off‟ sequence will continue until the deceleration returns to normal limits. BRAKE UNIT

MAXARET DRIVE WHEEL MAIN WHEEL SPRING TO PUSH DRIVE WHEEL ON TO MAIN WHEEL

Mechanical anti-skid unit Figure 48 Page 13-46

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engineering To brake

Base of Cam Pressure supply

Thrust plate

Thrust rod

A. - Normal Braking Condition From return brake

profile Wheel rim

Base of Cam profile 60

Pressure supply

o

B. - Anti-Skid Condition Operation of Rim-Driven Unit Figure 49 Main shaft Valve spring

Flywheel

Drive ring

Clutch friction pad

Thrust bearing Input shaft

Clutch cover Clutch plate spring Sun gear RingClutch gear Drive spring Planet gear

Valve

Main spring Valve thrust rod

Axle mounted anti-skid unit Figure 50 B1 Mod 11.13 Issue No

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uk engineering 13.20


AUTOBRAKING

Some modern aircraft have auto-braking systems. A selector switch on the instrument panel allows the pilot to select a deceleration rate that will be controlled automatically after landing. On landing the auto-braking system will smoothly apply the brakes to achieve the selected deceleration rate down to a complete stop without any further action from the aircrew. This allows the aircrew to concentrate on other activities during landing. The auto brake system utilises the normal anti-skid and brake units but instead of using pressure from the brake metering valve, hydraulic pressure is sent via solenoid valves which allow a pre-determined amount of pressure through the anti-skid valves to the brake units. 13.20.1 Selector Panel

The selector panel consists of a solenoid latched switch which will hold a selected position only if all the arming conditions for that setting are met. If the system cannot be armed the switch will automatically return to the DISARM position and a warning will illuminate on the local panel and centralised warning panel. The panel will have a number of settings that the pilot can select depending on the rate of deceleration that is required. 13.20.2 Auto-Brake Control Unit

Selection on the auto brake selector panel will send an electrical signal to the auto-brake control unit. The signal is processed by the control unit, which commands the solenoid valve to direct pressure to the brake units. The brake pressure must be gradually built up and released to prevent brake snatch and jerking. To prevent this a time delay and an electrical ramp are used. The time delay ensures that the aircraft is firmly on the ground before the system activates. The terminology used to indicate the auto-brake operation is: 

On Ramp – A gradual build up of brake pressure to the amount required for the selected rate of deceleration.  Off Ramp – A gradual decrease in pressure down to zero at the end of the landing run or cancellation of auto-brake.  Drop Out – Instantaneous pressure release to zero (go around mode).

Autobrake On-Ramp Figure 51 Page 13-48

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13.20.3 Auto Brake Solenoid Valve

These valves are electrically controlled, hydraulic valves that allow pressure to the brake units at a specific setting. The greater the deceleration rate the higher the setting. These valves are fitted just upstream of the anti-skid valves. The solenoid will open when all the arming conditions are met and the aircraft is weight on wheels. It is also the solenoid valves that immediately shuts on Drop Out. A solenoid servo valve modulates the brake pressure to regulate the deceleration rate. A pressure switch is connected to the DISARM warning light to monitor zero pressure when auto-brakes are armed. 13.20.4 System Operation

Once the aircraft lands and is weight on wheels the anti skid transducers send signals to the control unit. When the wheels have achieved a certain speed or after a pre-determined time delay the brakes will be applied “Up The Ramp”. Once the selected rate of deceleration is reached the auto-brake pressure is modulated to hold that rate. As the wheel speed slows down to more than the deceleration rate, the servo valve will close slightly reducing the brake pressure causing the wheel to speed up. Once the aircraft has come to a stop or the aircraft is below a certain speed the auto-brakes will switch off to enable the aircraft to taxi. 13.20.5 Auto Brake Termination

Auto-Brake can be cancelled at any time. Depending on the aircraft, the system can be over-ridden by: 

The pilot moving the selector lever to disarm or off.



The pilot using manual braking.

Auto-brake needs to be immediately cancelled if the pilot has to initiate a go round procedure. The following actions will cause immediate DROP OUT: 

The thrust levers are advanced from the idle gate.



The speed brake lever is moved to stow the speed brakes.


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engineering 13.21

STEERING

To improve the ground operation of aircraft nose wheel systems are used. These improve tyre life through less scrub, reduce brake wear, save fuel and engine life as brakes and engine thrust are no longer required to turn the aircraft. Most nose wheel steering systems use servo jack operated scissor links attached to a collar on the landing gear leg, the collar being driven by the servo jacks which rotates the nose wheel leg via the scissor links. Steering inputs to the servo jacks come from a tiller on the pilots side of the cockpit. Inputs can also come from the rudder pedals. Apart from mechanical steering systems there are three basic methods of operation: 

Single Servo Jack.

This system is used on smaller light aircraft (Figure 52). Both ends of the jack ram are attached to the landing gear leg. Fluid is directed to move the jack body along its ram. A cam and link assembly is attached to the jack body. Movement of the jack body operates the link which rotates the cam and turns the wheel. Action of the shock absorber is unaffected as the shock absorber is splined on to the steering shaft to allow the compression and extension of the absorber. JACK BODY PISTON LINK

CAM

SPLINED SHAFT

STRUT

AXLE

Single Servo Steering Figure 52 Page 13-50

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engineering 

Double Servo Jack.

Larger aircraft use a two servo jack system (Figure 53). The two jacks are fixed to a steering collar, which is free to rotate around the landing gear leg. The steering collar is attached to the upper scissor link. When the servo jacks are actuated they rotate the wheels and axle through the scissor link. assembly

Double Servo Jack Figure 53 

Rack and Pinion

Some aircraft use a rack and pinion steering system. Hydraulically operated racks rotate a pinion which rotates the wheel and axle. A mechanical linkage from the cockpit tiller operates a servo valve in a hydraulic metering valve. The servo valve when operated directs fluid to one side or the other of the rack piston. The rack then moves and rotates the pinion and turns the aircraft nose wheel in the required direction. B1 Mod 11.13 Issue No

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engineering 13.21.1 Steering Mechanisms

On some small aircraft the nose wheel is steered by direct linkages from the rudder pedals, or on small retractable landing gear aircraft, from the rudder pedals to a steering bar which locates against a steering arm on the landing gear leg. (Figure 54) Once the wheel is stowed the mechanism is ineffective.

STEERING BAR

STEERING ARMS

STEERING LUGS

Nose Wheel Steering Mechanism Figure 54 The nose wheels or tail wheels on light aircraft mat be steerable or castoring. A castoring nose wheel aircraft is steered by the independent use of the brakes and rudder inputs. Some light aircraft have limited tail wheel steering via a mechanism interlinked with the rudder pedals. The tail wheel will brake out if the turning circle is too small to allow the tail wheel to castor. Once centralised the tail wheel becomes steerable again. Some aircraft have a tail skid mechanism (Figure 55)

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Tail Skid Mechanism Figure 55 Inputs to the hydraulic control valves which direct pressure to the steering jacks are carried out by a mechanical system of cables, bellcranks, levers and gearboxes from the hand operated tiller and the rudder pedals. The input has a follow up action through interconnected links or cables which neutralise the nose wheel movement when the desired rate of turn has been achieved. Rudder pedals movement can also be inputted to the control valve, but this is usually restricted to a small degree of movement either side of the aircraft centre line. Rudder pedal steering is normally used on take off or landing and is isolated when the aircraft is airborne. 13.21.2 Nose Wheel Self Centring

It is important that when a steerable nose wheel is being retracted that the wheel is centred so that it fits into the wheel well to prevent any damage to the aircraft structure as well as the landing gear. This can be done by a centring cam inside the oleo strut. When the strut is compressed the piston cam disengages from the cylinder cam receptacle to allow the wheel to be steered. On take off when the strut extends the piston cam is forced into the cylinder receptacle to hold the wheel in the desired position for stowing. Double servo jacks can centralise the wheel by supplying pressure to a centralising jack. This is normally initiated by the weight-on-wheels micro-switches as the aircraft takes off.


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uk engineering



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B1 Mod 11.13 Issue No 01/02/11

JAR 66 CATEGORY B1

MODULE 11.15 OXYGEN

CONTENTS 15 OXYGEN ....................................................................................... 15-1 15.1 15.2 15.3

15.4

15.5

15.6

15.7

OXYGEN SYSTEMS GENERAL ................................................ 15-1 OXYGEN SAFETY PRECAUTIONS.................................................... 15-1 SYSTEM LAYOUT .......................................................................... 15-2 15.3.1 Cockpit System Layout ................................................. 15-2 15.3.2 Cabin System Layout .................................................... 15-4 15.3.3 Continuous Flow Oxygen System ................................. 15-5 15.3.4 Demand Type Oxygen System...................................... 15-6 15.3.5 Portable Oxygen Systems ............................................. 15-7 DROP OUT SYSTEM ...................................................................... 15-8 15.4.1 Pneumatically Operated PSU Flap ................................ 15-8 15.4.2 Electrically Operated PSU Flap ..................................... 15-9 SOURCES OF OXYGEN .................................................................. 15-9 15.5.1 Chemical Oxygen Generator ......................................... 15-9 15.5.2 Gaseous Oxygen Systems ............................................ 15-11 15.5.3 Charging Of Systems .................................................... 15-11 15.5.4 Oxygen Distribution ....................................................... 15-11 SUPPLY REGULATION ................................................................... 15-12 15.6.1 Diluter Demand Type Regulator .................................... 15-12 15.6.2 Continuous Flow Regulators. ........................................ 15-12 INDICATIONS AND WARNINGS ......................................................... 15-13


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JAR 66 CATEGORY B1 MODULE 11.15 OXYGEN


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JAR 66 CATEGORY B1

MODULE 11.15 OXYGEN

15 OXYGEN 15.1 OXYGEN SYSTEMS GENERAL If an aircraft is designed to fly at heights above, say, 8,000 feet, there must be some way in which we can maintain a comfortable environment for the crew and passengers to breathe normally. This is normally done by cabin pressurisation. If for whatever reason the pressurisation failed above this altitude an alternate but emergency source of breathable air must be supplied. This is normally by individual oxygen supplies from gaseous, liquid and chemical sources. Civil aircraft use the gaseous and chemical type, with the military using liquid. Some small, unpressurised aircraft only require oxygen occasionally and use a system that meters a continuous flow of oxygen; the amount based on the altitude flown. Aircraft that fly at altitudes above 18000 feet have a diluter demand system that also meters oxygen based on the altitude flown but directs it to the mask only when the user inhales. Aircraft flying at very high altitude where the outside air pressure is too low to force the oxygen into the lungs use pressure demand systems. These systems send oxygen to the mask under a slight positive pressure that forces the oxygen into the lungs. 15.2 OXYGEN SAFETY PRECAUTIONS The safety precautions associated with the use of oxygen are laid down in the aircraft maintenance manuals. Although oxygen is none flammable it will support combustion. If oil grease dust or metal particles are present a spontaneous explosion may occur. The following safety precautions must be adhered to: 1.

Keep oil and grease away. Oxygen equipment, hoses and fittings must not be handled with greasy hands or wearing greasy overalls.

2.

Keep oxygen away from fire. A small fire or spark will rapidly grow in an oxygen-enriched atmosphere.

3.

No smoking.

4.

Handle oxygen components carefully.

5.

Don’t mix oxygen

6.

Always follow any instructions given in manuals and/or on charging panels.


Page 15-1


When charging a gaseous system ensure: 1.

No refuelling operations are being carried out.

2.

No switching on or off electrical supplies.

3.

Adequate warning notices are in place i.e. oxygen charging in progress

4.

That there is no smoking or naked flames.

5.

That the aircraft is earthed

6.

That adequate fire fighting equipment is available.

15.3 SYSTEM LAYOUT The crew and passenger gaseous oxygen systems and their oxygen cylinders are usually independent of each other except for a common charging point and an over pressure relief facility. Both these systems provide for storage of the oxygen at high pressures and its delivery to the crew and passenger manifolds and outlets under low pressure. In general gaseous oxygen systems are used for the cockpit and chemically generated oxygen is used for the cabin. Some aircraft use gaseous systems for both the cockpit and cabin. 15.3.1 Cockpit System Layout The aircrew will have a mask for each occupant. These will be quick fitting and will be located inside boxes that are within easy reach. Figure 1 shows a typical cockpit layout

Typical Cockpit layout Figure 1

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MODULE 11.15 OXYGEN



Crew Mask

Crew oxygen masks contain a microphone implanted in the mask that is permanently connected; to allow communications to be maintained at all times. The end of the mask hose is connected to the supply regulator that regulates the oxygen flow to the mask. On some aircraft an inflatable harness is used to allow one handed fitting of the mask. The mask and harness is contained in a storage box shown in figure 2. When the mask is required the storage box release levers are squeezed together. The box doors are unlocked and the mask is withdrawn. A green oxygen on flag will appear. On withdrawal of the mask, the harness is automatically inflated. Once the mask and harness has been fitted the release levers are released and the oxygen that has inflated the harness is exhausted to atmosphere. The harness deflates and tightens on the crew members head. The storage box contains a test lever that can be operated to test the oxygen flow to the mask. When the system is operating correctly a blinker indication on the flow indicator turns green. There will also be a 100% selector button which when depressed will allow pure undiluted oxygen to be delivered to the mask.

Typical Storage Box and Mask Harness Figure 2


Page 15-3


15.3.2 Cabin System Layout On gaseous oxygen systems a ring main is provided from the storage bottles to the PSU`s. In chemically generated oxygen systems an oxygen module is located in each PSU. A typical layout is shown in figure 3.

Candle Type Oxygen Module Figure 3 

Passenger Mask

The passenger masks will be found within the Passenger Service Unit (PSU) and will be deployed by gravity on actuation of the drop out mechanism. Each seating position both in the cabin must have an easily fitted mask, which will be used by each occupant. Some aircraft do not have a mask for each person but have strategically placed masks in the PSU for the passengers to share. Some aircraft do not have “drop out” systems and the masks may have to be deployed manually by the cabin crew.

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MODULE 11.15 OXYGEN

These are normally simple cup shaped mouldings with an elasticated strap. The cup is designed to fit all sizes from babies to adults. A reservoir bag is fitted to the mask to store an immediate supply of oxygen.

Passenger Oxygen Mask Figure 4 15.3.3 Continuous Flow Oxygen System Continuous flow systems are usually used in passenger oxygen systems where oxygen is needed only occasionally. These systems are wasteful of oxygen but due to their simplicity are installed on most aircraft. The oxygen is carried in a high-pressure bottle. The pressure is regulated down to around 400 psi (depends on aircraft type) by a pressure reducing valve and the oxygen is metered by a pressure regulator to around 70 psi before it is delivered to the masks.


Page 15-5


A pressure relief valve is incorporated into the system to prevent damage in the event of a failure of the pressure-reducing valve. If the pressure is relieved through this valve a green blow out disc on the outside skin of the aircraft will blow giving a visual indication. As well as a visible blow out disc some aircraft also deploy a red streamer in an over pressure condition. 

Continuous Flow Masks

Continuous flow oxygen systems use re-breather type masks. These masks may be a simple transparent plastic re-breather bag. The mask is held loosely over the mouth and nose with an elastic band and oxygen continually flows into the bag through a plastic tube that is plugged into the mask outlet. When the user exhales the air that was in the lungs for the shortest period is the first out and fills the re-breather bag. The remaining air in the lungs is exhausted from the mast. Inhaling again the exhaled air in the bag is enriched with the oxygen supply and is re-breathed. More sophisticated continuous flow masks are used in pressurised aircraft. In the event of the loss of cabin pressure an automatic valve is turned on to send oxygen to into the passenger oxygen system. The oxygen pressure actuates the door actuator valve, which opens the overhead mask compartments. A mask drops down. When the passenger pulls the mask tube a lanyard operated rotary valve opens and starts the oxygen flow. The passenger places the mask over his mouth and nose and breathes normally. Valves mounted in the base plate of the mask allow some cabin air to enter the mask and allow exhaled air to leave. During inhalation the pure oxygen in the bag is taken into the lungs. When the bag is empty cabin air is taken in through the mask and mixes with the oxygen flowing through the tube. During exhale the air from the lungs leaves the mask through one of the valves while pure oxygen is flowing from the regulator into the bag ready for the next inhale. 15.3.4 Demand Type Oxygen System The cockpit crews of most commercial aircraft are supplied with oxygen through a diluter demand system. The system meters oxygen only when the user inhales and the amount of oxygen metered depends on the altitude of the aircraft. Almost all pressurised aircraft have a diluter demand type system for the aircrew and a continuous flow type system for the passengers. 

Pressure Demand Oxygen Regulator.

At altitudes above 40000 feet the oxygen in the air has such a low pressure that even the pure oxygen supply must be forced into the lungs. The low pressure from the users lungs are insufficient to draw in the oxygen. This is done under a slight positive pressure from the regulator.

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15.3.5 Portable Oxygen Systems At various positions in the cabin, there are located, portable oxygen sets for use by the cabin crew to allow them to check that the passengers have all got their masks on. These masks can also be used to help breathing in the case of fumes or smoke in the cabin. A slightly different type of oxygen set can also be found in the cabins of most passenger aircraft. These are called 'therapeutic' sets and are used for medical purposes when, for example, a passenger is having difficulty breathing. The set, which is illustrated overleaf, allows the passenger to receive an enriched or 100% oxygen supply, until they are feeling better or medical assistance is obtained after landing. A typical portable system is shown in Figure 5. It has two outlets, which might be therapeutic. A therapeutic outlet delivers more volume of oxygen than normal and is used to aid those passengers that may have breathing difficulties or heart conditions. The types of mask used with portable equipment, depends upon the designer's or the company's requirements.

Portable Walk Round Set Figure 5


Page 15-7


15.4 DROP OUT SYSTEM Drops out oxygen masks in each PSU are installed to ensure that there is an adequate supply of oxygen should the aircraft conditions require it. When the mask drops from the PSU the action of the passenger pulling the mask towards him automatically starts the oxygen flow. If there were not many passengers on board this would mean that only those masks that are pulled will have an oxygen flow preventing both excess oxygen waste and an oxygen enriched environment. The PSU is a hinged flap containing on its under face, reading lights, cold air vents, warning signs and cabin crew alert button. The masks are stored inside the flap panel (under the overhead bins). The masks will drop under the following conditions: 

Automatically when the cabin altitude reaches a pre-determined level (usually around 10000 feet).



When the aircrew selects oxygen. Drop out could be actuated electrically, pneumatically or mechanically.

In the case of a chemically generated supply the PSU flap is opened either electrically or mechanically. 15.4.1 Pneumatically Operated PSU Flap On the pneumatic door opening method a small plunger is fitted above each PSU flap. The doors are held closed by a spring loaded latch assembly. When oxygen is required and selected, oxygen pressure is directed to the over centre leaf spring and to the plunger. The pressure extends the plunger that pushes against and overcomes the latch assembly. The PSU flap opens under gravity deploying the oxygen masks in the process. Giving a sharp pull on the mask the flow control pin is withdrawn from its locating hole. The oxygen pressure overcomes the over centre leaf spring and directs oxygen to the mask. Should the cabin crew require to turn off the oxygen at each PSU a manual closing toggle is rotated which acts against the over centre leaf spring and cuts off the flow to the mask. If the PSU flap fails to open a manual actuation pin can be pushed to allow the flap to open.

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MODULE 11.15 OXYGEN

15.4.2 Electrically Operated PSU Flap When chemical oxygen generators are installed for the passengers to use, oxygen is not supplied until the masks are pulled. Therefore they cannot use the oxygen generated to open the PSU Flap. The flap is opened by an electrical solenoid. When oxygen is selected power is supplied to the solenoids which when energised operate a plunger. This plunger extends and operates a latch assembly. Operation of the latch assembly opens the PSU door and the masks fall under gravity. 15.5 SOURCES OF OXYGEN Most aircraft use gaseous oxygen as the primary source for the aircrew and a chemically generated source for the passengers. Some aircraft with oxygen generators fitted are having them replaced with a gaseous oxygen system due to the associated fire hazards. The main reason for using gaseous oxygen is the ease of handling and its availability at most airports, even though oxygen generated systems are more lightweight. The main disadvantage of a gaseous oxygen system is that the oxygen is stored at a high pressure, it reacts explosively with greases and oils and the storage bottles are very heavy. Bottles are made from high tensile steel, but on the more modern aircraft, Kevlar wrapped aluminium alloy, carbon fibre or plastics are used. They are painted either black with a white dome top or green (USA), and stencilled with Aviation Oxygen in white letters. 15.5.1 Chemical Oxygen Generator Oxygen generators or oxygen candles as they are sometimes known, is a convenient way to carry oxygen in an aircraft, when it may only be required in emergencies. They have a long shelf life and they are lightweight. The storage capacity is about three times that of a gaseous oxygen system. Once used they are easily replaced. A typical candle is shown in Figure 6. Sodium Chlorate and iron is mixed with a binding material and is then moulded into a solid block. The block is installed in an insulated stainless steel case. When oxygen is needed, pulling the oxygen mask withdraws a safety pin from the firing mechanism, and a spring-loaded percussion cap or an electrical squib igniter starts the sodium chlorate decomposing by chemical reaction. Enough heat is generated to start the reaction and then the heat of the reaction sustains itself. (It does not burn). As it decomposes it releases the oxygen at a pre-determined rate. The block will continue to react until the sodium chlorate is consumed. There is no way to cut off the process once it has started. The by product of the reaction, apart from the oxygen, is sodium chloride (salt) and ferrous oxide (rust)


Page 15-9


HEAT SHIELD

PERCUSSION CARTRIDGE


FILTER

FIRING PIN

DISTRIBUTION BLOCK IRON AND SODIUM CHLORATE CORE

ACTIVATION PIN

OUTLETS TO MASKS

LANYARD

Chemical Oxygen Generator Figure 6 The oxygen that is produced is proportional to the cross sectional area of the core and the rate of reaction. The rate of decomposition is determined by the concentration of iron in the core. The oxygen production is greater at initial reaction (larger cross sectional area) to provide high oxygen output during the initial few minutes of the emergency decent. Once generation has started core temperature is approximately 450 degrees F. The distributing and regulating system is self-contained. It consists of a manifold attached to one end of the stainless steel cylinder. The oxygen is filtered to remove any salt particles before it is supplied to the manifold. The manifold contains calibrated connections for a number of oxygen masks and they ensure an equal flow to each mask. Normal output from the generator is 10 psi and it is therefore not regulated prior to breathing. A pressure relief valve is also located on the casing to relive pressures in the generator above 50 psi. The disadvantage of the system is mainly the large amount of heat generated, which means that the generator must be well insulated from the airframe structure. Some aircraft that use oxygen generators are replacing them with gaseous oxygen systems due to associated the fire hazards.

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MODULE 11.15 OXYGEN

15.5.2 Gaseous Oxygen Systems Oxygen in a gaseous state is contained in storage cylinders the number and capacity of the cylinders depending upon the number of passengers and crew. The normal charge of the cylinders is usually 1800 psi and a capacity of 30 to 120 cubic feet. Cylinders normally have a manually operated shut off valve in the neck of the bottle to facilitate bottle removal. A direct reading pressure gauge is also fitted, as is an electrical transducer that sends pressure indication signals to the cockpit instrumentation. 15.5.3 Charging Of Systems Gaseous systems can be re-charged either at the aircraft, from portable, large capacity, bottle sets (oxygen trolley), or by removing the bottle itself, via quick release clamps and connections, and re-charging in a dedicated oxygen charging room or bay. Which system is used is dictated by the Airworthiness Authority of the country of registration, some of which allow 'on aircraft' charging, whilst most insist that all bottles are removed for re-charging remotely from the aircraft. With “on aircraft” charging, a regulated oxygen trolley supply is attached to the aircraft charging point. The hose is usually purged before connection to clear the hose of impurities and moisture. During the charging process temperatures are generated in the pipelines to the storage bottles. To dissipate this temperature thermal compensators are installed. These compensators are sintered bronze elements soldered inside the pipelines. They act as a heat sink to dissipate the heat whilst allowing the oxygen to flow to the storage bottles. With chemical oxygen systems, when the units themselves become life expired and due for return to their manufacturers, they are simply removed as a unit from the PSU. When they have been made safe, (usually by the fitting of a safety pin in the firing sear), they are returned to the factory. 15.5.4 Oxygen Distribution The supply pipes, in the high-pressure side of the system, from the storage bottle to the pressure-regulating valve, are made from stainless steel or copper based alloys. They are colour coded at each end with the words “breathing oxygen” and a black rectangular symbol on a white background. From the storage bottle the pressure is reduced to an acceptable level before being distributed to the passenger and crew compartments. As the maximum pressure to the masks will be 70psi, the distribution pipelines, from the pressure regulator valve are made from aluminium alloy or plastic.


Page 15-11


The distribution lines for the aircrew go from the storage bottle to the cockpit pressure regulator and the passenger lines go from the storage bottle, up the side walls and then along the roof. Each passenger service unit (PSU) where the masks are stowed are connected to the roof piping. Test connections are installed in the system to allow pressure gauges to be fitted during system testing. 15.6 SUPPLY REGULATION 15.6.1 Diluter Demand Type Regulator Oxygen flows into the regulator through the supply valve and when the user inhales the pressure inside the regulator decreases and the demand valve opens under action from the demand valve diaphragm allowing oxygen to flow to the mask. The aneroid capsule operated metering valve mixes cabin air with the oxygen. When the aircraft is flying at low altitudes the user gets mostly cabin air and a small amount of oxygen. As the altitude increases the aneroid capsule metering valve progressively reduces the amount of cabin air and increases the amount of oxygen supplied. At about 34000 feet the cabin air is shut off completely and pure oxygen is supplied. If there is smoke in the cockpit or if the pilot feels the need for pure oxygen the oxygen lever can be moved to the 100% position. The cabin air is shut off and the aneroid metering valve fully opens and only pure oxygen is supplied to the mask when the user inhales. If the regulator malfunctions the emergency lever can be operated. This opens the demand valve allowing a continuous flow of pure oxygen to the mask. 15.6.2 Continuous Flow Regulators.

There are automatic and manual continuous flow regulators. The automatic regulator contains and aneroid capsule that senses the aircraft altitude and meters the correct amount of oxygen for that altitude. The manual regulator has a control valve that allows the pilot to adjust the flow rate based on the altitude. A calibrated orifice in the mask outlet determines the amount of oxygen delivered to the mask.

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MODULE 11.15 OXYGEN

15.7


The systems are provided with an overpressure relief facility. This is normally a green coloured rupture disc. The disc will be located at the overboard discharge fitting which is flush with the aircraft’s skin. When the maximum cylinder pressure is exceeded the cylinder safety valve operates discharging the excess pressure into the overboard discharge line. The green disc ruptures, as the excess pressure escapes to atmosphere and a red (or yellow) indicator becomes visible. Some aircraft also deploy a red streamer from the fitting to make it instantly visible. On aircraft with oxygen generators fitted, once the generator has been activated, dolls eye indicators on the end casing turn from orange (or purple) to black. Some have heat sensitive tape wrapped around the outer casing. The tape changing colour when the generator has been activated. There will be various indications given on the oxygen panel and centralised warning panel (CWP) indicating faults within the system. A pressure gauge will be fitted which shows the pressures in the storage bottles. The gauge will have a green segment and a red segment. The green segment indicates the actual pressure in the system. The red segment will indicate that the bottle is empty or maybe that the shut off valve is closed. A low pressure switch is fitted in the system downstream of the storage bottle and will give an indication (LO PR) on the local panel if the pressure reduces below a pre-set figure. An associated (OXY) caution light will illuminate on the CWP and a single chime warning will also sound.

Indications and Warning Panels Figure 7


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MODULE 11.16 PNEUMATIC AND VACUUM

CONTENTS 16 PNEUMATIC AND VACUUM ....................................................... 16-3 16.1 16.2 16.3 16.4 16.5

GENERAL ..................................................................................... 16-3

SAFETY PRECAUTIONS ........................................................... 16-3 FULL PNEUMATIC SYSTEMS ................................................... 16-3 VACUUM SYSTEMS ........................................................................ 16-5 LOW PRESSURE PNEUMATIC SYSTEMS LAYOUT............................. 16-5 16.5.1 Engine Driven Air Pump ................................................ 16-5 16.6 AIR SUPPLY SOURCES .................................................................. 16-6 16.6.1 Engine Bleed Air ........................................................... 16-7 16.6.2 Compressors or Blowers. .............................................. 16-8 16.6.3 Auxillary Power Unit (APU) ........................................... 16-8 16.6.4 Ground Supply .............................................................. 16-9 16.7 PRESSURE CONTROL .................................................................... 16-10 16.7.1 Pressure Regulator ....................................................... 16-10 16.8 DISTRIBUTION ............................................................................... 16-11 16.8.1 Expansion Joints ........................................................... 16-12 16.9 INDICATIONS AND WARNINGS......................................................... 16-14 16.9.1 Overpressure ................................................................ 16-14 16.9.2 Overheat ....................................................................... 16-14 16.9.3 Duct Hot Air Leakage .................................................... 16-15 16.10 SYSTEM INTERFACES .................................................................... 16-15 16.10.1 Pneumatic Gyro Power systems ................................... 16-15 16.10.2 Backup High Pressure Pneumatic Systems .................. 16-16 16.10.3 Pneumatic De-Icing systems ......................................... 16-16 16.10.4 Air Conditioning And Pressurisation. ............................. 16-16 16.10.5 Air Driven Hydraulic Pumps. ......................................... 16-17 16.10.6 Pressurising Of Hydraulic Reservoirs. ........................... 16-17 16.10.7 Waste And Water Systems ........................................... 16-17 16.10.8 Pneumatic Stall Warning ............................................... 16-18


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16 PNEUMATIC AND VACUUM 16.1

GENERAL

Pneumatic systems are fluid power systems that use a compressible fluid, air. These systems are dependable and lightweight and because the fluid is air there is no need for a return system Some aircraft have only a low pressure pneumatic system to operate the gyro instruments, others use compressed air as an emergency backup for lowering the landing gear and operating the brakes in the case of hydraulic failure. Other aircraft have a complete pneumatic system that’s actuates the landing gear retraction, nose wheel steering, passenger doors and propeller brakes. 16.2 SAFETY PRECAUTIONS When working on bleed air systems, it is important to follow the precautions below: 

Bleed air is hot! Do not touch pipes and ducts.



Always replace seals, (normally crush seals), when replacing joints.



Tighten clamps to the torque figure quoted in the Maintenance Manual.



Never lever against ducts, as dents cause hot spots.



All duct supports and struts must not put any strain on to the duct.

16.3 FULL PNEUMATIC SYSTEMS The majority of aircraft use hydraulic or electrical power to operate landing gear systems, but some aircraft use air systems. Some advantages of using compressed air are: 

Air is universally available and in unlimited supplies.



Pneumatic system components are reasonably simple and lightweight.



No return lines are fitted resulting in a weight saving.



There is no fire hazard and the danger of explosion is slight.



Contamination is minimised by the use of filters.


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Figure 1 shows a typical high pressure pneumatic system, that uses air compressors driven from the engines accessory drive. The compressed air is discharged through a bleed valve to a pressure relief (unloading) valve. The bleed valve is held closed by oil pressure. In the event of oil pressure failure the bleed valve opens to offload the compressor. The pressure relief valve maintains system pressure at around 3000 psi. A shuttle valve in the line between the compressor and the main system makes it possible to charge the system from a ground source. When the engine is not running the shuttle valve slides over to isolate the compressor. SHUTTLE VALVE

GROUND CHARGING POINT

PRV BLEED VALVE WATER SEPARATOR

BLOW OUT DISC

AIR PUMP DESICCANT

NRV FILTER

EMEREGENCY BRAKE SYSTEM

PRIMARY AIR BOTTLE

ISOLATING VALVE EMERGENCY LANDING GEAR TO NORMAL SERVICES OFF

GAUGE PRV

AIR BOTTLE

PRESSURE REDUCING VALVE

A Typical Pneumatic System Figure 1 Moisture in a compressed air system will freeze as the air pressure drops when a component is actuated. To prevent this from happening, the water must be completely extracted from the air. A water separator is fitted which collects the moisture from the air onto a baffle and it is allowed to drain overboard. An electric heater prevents the water in the separator from freezing. After the air leaves the water separator any remaining moisture is removed as the air flows through a desiccant or chemical dryer. The air is then filtered before it enters main system.

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The air is then fed to each of the storage bottles, which provide the emergency air for several systems. A manually operated isolation valve allows the air supply to be shut off to so that maintenance can be carried out on the systems without having to discharge the storage bottles. The air is stored at maximum system pressure around 3000 psi to supply the landing gear and brakes in an emergency. A pressure reducing valve is fitted to reduce the air pressure down to the operating pressure that the majority of the components work at 9around 1000psi) ie landing gear normal operation, the passenger door, the propeller brake and the nose wheel steering. 16.4

VACUUM SYSTEMS

A supply of air at a negative pressure can be required for a number of purposes. The supply of vacuum to instruments for example, usually comes from either a small vacuum pump attached to the (piston) engine of the aircraft or from a venturi jet pump, which obtains its power via a tapping from the (jet) engine. The low pressure caused by the venturi draws in air to supply the system. Other requirements for a source of vacuum might be in a pneumatic de-icing system. This type of de-icing uses the inflation of flexible leading edge mats to break-off the ice, which has formed. To keep the de-icer boots, as they are called, in place, they are fed a negative pressure from a venturi, which ensures that the boots are sucked flat onto the wing leading edge, ensuring a smooth, aerodynamic surface. 16.5 LOW PRESSURE PNEUMATIC SYSTEMS LAYOUT These systems provide air for gyroscopic altitude and direction indicators and air to inflate the pneumatic de-icing boots. This compressed air is usually provided by a vane type engine driven air pump (Figure 2). 16.5.1 Engine Driven Air Pump On early aircraft engine driven air pumps were used primarily to evacuate the casings of air-driven gyroscopic instruments so they were more commonly known as vacuum pumps. On later aircraft the discharge air was used to inflate de-icing boots on control surfaces and are now more correctly called air pumps. There are two types of air pumps that are used, these are wet air pumps and dry air pumps.


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Vane Type Air Pump Figure 2 

Wet Air Pumps

Wet pumps have steel vanes that are lubricated and sealed with engine oil which is drawn in through the pump mounting pad and exhausted with the discharge air. This oil is removed from the discharge air with an oil separator before it is used for de-icing or driving the instruments. 

Dry Air Pumps

Dry air pumps were developed so that there was no oil in the discharge air and therefore there were no requirements for an oil separator. The pump vanes are made from carbon and are self lubricating. The main problem with this kind of pump is that the vanes are easily breakable by any contaminants that enters the pump. To prevent this form occurring the inlet air is filtered. 16.6 AIR SUPPLY SOURCES The source of air supply and arrangement of the system components depend on the aircraft type and system employed but in general one of the following methods may be used:

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16.6.1 Engine Bleed Air This is used in turbo jet aircraft in which hot air is bled of from the engine compressors to the cabin. Before the air enters the cabin it is passed through a pressure and temperature control system which reduces its pressure and temperature and is then mixed with ram air. Because of the great variation of air output available from ground to maximum flight rpm there is a need to maintain a reasonable supply of air during low rpm operation as well as restricting excessive pressures when operating at full speed. Two tappings are taken from the engine, one form the LP stages and one form the HP stages to maintain a reasonable pressure band at all engine speeds. Figure 3 shows a typical 2 stage bleed air system. At low engine rpm the LP air is of insufficient pressure for use in the pneumatic systems, so air will be tapped from the HP stages. When engine speed increases the LP air pressure will also increase and at a pre-determined pressure the HP air will be shut off and when operating at maximum engine speeds the air will be taken purely from the LP stages. In all normal stages of flight therefore the bleed air will come form the LP stages.

Typical Two Stage Bleed Air System Figure 3


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16.6.2 Compressors or Blowers. This is used by some turbo jet, turbo prop or piston engine aircraft, the compressors or blowers being either engine driven via an accessory drive, by bleed air or electric or hydraulic motors. The compressor inlet duct is connected to an air scoop and its outlet is connected to the pneumatic manifold. The unit is controlled by a shut off valve which is operated from the cockpit. When insufficient LP air pressure is available for the pneumatic systems at low engine speeds the aircrew will select the shut off valve to open. This will direct the LP air to drive the turbo compressor. A pressure regulator is incorporated to ensure a constant output at the required pressure. On large multi-engine aircraft only some of the engines will have a turbo compressor (Figure 4) which is normally mounted with its associated controls in an engine bay.

Turbo Compressor Figure 4 16.6.3 Auxillary Power Unit (APU) This provides an independent source of pressurised air. It is basically a small gas turbine engine that provides air and other service whilst the aircraft is on the ground with its main engines stopped. It is usually a self contained unit located in the tail section of the aircraft where it can be run safely (Figure 5). On some aircraft the APU can be started in flight and act as a back up source of air, hydraulics services in the event of a loss of an engine.

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Typical APU Setup Figure 5 16.6.4 Ground Supply For use on the ground when the engines are not running. This unit will run until the aircraft is independent of the trolley. The ground cart is basically a compressor driven by an engine, usually a diesel. The compressor output pressure is regulated to match the aircrafts system pressure. A quick release hose is connected from the cart to the aircraft service panel. The maximum aircraft systems pressure and operating instructions including safety precautions are detailed on the inside of the service access panel.

Ground Cart Control Panel Figure 6


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Instructions for operating the ground cart will be found on a panel on the carts control panel. Figure 6 shows a typical ground cart control panel. 16.7

PRESSURE CONTROL

In many bleed air systems the pressure is regulated only by the operation of the high pressure shut off valve. The range of pressure may be from 10psi at ground idle to 65 psi at take off power. Many modern aircraft use bleed air for many systems that are sensitive to pressure variations and therefore some form of regulation is required. The pressure regulator is a pneumatically operated valve which will give a predetermined output pressure form the engine bleed air system. The regulator may also perform as the shut off valve. This is then called a pressure regulating and shut off valve. 16.7.1 Pressure Regulator This valve operates on the principle of a balance between air pressure and spring pressures. Referring to Figure 7. Assuming the piston has an area of 1 square inch and is held in its seat by a spring that pushes with a 100 pounds force. The piston has a shoulder of 0.5 square inches and this area is acted on by a system air pressure of 1500psi. The cone shaped seat of the valve has an area of 0.5 square inches and is acted on by a reduced pressure of 200psi. A bleed orifice in the piston allows air pressure into the piston chamber. A relief valve being acted on by the reduced 200psi pressure and relief valve spring pressure, maintains the air pressure in the piston chamber at 750psi.


BLEED ORIFICE

PISTON PRESSURE IN

PISTON CONE TO SERVICES

Pressure Regulator Figure 7

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When the air supply is used by a pneumatic service, the reduced downline pressure of 200psi reduces further. This reduced pressure is now insufficient to keep the relief valve closed. The 750psi piston chamber pressure unseats the relief valve and reduces the piston chamber pressure. The reduced piston chamber pressure unseats the piston cone piston which allows the system pressure to bleed into the down lines. Once the downline pressure rises to 200psi, the piston cone and the relief valve re-seat and the system is once again in balance. 16.8

DISTRIBUTION

Distribution is achieved by ducting and pipelines that carry the charge air from the engine compressors to the various services that require air for their operation. Due to the heat of the bleed air any leakage of the ducts will cause an extreme temperature rise in the area of the leak with the possibility of fire or damage to the surrounding structure and equipment. Leak detection systems are therefore incorporated. Figure 9 shows a typical distribution layout.

Ducts Supports Figure 8 The ducting is made up of many sections for ease of maintenance and cheapness of replacement. They are constructed of thin wall material and clamped together with joints that allow for thermal expansion. Engine bleed air system ducts are manufactured from stainless steel and the ducts and pipelines are usually manufactured from titanium as they are able to withstand higher temperatures and are lighter in weight. The duct sections are supported throughout their length by clamps and tie rod attachments to the aircraft structure as shown in Figure 8.


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Bleed Air Distribution Manifold Figure 9 16.8.1 Expansion Joints Joints are assembled cold and when in use the temperatures int eh ducting can reach up tom 350 degrees F. Expansion devices must be incorporated into the systems to prevent any distortion or buckling of the ducts. This expansion can be allowed for in several ways. 

Pre-Stressed Joint

One method is to have the duct sections installed slightly shorter in length and allow them to expand with the heat to fit correctly. The ducts will be pre-stressed by the clamps when cold (Figure 10).

Pre-Stressed Joint Figure 10

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

Flexible Ball Joint

Another method is with a flexible ball joint fitting at the duct ends. The joint is designed to allow for slight flexing and misalignment as well as expansion. A flange on one end of the duct is connected to a bearing nut on the other and screwed together to form the joint (Figure 11). Shims are used to ensure adequate clearance is maintained for the expansion and flexing and a crush type metal seal is used to prevent air leakage at the joints.

Ball Joint Figure 11 

Cable Attachment Joint

The cable attachment type joint is used where large temperature changes exist, ie from cold soak at high altitudes to maximum working temperatures when the pneumatic system is selected on. This joint has bosses attached at each end of the duct. There are usually 3 short cables equally spaced around the duct (Figure 12). The cables have a swaged ball end fittings at one end and a swaged threaded fitting at the other. Each end is located in a bracket on the ducting. A seal is fitted around the duct before the ducts are connected. A nut is fitted on the threaded end and tightened. This pulls tightens the cables and seals the duct. A small gap is left at the seal ends to allow for expansion.


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Cable Joint Figure 12 16.9


Safety devices are fitted into pneumatic systems to prevent a possible overheat or overpressure which could cause severe damage to the air ducting or systems. 16.9.1 Overpressure Overpressure is usually caused by a malfunction of the high pressure shut off valve that remains open when the engine is operating at its maximum rpm. In most systems a pressure relief valve is fitted in the engine bleed air ducting which relieves excess pressures. The pressure relief valve may also work in conjunction with a pressure switch will close the high pressure shut off valve at a pre determined pressure. 16.9.2 Overheat Over temperature of the bleed air is prevented, by an electrical temperature sensor, downstream of the engine bleed air valve. When a pre determined temperature is reached the electrical sensor will signal the high pressure shut off valve to close. An overheat will be indicated to the aircrew on the CWP and associated control panel.

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16.9.3 Duct Hot Air Leakage Any ducting that includes joints is liable to leak under abnormal conditions. A duct protection system will include fire-wire elements around the hot zones such as engine air bleeds, air conditioning packs and auxillary power units if fitted. The sensing elements will be the thermistor type. As the temperature around the wire increases the resistance decreases until an electrical circuit is made. When the circuit is made a warning signal is sent to the cockpit central warning panel with associated caution/warning lights and aural chimes. The leaking duct may be isolated automatically or may require the pilot to take action to close off the air valves. The faulty system will then remain out of use. 16.10

SYSTEM INTERFACES

The pneumatic system interfaces with various other aircraft systems. Once the bleed air has been reduced in pressure to around 40 to 50 psi, most services have their own pressure and temperature controls, as well as generating their own warnings and indications to the CWP or system control panels in the cockpit. 16.10.1 Pneumatic Gyro Power systems The gyroscopes in pneumatic gyro instruments are driven by air impinging on cups cut in the periphery of the wheel. There are two methods of obtaining air to drive the instruments: 

Air Pump Suction

The air pump suction evacuates the instrument case and draws air in through a filter. The filtered air id directed through a nozzle and it strikes the driving cups to drive the gyro instrument. A suction relief valve regulates the suction to the correct value to drive the instrument and a suction gauge reads the pressure drop across the instrument. 

Dry Air Pump Pressure.

Since many aircraft fly at high altitudes where there is insufficient air pressure to drive the instruments another method must be used. The gyro instruments are driven by the air from the pressure side of a dry air pump. The air is filtered before it is taken into the air pump and is regulated before it flows through an in line filter to the instruments. After driving the instruments it is evacuated overboard.


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16.10.2 Backup High Pressure Pneumatic Systems On some aircraft, in case the hydraulic systems fail there must be provision for an emergency extension of the landing gear and application of the brakes. The system comprises of a pressurised cylinder which contains approximately 3000psi of compressed air or nitrogen. A shuttle valve (Figure 13) in the actuator line directs hydraulic fluid to the actuator for normal operation or compressed air/nitrogen for emergency operation.

PISTON PISTON

EMERGENCY AIR

HYDRAULIC FAILURE

AIR

HYDRAULIC PRESSURE

Shuttle Valve Operation Figure 13 16.10.3 Pneumatic De-Icing systems The compressed air system used for inflating de-icing boots uses wet air pumps. The oily air leaves the pump and passes through baffle plates in an oil separator. The oil collects on the baffles and drains down to a collector at the separator base and returned to the engine oil sump. Clean air leaves the separator and flows through the de-icing selector valve to a pressure regulating valve, where its pressure is reduced to the value needed for the boots. It then flows to the distribution sequencing valve. When the system is switched off the air is directed overboard. De-icing systems are dealt with in more detail in Module 11.12 Ice And Rain Protection. 16.10.4 Air Conditioning And Pressurisation. Bleed air supplies provide hot air to the air conditioning packs. The hot air passes through primary and secondary heat exchangers before it is mixed with cold air to provide conditioned air into the aircraft. As the hot air passes through the system it flows across a turbine which drives the system compressor.

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Bleed air is also used for cabin pressurisation. The air drives a compressor which pressurises the air before it is fed to the cabin. Some aircraft use a jet pump to pressurise the air. Th air passes through an inter cooler to reduce its temperature before entering the cabin. Air-conditioning systems are often protected by flow control valves, which double as shut-off valves in the case of a fault. 16.10.5 Air Driven Hydraulic Pumps. Some aircraft use hydraulic pumps operated by air turbines. These are driven by bleed air from the engines and the flow is controlled and modulated by a solenoid operated pressure regulator and shut off valve to maintain the turbine speed within set limits. The turbine is connected to the pump via a shaft and the air is exhausted to atmosphere from the turbine outlet. 16.10.6 Pressurising Of Hydraulic Reservoirs. Aircraft flying at altitudes in excess of 20000 feet require the hydraulic reservoir to be pressurised to prevent foaming of the fluid, due to the low ambient air pressure and to prevent pump cavitation. The bleed air is fed to a regulator/reducing valve which regulates the pressure supplied to the reservoir. A pressure relief valve is fitted to the system which vents any excess air pressure to atmosphere. 16.10.7 Waste And Water Systems The toilet systems fitted to larger aircraft use a vacuum to empty a number of toilets into a single collector tank. This saves having a self-contained tank, full of de-odorising fluid and the associated pumping mechanisms attached to each toilet assembly.

Vacuum Waste System Figure 14


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The flush operation consists of fresh water from the potable supply and, most importantly, the vacuum, which draws the waste into the collector tank. This is obtained by having the tank connected to the outside of the aircraft. Only at low levels, when the outside air pressure is insufficient, is a small vacuum pump called into operation. Figure 14 shows a typical vacuum toilet system. 16.10.8 Pneumatic Stall Warning

These systems are common on light aircraft. A slotted plate is mounted on the wing leading edge and its slot coincide with the stagnation point of the wing during normal flight. The slot is connected to a horn via a tube. When the angle of attack is sufficient to induce a stall the low air pressure is drawn into the tube and sounds the horn giving the pilot warning of an impending stall.

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MODULE 11.17 WASTE WATER

CONTENTS 17 WATER AND WASTE SYSTEMS ................................................ 17-3 17.1

17.2

WATER SYSTEMS ......................................................................... 17-3 17.1.1 Pressure Control ........................................................... 17-3 17.1.2 Water Distribution System ............................................. 17-4 17.1.3 Water Heating ............................................................... 17-5 17.1.4 Waste Water Collection And Drainage .......................... 17-6 17.1.5 Quantity Indication ........................................................ 17-6 17.1.6 Water Service Panel ..................................................... 17-7 WASTE SYSTEMS ......................................................................... 17-8 17.2.1 Removable Toilet Assemblies ....................................... 17-9 17.2.2 Liquid Flush Toilets ....................................................... 17-9 17.2.3 Vacuum Toilets ............................................................. 17-11 17.2.4 Corrosion Control .......................................................... 17-13


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17 WATER AND WASTE SYSTEMS 17.1 WATER SYSTEMS The term “Potable water” refers to drinking water. On aircraft it is used not only to supply water for drinking, but also the galleys and to provide hot and cold water to wash basins throughout the aircraft. A centralised water tank can feed any number of galleys and toilets through a gallery of pipes. This will speed servicing turnaround times when there need only be one main replenishment point. Potable water is Hyper-chlorinated to control bacteria and is carried out at set intervals. The major components in a potable water system are:      

A storage tank. Air pressure system to force water from the storage tank to the services. Distribution lines Filling system Quantity indication system Valves to drain the system

The tank is usually stored under the cabin floor in a cradle structure and is constructed from either fibreglass with metal bonded bands or stainless steel. The quantity and volume will be dictated by the number of passengers carried and the length of the time the aircraft is airborne. Aircraft that are expected to operate in cold climates may have heater blankets built in to the design to keep the tank and the replenishing panel free of ice. The tank assembly will incorporate a drain, filler connection, overflow connection an air pressure connection and outlet pipelines to the galley and toilets. 17.1.1 Pressure Control

The supply of air for the movement of water is, tapped from the bleed air supply of the engine compressor or the APU. Some aircraft, which require the ability to draw water when there is no air pressure (on the ramp), have an electrically powered air compressor that will provide a head of pressure to enable water to be drawn off at any time. The compressor may automatically start when the bleed air pressure drops below a pre-determined value. On aircraft using a compressor, a riser loop is incorporated to prevent water entering the compressor, the top of the loop being higher than the distribution ducting ensuring that the water goes to the distribution lines first. A pressure switch will control the compressor starting and stopping as the bleed air pressure varies. The distribution lines are connected to the tank drain, fill connection, overflow connection, air pressure connection and the supply lines to all of the galleys and lavatories.


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THE RISER LOOP PREVENTS THE WATER FROM SIPHONING BACK THROUGH THE COMPRESSOR

COMPRESSOR PRESSURE SWITCH

WATER TANK

PRV

NRV`S

FILTER

PRESSURE MANIFOLD

Water Tank Pressurisation System Figure 1 17.1.2 Water Distribution System A main water distribution line is taken from the water tank and is routed up into the cabin ceiling. Individual pipelines are routed from this pipeline to the toilets and galleys. The distribution lines are usually flexible hoses enclosed in an aluminium sheath.

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The flexible hose is normally insulated to prevent it from freezing. The outer sheath prevents any leakage from entering the cabin. Any leaking water will be directed to the lower fuselage through drain tubes where it can then be drained overboard. A quick release connection is located above each toilet and galley to enable the supply line to be disconnected for removal of the toilet or galley. On smaller aircraft the water tank may be located above the wash basin and galley areas and provides water to the systems under gravity. 17.1.3 Water Heating A water heater with a small capacity is installed in the supply piping under each lavatory sink and provides hot water to the hot water tap. The heater contains electrical elements in the base of the heater unit. On the side of the tank is a warning light, a control switch, an overheat re-set switch and a pressure relief valve. Normally the heater switch will be on and the light will be illuminated. A switch controller will regulate the water temperature to around 125 degrees F. If a malfunction occurs and the temperature increases to 190 degrees F the overheat switch will operate and switch off power to the heater unit. The power light will go out. After a cooling down period the heater will have to be manually reset by pressing the re-set button on the heater unit. A pressure relief valve will relieve pressures in excess of around 140 psi. the primary function of the relief valve is to relieve pressures caused by the water overheating. A typical water heating system is shown in Figures 2 and Figure 3. ON/OFF SWITCH

OVERHEAT RESET SWITCH

OVERHEAT SWITCH

POWER SUPPLY HEATING ELEMENTS

ELECTRICAL CONNECTOR

CYCLIC SWITCH

NEON INDICATOR

Heating System – Schematic Figure 2


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Heating System Figure 3 17.1.4 Waste Water Collection And Drainage Waste water collection and drainage depends on the aircraft. On some aircraft the water from the washbasins drains directly overboard while on others it drains into a soil tank and is used to flush the toilet system. Water drained overboard are drained through drain masts under the fuselage. These masts are normally electrically heated to prevent freezing and the forward motion of the aircraft ensures that the water is finely atomised as it leaves the aircraft. To test the drain mast heaters on the ground the hand is carefully used feeling for warmth. 17.1.5 Quantity Indication Some aircraft use a simple sight gauge by the side of the tank to indicate the level of the waste tank contents. On larger aircraft the tank will be fitted with a sensor to remotely signal the tank levels to the cabin crew. One method of indication is to use a gauge on the attendants panel and a corresponding gauge which is fed from the same float and electrical transmitter on the water service panel.

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Another method of indicating tank contents is to use a series of lights controlled by magnetic floats installed inside the tank. When the waste water level operates one of the magnetic floats a circuit is made and a corresponding light on the panel is illuminated. 17.1.6 Water Service Panel A water service panel will normally be found on the lower part of the fuselage, where it can be easily reached by the maintenance crew who have the job of replenishing the tank during the turnaround maintenance. The panel will probably contain most of the following items. A filling point, a drain/overflow point, a quantity indication, either in the form of an array of lights or a gauge unit and an external air connection. A typical servicing panel is shown in Figure 4 QUANTITY GAUGE

OPEN DRAIN CONNECTION

WATER MAIN CONNECTOR

WATER DRAIN HANDLE

WATER VENT VALVE OPEN

CLOSE D

WATER SYSTEM MAIN VALVE

WATER FILL CONNECTOR

CLOSE D

FILL CONNECTION WITH COVER Servicing Panel Figure 4 The filling point on the panel will allow the replenishing rig/ truck to fill the tank during a turnaround servicing, whilst the drain/overflow will show when the tank is full. When full any excess water overflows out of the overflow line. Once the water is seen from the overflow valve the fill/vent valve is closed to the vent position. The quantity indicator will allow the tank to be filled to a 'less-than-full' quantity, where the aircraft is, perhaps, on very short flight legs and the excess weight of the water that will not be used, is traded-off against fuel. The external air connection allows a ground air air source to be connected to allow the water to be moved, within the system, whenever there is no internal air pressure available. Figure 5 shows a typical potable water system.


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FILL LINE

FILL/VENT VALVE STACK PIPE

WATER TANK

WATER DRAIN VALVE

OVERFILL LINE

FILL POINT

QTY GAUGE

Replenishment System Figure 5 The water drain valve is manually operated and allows the tank contents to drain under gravity. When the tank is emptied the drain valve is manually re-set. The fill/vent valve can be manually or electrically operated and rotates the valve to the fill or vent position. Its operation may also electrically isolate the air compressor, if fitted during filling. The purpose of the vent valve is to prevent an air lock occurring in the wash basin taps by opening the tap lines to atmosphere. Modern aircraft have self venting taps which automatically relieve any air locks. 17.2 WASTE SYSTEMS The provision of aircraft toilets is an essential requirement for any aircraft carrying passengers over long distances. These toilets must be maintained and serviced with care, as the comfort and health of the passenger must be protected. They should be clean and odour free at all times. There are three main types of toilet fitted to aircraft. The type used will depend upon the number of passengers the aircraft can carry, and also the age of the aircraft. In all cases it is essential that all the relevant health precautions are observed during all forms of servicing carried out on these units. Due to the nature of the fluids carried in many toilets, protection must also be given to the structure of the aircraft to protect it from corrosion caused by these fluids.

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The three types of toilet are: 

Removable toilet assembly.



Liquid flush type.



Vacuum toilet assembly.

17.2.1 Removable Toilet Assemblies The removable, or 'carry out' toilet is of the simplest type of aircraft toilet. This unit is often referred to as an „Elsan‟, named after the original company which manufactured this type of toilet. It is simply a storage bin with a toilet seat fitted to the top and partially filled with a strong chemical deodorant. This type of toilet is removed from the aircraft and emptied into an approved disposal site. Once washed out, it is replenished with deodorant and re-fitted into the aircraft, using some form of quick release attachment such as 'pip' pins. You will only find this type of toilet fitted to short range small light aircraft. 17.2.2 Liquid Flush Toilets These are the most common type of toilet found in passenger aircraft, each toilet being a completely self contained quick release unit, having its waste collection tank mounted directly beneath the toilet bowl. The tank is normally made out of composites or plastics. Directly below the waste tank is a service panel. An illustration of a typical liquid flush toilet assembly is shown overleaf. The assembly shown in Figure 6, contains the following components, which will be found in most liquid flush toilet installations: 

Motor and Pump.



Filter.



Drain Valve



Rinse Ring.



Flush Line.



Air vent.


Page 17-9

JAR 66 CATEGORY B1


A Typical Liquid Flush toilet Figure 6 The bowl is constructed from stainless steel, but the tank units can be either the same material or fibreglass laminate. The capacity of the tank depends on both the duration of the flights and the number of passengers catered for by each unit. An average tank capacity figure is 20 gallons, (90 litres) of which 3 gallons, (13.5 litres) are a pre-charge of chemical; which contains disinfectant, dye and deodorant. This would be sufficient for about 100 uses, before emptying and recharging would be required. 

System Operation

When the flush button is pressed, the motor runs for a fixed time, usually around 10 seconds, which pumps the fluid through the bowl spray pipe in a swirling action. This action flushes the bowl contents into the tank, via a lightly sprung (loaded into the closed position) hinged separator. At the end of the 10-second cycle, the motor re-arms to run again, in the reverse direction, to ensure the filter does not become blocked with solid waste.

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17.2.3 Vacuum Toilets On an aircraft fitted with a number of liquid flush toilets, there were two major problems, the corrosion risk and the time taken to drain and replenish each individual toilet. Both of these problems are overcome by installing vacuum toilets. There are dry toilet modules installed at convenient locations, to suit the seating layout around the passenger cabin and connected to a central storage tank by pipelines. The vacuum toilet uses a waste container that has a negative pressure inside, (vacuum). This vacuum draws the waste from the bowl together with the clean flushing water and deposits it into the tank. On very large aircraft, more than one waste tank is used to overcome the problem of one tank filling up during the flight. The toilet systems fitted to larger aircraft use a vacuum to empty a number of toilets into a single collector tank. This saves having a self-contained tank, full of de-odorising fluid and the associated pumping mechanisms attached to each toilet assembly. The flush operation consists of fresh water from the potable supply and, most importantly, the vacuum, which draws the waste into the collector tank. This is obtained by having the tank connected to the outside of the aircraft. As the aircraft‟s speed increases the pressure at the connection drops which causes the waste to be drawn to the storage tank.

Vacuum Toilet System Figure 7


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At low speeds or low altitudes, when the pressure differential is insufficient to draw the waste to the storage tank, a small vacuum pump called a „vacuum generator‟, is operated by a pressure switch to provide the required pressure drop. Its normal range of operation is between sea level up to 16,000 ft. The illustration below shows a typical vacuum waste storage tank installation:

Waste Storage Tank Figure 8 

Emptying

Large aircraft usually hold waste in a storage tank that is emptied after the aircraft has landed. The task of emptying the tanks at the end of the flight usually rests with the specialist companies, sub-contracted to the airlines, they empty all waste tanks at particular airports. The tanks are emptied in one of two methods, gravity or suction. The gravity method empties the tank, after the hose of the toilet emptying vehicle has been connected, by simply operating the shut-off valve. Once the tank is emptied, it is flushed out and, depending on the type of tank, replenished with deodorising fluid. Suction requires both that the emptying vehicle has the correct equipment, (set at the correct suction value), and that the aircraft has ducting that is cleared for use with suction equipment. If the aircraft only has 'gravity' emptying ducting and piping, severe damage will be caused to much of the toilet equipment, if used with vacuum emptying equipment.

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17.2.4 Corrosion Control All the areas where toilet equipment is fitted must be protected against corrosion. The effect of many toilet chemicals on aluminium alloy aircraft structure is severe. All spillages must be neutralised and cleaned off as soon as possible, whilst thorough checks of all the areas of the aircraft that could be affected, must be inspected at regular intervals. Such areas could be the toilet floor itself and beneath that floor; the vicinity of the collector tank(s), around the draining/filling panel and anywhere else the corrosive fumes could affect the structure. Some toilet units are enclosed in an anti corrosion tank. Any leaks would be self contained within this tank. The tank would be connected to the drain lines. The toilet floors may be made from composite materials to reduces the likelihood of corrosion damage. All connections in the service panel are sealed off when the service panel is closed.


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JAR 66 CATEGORY B1



Page 17-14


JAR 66 CATEGORY B1 MODULE 11 BOOK 2 PART 1 AEROPLANE AERODYNAMICS, STRUCTURES & SYSTEMS

PART ONE CONTENTS 1

INSTRUMENT SYSTEMS (ATA 31) ............................................. 1-1 1.1 1.2

1.3

1.4 1.5 1.6

1.7 1.8

1.9

1.10

1.11 1.12 1.13

1.14

THE ATMOSPHERE ....................................................................... 1-1 1.1.1 STANDARD ATMOSPHERE ................................................ 1-3 PRESSURE INSTRUMENTS ............................................................. 1-4 1.2.1 AIR DATA INSTRUMENTS.................................................. 1-4 1.2.2 LOCATION OF PROBES AND STATIC VENTS ....................... 1-7 ALTIMETERS ................................................................................ 1-10 1.3.1 ANEROID BAROMETER .................................................... 1-10 1.3.2 FRICTION COMPENSATION ............................................... 1-13 1.3.3 TEMPERATURE COMPENSATION ....................................... 1-13 1.3.4 PRESSURE COMPENSATION............................................. 1-15 SERVO ASSISTED ALTIMETERS ..................................................... 1-18 1.4.1 GENERAL ....................................................................... 1-18 DIRECT SERVO ALTIMETER ........................................................... 1-19 1.5.1 DATUM PRESSURE SETTING ............................................ 1-22 PRESSURE REVERTING SERVO ALTIMETER .................................... 1-23 1.6.1 SERVO MODE –OPERATION ............................................. 1-25 1.6.2 STANDBY MODE – OPERATION ........................................ 1-25 1.6.3 DATUM PRESSURE SETTING ............................................ 1-26 CABIN ALTIMETERS ...................................................................... 1-26 AIRSPEED INDICATORS ................................................................. 1-28 1.8.1 SIMPLIFIED AIRSPEED INDICATOR .................................... 1-28 1.8.2 PITOT PRESSURE............................................................ 1-31 1.8.3 SPEED OF SOUND ........................................................... 1-32 1.8.4 MACHMETER .................................................................. 1-33 1.8.5 COMBINED SPEED INDICATOR .......................................... 1-35 1.8.6 PRESSURE OPERATED CSI ............................................. 1-36 1.8.7 SERVO OPERATED CSI ................................................... 1-37 VERTICAL SPEED INDICATORS ...................................................... 1-38 1.9.1 BASIC OPERATION .......................................................... 1-38 1.9.2 CALIBRATION .................................................................. 1-40 1.9.3 ALTITUDE & TEMPERATURE COMPENSATION .................... 1-41 GYROSCOPIC INSTRUMENTS ......................................................... 1-42 1.10.1 GYROSCOPIC PROPERTIES .............................................. 1-42 1.10.2 RIGIDITY......................................................................... 1-42 1.10.3 PRECESSION .................................................................. 1-43 1.10.4 PRECESSION .................................................................. 1-45 1.10.5 VERTICAL GYRO ............................................................. 1-46 GYRO HORIZON UNIT .................................................................... 1-48 VERTICAL REFERENCE UNIT (VRU) ............................................... 1-53 ATTITUDE DIRECTOR INDICATOR (ADI) .......................................... 1-54 1.13.1 WARNINGS ..................................................................... 1-56 1.13.2 ATTITUDE DISTRIBUTION ................................................. 1-56 1.13.3 ATTITUDE TRANSFER SWITCHING..................................... 1-58 STANDBY ATTITUDE INDICATORS .................................................. 1-59 1.14.1 DESCRIPTION AND OPERATION ........................................ 1-59

MOD 11 BOOK 2 PART 1 ISSUE 6 - 01/02/11

Part 1 Page 1


1.15 1.16 1.17 1.18 1.19 1.20 1.21 1.22 1.23 1.24 1.25

1.26

1.27

1.28 1.29 1.30

1.31

1.32

1.33

1.34 1.35 1.36

Part 1 - Page 2

1.14.2 RUNNING UP .................................................................. 1-60 1.14.3 ERECTION CONTROL ...................................................... 1-60 1.14.4 CAGING ......................................................................... 1-60 1.14.5 ATTITUDE INDICATION ..................................................... 1-60 STANDBY ATTITUDE INDICATOR H 341 .......................................... 1-62 1.15.1 DESCRIPTION ................................................................. 1-63 DIRECTION INDICATORS ............................................................... 1-65 TURN & SLIP INDICATOR .............................................................. 1-67 1.17.1 BANK INDICATION ........................................................... 1-69 TURN CO-ORDINATOR .................................................................. 1-71 HORIZONTAL SITUATION INDICATOR (HSI)..................................... 1-72 COLLINS 331A-8K HSI ................................................................ 1-74 1.20.1 WARNING FLAGS ............................................................ 1-76 ANGLE OF ATTACK (AOA) ........................................................... 1-77 STALL WARNING INDICATION ........................................................ 1-79 ELECTRONIC INSTRUMENT SYSTEMS ............................................. 1-81 ELECTRONIC FLIGHT INSTRUMENT SYSTEM ................................... 1-84 ELECTRONIC ATTITUDE DIRECTOR INDICATOR (EADI) ................... 1-84 1.25.1 FULL TIME EADI DISPLAY DATA ...................................... 1-86 1.25.2 PART TIME EADI DISPLAYS ............................................ 1-87 ELECTRONIC HORIZONTAL SITUATION INDICATOR (EHSI) .............. 1-89 1.26.1 FULL TIME EHSI DISPLAYS ............................................. 1-90 1.26.2 PART TIME EHSI DISPLAYS ............................................ 1-92 1.26.3 PARTIAL COMPASS FORMAT............................................ 1-93 1.26.4 MAP MODE .................................................................... 1-96 1.26.5 COMPOSITE DISPLAY ...................................................... 1-97 EFIS CONTROLLER ...................................................................... 1-98 1.27.1 DISPLAY CONTROLLER ................................................... 1-99 1.27.2 SOURCE CONTROLLER ................................................... 1-101 OTHER SYSTEM INDICATIONS ....................................................... 1-103 POWERPLANT INSTRUMENTATION ................................................. 1-103 FUEL CONTENTS GAUGE .............................................................. 1-103 1.30.1 RESISTANCE GAUGES..................................................... 1-103 1.30.2 CAPACITANCE QUANTITY INDICATORS ............................. 1-104 FUEL FLOW INDICATOR ................................................................ 1-106 1.31.1 FUEL FLOW TRANSMITTERS ............................................ 1-108 1.31.2 SYNCHRONOUS MASS FLOW FLOW-METER SYSTEM ......... 1-108 1.31.3 MOTORLESS MASS FLOW METER SYSTEM ....................... 1-109 PRESSURE INDICATORS................................................................ 1-111 1.32.1 PRESSURE CAPSULE DETECTION .................................... 1-112 1.32.2 BOURDON TUBE DETECTION ........................................... 1-113 OIL & FUEL TEMPERATURE INDICATORS ....................................... 1-115 1.33.1 RESISTIVE BULB SENSOR ............................................... 1-115 1.33.2 THERMOCOUPLE SENSOR ............................................... 1-116 ENGINE RPM INDICATORS ............................................................ 1-117 1.34.1 ENGINE SPEED GENERATOR ........................................... 1-119 EXHAUST TEMPERATURE INDICATING............................................ 1-121 ENGINE PRESSURE INDICATORS ................................................... 1-124



1.37 1.38 1.39 1.40

1.41 1.42 1.43 1.44 1.45

1.46

1.36.1 EPR FORMULA ............................................................... 1-125 VIBRATION INSTRUMENTS ............................................................. 1-126 ELECTRONIC INSTRUMENTS (ENGINE & AIRFRAME) ........................ 1-130 ENGINE INDICATING & CREW ALERTING SYSTEM (EICAS) ............. 1-130 1.39.1 DISPLAY UNITS ............................................................... 1-131 DISPLAY MODES .......................................................................... 1-135 1.40.1 OPERATIONAL MODE....................................................... 1-135 1.40.2 STATUS MODE ................................................................ 1-135 1.40.3 MAINTENANCE MODE ...................................................... 1-135 DISPLAY SELECT PANEL ............................................................... 1-137 1.41.1 DISPLAY SELECT PANEL OPERATION ............................... 1-138 ALERT MESSAGES ....................................................................... 1-139 MAINTENANCE CONTROL PANEL ................................................... 1-141 ELECTRONIC CENTRALIZED AIRCRAFT MONITORING ...................... 1-142 1.44.1 DISPLAY UNITS ............................................................... 1-142 ECAM DISPLAY MODES ............................................................... 1-143 1.45.1 FLIGHT PHASE RELATED MODE ....................................... 1-143 1.45.2 ADVISORY MODE ............................................................ 1-144 1.45.3 ECAM FAILURE MODE .................................................... 1-145 CONTROL PANEL ......................................................................... 1-151 1.46.1 ECAM CONTROL PANEL ................................................. 1-152


Part 1 Page 3



Part 1 - Page 4



1

INSTRUMENT SYSTEMS (ATA 31)

Aircraft instruments can, on initial observation, appear a bewildering mass of dials or 'TV ' type screens. The different types of instrumentation required fall into one of the following types: 1. Pressure instruments.

2. Gyroscopic instruments

3. Compasses.

4. Mechanical indicators 5. Electronic instruments

1.1 THE ATMOSPHERE A relatively thin layer of air called the atmosphere surrounds the earth. This extends upwards from the surface for a distance of about 250 miles and is composed mainly of nitrogen 78%, oxygen 21% plus 1% of other gases which includes amongst others, argon, carbon dioxide and helium. Under the gravitational effect of the earth, the atmosphere exerts a pressure upon the surface of the earth. This pressure, if measured at sea level, it is approximately 1.013bar (14.7lbf/in2), and reduces with height. The pressure reduction, is not linear, the rate of pressure reduction decreases with a rise in altitude to form an exponential curve. Temperature and water vapour within the air also affects the pressure of the air, and therefore the height at which a particular pressure can be measured. Figure 1 shows a Height/pressure graph. 65 60 55

AT 8,000ft 240mb

HEIGHT X 1000ft

50 45 40 35 30

AT 8,000ft 750mb

25

AT SEA LEVEL 1013mb

20 15 10 5 0 0

.100

.200

.300

.400

.500

.600

.700

.800

.900

1.000

AIR PRESSURE IN BARS

Height/Pressure Graph Figure 1


Page 1-1


Temperature change within the atmosphere can be divided into 3 bands, corresponding to the 3 layers or regions of the atmosphere: 1.

The Troposphere.

2.

The Stratosphere.

3.

The Chemosphere.

Figure 2 shows three bands of the atmosphere.

135

+22.473°

140,000ft

CHEMOSPHERE TEMPERATURE INCREASES AT APPROXIMATELY 2.256°C FOR AN INCREASE IN HEIGHT OF 1000ft

125 115 STRATOPAUSE 104,987ft

105

-56.5°

95 ALTITUDE FEET X 1000

85 75

STRATOPHERE - TEMPERATURE AT -56.5°C

UPPER LIMIT OF ICAO ISA 65,800ftt

65 55 45

TROPOPAUSE 36,090ft

35

-56.5°

25 TROPOSPHERE - TEMPERATURE DECREASES 1.98°C FOR AN INCREASE IN HEIGHT OF 1000ft

15 5 +15°

0 -50

-40

-30

-20

-10

0

10

20

30

TEMPERATURE (DEGREES C)

Atmosphere Bands Figure 2

Page 1-2



The height of these layers varies considerably with latitude and the season. It is assumed that the troposphere extends to a height of 36,090ft and has a temperature gradient falling at a linear rate to –56.5ºC at 36,090ft. The stratosphere is assumed to range from 36,090ft to 104,987ft and to have a constant temperature of –56.5ºC. Above this is the Chemosphere, extending to the limits of the atmosphere and which is assumed to have a temperature gradient, which initially rises approximately 2ºC for each 1000ft of altitude. For the purpose of aircraft pressure instruments, these higher levels are not important. 1.1.1 STANDARD ATMOSPHERE To be able to produce an instrument capable of accurately measuring aircraft height (and speed) using only the prevailing atmospheric pressure, requires that the instrument be calibrated and tested against a set of standard conditions. Standard atmospheres have been in use since 1800‟s. the early ones being based on very simple temperature laws. During WW1, these were found to be inadequate, this led to the development and the international acceptance in 1924 of the International Committee on Air Navigation (ICAN) standard. This standard was adopted by the International Civil Aviation Organisation (ICAO) in 1952. Advances in aircraft performance and the introduction of missiles highlighted the need for an increase in the altitude range of the standard atmosphere, the ICAO limit being 65,000ft. This introduced two further standards to supplement the ICAO standard, these being the Wright Air Development Centre (WADC) and the Air Research Development Command (ARDC). Table 1 shows the comparison of the standard atmospheres. Height in feet x 1000 0 10 20 30 40 50 60 70 80 90 100 110 120 130 140 150 160

ICAN 1013.25 696.91 465.63 301.89 187.61 115.81 71.79 44.36 -


Air Pressure in Millibars ICAO WADC 1013.25 1013.25 696.81 696.81 465.63 465.63 300.01 300.89 187.54 187.54 115.97 115.97 71.72 71.72 44.35 27.43 16.96 10.49 6.53 4.22 2.84 1.97 Table 1

ARDC 1013.25 696.91 465.63 300.89 188.23 115.97 71.716 44.438 27.425 17.067 10.820 6.981 4.5779 3.0476 2.0575 1.4650 0.9727

Page 1-3


1.2 PRESSURE INSTRUMENTS 1.2.1 AIR DATA INSTRUMENTS An Air Data system of an aircraft is one which the total pressure created by the forward motion of an aircraft, and the static pressure of the atmosphere surrounding it, are sensed and measured in terms of speed, altitude and rate of change of altitude. The measurement and indication of these three parameters may be achieved by connecting the appropriate sensors, either directly to mechanical-type instruments, or to a remotely-located Air Data Computer (ADC), which then transmits the data in electrical signal format to electro-mechanical or servo-type instruments. The basic Air Data Instruments display airspeed, altitude, Mach number and vertical speed. All are calculated from air pressure received from a Pitot/Static source. 1. Static air pressure, which is simply the outside air pressure at the instant of measuring. 2. Pitot pressure is the dynamic pressure of the air due to the forward motion of the aircraft and is measured using a tube, which faces the direction of travel.

Page 1-4



Figure 3 shows a Pressure head as fitted to aircraft to allow Pitot and Static pressures to the relevant indicators.

PITOT LINE

STATIC LINE

HEATER CONNECTION

FORWARD

PITOT PROBE

STATIC VENTS

Aircraft Pressure Head Figure 3 Indicated Airspeed (IAS), Mach No, Barometric Height (Height above sea level), and Vertical speed (Rate of climb/dive) are derived from the Pitot/Static inputs. 1. IAS = Pitot minus Static - (In knots). 2. Mach No = Pitot - Static divided by Static. 3. Baro Ht = Static - (In feet). 4. Vertical Speed = Change in Static pressure - (X 1000ft/min).


Page 1-5


Figure 4 shows typical aircraft static vent:

FUSELAGE

STATIC VENT

STATIC PIPE

Aircraft Static Vent Figure 4

Page 1-6



1.2.2 LOCATION OF PROBES AND STATIC VENTS The choice of probe/vent locations is largely dependent on the type of aircraft, speed range and aerodynamic characteristics, and as result there is no common standard for all aircraft. On larger aircraft it is normal to have standby probes and static vents. These are always located one on each side of the fuselage and are interconnected so as to balance out dynamic pressure effects resulting from any Yawing or side-slip motion of the aircraft. Figure 5 shows the location of probes and vents on a Boeing 737.

Boeing 737 Air Data Probe and Vent Location Figure 5 Pitot and static pressures are transmitted through seamless and corrosionresistant metal (light alloy) pipelines. Flexible pipelines are also used when connections to components mounted on anti-vibration mountings is required. In order for an Air Data System to operate effectively under all flight conditions, provision must also be made for the elimination of water that may enter the system as a result of condensation, rain, snow, etc. This will reduce the probability of “Slugs” of water blocking the lines. This provision takes the form of drain holes in the probes, drain taps and valves in the system‟s pipelines. The drain holes within the probes are of diameter so as not to introduce errors into the system.


Page 1-7


Methods of draining the pipelines varies between aircraft types and are designed to have a capacity sufficient to allow for the accumulation of the maximum amount of water that could enter the system between maintenance periods. Figure 6 shows a typical water drain valve.

ORANGE FLOAT INDICATOR

TRANSPARENT PLASTIC PIPE

DRAIN VALVE

BAYONET FITTING CAP

(SELF SEALING)

Water Drain Valve Figure 6 The three primary instruments in the Air Data System are: 1. Altimeter (Baro Ht). 2. Indicated Air Speed (IAS) Indicator. 3. Vertical Speed Indicator. The IAS is often combined to display Mach No as well as indicated airspeed and is referred to as the “Combined Speed Indicator”.

Page 1-8



Figure 7 shows the connection and equations for the primary Air Data instruments.

Air Data Instrumentation Figure 7


Page 1-9


1.3 ALTIMETERS 1.3.1 ANEROID BAROMETER In its simplest form, if a membrane or pressure sensitive capsule is to be used to measure pressure, it usually forms part of a sealed capsule. If the capsule is evacuated, the atmospheric pressure on the outside of the capsule will force the capsule into the chamber until its resistance is sufficient to support the atmospheric pressure. The greater the atmospheric pressure the greater the movement of the capsule, before a balance is attained, and vice versa. If a linkage mechanism is attached to the membrane, this movement can be transmitted to a pointer to reflect the movement of the capsule. This then is the principle upon which the aneroid barometer is based for the measurement of atmospheric pressure. Figure 8 shows a simplified aneroid barometer.

ATMOSPHERIC PRESSURE ATMOSPHERIC PRESSURE

PIVOT

CAPSULE STACK

Simplified Aneroid Barometer Figure 8

Page 1-10



Altitude measurements require little change in the basic instrument configuration to enable barometric pressure (atmospheric pressure) to be translated into aircraft altitude. Figure 9 shows a simplified mechanism of a directly operated capsule altimeter.

POINTER

AIRTIGHT INSTRUMENT CASE

EXTERNAL STILL AIR PRESSURE (STATIC)

CAPSULE STACK

Simplified Altimeter Figure 9 It consists of an airtight instrument case containing an evacuated capsule stack. The capsule stack is connected by a system of levers and gears to a pointer which, moves over a scale calibrated in feet. External still air (static) pressure is fed in to the instrument case so that as the aircraft climbs the pressure in the case falls, allowing the capsule to expand. This motion is then used by the system of levers and gears to drive the pointer over the dial. When the aircraft loses altitude, the reverse happens.


Page 1-11


Figure 10 shows the face of a barometric altimeter .

3 - POINTER ALTIMETER

SINGLEPOINTER ALTIMETER

Barometric Altimeter Figure 10

Page 1-12



1.3.2 FRICTION COMPENSATION Friction in the gearing of a simple altimeter cannot truly be compensated for, however, it is reduced as much as possible by careful design and meticulous attention to cleanliness and finish during manufacture. The rate of response of the instrument to capsule movement can be further improved, when considered necessary, by the use of a vibrator. This simply helps prevent the mechanism from sticking. 1.3.3 TEMPERATURE COMPENSATION Temperature affects the strength of the materials used in the manufacture of the capsules and springs, causing them to become stronger as temperature decreases. The overall effect of this is that with a drop in temperature the capsule stack will tend to extend, with the result that the instrument will over-read. Conversely, a rise in temperature causes the capsule stack to contract and the instrument under-read. There are two main methods employed to compensate for this temperature-induced variations in readings, both of which use a bimetallic element as the compensating mechanism. The first method used is to mount the capsule stack within a “U” shaped bimetallic bracket, the open end of which is connected to the top of the capsule stack by pins. The composition of the b-metallic brackets is arranged so that with a drop in temperature the limbs tend to move inwards, exerting a compressive force onto the capsule stack, in opposition to the tendency of the capsule stack to expand with a fall in temperature. Figure 11 shows the “U” bracket method of temperature compensation. DROP IN TEMPERATURE LIMBS MOVE INWARDS EXERTING A COMPRESSIVE FORCE ONTO THE CASULE STACK

CASULE STACK

BIMETAL “U” SPRING

Temperature Compensation – “U” Bracket Figure 11


Page 1-13


The second method employed to compensate for changes in temperature of the capsule stack is to introduce a bi-metallic link into the system of levers used to transmit capsule movement to the instrument‟s pointers. In this instance, a “U” shaped bi-metallic link has been introduced. This effectively alters the length of the linkage to compensate for the tendency of the capsule stack to expand or contract with changes in temperature. Figure 12 shows the bi-metallic compensating link method.

CAPSULE

BIMETAL COMPENSATING LINK

Bi-metallic Compensating Link Figure 12

Page 1-14



1.3.4 PRESSURE COMPENSATION All aircraft pressure operated altimeters, are calibrated to one of the standard atmospheres, and will provide an accurate altitude indication providing that the atmospheric pressure prevailing conforms to the standard atmosphere. Anyone who is familiar with the weather forecast on TV will realise that the atmospheric pressure is always changing at any given point, as well as varying from area to area. We are not concerned with the reasons why this happens, only the effect this has on the altimeter. Under standard conditions, at sea level with an ambient atmospheric pressure of 1013.25 millibars, an altimeter calibrated to the ICAO standard atmosphere would indicate zero feet. If the sea level pressure remains constant at 1013.15 millibars, the altimeter indications would correspond to the ICAO pressure standards. However, standard atmospheric conditions rarely prevail and variations in sealevel pressure will result in variations in the indicated altitude. For example, if the sea-level pressure falls to 1010 millibars, then the capsule stack would sense this decrease in pressure and expand, showing an error of +100ft. A corresponding change in sea-level pressure to 1016.55 millibars would cause an error reading of –100ft. At this height, one-millibar change in pressure corresponds to a 30ft change in altitude, but as altitude increases so does the error. This is shown in figure 13. ICAO STANDARD

PRESSURE DROP

PRESSURE RISE

+100ft ERROR

Ht 5,000ft 843.21

-100ft ERROR

SEA LEVEL 1013.25

SEA LEVEL 1010.00

SEA LEVEL 1016.55

A1

A2

A3

Pressure Compensation Figure 13 MOD 11 BOOK 2 PART 1 ISSUE 6 - 01/02/11

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The ICAO standard atmosphere also assumes a temperature of 15C at sea level and a temperature drop (lapse rate)of 1.98C per 1000ft up to 36,090ft, it then remains at a constant temperature of –56.5C. If the lapse rate differs from this assumption then even a correctly set altimeter will indicate an error when an aircraft flies into an area where air temperature is higher or lower than that expected. Assuming the same sea-level pressures, the pressure at a certain height over a column of cold air is less than the pressure over a column of warm air at the same height. This is because cold air is denser than warm air. Therefore, in these conditions, an altimeter will over-read in air colder than standard conditions and under-read in air warmer than standard conditions. To help overcome these problems, the altimeter is fitted with a mechanism which enables the instrument datum can be adjusted to the prevailing barometric pressure. This mechanism consists of a system of gears within the instrument, which is controlled by a knob on the face of the instrument. This knob, called the “Ground Pressure Setting Knob”, allows the instrument datum and therefore the indicator pointers be repositioned without affecting the capsule stack. At the same time, an indicator, usually calibrated in millibars, will rotate to display the instrument datum setting. This indicator, known as the “Baroscale”, can be displayed as a linear scale but more commonly displayed using a veeder counter viewed through an aperture in the indicator face. The altimeter may be adjusted by the ground engineers to the prevailing atmospheric pressure before take-off, but is more commonly adjusted by the flight crew, who will obtain information regarding the prevailing atmospheric pressure from flight maps and from the local Air Traffic Control (ATC) via the aircraft‟s VHF communication system. The information obtained in this way is given in the form of radio “Q” codes, the most important of which are: QFE – Airfield barometric pressure. Altimeters with the baroscale set to this will read zero feet when landing or taking-off at the airport for which the QFE was given. QNH – Actual sea-level barometric pressure. Altimeters with the baroscale set to this will indicate height above mean sea-level (MSL). QNE – Standard sea-level barometric pressure (1013.25). Altimeters with the baroscale set to 1013.25 will indicate “Standard Pressure Altitude”. QFE is normally set into the altimeter before take-off and on approach before landing at any particular airport. QNH is normally set into the altimeter when the aircraft is below 3,000. QNE set into the altimeter when the aircraft is above 3,000ft.

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STANDARD SETTING 1013.25 MILLIBARS SEA LEVEL

HEIGHT ABOVE AIRFIELD

QFE

HEIGHT ABOVE SEA LEVEL

QNH

FLIGHT LEVEL

QNE

Figure 14 shows the Radio “Q” Codes for atmospheric pressure.

Radio “Q” Codes Figure 14 MOD 11 BOOK 2 PART 1 ISSUE 6 - 01/02/11

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1.4 SERVO ASSISTED ALTIMETERS Despite the use of a vibrator mechanism to enhance its response, the basic altimeter becomes increasingly inaccurate with height. This results directly from the non-linear changes in atmospheric pressure, with changes in altitude. For example, the pressure drop from sea level to 1000ft is 36.08mb, whereas from 50,000ft to 51,000ft the pressure drop is only 5.44mb. In addition to the errors caused by friction, the reduced pressure changes as height increases also exaggerates the errors which result from capsule hysteresis and creepage. Hysteresis occurs when the capsule movement lags behind the pressure change causing the motion. Creepage is the tendency for the capsule to readjust itself without a pressure change occurring. 1.4.1 GENERAL The errors within the basic altimeter can be reduced to acceptable levels by minimising the work done by the capsule. This is achieved by interposing a servo-assistance mechanism between the capsule stack and the gearing mechanism. The other main difference between the servo assisted altimeter and the basic altimeter is the dial presentation. This consists of a single pointer moving over a scale calibrated from 0-1000ft in 50ft divisions and a veeder digital counter, which records height up to 99,950ft which again is displayed in 50ft increments. Two main methods are used to provide servo-assistance for the basic altimeter. 1.

Direct Servo-Control.

2.

Pressure Reverting Servo-Control.

Direct Servo-Control: The servomechanism is operated directly from the capsule stack, with no mechanical link between the capsule stack and the gearing. Consequently, there is no back-up mechanical operation of the instrument in the event of a failure. Pressure reverting Servo-Control: The servomechanism is controlled from a remote pressure sensor, and a mechanical connection between the capsule and the gearing is retained to allow reversion to mechanical operation (pressure reverting) in the event of a power failure.

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1.5 DIRECT SERVO ALTIMETER Referring to figure 15, the “I” bar of the transducer is connected to the capsule stack and pivoted so that “I” bar position will change as the capsule expand and contract in response to a change of altitude. The “E” core, whose position relative to the “I” bar is controlled by the servo-motor, is wound as a transformer, with the primary coil on the centre limb and the secondary coils wound in series opposition onto the outer limbs. The primary is supplied via a transformer from the aircraft‟s 115V 400Hz supply. Figure 15 shows the face of a direct reading Altimeter.

Direct Reading Altimeter Figure 15


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POINTER

Direct Reading Servo-Controlled Altimeter – Schematic Figure 16

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MILLIBAR COUNTERS

GROUND PRESSURE SETTING KNOB

WORM GEAR SHAFT

MILLIBAR ADJUSTING BAR

LEVER

CAM FOLLOWER OVERRUN SWITCH HEIGHT COUNTERS

WARNING FLAG

SOLENOID

MOTOR

SERVO AMP

TRANSFORMER

CAPSULES

115V 400Hz

Figure 16 shows a schematic diagram of a direct reading servo-controlled altimeter.


Providing the “I” bar is equidistant from the “E” bar limbs, the resultant output from the secondary coils will be zero. However, when a change of altitude occurs, the “I” bar will pivot to follow the capsule movement and consequently the air gaps between the outer limbs of the “E” bar and the “I” bar will become unequal. The magnetic flux in the outer limb with the smaller gap will increase and the induced voltage on that limb will also increase. The opposite effect occurs in the other outer limb. This results in an output voltage, the magnitude and phase of which depends upon the amount and direction of the movement of the “I” bar. This output voltage is fed via an amplifier to the control winding of a two-phase AC servo-motor. Figure 17 shows the operation of the “E” & “I” transducer for increases and decrease of height.

A LT I TU D E C O N ST A NT (L O W L E V E L )

A LT I TU D E R IS IN G (L O W L E V E L )

A .C . EX CI T AT I ON SU PP L Y


R E S U LT AN T W A V E FO RM

A LT I TU D E C O N ST A NT (H IG H L E V E L)


A LT I TU D E F A L L IN G (H IG H L E V E L)





“E” & “I” Bar Operation Figure 17


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The two-phase AC motor has its main winding supplied with a constant reference voltage from the transformer. When the “I” bar is displaced by the movement of the capsules, the resultant voltage output to the servo-motor control winding either lags or leads the reference voltage. This sets up a rotating field in the motor, which causes it to rotate in a direction such that the pointer and digital counter moves in the correct sense to indicate the increase or decrease in altitude. At the same time the servo-motor drives the cam and cam follower which re-positions the “E” bar to equalise the air gaps between the “E” bar cores and the “I” bar, thus reducing the transducer output to zero when the aircraft‟s height stabilises. As the motor only drives the indicator, any power failure will result in the indication remaining at the height shown when the power failed. For this reason a “Power Failure Warning Indicator” (PFWI) is fitted to the instrument. The PFWI takes the form of a spring-loaded flag, which is held out of view by solenoid action while the power is connected. Any power failure removes the supply from the solenoid, allowing the flag to be returned into view by the spring action. To prevent the servomotor overrunning and damaging the altimeter mechanism, an overrun limit switch is incorporated. When the cam reaches a predetermined position, a stud on the side of the cam makes contact with the limit switch, opens its contacts and disconnects the electrical supply from the altimeter. The servomotor stops and the PFWI comes into view. 1.5.1 DATUM PRESSURE SETTING As with the basic altimeter, a “Ground Pressure Setting Knob” (GPSK) is provided to allow the various “Q” codes to be set into the instrument. When this knob is rotated, the veeder counter is turned by the associated gear train to show the millibars set. Rotation of the knob also alters the setting of the millibar adjustment rod; this moves the millibar lever about its pivot causing the worm gear to move laterally. Movement of the worm gear shaft in this way rotates the differential gear, cam and cam follower, displacing the “E” bar relative to the “I” bar. An error signal is therefore generated and fed via the amplifier to the servomotor, driving the indicator gear train, the worm gear cam and cam follower and the “E” bar back to the zero output position. The altimeter now shows aircraft altitude with respect to the ground pressure set onto the baroscale.

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1.6 PRESSURE REVERTING SERVO ALTIMETER This type of altimeter is servo assisted with automatic reversion to mechanical operation from the capsule stack in the event of power or other failures. The servo-assistance takes the form of a control transformer (synchro), amplifier and a two phase drag cup motor connected to the gearing mechanism between the capsule stack and the indicator pointer and counter. Figure 18 shows the face of a pressure reverting altimeter.

Pressure Reverting Altimeter Figure 18 In the servo mode of operation, the altimeter is connected to a master altimeter or Air data Computer (ADC), which provides a signal so that the altimeter gives a corrected indication of the aircraft‟s altitude. When the standby mode is selected, or in the event of a failure, the altimeter will operate as an unassisted basic precision altimeter. A vibrator mechanism is also incorporated within the altimeter to help reduce the effects of friction when operating in the standby mode.


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MILLIBAR SETTING KNOB

SIGNAL

CAPSULE MECHANISM

ALTITUDE SIGNAL FROM ADC

CX

SYNCHRO

CONTROL TRANSFORMER

SERVO AMP

DRAG CUP MOTOR

ANTI BACKLASH GEAR

HEIGHT COUNTERS

MILLIBAR COUNTER

POINTER

Figure 19 shows a schematic of the pressure reverting altimeter.

Pressure Reverting Altimeter Figure 19

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1.6.1 SERVO MODE –OPERATION With power on, the altimeter functions in the standby mode until the altimeter is switched to servo motor by momentarily turning the mode selector switch to the “RESET” position. This via a self-maintaining relay circuit (nor shown), connects the power to the amplifier and drag cup motor circuits, and retracts the standby flag from view. A corrected altitude signal generated by a synchro transmitter in the master altimeter or ADC, is fed to the stator of the control transformer (CT). This gives an error signal related to the difference between the position of the stator‟s magnetic field and the position of the rotor coil. Provided these are aligned at 90º to each other a null error signal is produced. The rotor position is initially determined by capsule displacement. Provided the rotor position and the CT stator input signal position remain at 90º, no error signal is produced, however, when the rotor position is out of alignment with respect to the input signal position an error signal is produced. This error signal is fed to the amplifier and then fed to the control phase of the two-phase drag cup motor. The motor, which is connected to the altimeter‟s gearing mechanism now assists the capsules to drive the indicator to the correct reading and also to align the CT rotor to the nil error position stopping the motor. Thus as the CT rotor is always driven to the nil error position, the indications produced by the instrument reflect the input signal position generated by the master altimeter or ADC. 1.6.2 STANDBY MODE – OPERATION The altimeter is fitted with a failsafe detection circuit, which automatically returns the altimeter to the standby mode under any one of the following conditions: 1.

AC power failure.

2.

Servo Motor failure.

3.

Amplifier failure.

4.

Detection circuit failure.

Difference at sea-level between the input signal and standby altimeter of more than 4,000ft (difference increases with an increase of altitude). Under these conditions, the main AC supply is isolated, the standby flag drops into view and the vibrator is energised. In addition to the circumstances listed above, the standby mode can be selected by momentarily setting the mode selector switch to “STANDBY”. This interrupts the supply and allows the self-maintaining relay to de-energise thus isolating the main supply. This action completes the DC supply circuit for the vibrator and returns the standby flag into view.


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A change in static pressure resulting from a change in altitude causes the capsule to expand or contract. This motion is then used to drive the indicator pointer and drum counter to indicate barometric altitude. Although the motor and control transformer are permanently connected to the gearing, because of their small size and low friction, they impose negligible additional friction upon the system. 1.6.3 DATUM PRESSURE SETTING The “Q” codes can be set into the altimeter using the millibar setting knob. The knob when turned adjusts the millibar scale, the capsule position and, via bevel gear and worm drive the stator of the CT. Thus the rotation of the setting knob causes simultaneous adjustment of the millibar scale, the capsule mechanism, the pointer and counter and the CT stator. It is necessary to simultaneously adjust the CT stator with the CT rotor (via capsule mechanism) to ensure that inputs from the master altimeter or ADC are not affected. 1.7 CABIN ALTIMETERS In addition to the aircraft altimeters, most passenger aircraft also carry a cabin altimeter. This is to enable the flight crew to monitor the pressurisation of the cabin environment control system. This type of instrument is a single pointer instrument with a range of zero feet to 20,000ft. The instrument‟s case is unsealed (vented to cabin pressure) and is normally only proved with compensation for temperature fluctuation. As a consequence, it suffers from errors due to changes in atmospheric conditions from the standard atmosphere to which it is calibrated. In spite of this, the accuracy of the instrument is better than 500ft, which is sufficient for its normal application.

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Figure 20 shows a Cabin Altimeter and sectioned view.

EN D P L A TE

T E M P E R A TU R E C OM PEN SA TOR

F IL T E R C IR C L IP

M E C H AN IS M P L A TE H A N D S T A FF & P I N IO N A S S E M B L Y

R O C K IN G S H A F T P O IN T E R

& S E C TO R A S S E M B L Y

Cabin Altimeter Figure 20


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1.8 AIRSPEED INDICATORS Airspeed is displayed in two ways, in nautical miles per hour, knots (1 nautical mile = 6,080ft, 1.5 miles), or as a factor of the speed of sound, Mach (Mach 1 = speed of sound). This information can be displayed separately, using an Airspeed Indicator (ASI) displaying airspeed in knots and a Mach meter (MM) displaying airspeed relative to the speed of sound, or both displays can be combined into a single instrument. 1.8.1 SIMPLIFIED AIRSPEED INDICATOR When an aircraft is stationary (on the ground) all external surfaces are subjected equally to the prevailing atmospheric pressure. When the aircraft is in motion, there are changes in the pressures felt on its external surfaces and the aircraft experiences a build up of an additional pressure on its leading edges resulting from its passage through the air. For any given height, the build up of this pressure (known as dynamic pressure) is proportional to the speed of the aircraft. This pressure when sensed by a Pitot tube, and ducted to an instrument, can be used to measure aircraft speed. Pitot pressure alone cannot be used to accurately measure speed, since no allowances is made for the thinning of the air at altitude. This would if left uncorrected lead to an apparent (indicated) loss of airspeed as altitude is increased. Measuring the difference in pressure between the dynamic pitot pressure, and the static pressure used to measure altitude compensates for this apparent loss of speed.

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A IRS P E E D

PI VOT

A T MO SPH ER IC PR ES SU R E

PI T OT PR ES SU R E

C A PSU L E S T AC K

A T MO SPH ER IC PR ES SU R E

S T AT IC P R E S S U RE

G E A R ING

P IT O T P R E S S U RE

P O IN T E R

C AP S UL E

Simplified Airspeed Indicator Figure 21


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The Airspeed Indicator in its most simple form consists of a sealed instrument case with a capsule which has pitot pressure applied to its inside while static pressure is fed to the case. The movement of the capsule is due only to the effects of the dynamic pressure, which results directly from the aircraft‟s speed through the air. Figure 22 shows two types of simple airspeed indicators.

DOUBLE POINTER AIRSPEED INDICATOR

SINGLE POINTER AIRSPEED INDICATOR

Airspeed Indicators Figure 22 Page 1-30



1.8.2 PITOT PRESSURE The Pitot pressure as sensed by the Pitot tube, is the sum of the dynamic pressure and the static pressure and can be represented by the formula:

P = ½V2 + S Where

-

P = Pitot Pressure.

-

 = air density.

-

V = aircraft velocity.

-

S = Static Pressure.

It can be seen from the above formula that the actual dynamic pressure build-up increases as the square of the aircraft‟s speed increases whereas the movement of the capsule has a linear response to pressure change. If therefore, as is normally required, the instrument scale is to be linear with respect to speed, and not compressed or cramped at low speeds, the square law pressure rise must be compensated for within the indicator. This is normally achieved using a ranging spring assembly as shown in figure 23.

RANGING SCREWS

RANGING PLATE

RANGING SPRING CASULE

Ranging Assembly (Square Law Compensation) Figure 23


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1.8.3 SPEED OF SOUND When an aircraft flies at or near the speed of sound, shockwaves build up around the aircraft due to the increased resistance of the air to the passage of the aircraft. The effect of these shockwaves are such that the aerodynamic stability of the aircraft is affected, resulting in buffeting, loss of directional control and loss of lift. The severity of these effects when flying at, near or through the speed of sound (sound barrier), is different for each type of aircraft but is always severe enough for the pilot to be forewarned via instrumentation that he is approaching the speeds at which these effects can be expected. The problems associated with the speed of sound are aggravated by the fact that the speed of sound varies with air density (altitude & temperature), as altitude increases the speed of sound decreases. Hence the need for a Machmeter, which indicates the aircraft‟s speed in relation to the speed of sound. This is indicated as a Mach number, Mach 1 = speed of sound at the altitude at which the aircraft is flying. Mach number can be represented by the formula:

Mach Number =

TRUE AIRSPEED LOCAL SPEED OF SOUND

This can be derived from: TRUE AIRSPEED (P - S) ALTITUDE (S)

When referring to aircraft flying speeds with respect to the speed of sound, there are three distinct speed bands:

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1.

Subsonic – Speeds up to 0.75 Mach.

2.

Transonic – Speeds from 0.75 to 1.20 Mach.

3.

Supersonic – Speeds in excess of 1.20 Mach.



1.8.4 MACHMETER Figure 24 shows a typical Machmeter.

Machmeter Figure 24 To enable the Machmeter to indicate aircraft speed as a factor of local or ambient speed of sound, the airspeed as measured by the instrument is modified by altitude. This is accomplished by using a different airspeed capsule operating in conjunction with an aneroid altitude capsule. These two being housed within a single instrument and coupled together in such a way that the Mach number indicated is increased with an increase in the aircraft‟s airspeed and further increased with an increase in the aircraft‟s altitude.


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Figure 25 shows a schematic of the Machmeter.

ALTITUDE CAPSULE

RETAINING SPRING

PUSH ROD SECTOR

ROCKING ARM

POINTER

VERTICAL LINK

HIGH

PITOT ENTRY

PIVOT

LOW AIRSPEED CAPSULE

Machmeter Schematic Figure 25 As can be seen from the diagram in figure 25, an increase of aircraft speed causes the dynamic pressure P-S to increase and the airspeed capsule to expand. This motion is then transmitted via the vertical link, rocking arm and sector arm to the pointer; causing it to move up the Mach number scale. A rise in altitude causes the altitude capsule to expand, this motion is transmitted to the rocking arm, via the rocking arm pivot, moving the rocking arm towards the centre line of the sector arm pivot. The rocking arm therefore moves closer to the pivot of the sector arm. This action modifies and increases the effect of the airspeed capsule causing the indicated Mach speed to be increased.

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1.8.5 COMBINED SPEED INDICATOR As aircraft become more and more complex the demand for instrumentation is continually rising. This has resulted, where practical, in two or more instruments being combined into one. This practice has been particularly successful with respect to airspeed and Mach speed indications. Two different examples of this are shown in figure 26.

PRESSURE OPERATED CSI

SERVO OPERATED CSI

Combined Speed Indicators Figure 26


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1.8.6 PRESSURE OPERATED CSI Figure 27 shows the schematic layout for the pressure operated Machmeter.

AIRSPEED ROCKING SHAFT AIRSPEED DIAL

HAIRSPRING BI-METALIC LINK

TUNING BLOCK

STATIC SECTOR

AIRSPEED CAPSULE

POINTER

PITOT

MACH DISC

ALTITUDE CAPSULE HAIRSPRING

SECTOR AIRSPEED DIAL

HAIRSPRING BI-METALIC LINK ALTITUDE ROCKING SHAFT

Pressure Operated Machmeter – Schematic Figure 27 The construction of the pressure operated combined speed indicator is very similar to the Machmeter discussed earlier. The main difference is that the altitude capsule mechanism is not connected to the airspeed capsule mechanism. The airspeed capsule and pointer operate as a conventional ASI indicating the actual airspeed of the aircraft by pointer against the outer airspeed dial.

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The altitude capsule is connected to a disc located behind the ASI pointer and inside the ASI scale. The Mach scale is printed on this disc. An increase in altitude, causes the altitude capsule to expand driving the Mach scale disc counter clockwise, whilst an increase in airspeed causes the pointer to move clockwise. The result of this is that an increase of airspeed and/or altitude produces an increase in the Mach number reading on the innerscale against the ASI pointer. 1.8.7 SERVO OPERATED CSI This instrument has a conventional ASI mechanism combined with a servocontrolled digital Mach speed counter providing the dual display. The servomechanism usually receives its control signals from the ADC. Because the Machmeter part of the instrument is power operated the instrument is provided with a power failure warning indicator. This normally takes the form of a power failure warning flag, or shutters which obscure the Mach digital counters in the failed mode. There is also a second pointer on this type of CSI and is known as the “Velocity Maximum Operating” (Vmo) pointer. This is provided for the purpose of indicating the maximum safe speed of an aircraft over its operating altitude range; in other words, it is an indication of the critical Mach number. This instrument also has a command bug and associated setting know in the bottom left hand corner of the instrument. This is used to set a required airspeed value, which can be used as the datum for an autothrottle control system, or as a fast/slow speed indicator. There are also five external index pointers around the bezel, which are manually set to any desired reference speed, i.e. take off speeds V1 and V2.


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1.9 VERTICAL SPEED INDICATORS The vertical speed indicator, commonly known as the rate of climb indicator, provides the flight crew with an accurate indication of the rate at which the aircraft is changing height. This indication is very necessary when flying on instruments only, at night or in poor visibility. 1.9.1 BASIC OPERATION The rate of climb (Vertical Speed Indication) is a measure of an aircraft‟s rate of altitude change, both climbing and descending. The instrument used is a further adaptation of the differential pressure instrument (Cabin Pressure), however, this time the pressure fed into the instrument case and into the capsule is static pressure (atmospheric pressure). The difference being that pressure to the case is fed through a restrictor. This has the effect of greatly reducing the rate at which the pressure in the case can change, whilst allowing the capsule to respond rapidly to any change in pressure. Figure 28 shows a simplified Vertical Airspeed Indicator.

CALIBRATED CHOKE

STATIC PRESSURE

STATIC TUBE

CAPSULE

Simplified Vertical Airspeed Indicator Figure 28

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Referring to figure 28, the pressure flow into and out of the case is restricted by a calibrated choke, when the aircraft climbs, the pressure in the capsule falls, maintaining a balance with external (to the aircraft) air pressure. The pressure within the instrument case also falls, but is unable to escape at the same rate as that from the capsule, causing a pressure differential to occur. The pressure within the instrument case being the greater when compared to the capsule. This causes the capsule to contract, and by a series of linkages the indicator pointer to indicate the rate of climb. The faster the change of altitude the greater the differential pressure, which results in a greater contraction of the capsule and a further deflection of the instrument pointer to indicate a greater rate of climb. Upon descent, the capsule pressure becomes greater than that of the instrument case and the capsule expands, causing the pointer to indicate a descent. In level flight the two pressures are in balance and the pointer indicates zero Figure 29 shows a typical vertical speed indicator.

Vertical Speed Indicator Figure 29 The rate of climb/descent is indicated by a single pointer moving over a dial face, which is graduated in feet per minute. The dial face, which can have either linear or logarithmic graduations, conventionally has a zero point situated at the 9-oclock position. The indicator pointer moves clockwise over the face to indicate ascent and anti-clockwise to indicate descent.


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1.9.2 CALIBRATION Calibration is set during manufacture and cannot be adjusted during servicing and testing. Calibration of the instrument is achieved by two calibration springs, which act on the centre of the capsule via a calibration stem. The forces exerted by the calibration springs are modified during calibration by two rows of screws, one row bearing onto the top spring and the other the bottom spring. Adjustment of the screws varies the effective length of the spring, which dependant upon capsule position will control the capsules response to pressure change and will therefore modify the indications produced. The upper spring controls the expansion of the capsule (rate of descent) and lower spring controls the compression of the capsule (rate of ascent). Figure 30 shows the inside of a vertical speed indicator showing the calibration springs.

CALIBRATION SPRINGS ROCKING SHAFT MECHANISM

CALIBRATION SCREWS

BALANCE WEIGHT LINK CALIBRATION SCREWS

METERING UNIT

CALIBRATION BRACKET

STATIC CAPILIARY TUBE

CAPSULE

Vertical Speed Indicator – Calibration Spring Figure 30

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1.9.3 ALTITUDE & TEMPERATURE COMPENSATION The way in which air passes through a metering device varies with air density and with temperature. Since the metering unit is required to give a given pressure difference for any given rate of altitude change, it must compensate for changes in air temperature and the change in air density at different altitudes. 

Altitude Compensation

Compensation for altitude changes is obtained by a combination of two basic metering devices, an orifice and a capillary tube. The pressure difference across an orifice for a given rate of altitude change decreases as altitude increases and therefore produces a negative error. Whereas, the pressure difference across a capillary tube for a given rate of altitude change increases as altitude increases, and therefore produces a positive error. Thus, the two effects tend to cancel each other. 

Temperature Compensation

The viscosity of the air is proportional to temperature; viscosity falling with a drop in temperature. The effects of this is that the pressure difference across an orifice for a given rate of altitude change increases with a decrease in temperature. Conversely, the pressure differential across a capillary tube for a given rate of altitude change decreases as temperature decreases. Thus, during design, a correct combination of orifice and capillary tubes can be chosen which will provide a stable pressure differential over a wide range of altitude and temperature changes. Figure 31 shows the internal working of a metering unit. GASKETS

AIR FILTER CAPILLARY

STATIC INPUT

ORIFICE

CONNECTING TUBE TO CAPSULE

Metering Unit Figure 31


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1.10 GYROSCOPIC INSTRUMENTS A number of instruments depend on the use of gyroscopes for their correct operation. It is useful to know the basic principles of how they work, before describing, in some depth, what they do. 1.10.1 GYROSCOPIC PROPERTIES As mechanical device a gyroscope may be defined as a system containing a heavy metal wheel (rotor), universally mounted so that it has three degrees of freedom: Spinning freedom:

About an axis perpendicular through its centre (axis of spin XX).

Tilting Freedom:

About a horizontal axis at right angles to the spin axis (axis of tilt YY).

Veering Freedom:

About a vertical axis perpendicular to both the other two axes (axis of veer ZZ).

The three degrees of freedom are obtained by mounting the rotor in two concentrically pivoted rings, called inner and outer rings. The whole assembly is known as the gimbal system of a free or space gyroscope. The gimbal system is mounted in a frame so that in its normal operating position, all the axes are mutually at right angles to one another and intersect at the center of gravity of the rotor. The system will not exhibit gyroscopic properties unless the rotor is spinning. When the rotor is spinning at high speed the device becomes a true gyroscope possessing two important fundamental properties: 1. Gyroscopic Inertia (Rigidity). 2. Precession. 1.10.2 RIGIDITY The property, which resists any, force tending to change the plane of rotor rotation. It is dependent on: 1. The mass of the rotor. 2. The speed of rotation.

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1.10.3 PRECESSION The angular change in direction of the plane of rotation under the influence of an applied force. The change in direction takes place, not in line with the force, but always at a point 90º away in the direction of rotation. The rate of precession also depends on: 1. The strength and direction of the applied force. 2. The angular velocity of the rotor. Figure 32 shows a gyroscope.

Z FRAME

Y

X

ROTOR OUTER RING

X Y

INNER RING

Z

Gyroscope. Figure 32


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Figure 33 shows the characteristics of gyro rigidity.

A

B C

Gyro Rigidity Figure 33 Gyro A has its spin axes parallel with the Earth's spin axes, located at the North Pole. It could hold this position indefinitely. Gyro B has its spin axes parallel to the Earth's spin axes, but located at the Equator. As the Earth rotates, it would appear to continually point North. Gyro C is also situated at the Equator. As the Earth rotates, it appears to rotate about its axes, however it is the Earth that is rotating and not the gyro. This rigidity can be used in a number of gyro instruments including the directional gyro.

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1.10.4 PRECESSION If an external force is applied to a spinning gyro, its effect will be felt at 90 0 from the point of application, in the direction of gyro rotation. This is known as precession. It can be seen in Figure 34, that if a force is applied to the bottom of the rotating wheel, it will rotate about its horizontal axis. This property is not wanted in some instruments, such as directional gyros. The use of precession is used in turn indicators, which will be covered later.

DIRECTION OF ROTATION

PRECESSION RATE = APPLIED FORCE 90º IN THE DIRECTION OF SPIN

SPIN AXIS 90º APPLIED FORCE

DIRECTION OF PRECESSION

Gyro Precession Figure 34


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1.10.5 VERTICAL GYRO Figure 35 shows the effects on a free gyro in an aircraft circling the earth. As can be seen, it would only be perpendicular to the earth's surface at two points.

Behaviour of a Vertical Gyro Figure 35

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In order for the gyro to be used to indicate the aircraft's attitude, it has to be corrected to continually be aligned to the vertical. These corrections are very slow and gentle, since the amount of correction needed, for example, in a tenminute period is small. Figure 36 shows a vertical gyro corrected to the local vertical.

Corrected Vertical Gyro Figure 36


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Instruments that use either the rigidity or the precession of gyros are: 1. Gyro Horizon Unit. 2. Attitude Director Indicator. 3. Standby Horizon Unit. 4. Direction Indicator. 5. Turn and Slip Indicator. 6. Turn Co-ordinator. 1.11 GYRO HORIZON UNIT The Gyro Horizon Unit gives a representation of the aircraft‟s pitch and roll attitudes relative to its vertical axis. For this it uses a displacement gyroscope whose spin axis is vertical. Figure 37 shows a displacement gyro and the two axis of displacement.

ROLL

PITCH

Displacement Gyro Figure 37 Page 1-48



Indications of attitude are presented by the relative positions of two elements, one symbolising the aircraft itself, the other in the form of a bar stabilized by the gyroscope and symbolising the natural horizon. Figure 38 shows a typical Gyro Horizon Unit.

3

6

6

SPERRY

3

Gyro Horizon Unit Figure 38 The gimbal system is so arranged so that the inner ring forms the rotor casing and is pivoted parallel to an aircraft‟s lateral axis (YY1); the outer ring is pivoted at the front and rear ends of the instrument case, parallel to the longitudinal axis (ZZ1). The element symbolizing the aircraft may either be rigidly fixed to the case, or it may be externally adjustable for setting a particular pitch trim reference.


Page 1-49


Figure 39 shows the construction of the Gyro Horizon unit.

X OUTER RING

ROTOR

Y

Z1

SYMBOLIC AIRCRAFT

BALANCE WEIGHT

PIVOT POINT

Z Y1 ROLL POINTER & SCALE

X1

HORIZON BAR

Construction of a Gyro Horizon Unit Figure 39 In operation the gimbal system is stabilized so that in level flight the three axes are mutually at right angles. When there is a change in the aircraft‟s attitude, example climbing, the instrument case and outer ring will move about the YY1 of the stabilized inner ring.

Page 1-50



The horizon bar is pivoted at the side and to the rear of the outer ring and engages an actuating pin fixed to the inner ring, thus forming a magnifying lever system. The pin passes through a curved slit in the outer ring. In a climb attitude the pivot carries the rear end of the bar upwards so that it pivots about the stabilized actuating pin. The front end of the bar is therefore moved downwards through a greater angle than that of the outer ring, and since the movement is relative to the symbolic aircraft element, the bar will indicate a climb attitude. Figure 40 shows climb attitude operation.

X

Z

1

Z

HORIZON BAR

X

1

Climb Attitude operation. Figure 40


Page 1-51


Changes in the lateral attitude of an aircraft, i.e. rolling, displaces the instrument case about the axis (ZZ1), and the whole stabilized gimbal system. Hence, lateral attitude changes are indicated by movement of the symbolic aircraft element relative to the horizon bar, and also by relative movement between the roll angle scale and pointer. Figure 41 shows roll attitude operation.

X

Y

Y

1

BANK TO PORT DATUM X

1

Roll attitude operation Figure 41 Freedom of gimbal system movement is 360º for roll axis and 85º for the and pitch axis. The pitch scale is restricted by means of a resilient stop. This will prevent gimbal lock.

Page 1-52



1.12 VERTICAL REFERENCE UNIT (VRU) The VRU consists of an electrically-driven gyroscope spinning about a vertical axis. The gyro has full freedom of movement in roll, and plus 85 degrees, minus 85 degrees of freedom in pitch. It also has an erection system for maintaining the rotor spin axis vertical. The VRU contains two synchros for detecting movement about the roll (aileron) and pitch (elevator) axes of the aircraft, and also contains circuitry for maintaining the functional operation of its internal components. Figure 42 shows the Vertical Reference Unit (RU).

VIBRATION ISOLATION MOUNTS

PITCH ERECTION CUTOFF SWITCHES

FRAME DEHYDRATION PLUG

GYRO CASE

ROLL ERECTION CUTOFF SWITCHES

ELECTRICAL CONNECTION BONDING STRAP

GIMBAL RING

Vertical Reference Unit (VRU) Figure 42


Page 1-53


1.13 ATTITUDE DIRECTOR INDICATOR (ADI) The ADI presents a symbolic three-dimensional display of the aircraft‟s attitude, combined with lateral and vertical steering commands. The aircraft‟s attitude is displayed by the relationship of a stationary airplane symbol with respect to a moveable horizon line. The horizon line is carried on a sphere, which is servo driven in pitch and roll. The sphere is marked off in increments of 5 degrees, and is coloured blue to represent sky above the horizon line, and black or brown/orange to represent ground below the horizon line. The sphere is unbalanced in the roll axis so that on loss of power it rotates to approximately 90 degree left bank indication. Cross pointer bars are used to indicate flight director commands and are brought into view by operation of the flight director switches (FD BARS). The horizontal (pitch) bar indicates below the miniature airplane symbol to command pitch up attitude. The vertical (roll) bar indicates to the right center display to command right roll, and to the left of center of display to command left roll. Both bars are biased out of view when the FD BARS are off, but the FD flag will not appear unless a power loss is experienced. Aircraft position relating to a glideslope is given by a pointer moving over a vertical display. Aircraft position above the glideslope beam is indicated by the pointer being positioned below the glideslope scale index, and aircraft position below the glideslope beam is indicated by the pointer being positioned above the glideslope beam. The loss of the glideslope valid signal will cause the glideslope warning flag (GS) to come into view. The glideslope indicator and warning flag are mounted on the right hand side of the ADI presentation. Localiser deviation is indicated by lateral movement of the localiser pointer, and is a read on a fixed horizontal scale. The pointer indicates to the right of the fixed scale index if the aircraft is to the left of the localiser beam and to the left of the index if aircraft is to the right of the localiser beam. The loss of either the localiser valid input or tuned to localiser input will bias the localiser pointer from view. Loss of the localiser valid signal causes the localiser (LOC) flag to move into view. The localiser indicator is positioned at the bottom of the ADI display, above the inclinometer. Slip information is conventionally displayed on the ball type inclinometer mounted on the indicator at the bottom of the ADI display. Instantaneous testing of the sphere and flight director is accomplished by pressing the TEST switch. The sphere should indicate:

Page 1-54

(a)

10° ± 5° Pitch Nose up.

(b)

20° ± 5° Roll to the Right.

(c)

ATT and FD flag in view.

(d)

FD Bars Indicate Nose Up and Roll to the Right.



Figure 43 shows an Attitude Director Indicator (ADI)

GSL

FD 2

F

1

1

S

2

AT

T

RW

Y

TEST

Attitude Director Indicator (ADI) Figure 43


Page 1-55


1.13.1 WARNINGS 1. ATT Flag Indicates an internal failure of the ADI or a Gyro Attitude (VRU) failure. 2. FD Flag Indicates an internal failure of the command bars for any axis or flight director failure. 3. LOC Flag Indicates a loss of the localiser valid signal, or insufficient signal with index off scale. 4. Glideslope Flag Indicates loss of the localiser valid (G/S) signal with index off scale. 1.13.2 ATTITUDE DISTRIBUTION Figure 44 shows a block schematic of the attitude transfer switching circuit and shows the distribution of the attitude information. The transfer switching is drawn in the „NORMAL‟ position fed from 28V ESS DC. Switching allows either gyro to supply both ADI attitude displays and the autopilot. The flight data recorder and weather radar are hard wired to No 1 gyro. Primary outputs are used exclusively for the ADI attitude displays. Buffered secondary 3 wire outputs are used for the autopilot, FDR and ADI crossswitching. The latter arrangement prevents a faulty ADI being paralleled with the other ADI thus causing the loss of both. The instrument comparator monitor (ICM) provides comparison of the ADI attitude displays. A two wire roll signal is also fed to the ICM to increase the heading warning threshold in turns.

Page 1-56



115V AC No 2

115V AC ESS

VERT GYRO 2

PRIMARY 3 WIRE & VALID

3 WIRE & VALID

SECONDARY

3 WIRE & VALID

SECONDARY

PRIMARY 3 WIRE & VALID

VERT GYRO 1

Attitude Distribution Figure 44

Page 1-57

26V AC NO 2

26V AC ESS

26V AC NO 2

ALL ON 1

P/R

P

AUTOPILOT

ADI No 1

ADI No 2

AUTOPILOT (RACO)

WX RADAR

APDU

FDR

WX (ARINC 708)

P/R

P/R

R

P/R

P/R

ALL ON 2

P/R

ERROR

ROLL THRESHOLD MONITOR

INSTRUMENT COMPARATOR MONITOR (ICM)

COMPARARATOR RESOLVER COMPARARATOR RESOLVER

26V AC ESS

JAR 66 CATEGORY B1 MODULE 11 BOOK 2 PART 1

AEROPLANE AERODYNAMICS, STRUCTURES & SYSTEMS


1.13.3 ATTITUDE TRANSFER SWITCHING

VERT GYRO No 2

115V AC (No 2)

VERT GYRO No 1

115V AC (ESS)

N 1

A.I.D.S.

WX RX

N

2 1 AUTOPILOT

CAPTAIN

FROM HSI COMPASS (LH)

INST. COMP. MONITOR

FROM HSI COMPASS (RH)

2

FIRST OFFICER

ROLL ANGLE CUTOUT (RACO)

The attitude transfer switching comprises of two banks of relays operated by a three position guarded switch on the left hand instrument panel, such, that „ALLON-1‟ or „ALL-ON-2‟ operation can be achieved. Operation of the switch to „ALLON-2‟ energise the LH bank of relays and Vice-Versa. Figure 45 shows a schematic of the Attitude Switching.

Attitude Switching Figure 45 Page 1-58



1.14 STANDBY ATTITUDE INDICATORS The standby attitude indicator provides a continuous visual indication of the aircraft attitude in the pitch and roll axes. 1.14.1 DESCRIPTION AND OPERATION The standby attitude indicator display comprises a two-coloured drum supported in an outer gimbal, a roll marker mounted on the outer gimbal shroud and a roll scale and aeroplane index mounted on the front cover behind the dial glass. A white line dividing the two colours on the drum, blue representing the sky and dark orange representing the earth, represents the horizon. Attitude is indicated by the position of the drum relative to the aircraft symbol. A graduated scale on the drum, which can indicate 60 degrees of dive or 80 degrees of climb, indicates pitch angle. Roll angle is indicated by a white marker relative to the roll scale which is graduated at zero degrees and 10, 20, 30, 40, 50 and 60 degrees left and right of zero. A fast erection knob is provided on the bottom right-hand side of the instrument face and is a purely mechanical caging device. Figure 46 shows a Standby Attitude Indicator and its location.

Standby Attitude Indicator H 301 Figure 46


Page 1-59


1.14.2 RUNNING UP 28 V DC is applied to the indicator, which produces a three, phase 19 V, 400 Hz supply to the stator winding of the gyro. The stator becomes energised and the gyro rotor begins to run up. When it has reached 18,000 rpm a sensor operates the gyro flag on the upper right part of the display causing it to disappear from view. This indicates that the gyro has attained a usable airspeed and there is a power supply to the unit. 1.14.3 ERECTION CONTROL Erection control is achieved through a single-pendulum mechanical erector device, which basically slaves the gyro erector assembly to the local vertical. Should the gyro axis deviate from the vertical axis, it will be acted upon by the erector device to cancel out this deviation and return the gyro to the vertical axis. The erection control consists of a reduction gear, erector bob-weight and a moving pendulum. Energy from the gyro is taken through a reduction gear to drive a gearwheel integral with the erector bob-weight. An assembly consisting of the erector bob-weight and moving pendulum is driven about the same shaft. The erector bob-weight is also driven about the reduction gear shaft and rotates at a speed of approximately 40-rpm. The moving pendulum is driven between two limits called the stop and driving plates. When the shaft is aligned with vertical axis the pendulum and bob-weight are in the horizontal plane. The pendulum is then forced against its driving plate by the function torque of its bearings, which counteracts the driving effect of the bob-weight. If the shaft deviates from the vertical axis, the pendulum is no longer in the horizontal plane. It will move erratically, the effect of which will be to bring the shaft into alignment with the vertical axis. 1.14.4 CAGING As the gyro runs up to speed, the gyroscopic assembly may occupy any random position inside its casing. Caging to case datum may be rapidly achieved and without abruptness, by pulling the fast erection knob approximately thirty seconds after energising the gyro. This brings the gyroscopic assembly to the vicinity of the vertical axis and when the knob is released it is free to move and aligns itself precisely with the vertical axis. 1.14.5 ATTITUDE INDICATION When the gyro is erected and running at full speed and the aircraft is in a level flight attitude, then the horizontal line on the datum and the roll pointer (which are both attached to the gyro mechanism) are aligned with the aeroplane index and the roll scale datum respectively. Because the gyro axis remains at the local vertical due to the gravity sensitive erection control system, movement of the aircraft (and therefore the instrument dial carrying the pitch datum and roll datum) from the vertical is relative to the gyro. Aircraft movement in the pitch axis causes Page 1-60



a vertical displacement between the horizon line and the aeroplane index; movement about the roll axis causes a rotational displacement between the horizon line and the aeroplane index and also between the roll pointer and the roll scale datum. Figure 47 shows a simplified circuit for the Standby Attitude Indicator.

0.3 A 1A

28V DC EMERG/BATT

STATIC INVERTER 19V AC 400 Hz

GYRO

ROTOR SPEED SENSOR

5V AC INSTRUMENT LIGHTING

Standby Attitude Indicator – Circuit Figure 47


Page 1-61


1.15 STANDBY ATTITUDE INDICATOR H 341 The attitude indicator type H341 is an electrically operated gyroscopic horizon assembly that provides a visual presentation of the aircraft‟s flight attitude in the pitch and roll axes. It is fitted with crossed pointers that display ILS deviations, and with an inclinometer for providing slip indication. The instrument operates from the aircraft 28 V DC supply; the 400 Hz 3-phase AC supply for the gyroscope is provided by a built-in static inverter. Figure 48 shows the Standby Attitude Indicator and its location.

Standby Attitude Indicator H 341 Figure 48

Page 1-62



1.15.1 DESCRIPTION The attitude display comprises a two-coloured spherical drum mounted on pivots, a roll pointer registering against a roll scale, and an aircraft symbol, the horizon is represented by the intersection of the two colours of the sphere; these are blue and brown, denoting sky and earth respectively. Attitude is indicated by the position of the sphere relative to the aircraft symbol. Pitch angle is indicated by a graduated scale on the sphere, the indication is limited to 65 degrees in dive and 105 degrees in climb. Roll freedom is unlimited and roll angle is indicated by the position of the roll pointer relative to the roll scale. Power failure or insufficient gyro rotational speed is indicated by the appearance of a flag in the upper righthand portion of the dial presentation. The flag is coloured fluorescent red, with four superimposed diagonal black stripes. After the gyro commences to run up, a fast erection mechanism is used to bring it to the vertical position. This is brought into operation by pulling the knob on the front of the instrument and waiting for a few seconds until the horizon line stabilises at its datum position and the roll index reads zero. Localiser and glideslope pointers indicate ILS deviation and are driven from No. 1 VHF navigation system. LOC and G/S failure warning flags are driven out of view by external 28 V DC validity signals also emanating from NAV 1 receiver; the flags are in view when the validity signals are missing or do not conform. When power is applied to the NAV 1 receiver but it is not tuned to a localiser frequency, external bias voltages remove the LOC and G/S pointers and flags from view. Figure 49 shows the Standby Attitude Indicator internal circuit.


Page 1-63


1A 28V DC EMERG/BATT

STATIC INVERTER

GYRO

G/S SIGNAL

LOC SIGNAL

G/S VALIDITY

LOC VALIDITY

5V AC INSTRUMENT LIGHTING

Standby Attitude Indicator – Internal Circuit Figure 49

Page 1-64



1.16 DIRECTION INDICATORS This indicator was the first gyroscopic instrument to be introduced as a “Heading Indicator” and although for most aircraft currently in service it has been superseded by remote-indicating compass systems (see later). The instrument uses a horizontal axis gyroscope and, being non-magnetic, is used in conjunction with a magnetic compass. In its basic form, the outer ring of the gyro carries a circular card, graduated in degrees, and referenced against a lubber line fixed to the gyro frame. When the rotor is spinning, the gimbal system and card are stabilized so that, by turning the frame, the number of degrees through which it is turning may be read on the card. Figure 50 shows a Directional Indicator.

180

170

Directional Indicator Figure 50


Page 1-65


In the directional gyro, the rotor is enclosed in a case, or shroud, and supported in an inner gimbal which is mounted in an outer gimbal, the bearings of which are located top and bottom on the indicator case. The front of the case contains a cut-out through which the card is visible, and also a lubber line reference. The caging/setting knob is provided at the front of the case to set the indicator onto the correct heading (magnetic). When the setting the heading, the inner gimbal has to be caged to prevent it from precessing as the outer gimbal is rotated. Figure 51 shows the construction of a directional gyro.

Directional Gyro Figure 51

Page 1-66



1.17 TURN & SLIP INDICATOR This indicator contains two independent mechanisms: 1. A gyroscopically controlled pointer mechanism for the detection and indication of the rate at which an aircraft turns. 2. A mechanism for the detection and indication of slip/slide. A gimbal ring and magnifying system, which moves the pointer in the correct sense over a scale calibrated in what is termed “Standard Rates”, actuate the rate of turn pointer. Although they are not always marked on a scale, they are classified as follows: 1. Rate 1 - Turn Rate 180º per minute. 2. Rate 2 - Turn Rate 360º per minute. 3. Rate 3 - Turn Rate 540º per minute. 4. Rate 4 - Turn Rate 720º per minute. Figure 52 shows a typical Turn & Slip indicator.

2 MIN

Turn & Slip Indicator Figure 52


Page 1-67


For the detection of rates of turn, a rate gyroscope is used and is arranged in the manner shown in figure 53.

INPUT AXIS

FWD Y1

X

F Y

P

X1

Rate Gyro Turn Indicator Figure 53 It differs in two respects from the displacement gyro as it only has one gimbal ring and a calibrated spring restraining in the longitudinal axis YY1. When the indicator is in its normal operating position the rotor spin axis, due to the spring restraint, will always be horizontal and the turn pointer at the zero datum. With the rotor spinning, its rigidity will further ensure that the zero position is maintained.

Page 1-68



When the aircraft turns to the left about the vertical input axis the rigidity of the rotor will resist the turning movement, which it detects as an equivalent force being applied to its rim at point F. The gimbal ring and rotor will therefore be tilted about the longitudinal axis as a result of precession at point P. As the gimbal ring tilts, it stretches the calibrated spring until the force it exerts prevents further deflection of the gimbal ring. Since precession of a rate gyro is equal to its angular momentum and the rate of turn, then the spring force is a measure of the rate of turn. Actual movement of the gimbal ring from its zero position can, therefore, be taken as the required measure of turn rate. 1.17.1 BANK INDICATION In addition to the primary indication of turn rate, it is also necessary to have an indication that an aircraft is correctly banked for the particular turn. A secondary indicating mechanism is therefore provided, which, depends for its operation on the effect of gravitational and centrifugal forces. A method commonly used for bank indication is one utilising a ball in a curved liquid-filled glass tube as shown in Figure 26. In the normal level flight the ball is held at the center of the tube by the force of gravity. Let us assume the aircraft turns left at a certain airspeed and bank angle. The indicator case and the tube move with the aircraft and centrifugal force (CF) in addition to that of gravity acts upon the ball and tends to displace it outwards from the center of the tube. However, when the turn is executed at the correct bank angle and matched with airspeed, then there is a balanced condition between the two forces and so the resultant force (R) hold the ball in the center of the tube. If the airspeed were to be increased during the turn, then the bank angle and centrifugal force would also be increased. As long as the bank angle is correct for the appropriate conditions, the new resultant force will still hold the ball central. If the bank angle for a particular rate of turn is not correct (under-banked/overbanked), then the aircraft will tend to either skid or slip. In the skid condition the centrifugal force will be the greatest, whereas in the slip condition the force of gravity is greatest.


Page 1-69


Figure 54 shows bank indication for various aircraft bank conditions.

Bank Indications Figure 54

Page 1-70



1.18 TURN CO-ORDINATOR The final instrument in this group is the turn co-ordinator. Basically, its mechanism is changed slightly from the turn and slip indicator, so that it senses rotation about the longitudinal axis, (bank) as well as the vertical axis, (turn). This gives a more accurate indication to the pilot, of the turning of the aircraft. Figure 55 shows a Turn co-ordinator indicator.

TURN COORDINATION

L

R 2 MIN NO PITCH INFORMATION

Turn co-ordinator Indicator Figure 55


Page 1-71


1.19 HORIZONTAL SITUATION INDICATOR (HSI) The HSI consists of a servo-driven azimuth (compass) card, which is read in relation to a miniature aircraft symbol in the center of the display, and a lubber line at the 12 o‟clock position, the azimuth card being driven by the gyrocompass system. A glideslope pointer and scale on the right hand side of the indicator gives a conventional display of the aircraft with respect to the glideslope. The scale is in a three-position rotation display, the three positions being glideslope, when the scale is presenting ILS glideslope deviation. The center mark is a rectangle and the outer marks are dots. No 2 scale is presented to display vertical navigation display and No 3 scale shows vertical navigation failure flag. The course deviation bar represents the centerline of a selected VOR or a selected localiser course. Deviation from a selected course is indicated by the bar moving across a scale, which is represented by four white dots, two on either side of the center of the rotatable mask. Two windows in the course mask show indications of: 1. To-From a VOR station, (a solid triangle with a V annotation). 2. To-From a selected NAV co-ordinate (a solid triangle with an N annotation). 3. To-From a station with aircraft ILS selected (a half-blue and half-yellow flag). 4. Failure flag (orange and yellow striped flag).

One radio bearing pointer displays the bearing to the next WPT. The bearing pointer is a pink arrow. A window to the left of the heading dial displays an ALERT annunciator flag to indicate the proximity of a navigation reference point. On the top left and top right hand corners of the instrument are to windows labeled DIST (distance to waypoint) and GND SPD (ground speed) respectively. Three windows located in the lower left hand corner of the instrument are blank until one of the auxiliary servo monitors detects a persistent excessive null, at which time the ISM causes the appropriate servo symbol to come into view. A cursor consisting of two trapezoids indicates the selected heading on the heading dial. The heading select indicator is remotely positioned by the heading (HDG) knob on the navigation selector. A heading (HDG) flag will be displayed,

Page 1-72



and will cover the heading index at the 12 o‟clock position if the heading source fails, or if there is interrupted supply. Selected course is displayed on the heading dial by an orange dagger-shaped indicator, rotating in the centre of the heading dial. A similarly coloured pointer opposite the dagger-shaped indicator provides the reciprocal of selected course. The dagger and pointer, together with the airplane symbol, serve as the index for the course deviation indicator. The course select indicator is remotely positioned by means of the course setting knobs on the navigation selector. Figure 56 shows a Sperry RD700D HSI.

SELECTED COURSE CAPTURED

SELECTED HEADING CAPTURED

SELECTED WAYPOINT BEARING CAPTURED

VERTICAL FAIL FLAG

GLIDESLOPE

COURSE MASK ANNUNCIATION'S

Sperry RD700D HSI Figure 56


Page 1-73

VERTICAL NAVIGATION


1.20 COLLINS 331A-8K HSI The HSI consists of a servo-driven azimuth (compass) card, which is read in relation to a miniature aircraft symbol in the center of the display and a lubber line at the 12 o'clock position. The azimuth card is driven by the gyro compass system. A vertical track or glideslope deviation pointer and scale, on the right-hand side of the HSI, gives a conventional display of the aircraft with respect to the glideslope. The deviation scale is marked by five dash marks, one long dash mark in the center, two short dash marks above it and two short dash marks below it. The vertical track or glideslope deviation pointer is such that when the aircraft is on the glidepath the pointer is in the central position on the scale. If the aircraft is off the glideslope, the pointer will move to indicate whether the deviation is up or down and the amount of movement indicates the extent of the deviation. The course deviation bar represents the centerline of a selected VOR or localizer course. The course deviation scale is marked by five dots, the center one being enclosed in a small circle. If the aircraft moves off course, the deviation bar will move to indicate whether left or right of selected course, and the amount of deviation. A To-From pointer is used when the navigation receiver is tuned to, and receiving a VOR signal. The to-from pointer indicates whether the selected course is "To" (pointer up) or "from" (pointer down) the received signal. When the selected course is the same as the selected VOR radial, and the aircraft is heading towards the signal course, a "to" indication is given. When the selected course is the same as the selected VOR radial and the aircraft is flying away from the signal course, a "From" indication is given. An RNAV bearing pointer indicates the direction to the active waypoint. When not in the RNAV mode, the pointer is biased to the 6 o'clock position. Two digital LCD displays in the top left-hand and right-hand corners of the HSI indicate the distance to go to the next waypoint (MILES) and the groundspeed of the aircraft (GND SPEED). The brightness of the two displays can be adjusted using the HSI & RA DIM control located on the main instrument panel.

Page 1-74



Figure 57 shows a Collins 331A-8K HSI.

Collins 331A-8K HSI Figure 57


Page 1-75


1.20.1 WARNING FLAGS 1. MAG annunciator. Is displayed to show that the information is magnetic heading. 2. HEADING warning flag. HEADING warning flag comes into view and covers the MAG annunciator if the heading information becomes unreliable. 3. Navigation warning flag. Comes into view if navigation data is (orange with white stripes) missing, or unreliable, when the receiver is tuned to a VOR station. 4. VERT warning flag (GS). Comes into view when glideslope data is missing or unreliable.

Page 1-76



1.21 ANGLE OF ATTACK (AOA) Apart from the main flight instruments, one item of information that the pilot needs to know at various stages of flight is the angle of attack. Earlier aircraft had a range of devices that gave the pilot indication of an approaching stall, which was an essential indicator but knowing the angle of attack has become an essential part of flying modern, larger aircraft. The simplest forms of angle of attack indicators are the Angle of Attack probe and the stall vane. The probe consists of a hinged-vane-type sensor mounted in the leading edge of a wing so that the vane protrudes into the airstream. Figure 58 shows an Angle of Attack vane sensor.

ELCTRICAL CONNECTION FWD

HINGED VANE

SYNCHRO

FUSELAGE SKIN

INDEX PINS

Angle of Attack Vane Sensor Figure 58


Page 1-77


In normal level flight conditions, the airstream maintains the vane in a parallel position. If the aircraft‟s attitude changes such that the AOA increases, then by definition, the airflow will meet the leading edge at an increasing angle, and so cause the vane to be deflected. Figure 59 shows the detection of the AOA.

A330

ANGLE OF ATTACK

VANE ARM ANGLE OF ATTACK TRANSDUCER

AIRCRAFT LONGITUDINAL AXIS

 FLIGHT PATH

AIRFLOW

Detecting AOA Figure 59 When the AOA reaches that which the warning unit has been pre-set, the vane activates a circuit to activate the stick shaker on the control column (Indicating the aircraft is approaching a stall).

Page 1-78



1.22 STALL WARNING INDICATION

VANE SENSOR SYNCHRO SUPP

HTR SUPP

K1 GND FLT

115V 400Hz

28V DC

WOW SW

STICK SHAKER

M SS1 FLAP POSITION TRANSMITTER

AOA SIGNAL

BIAS “OFF”

DEMODULATOR

Figure 60 shows the “Stall Warning System”

Stall Warning System Figure 60


Page 1-79


The system in figure 60 consists of a precision counter-balanced aerodynamic vane, which positions a synchro. The vane is protected against ice formation by an internal heating element. Also, since the pitch attitude of an aircraft is changed by the extension of the flaps, the sensor synchro is also interconnected with a synchro within the transmitter of the flap position indicating system, in order to modify the AOA signal output as a function of the flap position. Stick shaking is accomplished by a motor which is secured to the control column and drives a weighted ring that is deliberately unbalanced to set up vibrations of the column, to simulate the natural buffeting associated with a stalled condition. Figure 61 shows a stick-shaker installation.

MOUNTING BRACKET

STICK-SHAKER MOTOR

Stick-Shaker Installation Figure 61

Page 1-80



1.23 ELECTRONIC INSTRUMENT SYSTEMS Modern technology has enabled some significant changes in the layout of flight instrumentation on most aircraft currently in service. The biggest change has been the introduction of Electronic Instrument systems. These systems have meant that many complex Electro-mechanical instruments have now been replaced by TV type colour displays. These systems also allow the exchange of images between display units in the case of display failures. There are many different Electronic Instrument Systems, including: 1. Electronic Flight Instrument System (EFIS). 2. Engine Instrumentation & Crew Alerting System (EICAS). 3. Electronic Centralised Aircraft Monitoring (ECAM). Figure 62 shows a typical flight deck layout of an Airbus A320.

COMBINED AIRSPEED INDICATOR

EADI

ALTIMETER

BASIC “T” GROUPING WITH ELECTRONIC FLIGHT INSTRUMENTS

RADIO MAGNETIC INDICATOR

EHSI

EFIS PFD

EFIS ND

VERTICAL SPEED INDICATOR

ECAM ENGINE WARNINGS

EFIS ND

EFIS PFD

ECAM SYSTEMS

GLASS FLIGHTDECK - AIRBUS A320

Flight Deck Electronic Instrumentation Layout Figure 62


Page 1-81


The Electronic Instrument System (EIS) also allows the flight crew to configure the instrument layout by allowing manual transfer of the Primary Flight Display (PFD) with the Navigation Display (ND) and the secondary Electronic Centralized Aircraft Monitoring (ECAM) display with the ND. Figure 63 shows the switching panel from Airbus A320.

ATT HDG

AIR DATA

NORM CAPT 3

E/S DMC

NORM F/O 3

CAPT 3

ECAM / ND XFR

NORM F/O 3

CAPT 3

NORM F/O 3

CAPT

F/O

A320 EIS Switching Panel Figure 63

Page 1-82



As well as a manual transfer, the system will automatically transfer displays when either the PFD or the primary ECAM display fails. The PFD is automatically transferred onto the corresponding ND, with the ECAM secondary display used for the primary ECAM display. The system will also automatically transfer the primary ECAM information onto the ND if a double failure of the ECAM display system occurs. Figure 64 shows a block schematic of the EIS for the Airbus 320.

DISPLAY MANAGEMENT SYSTEM DMS No 1



Electronic Instrument System (EIS) Figure 64


Page 1-83


1.24 ELECTRONIC FLIGHT INSTRUMENT SYSTEM As in the case of conventional flight instrument systems, a complete EFIS installation is made up of left (Captain) and right (First Officer) systems. Each system comprises: 1. Electronic Attitude Director Indicator (EADI). 2. Electronic Horizontal Situation Indicator (EHSI). 3. Display Control Panel. 4. Symbol Generator. The EADI and EHSI can either be positioned side by side or vertically top and bottom. Normally the EADI is positioned on the top or on the onside position.

1.25 ELECTRONIC ATTITUDE DIRECTOR INDICATOR (EADI) The EADI displays traditional attitude information (Pitch & Roll) against a twocolour sphere representing the horizon (Ground/Sky) with an aircraft symbol as a reference. Attitude information is normally supplied from an Attitude Reference System (ARS). The EADI will also display further flight information, Flight Director commands right/left to capture the flight path to Waypoints, airports and NAVAIDS and up/down to fly to set altitudes. Information related to the aircraft‟s position with respect to Localizer (LOC) and Glideslope (GS) beams transmitted by an ILS. Auto Flight Control System (AFCS) deviations and Autothrottle mode, selected airspeed (Indicated or Mach No) Groundspeed, Radio Altitude and Decision Height information.

Page 1-84



Figure 65 shows a typical EADI display

Electronic Attitude Director Indicator (EADI) Display Figure 65


Page 1-85


The EADI has two display formats: 1. Full Time EADI Display (Data which is always present). 2. Part Time EADI Display (Data which is only present when active). 1.25.1 FULL TIME EADI DISPLAY DATA Attitude Sphere:

Moves with respect to the aircraft symbol to display actual pitch and roll attitude.

Pitch Attitude:

The pitch attitude display has white scale reference marks at 5, 10, 15, 20, 30, 40, 60 and 80 on the sphere.

Roll Attitude:

Displays actual roll attitude through a moveable index and fixed scale reference marks at 0, 10, 20, 30, 45, 60 and 90.

Aircraft Symbol:

Serves as a stationary symbol of the aircraft. Aircraft pitch and roll attitudes are displayed by the relationship between the fixed miniature aircraft and the moveable sphere.

Flight Director Cue:

Displays computed commands to capture and maintain a desired flight path. Flying the aircraft symbol to the command cue satisfies the commands.

Fast/Slow Display:

The pointer indicates fast or slow error provided by an angle-of-attack, airspeed or alternative reference system.

Inclinometer:

The EADI uses conventional inclinometer, which provides the pilot with a display of aircraft slip or skid, and is used as an aid for coordinated maneuvers.

Page 1-86



Attitude Source Annunciation:

The selected attitude source is not annunciated if it is the normal source for that indicator. As other attitude sources are selected, they are annunciated in white at the top left-hand side of the EADI. When the pilot and co-pilot sources are the same, then the annunciation is amber.

1.25.2 PART TIME EADI DISPLAYS Several displays are in view only when being used. When not in use, these displays are automatically removed from the EADI. Radio Altitude:

Displayed by a four-digit display from –20 to 2500 feet. Display resolution between 200 and 2500 feet is in 10-foot increments. The display resolution below 200 is 5 feet. The display disappears for altitudes above 2500 feet (Radio Altitude max altitude is 2,500 feet).

Decision Height:

Decision Height is displayed by a three-digit display. The set range is from 0 to 990 feet in 10-foot increments. The DH display may be removed by rotating fully counterclockwise the DH set knob.

Note; when the Radio Altimeter height is 100 feet above the DH, a white box appears adjacent to the radio altimeter display. When at or below the DH, an Amber DH will appear inside the white box. Flight Director Mode Annunciators:

Flight director vertical and lateral modes are annunciated along the top of the EADI. Armed vertical and lateral modes are annunciated in white to the left of the captured vertical and lateral mode annunciators. Capture mode annunciators are displayed in green and are located on the top center for lateral modes and in the top right corner for vertical modes. As the mode's transition from armed to capture, a white box is drawn around the capture mode annunciator for 5 seconds.


Page 1-87


Marker Beacon:

Displayed above the Radio Altimeter height information. The markers are of a specified colour of: Blue

-

Outer Marker.

Amber

-

Middle Marker.

White

-

Inner Marker.

Rising Runway:

a miniature rising runway displays Absolute altitude reference above the terrain. It appears at 200 feet, and contacts the aircraft symbol at touchdown (0 feet).

Rate-of-Turn:

Pointer and scale at the bottom of the display indicates rate or turn. Used with the inclinometer, will enable coordinated turns to be achieved.

Glide Slope:

By tuning to an ILS frequency, the Glide Slope information will be displayed. Aircraft displacement from the Glide Slope beam centerline is then indicated by the relationship of the aircraft to the Glide Slope pointer. The letter “G” inside the vertical scale pointer identifies the information as Glide Slope deviation. When tuning to other than an ILS frequency, the Glide Slope display is removed.

Expanded Localizer:

By tuning to an ILS frequency, the Rate-of-Turn display is replaced by the expanded Localizer display. When tuning to other than an ILS frequency, the expanded localizer display is replaced by the Rate-ofTurn display.

Vertical Navigation Display:

Page 1-88

The deviation pointer indicates the VNAV‟s computed path center to which the aircraft is to be flown. In this mode, the letter “V” inside the vertical scale pointer identifies the information as VNAV deviation.



1.26 ELECTRONIC HORIZONTAL SITUATION INDICATOR (EHSI) The EHSI presents a selectable, dynamic colour display of flight progress and plan view orientation. The EHSI has a number of different modes of operation, these are selectable by the flight crew and the number will be dependent on the system fitted.

Figure 66 shows an EHSI display.

Electronic Horizontal Situation Indicator (EHSI) Display Figure 66


Page 1-89


The EHSI has two display formats: 1. Full Time EADI Display (Data which is always present). 2. Part Time EADI Display (Data which are only present when active). 1.26.1 FULL TIME EHSI DISPLAYS Aircraft Symbol:

The aircraft symbol provides a quick visual cue as to the aircraft‟s position in relation to the selected course and heading, or actual heading.

Heading Dial:

Displays the heading information on a rotating heading dial graduated in 5 increments. Fixed heading indexes are located at each 45 position.

Heading “Bug” & Heading Readout:

Course Deviation Indicator:

Select Course Pointer & Course Readout:

Page 1-90

The notched heading bug is positioned around the rotating heading dial by the remote heading select knob on the Display Controller. A digital heading select readout is also provided for convenience in setting the heading bug. Heading select error information from the heading bug is used to fly to the bug.

The course deviation bar represents the centerline of the selected navigation or localizer course. The aircraft symbol pictorially shows the aircraft position in relation to the displayed deviation.

Course pointer is positioned inside the heading dial by the remote select knob on the Display Controller. Course error information from the course select pointer is used to fly the selected navigation path. A digital course select readout is provided for convenience in setting the select course pointer.



Distance Display:

Navigation Source Annunciators:

Time-to-Go/Ground Speed:

The distance display indicates the nautical miles to the selected DME station or LRN Waypoint. Depending on the equipment, the distance will be displayed in a 0 to 399.9 NM or a 0 to 3999 NM format. An Amber “H” adjacent to the distance readout indicates DME Hold. This will indicate to the crew that DME information is from the previous VOR/DME beacon, and not the one providing VOR bearing.

Annunciation of the navigation source is displayed in the upper right hand corner. Long range navigation sources such as INS, VLF, RNAV and FMS are displayed in blue to distinguish them from short-range sources, which are annunciated in white.

Either Time-to-Go or Groundspeed can be displayed, selected via the Display Controller. Ground Speed is calculated using the LRN, if fitted. If no LRN, then the EFIS uses the DME distance to calculate Ground Speed.

Drift Angle Bug:

The drift angle bug with respect to the lubber line represents drift angle left or right of the desired track. The drift angle bug with respec to the compass card represents actual aircraft track. The bug is displayed as a magenta triangle that moves around the outside of the compass card.

Desired Track:

When LRN is selected, the Course Pointer now becomes the Desired Track Pointer. The position of the desired Track Pointer is controlled by the LRN. A digital display of desired track (DRAK) is displayed in the upper left-hand corner.

TO-FROM Annunciator: An Arrowhead in the center of the EHSI indicates whether the selected course will take the aircraft TO or FROM the station or Waypoint. The TO-FROM annunciator is not in view during ILS operation.

Heading Source Annunciation:

At the top center of the EHSI is the heading source annunciator.


Page 1-91


Heading SYNC Annunciator:

The heading SYNC annunciator is located next to the upper left corner and indicates the state of the compass system in the slaved mode. The bar represents commands to the compass gyro to slew to the indicated direction (+ for increased heading and 0 for decreased heading). Heading SYNC is removed during compass FREE mode and for LRN derived heading displays.

1.26.2 PART TIME EHSI DISPLAYS Vertical Navigation Display:

Glide Slope Deviation:

Bearing Pointer Source Annunciators:

Elapsed Time Annunciation:

Page 1-92

The vertical navigation display comes into view when the VNAV mode on the flight director is selected. The deviation pointer then indicates the VNAV‟s computed path center to which the aircraft is to be flown. In this mode the letter “V” inside the scale pointer identifies the deviation display. The Glide Slope display comes into view when a VHF NAV source is selected and the NAV source is tuned to an ILS frequency. The deviation pointer then indicates the Glide Slope beam center to which the aircraft is to be flown. The letter “G” inside the scale pointer identifies the deviation display.

The bearing pointers indicate relative bearing to the selected NAVAID. Two bearing pointers are available and can be tuned to either VOR or ADF NAVAIDs. If no NAVAIDs are selected then the pointers and annunciators are removed. The bearing source annunciators are colour and symbol coded with the bearing pointers.

When in the Elapsed Time (ET) mode, the ET display can read minutes and seconds or hours and minutes. The hour/minute mode will be distinguishable from the minute/second mode by an “H” on the left of the digital display.



1.26.3 PARTIAL COMPASS FORMAT The partial compass mode displays a 90 ARC of compass coordinates. The Partial mode allows other features such as MAP and Weather Radar displays to Be selected. Figure 67 shows a Partial EHSI display (Compass Mode).

EHSI Partial Compass Mode Display Figure 67 Wind Vector Display:

Wind information is displayed in any partial format. The wind information can be shown as magnitude and direction or as head/tail component and cross wind component, type used is determined on installation of EFIS. In both cases, the arrow shows the direction and the number indicates the velocity of the wind (in knots). Wind information is calculated from the LRN.


Page 1-93


Range Rings:

Range rings are displayed to aid in the determining the position of radar returns and NAVAIDs. The range ring is the compass card boundary and represents the selected range on the Radar.

NAVAID Position:

NAVAID position can be selected during MAP mode. The source of the NAVAID position marker is selected and annunciated in conjunction with the associated bearing source and is colour coded.

Weather Information:

Weather information from the Radar can be displayed in partial compass mode. Weather Radar data is presented in the following colours:

Page 1-94

1.

Black

-

No storm.

2.

Green

-

Moderate storm.

3.

Yellow

-

Less severe storm.

4.

Red

-

Severe storm.



Figure 68 shows an EHSI partial format with Weather Radar information.

EHSI Weather Radar Display Figure 68


Page 1-95


1.26.4 MAP MODE The MAP mode will allow the display of more navigational information in the partial compass mode. Information on the location of Waypoints, airports, NAVAIDs and the planned route can be overlaid on the compass mode. Weather information can also be displayed in the MAP mode to give a very comprehensive display. Figure 69 shows an EHSI MAP mode display.

EHSI MAP Mode Display. Figure 69

Page 1-96



1.26.5 COMPOSITE DISPLAY In the event of a display unit failure, the remaining good display can display a “Composite Display”. This display is selected via the Display Controller and is basically a display consisting elements from an EADI and EHSI display. Figure 70 shows a typical composite display.

EFIS Composite Display Figure 70


Page 1-97


1.27 EFIS CONTROLLER Allows the crew to select the required display configuration and what information is to be displayed. Both Captain and Co-Pilot have their own display controller‟s. The controllers have two main functions: Display Controller:

Selects the display format for EHSI as either FULL, ARC, WX or MAP.

Source Select:

Selects the system that will provide information required for display. The source information will be VOR, ADF, INS, FMS, VHF and NAV.

EFIS Display Controllers are shown at Figure 71.

DISPLAY SELECT BUTTONS

FULL ARC

GS TTG

WX

CRS

DIM

ET

DH

MAP

BOT

SC CP

REV

HDG

TOP

TEST

RASTER DIM

DISPLAY CONTROLLER

SOURCE SELECT BUTTONS

NAV

VLF

FMS

INS 1

INS 2

HDG

ATT

VOR 2

ADF 2 ADF 1

VOR 1 ADF 2

AUTO

ADF 1

OFF

OFF

BRG

BRG

SOURCE SELECT CONTROLLER

EFIS Display and Source Controllers Figure 71 Page 1-98



1.27.1 DISPLAY CONTROLLER FULL/ARC:

The FULL/ARC button is used to change the EHSI display from full compass rose display to a partial compass display format. Successive pushes of the button change the display format back and forth between FULL and ARC.

WX (Weather):

The WX button is used to call up weather radar returns on the partial compass display. If the EHSI is in the FULL display format, selecting the WX display will automatically select the ARC format. A second push of the WX button will remove the weather information but the ARC format will remain.

GS/TTG:

By pressing the GS/TTG button, Groundspeed or the Time-to-GO will alternately be displayed in the lower right corner of the EHSI.

ET:

By pressing the ET button, Elapsed time is displayed. If the ET button is pressed again, it will zero the displayed time. The sequence is: 1. Zero. 2. Start. 3. Stop.

MAP:

By pressing the MAP button, the full compass display is changed to the partial compass display, with active Waypoints displayed. Also VOR/DME ground station positions will be displayed.

SC/CP:

By pressing the SC/CP button, the flight director command cues are toggled back and forth from single cue (SC) configuration to cross pointer (CP) configuration.


Page 1-99


REV:

In the event of an EADI/EHSI display failure, the REV button may also be used to display a composite format on the remaining good display. The first push of the button will blank the EHSI and put the composite display onto the EADI. The second push blanks the EADI and puts the composite display onto the EHSI. A third push will return EHSI/EADI to normal.

CRS Select Knob:

Rotation of the Course select knob allows the course pointer on the EHSI to be rotated to the desired course.

DIM:

Rotation of the outer concentric DIM knob allows the overall brightness of the EADI, EHSI to be adjusted. After the reference levels are set, photoelectric sensors maintain the brightness level over various lighting conditions.

DH:

Rotation of the inner concentric DH knob allows the Decision Height, displayed on the EADI, to be adjusted. If the knob is rotated fully counterclockwise, the DH display is removed.

TEST:

By pressing the TEST button, the displays will enter the test mode. In the test mode, flags and cautions are presented along with a check of the flight director mode annunciations. If the test is successful a “PASS” is displayed. If the test is unsuccessful then an “FD FAIL” is annunciated.

RASTER DIM TOP/BOT: Rotation of the outer (Bottom display) and inner (Top display) concentric knobs adjusts the raster scan display (Weather Radar and Attitude Sphere). HDG:

Page 1-100

Rotation of the heading select knob allows the heading select bug to be rotated to the desired heading.



1.27.2 SOURCE CONTROLLER Used to select the available sources of heading, attitude, bearing and navigational information for display. Since each aircraft is different, the source controller is normally tailored to fit each need. NAV:

This button is used to control the source of VHF NAV display information. Each push of the button will toggle the source between pilot and copilot‟s NAV information. VHF systems include DME, ILS and VOR.

LRN:

Long Range Navigation selections depend on the systems available. These include INS, VLF and FMS systems.

ATT:

Attitude button selects the source of attitude information. Each push of the button will select a different source for display. Not available to all aircraft.

BRG:

This knob allows the selection of VOR and ADF bearings to be displayed. The selected source is annunciated on the left-hand side of the display and the bearing to the selected beacon via two bearing pointers.


Page 1-101


The EFIS comprises the following units: 1. Symbol Generator (SG). 2. Display units X 2 (EADI & EHSI). 3. Control Panel. 4. Remote Light Sensor. Figure 72 shows the EFIS units and signal interface in block schematic form. Honeywell GS

ATT 2 AOA F

20

20

10

10

10

10

G GS

WX

TTG DIM

CRS

DH

SC

MAP

BOT

REV

CP

TOP

S CMD M .99 200DH

HDG

TEST RASTER DIM

AIR DATA COMP NAV

VLF

FMS

INS 1

INS 2

ADF 2 AUTO

ATT

HDG

I

140RA

Honeywell

VOR 2

CRS +0

OFF

N 33

H 2.1 NM 3

30

BRG

BRG

NAV 1

345

ADF 1

OFF

DH

6

VOR 1

ADF 1

E 1 2

INERTIAL REF SYSTEM

20

EFIS SG No 1

VOR 1 ADF 2

ADF 1

20

W 24

ARC

ET

21

S

HDG

NAV AID ILS/VOR

15

FULL

GSPD

013

130 KTS

EFIS SG No 3 RAD ALT Honeywell GS

ATT 2

WEATHER RADAR

AOA F

20

20

10

10

10

10

G S CMD M .99 200DH

DME FULL ARC

DIM

CRS

FMS

GS TTG

WX

ET

DH

MAP

BOT

SC CP

REV

TOP

20

20 DH

140RA

HDG

TEST RASTER DIM

EFIS SG No 2

AFCS

Honeywell FMS

INS 1

INS 2

CRS

ATT

HDG

NAV 1

345 +0

AUTO

VOR 1

BRG ADF 1

HDG

E 1 2

OFF

BRG

S

013

EFIS Block Schematic Figure 72

Page 1-102

H 2.1 NM 3

30

ADF 1

OFF

N 33

VOR 2

6

VOR 1 ADF 2

W 24

ADF 2 ADF 1

21

VLF

15

GPWS

NAV


GSPD

130 KTS


1.28 OTHER SYSTEM INDICATIONS There are endless different instrument displays, which show the pilot's or flight engineer, the condition of the aircraft's many systems, the range of instruments depending on the size of the aircraft. On earlier airliners there could have been dozens of instruments on the panels to pass on information regarding, for example, oil temperature & pressure, cabin altitude, hydraulic oil quantity, electrical power being used, etc. 1.29 POWERPLANT INSTRUMENTATION Information required by the flight crew to enable them to monitor the engines include: 1. Fuel Contents. 2. Fuel Flow. 3. Engine RPM. 4. Engine Temperature. 5. Engine pressure. 1.30 FUEL CONTENTS GAUGE Most modern aircraft have a number of fuel tanks within the wing structure and each individual tank's contents must be known. There are two main methods of indicating fuel contents: 1. Resistance Gauges. 2. Capacitance Quantity Indicators. 1.30.1 RESISTANCE GAUGES This type of gauge tends to found on smaller aircraft. It has a float in the fuel tank that is connected to a variable resistor. As the fuel level changes, the float will move, thus changing the resistance, which in turn will alter the current flow through a DC circuit, which in turn will operate a meter indicating fuel contents.


Page 1-103


Figure 73 shows a simplified resistance gauge.

INDICATOR

N S

TANK RESISTOR

+ DC POWER

FUEL TANK

Resistance Gauge Figure 73 1.30.2 CAPACITANCE QUANTITY INDICATORS This has the advantage over other quantity systems in that it can give accurate readings in very large or unusually shaped tanks. The probes within the fuel tank are actually capacitors. The two plates of the capacitor will be separated by fuel on the lower end and air on the upper end. Since fuel and air have different dielectric constant values, the amount of capacitance will change as the fuel level rises and falls. The probes will then send signals to the flight deck gauges to indicate fuel contents. This system usually includes a totalizer, which will give a reading of the total fuel on board. Some fuel systems will also include indications of fuel used since take-off.

Page 1-104



Figure 74 shows a circuit of a capacitance quantity system.

TANK UNIT

EMPTY

IS

LOOP A

IB

LOOP B

REF C FULL

2 - PHASE MOTOR

DISCRIMINATION STAGE

AMPLIFIER STAGE INDICATOR

REF PHASE

AMPLIFIER UNIT

Capacitance Quantity Indicating System Figure 74


Page 1-105


1.31 FUEL FLOW INDICATOR Although the amount of fuel consumed during a given flight may vary slightly between engines of the same type, fuel flow does provide a useful indication of the satisfactory operation of the engine and the amount of fuel being consumed during flight. A typical system consists of a fuel flow transmitter, which is fitted into the low pressure fuel system, and an indicator, which shows the rate of fuel flow and the total fuel used in pounds per hour. The transmitter measures the fuel flow electrically and an associated electronic unit gives a signal to the indicator proportional to the fuel flow. Figure 75 shows a fuel flow transmitter & indicator.

Fuel Flow Transmitter & Indicator Figure 75

Page 1-106



In some aircraft, a combined fuel flow/pressure indicator is used. This type of indicator usually has two pointers moving over the double-scale, one pointer the right engine and the other the left engine. On most large passenger aircraft, a dedicated fuel flow indicator is used for each engine. Figure 76 shows two types of fuel flow indicators.

TWIN ENGINED COMBINED FUEL FLOW/PRESSURE INDICATOR

SINGLE ENGINE FUEL FLOW INDICATOR

Fuel Flow Indicators Figure 76


Page 1-107


1.31.1 FUEL FLOW TRANSMITTERS There are two types of fuel flow transmitters currently in use: 1. Synchronous Mass Flow Flow-meter System. 2. Motorless Mass Flow Meter System. 1.31.2 SYNCHRONOUS MASS FLOW FLOW -METER SYSTEM This system measures mass flow rather than volume. In this way, it compensates for fuel temperature in its readout. The system also measures in pounds per hour. Figure 77 shows a schematic diagram of the Synchronous Mass Flow Flow-meter System.

CALIBRATED RESTRAINING SPRINGS

TURBINE

DECOUPLING DISK

IMPELLER FUEL FLOW

IMPELLER MOTOR

FLUID PASSAGE FLUID PASSAGE 115V 400Hz

TRANSMITTER

MOTOR CIRCUIT

INDICATOR

Synchronous Mass Flow Flow-meter System Figure 77

Page 1-108



Referring to figure 77, fuel enters the transmitter impeller, which is rotated at a constant 60rpm by the synchronous impeller motor. The temperature of the fuel will determine its volume and the amount of force to be created by the action of the impeller. The turbine is twisted against its retaining springs by the mass flow force created by impeller movement. The mass flow electrical transmitter arrangement works on the principle of a torque synchro. 1.31.3 MOTORLESS MASS FLOW METER SYSTEM The motorless flow meter represents the latest in electronic solid-state fuel measuring systems. It is small in size and accounts for variables such as fuel temperatures and specific gravity with an accuracy of 1% as opposed to 2% for motor driven flow meters. Almost all the large turbine powered aircraft are configured with the motorless type, pound per hour fuel flow meter system. Figure 78 shows a schematic of the Motorless Mass Flow Meter.

DRUM PICK-OFF COIL 1

t

PICK-OFF COIL 2

DRIVE

FUEL FLOW

IMPELLER MAGNET ONE

SPRING

MAGNET TWO

Motorless Mass Flow Meter Figure 78


Page 1-109


Referring to figure 78, The flow meter transmitter converts the rate into two electronic signals. The signals are created as the flowing fuel gives an angular displacement to two continuously rotating magnets. The magnets induce electronic impulses into stationary coils and the time difference is used as a measure of the mass flow rate. The fuel enters from the drive end and rotates the drum containing magnet 1 and the drive shaft. The spring connects the drive shaft to the impeller containing magnet 2. As the magnets rotate, the pick-off coils receive current pulses, the first pulse occurring at pick-off coil 1. Then as the spring deflects in proportion to fuel flow, magnet 2 turns with the impeller and induces a current pulse with a time lag into pick-off coil 2. The greater the mass flow, the greater the spring deflection and angular difference between the magnets. The time displacement which, results is directly proportional to mass flow rate in this motorless transmitter design. The indicator contains electronic circuits, which convert the time difference to a pound per hour readout.

Page 1-110



1.32 PRESSURE INDICATORS It is essential for the correct and safe operation of the engine that accurate indication is obtained of both the temperature and pressure of the engine oil and fuel supply. Figure 79 shows a fuel pressure indicator and an engine oil pressure indicator.

FUEL PRESSURE INDICATOR

ENGINE OIL PRESSURE INDICATOR

Fuel & Oil Pressure Indicators Figure 79 MOD 11 BOOK 2 PART 1 ISSUE 6 - 01/02/11

Page 1-111


There are two methods of detecting the pressure, these are: 1. Pressure Capsule detection. 2. Bourbon Tube detection. 1.32.1 PRESSURE CAPSULE DETECTION This type of indicator utilizes a “pressure capsule” or “diaphragm”. Like the bourbon tube, a diaphragm type pressure indictor is attached to a capillary tube, which attaches to the fuel system and carries pressurised fuel to the diaphragm. As the diaphragm becomes pressurised it expands, causing an indicator pointer to rotate. Figure 80 shows a pressure capsule type fuel pressure indicator.

DIAPHRAGM

Capsule Type Pressure Indicator Figure 80

Page 1-112



1.32.2 BOURDON TUBE DETECTION The bourdon tube is made with a metal tube that is formed in a circular shape with a flattened cross-section. One end is open while the other is sealed. The open end of the bourdon tube is connected to a capillary tube containing the pressurised medium. As the pressurised medium enters the bourdon tube, the tube tends to straighten. Through a series of gears, this movement is used to move the indicating pointer on the instrument face. Figure 81 shows a Bourdon tube mechanism.

POINTER STAFF

BOURDON TUBE

ANCHOR POINT GEARING

Bourdon Tube Mechanism Figure 81


Page 1-113


The most common method used on modern passenger aircraft is a pressure transmitter and indicator. The operation is that oil/fuel pressure acting on a bourdon tube within the transmitter moves an electromagnet core. This movement is then transmitted to the indicator via a torque synchro system, moving a pointer over the calibrated pressure scale. Figure 82 shows a schematic of this system.

PRESSURE INPUT 26 V AC

ENGINE FIREWALL

BOURDON TUBE

FLIGHTDECK PRESSURE INDICATOR PRESSURE TRANSMITTER

Pressure Transmitter and indicator Figure 82 .

Page 1-114



1.33 OIL & FUEL TEMPERATURE INDICATORS There are two main types of temperature sensors used in the oil & fuel temperature measurement, these are: 1. Resistive Bulb Sensor. 2. Thermocouple Sensor. 1.33.1 RESISTIVE BULB SENSOR Oil & fuel temperatures are sensed by a temperature sensitive element (resistive BULB), fitted in the oil and fuel system. A temperature sensor and indicator is shown in figure 82.

28V DC

TEMPERATURE BULB (RESISTIVE TYPE) OIL TEMPERATURE INDICATOR

CONNECTOR PINS

MICA INSULATOR

MICA CORE

NICKEL WINDING ON MICA CORE

COMPENSATING COIL

Temperature Sensor & Indicator Figure 82


Page 1-115


The indicator contains a Wheatstone Bridge circuit with the temperature sensor as the variable resistance. A change rise in temperature causes a rise in the resistance value and, consequently, unbalances the bridge network with a corresponding flow of current at the indicator. The indicator pointer is deflected by an amount equivalent to the temperature change and this is recorded on an indicator calibrated in degrees centigrade. 1.33.2 THERMOCOUPLE SENSOR The advantage of the thermocouple sensor over the resistive bulb type is that it requires no power from the aircraft electrical system to operate, It is selfcontained and self-generating circuit. It derives its power from a pair of dissimilar metals, iron and constantan, which when heated at the hot junction, produces a millivoltage and causes a current flow through the meter. Figure 83 shows the thermocouple sensor and indicator,

CONSTANTAN (-) (YELLOW)

IRON (+) (BLACK)

THERMOCOUPLE HOT JUNCTION

OIL TEMPERATURE INDICATOR

Thermocouple & Indicator Figure 83

Page 1-116



1.34 ENGINE RPM INDICATORS All engines have their rotational speed (rpm) indicated, on two spool or triple spool engines, the high pressure assembly speed (N3) is always indicated; in most cases, additional indicators show the speed of the low pressure (N1) and intermediate pressure (N2) assemblies. Engine speed indication is electrically transmitted from a small generator driven by the engine to an indicator that shows the speed as a percentage of the maximum engine speed. Figure 84 shows the compressor speeds for a triple spool engine.

INTERMEDIATE PRESSURE INTERMEDIATE SPEED N2 COMPRESSOR

LOW PRESSURE LOW SPEED N1 COMPRESSOR

HIGH PRESSURE HIGH SPEED N3 COMPRESSOR

Compressor Spool Speeds Figure 84 The engine speed is often used to assess the engine thrust, but it does not give an absolute indication of the thrust being produced because inlet temperature and pressure conditions affect the thrust at a given engine speed.


Page 1-117


Figure 85 shows two types of engine rpm indicators.

ENGINE RPM INDICATOR

N1 PERCENTAGE INDICATOR

Engine Speed Indicators Figure 85

Page 1-118



1.34.1 ENGINE SPEED GENERATOR The Engine speed generator supplies a three-phase alternating current, the frequency of which is dependent upon engine speed. The generator output frequency controls the speed of a synchronous motor in the indicator, and rotation of a magnet assembly housed in a drum or drag-cup induces movement of the drum and consequent movement of the indicator pointer. Figure 86 shows an engine speed generator & indicator.

SYNCHRONOUS MOTOR FIELD POINTER YOKE SPOOL DRIVE

FLUX COUPLING SPRING

INDICATOR

GENERATOR

GENERATOR FIELD GENERATOR OUTPUT

Engine Speed Generator & Indicator Figure 86 Where there is no provision for driving a generator, a variable-reluctance speed probe, in conjunction with a phonic wheel, may be used to induce an electric current that is amplified and then transmitted to an indicator. This method can be used to provide an indication of rpm, without the requirement for a separately driven generator, with its associated drives, thus reducing the number of components and moving parts of the engine.


Page 1-119


Figure 87 shows a variable-reluctance speed probe and phonic wheel system.

SPEED SIGNAL TO AMPLIFIER

COMPRESSOR CASE

DRIVE SHAFT

SPEED PROBE

PHONIC WHEEL

Variable-Reluctance Speed Probe & Phonic Wheel Figure 87 The speed probe is positioned on the compressor casing in line with the phonic wheel, which is a machined part of the compressor shaft. The teeth on the periphery of the wheel pass the probe once every revolution and induce an electric current by varying the magnetic flux across a coil in the probe. The magnitude of the current is governed by the rate of change of the magnetic flux and is thus directly related to the engine speed. In addition to providing an indication of rotor speed, the current induced at the speed probe can be used to illuminate a warning lamp on the instrument panel to indicate to the flightcrew that a rotor assembly is turning. This is particularly important at engine start, because it informs the flightcrew when to open the fuel cocks to allow fuel to the engine.

Page 1-120



1.35 EXHAUST TEMPERATURE INDICATING The temperature of the exhaust gases is always monitored closely during engine operation, especially during the starting cycle when overheat damage is most prevalent. Hot section temperature is considered the most critical of all engine operating parameters because an out of limits condition can render an engine unairworthy in a matter of seconds. There are a number of different locations that the exhaust temperature can be measured and thus a number of different indicators such as: 1. Turbine Inlet Temperature (TIT) – Indicates the temperature is being monitored forward of the turbine wheel(s). 2. Interstage Turbine Temperature (ITT) – Indicates the temperature is being taken at some intermediate position between multiple turbine wheels. 3. Exhaust Gas Temperature (EGT) – Indicates the temperature is being taken aft of the turbine wheels. 4. Turbine Outlet Temperature (TOT) – Indicates the temperature is being taken aft of the turbine wheels. Figure 88 shows a typical EGT indicator.

OVER TEMP LIMIT BUG OVERTEMP WARNING LIGHT

EGT Indicator Figure 88 MOD 11 BOOK 2 PART 1 ISSUE 6 - 01/02/11

Page 1-121


Each type of EGT system consists of several thermocouples spaced at intervals around the circumference of the engine exhaust section casing. The EGT indicator in the cockpit displays the average temperature measured by the individual thermocouple probes. The thermocouple probes consist of two wires of dissimilar metals that are joined together inside a metal guard tube. Transfer holes in the tube allow the exhaust gas to flow across the junction. The metals from which the thermocouple wires are made are usually nickel-chromium and nickel-aluminium alloys. Figure 89 shows a thermocouple with figure 90 showing a typical thermocouple harness.

NICKEL ALUMINIUM WIRE

NICKEL CHROMIUM WIRE

TRANSFER HOLES

Thermocouple Figure 89

Page 1-122



TO GAS TEMPERATURE CONTROL SYSTEM

AIR INTAKE THERMOCOUPLE

JUNCTION BOX

JET PIPE THERMOCOUPLES


Thermocouple Harness Figure 90


Page 1-123


1.36 ENGINE PRESSURE INDICATORS The Engine Pressure Ration (EPR) system has for many years been the most widely used thrust indicating system for aircraft flight deck indication. The EPR is used as a performance (thrust) setting instrument on many flight decks. The EPR is ratio of two engine pressures: Turbine discharge total pressure and compressor inlet pressure. Each manufacturer uses a slightly different engine station numbering system, and engine stations are a means of identifying engine pressure ratio tap off points. For example Pratt & Whitney uses station 2 (Pt2) and station 5 (Pt5), to identify the engine pressure ratio tap-off points of singlespool engines. They also use stations 2 and 7 (Pt2) & (Pt7), to identify the engine pressure ratio tap-off points of a dual-spool engine. Figure 91 shows a typical EPR indicator.

EPR Indicator Figure 91

Page 1-124



1.36.1 EPR FORMULA

= 1.3 19.11 14.70 EPR =

Pt2 = 14.70 PSI (ABSOLUTE)

Pt2 PROBE

PRESSURE RATIO TRANSMITTER

EPR INDICATOR

Pt7 MANIFOLD

Pt7 = 19.11 PSI (ABSOLUTE)

The following example is of a Pratt & Whitney JT12 engine EPR cockpit indications. When turbine discharge pressure is 19.11 pounds per square inch – absolute and the compressor inlet pressure is 14.7 pounds per square inch – absolute, the EPR will be 1.3. Figure 92 shows the EPR system and the calculation of the example in this paragraph.

EPR System Figure 92


Page 1-125


1.37 VIBRATION INSTRUMENTS A turbo-jet engine has an extremely low vibration level and a change of vibration, due to an impending or partial failure, may pass without being noticed. Many engines are therefore fitted with vibration indicators that continually monitor the vibration level of the engine. The indicator is usually a milliammeter that receives signals through an amplifier from engine mounted transmitters. Figure 93 shows a vibration transmitter and indicator.

ENGINE VIBRATION MEASURED IN MILS (THOUSANDTHS) OF INCHES

VIBRATION TRANSMITTER

Vibration Transmitter & Indicator Figure 93

Page 1-126



The vibration level recorded on the indicator is the sum total of vibration felt at the pick-up. A more accurate method differentiates between in the frequency ranges of each rotating assembly and so enables the source of vibration to be isolated. This is particularly important on m multi-spool engines. A crystal-type vibration transmitter, giving a more reliable indication of vibration, has been developed for multi-spool engines. A system of filters in the electronic circuit to the indicator makes it possible to compare the vibration source. A multiple selector switch enables the pilot to select a specific area to obtain a reading of the level of vibration. Figure 94 shows a multiple-selector vibration indicator.

Multiple-Selector Vibration Indicator Figure 94


Page 1-127


FUEL PRESSURE OIL TEMPERATURE

OIL PRESSURE

ENGINE PRESSURE RATIO

ENGINE VIBRATION

COMPRESSOR SPEED

MONITOR

FUEL FLOW

EXHAUST GAS TEMPERATURE

CORE SPEED

Figure 95 shows the functional diagram of a large engine indicating system.

Large Engine Indicating System Figure 95

Page 1-128



Figure 96 shows a typical LED type electronic engine instrumentation group for a four engine aircraft.

LED Electronic Instrument group Figure 96


Page 1-129


1.38 ELECTRONIC INSTRUMENTS (ENGINE & AIRFRAME) With the introduction of the "Glass Cockpits", most traditional gauges, instruments and warning lights have been replaced by fully electronic display systems. There are different types of display systems available, the two main ones being: 1. Engine Instrument and Crew Alerting System (EICAS). 2. Electronic Centralized Aircraft Monitoring (ECAM). 1.39 ENGINE INDICATING & CREW ALERTING SYSTEM (EICAS) The basic system comprises two display units, a control panel and two computers supplied with analog and digital signals from the engine and system sensors. The computers are designated “Left” and “Right” and only one is in control of the system at any one time, the other is held in standby. In the event of a failure, it may be switched in either manually or automatically. Operating in conjunction with the system are discrete caution and warning lights, standby engine indicators and a remotely-located panel for selecting maintenance data display. The system provides the flight crew with information on primary engine parameters (Full-time), with secondary engine parameters and advisory/caution/warning alert messages displayed as required.

Page 1-130



1.39.1 DISPLAY UNITS These units provide a wide variety of information relevant to engine operation, and operation of other automated system. The operation of these displays is the same as those in the EFIS as previously described. The upper unit displays primary engine parameters, i.e. N1 speed, EGT, and warning and caution messages. The lower unit displays secondary parameters, i.e. N2 speed, fuel flow, oil quantity, pressure and temperature. In addition, the status of non-engine systems e.g. flight control surface position, hydraulic system, APU, etc., can be displayed. On the upper unit, a row of Vs will appear when secondary information is being displayed on the lower unit. Seven colors are produced by the CRTs for displaying information. Table 1 shows the colors and description of there uses.

Colour White Red Green Blue Yellow Magenta Cyan

Description All scales, normal operating range of pointers, digital readouts. Warning messages, maximum operating limit marks on scales, and digital readouts. Thrust mode readout and selected EPR/N1 speed marks or target cursors. Testing of system only. Caution and advisory messages, caution limit marks on scale, digital readouts During in-flight engine starting, and for cross bleed messages. Names of all parameters being measured (e.g. N1, oil pressure, TAT, etc.) and status marks or cues. Table 1


Page 1-131


Figure 97 shows layout of the EICAS Displays.

CAUTION RESET CANCEL

0 1

SBY

1013 2

8 X 100 ft

7

UPPER DISPLAY (PRIMARY)

3

3 5 0 00 5

6

4

LOWER DISPLAY (SECONDARY) -

COMPUTER BRT

DISPLAY

ENGINE STATUSEVENT RECORD

THRUST REF SET BOTH

L AUTO R

L

R

MAX IND RESET

EICAS Primary and Secondary Display Formats Figure 97

Page 1-132



Figure 98 and 99 show display formats for primary and secondary displays.

CAUTION

TAT 15°c 0.0

0.0

10

CANCEL RECALL

6

10 2

6

2

N1 0

0

EGT

VVVVVVV

Primary EICAS Display Figure 98


Page 1-133


50

50

OIL

PRESS

120

120

OIL

TEMP

18

18

OIL

88

88.00 N2 86

86

N3 4.4

4.4

QTY

N1

FAN

3.1

1.9

FF

VIB

Secondary EICAS Display Figure 99

Page 1-134



1.40 DISPLAY MODES EICAS is designed to categorize displays and alerts according to the function and usage. For this purpose there are three modes of displaying information: 1. Operational (selected by the flight crew). 2. Status (selected by the flight crew). 3. Maintenance (ground use only and selected via the maintenance panel). 1.40.1 OPERATIONAL MODE This mode displays the engine operating information and any alerts required to be actioned by the crew in flight. Normally only the upper display unit presents information: the lower one remains blank and can be selected to display secondary information as and when required. 1.40.2 STATUS MODE When selected this mode displays data to determine the dispatch readiness of an aircraft, and is closely associated with details contained in the aircraft‟s Minimum Equipment List. The display shows the positions of the flight control surfaces in the form of pointers registered against vertical scales, selected sub-system parameters, and equipment status messages on the lower display unit. Selection is normally done on the ground, either as part of the pre-flight checks of dispatch items, or prior to shutdown of electrical power to aid the flight crew in making entries in the aircraft‟s Technical log. Figure 100 shows an example of a status page. 1.40.3 MAINTENANCE MODE This mode provides maintenance engineers with information in five different display formats to aid them in fault finding and verification testing of major subsystems.


Page 1-135


HYD QTY

L 0.99

C 1.00

HYD PRESS

2975

3010 3000

APU

EGT 440

OXY PRESS

RPM 103

0.0

R 0.98

FF

0.0

OIL 0.75

1750

RUD

AIL ELEV AIL

EICAS Status Page Figure 100

Page 1-136



1.41 DISPLAY SELECT PANEL To control the operation of the EICAS, a control panel is situated on the center pedestal. Figure 101 shows a typical EICAS control panel.

COMPUTER

DISPLAY

BRT BRT

ENGINE

STATUS

EVENT RECORD

BAL

L AUTO R

L BOTH R

MAX IND RESET

EICAS Control Panel Figure 101


Page 1-137


1.41.1 DISPLAY SELECT PANEL OPERATION Engine Display Switch:

This is a push type switch for removing or presenting the display of secondary information on the, lower display.

Status Display Switch:

This is a push type switch for removing or presenting the status page on the lower display.

Event Record Switch:

Normally, there is an auto event function, this will automatically record any malfunctions as they occur. The push switch enables manual event marking so that the crew can record a suspect malfunction for storage in a non-volatile memory. This data can be retrieved from the memory and displayed by ground engineers by operating the ground maintenance panel. This manual switch can also be used for activating the recording of fault data, either in the air or on the ground, on the Environmental Control system, Electrical Power system, Hydraulic system and APU.

Computer Select Switch: In the “AUTO” position it selects the left or primary computer and automatically switches to the other in the event of a failure. The other positions are for manually selecting either the right or left computers. Display Brightness:

Thrust Reference Set Switch:

Max Indicator Reset:

Page 1-138

Controlled by the inner knob for the display intensity, the outer for display brightness.

Pulling and rotating the inner knob positions the reference cursor on the thrust indicator display (either EPR or N1) for the engines, which are selected by the outer knob. If any of the measured parameters e.g. Oil Pressure, EGT etc. and if they exceed normal operating limits, this will be automatically alerted on the display units. The purpose of the reset button is to clear the alerts from the display when the excess limits no longer exist.



1.42 ALERT MESSAGES The system will continually monitor a large number of inputs (400+) from engine and airframe systems. If a malfunction is detected then the appropriate alert message is annunciated on the upper display. Up to 11 messages can be displayed and are at the following levels: LEVEL A - Warning:

Requiring immediate corrective action and are displayed in “RED”. Master warning lights are also activated and aural warnings from the Central Warning System are given.

LEVEL B - Caution:

Requiring immediate crew awareness and possible action. They are displayed in “AMBER”. An aural tone is also repeated twice.

LEVEL C - Advisory:

Requiring crew awareness, displayed in “AMBER”. There are no caution lights or aural tones associated with this level.


Page 1-139


Figure 102 shows a display with the three different types of alert messages Displayed.

LEVEL A WARNING

LEVEL B CAUTION

LEVEL C ADVISORY

TAT 15°c APU FIRE R ENGINE FIRE CABIN ALTITUDE C SYS HYD PRESS R ENG OVHT AUTOPILOT C HYD QTY R YAW DAMPER L UTIL BUS OFF

70.0

110.0

10 6

10 2

6

2

N1 999

775

EGT

VVVVVVV

Upper EICAS Display – Alert Messages Figure 102

Page 1-140



1.43 MAINTENANCE CONTROL PANEL This panel is used by maintenance engineers for the purpose of displaying maintenance data stored within the system‟s computer memories. Figure 103 shows a typical maintenance control panel.

PERFORMANCE AND AUXILLIARY POWER UNIT FORMATS ENVIRONMENTAL CONTROL SYSTEM AND MAINTENANCE MESSAGE FORMATS

ELECTRICAL AND HYDRAULIC SYSTEM FORMAT

EVENT READ

EICAS MAINT DISPLAY SELECT

ECS

ELEC

PERF

MSG

HYD

APU

CONF MCDP

CONFIGURATION AND MAINTENANCE CONTROL/DISPLAY PANEL

SELECTS DATA FROM AUTO OR MANUAL EVENT IN MEMORY

AUTO

MAN

REC

ERASE

ENG EXCD

ENGINE EXCEEDANCES

TEST

BITE TEST SWITCH FOR SELF-TEST ROUTINE

ERASES STORED DATA CURRENTLY DISPLAYED RECORDS REAL-TIME DATA CURRENTLY DISPLAYED (IN MANUAL EVENT)

Maintenance Control Panel Figure 103


Page 1-141


1.44 ELECTRONIC CENTRALIZED AIRCRAFT MONITORING ECAM differs from EICAS in that the data displayed relate essentially to the primary systems of the aircraft and are displayed in checklist and pictorial or synoptic format. 1.44.1 DISPLAY UNITS These can be mounted either side-by-side or top/bottom. The left-hand/top unit is dedicated to information on the status of the system; warnings and corrective action in a sequenced checklist format, while the right-hand/bottom unit is dedicated to associated information in pictorial or synoptic format. Figure 104 shows the layout of ECAM displays.

350

400

8 4

300

MACH

60 1 0 9

80

250

120 IAS KNOTS

240 220

200

140 180

5

LDG GEAR GRVTY EXTN

5

RESET OFF DOWN

ECAM Display Layout Figure 104

Page 1-142



1.45 ECAM DISPLAY MODES There are four display modes, three of which are automatically selected and referred to as phase-related, advisory (mode and status), and failure-related modes. The forth mode is manual and permits the selection of diagrams related to any one of 12 of the aircraft‟s systems for routine checking, and also the selection of status messages provided no warnings have been triggered for display. Selection of the displays is by means of a system control panel. See Figure 81 1.45.1 FLIGHT PHASE RELATED MODE In normal operation the automatic flight phase-related mode is used, and the displays will be appropriate to the current phase of aircraft operation, i.e. Preflight, Take-off, Climb, Cruise, Decent, Approach, and post landing. Figure 105 shows display modes. The upper display shows the display for pre-take off, the lower is that displayed for the cruise.

E N G IN E 10

5

8 7. 0

5

10

F .US E D

6 5. 0

N1

1530

F O B : 1 4 0 0 0 KG

%

KG

10

6 50

5

EG T

QT Y

10

S

4 80

F LA P

80 1500

% FF K G /H

N O S M O K IN G : S E A T B E LT S :

1 1 .5

V IB 1 .2

(N 2 ) 1 .3

1 1 .5

A IR

ºC N2

F

(N 1 ) 0 .9

1530

O IL 5

V IB 0 .8

L DG E LE V A U TO

500F T

F UL L

8 0 .2

C AB V /S

1 5 00

C K PT 2 0

FWD 2 2

A FT 23

24

22

24

C AB A LT FT 4150

ON ON

S P LR S :

F UL L

F LA P S :

F UL L

F T/M IN

250

L D G I N H I B IT A PU B LEED

T AT +1 9 ºC S A T + 17 ºC

E C A M U P P E R D IS P L A Y

G .W . 6 0 3 0 0 K G 2 3 H 56

C .G . 2 8 .1 %

E C A M L O W E R D IS P L A Y - C R U IS E

ECAM Upper and Lower Display (Cruise Mode) Figure 105


Page 1-143


1.45.2 ADVISORY MODE This mode provides the flight crew with a summary of the aircraft‟s condition following a failure and the possible downgrading of systems. Figure 106 shows an advisory message following a Blue Hydraulic failure.

10

5

87.0

650

ADVISORY MESSAGES

80 1500

65.0

N1 %

10

5

10

5

FOB : 14000KG 10

5

EGT ºC

S

480

N2 %

80.2

FF KG/H

1500

F

FULL

HYD B RSVR OVHT B SYS LO PR

FAILURE MESSAGES

FLAP

FLT CTL SPOILERS SLOW

1 FUEL TANK PUMP LH

ECAM Advisory Mode Figure 106

Page 1-144



1.45.3 ECAM FAILURE MODE The failure-related mode takes precedence over the other modes. Failures are classified in 3 levels LEVEL 3: WARNING This corresponds to an emergency configuration. This requires the flight crew to carry out corrective action immediately. This warning has an associated aural warning (fire bell type) and a visual warning (Master Warning), on the glare shield panel. LEVEL 2: CAUTION This corresponds to an abnormal configuration of the aircraft, where the flight crew must be made aware of the caution immediately but does not require immediate corrective action. This gives the flight crew the decision on when action should be carried. These cautions are associated to an aural caution (single chime) and a steady (Master Caution), on the glare shield panel. LEVEL 1: ADVISORY This gives the flight crew information on aircraft configuration that requires the monitoring, mainly failures leading to a loss of redundancy or degradation of a system, e.g. Loss of 1 FUEL TANK PUMP LH or RH but not both. The advisory mode will not trigger any aural warning or attention getters but a message appears on the primary ECAM display.


Page 1-145


Figure 107 – 111 shows the 12-system pages and status page available.

C ON D

TE M P ºC

C AB P R E S S

L DG E LE V

AP PS I

V/ S F T /M IN

A LTN M O DE F AN

2

8 0

F W D 22

24

22

C

H

C

A FT 2 3

0

4 .1

C

10

1150 0

2

4150

DN

24 H

50 0 F T C A B AL T FT

UP

F AN

C KP T 2 0

MAN

M AN

SY ST 1

H

SY ST 2

SA F ET Y

VE NT HOT

IN L ET

EX T RA C T

A IR

PA C K 1

T AT +1 9 ºC S A T + 17 ºC

2 3 H 56

G .W . 6 0 3 0 0 K G

T AT +1 9 ºC

C .G . 2 8 .1 %

S A T + 17 ºC

A IR C O N D IT IO N IN G S Y S T E M P A G E

PA C K 2

G .W . 6 0 3 0 0 K G 2 3 H 56

C .G . 2 8 .1 %

P R E S S U R IS A T IO N S Y S T E M P A G E

ECAM System Displays Figure 107 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.

Page 1-146



BAT 1 28 V 15 0A

E LE C DC

D C B AT

F/ CT R

BAT 2 28 V 15 0A

1

DC

G B Y

2

D C ES S TR 1 28 V 15 0A

AC 1

GE N 1 26 % 11 6V 40 0H Z

ES S TR 28 V 13 0A

EM ER G GEN 11 6V 40 0H Z

A C ESS

A PU 26 % 11 6V 40 0H Z

T AT +1 9 ºC S A T + 17 ºC

2 3 H 56

SP D BR K

TR 2 28 V 15 0A

L A IL B G

P IT CH TR IM

R A IL G B

G Y

3 .2 º UP

AC 2

EX T PW R 11 6V 40 0H Z

R UD G B Y

L EL EV B G

GE N 2 26 % 11 6V 40 0H Z

G .W . 6 0 3 0 0 K G

T AT +1 9 ºC

C .G . 2 8 .1 %

S A T + 17 ºC

E L E C T R IC A L S Y S T E M P A G E

R EL EV Y B

G .W . 6 0 3 0 0 K G 2 3 H 56

C .G . 2 8 .1 %

F L IG H T C O N T R O L S Y S T E M P A G E

ECAM System Displays Figure 108 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.


Page 1-147


FU E L K G

F .U SE D 1

1550

F .U SE D 2

A PU

H YD

1550

F OB

GR EE N

3000

L E FT

1 07 5 0

5 60 0

T AT +1 9 ºC S A T + 17 ºC

Y E LL O W

2 3 H 56

PSI

3000

PSI

3000

R IG H T

C TR

5 50

B LU E

2 87 5 0

1 07 5 0

5 50

G .W . 6 0 3 0 0 K G

T AT +1 9 ºC

C .G . 2 8 .1 %

S A T + 17 ºC

FUE L SYSTE M P AG E

G .W . 6 0 3 0 0 K G 2 3 H 56

C .G . 2 8 .1 %

H Y D R A U L IC S Y S T E M P A G E

-ECAM System Displays Figure 109 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.

Page 1-148



B LE E D

WHEEL

2 0 ºC

2 4 ºC C

C

H R AM A IR

5 0 ºC

170 1

ºC REL

140

140

2

3

ºC REL

LO

HI

140 4 2

GN D A PU

AUTO BRK

23 H 56

2 3 0 ºC LO

HI

1

TAT +19 ºC SAT +17 ºC

H

LP

G.W. 60300 KG

T AT +1 9 ºC

C.G. 28.1 %

S A T + 17 ºC

LANDING GEAR/WHEEL/BRAKE SYSTEM PAGE

HP

HP

LP

G .W . 6 0 3 0 0 K G 2 3 H 56

C .G . 2 8 .1 %

A IR B LE ED SY STE M PA GE

ECAM System Displays Figure 110 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually. The Gear/Wheel page is displayed at the related flight phase.


Page 1-149


A PU

O X Y 1 85 0 P S I

D OOR

A PU ARM

ARM

A V I O N IC

26% B LE E D

116 V

C A BIN

35 PSI

400 HZ

FW D C OM PT C A RG O

N 10 ARM

EM ER

ARM

E X IT

0

%

80

F LA P O P E N

C A RG O EG T

B U LK ARM

ARM

5

7

ºC

C A BIN 3

T AT +1 9 ºC S A T + 17 ºC

580

T AT +1 9 ºC 2 3 H 56

S A T + 17 ºC

C .G . 2 8 .1 %

D O O R /O X Y S Y S T E M P A G E

2 3 H 56

C .G . 2 8 .1 %

AP U SY STEM PAG E

ECAM System Displays Figure 111 Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually. Related flight phase.

Page 1-150



1.46 CONTROL PANEL The layout of the control panel is shown in Figure 112.

DISPLAY ON & BRIGHTNESS CONTROL

DISPLAY ON & BRIGHTNESS CONTROL

SGW SELECT SWITCHES

1

TOP DISPLAY

OFF

ECAM

SGU

2

FAULT

FAULT

OFF

OFF

BOTTOM DISPLAY

BRT

OFF

BRT

MESSAGE CLEARANCE SWITCH CLR

STS

RCL

STATUS MESSAGE SWITCH

RECALL SWITCH

ENG

HYD

AC

DC

BLEED

COND

PRESS

FUEL

APU

F/CTL

DOOR

WHEEL

SYSTEM SYNOPTIC DISPLAY SWITCHES

ECAM Control Panel Figure 112


Page 1-151


1.46.1 ECAM CONTROL PANEL SGU Selector Switches: Controls the respective symbol generator units. Lights are off in normal operation of the system. The “FAULT” caption is illuminated amber if the SGU‟s internal self-test circuit detects a failure. Releasing the switch isolates the corresponding SGU and causes the “FAULT” caption to extinguish, and the “OFF” caption to illuminate white. System Synoptic Display Switches: Permit individual selection of synoptic diagrams corresponding to each of the 12 systems, and illuminate white when pressed. A display is automatically cancelled whenever a warning or advisory occurs. CLR Switch: Light illuminates white whenever a warning or status message is displayed on the left-hand display unit. Press to clear messages. STS Switch: Permits manual selection of an aircraft‟s status message if no warning is displayed. Illuminates white when pressed also illuminates the CLR switch. Status messages are suppressed if a warning occurs or if the CLR switch is pressed. RCL Switch: Enables previously cleared warning messages to be recalled provided the failure conditions which initiated the warnings still exists. Pressing this switch also illuminates the CLR switch. If a failure no longer exists the message “NO WARNING PRESENT” is displayed on the left-hand display.

Page 1-152



PART TWO CONTENTS 2

AVIONICS SYSTEMS ................................................................... 2-1 2.1 2.2

2.3 2.4 2.5 2.6 2.7 2.8 2.9 2.10 2.11 2.12 2.13 2.14 2.15

2.16

2.17 2.18 2.19 2.20 2.21 2.22 2.23 2.24

2.25 2.26

AUTOMATIC FLIGHT ...................................................................... 2-1 AUTOPILOT SYSTEM ..................................................................... 2-1 2.2.1 Error Sensing ................................................................ 2-2 2.2.2 Correction ..................................................................... 2-2 2.2.3 Follow-Up ...................................................................... 2-2 2.2.4 Command ..................................................................... 2-2 AUTOPILOT INTERLOCKS .............................................................. 2-4 SERVOMOTORS ............................................................................ 2-6 SINGLE AXIS CONTROL SYSTEM ................................................... 2-8 TWO-AXIS SYSTEM ....................................................................... 2-8 THREE-AXIS SYSTEM.................................................................... 2-8 SENSING ATTITUDE CHANGES....................................................... 2-8 AUTOPILOTS & FLIGHT DIRECTOR SYSTEMS .................................. 2-10 ALTITUDE HOLD SYSTEM .............................................................. 2-11 AIRSPEED HOLD........................................................................... 2-12 ALTITUDE ALERTING SYSTEM ....................................................... 2-12 CONTROLS AND SELECTORS ......................................................... 2-13 AUTOMATIC FLIGHT DIRECTOR SYSTEM (AFDS) ............................ 2-15 MODE CONTROL PANEL ............................................................... 2-21 2.15.1 Power supplies A and B ................................................ 2-21 2.15.2 Microprocessor A and B ................................................ 2-21 2.15.3 Push-Button and Toggle Switches ................................ 2-21 2.15.4 Fluorescent Tube Control .............................................. 2-21 2.15.5 Liquid Crystal Displays & Control Knob Encoders ......... 2-22 2.15.6 AFDS Disconnect Switches ........................................... 2-23 AUTOPILOT FLIGHT DIRECTOR COMPUTER (AFDC) ....................... 2-25 2.16.1 Input Signal Selection ................................................... 2-25 2.16.2 AFDC Processors ......................................................... 2-25 PRIMARY FLIGHT COMPUTER (PFC) .............................................. 2-26 COMMUNICATIONS ........................................................................ 2-27 RADIO WAVES ............................................................................. 2-27 WAVELENGTH & FREQUENCY ....................................................... 2-28 2.20.1 Frequency Bands .......................................................... 2-29 CARRIER WAVE............................................................................ 2-29 AMPLITUDE MODULATION (AM) .................................................... 2-30 FREQUENCY MODULATION (FM).................................................... 2-31 RADIO WAVE PROPAGATION ......................................................... 2-33 2.24.1 Ground Wave Propagation ............................................ 2-34 2.24.2 Sky Wave Propagation .................................................. 2-34 2.24.3 Space Wave Propagation.............................................. 2-34 ANTENNAS ................................................................................... 2-34 MICROPHONES (MIC) ................................................................... 2-37


Part 2 - Page 1


2.27 2.28 2.29 2.30 2.31 2.32 2.33 2.34 2.35 2.36 2.37 2.38 2.39

2.40

2.41 2.42

2.43

2.44

2.45 2.46

2.47

Page 2

2.26.1 Carbon Microphone ...................................................... 2-37 2.26.2 The Crystal Microphone ................................................ 2-38 2.26.3 Moving Coil Microphone ............................................... 2-39 2.26.4 Electrostatic Microphone .............................................. 2-40 EARPHONES ................................................................................ 2-41 VHF RADIO COMMUNICATION ....................................................... 2-44 2.28.1 VHF Control Panel ........................................................ 2-45 AUDIO CONTROL PANEL .............................................................. 2-46 VHF TRANSCEIVER ...................................................................... 2-47 2.30.1 Control .......................................................................... 2-47 VHF COMMUNICATION ANTENNA .................................................. 2-48 SERVICE INTERPHONE .................................................................. 2-49 2.32.1 Attendant Interphone Handsets .................................... 2-50 FLIGHT & GROUND CREW CALL SYSTEM ...................................... 2-51 PASSENGER ADDRESS SYSTEM (PA) ............................................ 2-52 AUDIO INTEGRATION SYSTEM ....................................................... 2-53 CONTROL WHEEL MIC SWITCH ..................................................... 2-54 OPERATION OF VHF COMMUNICATION SYSTEM ............................ 2-56 HIGH FREQUENCY (HF) RADIO COMMUNICATION ........................... 2-57 2.38.1 HF Communication Control Panel................................. 2-59 SELECTIVE CALLING SYSTEM (SELCAL) ...................................... 2-61 2.39.1 SELCAL Control Panel ................................................. 2-62 2.39.2 SELCAL Decoder ......................................................... 2-63 COCKPIT VOICE RECORDER (CVR) ............................................... 2-64 2.40.1 Voice Recorder Control Panel ...................................... 2-66 2.40.2 Voice Recorder Unit...................................................... 2-67 2.40.3 Underwater Locator Device .......................................... 2-67 NAVIGATION SYSTEMS ................................................................. 2-68 VERY HIGH FREQUENCY OMNI RANGE (VOR) ............................... 2-68 2.42.1 VOR Operation ............................................................. 2-69 2.42.2 Deviation Calculations .................................................. 2-72 2.42.3 VOR Aerial Locations ................................................... 2-73 DISTANCE MEASURING EQUIPMENT (DME) ................................... 2-76 2.43.1 DME Operation ............................................................. 2-79 2.43.2 DME Controller ............................................................. 2-81 INSTRUMENT LANDING SYSTEM (ILS)............................................ 2-82 2.44.1 ILS Operation ............................................................... 2-83 2.44.2 Antennas ...................................................................... 2-87 2.44.3 LOC/GS Operation ....................................................... 2-88 MARKER BEACON SYSTEM (MBS) ................................................ 2-90 AUTOMATIC DIRECTION FINDER (ADF).......................................... 2-92 2.46.1 Loop Aerial ................................................................... 2-93 2.46.2 Station Line .................................................................. 2-94 2.46.3 Sensing the Correct Null ............................................... 2-95 AIR TRAFFIC CONTROL RADIO BEACON SYSTEM (ATCRBS) ......... 2-101 2.47.1 Transponders ............................................................... 2-101 2.47.2 ATCRBS Control Panel ................................................ 2-103 2.47.3 Mode A ......................................................................... 2-103



2.48

2.49

2.50

2.51

2.52

2.47.4 Mode C ......................................................................... 2-104 MODE S TRANSPONDERS ............................................................. 2-107 2.48.1 Mode S Interrogation & Replies..................................... 2-107 2.48.2 Discrete Addressing ...................................................... 2-107 2.48.3 Operation ...................................................................... 2-108 TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM ................... 2-110 2.49.1 TCAS Introduction ......................................................... 2-110 2.49.2 The TCAS II System ..................................................... 2-112 2.49.3 Aural Annunciation ........................................................ 2-115 2.49.4 Performance Monitoring ................................................ 2-119 2.49.5 TCAS Units ................................................................... 2-119 2.49.6 Self Test........................................................................ 2-121 2.49.7 Data Loader Interface ................................................... 2-122 INERTIAL NAVIGATION SYSTEM (INS) ............................................. 2-123 2.50.1 General Principle........................................................... 2-124 2.50.2 INS Operation ............................................................... 2-126 2.50.3 Earth Rate Compensation ............................................. 2-130 2.50.4 Vehicle Rate Compensation .......................................... 2-131 2.50.5 Alignment ...................................................................... 2-135 2.50.6 The Navigation Mode .................................................... 2-135 2.50.7 Strapdown Inertial Navigation ....................................... 2-136 2.50.8 Laser Ring Gyro (LRG) Operation ................................. 2-138 2.50.9 Mode Select Unit (MSU)................................................ 2-140 2.50.10 Mode Select Unit Modes ............................................... 2-141 2.50.11 MSU Annunciators ........................................................ 2-142 2.50.12 Inertial System Display Unit (ISDU) ............................... 2-143 2.50.13 Keyboard ...................................................................... 2-144 2.50.14 Display .......................................................................... 2-144 2.50.15 System Display Switch (SYS DSPL) ............................. 2-144 2.50.16 Display Selector Switch (DSPL SEL)............................. 2-144 2.50.17 Dimmer Knob ................................................................ 2-145 2.50.18 Inertial Reference Unit (IRU) ......................................... 2-145 2.50.19 IRS Alignment Mode ..................................................... 2-147 2.50.20 Gyro Compass Process ................................................ 2-147 2.50.21 Initial Latitude ................................................................ 2-147 2.50.22 Alignment Mode ............................................................ 2-147 RADIO MAGNETIC INDICATOR (RMI) .............................................. 2-150 2.51.1 Dual Distance Radio Magnetic Indicator (DDRMI) ......... 2-151 2.51.2 DDRMI Principle............................................................ 2-151 GLOBAL POSITIONING SYSTEM (GPS) ........................................... 2-154 2.52.1 Space Segment ............................................................ 2-154 2.52.2 Control Segment ........................................................... 2-155 2.52.3 Operation ...................................................................... 2-156 2.52.4 Signal Structure ............................................................ 2-158 2.52.5 Time Measurements ..................................................... 2-158 2.52.6 Position Fixing ............................................................... 2-160 2.52.7 Ionospheric Propagation Error....................................... 2-161 2.52.8 Derived Information ....................................................... 2-162 2.52.9 Navigation Management ............................................... 2-162 2.52.10 Boeing 777 GPS ........................................................... 2-164 2.52.11 GPS Modes of Operation .............................................. 2-165 2.52.12 Acquisition Mode ........................................................... 2-165


Part 2 - Page 3


2.53

2.54

2.55

2.56

2.57

2.58

Page 4

2.52.13 Navigation Mode ........................................................... 2-166 2.52.14 Altitude Aided Mode...................................................... 2-166 2.52.15 Aided Mode .................................................................. 2-167 2.52.16 Receiver Autonomous Integrity (RAIM) ......................... 2-169 2.52.17 Differential GPS ............................................................ 2-170 COMPASS SYSTEMS ..................................................................... 2-171 2.53.1 Direct Reading Compass .............................................. 2-171 2.53.2 Remote Reading Compass (Magnet Gyro) ................... 2-173 2.53.3 Flux valve (Detector Unit) ............................................. 2-174 2.53.4 Control Panel ................................................................ 2-175 2.53.5 Synchronisation Annunciator ........................................ 2-175 2.53.6 Synchronisation Knob ................................................... 2-175 2.53.7 Slaved/DG Switch ......................................................... 2-175 2.53.8 System Test.................................................................. 2-177 2.53.9 Gyro Unit ...................................................................... 2-177 2.53.10 Servo System ............................................................... 2-178 2.53.11 Slaving loop .................................................................. 2-180 RADIO ALTIMETER ....................................................................... 2-181 2.54.1 Basic Principles ............................................................ 2-181 2.54.2 Radio Altimeter Antenna ............................................... 2-184 2.54.3 Testing.......................................................................... 2-186 WEATHER RADAR ........................................................................ 2-187 2.55.1 Principle Of Operation .................................................. 2-190 2.55.2 Scanner Stabilization .................................................... 2-192 2.55.3 Weather Radar Installation ........................................... 2-194 2.55.4 Test Mode .................................................................... 2-197 2.55.5 Radome ........................................................................ 2-199 GROUND PROXIMITY WARNING SYSTEM ......................................... 2-200 2.56.1 System Operation ......................................................... 2-204 2.56.2 Ground Proximity Warning Computer ........................... 2-205 2.56.3 GPWS Control Panel .................................................... 2-206 2.56.4 Warning Lights.............................................................. 2-206 2.56.5 GPWS Bite Operation ................................................... 2-208 2.56.6 BITE Tests .................................................................... 2-208 2.56.7 Fault Recording ............................................................ 2-208 ENHANCED GROUND PROXIMITY WARNING SYSTEM ...................... 2-210 2.57.1 Controlled Flight Into Terrain (CFIT) ............................. 2-211 2.57.2 Terrain Alerting & Display (TAD) ................................... 2-214 2.57.3 Envelope Modulation .................................................... 2-216 2.57.4 Terrain Look Ahead Alerting ......................................... 2-217 2.57.5 Terrain Clearance Floor (TCF) ...................................... 2-218 2.57.6 TCF/TAD Control .......................................................... 2-220 2.57.7 EGPWS Interface ......................................................... 2-221 2.57.8 System Activation ......................................................... 2-223 2.57.9 Self Test ....................................................................... 2-223 AIR DATA SYSTEM (ADS) ............................................................ 2-226 2.58.1 Total Air Temperature Probe ........................................ 2-227 2.58.2 Location Of Probes And Static Vents ............................ 2-228 2.58.3 Air Data Computer (ADC) ............................................. 2-231 2.58.4 Altitude Module ............................................................. 2-233 2.58.5 True Airspeed/Indicated Airspeed Vs Altitude ............... 2-234 2.58.6 Air Data Computer (ADC) ............................................. 2-235



2.58.7 2.58.8

Digital Air Data Computer (DADC) ................................ 2-236 Definitions and Abbreviations ........................................ 2-237


Part 2 - Page 5



Page 6



2

AVIONICS SYSTEMS

2.1 AUTOMATIC FLIGHT The Automatic Flight Control Systems or AFCS, in modern jet transports, are all uniquely tailored to the specific aircraft, but all share common features. For example, the flight aerodynamics of a DC-9 are different from those of a Boeing 747 but both aircraft would most likely require an "attitude hold" mode of operation. In this case, the attitude hold feature is common to both autopilot designs, but gains in the two autopilots will differ to accommodate the differences in the aerodynamics of each aircraft. Each AFCS receives attitude and heading signals from a vertical and directional gyro and has its own rate gyro/accelerometer system to develop attitude and flight path stabilization signals. The AFCS computers comprise an electronic "brain" that receives signals from its "senses" to compute the proper responses and provides outputs to electric and/or hydraulic actuators, which move the aircraft's control surfaces. 2.2 AUTOPILOT SYSTEM Today's modern autopilots are designed to provide pitch, roll, and yaw axis stabilization around the pilot's desired reference attitude. To do this, the autopilot system must detect changes in aircraft attitude and respond to those changes more quickly and smoothly than its human counterpart. For an autopilot to maintain this stability, it must: 1. Know what the pilot's desired aircraft attitude is. 2. Know what the actual aircraft attitude is. 3. Compares the two and produce a control signal if there is a difference or error: 4. Use the control signal to correct for the difference or error and Control the speed of the correction. The human pilot controls the aircraft by detecting a change in aircraft attitude by one of his senses. His brain then computes the necessary corrective action required and transmits a signal to his muscles to move the flight controls. Again his senses will detect that corrective action has taken place and he will move the flight controls back to where they started. A typical autopilot would have to do all that the human pilot does, but would do it through electronic or electrohydraulic devices.


Page 2-1


The autopilot is divided into four main parts: 2.2.1

Error Sensing

Determines when the flight condition of the aircraft is differing from that commanded by the pilot. Almost all-modern aircraft use a gyro of some type for this purpose, and there are two ways that the error signal can be generated, either by attitude gyros or rate gyros. The attitude gyros only detect how far the aircraft is away from the settings; the rate gyros detect the rate at which it is deviating and, hence, are more accurate. 2.2.2

Correction

This is the correcting input, sent to the actuators connected into the flying control systems. This input is simply the command from the autopilot to reverse the movement of the aircraft away from its set course. It does not have any idea of when to stop the correction; this is the job of the follow-up mechanism. 2.2.3

Follow-Up

Is the detection mechanism, which senses that the aircraft is righting itself, under the commands from the correction part of the autopilot. The mechanism reduces the correction input as it nears the original selected position and, by the time the aircraft is level, there will be no correcting input to the actuators. 2.2.4

Command

The command system is incorporated to allow the pilot to dictate which heading, height, speed or rate of climb he wants the aircraft to follow. This can be a simple 'Heading Hold' system which is controlled by a "bug" on the compass, which the pilot sets with a knob on the instrument. Alternatively, the system 'Mode Control Panel' can have many different parameters commanded by the pilot, such as autopilot modes, altitude, and vertical speed and airspeed/mach number modes.

Page 2-2



Figure 1 shows a block schematic of a typical autopilot.

PITCH SERVO VERTICAL GYROSCOPE

AUTOPILOT COMPUTER

AIRCRAFT TRIM SYSTEM

COMPASS GYROSCOPE

AUTOPILOT CONTROLLER

ROLL SERVO

AIR DATA COMPUTER YAW SERVO

Basic Autopilot Figure 1 The sensors take the place of our pilot's "senses" to detect various changes in aircraft attitude. This information is fed to the computer, which calculates the size of its output signal and which axis to send it on. The controller turns the autopilot on and off and provides other system inputs not discussed here. Finally we come to the loads which are the muscle of our system and move the aircraft's flight control surfaces in response to the output signal of the computer. As the aircraft responds to these signals, the sensors, through aerodynamic feedback, detect the attitude change and tell the computer when the aircraft is back where it should be.


Page 2-3


2.3 AUTOPILOT INTERLOCKS Before an automatic control system can be engaged with an aircraft's flight controls, certain preliminary operating requirements must be fulfilled to ensure that the system is in a condition whereby it may safely take control of the aircraft. The principal requirements are that the connections between system power supplies, the elements comprising the system and the appropriate signal and engage circuits are electrically complete. It is the practice, therefore, to incorporate within any automatic control system, a series of switches and/or relays, known as interlocks, which operate in a specific sequence to ensure satisfactory engagement, and the coupling of input signals from outer loop control elements. Figure 2 shows the interlock circuit.

A/P DISCONNECT

CAPT

A.C - D.C.

MACH TRIM

PITCH TRIM

ATT REF

F/O

K1

ENGAGE RELAY

YAW DAMP

MAN AUTO

OFF

AUTO PILOT

ENGAGE RELAY K2

OFF

SERVO RUDDER CLUTCH SERVOS ELEVATOR AILERON CLUTCHES

OFF 28V DC

Interlock Circuit Figure 2

Page 2-4



The number of interlocks incorporated in any one system varies considerably according to the control capability of that system. The signals from the pitch roll and yaw gyros or computers are at their respective servos, but cannot impose their influence until the clutches are engaged. In the yaw damper only position, K1 relay will close and energize the rudder servo clutch and engage it to accept signals from a gyro or computer to move the rudder. With No 1 switch in the autopilot position, it will energize K1 relay subject to all the interlock switches being made, switch No 2 will now engage the aileron and elevator clutches, and switch No 3 will pick a voltage from switch No 2 and energize the rudder clutch. So in the yaw damper position, it is yaw damper only, and in the autopilot position it is yaw, pitch and roll engagement. In the yaw damper switch position, only the pilot's disconnect and power valid's are needed; in the full autopilot condition all switches must be made. Here is a review of the interlock switches. Firstly, the autopilot disconnects; either the Captain's or the First Officer's switch will disconnect the autopilot. The switches are usually located on the control column. The power ac and dc valid's are qualifying that power is available and any loss of power will disconnect. Mach trim has to be engaged in this case. In some systems, Mach trim is on all the time, whether the autopilot is engaged or not, and in others it is disengaged when the autopilot is engaged. The pitch trim switch is qualifying that auto pitch trim is available in the autopilot mode. The reason that this is important is that if there is a mis-trim, the autopilot can compensate for that situation until the autopilot is disconnected, either through malfunction or deliberate action, or the aircraft could nose up/down rather dramatically. The attitude reference switch is checking that the valid's from the vertical gyro are all correct, and the attitude references are available for the autopilot. On more sophisticated systems, there are other interlock switches, for example air data computer, compass system, hydraulic pressure monitoring and radio altimeters. Some systems, apart from those that use electrical interlocks, do not use an electrical servomotor. Instead they use a hydraulic servo.


Page 2-5


2.4 SERVOMOTORS The power output element of any automatic flight control system consists of servomotors, or servo-actuators as they are sometimes called, connected into the aircraft's primary flight control system circuits; the number of servomotors employed is governed by the number of control loops required. In addition to the actuation of primary flight controls, servomotors may also be used, in some cases, for the actuation of the secondary flight controls provided for trimming purposes and for yaw damping. In general, servomotors operate on either electro-pneumatic, electromechanical, or electro-hydraulic principles, the choice, and constructional features adopted in applying such principles being dependent on the type of automatic control system, and on the methods adopted for actuation of the primary flight control surfaces. Servomotors may be connected either in series or in parallel with the normal flight control system of an aircraft. A series-connected servomotor is one, which moves the flight control surfaces without moving the pilot's controls, while a parallel-connected servomotor moves both the control surfaces and the pilot's controls. Servomotors may utilise either direct current or alternating current, depending on the individual systems. Motor type ranges from dc permanent magnet to ac twophase or hysteresis type. The closed loop servo technique can be applied as a means of achieving automatic flight control of an aircraft. A functional diagram is at Figure 3. This forms the basic control system for all classes of automatic flight control systems. This system controls what is termed “Inner Loop” stabilization.

Page 2-6



AUTOPILOT MODE SELECT

FEEDBACK

SIGNAL PROCESSING ATTITUDE SENSING

ERROR SENSING

PILOT’S DEMANDS MANUAL FLIGHT CONTROLS

AERODYNAMICS

SERVOMOTORS (ACTUATORS)


Inner Loop Stabilization Figure 3 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-7


The number of control loops, or channels, comprising an automatic flight control system is dependent on the number of axes about which control is to be effected and in this connection it is usual to classify systems as: 1. Single Axis System. 2. Two-Axis System. 3. Three-Axis System. 2.5 SINGLE AXIS CONTROL SYSTEM In the single axis system, control is normally about the “Roll” axis. The control surfaces forming part of this system are therefore the “Ailerons”. It is found on small aircraft to provide lateral stabilization (wing levelling). 2.6 TWO-AXIS SYSTEM In the two-axis system, control is normally about the “Roll” and “Pitch” axes. The control surfaces forming part of this system are therefore the “Ailerons” and “Elevators”. These are found on medium sized aircraft and provide a means of automatically controlling the aircraft’s heading and altitude. 2.7 THREE-AXIS SYSTEM In the three-axis system, control is about all three axes (Pitch, Roll and Yaw). These systems are designed to meet the requirements for stabilization and control of high performance category aircraft, and have a large number of modes of operation. 2.8 SENSING ATTITUDE CHANGES Under automatically controlled flight conditions, the sensing of all changes in the aircraft’s attitude is accomplished by referencing them against some form of stabilized device. The device universally adopted for this purpose, from the earliest types of control system to those now current, has been the gyroscope. In addition to the gyro, it is also the practice in many cases to adopt a pendulous device which although not purely stabilizing in function, can serve as a “back-up” to a gyro by sensing short-term attitude changes brought about by the effects of accelerations, vertical speed changes, and by side-slip.

Page 2-8




PITCH AXIS

ROLL RATE GYRO


YAW AXIS

YAW RATE GYRO

PITCH RATE GYRO


ROLL AXIS

Figure 4 shows the gyro configuration for a three-axis automatic control system.

Three-axis Automatic Control System. Figure 4 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-9


2.9 AUTOPILOTS & FLIGHT DIRECTOR SYSTEMS Once the controller has been selected, and activated, the aircraft is controlled by the Flight Director/Autopilot System. Rate gyros detect any movement of the aircraft from the selected flight datum and will output a signal proportional to the disturbance and in the opposite sense. The gyro output, along with other signals from associated systems, are processed in the Flight Director/Autopilot Computer, which in turn will give flight director information and or outputs to move the control surfaces to bring the aircraft onto the correct flight datum. Figure 5 shows a schematic of a Flight Director/Autopilot System.

AERODYNAMIC RESPONSE

FLIGHT DIRECTOR COMMAND BAR

FLIGHT DIRECTOR ENGAGED

GYRO INPUT MODE SELECT NAV AIDS INPUT HEADING INPUT

PILOT’S INPUT AUTOPILOT COMPUTER

FEEDBACK

ALTITUDE INPUT

SERVO

AUTOPILOT ENGAGED

Flight Director/Autopilot System Figure 5

Page 2-10



2.10 ALTITUDE HOLD SYSTEM We know that any change in the aircraft's attitude will be detected by the Autopilot system. This system alone will not be able to detect a pure vertical displacement of the aircraft. To maintain an aircraft at a selected altitude we require further sensing elements. The purpose of the Altitude Hold system is to maintain the aircraft at a selected height. The pilot will select "ALT" on the Flight Mode Panel (FMP) and the system will maintain that altitude. The sensing element consists of a pressure transducer, similar to that in the Air Data System. Any change in the static pressure will be felt and an output produced, this output will be fed to the pitch channel of the autopilot system to adjust the aircraft's altitude. A simplified Altitude Hold system is shown at Figure 6.

ANEROID CAPSULE

STATIC

CONTROL MOTOR ERROR AMP

CHASER MOTOR ALT HOLD SELECT

REF TO PITCH CONTROL CHANNEL

Altitude Hold System Figure 6


Page 2-11


2.11 AIRSPEED HOLD Since airspeed hold sensors are used in conjunction with altitude hold sensors, the methods of transmitting error signals are of a common nature. The only difference is that whereas an altitude sensor measures only static pressure changes, an airspeed sensor is required to measure Static and Pitot pressures. 2.12 ALTITUDE ALERTING SYSTEM The Altitude Alerting System allows the pilot to make changes to the aircraft's altitude and provide alerts to the pilot when the selected altitude is reached. The pilot sets the required altitude, from 0 - 50,000 feet, in steps of 10 feet, on the Flight Mode Panel (FMP). The altitude alerter gives the pilot an alert when the aircraft approaches the selected altitude, entry alert ("C" Chord) and illuminates a warning lamp. The system will then alert the pilot when the aircraft does not follow the selected altitude with an exit alert ("C" Chimes) and illuminates a warning lamp. Figure 7 shows the different alerts.

1000 feet

EXIT ALERT ON

ENTRY ALERT ON “C” CHORD

ENTRY ALERT OFF

250 feet

SELECTED HEIGHT

250 feet

EXIT ALERT ON “C” CHIME

1000 feet

Alert Levels Figure 7 Page 2-12



2.13 CONTROLS AND SELECTORS Figure 8 shows the controls from a BAe 146 aircraft.

S P L IT A /P

M O D E

GSL

A LT

A /P

VS

M A CH

N AV 1

V - NA V

N AV 2

IA S

0 6 8

2 4 6

C O U RS E S E L

C O U RS E

T UR B V /L

B -L O C

L -N AV

H DG

H DG

M OD E SELE CTO R

N A V IG A T IO N S E L E C T O R

R UD

EL EV

T EST

P IT CH

R OL L L

A L T

A LT AR M

R

D OW N

A LT S E L

2 5 9.0 0 F EET YD

S E L

YD 1 YD 2

A L T IT U D E S E L E C T O R

A /P

IN

UP

A U T O P IL O T S E L E C T O R

BAe 146 Autopilot controllers Figure 8 a)

MODE SELECTOR : Is mounted on the glare-shield and contains the push button switches for the selected mode. Hidden legends are used so that the button appears blank, until a mode is selected when a white triangle is illuminated. Engagement of the autopilot is indicated by a green triangle on the AP button at the top of the panel. In essence the bottom row selects lateral modes and the middle row selects vertical modes.

b)

NAVIGATION SELECTOR : Mounted on the glare-shield and contains a large rotary switch labeled NAV 1-SPLIT-NAV 2. This selects the distribution of radio navigation information to the autopilot and to the pilot’s flight instruments.


Page 2-13


The autopilot and flight directors use the information that is displayed on the captain's HSI. With SPLIT selected NAV 1 supplies HSI 1 and NAV 2 supplies HSI 2. If NAV 1 is selected then both HSIs are supplied from NAV 1 and a NAV 2 selection supplies both HSIs from NAV 2. The COURSE selector knobs allow rotation of the course pointer on the HSIs. A HDG knob provides remote selection of the heading cursor on both HSIs. Two ratios are available, coarse and fine. c)

ALTITUDE SELECTOR : Mounted on the glare-shield this contains a fivefigure readout; the last two figures are fixed zeros. A mode select button labeled ALT ARM allows arming of the selected altitude. The 'armed' state is indicated by a white triangle. A press to TEST switch allows warning altitudes to be checked against the altitude set on the captain's altimeter.

e)

AUTOPILOT CONTROLLER : Mounted on the center console and contains the autopilot (AP) and the yaw damper YD engage buttons. These also indicate engagement by green illuminated IN for the AP, and YD1/YD2 for the yaw damper. PITCH and ROLL controls and associated out of trim indicators (ELEV and RUD) are also found on the controller.

Page 2-14



2.14 AUTOMATIC FLIGHT DIRECTOR SYSTEM (AFDS) The Boeing 777 AFDS is used as an example in this module. The purpose of the AFDS is to automatically control the aircraft’s attitude and to supply indications to the flight crew in order for them to manually control the aircraft’s attitude. The autopilot controls the aircraft’s attitude through: Takeoff (Flight Director only), Climb, Cruise, Descent, Approach, Go-around and Autoland. In the Flight Director mode, the director bars (horizontal/vertical) show on the Primary Flight Displays (PFD). The bars are used as guides to control the attitude of the aircraft. Figure 9 shows the Primary Flight Display (PFD).

HOLD

LNAV

VNAV

LOC

G/S

5100 5200

200

A/P 180 160

20

20

10

10

3

1

4800

14 2 120

5000

6 2

10

10

20

20

1

4600

2 6

100 4400

Primary Flight Display Figure 9


Page 2-15


HOLD VS/FPA

UP

DOWN

+3288 V/S

FPA V/S HOLD

SEL

(b)

5

AUTO

A/P DISENGAGE

A/T OFF

F/D ON

A/P

OFF

F/D ON

A/P

HOLD

25 BANK LIMIT

238

(a)

HDG

TRK HDG

FLCH

V-NAV

OFF

A/T ARM R L

CLB ON

IAS

288

FLCH

A/T

OFF

CLB ON

IAS

IAS

MACH

L-NAV

A/P DISENGAGE V-NAV

L-NAV

288

MACH IAS R L

A/T ARM

AUTO

ALTITUDE

17000

1000

(c)

OFF APP

HOLD

UP

BANK LIMIT

SEL AUTO

5

HDG

HDG

238

TRK

25

DOWN

V/S

V/S

FPA

+3288

VS/FPA

AUTO

ALTITUDE

17000

LOC

1000

APP

LOC

A/P

F/D ON

OFF

A/P

F/D ON

The mode select panel is the primary interface between the flight crew and AFDS. Other flight crew inputs to the AFDS are; the disconnect switches and the Go-around (GA) switches. Figure 10 shows the Auto Flight Director System (AFDS) mode control panel for the Boeing 777 aircraft.

Boeing 777 – AFDS Mode Control Panel Figure 10

Page 2-16



With reference to Figure 10a: A/P Engage Switch – Captain’s autopilot engage button, shows white when engaged. A/T ARM – Left and right autothrottle arm switches. F/D Switch – Allows the selection of the Flight Director bars for display on the PFD. CLB CON Switch – Climb continuous thrust switch. A/T Switch – Engages the autothrottle system. L-NAV Switch – Engages lateral navigation mode. V-NAV – Engages vertical navigation mode. FLCH – Flight Level change engage switch. IAS/MACH Window – Shows the selected IAS/MACH as selected using the IAS/MACH select knob. IAS/MACH Switch – Selects either IAS or MACH as the reference for speed hold mode.


Page 2-17


With reference to Figure 10b: A/P DISENGAGE Bar - There are three toggle switches under the disengage bar. The left switch controls the left AFDS only and the right switch controls the right AFDS. The center switch controls the center AFDS. The center AFDC cannot do a single autopilot engagement because it does not connect to any back drive unit. It is there only as a back up for the left or right. The bar is normally in the up position. Pushing the bar down will disengages all the AFDS. Figure 11 shows the operation of the disengage bar.

DISENGAGE BAR UP (ALL THREE AFDS ENGAGED)

DISENGAGE BAR DOWN (ALL THREE AFDS DIS-ENGAGED)

DISENGAGE BAR UP RIGHT AFDS SWITCH DOWN (ONLY LEFT AND CENTER AFDS ENGAGED)

Disengage Bar Operation Figure 11

Page 2-18



Light Sensor – A photo light sensor on the MCP front panel monitors ambient lighting. It controls the brightness of the LCD’s on the mode panel. HDG/TRK Switch – This switch controls the reference for the Heading/Track window. HDG/TRK Window – The window shows heading or track angle in increments of one degree. The window range is from 001 to 360. At AFDS power-up, the window shows 360. Heading/Track Selector Switch – This control has two concentric selectors and one push-button. The outer selector controls the bank angle, the inner selector controls the value of heading/track required. The inner selector (push-button) selects between Heading, or Track select modes. HOLD Push Button – Engages the AFDS into Heading/Track hold mode. Vertical Speed/Flight Path Angle (V/S/FPA) reference Switch – This switch controls the reference for the vertical speed/flight path angle window. Vertical Speed/Flight Path Angle Window - The window shows vertical speed value (range is +6000 fpm to –8000 fpm). The flight path angle is +9.9 to –9.9. Vertical Speed/Flight Path Angle selector – Rotate the selector up to decrease the value and down to increase the value. VS/FPA Push Button – Engages the VS/FPA mode. Altitude Window – Has a range from 0 to 50,000ft. The increment is variable. The set altitude is also the altitude alert value for the caution and warning system. At AFDS power-up the display is set at 1,000ft. Altitude Selector – The control has two concentric selectors. The inner selector changes the reference altitude in the window. If the selector is pushed while in VNAV, this will activate the altitude intervention. The outer selector changes the window increment. With it selected to 1000 position, the inner selector changes the window at 1000 feet/detent. With the outer selector in the AUTO position, the window change rate is 100 feet/detent. HOLD Push Button – Engages the Altitude hold mode.


Page 2-19


With reference to Figure 10c: A/P Engage Switch – First Officer’s autopilot engage button, shows white when engaged. F/D Switch – Allows the selection of the Flight Director bars for display on the PFD. LOC Push Button – Engages the ILS LOC mode. Captures and holds the aircraft to a Localizer flight path. APP Push Button – Engages the Approach mode. Captures and holds the aircraft to a Glideslope (vertical descent) flight path. Figure 12 shows a block schematic of the Mode Control Panel.

POWER SUPPLY “A”

INPUT/OUTPUT SIGNAL PROCESSING

MICROPROCESSOR “A”

A/T HDG 5 AUTO

LIGHTS

F/D ON

238 SEL

25 BANK LIMIT

LCD DIPLAY & KNOBS

OFF V-NAV

PUSH SW

AFDC “L”

ARINC 429 TX

AFDC “L”/”C” & AIMS

ARINC 429 RX

AFDC “C”

ARINC 429 TX

AFDC “R” & AIMS

ARINC 429 RX

AFDC “R”

A/T ARM L

R

OFF

ON-OFF SWITCHES

MICROPROCESSOR “B”

POWER SUPPLY “A”

ARINC 429 RX

INPUT/OUTPUT SIGNAL PROCESSING

Mode Control Panel Schematic Figure 12

Page 2-20



2.15 MODE CONTROL PANEL 2.15.1 Power supplies A and B They receive 28V dc from the left and right 28V dc buses. The MCP functions with either source. Power supplies A and B supply +12V, -12V and +5V dc to their respective microprocessors and logic circuits. They also supply power supply C, which is part of the fluorescent tube control. 2.15.2 Microprocessor A and B There are two separate microprocessors within the MCP. Microprocessor A receives data from the left and center AFDC and microprocessor B receives data from the right and center AFDC. Microprocessor sends data to the left and center AFDC, microprocessor B sends data to the right AFDS. To make sure all three AFDC use data from one MCP processor, all AFDS use the microprocessor data sent to the master AFDS. When the left AFDS is master, microprocessor A writes to the LCD displays. The right and center AFDC receive microprocessor A data through the AFDC cross-channel buses. When the right AFDS is master, microprocessor B writes to the LCD displays. The left and center AFDC receive microprocessor B data through the AFDC cross-channel buses. 2.15.3 Push-Button and Toggle Switches Each push-button and toggle switch has two sets of contacts. One set connects to microprocessor A and one set connects to microprocessor B. The LED annunciators in the push-button switches also connect to each microprocessor. 2.15.4 Fluorescent Tube Control The fluorescent tube control circuit supplies current to drive the tube and the tube heater. There is a heater coil around the tube, which will operate when the temperature is <40F.


Page 2-21


2.15.5 Liquid Crystal Displays & Control Knob Encoders There are four LCD windows, which show reference values. The values change when the selector is rotated or when the AFDC command a new reference. Microprocessor A and B drive each LCD. Each selector connects to the two encoders and each encoder sends data to the on-side microprocessor. Figure 13 shows the LCD illumination.

FLUORESCENT TUBE

LCD (TYPE) DISPLAY

LCD Illumination Figure 13

Page 2-22



2.15.6 AFDS Disconnect Switches The autopilot disconnect switches are on the outboard side of each control wheel. The switch is a push-button type with multiple contacts. These switches manually disconnect all AFDCs. Figure 14 shows the Captain’s Control wheel A/P disconnect switch.

ELEVATOR TRIM SWITCHES

AUTOPILOT DISCONNECT BUTTON

A/P Disconnect Switch Figure 14


Page 2-23

Page 2-24

DISC

TO/GA

SECONDARY ATTITUDE & AIR DATA REFERENCE UNIT

AIR DATA INERTIAL REFERENCE UNIT

A/C SENSORS

AIMS ON

CLB

A/T

OFF

R

IAS

IAS

288

MACH

FLCH

V-NAV

L-NAV

ACTUATOR CONTROL ELECTRONICS

PRIMARY FLIGHT COMPUTER

AUTOMATIC FLIGHT DIRECTOR COMPUTER (AFDC)

Collins

MODE CONTROL

OFF

F/D ON

A/P

L

A/T ARM

AUTO

5

HOLD

SEL

238 25

TRK

LIMIT

BANK

V/S

PCU

FPA

+3288

V/S

POSITION TRANSDUCERS

UP

DOWN

BACKDRIVE ACTUATORS

NAVIGATION SENSORS

A/P DISENGAGE

HDG

HDG

VS/FPA

APP

LOC

OFF

F/D ON

A/P

FLIGHT DECK CONTROLS

HOLD

1000

ELEVATOR AILERON & RUDDER

AUTO

ALTITUDE

17000



Figure 15 shows the AFDS block schematic diagram

AFDS Block Diagram Figure 15



2.16 AUTOPILOT FLIGHT DIRECTOR COMPUTER (AFDC) There are three AFDC within the AFDS each containing the following: 1. ARINC 629 Input/output modules (I/O). 2. Discrete Input/output (I/O) module. 3. Processor A. 4. Processor B. 5. Processor C. 6. Power supply module. 2.16.1 Input Signal Selection Each ARINC and discrete input/output module monitors and selects input signals. Each I/O module monitors the validity of the signal first, if validity check is good, the I/O section selects the signal by one of the following methods: 1. Mid value selection which uses the middle value of the three signals. Radio Altitude (RA) and ILS are examples of signals selected using this method. 2. Priority selection for signals with two sources (left/right). Example; the Air Data Inertial reference Unit (ADIRU) is the normal source of air/inertial data. If the ADIRU fails, the AFDC selects the Secondary Attitude Air Data reference Unit (SAARU). 3. Forced selection for Aeroplane Information Management System (AIMS) data. AIMS tell the AFDC which signal to use. 2.16.2 AFDC Processors The AFDC has three processors (A,B and C). Processor A and B receive digital backdrive commands from the Primary Flight Computers (PFC). They convert the digital backdrive signals into analogue signals for output to the backdrive actuators. Processor C calculates the autopilot and flight director control laws, test and data loading, engage/disengage logic and failure detection/fault response monitoring.


Page 2-25


2.17 PRIMARY FLIGHT COMPUTER (PFC) The PFC receives commands from the AFDC. The PFC calculates and sends surface position digital commands to the Actuator Control Electronics (ACE). The ACE converts these signals to analogue and sends the signal to the Power Control Units (PCU). The PCU move the flight surface, sending positional feedback signal to the ACE, which converts the feedback signals into digital and sends it to the PFC. The PFC then calculates and sends the digital feedback signals to the AFDC. The AFDC converts the signals into analogue, and sends these signals to the backdrive actuators, which moves the control column, control wheels and rudder pedals. Figure 16 shows the Primary Flight Computer/Actuator Control Electronics Block diagram.

ANALOG ANALOG POWER CONTROL UNIT POSITION TRANSDUCER BACKDRIVE ACTUATORS

PRIMARY FLIGHT COMPUTER

ACTUATOR CONTROL ELECTRONICS

CONTROL SURFACE

ANALOG

FLIGHT CONTROL - ARINC 629 BUS (X3)

AFDC

AIMS

ADIRU

SAARU

PFC/ACE Block Diagram Figure 16

Page 2-26



2.18 COMMUNICATIONS The word "Communication" is defined as the exchange of information of any kind, by any means and involves the transfer of meaningful information from one location (the sender or transmitter), to another location (the destination or receiver). Radio Communication equipment in aircraft is primarily for the purpose of communicating with Air Traffic Control (ATC), and other ground stations. It can also be used for communicating with other aircraft and internally to speak with cabin crew and passengers. 2.19 RADIO WAVES Radio signals emanate from the antenna of a transmitter partly in the form of “Electromagnetic waves”. During radio transmission, the antenna in addition to the electromagnetic field also generates an electric field. The two fields radiate from the antenna at the speed of light, which is approximately 186,300 Mls/sec (300,000,000 mtr/sec). Since radio waves travel at the speed of light, as soon as the transmitter starts to transmit, its signal may be detected instantly hundreds or thousands of miles away, depending on the power of the transmitter and the nature of the wave being transmitted. The transmitter typically radiates an electromagnetic signal in a 360º pattern from the antenna. Figure 17 shows the effect of a radio wave being transmitted from an Omni-directional aerial.

AERIAL

Radio Wave Transmission Figure 17


Page 2-27


2.20 WAVELENGTH & FREQUENCY The length of a radio wave depends on its frequency. Like an ac sine wave, the wave emanating from an antenna increases to a maximum in one direction, drops to zero, and then increases to maximum in the opposite direction. The wavelength, indicated by the Greek letter lambda (), is the distance from the crest of one wave to the next. Since the wave travels at the speed of light (300,000,000 mtr/sec) the wavelength in metres is equal to 300,000,000 divided by the number of cycles per second (hertz). Figure 18 shows the relationship between wavelength and frequency.



WAVELENGTH = CREST TO CREST (MTR)

RADIO WAVE CYCLE FREQUECNY = NUMBER OF CYCLES PER SECOND

WAVELENGTH =

VELOCITY FREQUENCY

VELOCITY = SPEED OF LIGHT ( 300,000,000 METRES PER SECOND)

FREQUENCY =

VELOCITY WAVELENGTH

Wavelength and Frequency Figure 18

Page 2-28



2.20.1 Frequency Bands Frequencies utilized in various types of radio systems range from 3 KHz to as high as 30 GHz. The frequencies are divided into seven bands, and these bands are assigned to certain types of operation. Table 1 shows the various types of bands. Designation Very Low Frequency (VLF) Low frequency (LF) Medium Frequency (MF) High Frequency (HF) Very High Frequency (VHF) Ultra High frequency (UHF) Super High frequency (SHF)

Frequency range 3 – 30 KHz 30 – 300 KHz 300 – 3000 KHz 3 – 30 MHz 30 – 300 MHz 300 – 3000 MHz 3 – 30 GHz

Wavelength 1000,000 – 10,000 Mtr 10,000 – 1,000 Mtr 1,000 – 100 Mtr 100 – 10 Mtr 10 – 1 Mtr 1 Mtr – 10 cm 10 cm – 1 cm

Frequency Bands Table 1 Above these radio frequencies lie the various light frequencies. Infrared and white light are currently being used for some information transmission at frequencies between 109 – 1011 KHz. Below the radio frequencies are the audible sound waves, ranging from 20 Hz to 15 KHz. The audio frequency range for radio transmission is between 300 Hz – 3 KHz and is known as “Commercial Quality Speech”. Without special techniques, the transmission of these low frequencies would cause two major problems: 1. High power would be required to transmit them. 2. All radio transmissions would interfere with each other. 2.21 CARRIER WAVE The energy that carries the intelligence of a radio signal is called the “Carrier Wave”. The frequency of this carrier wave may be only a few hundred kilohertz (VLF) or several thousand megahertz (UHF). Carrier waves are usually in the “Radio Frequency” (RF) range, which is in excess of 20 KHz. Frequencies below 20 KHz are in the “Audio Frequency” (AF) range.


Page 2-29


In order to carry intelligence, a RF carrier wave must be modulated. This means its form and characteristics are changed by means of some form of signal impressed onto it. There are two methods used for modulating the carrier wave, these are: 1. Amplitude Modulation (AM). 2. Frequency Modulation (FM). 2.22 AMPLITUDE MODULATION (AM) In amplitude modulation, the audio signal is mixed in a “Modulator” with the higher carrier frequency. The audio affects the amplitude of the carrier frequency as shown in figure 19.

CARRIER FREQUENCY

AUDIO FREQUENCY

TRANSMITTED AMPLITUDE MODULATED SIGNAL Amplitude modulation Figure 19

Page 2-30



2.23 FREQUENCY MODULATION (FM) This type of modulation provides a signal that is much less affected by interference than an AM signal. Frequency modulation is accomplished by varying the frequency of the carrier in accordance with the audio signal desired. Figure 20 shows how frequency modulation affects the carrier wave.

AUDIO WAVE

FREQUENCY MODULATED SIGNAL

Frequency Modulation (FM) Figure 20


Page 2-31


Figure 21 shows the relationship of frequency modulation with different amplitudes of audio frequency.

1KHz AUDIO

HALF MAX AMPLITUDE

MAX AMPLITUDE

10.5KHz 9.5KHz

11KHz 9KHz

10KHz CARRIER

FM Modulation Figure 21

Page 2-32



2.24 RADIO WAVE PROPAGATION The carrier wave emitted by a radio transmission antenna may be broken into three different propagation categories: 1. Ground wave. 2. Space wave. 3. Sky wave. Figure 22 shows the different propagation paths.

Y

W

A

VE

IONOSPHERE

SK

S

C PA

AC

E

A EW

VE

GR

OU

ND

W

AV

E

SP

E

V WA

Radio Wave Propagation Figure 22


Page 2-33


2.24.1 Ground Wave Propagation Ground waves tend to be held near the earth’s surface and bend with the curvature of the earth. Ground waves travel a distance limited by the transmitter’s output power, antenna design, local terrain, and current weather conditions. Typically, a relatively powerful transmitter is capable of sending ground waves a distance of 1,000 miles. 2.24.2 Sky Wave Propagation Sky waves tend to travel in straight lines, but may also be reflected off the ionosphere layer in order to reach the receiver. Because of this method, sky waves may produce a skip zone, where no reception is possible. Neither the lineof-sight nor the reflected wave can be received in the skip zone. The ionosphere’s density and distance from the earth determine the skip-zone range and the exact frequencies that are reflected. The ionosphere is a layer of ionized gases that surround the earth at an altitude of between 20 – 250 miles, varying with the time of day, season and location. The density of this layer is also affected by the sun’s solar flare activities. All these factors will determine the frequencies that are reflected and their angle of reflection off the ionosphere. 2.24.3 Space Wave Propagation Space wave frequencies have a short wavelength, which allows them to penetrate the ionosphere. Because of this, space waves are limited to line-ofsight reception only. This method is used to communicate with satellites (SAT COMM/GPS). 2.25 ANTENNAS An antenna is a specially designed conductor that accepts energy from a transmitter and radiates it into the atmosphere. During reception, an antenna acts as a device that receives an induced current from passing electromagnetic waves. This induced current is then sent to the radio receiver circuitry. Where transmitters and receivers are built into one unit, often called a “Transceiver”, the same antenna is used for both transmitting and receiving. The size and design of antennas vary in accordance with the frequency or frequencies of signals being handled. As frequencies increase the wavelengths decrease and the length of the antenna must be matched as closely as possible to the wavelengths of the carrier waves. On aircraft the size of the antenna is normally ¼. Most aircraft communication antennas are of the “Blade” type. The radiating surface is ¼ and is protected by a polyurethane rubber coating. They are generally termed “Broad band” antennas, meaning they will receive a wide range of frequencies. The required frequency is filtered out from all the others by circuitry within the transceiver. Figure 23 shows a typical VHF blade antenna.

Page 2-34



ANTENNA CAP

ANTENNA

EROSION BOOT

AIRCRAFT SKIN

MOUNTING SCREWS

TUNING CABLE ANTENNA CABLE

VHF Antenna Figure 23


Page 2-35


RECEIVER

ANTENNA COUPLER

ANTENNA

Figure 24 shows an Antenna Coupling unit.

Antenna Coupling Unit Figure 24

Page 2-36



2.26 MICROPHONES (MIC) The purpose of the mic is to convert sound energy into electric energy. This process uses the dynamic energy of the sound wave produced by the pilot. The sound strikes a diaphragm, and the sound energy is converted into mechanical energy. This mechanical energy is then converted into electric energy. There are four common aircraft microphones: 1. The Carbon Microphone. 2. The Crystal Microphone. 3. The Moving Coil Microphone. 4. Electrostatic Microphone. 2.26.1 Carbon Microphone The carbon mic contains tiny carbon granules compressed in a sealed chamber. The voice diaphragm vibrates the carbon chamber, changing the resistance of the carbon granules. A current that passes through the granules changes in amplitude as the sound wave moves the diaphragm. Figure 25 shows the operation of a carbon microphone.

CERAMIC CUP

CARBON GRANULES

CONDUCTING SURFACE

DIAPHRAGM INSULATED PLUNGER

Carbon Microphone Figure 25


Page 2-37


2.26.2 The Crystal Microphone The crystal mic is a voltage generator which utilizes the piezoelectric properties of a quartz crystal. When the crystal is subjected to mechanical pressure it develops a potential across two of its faces. This potential is dependent on the pressure exerted on the crystal. This in turn produces an output, which corresponds exactly to the applied pressure wave. Figure 26 shows a crystal microphone operation.

DIAPHRAGM

ELECTRODES

CRYSTAL

Crystal Microphone Operation Figure 26

Page 2-38



2.26.3 Moving Coil Microphone The moving coil microphone is again a type of voltage generator, this time working on the electromagnetic-induction principle. The diaphragm is attached to a coil, which is free to move in or out of a strong magnet. Movement of the diaphragm causes the coil to cut the magnetic flux and a voltage is induced into the coil. Figure 27 shows the operation of the moving coil microphone.

MAGNET

COIL

Moving Coil Microphone Operation Figure 27


Page 2-39


2.26.4 Electrostatic Microphone An electrostatic microphone is similar to the carbon microphone in that it controls the power taken from a dc supply. The principle is that of varying the value of a capacitor by altering the distance between the capacitor plates. When the diaphragm moves under the influence of a sound pressure wave the gap between the plates varies and alters the capacitance. The current varies directly as the charge across the microphone alters. Figure 28 shows the operation of the electrostatic microphone.

DIAPHRAGM

AIRGAP

MOVEABLE PLATE

FIXED PLATE

Electrostatic Microphone Operation Figure 28

Page 2-40



2.27 EARPHONES The earphone is a transducer that converts electrical waves into sound (pressure) waves. The waveform of the sound wave should be identical to the electrical wave in all factors but amplitude. Earphones therefore, merely perform the reverse process of a microphone. The same principles apply to earphones as they did for microphones. Figures 29 - 31 show the different types of earphones and microphones found on modern aircraft.

HEADBAND

PRESS-TO-TALK SWITCH

MOUTHPIECE

EARPIECE ONLY USED TO MONITOR AUDIO NO TALK FACILITY

Headset and Hand Held Microphone Figure 29


Page 2-41


HEADBAND

EAR PIECE TRANSDUCERS AUDIO TUBES AMPLIFIER BOOM MIC

JACK PLUG

Headset and Hand Held Microphone Figure 30 HEAD RESTRAINER MASK

COMMUNICATION JACK PLUG

OXYGEN CONNECTION

Emergency Oxygen Mask Figure 31

Page 2-42



DETECTOR & DEMODULATOR RADIO FREQUENCY OSCILLATOR

MICROPHONE

RADIO FRQUENCY AMPLIFIER

AUDIO FREQUEUNCY AMPLIFIER

LOUDSPEAKER

AUDIO FREQUEUNCY AMPLIFIER

MODULATOR

MODULATED RADIO FREQUENCY AMPLIFIER

RADIO FREQUENCY AMPLIFIER

Figure 32 shows a simplified block schematic diagram of a basic radio system.

Simplified Radio System Figure 32


Page 2-43


2.28 VHF RADIO COMMUNICATION Aircraft communication systems normally use the VHF wave band within a 118.000 MHz to 136.000 MHz range. Within this range the channel spacing was previously 25 KHz, but because of the high demand for more channels it is currently being reduced to 8.33 KHz. The VHF system provides short-range (space wave) voice communication between: 1.

The aircraft and ground stations.

2.

Aircraft to aircraft.

All modern aircraft have at least two VHF systems, on the larger aircraft, there is also a third system fitted. Each VHF communication system receives RF energy via its antenna, processes the RF signal and sends the resulting AF to the digital audio control system, and the SELCAL (see later). During transmission, microphone audio from the flight compartment is processed by the VHF communication system and the RF energy is transmitted via the antenna. Control of the frequency selection is provided on a VHF Communication control panel. Figure 33 shows a VHF Radio system block schematic.

AUDIO OUT TO INTERPHONE VHF AERIAL RF IN & OUT MICROPHONE INTERPHONE

VHF COMMUNICATION TRANSCEIVER 1ST OFF

VHF CONTROL PANEL

CAPT

VHF Radio System Figure 33

Page 2-44



2.28.1 VHF Control Panel The purpose of the VHF communication control panel is to provide frequency selection (tuning), frequency transfer, and testing of the associated VHF transceiver. There are two sets of concentric frequency select knobs. On each set, the out knobs select the 2nd and 3rd digits and the inner knobs select the 4th and 5th digits. Above each set of knobs is a frequency select readout for displaying the selected frequency. A two position VHF COMM TRF (transfer) switch allows the selection of one of the pre-selected frequencies. The unselected frequency window has a bar obscuring the readout. The COMM TEST switch is a push button switch that enables confidence testing of the receiving circuits in the system. Figure 34 shows a VHF control panel.

VHF COMM

120.60

118.30 TFR COMM

TEST

VHF Control Panel Figure 34


Page 2-45


2.29 AUDIO CONTROL PANEL The audio control panels provide microphone selector pushbuttons and listen switches for the VHF communication systems. The mic selector pushbuttons connect the microphone to the desired VHF transceiver. The audio volume controls allow the selection of audio from the transceivers to heard over the flight compartment speakers or headphones. Figure 35 shows an audio control panel.

MIC SELECTOR

1 - VHF - 2

HF - 1

1 - NAV - 2

SERV INT

INOP

1 - ADF

MASK R/T

FLT INT

PA

INOP

MKR

V

B

R

I/C BOOM

Audio Control Panel Figure 35

Page 2-46



2.30 VHF TRANSCEIVER The purpose of the transceiver is to transmit and receive RF signals for voice and data communication. It is a solid-state device with a minimum transmit power output of 20 watts. 2.30.1 Control The transceiver is tested using front panel controls. The “squelch disable” pushbutton allows the testing of the receiver section of the transceiver. An amber “transmit monitor” lamp illuminates whenever the transmitted output power Exceeds 10 watts. There are also “phone” and “mic” jacks available for the monitoring of the receiver and transmitter. Figure 36 shows a VHF transceiver.

ON WHEN TRANSCEIVER POWER > 10W

TESTS THE RECEIVER SECTION OF THE TRANSCEIVER SQUELCH DISABLE

TRANSMIT POWER

PHONE

MIC

MONITORING OF AUDIO OUTPUT

OPERATION OF TRANSMITTER

VHF Transceiver Figure 36


Page 2-47


2.31 VHF COMMUNICATION ANTENNA The purpose of the VHF antenna is to radiate and intercept radio signals in the VHF frequency range (118.00 – 136.00 MHz). The No 1 VHF antenna is located on top of the fuselage and VHF No 2 antenna is on the forward underside of the fuselage. Figure 37 shows the VHF antenna location on a Boeing 737 aircraft.

UPPER VHF AERIAL (NO 2 SYSTEM)

OPTION FOR 3RD VHF SYSTEM

LOWER VHF AERIAL (NO1 SYSTEM)

VHF Antenna Location (Boeing 737 aircraft) Figure 37

Page 2-48



2.32 SERVICE INTERPHONE

V

BOOM

MASKS

1 - ADF

ON

FORWARD ATTENDANT’S STATION

I/C

LIGHT NORM CALL NOT IN USE

ON

SERVICE

NOSE WHEELWELL

FLIGHT

PILOT

INTERPHONE EXTERNAL POWER

EXTERNAL POWER PANEL

FLIGHT COMPARTMENT

HEADSET

SPEAKER

MIC

PILOT’S CONTROL STAND

R/T

1 - NAV - 2

OFF

SERVICE INTERPHONE

AFT ATTENDANT’S STATION

R

MKR INOP

B

PA FLT INT SERV INT INOP

HF - 1 1 - VHF - 2

MIC SELECTOR

AUDIO SELECTOR PANEL

AUDIO ACCESSORY UNIT

The service interphone system allows communication between the flight crew, cabin attendants, ground crew or maintenance personnel. Jacks for plug-in microphone and headsets are installed at various locations in the aircraft. These jacks allow ground personnel to communicate with each other. An on/off switch on the aft overhead panel on the flight deck controls the external jacks. Handsets are available at the forward and aft attendant’s panels. Figure 38 shows the layout of the Boeing 737 interphone system.

Boeing 737 Interphone System Figure 38 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-49


2.32.1 Attendant Interphone Handsets The handsets provide the facility for introducing microphone audio into the system and for listening to audio from the systems other stations. The handsets resemble the common hand-held telephone receiver. A pushbutton switch is located on the grip to activate the mic. Handsets are permanently installed at each attendant’s station. Figure 39 shows the forward attendant’s interphone panel as fitted to the Boeing 737 aircraft.

MUSIC SYSTEM VOLUME CONTROL

LIGHTS

ATTENDANT’S CALL LIGHT

MUSIC

CAPTAIN’S CALL LIGHT

CALL SYSTEM

CAPTAIN ATTENDANT RESET

HANDSET HANDHELD MICROPHONE

Forward Attendant’s Interphone Panel Figure 39

Page 2-50



2.33 FLIGHT & GROUND CREW CALL SYSTEM The crew call system is a three-way alerting system that signals crew members to use the interphone system. The three types of crew call are:  Captain’s Call – A “Hi tone” chime sounds once and the Captain’s CALL light illuminates. This advises the flight crew that a call has been initiated from the attendant’s panel or ground crew panel.  Attendant Call – A “Two-tone” chime sounds and the “Pink Master Call” lights illuminates when an attendant’s call is initiated from the flight compartment or either attendant’s panels. The lights reset at the attendant’s panels.  Ground Crew Call – When the ground crew call is initiated in the flight compartment, a call horn sounds in the nose wheel well. Figure 40 shows the ground crew’s interphone panel, which is located in the nose wheel bay.


FLIGHT

PILOT

SERVICE NOSE WHEELWELL ON NORM

NOT IN USE

CALL LIGHT

Ground Crew’s Interphone Panel Figure 40


Page 2-51


2.34 PASSENGER ADDRESS SYSTEM (PA) The passenger address (PA) system provides a means of transmitting flight crew announcements, boarding music and chime signals to the passenger cabin. Audio inputs from the pilot’s, attendants and tape reproducer are prioritized by the PA amplifier. The priority is: 1. Pilots. 2. Cabin attendants. 3. Pre-recorded announcements. 4. Boarding music. The audio with the highest priority is amplified and distributed to the passenger cabin speakers, attendant’s speakers and audio integration. Figure 41 shows the layout of the PA system for the Boeing 737 aircraft.

PA AMPLIFIER PA SPEAKER PA MIC PA SPEAKERS

PA System (Boeing 737 aircraft) Figure 41

Page 2-52



2.35 AUDIO INTEGRATION SYSTEM

BOOM

MASKS I/C

OXYGEN MASK SPEAKER

HEADSET & BOOM MIC

HEADSET

CONTROL SWITCHES

HAND MIC

R/T

1 - NAV - 2

1 - VHF - 2

AUDIO SELECTOR PANEL

V

INOP 1 - ADF

INOP HF - 1

MIC SELECTOR

SERV INT

FLT INT

B

R

MKR

PA

Provides the flight crew with a means of controlling all radio communications, interphone and PA selection and Navigation receiver’s audio signals. Both pilots have their own individual system and control panel. Figure 42 shows a block schematic of the audio integration system.

Audio Integration System Figure 42


Page 2-53


2.36 CONTROL WHEEL MIC SWITCH The purpose of the INT/MIC switch is to provide PTT input for the boom or oxygen mask microphones. The switch is a three-position switch on the outboard horn of the captain’s and first officer’s control wheel. In the MIC position, mic audio is directed to the selected communication system. In the INT position, mic audio is connected directly to the flight interphone system. Figure 43 shows the control wheel INT/MIC switches of a Boeing 737 aircraft

MIC/INT SWITCH PRESS-TO-TALK

MICROPHONE

INTERPHONE

Control Wheel INT/MIC Switches Figure 43

Page 2-54


120.60

120.60


118.30

TEST

COMM

TFR

118.30

VHF COMM

TEST

COMM

TFR

VHF COMM

R/T

I/C I/C

1 - NAV - 2

1 - NAV - 2

R/T

HF - 1

1 - VHF - 2

1 - VHF - 2

SERV INT INOP

BOOM

MASK BOOM

MASK

1 - ADF

1 - ADF

HF - 1

INOP

V

INOP

PA

V

B R B

MKR

FLT INT

INOP

FLT INT SERV INT

MIC SELECTOR

MIC SELECTOR

R

MKR

PA

NO 2 VHF COMM TRANSCEIVER

NO 1 VHF COMM TRANSCEIVER



Figure 44 shows the layout for a two VHF communication system.

VHF Communication System Figure 44

Page 2-55


2.37 OPERATION OF VHF COMMUNICATION SYSTEM To operate the system, apply power to the transceiver and allow a short warm up period. Select the VHF comm to be used (1 or 2) on the audio selector panel. Select the frequency of the ground station that is going to be used and 'listen out' to ensure that no other transmissions are taking place. With the microphone close to the mouth, key the transmitter with the ac PTT switch and speak slowly and clearly into the microphone to establish 2-way communication with the ground station. Identify your position by airline and aircraft registration using the standard phonetic alphabet; 1. A = alpha. 2. B = Bravo etc Once you have finished speaking, release the PTT switch and listen for the reply, ensuring that it is loud and clear. Complete the check by confirming to the ground station the receipt of the reply. Some important points to note are: 1.

DO NOT TRANSMIT ON 121.50 MHz. This is a recognized emergency/distress channel.

2.

DO NOT TRANSMIT WHILST REFUELLING IS TAKING PLACE.

3.

DO NOT INTERRUPT ATC-AIRCRAFT COMMUNICATION.

Page 2-56



2.38 HIGH FREQUENCY (HF) RADIO COMMUNICATION The HF communication system (HF COMM) is for communication between the aircraft and ground stations. The ionosphere reflects the frequencies in the HF band, so the line of sight does not limit the reception range of the system. That is why the HF COMM is suitable for long range, worldwide communication. The frequency range of the system is 2 to 29.999 MHz. Frequency selection is made in 1 KHz steps, so there are 28000 channels available. There are two modes of operation. These modes are: 1. AM – AMPLITUDE MODULATION. 2. SSB – SINGLE SIDE BAND. In the AM mode the system transmits a carrier with amplitude modulation. In the SSB mode the carrier and the lower side band is removed. The system only transmits the upper side band (USB). Figure 45 shows a block schematic of a HF system.

HF COMM CONTROL PANEL

TUNING & CONTROL

POWER & CONTROL

RECEIVE TRANSMIT

HF COMM ANTENNA COUPLER

MIC KEY AUDIO OUT

HF COMM TRANSCEIVER

SELCAL SYSTEM

HF Communication Figure 45


Page 2-57


A HF aerial is of quite a different technology compared with most other aerials for two reasons. One is the power output of the Transmitter (400 watts) the other reason is that the quarter wavelength (/4) distance is about 40 metres at 2 MHz but only 2.5 metres at 30 MHz, so broadband aerials are not possible. Instead all HF aerials are fed from an aerial coupling unit to attempt to electrically lengthen or shorten the aerial for optimum matching, especially to the transmitter. Figure 46 shows the location of the HF antenna and coupling unit.

HF ANTENNA HF COUPLER UNIT

HF Antenna & Coupling Unit Locations Figure 46

Page 2-58



To get optimal power transfer from transmitter to antenna, the antenna impedance must be the same as the transmitter output impedance (50-Ohm). For each frequency the impedance of an antenna is different. Since the antenna on the aircraft has a fixed length, it is only suitable for one frequency. The antenna coupler tunes filters to adapt the antenna impedance for each different frequency to the transmitter output impedance. 2.38.1 HF Communication Control Panel The purpose of the HF Control Panel is to enable frequency selection, mode control and RF sensitivity adjustment. There are four frequency select controls for MHz, 100 KHz, 10 KHz and 1 KHz frequencies. The function selector allows selection of the operating mode as either: 1. OFF. 2. Upper Side Band (USB). 3. Lower Side Band (LSB). 4. Amplitude Modulation (AM). Note: LSB is reserved for military operations and normally, civil aircraft have the facility to select USB.


Page 2-59


Figure 47 shows a HF communication control panel.

SELECTED FREQUENCY DISPLAY

MEGA HERTZ SELECTOR

2 LSB

.0

KILO HERTZ SELECTOR

0

0

AM

USB OFF RF SENS

FUNCTION SELECTOR SWITCH

100 KILO HERTZ SELECTOR

10 KILO HERTZ SELECTOR

RECEIVER GAIN CONTROL

HF Communication Control Panel Figure 47

Page 2-60



2.39 SELECTIVE CALLING SYSTEM (SELCAL) The selective calling (SELCAL) system allows a ground station to call an aircraft or group of aircraft using HF and VHF communications without the flight crew having to continuously monitor the ground station’s frequency. A coded signal is transmitted from the ground station and received by the aircraft’s HF or VHF transciever tuned to the appropriate frequency. The output code is fed to a SELCAL decoder, which, activates aural and visual alerts if and only if the received code corresponds to the code, selected in the aircraft. Figure 48 shows the SELCAL System layout.

SE

A LC

L

C

VO

E OD

IC E

CO

M

M

UN

IC

AT

IO

N

AIRLINE DISPACH COMMUNICATION ARINC REMOTE STATION

ARINC VOICE STATION

SELCAL Operation Figure 48


Page 2-61


There are a total of 10920 codes available and these codes are assigned to airline organisations, who in turn assign codes to their individual aircraft either on a flight number or aircraft registration basis. 2.39.1 SELCAL Control Panel The SELCAL control panel consists of SELCAL warning lamps annotated to the associated radio system, i.e. VHF 1, VHF 2, HF 1 and HF 2. It also provides a means of resetting the SELCAL, thus cancelling the visual and audio indications. The panel also has a self-test button to allow testing of the SELCAL system. Normally located along with the control panel is the SELCAL code selection panel, this is used to set the aircraft’s SELCAL code. Figure 49 shows a SELCAL selector panel.

SELCAL CODE SELECTION PANEL

SELCAL ANNUNCIATOR PANEL

SELCAL Selector Panel & Code Select Panel Figure 49

Page 2-62



2.39.2 SELCAL Decoder The SELCAL decoder determines if the aircraft’s four-letter code has been received and produces alert signals in the form of indicators on the SELCAL control panel and audio tones to the audio system. The alerts are cancelled by pressing the corresponding alert light on the SELCAL control panel. A self-test of the alerts lights and audio warnings is carried out using the self-test button on the SELCAL control, panel. Figure 50 shows a block schematic of the SELCAL system.

LAMP DRIVES 5 WIRE RESET

CODE SELECT

TEST

AUDIO SYSTEM

CHIME SWITCH

VHF 1

LAMP SWITCH LAMP SWITCH

VHF 2

DECODER

VHF 3

LAMP SWITCH LAMP SWITCH

VHF 4 VHF 5 INTERRUPTER CIRCUIT

LAMP SWITCH

ELECTRICAL SUPPLY

SELCAL Block Schematic Figure 50


Page 2-63


2.40 COCKPIT VOICE RECORDER (CVR) The CVR records the last 30 minutes of the flight deck audio on continuous magnetic tape. All voice communication is recorded. Operation is automatic from engine start until five minutes after engine shutdown. The CVR receives sound from the flight compartment and audio signals from the digital audio control system. The voice recorder continuously records the sound and audio. Sensing of the “aircraft-on-ground” and “parking-brake-set” is used to permit bulk erasure of the voice recording. The system records on four channels: 1. Channel 1 - Records the third crew member’s summed microphone and telephone audio or passenger address system audio. 2. Channel 2 – Records the First Officer’s summed microphone and telephone audio. 3. Channel 3 – records the Captain’s summed microphone and telephone audio. 4. Channel 4 – records the control panel area microphone audio.

Page 2-64



ERASE

METER

MIC

4

3

2

1

A/C ON GROUND

TEST

RECORDING HEADS

PARKING BRAKE ON

ERASE HEAD

TAPE AND MOTOR

Figure 51 shows a block schematic of the Cockpit Voice Recorder System.

Cockpit Voice Recorder System Figure 51


Page 2-65


2.40.1 Voice Recorder Control Panel The control panel allows remote monitoring and testing of the voice recorder unit detects flight compartment sounds and conversations. It also controls bulk erasure of the recording tape. It contains an area microphone (capacitive) which senses compartment audio. Pressing the erase button for a minimum of 2 seconds erases the tape. This is only possible when the aircraft is on the ground and the parking brake is set. Pressing the TEST switch tests all 4 recording channels in sequence. The meter indicates green during TEST if the test tone is recorded at a sufficient level. The headset jack is used to monitor all 4 recorded channels. Figure 52 shows a CVR control panel.

COCKPIT VOICE RECORDER

0

TEST

2 4

6 8

10

ERASE

HEADSET 600 OHMS

CVR Control Panel Figure 52

Page 2-66



2.40.2 Voice Recorder Unit The voice recorder unit makes a 30 - minute recording of four audio channels on a continuous polyamide tape. The recorder is shock and heat resistant and contains an underwater locating beacon. It has a TEST switch to initiate an internal test signal to be recorded. A phone jack monitors the recording as it is being recorded. The Status indicator provides monitoring of the tape transport operation and the recorded signal during test. Figure 53 shows a voice recorder unit.

TEST

PHONE

STATUS IND

UNDERWATER LOCATING DEVICE

BATTERY LIFE LABEL

Voice Recorder Unit Figure 53 2.40.3 Underwater Locator Device The underwater locating device is a battery operated acoustical beacon that is activated when the unit is submerged in water. The unit provides a usable signal for 30 days. The battery replacement date decal is located on the front of the device.


Page 2-67


2.41 NAVIGATION SYSTEMS Aircraft navigation is simply a matter of knowing the direction in which we are flying and our position in relation to the earth's surface. Navigation through the air is a relatively simple matter of calculating the distance travelled in a given period of time. In today's modern aircraft we have various methods of assisting the crew to navigate their aircraft safely form A to B. These navigational aids are as follows: 1. Very High Frequency - Omni-Range - VOR. 2. Distance Measuring Equipment - DME. 3. Instrument Landing System - ILS. 4. Marker Beacon System MBS. 5. Automatic Direction Finder - ADF. 6. Air Traffic Control - ATC. 7. Traffic Alert & Collision Avoidance System – TCAS. 8. Inertial Navigation System - INS. 9. Radio Magnetic Indicator - RMI. 10. Global Positioning System - GPS. 11. Compass Systems 12. Radio Altimeter System – RADALT. 13. Weather Radar. 2.42 VERY HIGH FREQUENCY OMNI RANGE (VOR) VOR is an international standard navigational beacon system enabling a number of aircraft to receive signals from a ground station and determine the bearing to the station, with respect to magnetic North. This is possible because the VOR ground station, or transmitter, continually broadcasts an infinite number of directional radio beams or radials. The VOR signal received in an aircraft is used to operate a visual indicator from which the pilot determines the bearing of the VOR station with respect to the aircraft.

Page 2-68



2.42.1 VOR Operation The VOR operates in the frequency band from 108.00 to 117.95 MHz. The VOR ground-station transmits a combination of signals in all directions (omnidirectional). The VOR ground station modulates two signals of 30 Hz each on the carrier. One 30 Hz signal is the reference signal and the other is the variable signal. The phase shift between the reference signal and the variable signal depends on the radial over which the two signals are transmitted. The radial in the magnetic north direction has a phase shift of 0 degrees, the radial in the magnetic east direction (90 degrees) has a phase shift of 90 degrees, the radial in the magnetic south direction (180 degrees) has a phase shift of 180 degrees, etc. In this way the VOR ground station identifies each radial with the phase shift between the reference and the variable signal. Figure 54 shows a VOR ground station and corresponding transmitted frequencies.

000º

VOR RADIAL = 45° VOR BEARING = 225°

090º

270º VOR BEACON

180º VOR Ground Station Operation Figure 54


Page 2-69


The radial information is transmitted from the ground station to the aircraft. When the VOR system in the aircraft detects the phase shift between the reference and the variable signal it knows on which radial the aircraft flies. For the bearing information (opposite direction from aircraft to ground station) the VOR system adds 180 degrees. In this way the bearing to the station depends on the detected radial from the station. The phase shift between the reference and the variable signal identifies a radial with respect to the magnetic north. The bearing, which is a result from the detected radial, has therefore also a relation with the magnetic north. So the bearing output from the VOR system is a MAGNETIC bearing output. This information is displayed on a Radio Magnetic Indicator (RMI). Figure 55 shows a schematic of the VOR system.

RECEIVED VOR SIGNAL

30Hz AM DETECTOR

INTERMEDIATE FREQUENCY

RF

IF

AUDIO DETECTOR

ROTATION

TO PHASE DETECTOR

RADIO FREQUENCY

30Hz FM DETECTOR

REFERENCE

RMI VOR Indications Figure 56

Page 2-70



Figure 57 shows VOR control panel from a BAe 146 aircraft.

1 1 1 . 25 1 1 1 . 30 STBY

ACT

PRE

VOR/DME

NORM

ON

TEST

HOLD

I L S T E S T

DME

SPILT NAV 1

NAV 2

068

246

COURSE

COURSE

VOR Control & Course select Panels (BAe 146) Figure 57


Page 2-71


2.42.2 Deviation Calculations

VOR RADIAL POSITION WRT AIRCRAFT

SELECTED VOR RADIAL

“TO” FLAG IN VIEW

SELECTED COURSE

“FROM” FLAG IN VIEW

For lateral guidance in airways, the pilots can select a VOR course on the VOR control panel. The deviation from the selected course is calculated in the systems, which show course and deviation (EFIS) or use it for guidance (AFCAS). Also calculated from the difference between received radial and selected course is the information if the aircraft flies to or from a VOR station. The navigation display shows selected course, deviation, and the to-from information. Figure 58 shows HSI indications for a selected VOR course.

E

12 15

3

6

21

33

N

S

30 MILES

W

0 0 0

24

HSI Indications for a VOR Course Figure 58

Page 2-72



2.42.3 VOR Aerial Locations Figure 59 shows the location of the VOR aerials on a Boeing 737 and a Fokker 100 aircraft.

VOR AERIAL LOCATED ON TOP OF VERTICAL STABILISER

VOR AERIAL LOCATED ON EITHER SIDE OF VERTICAL STABILISER

VOR Aerial Locations (B737 & F100 Aircraft) Figure 59


Page 2-73


VOR operates in the VHF band responding to horizontally polarised transmissions. It shares its frequency range with the Localiser facility of the ILS and in so doing often shares the aerial system and much of the receiver unit. The aerial can be mounted on either side of the fin and much be Omni directional to receive VOR/ILS signals from all directions or flush mounted on either side of the nose section. Figure 60 shows the construction of a typical VOR/ILS aerial system.

No 2 SYSTEM

No 1 SYSTEM

No 1 SYSTEM

No 2 SYSTEM

VOR/ILS Aerial Figure 60

Page 2-74



I L S

TEST

ON HOLD

PRE

VOR/DME

W 3 0

24

15

3

VOR

COURSE

W

30 MILES

NAV 1

24

000

COURSE

21

33

N

S

SPILT

15

246

12

068

E

3

6

NAV 2

NAV SELECTOR

A D

F

21

33

N

S

COMPASS HEADING

VOR CONTROLLER

VOR RECEIVER

12

VOR

E

6

F

A D

STBY

DME

NORM

1 1 1 . 25 1 1 1 . 40

ACT

AUDIO

T E S T

VOR ANTENNA

Figure 61 shows a VOR system block diagram.

VOR Block Diagram Figure 61 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-75


2.43 DISTANCE MEASURING EQUIPMENT (DME) The Distance Measuring Equipment (DME) system gives distance information from the aircraft to the DME ground station. The system interrogates the ground station and the ground station gives a reply on every interrogation. The system then detects the time-delay between the transmitted interrogation and the received reply and from the time-delay the distance is calculated. Figure 62 shows the principle of DME operation.

DME Operation Figure 62

Page 2-76



The DME system operates in the UHF band and interrogates the ground stations in the frequency range from 1025 MHz to 1150 MHz. Within this frequency range the following ground stations are interrogated: 1. DME - Gives a reply on every DME interrogation. 2. VOR/DME - Combination of VOR and DME station and gives VOR bearing and distance replies. 3. ILS/DME - Combination of ILS and DME station and gives ILS guidance and distance replies. 4. MLS/DME - Combination of Microwave Landing System (MLS) and DME station. 5. TACAN - Military station for bearing and distance information for military aircraft. The civil aircraft use only the distance replies from these stations. 6. VOR/TAC - Combination of VOR and TACAN station and gives VOR bearing and distance replies. In addition to the distance reply, identification tones (1350Hz) are received from the ground station and may be heard as Morse code by the aircrew through headsets.


Page 2-77


Figure 63 shows the location of DME antennas.

DME No 1 DME No 2

BROADBAND L-BAND AERIAL

DME antenna Location Figure 63

Page 2-78



Figure 64 shows the principle of operation of the DME.

DME TRANSPONDER

AIRCRAFT

RX

TX TIMING INTERROGATION

RANGE CIRCUIT

50 SEC DELAY REPLY

TX

RX DISTANCE OUTPUT

DME Operation Figure 64 2.43.1 DME Operation In the DME system, the airborne unit transmits a 2-pulse group to the ground station at a random rate of 150 pulse pairs a second. After a 50 second delay, the ground station retransmits the pulse groups. Pulses are sent at one frequency and received at a different frequency, using the same antenna. Since many aircraft are using the DME facility, the aircraft equipment must be capable of selecting only those pulses that are replies from their own interrogations. A “Search and Track” circuit within the airborne equipment achieves the selection.


Page 2-79


The Search and Track circuit receives all DME replies and examines them to determine which ones have a regular time relation with respect to the transmitted signals. When the search circuit determines which received pulses are due to its own interrogations, the tracking unit locks onto them. At the same time, the pulse rate is greatly reduced; this in turn reduces the interrogation/replies at the ground station.

VHF/NAV CONTROL PANEL SUPPRESSOR BUS

TRANSMITTER

CONTROLLED VARIABLE DELAY

RCVR/XMITTR FREQUENCY SYNTHESISER

NAUTICAL MILES

1 2 3.5

DME

DELAYED TX LOCK ON

DUPLEXER

RECIEVER

AUDIO IDENTIFICATION

MATCHING CIRCUITS

FLIGHT INTERPHONE SYSTEM

INDICATOR

Figure 65 shows a basic DME system.

Basic DME System Figure 65 Page 2-80



2.43.2 DME Controller Figure 66 shows a NAV/VHF controller from a BAe 146 aircraft.

NAV/VHF Controller Figure 66 All DME frequencies are paired with either VOR or ILS system frequencies. When these system frequencies are selected, the associated DME facility will be automatically be selected. DME frequency range is 960 to 1215 MHz.


Page 2-81


2.44 INSTRUMENT LANDING SYSTEM (ILS) The purpose of the ILS is to provide approach information to the pilot when, due to weather, the runway is obscured from view. A typical system will allow the pilot to bring the aircraft to within ½ mile of the runway and less than 200ft above the runway without external visual reference. At these heights (Decision Height), the pilot must have visual on the runway and surrounding environment in order to continue the landing process. If the runway cannot be identified then a missed approach procedure is carried out. Aircraft will then be flown around the circuit for another attempt at landing. Aircraft are fitted with ILS in three categories, these are: Cat I - Operation down to a minimum of 200ft-decision height and runway visual range of 800m with a high probability of approach success. Cat II - Operation down to a minimum below 200ft decision height and runway visual range of 800m, and to as low as 100ft decision height and runway visual range of 400m with a high probability of approach success. Cat III - Three options A, B and C. A - Operation down to and along the surface of the runway, with external reference during final phase of the landing down to runway visual range minimum of 200m. B - Operation to and along the surface of the runway and taxiways, with external reference during final phase of the landing down to runway visual range minimum of 50m. C - Operation to and along the surface of the runway and taxiways without external visual reference.

Page 2-82



Figure 67 shows the different ILS categories.

CAT 1 200

CAT 2

A

B

C

100

CAT 3 800

600

400

200

50

RUNWAY VISUAL RANGE (METRES)

ILS Categories Figure 67 2.44.1 ILS Operation The ILS gives horizontal and vertical guidance in the approach to a runway. The system uses two radio signals: 1. The localizer for lateral guidance. 2. The Glideslope for vertical guidance. The localizer signal comes from a transmitter located at the end of the runway that operates in the frequency range from 108.000 - 111.95 MHz. The localizer transmits two beams one on the right side of the runway centerline and one on the left side of the runway centerline. The beam on the right side has a 150 Hz modulation; the one on the left side has a 90 Hz modulation. When the aircraft flies over the extended centerline to the runway it receives both signals with an equal strength. When the aircraft deviates from the centerline there is a difference in signal strength. The system measures the deviation from the center line by comparing the strength of these 90 Hz and 150 Hz modulation signals. MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-83

0


The Glideslope signal comes from a transmitter at the beginning of the runway that operates in the frequency range from 329.3 MHz to 335 MHz. The Glideslope transmits two beams to give vertical guidance over the glidepath. The glidepath has an angle of approximately 3°. The Glideslope beams are just like the localizer, modulated with 90 Hz and 150 Hz. The 90 Hz modulated beam is above and the 150 Hz modulated beam is below the 3° glidepath. The system measures the deviation from the difference in signal strength between the 90 Hz and 150 Hz modulation signals. Figures 67 and 68 show the localizer and Glideslope principles respectively.

Page 2-84



LOC TX

4º ON LONG RUNWAY 5º ON SHORT RUNWAY

700 ft WIDE AT THRESHOLD

2 DOT ENVELOPE (COURSE WIDTH)

DOTS ON HSI LATERAL DEVIATION


Localizer Principle Figure 67


Page 2-85


1,000 ft 1,000 ft

50 ft

100 ft

14 ft

3º

28 ft

2 DOT ENVELOPE COURSE WIDTH 1.4º

3,000 ft


Glideslope Principle Figure 68

Page 2-86



2.44.2 Antennas

GLIDESLOPE

VOR/LOC No 2 LOCATED ON THE OTHER SIDE

VOR/LOC AERIAL

VOR/LOC No 2 LOCATED ON THE OTHER SIDE

GLIDESLOPE No 1 & No 2

VOR/LOC AERIAL

Figure 69 shows the location of the antennas.

ILS Antenna Location (BAe 146 & F50 aircraft) Figure 69


Page 2-87


2.44.3 LOC/GS Operation Figure 70 shows a diagram for the LOC signal detection and display.

RF AMP WARNING FLAG IF AMP

DETECTOR

SUM

90Hz FILTER

250mV OUT OF VIEW

150Hz FILTER

DIFFERENCE

DEVIATION BAR

LOC Signal detection and Display Figure 70 The receiver of the Glideslope and Localiser operate in the same manner and include conventional “Radio Frequency (RF)”, “Intermediate Frequency (IF)” and Audio Frequency (AF)” stages. The output of the AF detector stage is the 90Hz and 150Hz signals. These are separated in there respective filters. The two signals are 180 out of phase and so oppose each other. The two signals are first summed together, and if the result is more than 250mV, the LOC/GS flag will be out of view (ILS valid). If the result of the summing is less than 250mV, the LOC/GS flags will remain in view (ILS invalid). If the 90Hz and 150Hz signals have the same amplitude, they cancel each other out in the difference circuit. This produces a 0V output to the deviation bar that is basically a centre reading dc voltmeter. With the output 0V the deviation bar will be central indicating the aircraft is positioned on the extended runway centerline (LOC) or on the glideslope.

Page 2-88



If the aircraft is positioned in the 90Hz signal lobe, then the amplitude of the 90Hz signal will be strongest. This will give a fly right signal (LOC) or fly down signal (G/S). If the difference is -75mV, the deviation bar will be located on the first dot right, if the difference is -150mV or more, then the deviation bar will be located on the second dot right. If the 150Hz is the stronger signal, then the voltage produced will be positive. This will give either fly left (LOC) or fly up (G/S). Because the result of the difference circuit is either a +dc half-cycle or –dc halfcycle, the signals are condensed using the capacitor, which will produce a steady dc signal. These condenser capacitors also damp the deviation bar movement.


Page 2-89


2.45 MARKER BEACON SYSTEM (MBS) In order to inform the pilot as to the aircraft's progress, during an ILS approach, along the centerline and Glideslope, there is a marker system. The markers are normally annotated as follows: 1. Outer marker. 2. Middle Marker. 3. Inner marker. Note: With Category II & III, ILS the inner marker is virtually non existent. The marker beacons transmit at a certain frequency to identify it and in a fan shaped pattern. They will also illuminate certain colour warning lamps within the flightdeck to inform the pilot of reaching the marker. Figure 71 shows the layout of the marker system for an ILS approach.

3000 Hz WHITE INDICATOR 1300 Hz AMBER INDICATOR

400 Hz BLUE INDICATOR

MORSE

MORSE MORSE

INNER MARKER

MIDDLE MARKER

OUTER MARKER

Marker Beacon System Figure 71

Page 2-90



The inner marker is not normally used with ILS, but is now used as an “Airways” marker, used for enroute navigation or as holding points above an airport. Airways markers are identified when the white light comes on and a 3,000Hz tone is heard. Outer and Middle markers are associated with the ILS. The outer marker is usually located directly below the point where an aircraft on a localizer course should intersect the Glideslope and start descending. An outer marker is identified when the blue light comes on and a 400 Hz tone is heard. The middle marker is located near the runway, usually under a point on the glidepath where a descent could be discontinued. The middle marker is identified when the amber light comes on and a 1,300 Hz tone is heard. A 75MHz carrier modulates all marker frequencies. Figure 72 shows the system layout.

75 MHz FILTER

MARKER BEACON SYSTEM

AUDIO AMPLIFIER

AUDIO (MORSE)

3000 Hz FILTER AMP

RF AMP & DETECTOR

MARKER

INNER

1300 Hz FILTER AMP

MIDDLE

400 Hz FILTER AMP

OUTER

HIGH

LOW

SENSITIVITY SWITCH

Marker Beacon System Layout Figure 72


Page 2-91


2.46 AUTOMATIC DIRECTION FINDER (ADF) The ADF system detects the direction to a Non Directional Beacon (NDB) and receives audio identification from the NDB. The ADF system shows the direction to the NDB on the instruments with the bearing pointer. The ADF system operates in the frequency range of 190 to 1750 KHz. The NDB ground station transmits an AM (Amplitude Modulated) signal in circular pattern in all directions. The radio energy induces RF (Radio Frequency) signals in a combined loop and sense antenna. The receiver antenna signals are measured in an ADF receiver and calculated to give relative station bearing. Figure 73 shows the operation of ADF.

MAGNETIC NORTH ADF2

ADF 1

300°

HEADING 30°

RF S AD IGN F 2 AL ST S F AT RO IO M N

60°

M RO F S N AL TIO N A SIG ST RF DF1 A

RELATIVE BEARINGS

ADF Operation Figure 73

Page 2-92



2.46.1 Loop Aerial A 'loop aerial', is very sensitive to its directional position, meaning that when it is pointing towards the transmitter, it receives a null signal but when pointing away from the transmitter, it receives a strong signal. This ability is used to automatically find the direction of the transmitter, relative to the aircraft heading and is displayed on the Radio Magnetic Indicator, RMI. Figure 74 shows the operation of a loop antenna. LOOP AERIAL AT 90° TO SIGNAL

NO CURRENT

INCREASING CURRENT

NO CURRENT

INCREASING CURRENT

NO CURRENT

RADIO TRANSMITTER

Loop Antenna Figure 74


Page 2-93


2.46.2 Station Line By turning the loop aerial to either of its two null positions the directions of a line joining the receiver with the transmitting station can be determined. This is called the “Station Line” and is shown in Figure 75.

A

B

Station Line Figure 75 As there are two nulls, 180 apart, the transmitter could be towards “A” or “B” in figure 75, causing ambiguity. To resolve this ambiguity it is necessary to change the directional properties of the aerial system. This is achieved by introducing a second aerial which combines its horizontal polar diagram with that of the loop aerial which produces a new heart shaped polar diagram called a “Cardioid”. Figure 76 shows the resultant Cardioid polar diagram from the loop and sense aerials.

Page 2-94



LOOP POLAR DIAGRAMS

SENSE POLAR DIAGRAM

CARDIOID POLAR DIAGRAM

Cardioid Polar Diagram Figure 76 2.46.3 Sensing the Correct Null The signal from the ADF transmitter induces a voltage into the loop using the magnetic component of the signal. The sense aerial has a voltage induced by the electric component of the signal. This produces 90 phase shift between the loop and sense aerial voltages. Whether the sense voltage leads or lags the loop voltage depends on which side of the station line the signal is coming from. Given a means of rotating the loop and switching the sense aerial into and out of the receiver input and a means of reversing the polarity of the loop signal to produce a Cardioid either to the right or left of a relative bearing pointer, aural sensing can be carried out. The bearing pointer is positioned along the loop axis in one direction as shown if figure 77.


Page 2-95


BEARING POINTER ALONG LOOP AXIS IN ONE DIRECTION

LEFT CARDIOID

RIGHT CARDIOID

ADF AERIAL

Bearing Point and Cardiods Figure 77 On tuning to the ADF beacon and listening to the audio signal, the loop is turned until a minimum signal is received. The loop is then offset in one direction by 15 - 20. The sense aerial is now switched in for the right Cardioid and the loudness of the audio noted. The loop is then reversed to give a left Cardioid, again the loudness is noted. The sense aerial is now switched out and the loop is tuned in the direction of the Cardiods, which gave the loudest signal. The first null the loop aerial reaches will be the correct one and the pointer will now be pointing the ADF transmitter the system is tuned to. Figure 78 shows this operation.

Page 2-96



ADF BEACON

ADF BEACON

RIGHT IS LOUDEST

LEFT IS NOT AS LOUD

POINTING TO WRONG NULL TURN TOWARDS LOUDEST FOR CORRECT NULL

LOOP OFFSET BY 15° - 20°

Calculating the Correct Null Figure 78


Page 2-97


In modern aircraft, the loop aerials are more streamlined and do not physically rotate (rotated electronically). Figure 79 shows the location of the ADF antenna.

NO 2 SENSE ANTENNA & COUPLER

NO 2 LOOP ANTENNA

NO 1 LOOP ANTENNA

NO 1 SENSE ANTENNA & COUPLER

ADF Antenna Location Figure 79

Page 2-98



Figure 80 shows a block schematic of the ADF system.

SENSE ANTENNA

FWD 90° SHIFT

A

BALANCED MODULATOR

RS

AUDIO

AUDIO DETECTOR

MIXER

47Hz OSC

47Hz FILTER

B ADF ANTENNA

TX

M

PHASE DETECTOR & MODULATOR

47Hz or ADF RECEIVER

RMI

AC SUPPLY

TR

ADF System Figure 80 The fixed loop antenna is preferred because it is more trouble free, due to fewer moving parts. The fixed loop consists of two loops orientated at 90 to each other. Each loop is connected to an individual stator of a receiving resolver within the ADF receiver. If the received station is directly ahead of the aircraft, loop “A” will have maximum signal and the “B” loop will have a null signal. In this case the resolver will see a null signal due to the orientation of the resolver rotor. If the aircraft is positioned to the right of the station, then the “A” will see a null and the “B” will see maximum signal. Intermediate positions of the received station would result in intermediate positions of the resultant filed in the stator of the resolver.


Page 2-99


Figure 81 shows a typical ADF control panel.

ADF ANT

ADF 1

OFF

191. 5

TEST

ADF 2

1231.5

ADF 2 BFO

1

2

NORM ADF ANT

OFF

TEST

ADF 1

ADF Control panel Figure 81

Page 2-100



2.47 AIR TRAFFIC CONTROL RADIO BEACON SYSTEM (ATCRBS) 2.47.1 Transponders Transponders are not exactly navigation equipment, but are a "Means of Identification". In the past, a radar controller watching his scope, would only know if the 'blip' that he saw on his screen was the aircraft he was 'working', (handling), if it identified itself by carrying out a turn at the controller's request. With the ATC system, the controller can identify the aircraft by interrogating it. The ground control has two types of radar with which to control air traffic: 1. Primary Radar. 2. Secondary Radar. The primary radar provides the ground station operator with a symbol on his surveillance radarscope for every aircraft in his area. It is a reflection type of radar system not requiring any response from the aircraft. The secondary radar system uses what is called an “ATC Transponder” in the aircraft. The transponder is a transmitter/receiver, which transmits in response to an interrogation from the ground station secondary surveillance radar system. The primary and secondary radar antennas are mounted on the same rotating mounting, and therefore both always look in the same direction at the same time. The aircraft’s transponder reply can also include a special code, which identifies that particular aircraft on the scope. If the pilot receives instructions from the ground station to do so he presses his “Ident” button on his control panel. This causes the display on the radarscope to change thus identifying the aircraft to the controller. The transponder can also transmit the aircraft’s altitude, which can be displayed to the ground controller.


Page 2-101


Figure 82 shows the operation of the ATCRBS.

GROUND SURVEILLANCE RADAR

ATC RADAR ANTENNAS

ATC RADAR TRANSMITTER/ RECEIVER

ATCRBS Operation Figure 82

Page 2-102



2.47.2 ATCRBS Control Panel The ATCRBS control panel allows the flight crew to select ATC 1 or 2, mode of operation and ident code select. Figure 83 shows a typical ATCRBS control panel.

MODE

STBY

A

1

B

2

2567 ALT RPTG

ALT IDENT

1

2 OFF

SOURCE

ATCRBS Control Panel Figure 83 The ground station transmits its interrogation pulse on 1030 MHz as a three-pulse signal. The space between the first and third pulse signifies the mode reply required. The system operates in four modes, these are: 1. Mode A - Identify. 2. Mode B - Obsolete. 3. Mode C - Pressure Altitude. 4. Mode D – Unassigned. 2.47.3 Mode A Operating mode for normal operation. The transponder is ready to respond to ATC any interrogations and replying with a unique identification code. The pulse spacing is 8sec.


Page 2-103


2.47.4 Mode C Altitude reporting capability of the transponder. The aircraft's Air Data System will supply altitude information for use in Mode C replies. This allows the ground controller, to not only identify an aircraft but also to ascertain its altitude, so he can guide it safely through his allocated airspace. The pulse spacing is 21sec. Figure 84 shows the interrogation pulses for mode A & C.

8 SEC

P1

P3 P2

MODE “A” IDENTITY ONLY 21 SEC

P1

P3 P2

MODE “C” IDENTITY & ALTITUDE

Mode A & C Interrogation Pulses Figure 84 Once the aircraft’s transponder has received an interrogation, it will reply with either Mode A or C (1090 MHz). One problem to overcome with this system is an aircraft replying to interrogations when not being illuminated by the primary radar. To overcome this, a suppression pulse is transmitted (P2). If the amplitude of this pulse is equal/greater than P1, the aircraft will not reply to the interrogation. Figure 85 shows the operation of the suppression signal.

Page 2-104



P1

OMNI DIRECTIONAL ANTENNA (P2)

P2

NO REPLY

P3

SIDELOBES

P1

ROTATION

P2

REPLY

P3

DIRECTIONAL ANTENNA MAIN BEAM (P1, P2 & P3)


ATCRBS Suppression Figure 85 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-105


Figure 86 shows the basic layouts of the ATCRBS.

R EP LY A N D FA U LT LIGH T C ON TR OL

A IR D A TA C O M P U TE R NO 1

A LT R PTG ON

N O 1 A TC T RA N S P O N DE R N0 1 A TC AE R IAL

N O 1 EN A B LE

A TC R BS D UA L C O N TR O L

M OD E, 40 9 6, ID EN T SU P P N O 2 EN A B LE

N O 2 A TC T RA N S P O N DE R

A IR D A TA C O M P U TE R NO 2

A LT R PTG ON

N0 2 A TC AE R IAL

R EP LY A N D FA U LT LIGH T C ON TR OL

V ID E O SU PP

C O M P A RA T O R P1 - P2

R E CE IV E R M O D E A or C

D E CO DE R

A TC R BS T RA N S P O N DE R

D IP LE X E R 1 0 3 0 M Hz

M OD E SW SEL F T EST C IRC U ITS

S E L FT E S T 4096

T RA N S M ITT E R 1 0 9 0 M Hz

M O D U LA TO R 1 0 9 0 M Hz

ID E NT

E N CO DE R

E N CO DE D H E IG H T

ATCRBS Block Schematic Figure 86

Page 2-106



2.48 MODE S TRANSPONDERS After 1989, a completely new type of ATC system was introduced. This system is called mode S (mode select). The new interrogators and transponders are called ATCRBS/mode S because they are capable of working with the old ATCRBS equipment or with new mode S equipment. For the present time, there will be ATCRBS only equipped aircraft sharing airspace with ATCRBS/mode S equipped aircraft. On the ground, most of the stations are ATCRBS-only, but there will be a gradual phasing in of ATCRBS/mode S ground stations. Both types of station can interrogate either type of transponder, and both types of transponder can respond to either type of ground station. TCAS-equipped aircraft interrogate both ATCRBS and ATCRBS/mode S equipped aircraft just as an ATCRBS/mode S ground station would do. At some point in the future, all ATCRBS-only equipment will be phased out for commercial aviation. All ground stations and aircraft will then operate in mode S only. The mode S ATC system enables ground stations to interrogate aircraft as to identification code and altitude just as the ATCRBS system does. These interrogations, however, are only part of a larger list of (up-link and downlink) formats comprising the mode S data link capacity. One of the most important aspects of mode S is the ability to discretely address one aircraft so that only the specific aircraft being interrogated responds, instead of all transponder-equipped aircraft within the range of the interrogator. 2.48.1 Mode S Interrogation & Replies The ATCRBS/Mode S system operates in a way similar to ATCRBS. As a transponder equipped aircraft enters the airspace, it receives either a Mode S only all-call interrogation or an ATCRBS/Mode S all-call interrogation which can be identified by both ATCRBS and Mode S transponders. ATCRBS transponders reply in Mode A and Mode C, while the Mode S transponder replies with a Mode S format that includes that aircraft's unique discrete 24-bit Mode S address. The Mode S only all-call is used by the interrogators if Mode S targets are to be acquired without interrogating ATCRBS targets. 2.48.2 Discrete Addressing The address and the Location of the Mode S aircraft is entered into a roll-call file by the Mode S ground station. On the next scan, the Mode S aircraft is discretely addressed. The discrete interrogations of a Mode S aircraft contain a command field that may desensitise the Mode S transponder to further Mode S all-call interrogations. This is called Mode S lockout. ATCRBS interrogations (from ATCRBS only interrogators) are not affected by this lockout. Mode S transponders reply to the interrogations of an ATCRBS interrogator under all circumstances.


Page 2-107


TCAS separately interrogates ATCRBS transponders and Mode S transponders. During the Mode S segment of the surveillance update period, TCAS commences to interrogate Mode S intruders on its own roll-call list. Because of the selective address features of the Mode S system, TCAS surveillance of Mode S- equipped aircraft is straightforward. Figure 87 shows “Mode S” operation.

TRANSPONDER REPLY 1090MHz

INTERROGATION 1030MHz

PRIMARY RADAR ECHO

PRIMARY SURVEILLANCE RADAR (PSR)

SECONDARY SURVEILLANCE RADAR (SSR) ATC RADAR SCOPE ROLL CALL AIRPLANE 1 AIRPALNE 2 AIRPLANE 3

GROUND LINK

NEIGHBORING AIRSPACE CONTROLLER (MODE S)

Mode S Operation Figure 87 2.48.3 Operation As a Mode S aircraft flies into the airspace served by another Mode S interrogator, the first Mode S interrogator may send position information and the aircraft's discrete address to the second interrogator by way of ground lines. Thus the need to remove the lockout may be eliminated, and the second interrogator may schedule discrete roll-call interrogations for the aircraft. Because of the discrete addressing feature of Mode S, the interrogators may work at a lower rate (or handle more aircraft).

Page 2-108



NO REPLY P6 P2 P1

P5

P6 P2 P1

P1

P2

P5

P3

P4

REPLY

NO REPLY

ONLY MODE S DISCREETLY ADDRESSED REPLIES

REPLY MODE S (UNLESS LOCKED OUT)

REPLY MODE S (UNLESS LOCKED OUT)

NO REPLY REPLY P3 P2 P1

P2 P1

INTERROGATION PULSE

P3

P4

REPLY

ATCRBS TRANSPONDERS

REPLY ATCRBS

ATCRBS/MODE S TRANSPONDERS

In areas where Mode S interrogators are not connected by way of ground lines, the protocol for the transponder is for it to be in the lockout state for only those interrogators that have the aircraft on the roll-call. If the aircraft enters airspace served by a different Mode S interrogator, the new interrogator may acquire the aircraft via the replay to an all-call interrogation. Also, if the aircraft does not receive an interrogation for 16 seconds, the transponder automatically cancels the lockout. Figure 88 shows the different types of interrogation pulses for ATCRBS and Mode S systems

ATCRBS & Mode S Interrogation Signals Figure 88


Page 2-109


2.49 TRAFFIC ALERT AND COLLISION AVOIDANCE SYSTEM 2.49.1 TCAS Introduction TCAS is an airborne traffic alert and collision avoidance advisory system, which operates without support from ATC, ground stations. TCAS detects the presence of nearby intruder aircraft equipped with transponders that reply to Air Traffic Control Radar Beacon Systems (ATCRBS) Mode C or Mode S interrogations. TCAS tracks and continuously evaluates the threat potential of intruder aircraft to its own aircraft and provides a display of the nearby transponder-equipped aircraft on a traffic display. During threat situations TCAS provides traffic advisory alerts and vertical maneuvering resolution advisories to assist the flight crew in avoiding mid-air collisions. TCAS I provides proximity warning only to assist the pilot in the visual acquisition of intruder aircraft. It is intended for use by smaller commuter and general aviation aircraft. TCAS II provides traffic advisories and resolution advisories (recommended escape maneuvers) in a vertical direction to avoid conflicting traffic. Airline, larger commuter and business aircraft will use TCAS II equipment. TCAS III Still under development, will provide traffic advisories and resolution advisories in the horizontal as well as the vertical direction to avoid conflicting traffic. The level of protection provided by TCAS equipment depends on the type of transponder the target aircraft is carrying. It should be noted that TCAS provides no protection against aircraft that do not have an operating transponder.

Page 2-110



Table 1 shows levels of protection offered by the transponder carried by individual aircraft. TCAS I

OWN AIRCRAFT TCAS II

TCAS III

TA

TA

TA

Mode C Or Mode S XPDR

TA

TA VRA

TA VRA HRA

TCAS I

TA

TCAS II

TCAS III

TARGET AIRCRAFT EQUIPMENT

Mode A XPDR Only

TA VRA

TA VRA HRA

TA

TA VRA TTC

TA VRA HRA TTC

TA

TA VRA TTC

TA VRA HRA TTC

TA – TRAFFIC ADVISORY VRA - VERTICAL RESOLUTION ADVISORY HRA - HORIZONTAL RESOLUTION ADVISORY TTC - TCAS – TCAS COORDINATION

Table 1


Page 2-111


2.49.2 The TCAS II System TCAS II provides a traffic display and two types of advisories to the pilot. One type of advisory, called a traffic advisory (TA) informs the pilot that there are aircraft in the area, which are potential threats to his own aircraft. The other type of advisory is called a resolution advisory (RA), which advises the pilot that a vertical corrective or preventative action is required to avoid a threat aircraft. TCAS II also provides aural alerts to the pilot. Figure 89 shows TCAS protection area.

TCAS Protection Area

Page 2-112



Figure 89 When a Mode S or Mode C intruder is acquired, TCAS begins tracking the intruder. Tracking is performed by repetitious TCAS interrogations in Mode S and Mode C. When interrogated transponders reply after a fixed delay. Measurement of the time between interrogation transmission and reply reception allows TCAS to calculate the range of the intruder. If the intruder's transponder is providing altitude in its reply, TCAS is able to determine the relative altitude of the intruder.

AURAL ALERT

MODE S/TCAS CONTROLLER TA/RA

DATA BUS

TA/RA OMNI DIRECTIONAL ANTENNA

DIRECTIONAL ANTENNA

TCAS COMPUTER UNIT

RADAR ALTIMETER

BAROMETRIC ALTIMETER

MODE S TRANSPONDER UNIT

OMNI DIRECTIONAL ANTENNA

OMNI DIRECTIONAL ANTENNA

Figure 90 shows a block schematic diagram of the TCAS system

TCAS System Block Schematic Figure 90


Page 2-113


Transmission and reception techniques used on TCAS directional aerials allows TCAS to calculate the bearing of the intruder. Based on closure rates and relative position computed from the reply data, TCAS will classify the intruders as non-threat, proximity, TA, or RA threat category aircraft. If an intruder is being tracked, TCAS displays the intruder aircraft symbol on an electronic VSI or joint-use weather radar and traffic display. Alternatively in some aircraft the TCAS display will be on the EFIS system. The position on the display shows the range and relative bearing of the intruder. The range of TCAS is about 30 NM in the forward direction. Figure 91 shows TCAS TA and RA calculations.

SURVEILLANCE

OWN AIRCRAFT

TRACK & SPEED

TRACKING

TARGET AIRCRAFT

BEARING & CLOSING PEED

TRAFFIC ADVISORY (TA) RANGE TEST

ALTITUDE TEST THREAT DETECTION (RA)

SENSE SELECTION

CLIMB DECENT

RA TCAS/TCAS CO-ORDINATION

RATE OF CLIMB/DECENT

STRENGTH SELECTION

RA

TA

ADVISORY ANNUNCIATION (TA/RA)

AIR GROUND COMMUNICATION

ATC

TCAS RA and TA Calculations Figure 91

Page 2-114



2.49.3 Aural Annunciation Displayed traffic and resolution advisories are supplemented by synthetic voice advisories generated by the TCAS computer. The words "Traffic, Traffic" are annunciated at the time of the traffic advisory, which directs the pilot to look at the TA display to locate the intruding aircraft. If the encounter does not resolve itself, a resolution advisory is annunciated, e.g., "Climb, Climb, Climb". At this point the pilot adjusts or maintains the vertical rate of the aircraft to keep the VSI needle out of the red segments. Figure 92 gives an overview of TCAS air-to-air operation.

AIRCRAFT 2 TCAS AIRCRAFT 2 RECIEVES SQUITTER AND ADDS AIRCRAFT 1 TO ITS ROLL CALL, THEN INTERROGATES AIRCRAFT 1 (TCAS 1030 MHz)

AIRCRAFT 2 TRANSMITS ATCRBS ALL CALL (1030 MHz) AIRCRAFT 3 RESPONDS MODE C (1090 MHz)

AIRCRAFT 3 ATCRBS ONLY AIRCRAFT 1 MODE S ONLY

AIRCRAFT 1 TRANSMITS OMNIDIRECTIONAL SQUITTER SIGNALS (MODE S 1090 MHz) ALL 3 AIRCRAFT REPLY TO INTERROGATIONS FROM GROUND STATION (1090 MHz) GROUND STATION TRANSMITS INTERROGATIONS AT (1030MHz)

NOTE:

TCAS OPERATION IS COMPLETELY INDEPENDENT OF GROUND STATION OPERATION

TCAS Air-to-Air Operation Figure 92


Page 2-115


Figure 93 shows typical Electronic VSI - TCAS indications.

Honeywell

1

2

4

.5

6

+03 -05

0 -03

6

.5

1

2

4

Electronic VSI - TCAS indications Figure 93

Page 2-116



Figure 94 shows examples of TCAS warnings as displayed on EADI.

HOLD

LNAV

VNAV

LOC 110.90

142

DME

G/S

VERTICAL SPEED LINE

DH150

25.3

2400

CMD

5200

180 5000 160

10

6 2

10

1

14

4800 REF

120

1

10

10

4600

2 6

MDA

CRS 123

4700

100

FLY OUT OF AREA

STD 117

MAG

29.86IN 750

RA FLIGHT BOUNDARY (RED)

6 2 1

GREEN SEGMENT

1

RED SEGMENT

2 6

VERTICAL SPEED LINE

TCAS Warnings EADI Display Figure 94


Page 2-117


Displayed traffic and resolution advisories are supplemented by synthetic voice advisories generated by the TCAS computer. The words "Traffic, Traffic" are annunciated at the time of the traffic advisory, which directs the pilot to look at the TA display to locate the traffic. If the encounter does not resolve itself, a resolution advisory is annunciated. The aural annunciation’s listed in Table 2 have been adopted as aviation industry standards. The single announcement "Clear of Conflict" indicates that the encounter has ended (range has started to increase), and the pilot should promptly but smoothly return to the previous clearance. Traffic Advisory: TRAFFIC, TRAFFIC Resolution Advisories: Preventative: MONITOR VERTICAL SPEED, MONITOR VERTICAL SPEED. Ensure that the VSI needle is kept out of the lighted segments. Corrective: CLIMB-CLIMB-CLIMB. Climb at the rate shown on the RA indicator: normally 1500 fpm. CLIMB.CROSSING CLIMB-CLIMB, CROSSING CLIMB. As above except that it further indicates that own flightpath will cross through that of the threat. DESCEND-DESCEND-DESCEND. Descend at the rate shown on the RA indicator: normally 1500 fpm. DESCEND, CROSSING DESCEND-DESCEND, CROSSING DESCEND. As above except that it further indicates that own flight path will cross through that of the threat. REDUCE CLIMB-REDUCE CLIMB. Reduce vertical speed to that shown on the RA indicator. INCREASE CLIMB-INCREASE CLIMB. Follows a "Climb" advisory. The vertical speed of the climb should be increased to that shown on the RA indicator nominally 2500 fpm. INCREASE DESCENT-INCREASE DESCENT. Follows a "Descend" advisory. The vertical speed of the descent should be increased to that shown on the RA indicator: nominally 2500 fpm. CLIMB, CLIMB NOW-CLIMB, CLIMB NOW. Follows a "Descend" advisory when it has been determined that a reversal of vertical speed is needed to provide adequate separation. DESCEND, DESCEND NOW-DESCEND. DESCEND NOW. Follows a "Climb" advisory when it has been determined that a reversal of vertical speed is needed to provide adequate separation. Table 2

Page 2-118



2.49.4 Performance Monitoring It is important for the pilot to know that TCAS is operating properly. For this reason a self-test system is incorporated. Self-test can be initiated at any time, on the ground or in flight, by momentarily pressing the control unit TEST button. If TA's or RAs occur while the self-test is activated in flight, the test will abort and the advisories will be processed and displayed. When self-test is activated an aural annunciation "TCAS TEST" is heard and a test pattern with fixed traffic and advisory symbols appears on the display for eight seconds. After eight seconds "TCAS TEST PASS" or "TCAS TEST FAIL" is aurally announced to indicate the system status. 2.49.5 TCAS Units Figure 95 shows a typical Mode S/TCAS control unit.

XPDR FAIL

ATC C A S

7777

IDENT

ALT RPTG 1

OFF

XPDR ON STBY

TA /RA

TCAS TEST

2

TA

AS TC

TA DSPLY AUTO OFF ON

XPDR 1

2

Dual Mode S Control Unit Figure 95


Page 2-119


The controls operate as follows: (1)

Transponder Code Display This shows the ATC code selected by the two dual concentric knobs below the display. The system select switch (XPDR 1-2) controls input to the display. Certain fault indications are also indicated on the display. "PASS" will show after a successful functional test and "FAIL" will show if a high level failure is detected under normal operating conditions. Also shown is the active transponder by displaying ATC 1 or 2.

(2)

Mode Control Selector Switch This is a rotary switch labeled STBY-ALT RPTG OFF-XPNDR-TA-TA/RA. The TCAS system is activated by selecting traffic advisory (TA) or traffic and resolution advisory (TA/RA). When STBY is selected both transponders are inactive. In the ALT RPTG OFF position the altitude data sources are interrupted preventing the transmission of altitude.

(3)

ABV-N-BLW Switch This selects the altitude range for the TCAS traffic displays. In the ABV mode the range limits are 7,000 feet above and 2,700 feet below the aircraft. In the BLW mode the limits are 2,700 feet above and 7,000 feet below. When normal (N) is selected the displayed range is 2,700 feet above and below the aircraft.

(4)

Traffic Display Switch When AUTO is selected the TCAS computer sets the displays to "pop-up" mode under a traffic/resolution advisory condition. In MAN the TCAS displays are constantly activated advising of any near by traffic.

(5)

Range Switch This selects different nautical mile traffic advisory horizontal range displays.

(6)

IDENT Push-button When pushed causes the transponder to transmit a special identifier pulse (SPI) in its replies to the ground.

(7)

Flight Level Push-button (FL) This is used to select between relative and absolute attitude information.

Page 2-120



Figure 96 shows a TCAS & Mode S computers.

ATC TPR/MODE S Honeywell

BENDIX/KING

RT-950 TCAS COMPUTER UNIT

TPR "SELF TEST" Replace TCAS CU if ONLY the red TCAS Fail lamp is on during any status display (following the lamp test). When additional lamps are on, correct indicated subsystem PRIOR to replacement of TCAS CU.

TCAS PASS

TA DISP

ALT

TCAS FAIL

RA DISP

DATA IN

TOP ANT

RAD ALT

TOP

BOT ANT

XPDR BUS

BOT

HDG

ATT

DATA LOADER

TCAS MAINT

PUSH TO TEST

RESERVED RESERVED

BITE

TEST

TCAS COMPUTER MODE S COMPUTER

Honeywell TCAS & Mode S Computers Figure 96 2.49.6 Self Test If the test button is momentarily pressed fault data for the current and previous flight legs can be displayed on the front panel annunciators. When the TEST is initially activated all annunciators are on for 3 seconds and then current fault data is displayed for 10 seconds, after which the test terminates and all annunciators are extinguished. If the test button is pressed again during the 10-second fault display period the display is aborted and a 2-second lamp test is carried out. The fault data recorded for the previous flight leg is then displayed for 10 seconds. This procedure can be repeated to obtain recorded data from the previous 10 flight legs. If the test button is pressed to display fault data after the last recorded data all annunciators will flash for 3 seconds and then extinguish.


Page 2-121


2.49.7 Data Loader Interface Software updates can be incorporated into the computer via a set of ARINC 429 busses and discrete inputs. These allow an interface to either an Airborne Data Loader (ADL) through pins on the unit's rear connector or to a Portable Data Loader (PDL) through the front panel "DATA LOADER" connector. The computer works with either ARINC 603 data loader low speed bus or ARINC 615 high-speed bus. A personal computer (PC) can be connected to the front panel "DATA LOADER" connector. This allows the maintenance log and RA event log to be downloaded to the PC via an RS 232 interface.

Page 2-122



2.50 INERTIAL NAVIGATION SYSTEM (INS) The modern inertial navigation system is the only self-contained single source for all navigation data. After being supplied with initial position information, it is capable of continuously updating extremely accurate displays of the aircraft’s: 1.

Position.

2.

Ground Speed.

3.

Attitude.

4.

Heading.

It can also provide guidance and steering information for the auto pilot and flight instruments. Figure 97 shows a representation of Inertial Navigation principal. Navigation Triangle

TRK

CK RA ED T E ’S FT DSP A N R RC O U AI G R &

DRIFT

HDG

EAST/WEST VELOCITY (VE)

Basic Navigation triangle Figure 97


PRESENT POSITION

VELOCITY NORTH/SOUTH (VN)

AI R & CRA AI F RS T’ PE S H ED EA (A DIN DC G )

WIND SPEED & DIRECTION

Page 2-123


2.50.1 General Principle In order to understand an inertial navigation system we must consider both the definition of “Inertia” and the basic laws of motion as described by Sir Isaac Newton. Inertia can be described as follows: 1. Newton’s first law of motion states: “A body continues in a state of rest, or uniform motion in a straight line, unless acted upon by an external force”. 2. Newton’s second law of motion states: “The acceleration of a body is directly proportional to the sum of the forces acting on the body.” 3. Newton’s third law states: “For every action, there is an equal and opposite reaction”.

With these laws we can mechanise a device which is able to detect minute changes in acceleration and velocity, ability necessary in the development of inertial systems. Velocity and distance are computed from sensed acceleration by the application of basic calculus. The relationship between acceleration, velocity and displacement are shown in figure 98.

Page 2-124



TIME

DISTANCE IN FEET

VELOCITY FEET PER SECOND

ACCELERATION FEET PER SECOND PER SECOND


Acceleration, Velocity and Distance Graphs. Figure 98 Note; Velocity changes whenever acceleration exists and remains constant when acceleration is zero.


Page 2-125


2.50.2 INS Operation The basic measuring instrument of the inertial navigation system is the accelerometer. Two accelerometers are mounted in the system. One will measure the aircraft’s accelerations in the north-south direction and the other will measure the aircraft’s accelerations in the east-west direction. When the aircraft accelerates, the accelerometer detects the motion and a signal is produced proportional to the amount of acceleration. This signal is amplified, current from the amplifier is sent back to the accelerometer to a torque motor, which restores the accelerometer to its null position. The acceleration signal from the amplifier is also sent to an integrator, which is a time multiplication device. It starts with acceleration, which is in feet per second squared (feet per sec per sec) and end up after multiplication by time with velocity (feet per second). The velocity signal is then fed through another integrator, which again is a time multiplier, which gives a result in distance in feet. So from an accelerometer we can derive: 1.

Ground Speed.

2.

Distance Flown.

If the computer associated with the INS knows the latitude and longitude of the starting point and calculates the aircraft has travelled a certain distance north/south and east/west it can calculate the aircraft’s present position.

Page 2-126



PRESENT POSITION START POSITION

ACCELEROMETER

START POSITION

MASS

RECENTRING (FEEDBACK)

COMPUTER

PRESENT POSITION

1ST

INTERGRATORS

DISTANCE FLOWN

2ND

VELOCITY GROUNDSPEED

DESTINATION

DISTANCE

Figure 99 shows INS Operation.

INS Operation Figure 99


Page 2-127


To accurately compute the aircraft’s present position, the accelerometer must be maintained about their sensing axes. To maintain the correct axes, the accelerometers are mounted on a gimbal assembly commonly referred to as the platform. The platform is nothing more than a mechanical device, which allows the aircraft to go through any attitude change at the same time maintaining the accelerometers level. The inner element of the platform contains the accelerometers as well as gyroscopes to stabilize the platform. The gyros provide signals to motors, which in turn control the gimbals of the platform. Figure 100 shows an Inertial Platform (IP).

AZIMUTH AXIS

ROLL AXIS

PITCH AXIS

Inertial Platform Figure 100

Page 2-128



We can also measure the angular distance between the aircraft and the platform in the three axes, giving us the aircraft’s pitch, roll and heading angles. These can be used in the navigation computations and also give heading and attitude information to the relative systems. The gyro and accelerometer are mounted on a common gimbal. When this gimbal tips off the level position, the spin axis of the gyro remains fixed. The case of the gyro moves with the gimbal, and the movement is detected by a signal pick-off within the gyro. This signal is amplified and sent to the gimbal motor, which restores the gimbal back to the level position. Figure 101 shows the operation of gyro stabilization.

INPUT AXIS RATE GYROSCOPE OUTPUT AXIS

PLATFORM

AMPLIFIER

GEARS

MOTOR

TACHO GEN

Gyro Stabilization Figure 101


Page 2-129


2.50.3 Earth Rate Compensation The INS gyro operates on the principle of gyroscopic inertia, which is the characteristic of a rotating mass to resist any forces, which tend to change the direction of its spin axis. Because the earth rotates in space, the spaceorientated gyro appears to rotate with respect to an earth bound observer. This makes the gyro unsuitable for use as an earth-fixed reference unless the gyro is deliberately torqued to rotate at a rate proportional to the earth’s rotational rate (earth rate = 15º/hour). When torqued in this manner, the spin axis appears stationary, and the gyro is effectively slaved to the earth’s co-ordinate system. Figure 102 shows the calculations of earth rate for the north and vertical gyros. E A RT H RA TE = (1 5 °/H R ) X C O S LA TIT UD E

E A R TH R AT E C O M P E N S A T IO N

= 15 D EG /H R

9 0°

A T 0 ° L A T IT U D E N O R TH G Y R O


4 5°

= 0 D E G /HR

A T 9 0 ° L A T IT U D E




= 1 0 . 6 º /H R

0°


E A RT H RA TE = (1 5 °/H R ) X S IN L AT ITU D E


9 0° = 0 D E G /HR

A T 0 ° L A T IT U D E V E RT ICA L G Y R O


4 5°

= 15 D EG /H R





= 1 0 . 6 º /H R

0°


North Gyro Earth Rate Calculation Figure 102

Page 2-130



2.50.4 Vehicle Rate Compensation These corrections are used to keep the platform horizontal and pointing to north. The aim is to cancel out the apparent movement of the gyro as the aircraft moves over the earth’s surface. These corrections are applied to all three gyros as torque to the gyro torque motor, the amount of torque being dependant on the direction of the aircraft movement over the earth’s surface. 

Aircraft Moving North

This will cause the platform to move away from its horizontal attitude. This effect is corrected by applying a signal to the East gyro’s torque motor. The strength of the signal is dependant on the angular rate of change which is found out by the following formula:

=

AIRCRAFT’S VELOCITY EARTH’S RADIUS

Figure 103 shows the vehicle rate corrections for an aircraft travelling North.

Aircraft Travelling North Figure 103


Page 2-131




Aircraft Moving East

This again causes the platform to move away from its horizontal attitude. This is corrected by applying a signal to th torque motor of the North gyro. Figure 104 shows the vehicle rate corrections for an aircraft travelling East with the North axis rotated.

Aircraft Travelling East (North Axis Rotated) Figure 104

Page 2-132



When moving east at any latitude other than the equator, the movement also causes the platform to move away from pointing north. To correct this, we apply a signal to the torque motor on the vertical gyro. The size of the signal is dependant on the latitude and the sped of the aircraft. Figure 105 shows the vehicle rate corrections for an aircraft travelling East with the vertical axis rotated.

Aircraft Travelling East (Vertical Axis Rotated) Figure 105


Page 2-133

Page 2-134

P / O

P / O

P / O

E GYRO

V GYRO

N GYRO

T / M

T / M

T / M

(VEHICLE RATE CORRECTION)


(EARTH RATE CORRECTION)


(EARTH RATE CORRECTION)

E = (RATE OF CHANGE LAT)

V = (15º/Hr + RATE OF CHANGE LONG) SIN LAT

N = (15º/Hr + RATE OF CHANGE LONG) COS LAT



Figure 106 shows the Earth rate & Vehicle rate corrections.

Earth & Vehicle Rate Corrections Figure 106



2.50.5 Alignment The accuracy of an INS is dependent on the precise alignment of the inertial platform to a known reference (True North), with respect to the latitude and longitude of the ground starting position at the time of “Starting Up” the system. The inertial system computer carries out a self-alignment calibration procedure over a given period of time before the system is ready to navigate the aircraft. The computer requires the following information prior to alignment so that it can calculate the position of “True North”: 1. Aircraft’s Latitude Position. 2. Aircraft’s Longitude Position. 3. Aircraft’s Magnetic Heading (from Mag Heading System). The alignment procedure can only be carried out on the ground, during which the aircraft must not be moved. Once started the alignment procedure is automatic 2.50.6 The Navigation Mode In the navigation mode the pitch, roll attitude and the magnetic heading information is updated mainly with the attitude changes sensed by gyros. Because the IRS is aligned to true north a variation angle is used to calculate the direction to magnetic north. Each location on earth has its own variation angle. All variation angles between the 73 North and 60 South latitude are stored in the IRS. The present position is updated mainly with accelerations sensed by the accelerometers. The accelerations are corrected for the pitch and roll attitude and calculated with respect to the true north direction.


Page 2-135


2.50.7 Strapdown Inertial Navigation As already discussed, inertial navigation is the process of determining an aircraft’s location using internal inertial sensors. Unlike in the gimballed system, in a strapdown system the accelerometers and gyros are mounted solidly to the aircraft’s axis. There are no gimbals to keep the sensors level with the earth’s surface, so that one sensor is always on the aircraft’s longitudinal axis, one on the lateral axis and one on the vertical axis. Likewise the gyros are mounted such that one will detect the aircraft’s pitch, another the roll and the third the aircraft’s heading. The accelerometer produces an output that is proportional to the acceleration applied along the sensor’s input axis. A microprocessor integrates the acceleration signal to calculate a velocity and position. Although it is used to calculate velocity and position, acceleration is meaningless to the system without additional information. Example: Consider the acceleration signal from the accelerometer strapped to the aircraft’s longitudinal axis. It is measuring the forward acceleration of the aircraft, however, is the aircraft accelerating north, south, east, west, up or down? In order to navigate over the surface of the earth, the system must know how its acceleration is related to the earth’s surface. Because the accelerometers are mounted on the aircraft’s longitudinal, Lateral and vertical axes of the aircraft, the IRS must know the relationship of each of these axes to the surface of the earth. The Laser Ring Gyros (LRGs) in the strapdown system make measurements necessary to describe this relationship in terms of pitch, roll and heading angles. These angles are calculated from angular rates measured by the gyros through integration e.g. Gyro measures an angular rate of 3/sec for 30 seconds in the yaw axes. Through integration, the microprocessor calculates that the heading has changed by 90 after 30 seconds.

Given the knowledge of pitch, roll and heading that the gyros provide, the microprocessor resolves the acceleration signals into earth-related accelerations, and then performs the horizontal and vertical navigation calculations. Under normal conditions, all six sensors sense motion simultaneously and continuously, thereby entailing calculations that are substantially more complex than a normal INS. Therefore a powerful, high-speed microprocessor is required in the IRS in order to rapidly and accurately handle the additional complexity.

Page 2-136



LONGITUDE

VECTOR SOLVER LATITUDE PITCH ROLL YAW

GYROS

B MATRIX

COORDINATE CONVERTER

ACCELEROMETERS

ALTIMETER

POSITION COMPUTER

Figure 107 shows the block schematic of the Strap-Down inertial Navigation system.

Strap-Down Inertial Navigation System Figure 107


Page 2-137


2.50.8 Laser Ring Gyro (LRG) Operation Laser Ring Gyros (LRG) are not in fact gyros, but sensors of angular rate of rotation about a single axis. They are made of a triangular block of temperature stable glass. Very small tunnels are precisely drilled parallel to the perimeter of the triangle, and reflecting mirrors are placed in each corner. A small charge of Helium-neon gas is inserted and sealed into an aperture in the glass at the base of the triangle. When a high voltage is run between the anodes and the cathode, the gas is ionized, and two beams of light are generated, each travelling around the cavity in opposite directions. Since both contrarotating beams travel at the same speed (speed of light), it takes the exact same time to complete a circuit. However, if the gyro were rotated on its axis, the path length of one beam would be shortened, while the other would be lengthened. A laser beam adjusts its wavelength for the length of the path it travels, so the beam that travelled the shortest distance would rise in frequency, while the beam that travelled the longer distance would have a frequency decrease. The frequency difference between the two beams is directly proportional to the angular rate of turn about the gyro’s axis. Thus the frequency difference becomes a measure of rotation rate. If the gyro doesn’t move about its axis, both frequencies remain the same and the angular rate is zero.

Page 2-138



Figure 108 shows a Laser Ring Gyro.

FRINGE PATTERN

CORNER PRISM

Laser Ring Gyro (LRG) Figure 107


Page 2-139


2.50.9 Mode Select Unit (MSU) The mode select unit controls the mode of operation of the IRS. There are two types in common use: 1.

Six Annunciator MSU.

2.

Triple-Channel MSU.

The six-annunciator MSU provides mode selection, status indication and test initiation for one Inertial Reference Unit (IRU). Figure 109 shows six-annunciator MSU and Figure 110 shows a triple-channel MSU.

LASEREF

NAV ATT

ALIGN OFF

ALIGN

FAULT

NAV RDY

NO AIR

ON BATT

BATT FAIL TEST

IRS Six-Annunciator MSU Figure 109

Page 2-140



NAV

NAV

NAV ATT

ATT

ATT

ALIGN

ALIGN

ALIGN

OFF

OFF

OFF

SYS 1

SYS 2

SYS 3

ALIGN

ALIGN

ALIGN

ON BATT

ON BATT

ON BATT

BATT FAIL

BATT FAIL

BATT FAIL

FAULT

FAULT

FAULT

TEST

IRS Triple-Channel MSU Figure 110 2.50.10

Mode Select Unit Modes

IRS Modes or set by setting the MSU mode select switch as follows: OFF-TO-ALIGN The IRU enters the power-on/built-in test equipment (BITE) submode. When BITE is complete after approximately 13 seconds, the IRU enters the alignment mode. The IRU remains in the alignment mode until the mode select switch is set to OFF, NAV or ATT. The NAV RDY annunciator illuminates upon completion of the alignment. OFF-TO-NAV The IRU enters the power-on/built-in test equipment (BITE) submode. When BITE is complete after approximately 13 seconds, the IRU enters the alignment mode. Upon completion of the alignment mode the system enters the navigation mode.


Page 2-141


ALIGN-TO-NAV The IRU enters navigate mode from alignment mode upon completion of alignment. NAV-TO-ALIGN The IRU enters the align downmode from the navigate mode. NAV-TO-ALIGN-TO-NAV The IRU enters the align downmode and after 30 seconds, automatically reenters the navigate mode. ALIGN-TO-ATT or NAV-TO-ATT The IRU enters the erect attitude submode for 20 seconds, during which the MSU ALIGN annunciator illuminates. The IRU then enters the attitude mode. 2.50.11

MSU Annunciators

ALIGN – Indicates that the IRU is in the alignment mode. A flashing ALIGN annunciator indicates in-correct LAT/LONG entry, excessive aircraft movement during align. NAV RDY Indicates that the alignment is complete. FAULT Indicates an IRS fault. ON BATT Indicates that the back-up battery power is being used. BATT FAIL Indicates that the back-up battery power is inadequate to sustain IRS operation during back-up battery operation (less than 21 volts). NO AIR Indicates that cooling airflow is inadequate to cool the IRU.

Page 2-142



2.50.12

Inertial System Display Unit (ISDU)

The ISDU selects data from any one of three IRUs for display and provides initial position or heading data to the IRUs. Figure 111 shows an ISDU.

Honeywell

LASEREF

DSPL SEL P/POS TK/GS TEST

WIND HDG/STS

W 4 1

BRT

SYS DSPL 2 1

1

3

7 W 4 ENT

7

OFF

N 2 H N 5 2

S H 8 5 S0 8

3 E 36 E9 6 CLR

9

Inertial System Display Unit (ISDU) Figure 111


Page 2-143


2.50.13

Keyboard

The keyboard is used to enter latitude and longitude in the alignment mode or magnetic heading in the attitude mode. The ISDU then sends the entered data simultaneously to all IRUs when ENT pressed. The keyboard contains 12 keys, five of the 12 keys are dual function: N/2, W/4, H/5,E/6 AND S/8. A dual function key is used to select either the type of data (latitude, longitude or heading) or numerical data to be entered. Single function keys are used to select only numerical data. The CLR (clear) and ENT (enter) keys contain green cue lights which, when lit indicate that the operator action is required. CLR is used to remove data erroneously entered onto the display; ENT is used to send data to the IRU. 2.50.14

Display

The 13-digit alphanumeric spilt display shows two types of navigation data at the same time. The display is separated into one group of 6 digits (position 1 through 6) and one group of 7 digits (positions 7 through 13). Punctuation marks (located in positions 3,5,6,10,12,and 13) light when necessary to indicate degrees, decimal points, and minutes. 2.50.15

System Display Switch (SYS DSPL)

The SYS DSPL switch is used to select the IRU (position 1,2 or 3) from which the displayed data originates. If the switch is set to OFF, the ISDU cannot send or receive data from any of the 3 IRUs. 2.50.16

Display Selector Switch (DSPL SEL)

The DSPL SEL switch has five positions to select data displayed on the ISDU. TEST – Selects a display test that illuminates all display elements and keyboard cue lights to allow inspection for possible malfunctions. The DSPL SEL switch is spring loaded and must be help in this position. TK/GS – Selects track angle in degrees on the left display and ground speed in knots on the right. PPOS – Selects the aircraft’s present position as latitude on the left display and longitude on the right. Both latitude and longitude are displayed in degrees, minutes, and tenths of a minute. WIND – Selects wind direction in degrees on the left display and wind speed in knots on the right display.

Page 2-144



HDG/STS – Selects heading or alignment status for display, depending upon the current IRU mode. Heading is displayed in degrees and tenths of degrees, and time-to-alignment completion is displayed in minutes and tenths of minutes. In the alignment mode, the ISDU displays alignment status (time to NAV ready) in the right display. In the NAV mode, the ISDU displays true heading in the left display. In the attitude mode, the ISDU displays magnetic heading in the left display and ATT in the right display. 2.50.17

Dimmer Knob

The dimmer knob is mounted on, on operates independently of, the DSPL SEL switch. As the dimmer knob is rotated clockwise, the display brightens. 2.50.18

Inertial Reference Unit (IRU)

The IRU is the main electronic assembly of the IRS. The IRU contains an inertial sensor assembly, microprocessors, and power supplies and aircraft electronic interface. Accelerometers and LRG in the inertial sensor assembly measure acceleration and angular rates of the aircraft. The IRU microprocessors performs computations required for: 1.Primary Attitude. 2.Present Position. 3.Inertial Velocity Vectors. 4.Magnetic and True North Reference. 5.Sensor Error Compensation. The power supplies receive a.c. and d.c. power from the aircraft and back-up battery. It supplies power to the IRS, and provides switching to primary a.c. and d.c. or backup battery power The aircraft electronic interface converts ARINC inputs for use by the IRS. The electronic interface also provides IRS outputs in ARINC formats for use by associated aircraft equipment. A fault ball indicator and a manual “Interface Test” switch are mounted on the front of the IRU and are visible when the IRU is mounted in an avionics rack.


Page 2-145


Figure 112 shows an IRU

Inertial Reference Unit

INTERFACE TEST

Inertial Reference Unit Figure 112

Page 2-146



2.50.19

IRS Alignment Mode

During alignment the inertial reference system determines the local vertical and the direction of true north. 2.50.20

Gyro Compass Process

Inside the inertial reference unit, the three gyros sense angular rate of the aircraft. Since the aircraft is stationary during alignment, the angular rate is due to earth rotation. The IRU computer uses this angular rate to determine the direction of true north. 2.50.21

Initial Latitude

During the alignment period, the IRU computer has determined true north by sensing the direction of the earth’s rotation. The magnitude of the earth’ rotation vector allows the IRU computer to estimate latitude of the initial present position. This calculated latitude is compared with the latitude entered by the operator during initialization. 2.50.22

Alignment Mode

For the IRU to enter ALIGN mode, the mode select switch is set to either the ALIGN or NAV position. The systems software performs a vertical levelling and determines aircraft true heading and latitude. The levelling operations bring the pitch and roll attitudes to within 1 accuracy (course levelling), followed by fine levelling and heading determination. Initial latitude and longitude data must be entered manually either via the IRS CDU or the Flight Management System CDU. Upon ALIGN completion, the IRS will enter NAV mode automatically if the mode select switch was set to NAV during align. If the mode select switch was set to ALIGN, the system will remain in align until NAV mode is selected. The alignment time is approximately 10 minutes.


Page 2-147

Page 2-148

S Y S T E M S

A I R C R A F T

IRU 1


IRU 3


IRU 2


INTERFACE TEST

INTERFACE TEST

INTERFACE TEST

7

ENT

S 0 8

H 8 5

S

2

H N 5

N 2

FAULT

FAULT

ATT

ON BATT

ALIGN

SYS 3

NAV

9

CLR

E9 6

E 6 3

3

FAULT

BATT FAIL

OFF

ALIGN

MODE SELECT UNIT

ON BATT BATT FAIL

ATT

ON BATT

SYS 2

NAV

ALIGN

OFF

ALIGN

SYS 1

ATT

ALIGN

NAV

BATT FAIL

OFF

ALIGN

3

4

1

7 W

W 4 1

1

2

HDG/STS

WIND

SYS DSPL

BRT

P/POS

LASEREF

INERTIAL SYSTEM DISPLAY UNIT

OFF

TEST

TK/GS

DSPL SEL

Honeywell

TEST



Figure 113 shows a block schematic of a three IRU inertial system.

IRS Block Schematic Figure 113



THRUST MANAGEMENT COMPUTER

YAW DAMPER

FLIGHT DATA ACQN UNIT

AIR DATA COMPUTER

EHSI/EADI VSI RDMI

ANTI-SKID AUTOBRAKE SYSTEM

WEATHER RADAR

IR MODE PANEL

INERTIAL REFERENCE UNIT

FLIGHT MANAGEMENT COMPUTER

GROUND PROXIMITY WARNING

FLIGHT CONTROL COMPUTERS

Figure 114 shows a block schematic of the interface of the IRS with the aircraft’s avionics systems.

IRS Interface – Block Schematic Figure 114 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-149


2.51 RADIO MAGNETIC INDICATOR (RMI) The radio magnetic indicator is a very useful navigation tool due to its ability to display several different pieces of information simultaneously. Primarily, the circular rotating 'card' is a self- correcting compass which is much more accurate that the older, floating magnet type of compass. Secondly, the displays from the ADF or VOR units can be displayed on top of the card, using two pointers, one single and one double. This allows the pilot to see, in one instrument, his heading and the orientation of up to two ground stations, relative to the aircraft, using two different navigation systems. Figure 115 shows an RMI display and aircraft position with respect to an ADF and VOR station.

AIRCRAFT HEADING MAGNETIC NORTH

BEARING TO VOR 2 BEACON

BEARING TO ADF 1 BEACON

N

3

30

33

24

12

S

21 VOR

A D F

15

F

E

W

6

A D

VOR

RMI Display Figure 115

Page 2-150



2.51.1 Dual Distance Radio Magnetic Indicator (DDRMI) The Dual-Distance Radio-Magnetic Indicator (DDRMI) is an instrument that gives indications for various navigation systems: 1.

Magnetic heading from Compass system.

2.

Bearings from VORs or ADFs.

3.

Distances from the DMEs.

2.51.2 DDRMI Principle Figure 116 shows the principle operation of the DDRMI system.

NM

VOR NO 2

DM

ED

IST

AN CE =

75 .5

BEARING TO VOR NO 2 = 30º

DME DIS

65.5 NM TANCE =

BEARING TO VOR NO 1 = 87º

VOR NO 1

DDRMI Operation Figure 116


Page 2-151


Figure 111 shows a DDRMI indications resulting from the situation in figure 117.

65.5

75.5

DME - 1

DME - 2

N

3

24

12

S

21 ADF

15

V O R

E

W

6

30

33

V O R

ADF

DDRMI Indication Figure 117

Page 2-152



CAPT’S DDRMI

DME SYSTEM NO 1

COMPASS SYSTEM NO2

ADF SYSTEM NO 1

VOR SYSTEM NO 1

ADF SYSTEM NO 2

VOR SYSTEM NO 2

COMPASS SYSTEM NO 1

DME SYSTEM NO 2

F/O’S DDRMI

Figure 118 shows a block schematic of the DDRMI system and the source of all displayed data.

DDRMI Schematic Figure 118


Page 2-153


2.52 GLOBAL POSITIONING SYSTEM (GPS) GPS is a space based radio navigation system, which provides worldwide, highly accurate three-dimensional position, velocity and time information. The overall system is divided into three parts. 1.Space Segment. 2.Control Segment. 3.User Segment. 2.52.1 Space Segment Consists of 24 satellites (21 active + 3 spare), in six orbital planes with 4 satellites in each orbit. They are orbiting the earth every 12 hours at an approximate altitude of between 11,000nm – 12,500nm. The orbits are such that a minimum of 6 satellites are in view from any point on the earth. This provides redundancy, as only 4 satellites are required for three-dimensional position. Figure 119 shows the Space Segment.

GPS Space Segment Figure 119 Page 2-154



2.52.2 Control Segment This is a ground station that controls all satellites and is made up of: 1.

Master Control Station.

2.

Monitor Stations.

The Master Control Station is located at Colorado, USA, and is responsible for processing satellite-tracking information received from the five Monitor Stations. The Control Segments monitor the total system performance, corrects satellite position and re-calibrates the on-board atomic time standards as necessary. The Monitor Stations are located to provide continuous "ground" visibility of every satellite. Three of the five monitor stations have ground antennas, which are used to upload data to the satellites. Figure 120 shows the location of the Control Segment.

COLORADO SPRINGS

HAWAII

KWAJALIEN

ASCENSION DIEGO GARCIA

MASTER CONTROL

MONITOR STATION

GROUND ANTENNA

GPS Control Segment Figure 120


Page 2-155


2.52.3 Operation GPS operates by measuring the time it takes a signal to travel from a satellite to a receiver on-board the aircraft. This time is multiplied by the speed of light to obtain the distance measurement. This distance results in a Line Of Position (LOP). Figure 121 shows GPS LOP.

LINE OF POSITION (LOP)

GPS Line of Sight (LOP) Figure 121 The satellites transmit a signal pattern, which is computer generated, in a repeatable random code. The receiver on the aircraft also generates the same code and the first step in the process of using GPS data is to synchronies these two codes. The receiver will receive the LOPs from three different satellites and uses this information to establish synchronization. The receiver is programmed to receive signals that intersect the same point, if they don’t, then the two codes are not synchronized. The receiver will now add or subtract time from its code to establish the LOPs intersecting the same point and thus synchronize its code with the one from the satellite.

Page 2-156



Figure 122 shows the principle of code synchronisation.

R EC EIV ER C OD E N OT SY NC H R ON ISE D W IT H T HE SA T ELL IT E CO DE W ILL GIV E T W O/T H R EE PO SSI BLE POS IT ION S

R EC EIV ER AD D S/ SU BT R AC T S T IM E F R OM I TS C OD E T O ES T ABL ISH T H E LOP S IN T ER SE CT I NG T H E SA ME PO IN T

Code Synchronisation Figure 122


Page 2-157


2.52.4 Signal Structure GPS satellites transmit on 2 frequencies in 2 modes in the UHF band. The 2 modes are: 1.

Precision Mode (P).

2.

Coarse/Acquisition Mode (C/A).

The P code is for military use only. Both codes transmit signals in a "Pseudo Random Code" at a certain rate. 2.52.5 Time Measurements Once the GPS receiver has synchronized with the satellite code, it can then measure the elapsed time since transmission by comparing the phase shift between the two codes. The larger the phase shift, the longer the length of time since transmission. The length of time since transmission times the speed of light equals distance.

Page 2-158



SIGNAL RECEIVED FROM SATELLITE

TIME DELAY = RANGE

SIGNAL TRANSMITTED FROM SATELLITE

Figure 123 shows code synchronization and time measurements.

Code synchronization and Time measurement Figure 123


Page 2-159


2.52.6 Position Fixing If we know our distance from a specific point in space (satellite), then it follows that we are located somewhere on the surface of a sphere, with its radius of that distance. The addition of a second satellite and a second distance measurement further refines the position calculation as the two LOPs intersect each other. The addition of a third distance measurement from a third satellite further refines the position calculation as we now have three LOPs intersecting at a specific point in space. This point in space represents the distance measured between the aircraft and the three satellites. Figure 124 shows the process of position fixing.

AIRCRAFT’S VERTICAL POSITION

AIRCRAFT’S HORIZONTAL POSITION

GPS Position Fixing Figure 124

Page 2-160



2.52.7 Ionospheric Propagation Error The ionosphere refracts UHF satellite transmission in the same way it refracts VLF, L.MF and HF transmissions, only to a lesser degree. Since a refracted signal has a greater distance to travel than a straight signal, it will arrive later in time, causing an error in the distance measurement. The ionosphere refracts signals in an amount inversely proportional to the square of their frequencies. This means that the higher the frequency, the less the refraction and hence the less error induced in the distance measurement. Since the GPS satellites transmit two different UHF frequencies (1575.42 MHz and 1227.60 MHz), each frequency will be affected by the ionosphere differently. By comparing the phase shift between the two frequencies, the amount of ionosphere distortion can be measured directly. By knowing the amount of distortion that is induced, the exact correction factor can be entered into the computer and effectively cancel ionosphere propagation error. Figure 125 shows the principle of Ionospheric Propagation Errors.

Ionospheric Propagation Errors Figure 125


Page 2-161


2.52.8 Derived Information Although the GPS is primarily a position determining system, it is possible to derive certain data by taking into account the change in position over time. Actual track can be obtained by looking at several position fixes. Ground speed can be calculated by measuring the distance between two fixes. Drift angle can be obtained by comparing the aircraft’s heading, with the actual track of the aircraft. GPS is able to produce all the derived data commonly associated with existing long-range navigation systems such as INS. 2.52.9 Navigation Management A typical GPS provides Great Circle navigation from its present Position direct to any waypoint or via a prescribed flight plan. When necessary, a new route can be quickly programmed in flight. Up to 999 waypoints and up to 56 flight plans are retained by the GNS-X when power is turned off or interrupted. Selection of waypoints or of the leg to be flown is not necessary to determine aircraft position; however, when these are provided, the GNS-X computes and displays on the Colour Control Display Unit all pertinent navigation data including: Greenwich Date and Mean Time. Present Position Coordinates. Magnetic Variation. Stored Waypoint Coordinates. Stored Flight Plans. Departure Time/Time at last Waypoint. Bearing to Waypoint. Distance to Waypoint. Estimated Time to Waypoint (ETE).

Estimated Time of Arrival (ETA). Wind Direction and Speed. Desired Track. Drift Angle. Ground Speed. Track Angle. Crosstrack Distance. HSI/CDI/RMI Course Display.

The computer determines the composite position based on sensor position/velocity. Plotting multiple moving position points allows determination of Track Angle and the rate of change of position equals groundspeed. Drift Angle becomes available with the Heading input, and with a True Airspeed (TAS) input allows calculation of the Wind direction and speed. The computer is constantly processing all available inputs. The displays of Present Position, Distance-to-Go, and Crosstrack as well as the displays of Track Angle, Drift Angle, Groundspeed, Wind, and Estimated Time Enroute are updated at periodic intervals.

Page 2-162



Figure 126 shows the block schematic of a GNS-X system.

GP S A N TE N NA

E FIS

N AV I G A TIO N PR OC ESSOR U NI T

A DC

A UT O P IL O T

C OM PA SS

A UT O P IL O T M O D E S E L E CT

M U LT IF U N C T I O N C O N TR O L D IS P L A Y U N IT

R TE

D EP ARR

LE GS

H OLD

PR OG

VN A V

A TC

T IT L E F IEL D

L EF T F IE LD

R IGH T F IE LD

SC R A T CH PA D

BRT

/

C LR

D IM

PR E V

N EX T

M EN U

D A TA

EX EC

1

2

3

A

B

C

D

E

F

G

4

5

6

H

I

J

K

L

M

N

7

8

9

O

P

Q

R

S

T

U

V

W

X

Y

Z

SP

0

+

/

GNS-X CONTROL & DISPLY UNIT GNS-X System Figure 126


Page 2-163


2.52.10

Boeing 777 GPS

The Boeing 777 has two independent GPS, which are used to calculate the following: 1. Aircraft’s Latitude. 2. Aircraft’s Longitude. 3. Aircraft’s Altitude. 4. Aircraft’s Groundspeed. 5. Accurate Time. Figure 127 shows the system layout.

LEFT GPS SENSOR UNIT

RIGHT GPS SENSOR UNIT

GPWC

RIGHT GPS ANTENNA

CHR

DATE

60

AIR DATA INERTIAL REFERENCE UNIT ADIRU X 3

DAY. MON . YR

50

23 : 59

10

GMT

DIGITAL CLOCK X2

ET/CHR

45

20

30

RUN HLD

ET

G

RUN HLD

99 : 59

RESET

FS D

629 DATA BUS X 3 AIMS CABINET X 2

Boeing 777 GPS Figure 127 Page 2-164

SS MT M

LEFT GPS ANTENNA



The two sensor units receive GPS satellite signals from their respective antennas and calculate the aircraft’s position and accurate time. This data is sent to the Aircraft Information Management System (AIMS) cabinets and the Ground Proximity Warning Computer (GPWC). The Flight Management Computing system uses the AIMS GPS data to calculate the aircraft’s position for use in its navigation calculations. The AIMS cabinets also send GPS data to the Air Data Inertial Reference Units (ADIRU) which is used to calibrate the inertial sensors, thus decreasing any inertial reference drift. GPS time goes to the Universal Time Co-ordinated function (UTCF) within the AIMS, the AIMS also outputs time data to the flight deck clocks. 2.52.11

GPS Modes of Operation

The Boeing 777 GPS operates in the following modes:

2.52.12

1.

Acquisition Mode.

2.

Navigation Mode.

3.

Altitude Aided Mode.

4.

Aided Mode.

Acquisition Mode

The GPS sensor units look for and lock onto the satellite signals. The sensors must find at least 4 satellites before it can start to calculate GPS data. Whilst the sensor is in the acquisition mode, itreceives the following data from the Flight Management system: 1.

Aircraft’s Present Position.

2.

Aircraft’s Velocity.

3.

Time & Date.

The GPS sensor unit uses this data to calculate which satellites are available at the current aircraft’s position, allowing the sensor unit to receive the signals from those satellites available and which ones may be used for navigation calculations.


Page 2-165


2.52.13

Navigation Mode

Once the GPS sensor has acquired and locked onto at least 4 satellites it will enter the navigation mode. In this mode the sensor unit it will compute the GPAS data. If during the Navigation mode the GPS accuracy is not within 16NM of the actual aircraft’s position, the sensor output will go into “None Computed Data” (NCD). 2.52.14

Altitude Aided Mode

With 4 satellites available, the GPS sensor stores the difference between the Air Data Inertial Reference Unit (ADIRU) altitude and the GPS altitude. When the GPS sensor is only receiving signals from 3 satellites, it will use this stored data so that it can estimate the GPS altitude. During this phase the GPS sensor will use the aircraft’s altitude from the ADIRU and the length of the earth’s radius as the fourth range required for GPS altitude calculations. Figure 128 shows the Altitude Aided Mode.

Altitude Aided Mode Figure 128

Page 2-166



2.52.15

Aided Mode

The GPS sensor enters the “Aided Mode” during short periods (Less than 30 seconds) of bad satellite coverage. An example of bad satellite coverage is poor satellite geometry, where at least 4 satellites are available but they are not spread out far enough so the GPS sensor unit can make an accurate position fix. In the aided mode, the GPS sensor unit receives altitude, heading and groundspeed from the Flight Management System (FMS). The GPS sensor unit uses this data to go back into Navigation mode when there is good satellite coverage again. During the Aided Mode the GPS sensor unit output is once again “Non Computed Data (NCD)


Page 2-167


C

ARE 4 SATS AVAIL?

NO

BAD SAT COVERAGE?

NO ALTITUDE AIDED MODE

30 SEC PASSED?

ARE THERE ONLY 3 SATS AVAIL?

NO

BAD SAT COVERAGE?

NAVIGATION MODE

YES

B

NO

NO

C

ARE 4 SATS AVAIL?

ACQUISITION MODE

POWER-UP

A

YES

YES

NO

NO

AIDED MODE

ARE 4 SATS AVAIL?

YES

YES

B

A

YES

Figure 129 shows the Boeing 777 GPS modes of operation.

GPS Modes of Operation Figure 129 Page 2-168



2.52.16

Receiver Autonomous Integrity (RAIM)

The purpose of the RAIM is to monitor the status of the satellites that the GPS sensor unit is using for its navigation calculations. The output of the RAIM function is an estimate of the GPS position error. The RAIM value goes to the Flight Management System (FMS) and is used by the FMS to determine if the GPS data can be used for navigation. Figure 130 shows the operation of RAIM.

1 5

2

SATELLITE CURRENTLY MONITORED

4

3

RAIM Operation Figure 130


Page 2-169


2.52.17

Differential GPS

The accuracy of the GPS is typically 15 – 25 metres in 95% of the position fixes available. The USA Department of Defence degrades this accuracy for security reasons to 100 metres in 95% of the position fixes. However, this error can be further reduced to almost zero by the use of “Differential GPS”. If GPS receivers are placed on the ground in known locations (Latitude Longitude), the exact errors of the GPS satellites can then be calculated by comparing the known position of the receivers against the GPS satellites calculated position. This error is then transmitted to other receivers who use it to correct the GPS errors and thus have a more accurate position fix. Figure 131 shows the operation of differential GPS.

ERROR CALCULATION

ERROR TRANSMISSION

Differential GPS Figure 131

Page 2-170



2.53 COMPASS SYSTEMS 2.53.1 Direct Reading Compass This type of compass comprises a magnet system in a liquid filled bowl. In this type the compass card is attached to single angular cobalt steel magnet which is suspended in a sapphire cup by an iridium tipped pivot. Figure 132 shows a common type of direct reading compass. MOUNTING PLATE

HORIZONTAL (“B” & “C”) CORRECTORS

FILLER PLUG

BELLOWS

BOWL

MAGNET SYSTEM

STEM & BRACKET ASSEMBLY

Direct Reading Compass Figure 132

Damping is achieved by filling the compass bowl with a mineral liquid or alcohol, which has a low viscosity, low freezing point, high resistance to corrosion and does not discolour. The compass is also given buoyancy by the liquid and this reduces wear on the pivots. The compass liquid expands and retracts with changes in temperature and this has undesirable effects. To compensate for this, a bellows or corrugated diaphragm is fitted. Note:

B and C correctors are for East-West, North-South errors respectively.


Page 2-171


On some modern aircraft the direct reading compass is stowed on the center windscreen strut, only being used in an emergency. They also have to have some sort of lighting; this lighting is operated by dc and does not effect the compass operation. Figure 134 shows two types of compass fitted to modern aircraft.

FIXED COMPASS

HINGED COMPASS

Direct reading Compass Figure 134

Page 2-172



2.53.2 Remote Reading Compass (Magnet Gyro) The (magnetic gyro) compass system provides the flight crew with magnetic heading information. A compass card in the radio magnetic indicators (RMI's) on the instrument panel displays the heading, which must be read against a reference point or a lubber line. The compass heading is controlled by a directional gyro, which has a stable direction. For proper orientation of the system with the earth’s magnetic field and to correct for gyro drift, a flux valve is used. The flux valve senses the direction of the earth magnetic field. Figure 135 shows the layout of a basic system.

115v 400 Hz

B

+

_

6 E

12

21 VOR

S 15

A D F

_ +

3

W 3 0 24

N 33

C

A D F

VOR

SLAVED

DG

VOR/ADF

SYNC

Remote Reading Compass System Figure 135


Page 2-173


2.53.3 Flux valve (Detector Unit) A flux valve, or detector unit, senses the angle of the horizontal component of the Earth's magnetic field with respect to the aircraft's heading, and gives a long-term stable signal to monitor the gyro controlled master shaft. The detector unit can best be described as a North sensing device which is capable of detecting the direction of the horizontal component of the Earth's field and transmitting it to other components. It is similar to a CX in a synchro control system. Figure 136 shows a Detector Unit and internal circuit.

SIDE VIEW

LAMINATED COLLECTOR HORNS

A

A

AC POWER

B

EXCITER COIL

C TOP VIEW

B C

SECONDARY PICK-OFF COILS

DETECTOR UNIT CIRCUIT

Flux Valve Construction Figure 136

Page 2-174



2.53.4 Control Panel The control panel consists of the following: 1.

Synchronisation Annunciator.

2.

Synchronisation Knob.

3.

Slaved/DG Switch.

2.53.5 Synchronisation Annunciator This indicates the synchronisation between the DU heading and the gyro heading. If there is a discrepancy between the two headings then the indicator will show either a “DOT” or a “CROSS”. 2.53.6 Synchronisation Knob This allows for manual synchronisation of the DU/gyro headings. The Synchronisation Knob has two directions (DOT & CROSS), moving the synchronisation knob in the direction indicated by the synchronisation indicator will ensure the system will indicate the correct heading. 2.53.7 Slaved/DG Switch The compass systems normal operation mode is the slaved mode, where the DU/gyro headings are slaved together (DU will precess the gyro when an error occurs between the two detected headings. In the DG mode, the DU is removed from the system and the compass operates as a “Directional Gyro”. This mode is used more in maintenance when aircraft heading is required. If used in flight there is a possibility that the heading indication will drift due to gyro drift.


Page 2-175


Figure 137 shows the Compass Control Panel.

SYNCHRONISATION ANNUNCIATOR

SLAVED/DG SWITCH

COMPASS DG HDG SLEW

COMPASS SLAVE SLAVED

SYNCHRONISATION SWITCH

Compass Control Panel Figure 137 The synchro transmitters in the RMI and in the directional gyro unit are used to "transmit" the heading information to the following systems:

Page 2-176

1.

Autopilot.

2.

Flight director system (if installed).

3.

Horizontal situation indicators (HSI's).

4.

Flight data recorder system (if installed).

5.

VHF NAV receivers.



2.53.8 System Test Three items can be tested in the compass system: The sensitivity of the cross-dot annunciator. When the card is moved 5° away from the synchronised position with the manual synchronisation knob the cross or the dot must be completely visible. The slaving speed when the card is moved 10° away from the synchronised position the automatic slaving system should move the card to the synchronised position within 10 minutes (min. slaving speed 1°/min). The directional gyro drift with the slaving cut-out switch in DG the gyro drift should not exceed (3.75 x sine attitude + 1.75° per 15 minutes). 2.53.9 Gyro Unit The basic element of this compass system is a directional gyro. When the system is supplied with 115-V AC the gyro starts to rotate and becomes a stable element, which means that its direction (heading) in space is fixed.


Page 2-177


2.53.10

Servo System

A servo loop between the gyro and the compass card in the RMI ensures that any change of aircraft heading causes a corresponding rotation of the compass card, but in the opposite direction. The servo loop comprises a synchro transmitter (Tx), a control transformer (CT), a servo amplifier and a servomotor. The rotor of the synchro transmitter points in the same direction as the gyro. The error signal is applied via the servo amplifier to the motor. The motor in its turn drives the compass card and the rotor of the control transformer. When the latter rotates, the error signal reduces to zero and the motor stops rotating. The rotor is powered with a 400-Hz signal, which causes a 400-Hz magnetic field. This magnetic field produces 3 voltages in the stator windings of the synchro transmitter. The 3 voltages in the control transformer cause a resulting magnetic field. The rotor of the control transformer produces an error signal any time the rotor of the control transformer is not perpendicular with the direction of the resulting magnetic field. The error signal is applied via the servo amplifier to the motor. The motor in its turn drives the compass card and the rotor of the control transformer. When the latter rotates, the error signal reduces to zero and the motor stops rotating. If the aircraft changes heading, the direction of the 400-Hz magnetic field in the synchro transmitter changes with respect of the stator windings and therefore the direction of the resulting magnetic field in the control transformer changes too. An error signal is now present and after amplification the "heading" of the compass card and the rotor of the control transformer changes accordingly and the compass card reads the new aircraft heading.

Page 2-178



COMPASS CARD

TORQUE MOTOR

TM ANNUNCIATOR

SYNCH KNOB

DU EXCITATION (AC) DETECTOR UNIT (FLUX VALVE)

CT

SLAVING AMP

HEADING SHAFT

SERVO AMP

M

CT

VELOCITY FEEDBACK

TG

TACHO GENERATOR

GYRO CASE

OUTER RING

GYRO

CX

INNER RING

26V AC 400 Hz

Figure 138 shows the Compass System Schematic.

Compass System Schematic Figure 138 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-179


2.53.11

Slaving loop

To obtain a common reference for all aircraft, use is made of the earth magnetic field the direction of which is detected by a flux valve. A second control transformer in the RMI compares the compass card reading with the direction of the earth magnetic field. Any difference between these two causes an error signal at the output of the rotor of the control transformer. The error signal is amplified in a slaving amplifier and this signal drives a torque motor in the directional gyro unit. The torque motor changes the position of the stable element and of the rotor of the synchro transmitter. As described for the servo loop, the compass card and rotors of both control transformers rotate accordingly until the error signals have been reduced to zero. The amplified error signal at the output of the slaving amplifier also drives a cross-dot annunciator. Either a cross or a dot indicates any unsynchronised condition of the compass system. The cross-dot annunciator can be used to manually synchronise the compass system by turning the manual synchronising knob on the control panel in the cross or dot direction until the cross or dot has disappeared.

Page 2-180



2.54 RADIO ALTIMETER Radio altimeters are carried in virtually all aircraft outside the general aviation sector. The outputs from the system play a vital role in the operation of automatic landing and ground proximity warning systems. Because the radio altimeter comes into play at a critical part of the flight, when the aircraft is close to the ground, the serviceability and accuracy are perhaps more important than with any other radio system. Barometric altitude is the altitude of the aircraft as a function of change in air pressure. Since it involves a measure of the change in pressure it is the altitude of the aircraft above the level at which a certain air pressure exists. For aircraft flying above about 3000 feet the usual reference pressure is 1013.25 millibars (mb) or 29.92 inches of mercury (in Hg). This is known as the mean sea level (msl) pressure. The actual pressure at sea level is unlikely to be exactly 1013.25 mb; hence the altimeter will not be reading the aircraft's height above sea level, let alone the ground. Radio altitude, on the other hand, is always the height above the ground regardless of air pressure or indeed the terrain the aircraft is flying over. It follows that radio altitude is more useful at low levels, in particular when in the landing phase or to give ground collision warning. 2.54.1 Basic Principles Radio altimeters are primary radar systems that transmit RF energy and time how long it takes before an echo is received. The radio altimeter target is always the ground immediately below the aircraft. The transmitted beam is broadly directional, pointing straight down, so for moderate bank and pitch angles part of the beam will be vertical. Figure 139 illustrates the idea showing dual aerial working.


Page 2-181


MODULATOR

BEAT FREQUENCY COUNTER

INDICATOR

TRANSMITTER

RECEIVER/MIXER

Radio Altimeter Operation Figure 139 The system transmits a continuous wave, constant amplitude; frequency modulated carrier at 4,300 MHz. The depth of modulation is 50 MHz, so the transmission is continuously varied between 4,250 MHz and 4,350 MHz. With an aircraft flying over the ground, there is a difference in the frequency of the reflected signals seen by the receiver and the transmitted frequency at the same instant. This difference is due to the distance the radio wave has had to travel from the transmitter antenna to the ground and back to the receiver antenna. For each foot of transmitted distance, there is a frequency change of approximately 10 cycles. Since the transmission must travel to the ground and back again to the receiver, the frequency change per foot of aircraft altitude is approximately 20 cycles. For example, if the aircraft is 1,000ft above the ground, there will be 20,000 cycles of frequency difference between the transmitter frequency and the received frequency.

Page 2-182



In the receiver mixer, the transmitted and received frequencies are mixed and the beat frequency (difference) is counted in the counter. The beat counter converts the frequency difference to an analog dc voltage whose amplitude is a function of aircraft altitude above the ground. A servo system in the indicator drives the indication to a position corresponding to the amplitude of the dc voltage received from the beat counter. Figure 140 shows two types of indicator used.

0

DIAL INDICATOR RIBBON INDICATOR

Radio Altimeters Figure 140 On the ribbon type indicator, the aircraft reference symbol remains fixed in the centre while the tape is driven behind it. Different tape colours are used to give an instant indication of the approximate height. The flag, when activated, partially obscures the aircraft symbol. A manually set altitude trip (decision height) is provided. By means of a DH Index control a marker or bug can be set to any desired height. If the aircraft is flying below the DH bug setting the DH lamp will be illuminated to give a warning.


Page 2-183


2.54.2 Radio Altimeter Antenna The antennas are so designed so that as long as the roll angle does not exceed 30, and the pitch attitude is not more than 20, the altitude indication remains correct. If these limits are exceeded, then the altitude indications would be excessive. These high values would not be maintained very long, so do not present a problem. Figure 141 shows the effects of aircraft roll on the operation of the Radio Altimeter system.

FAN BEAM TRANSMISSION

30° 30° SHORTEST RETURN PATH

30°

SHORTEST RETURN PATH

Roll Angle Effect on Radio Altimeter System Figure 141

Page 2-184



Radio Altimeter systems are called “Low Range” because they are not intended to operate at aircraft altitudes above the ground greater than 2,500ft. It is used mostly during final approach. When making a CAT II approach, the radio altimeter notifies the crew when the aircraft is 100 feet above the extended runway. This is the point at which the flight crew must be able to see the runway to land and is called the “Decision Height”. The decision height may be selected above 100ft as required. Figure 142 shows a Radio Altimeter antenna and its location.

RX ANTENNAS

TX ANTENNAS

HORN ANTENNA

Radio Altimeter Antenna & Location Figure 142


Page 2-185


2.54.3 Testing The Radio Altimeter system may be tested from the transceiver or other areas depending on aircraft type. When the test switch on the transmitter/receiver is operated, the integral test lights are tested. A test altitude of 40 ft is given and lights displayed, 'SYS OK' for serviceable, 'RT' or 'ANT' for a fault, give test results. Figure 143 shows a radio Altimeter Transceiver.

TEST

SYSTEM OK R/T UNIT ANT IND

Radio Altimeter Transceiver Figure 143 Page 2-186



2.55 WEATHER RADAR Weather Radar is designed to detect turbulent conditions so as to allow the pilot to avoid areas, which could cause an uncomfortable flight for the passengers and even the possibility of structural damage to the aircraft. At present there is no direct method of detecting turbulence. The Weather Radar, therefore, relies on detecting the conditions associated with turbulence. A vast amount of water exists in the air in one of three forms: vapour, liquid or solid, the form taken depending, most importantly, on the temperature, and the number of microscopic particles in the air. Calm conditions mean that the water droplets in the air are very small and float gently around, their weight being balanced by air resistance so that they do not fall to the ground. In turbulent conditions the water droplets or ice particles are thrown around, collide and stick together. Eventually they become large and heavy enough to fall to earth. The more violent the turbulence the larger the droplets will become before falling, particularly so where there is an updraft of air. If a water droplet is large enough it will scatter incident electro-magnetic waves, with some of the scattered energy being in the direction of the transmitter-receiver. Primary radar, therefore, can be used to detect water droplets and ice particles. The smaller the wavelength of the incident waves the smaller the water droplets that will scatter energy. This also applies to ice particles but the situation is complicated by the form of the ice i.e.: snowflakes, hailstones or sleet. The larger the droplet, the more energy is scattered and a cloud mass with large droplets will give rise to strong signals. Strong signals from a cloud, therefore, suggest turbulent conditions. Sometimes strong signals are received from one region of a cloud and small signals from an adjacent region. In this case a high rainfall gradient exists with strong clearly defined updrafts in the region of the strong signals.


Page 2-187


TRANSMITTED ENERGY

SELECTED RANGE

SCAN ANGLE

RFLECTED ENERGY

Figure 144 shows the operation of weather radar.

Weather Radar Operation Figure 144 Page 2-188



The part of a cloud, which gives the strong radar returns, is known as the storm cell. The closer the storm cell is to the edge of the detected cloud, the higher the rainfall gradient and the worse the conditions are likely to be. The function of Weather Radar is to detect and display conditions involving storm cells and rainfall gradients in such a way as to allow the operator to assess the probability of turbulence associated with such conditions. Figure 145 shows a typical weather radar scanner.

AZIMUTH GEAR

REFLECTOR

ELEVATION GEAR AERIAL

GIMBAL WAVEGUIDE

Weather Radar Scanner Figure 145


Page 2-189


2.55.1 Principle Of Operation The radar antenna is installed in the nose of the aircraft behind a radome. The radar transmitter/receiver transmits high-energy pulses via the antenna to the area in front of the aircraft. The weather radar system moves the antenna from the left to the right and back again so pulses are transmitted in a wide area in front of the aircraft. The weather circumstances in front of the aircraft (rain density in the clouds) reflect the transmitted pulses back to the weather radar system. The weather radar antenna receives the reflected pulses. The weather radar receiver converts the received pulses into a picture, which represents the weather circumstances. To produce this picture the weather radar system makes use of: The strength of the reflected pulses which depends on the amount of rain in the clouds. (The weather radar system converts the strength of the reflected pulses into a colour). The time delay between the transmission and reception of the pulses. (The weather radar system converts the time delay into the distance between the aircraft and the weather circumstances). The azimuth angle of the antenna. (The weather radar system uses the azimuth angle of the antenna to position the weather information on the display. A secondary function of the weather radar system is to show a ground terrain map of the area ahead of the aircraft. Therefore the crew tilts the antenna down with help of the tilt knob on the weather radar control panel. Because the cloud returns are different from the ground returns the weather radar system is switched over to the MAP mode for the correct interpretation of the ground returns. The radar must display three things: the range, the bearing and the signal intensity of the cloud. The display device best suited to showing all three of the above in an easily assimilated form is the cathode ray tube (CRT) used as a plan position indicator (PPI).

Page 2-190



Figure 146 shows a typical weather radar display unit.

GAIN

TILT SB/T

WX

WX/T

RCT

GCR

MAP

UP DWN

MIN

MAX

VAR

10

20

40

80

160

DISPLAY

OFF

320

SEC

MARKER

MAX

OFF

MAX

FRZ

FRZ

LEFT

RIGHT

INOP

ALRT

Weather Radar Display Unit Figure 146


Page 2-191


2.55.2 Scanner Stabilization As the aircraft pitches and rolls, the scanner will also pitch and roll. The weather radar is designed to scan directly in front of the aircraft, so as the aircraft pitches and roll; the scanner must pitch and roll in the opposite direction to that of the aircraft. The scanner is therefore mounted on a stabilized platform, which is maintained at a constant attitude with respect to the horizon. Stabilization is derived from the Inertial Reference System (IRS). The pitch and roll stabilization is completely independent system. Each having a separate motor, giving freedom of rotary movement in both pitch and roll. There must also be a freedom of movement in azimuth for scanning to port and starboard. Therefore three rotating joints are required in the scanner waveguide assembly. Figure 147 shows stabilization for pitch and roll.

ROLL ANGLE

PITCH ANGLE

AZIMUTH ANGLE 0°

AZIMUTH ANGLE 90°

NO STABILIZATION REQUIRED

NO STABILIZATION REQUIRED

ROLL ANGLE

PITCH ANGLE

AZIMUTH ANGLE 90°

AZIMUTH ANGLE 90°

WITH NO STABILIZATION WITH NO STABILIZATION

ROLL ANGLE

PITCH ANGLE

AZIMUTH ANGLE 90°

AZIMUTH ANGLE 90°

STABILIZED

STABILIZED

Roll/Pitch Stabilization Figure 147 Page 2-192



40

WX/T

80

RCT

160

GCR

TILT

UP

ALRT

MAX

RIGHT

OFF

MARKER

SEC

DWN

INOP

320

MAP

LEFT

INDICATOR CONTROL

20

WX

FRZ

MAX

10

SB/T

FRZ

OFF

MAX

VAR

DISPLAY

MIN

GAIN

POWER SUPPLIES

VIDEO

POWER

CONTROL

28V D.C.

115V A.C.

ROLL

SWEEP

TRANSMITTER -RECEIVER

PITCH

IRS

STAB

POWER

WAVEGUIDE

ANTENNA ASSEMBLY



Figure 148 shows a block schematic of a radar system.

Radar System Schematic Figure 148

Page 2-193


2.55.3 Weather Radar Installation

WX RX/TX

WAVEGUIDE SECTION 3 30" FLAT PLATE ANTENNA

WAVEGUIDE SECTION 2

WAVEGUIDE SECTION 1

ANTENNA PEDESTAL

Figure 149 shows the type of weather radar fitted to modern aircraft.

Weather Radar Installation (B737) Figure 149 Page 2-194



The weather radar system has a dedicated control panel for selection of the required mode of operation. This type of system uses the EFIS ND to display the weather information. Figure 150 shows the weather display on the EFIS ND on a Boeing 737 aircraft.

WX WEATHER RETURN

+10 14

MODE ANNUNCIATION

13

40

TILT ANGLE

1/2 RANGE INDICATION

EFIS ND Weather Display (B737 Aircraft) Figure 150


Page 2-195


The display is controlled via the EFIS control panel. Figures 151 and 152 shows an EFIS control panel and a weather radar control panel.

HSI

RANGE

EXP

ADI DH REF

150

80 160

VOR/ ILS

NAV

40

VOR/ ILS

MAP

NAV

CTR MAP

FULL

320

20 10

PLAN

WXR

ON

RST MAP BRT

VOR/ADF

ON

NAV AID

ARPT

ON

RTE DATA

ON

ON

WPT

ON

EFIS Control Panel Figure 152

MODE TEST

-7

WX

WX+T

MAP

10

-6

MIN

15

5

-5 -4

UP

TILT

0

GAIN -3 CAL

DN

-2

5

MAX -1

15 10

IDNT

STAB

Weather Radar Control Panel Figure 153 Page 2-196



2.55.4 Test Mode With the mode selector on the control panel in the TEST position the transmitter is on for 1 second for a transmitter test. For the remainder of the test the transmitter is off. A test picture is 'painted' on the EFIS. A test sweep of + 15° up and -15° down is carried out by the antenna. At the end of the test the antenna centralizes at 0°. The following precautions are to be observed for Boeing 737 weather radar ground operation: If the radar system is to be operated while the aircraft is on the ground, direct the nose of the aircraft such that a 240-degree forward sector is free of large metallic objects (hangars, other aircraft). Tilt the antenna upward 15 degrees and prevent personnel from standing closer than 10 feet to the 240-degree forward section of the aircraft. The receiver may be damaged as a result of strong returns from nearby metallic objects.


Page 2-197


120°

R = 6 MTRS

120°

RED/WHITE ROPE

WARNING SIGN

Figure 154 shows safety areas and boundary marks to be displayed during ground operation of weather radar.

Safety Areas and Boundary Marks Figure 154 Page 2-198



2.55.5 Radome The radome is an aerodynamically shaped nose cone made of a dielectric material, which can have an overriding effect on the weather radar system’s performance. The radome should transmit 90% of the incident energy, posses structural strength, protect against erosion, prevent spark discharge of static and protect against lightning strikes. Structural strength comes from how the radome is constructed. Normally they are of the sandwich type, consisting of a honeycomb structure supported on each side by a thin skin of laminated glass fibre. Anti static/erosion is overcome by coating the nose area with a polyurethane material. This material is bonded onto the radome, but must not be too thick so as to effect the transmission of energy. The radome is also coated with an anti static paint containing small graphite particles. Lightning strike protection takes the form of metal strips bonded to the surface of the radome and painted over. The strips run from the nose of the radome to the bulkhead, where good electrical bonding must be achieved so that any lightning strikes are dissipated in the airframe with minimum damage. Figure 155 shows the construction of a radome.

Radome Construction Figure 155


Page 2-199


2.56 GROUND PROXIMITY WARNING SYSTEM The purpose of the Ground Proximity Warning System (GPWS) is to alert the flight crew to the existence of an unsafe condition due to terrain proximity. The various hazardous conditions that may be encountered are divided into 7 Modes. These are: 1. Mode 1 - Excessive Descent Rate. 2. Mode 2 - Excessive Closure Rate (with respect to rising terrain). 3. Mode 3 - Excessive Altitude Loss (during climb-out after take-off). 4. Mode 4 - Insufficient Terrain Clearance (when not in landing configuration). 5. Mode 5 - Excessive Deviation below the Glideslope (ILS Landing). 6. Mode 6 - Descent Below selected Decision Height. 7. Mode 7 – Windshear. Figures 156 - 162 show schematics of each of the above modes.

“SINK RATE” WHOOP! WHOOP! PULL-UP

GPWS Mode 1 Figure 156 Page 2-200



“TERRAIN” “TERRAIN” “TERRAIN” “TERRAIN”

WHOOP! WHOOP! PULL-UP

GPWS Mode 2 Figure 157

“DON’T SINK”

GPWS Mode 3 Figure 158 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-201


“TOO LOW GEAR…...”


“GLIDESLOPE” “GLIDESLOPE


Page 2-202



“MINIMUMS” “MINIMUMS”

DECISION HEIGHT


STRONG DOWNDRAFT

HEADWIND

TAILWIND “WINDSHEAR” “WINDSHEAR”



Page 2-203


2.56.1 System Operation The main component of the system is the GPWS computer. It receives information from other aircraft systems (Baro/Rad Alt Ht, speed, etc.). From these inputs, the computer makes calculations to determine if the aircraft is in danger of contacting the terrain below. GPWS only operates within the Rad Alt range (50' to 2,500'). Figure 163 shows a block schematic diagram of a typical GPWS.

EFIS SYMBOL GENERATORS

PULL UP DATA & LOGIC INPUTS SYSTEM TEST

GROUND PROXIMITY WARNING COMPUTER

BELOW G/S P - INHIBIT

INOP

EADI PFD

EADI PFD

CAPT

F/O

GPWS CONTROL PANEL

RADIO ELECTRONICS UNIT

GPWS Block Schematic Figure 163

Page 2-204



2.56.2 Ground Proximity Warning Computer The GPWC establishes the limits for the GPWS modes and compares the aircraft’s flight and terrain clearance status against established mode limits. If the aircraft is found to have entered a GPWS mode, the computer issues appropriate warning or alerting signals. The computer also stores failure data in a nonvolatile memory for display on a front panel window on the GPWC. Figure 164 shows a GPWC and Control panel.

STATUS/HISTORY PRESENT STATUS

GROUND PROXIMITY

FLIGHT HISTORY

INOP

CAUTION OBSERVE PRECAUTIONS FOR HANDLING ELECTROSTATIC SENSITIVE DEVICES

FLAP/GEAR INHIBIT

NORMAL

SYS TEST

CONTROL PANEL

GROUND PROXIMITY WARNING COMPUTER

GPWC and GPWS Control Panel Figure 164


Page 2-205


2.56.3 GPWS Control Panel The GPWS control panel provides the flight crew with visual indications of GPWS operation; self-test capability and flap/gear inhibit capability. INOP Light Amber “INOP” light is illuminated when a computer or input signal malfunction is detected, or a GPWS self-test is being performed. Flap/Gear Inhibit This switch is a two-position toggle switch; guarded and safety-wired in the “NORMAL” position. When it is placed in the “INHIBIT” position, Modes 2,3 and 4 are inhibited. Self Test Switch This switch is used to initiate a GPWS self-test. A self-test can be conducted on the ground or in-flight. 2.56.4 Warning Lights Two warning lights are provided to give visual indication of ground proximity warnings. These are: 1.

PULL-UP.

2.

BELOW G/S.

A “WINDSHEAR” warning message (displayed on the EFIS PFD) provides visual indication of a Windshear condition. The red PULL-UP light illuminates when Mode 1,2,3 or 4 flight path is detected. The amber BELOW G/S warning light illuminates when glide slope deviation becomes excessive. Pressing the BELOW G/S switch inhibits the warning.

Page 2-206



Figure 165 shows a PFD with Windshear annunciation.

MCP SPD

CLMB

HDG SEL

V NAV

10 10

180 160

150 140

10 10

120 WINDSHEAR

GS 173

DH 350 RA 1620

Primary Flight Display (Windshear) Figure 165


Page 2-207


2.56.5 GPWS Bite Operation The purpose of the BITE is to perform an internal check of the GPWC functions, to record past faults that occur during the last ten flights, and to annunciate system status information. 2.56.6 BITE Tests The BITE function carries out three BITE tests: Continuous Test – Performed during each program loop. This checks the CPU operation and data input integrity for shorts to ground or open circuits. The ADC, IRS, ILS and RAD ALT systems and internal power supplies are also monitored for valid data. Periodic Test – Tests requiring excessive processing time are subdivided into small segments. Tests on the individual segments are performed sequentially, one segment during each program loop. Periodic tests include checks on the processor instruction sets, program memory contents, RAM addressing and storage functions, voice memory addressing and contents, parity of received data and the ability to read the data. Event-Initiated Tests – These are performed during or after a specific event has occurred. They include resetting the program a fraction of a second prior to a power supply failure. Checksumming the data stored in the non-volatile fault memory at power up. Checksumming the data written after entering data and sampling and storing program pin status at power up. Restarting the CPU at a known location in the program after loss of CPU. 2.56.7 Fault Recording Faults are recorded in a non-volatile fault memory by flight segments. The beginning and the end of each flight segment are identified using radio altitude, IAS and Mode 3 – 4 transitions. Up to 24 faults may be recorded during each flight segment.

Page 2-208



FMC

LS

LS

MODE CONTROL

IRU

PP/TKE

COURSE SELECT

FLAPS/AOA

HS

HS

PP/TKE/ROLL PITCH/ACCL

STALL WARNING

IAS/ALT ALT RATE

RAD ALT HT

LOC/GS

LS

LS

ADC

RAD ALT

ILS

FLAP POS GEAR POS

FDAU

G/S GPEW W/S

NORMAL

INHIBIT

INOP


PULL UP

MONITOR

G/S WARNING

G/S INHIBIT

GPWS WARNING

WINDSHEAR

CAPT PFD

GEAR POSITION SWITCHES

FLAP POSITION SWITCHES


PULL UP

F/O PFD



Figure 166 shows GPWS block schematic for the Boeing 737 aircraft.

GPWS Block Schematic (B737) Figure 166

Page 2-209


2.57 ENHANCED GROUND PROXIMITY WARNING SYSTEM The EGPWS contains all the modes as with the standard GPWS with some additional features. The system contains a worldwide terrain database, an obstacle database and a worldwide airport database, and using this extra data enables the system to give an Enhanced GPWS. The additional features are as follows: Terrain alerting and display (TAD) - This provides a graphic display of the surrounding terrain on the Weather Radar Indicator, EFIS or a dedicated GPWS display. Based on the aircraft’s position and the internal database (terrain topography), all terrain that is above or within 2000 feet below the aircraft’s altitude is presented on the system display. This feature is an option, enabled by program pins during installation. Peaks – Is a TAD supplemental feature providing additional terrain display features for enhanced situational awareness, independent of the aircraft’s altitude. This includes digital elevations for the highest and lowest displayed terrain, additional elevation (colour) bands, and a unique representation of 0 MSL elevation. This feature is an option enabled by program pins during installation. Obstacles – This feature utilizing an obstacle database for obstacle conflict alerting and display. EGPWS caution and warning visual and audio alerts are provided when a conflict is detected. Additionally, when TAD is enabled, Obstacles are graphically displayed similar to terrain. This feature is an option, enabled by program pins during installation. Terrain Clearance Floor – This feature adds and additional element of protection by alerting the flight crew of possible premature descent. This is intended for non-precision approaches and is based on the current aircraft position relative to the nearest runway. This feature is enabled with the TAD feature. Geometric Altitude – Based on the GPS altitude, this is a computed pseudobarometric altitude designed to reduce or eliminate altitude errors resulting from temperature extremes, non-standard pressure altitude conditions, and altimeter miss-sets. This ensures an optimal EGPWS alerting and display capability. Note; Some of these features have been added to the EGPWS as the system evolved and are not present in all EGPWS part numbers.

Page 2-210



2.57.1 Controlled Flight Into Terrain (CFIT) Because the overwhelming majority of “Controlled Flight Into terrain” accidents occur near to an airport, and the fact that aircraft operate in close proximity to terrain near an airport, the terrain database contains higher resolution grids for airport areas. Lower resolution grids are used outside airports areas where aircraft enroute altitude make CFIT accidents less likely and terrain feature detail is less important to the flight crew. With the use of accurate GPS and FMS information, the EGPWS is provided aircraft’s present position, track, and ground speed. With this information the EGPWS is able to present a graphical plan view of the aircraft relative to the terrain and advise the flight crew of any potential conflict with the terrain or an obstacle. Conflicts are recognised and alerts are provided when terrain violates specific computed envelope boundaries on the projected flight path of the aircraft. Alerts are provided in the form of visual light annunciation of a caution or warning, audio enunciation based on the type of conflict, and colour enhanced visual display of the terrain or obstacle relative to the forward look of the aircraft. Figure 167 shows Terrain/Obstacle database.

OBSTACLES SURVEY POINTS ABOVE SEA LEVEL

MEAN SEA LEVEL

Terrain/Obstacle Database Figure 167


Page 2-211


50% YELLOW

25% YELLOW

50% GREEN

16% GREEN

REF ALTITUDE +1000

REF ALTITUDE -250/500

REF ALTITUDE -1000

REF ALTITUDE -2000

REFERENCE ALTITUDE

MIN ELEVATION No

50% RED REF ALTITUDE +2000

MAX ELEVATION No

Figure 168 shows a graph on when caution and warning alerts are triggered.

Terrain Caution/Warning Graph Figure 168

Page 2-212



Table 3 shows the different Terrain/Obstacle threat levels and the colour indication present with TAD and Peaks selected.

Colour Solid Red Solid Yellow 50% Red Dots

Indication Terrain/Obstacle threat warning. Terrain/Obstacle threat warning. Terrain/Obstacle that is more than 2000 feet above the aircraft. 50% Yellow Dots Terrain/Obstacle that is between 1000 and 2000 feet above the aircraft’s attitude. 25% Yellow Dots Terrain/Obstacle that is 500 (250 with gear down) feet below to 1000 feet above the aircraft’s altitude. Solid Green Shown only when no red or yellow (Peaks Only) Terrain/Obstacle areas are within range on the display. Highest terrain/obstacle not within 500 (250 with gear down) feet of the aircraft’s altitude. 50% Green Dots Terrain/Obstacle that is 500 (250 with gear down) feet below to 1000 below the aircraft'’ altitude. 50% Green Dots Terrain/Obstacle that is in the middle elevation (Peaks Only) band when there is no red or yellow terrain areas within range on the display. 16% Green Terrain/Obstacle that is 1000 to 2000 feet below the aircraft’s altitude. 16% Green Terrain/Obstacle that is the lower elevation band (peaks Only) when there is no Red or Yellow terrain areas within range on the display. Black No significant Terrain/Obstacle 16% Cyan Water at Sea Level Elevation (0 feet MSL) Magenta Dots Unknown terrain. No terrain data in the database for the magenta area shown.

Terrain/Obstacle Threat Levels Table 3


Page 2-213


Figure 169 shows a Weather Radar Display used for EGPWS displays.

40

20

160

10 OFF CAUTION TERRAIN (YELLOW)

DIM

80

320 WARNING TERRAIN (RED)

RANGE

RNG 20

TERRAIN (GREEN)

TERR

EGPWS Display Figure 169 2.57.2 Terrain Alerting & Display (TAD) With a compatible EFIS or Weather Radar display, the EGPWS TAD feature provides an image of the surrounding terrain represented in various colours and intensities. There are two types of TAD display depending on the options selected: Standard TAD – Provides a terrain image only when the aircraft’s altitude is 2000 feet or less above the terrain. Peaks – Enhances the standard display characteristics to provide a higher degree of terrain awareness independent of the aircraft’s altitude. In either case, terrain and obstacles (if enabled) forward of the aircraft are displayed. Note; Obstacles are presented on the display as terrain, using the same colour scheme. Peaks and Obstacle functions are enabled by EGPWS program pin selection.

Page 2-214



16% BETWEEN 1000 - 2000 FEET

50% UPTO 1000 FEET

REF ALT

> 500 Ft

SOLID GREEN WHEN NO RED OR YELLOW

Figure 170 shows the “Peaks” function of EGPWS.

EGPWS “Peaks” Function Figure 170


Page 2-215


2.57.3 Envelope Modulation This special feature utilizes the internal database to tailor EGPWS alerts at certain geographical locations to reduce nuisance warning and provide added protection. Due to terrain features at or near certain specific airports around the world, in the past, normal operations have resulted in nuisance or missed alerts at these locations. With the introduction of accurate position information and a terrain and airport database, it is possible to identify these areas and adjust the normal alerting process to compensate for the condition. An EGPWS Envelope Modulation feature provides improved alert protection and expanded alerting margins at identified key locations throughout the world. This feature is automatic and requires no flight crew action. Modes 4,5, and 6 are expanded at certain locations to provide alerting protection consistent with normal approaches. Modes 1,2 and 4 are desensitized at other locations to prevent nuisance warnings that result from unusual terrain or approach procedures. In all cases, very specific information is used to correlate the aircraft position and phase of flight prior to modulating the envelopes. Figure 171 shows the Envelope Modulation function.

ENVELOPE MODULATION AREA

Envelope Modulation Figure 171 Page 2-216



2.57.4 Terrain Look Ahead Alerting Another enhancement provided by the internal terrain database, is the ability to look ahead of the aircraft and detect terrain or obstacle conflicts with greater alerting time. This is accomplished (when enabled) based on the aircraft position, flight path angle, track and speed relative to the terrain database image forward of the aircraft. Through sophisticated look ahead algorithms, both caution and warning alerts are generated if terrain or an obstacle conflict with “Ribbons” projected forward of the aircraft. Figure 172 shows the Terrain Look Ahead Alerting function.

WARNING (TYPICALLY 30 SEC AHEAD OF TERRAIN)

CAUTION (TYPICALLY 60 SEC AHEAD OF TERRAIN)

Terrain Look Ahead Alerting Figure 172


Page 2-217


laterally (more if turning). The look-ahead and up angles are a function of the aircraft flight path angle, and the look-ahead distance are a function of the aircraft’s altitude with respect to the nearest runway. This relationship prevents undesired alerts when taking off and landing. The look-ahead distance is a function of the aircraft’s speed and distance to the nearest runway. A terrain conflict intruding into the caution ribbon activates the EGPWS caution lights and the aural message “CAUTION TERRAIN, CAUTION TERRAIN” or “TERRAIN AHEAD, TERRAIN AHEAD”. The caution alert is given typically 60 seconds ahead of the terrain conflict and is repeated every seven seconds as long as the conflict remains within the caution area. When the warning ribbon is intruded, typically 30 seconds ahead of the terrain, EGPWS warning lights activate and the aural message “TERRAIN, TERRAIN, PULL UP” is enunciated with “PULL UP” repeating continuously while the conflict is within the warning area. Note; the specific aural message provided is established during the initial installation of the EGPWS and is a function of whether or not the terrain features are enabled and the selected audio menu (via program pins). 2.57.5 Terrain Clearance Floor (TCF) The TCF function enhances the basic GPWS Modes by alerting the flight crew of a descent below a defined “Terrain Clearance Floor” regardless of the aircraft’s configuration. The TCF alert is a function of the aircraft’s RAD ALT and distance (calculated from Lat/Long position) relative to the center of the nearest runway in the database. TCF alerts result in the illumination of the EGPWS caution lights and the aural message “TOO LOW TERRAIN”. The audio message is provided once when initial envelope penetration occurs and again only for an additional 20% decrease in RAD ALT altitude. The EGPWS caution lights will remain on until the TCF envelope is exited.

Page 2-218



The TCF envelope is shown in Figure 173

½ R U N WA Y LE NG TH

EN V ELOP E B IA S FA C TOR

15N M

12N M

4N M

15 N M

12 N M

4 NM

E N V E LO P E B IA S “TO O LOW TE RR A IN ” “TO O LOW TE RR A IN ”

4 0 0 ' - 7 0 0 ' ft

F AC T O R

0 ' - 4 0 0 ' ft

T E R RA IN C L E A RA N C E F L O O R

Terrain Clearance Floor Figure 173


Page 2-219


2.57.6 TCF/TAD Control The EGPWS TCF and TAD functions are available when all required data is present and acceptable. Aircraft position and numerous other parameters are monitored and verified for adequacy in order to perform these functions. If determined invalid or unavailable, the system will display “TERRAIN INOPERATIVE” or unavailable annunciation’s and discontinue the terrain display if active. TAD/TCF functions may be inhibited by manual selection of a cockpit “TERRAIN INHIBIT SWITCH”. Note; neither loss nor inhibited TAD/TCF effects the basics GPWS functions Modes 1 –7. Figure 174 shows EGPWS control switches and annunciations.

GND PROX G/S INHIBIT

FLAP OVRD

GEAR OVRD

G/S INHB

OVRD

OVRD

GND PROX

TERR OVRD

OVRD

EGPWS Control Switches & Annunciation Figure 174

Page 2-220



2.57.7 EGPWS Interface The EGPWS uses various input signals from other on-board systems. The full compliment of these other systems depends on the EGPWS configuration and options selected. The basic enhanced facilities require: 1. Altitude (RAD ALT/GPS/IRS). 2. Airspeed (IAS/TAS). 3. Attitude (IRS). 4. Glideslope (ILS). 5. Present Position (FMS/IRS/GPS). 6. Flap/Gear Position. 7. The Windshear function requires additional information of: a)

Accelerations (IRS).

b)

Angle of Attack.

c)

Flap Position.

Inputs are also required for discrete signals. These discrete inputs are used for system configuration, signal/status input and control input functions. EGPWS program pins are utilized to inform the system of the type of aircraft and interface that is in use. These are established during EGPWS installation. Discrete signals also include signals for “Decision Height”, Landing Flaps” selected, display range and status discrete such as RAD ALT/ILS valid. EGPWS provides both visual and audio outputs. The visual outputs provide discrete alert and status annunciations and display terrain video on a compatible CRT screen. Audio enunciations are provided (via the aircraft’s interphone system) at specific alert phases.


Page 2-221

Page 2-222

DADC IRS GPS FMS RAD ALT

AIRCRAFT SENSORS

CONTROL DISCRETE INPUTS

P R O C E S S I N G

I N P U T

EGPWC

WINDSHEAR DETECTION & ALERTING ALGORITHMS

TERRAIN CLEARANCE FLOOR ALGORITHMS

TERRAIN AWARENESS & OBSTACLE ALERTING & DISPLAY ALGORITHMS

AURAL CALLOUTS

GPWS ALGORITHMS

P R O C E S S I N G

O U T P U T

TERRAIN DISPLAY DATA

WARNING/ CAUTION LAMPS

AUDIO ALERT MESSAGES



Figure 175 shows EGPWS system schematic.

EGPWS System Figure 175



2.57.8 System Activation The EGPWS is fully active when the following systems are powered and functioning normally: 1.

EGPWS.

2.

RADIO ALTIMTER.

3.

AIR DATA SYSTEM

4.

ILS (Glideslope).

5.

GPS/FMS or IRS (PP).

6.

GEAR/FLAPS.

7.

WEATHER RADAR/EFIS DISPLAY.

In the event that the required data for a particular function is not available, then that function is automatically inhibited and annunciated (e.g. if PP data is not available or determined unacceptable, TAD/TCF is inhibited, any active terrain display is removed and “TERR INOP” indicated on CRT display. 2.57.9 Self Test The EGPWS provides a Self-Test Capability for verifying and indicating intended functions. This Self-Test capability consists of six levels to aid testing and troubleshooting the EGPWS. These six levels are: Level 1 - GO/NO GO Test. Provides an overview of the current operational functions and an indication of their status. The flight crew as part of their “PreFlight test” carries out this test. Level 2 - Current Faults. Provides a list of internal and external faults currently detected by the EGPWC. Level 3 – EGPWS Configuration. Indicates the current configuration by listing the EGPWS hardware, software, databases and program pin numbers detected by the EGPWC. Level 4 – Fault History. Provides an historical record of the internal and external faults detected by the EGPWC. Level 5 – Warning History. Provides an historical record of the alerts given by the EGPWS. Level 6 – Discrete Test. Provides audible indication of any change to a discrete input state.


Page 2-223


Note: Level 2 – 6 tests are typically used for installation checkout and maintenance operations. Figure 176 shows TAD/TCF display test pattern.

20

40

160

10 OFF

DIM

80

320 RANGE

RNG 160

TERR ST

TAD/TCF Test Display Figure 176

Page 2-224



Figure 177 shows a EGPWS computer as fitted to the Boeing 777.

EGPWS Computer (B777) Figure 177


Page 2-225


2.58 AIR DATA SYSTEM (ADS) Air Data systems depend upon Pitot and Static pressure sensing, as well as temperature sensing. Static air pressure is the pressure of the outside air at the location of the aircraft. Pitot pressure is the dynamic pressure caused by the forward motion of the aircraft. Temperature sensing is required to calculate the Total/Static Air Temperature (TAT/SAT), and for calculating True Air Speed (TAS). Figure 178 shows a Pitot/Static probe.

STATIC LINE NO1 PITOT LINE

HEATER CONNECTION

STATIC LINE NO2 PITOT

STATIC PORTS

Pitot/Static Probe Figure 178 In a parked aircraft, the pitot and static pressures are equal. In a moving aircraft, the pitot pressure is greater because additional pressure is developed at the forward end of the tube by its motion through the air. Altitude is calculated on the basis of static air pressure, and airspeeds are calculated on the basis of the difference between pitot and static pressures. Since a pitot/static probe is, under certain conditions, subjected to icing, it is necessary to have available a heater to melt the ice which would block the ports. Flush static ports may also be heated, if required.

Page 2-226



2.58.1 Total Air Temperature Probe Total Air temperature (TAT), is the static air temperature plus the rise in temperature created due to the pitot effect. TAT is of great importance in setting the operating conditions of a jet engine, since the temperature of the air entering the engine is static air temperature increased by the pitot factor. It is also possible to derive Static Air Temperature (SAT), from TAT and pitot pressure information. Figure 179 shows a basic Total Air Temperature (TAT) probe.

AIR FLOW

SENSING ELEMENT

METERED ORIFICE (VERY SMALL)

ELECTRICAL CONNECTION ELECTRICAL SIGNAL OF TEMPERATURE

Total Air Temperature Probe Figure 179 The TAT probe is constructed similar to a pitot probe. It is however, more complicated due to the need for providing de-icing heat. The air entering the TAT probe must be shielded from the de-icing heat. The TAT probe has metered orifices to allow the air to flow through the sensing element.


Page 2-227


2.58.2 Location Of Probes And Static Vents The choice of probe/vent locations is largely dependent on the type of aircraft, speed range and aerodynamic characteristics, and as result there is no common standard for all aircraft. On larger aircraft it is normal to have standby probes and static vents. These are always located one on each side of the fuselage and are interconnected so as to balance out dynamic pressure effects resulting from any Yawing or side-slip motion of the aircraft. Figure 180 shows the location of the pitot/static sensing elements on a Boeing 737 aircraft.

Boeing 737 Pitot/Static Locations Figure 180

Page 2-228



Pitot and static pressures are transmitted through seamless and corrosionresistant metal (light alloy) pipelines. Flexible pipelines are also used when connections to components mounted on anti-vibration mountings are required. In order for an Air Data System to operate effectively under all flight conditions, provision must also be made for the elimination of water that may enter the system as a result of condensation, rain, snow, etc. This will reduce the probability of “Slugs” of water blocking the lines. This provision takes the form of drain holes in the probes, drain taps and valves in the system’s pipelines. The drain holes within the probes are of diameter so as not to introduce errors into the system. Methods of draining the pipelines varies between aircraft types and are designed to have a capacity sufficient to allow for the accumulation of the maximum amount of water that could enter the system between maintenance periods. Figure 181 shows a typical water drain valve.

ORANGE FLOAT INDICATOR

TRANSPARENT PLASTIC PIPE

DRAIN VALVE

BAYONET FITTING CAP

(SELF SEALING)

Water Drain Figure 181


Page 2-229


LOWER PRESSURE HEADS

UPPER

PRESSURE HEADS LOWER UPPER

PC

A/S 2

MS 1

DIFF PRESS

A/S 1

MS 2

ADC 2

ADC 1

FLT REC

Figure 182 shows a typical air data system for a large aircraft.

PITOT

PITOT

F/O

CAPT

IAS

VS

ALT

ALT

VS

STATIC

IAS

STATIC

Air Data System Figure 182 Page 2-230



2.58.3 Air Data Computer (ADC) The ADC receives Pitot/Static pressure and an electrical signal representing temperature. It uses these inputs to calculate the following parameters: 1.

Indicated Air Speed (IAS).

2.

True Airspeed (TAS).

3.

Speed of Sound (Mach).

4.

Altitude.

5.

Rate of change of Altitude.

6.

Total Air Temperature (TAT).

7.

Static Air Temperature (SAT).

The system supplies air data information to any system in either digital or analog form.


Page 2-231


TRUE AIRSPEED STATIC AIR TEMP

TOTAL AIR TEMPERATURE

MACH PITOT PRESSURE

INDICATED AIRSPEED STATIC PRESSURE

TRUE AIRSPEED

AUTO THROTTLE MACH HOLD (AUTOPILOT)

AUTO THROTTLE AIRSPEED HOLD (AUTOPILOT) FDR

ALTITUDE

VSI, FDR ATC TRANSPONDER CABIN PRESSURE ALT HOLD (AUTOPILOT) ALT RATE (AUTOPILOT)

Figure 183 shows a block schematic of an Air data Computer.

Air Data Computer Figure 183 The four blocks within the computer represent modules, which are capable of supplying the information and controls indicated with the information required. Page 2-232



2.58.4 Altitude Module The module contains a capsule for measuring the static pressure and an “E & I” pick off. The “E & I” pick-off changes the capsule movement into an electrical signal. This electrical signal is amplified and then fed to a motor servo system. The motor will drive a gear train to move the output devices to give the correct altitude reading. Figure 184 shows the altitude module of an ADC.

AC EXCITATION

M STATIC PRESSURE ALTITUDE RATE

G

GEAR TRAIN ALTITUDE DIGITAL ENCODER

CT

ATC TRANSPONDER

ALTITUDE 0V

+10V

Altitude Module Figure 184 If the aircraft is parked, or on a holding altitude, the static pressure will be constant. The servo motor will have driven until the force exerted by the spring balances the force exerted by the evacuated bellows, and the “E” pick-off armature has been moved to its null position. Any position other than the null causes the servo-motor to run in one direction or the other. The tacho-generator gives position feedback and also gives an output off altitude rate.


Page 2-233


2.58.5 True Airspeed/Indicated Airspeed Vs Altitude True Airspeed (TAS) and Indicated Airspeed (IAS) are the same value at sea level; however, as altitude increases, holding the same IAS results in an increasing TAS. For example; 400kts IAS at sea level becomes 450kts TAS at 10,000ft, and approximately 550kts TAS at 20,000ft. Figure 185 shows the relationship between IAS/TAS/Mack with an increase of altitude.

.90 .885

50

MACH NO

40 IAS

36,000ft

350 30

400 21,000ft

20

ALTITUDE 1,000ft 10

0 250

300

350

400

450

500

550

600

TAS/IAS/Mach Vs Altitude Figure 185 In the diagram above the Mach No lines are drawn on the basis of a standard day temperature chart; It can be seen that .90 Mach at sea level would be 600kts TAS. From above 36,000ft, .90 Mach equals only 525kts TAS. If an aircraft is limited to Mach .885/IAS 390kts, it could fly at 390kts up to 21,000ft. Above this height, the IAS would have to decrease to ensure that the maximum Mach No of .885 is not exceeded.

Page 2-234



2.58.6 Air Data Computer (ADC)

MACH MODULE

M POT

ALT (ALT MODULE)

POT MACH

TAS MODULE

POT

IAS MODULE

ALT RATE

ALT (MACH MODULE)

E PICK-OFF

ALT MODULE

G

POT

M

ENCODER CT

CT

ATCRBS ALT HOLD

ALT

ALT CABIN PRESS

E PICK-OFF

M

CT

IAS

POT

CT

TAS

M

TAT

POT

Figure 186 shows the four modules of an ADC.

Air Data Computer Modules Figure 186 MOD 11 BOOK 2 PART 2 ISSUE 6 - 01/02/11

Page 2-235


2.58.7 Digital Air Data Computer (DADC) The DADC uses digital computing and electronic circuits rather than the servo motor system to calculate the outputs. Analog inputs are converted into digital for computation. The outputs required are then either converted back to analog or left as digital signals and output via the ARINC 429 or 629 data busses.

IAS HOLD ENGAGE

ALT HOLD ENGAGE

MACH HOLD ENGAGE

TEMP

BRIDGE CCTS & MULTIPLEXER

PITOT TRANDUCER PITOT

MILLIBAR SET

STATIC TRANDUCER STATIC

I/P CCTS

CENTRAL PROCESSOR & MEMORY

A/D

D/A

O/P

DIGITAL O/Ps

ANALOG O/Ps

Figure 187 shows a DADC block schematic.

Digital Air Data Computer (DADC) Figure 187 Page 2-236



2.58.8 Definitions and Abbreviations Static pressure (Ps) - Is the ambient atmosphere pressure, which acts on the surface of a body in rest. Total pressure (Pt) or pitot pressure - Is the sum of static pressure and the impact pressure and is the total force which acts on the surface of a body in motion. Impact pressure - Is the force we need to stop moving air. It is the actual pressure, which a body in motion feels. It is equal to total pressure minus static pressure (Pt-Ps). The relation impact pressure - speed is not linear due to the compressibility and changes in density of the air. The computer software has the formula to change impact pressure into airspeed. Pressure altitude - Is the altitude in a standard atmosphere. We do not take in account the variations in pressure or temperature, which occur on earth when the weather changes. A standard atmosphere is equal to 29.921 inches of mercury (in Hg) or 1013.25 milibars (mB) at sea level. Baro corrected altitude - Is the pressure altitude corrected for QFE or QNH barometric correction signals (this requires a manual input for the computer from an altimeter set panel). Altitude rate - The change in altitude in ft per min. We also call this signal vertical speed, vertical rate, rate of climb or baro rate. Computed airspeed (CAS) - Is an air data function related to impact pressure. The computer corrects the airspeed for instrument errors and installation errors. Mach number (Ma) - Is the ratio between the true airspeed and the speed of sound. Maximum operational speed (Vmo/Mmo) - Vmo is the maximum safe airspeed an aircraft can fly with no excessive stress on its structure. The Mmo is the maximum safe mach number an aircraft can fly without the negative effects caused by subsonic shock waves. The aircraft manufacturer specifies the Vmo/Mmo and this value is pre-programmed in the ADC. Overspeed warning - A discrete signal from the ADC, present at Vmo/Mmo. Angle Of Attack (AOA) - Is the angle between the longitudinal axis of the aircraft and the flight path of the aircraft measured from the aircraft’s centre of mass. Corrected angle of attack - Is the local AOA from the AOA transducers corrected for errors as a function of machnumber.


Page 2-237


True airspeed (TAS) - Is the speed of the aircraft with respect to the ambient air through which it flies. It is airspeed (impact pressure) corrected for compressibility and density. This depends on altitude and temperature. Total Air Temperature (TAT) - To measure the outside air temperature we install a sensor outside the aircraft. When we fly the sensor is in an airstream. The airstream hits the sensor, comes to a stop, rises in pressure and therefore rises in temperature. The air temperature plus the temperature rise is called total air temperature. Static Air Temperature (SAT) - Is the temperature of the undisturbed ambient air (TAT corrected for speed).

Page 2-238



PART THREE CONTENTS 3

ELECTRICAL POWER ................................................................. 3-1 3.1

3.2

3.3

3.4

3.5

3.6

3.7

BATTERIES .................................................................................. 3-1 3.1.1 Primary Cell/Secondary Cell ......................................... 3-1 3.1.2 Cell Voltage And Capacitance ....................................... 3-1 3.1.3 Lead Acid Cell ............................................................... 3-3 3.1.4 Chemical Action ............................................................ 3-4 LEAD/ACID BATTERIES ................................................................. 3-5 3.2.1 Cell Characteristics ....................................................... 3-6 3.2.2 Capacity of Batteries ..................................................... 3-7 3.2.3 State of Charge ............................................................. 3-8 BATTERY CHARGING (WORKSHOP) ................................................ 3-9 3.3.1 Preparation for Charge .................................................. 3-9 3.3.2 Charging The Battery .................................................... 3-9 3.3.3 Completion of Charge ................................................... 3-10 3.3.4 Capacity Test ................................................................ 3-11 3.3.5 Fully Discharged Condition ........................................... 3-12 NICKEL/CADMIUM (NI/CD) CELL ................................................... 3-12 3.4.1 Chemical Action ............................................................ 3-13 3.4.2 Nickel Cadmium Batteries ............................................. 3-15 3.4.3 Thermal Runaway ......................................................... 3-16 3.4.4 Causes.......................................................................... 3-16 NI/CD BATTERY CHARGING .......................................................... 3-17 3.5.1 Cell Caps ...................................................................... 3-17 3.5.2 Voltmeter ...................................................................... 3-17 3.5.3 Cell Shorting Links ........................................................ 3-17 3.5.4 Battery Characteristics .................................................. 3-18 3.5.5 Battery Characteristics .................................................. 3-19 3.5.6 Preparation For Charge................................................. 3-19 3.5.7 Constant Current Charging ........................................... 3-20 3.5.8 Charging Rate ............................................................... 3-20 3.5.9 Action Prior To Charge .................................................. 3-20 METHODS OF CONSTANT CHARGING ............................................. 3-21 3.6.1 Method 1 ....................................................................... 3-21 3.6.2 Method 2 ....................................................................... 3-22 3.6.3 Completion Of Charge .................................................. 3-23 BATTERY TESTING ....................................................................... 3-23 3.7.1 Capacity Test ................................................................ 3-23 3.7.2 Capacity Test ................................................................ 3-23 3.7.3 Capacity Recycling........................................................ 3-24 3.7.4 Deep Discharge ............................................................ 3-24 3.7.5 Cell Balancing ............................................................... 3-25 3.7.6 Voltage Recovery Check ............................................... 3-26 3.7.7 Storage ......................................................................... 3-26 3.7.8 Ready For Service ........................................................ 3-26 3.7.9 Long Term..................................................................... 3-26 3.7.10 Facts And Figures ......................................................... 3-26 3.7.11 Aircraft Charging Systems ............................................. 3-27 3.7.12 Constant Current Mode ................................................. 3-27


Page 1


3.8

3.9

3.10

3.11 3.12 3.13 3.14 3.15

3.16 3.17 3.18 3.19 3.20

3.21

3.22

Page 2

3.7.13 Constant Voltage Mode ................................................ 3-27 3.7.14 Charger Isolation .......................................................... 3-28 TYPICAL AIRCRAFT BATTERY SYSTEM .......................................... 3-29 3.8.1 Parallel/Series Batteries ............................................... 3-31 3.8.2 Aircraft Battery Charger Units ....................................... 3-32 3.8.3 DC-10 Charger Unit ...................................................... 3-34 3.8.4 Operation...................................................................... 3-34 3.8.5 Charging Unit................................................................ 3-34 3.8.6 Boeing 373 Charger System ......................................... 3-35 DC POWER GENERATION ............................................................. 3-38 3.9.1 DC Generator ............................................................... 3-38 3.9.2 Fixed Winding Arrangement ......................................... 3-39 VOLTAGE REGULATION ................................................................ 3-40 3.10.1 Vibrating Contact Type Regulator ................................. 3-40 3.10.2 Carbon Pile Voltage Regulator ..................................... 3-41 3.10.3 Transistorised Voltage Regulation ................................ 3-43 REVERSE CURRENT CUT-OUT RELAY ........................................... 3-44 CURRENT LIMITER ....................................................................... 3-45 THREE UNIT CONTROL PANEL ...................................................... 3-47 PARALLEL & LOAD SHARING ........................................................ 3-49 AC POWER GENERATION ............................................................. 3-53 3.15.1 Brushless Generators ................................................... 3-53 3.15.2 Constant Speed Drive (CSD) Unit................................. 3-54 3.15.3 Variable Displacement Unit........................................... 3-55 3.15.4 Control Cylinder ............................................................ 3-55 3.15.5 Governor ...................................................................... 3-55 3.15.6 Fixed Displacement Unit ............................................... 3-55 3.15.7 Differential Gear Unit .................................................... 3-55 3.15.8 CSD Operation ............................................................. 3-56 3.15.9 Underdrive Phase ......................................................... 3-56 3.15.10 Overdrive Phase ........................................................... 3-57 3.15.11 CSD Disconnection....................................................... 3-59 INTEGRATED DRIVE GENERATOR (IDG)......................................... 3-60 FIELD EXCITATION ....................................................................... 3-61 VOLTAGE REGULATION ................................................................ 3-61 3.18.1 Operation...................................................................... 3-61 VARIABLE-SPEED CONSTANT-FREQUENCY POWER SYSTEMS ........ 3-64 FREQUENCY-WILD SYSTEMS ........................................................ 3-66 3.20.1 Frequency-Wild Generator Construction ....................... 3-66 3.20.2 Operation...................................................................... 3-67 3.20.3 Stator Assembly ........................................................... 3-67 3.20.4 Rotor Assembly ............................................................ 3-67 3.20.5 Generator Cooling ........................................................ 3-67 3.20.6 Frequency Wild Generator Excitation ........................... 3-69 THREE PHASE GENERATOR .......................................................... 3-70 3.21.1 Interconnection Of Phases............................................ 3-71 3.21.2 Star Connection ............................................................ 3-71 3.21.3 Delta Connection .......................................................... 3-72 BUSBARS .................................................................................... 3-73 3.22.1 Busbar Systems ........................................................... 3-73



3.23

3.24

3.25 3.26

3.27 3.28

3.29

3.30 3.31 3.32 3.33

3.22.2 Split Bus-Bar A.C. Generation System .......................... 3-75 3.22.3 Bus-Bar Supply Priority ................................................. 3-78 3.22.4 Parallel Electrical System .............................................. 3-81 3.22.5 Split Parallel Electrical System ...................................... 3-82 GENERATOR CONTROL UNITS (GCU) ............................................ 3-83 3.23.1 Power Distribution System Control ................................ 3-83 3.23.2 Current Transformers .................................................... 3-85 3.23.3 Generator Control & Protection ..................................... 3-87 3.23.4 Under-voltage & Reverse Phase Sequence Unit ........... 3-89 3.23.5 Time Delay Activation ................................................... 3-91 3.23.6 Abnormal Frequency Protection .................................... 3-92 3.23.7 Differential Current Protection ....................................... 3-93 3.23.8 Over-Current Protection ................................................ 3-95 3.23.9 GCU Operation ............................................................. 3-96 TRANSFORMERS .......................................................................... 3-98 3.24.1 Voltage transformers ..................................................... 3-98 3.24.2 Transformer Ratings ..................................................... 3-100 3.24.3 Transformer Rectifier Units (TRU) ................................. 3-101 ROTARY INVERTER ....................................................................... 3-105 3.25.1 Static Inverter ................................................................ 3-107 BOEING 737 ELECTRICAL SYSTEM ................................................ 3-109 3.26.1 Controls & Indications ................................................... 3-110 3.26.2 Boeing 737 P5-13 Electrical Panel ................................ 3-111 3.26.3 Boeing 737 P5-5 Electrical Panel .................................. 3-112 3.26.4 B737 Electrical Power Distribution ................................ 3-118 3.26.5 Operation ...................................................................... 3-119 3.26.6 Generator Feeder Lines ................................................ 3-120 3.26.7 Boeing 737 D.C. Power ................................................. 3-123 B747 GENERATING SYSTEM ......................................................... 3-124 3.27.1 Operation ...................................................................... 3-125 LOAD SHARING ............................................................................ 3-126 3.28.1 Real Load Division ........................................................ 3-127 3.28.2 Reactive Load Division .................................................. 3-130 EMERGENCY AC POWER GENERATION .......................................... 3-136 3.29.1 Standby Generator ........................................................ 3-136 3.29.2 Auxiliary Power Unit (APU) Boeing 737......................... 3-136 3.29.3 Ram Air Turbine (RAT).................................................. 3-137 3.29.4 Emergency Pump.......................................................... 3-138 3.29.5 Emergency Generator ................................................... 3-138 3.29.6 Generator And Pump .................................................... 3-138 3.29.7 Extended Twin Engined Operations (ETOPS) ............... 3-138 EXTERNAL/GROUND POWER ......................................................... 3-139 DC EXTERNAL POWER ................................................................. 3-139 3.31.1 External DC – Multiple Busbar System.......................... 3-141 AC EXTERNAL POWER ................................................................. 3-143 3.32.1 A.C. External Power Circuit ........................................... 3-145 B747 ELECTRICAL SYSTEM .......................................................... 3-148 3.33.1 Normal Operation .......................................................... 3-148 3.33.2 Ground Handling and Ground Service Systems ............ 3-148 3.33.3 Main Standby System ................................................... 3-149 3.33.4 APU Standby Power System ......................................... 3-150


Page 3


3.34

3.35 3.36 3.37 3.38

Page 4

3.33.5 Electrical System Control Module ................................. 3-152 3.33.6 Electrical Synoptic EICAS Display ................................ 3-154 3.33.7 DC Distribution ............................................................. 3-156 CIRCUIT PROTECTION .................................................................. 3-158 3.34.1 Fuses ........................................................................... 3-159 3.34.2 Current Limiters ............................................................ 3-159 CIRCUIT BREAKERS ..................................................................... 3-161 REVERSE CURRENT CUT-OUT RELAY ........................................... 3-163 3.36.1 Operation...................................................................... 3-166 OVERVOLTAGE PROTECTION ........................................................ 3-167 3.37.1 Operation...................................................................... 3-168 SOLID STATE OVERVOLT PROTECTION.......................................... 3-169 3.38.1 List of Abbreviations ..................................................... 3-171


JAR 66 CATEGORY B1 CONVERSION COURSE MODULE 11 AEROPLANE AERODYNAMICS, STRUCTURES & SYSTEMS

3

MODULE 11.6 ELECTRICAL POWER

ELECTRICAL POWER

3.1 BATTERIES In almost all aircraft electrical systems a battery has the following principal functions: To help maintain the dc system voltage under transient high load current. To supply power for short-term heavy loads when generator, or ground power, is not available: e.g. engine starting. To supply power for essential services, under emergency conditions. 3.1.1 Primary Cell/Secondary Cell A battery is a device for converting chemical energy into electrical energy and is made up of a number of primary or secondary cells. As a primary cell discharges, i.e. supplies electrical energy, the chemical action destroys the cell and it cannot be re-formed, i.e. charged. As a secondary cell discharges, the chemical action converts the cell material into other forms and these can be converted into the original material, i.e. charged. Therefore secondary cells can be discharged and charged during the 'life' of a battery. Secondary cells are used in aircraft batteries of which there are two types 1.

Lead - Acid (L/A)

2.

Nickel - Cadmium (Ni/Cd)

3.1.2 Cell Voltage And Capacitance Each cell gives a voltage: 1.

The nominal voltage of a L/A cell is 2 volts

2.

The nominal voltage of a Ni/Cd is 1.2 volts.

Each cell has capacity, a measure of current it is capable of delivering over 1 hour. The unit is Ampere-Hour (AH).


Page 3-1


If cells are connected in series, the total voltage across the arrangement is the sum of each cell voltage. The capacity is as for one cell. If cells are connected in parallel, the total voltage is as for one cell. The total capacity of the arrangement is the sum of each cell capacity. Figure 1 shows the connection of cells.

SERIES CONNECTION

PARALLEL CONNECTION

Cell Connection Figure 1

Page 3-2




3.1.3 Lead Acid Cell The cell consists of a positive electrode and a negative electrode, each made up of a group of lead-antimony alloy grid plates; the positive plates have lead peroxide paste (Pb 02) forced in as the active material and the negative plates have pure spongy lead (pb) forced into them. Figure 2 show the arrangement of plates in a cell.

POSITIVE PLATE GROUP

NEGATIVE PLATE GROUP

C E LL C O N NE C TO R

VEN T C A P

S E P A RA TO R P R O TE C TO R

P LA TE S TR AP

C E LL C OVER

TE R M IN A L POST S

S E P A RA TO RS

P LA TE

Cell Arrangement Figure 2 Note that there are more negative plates than positive plates. This is because positive plates may buckle under discharge; negative plates do not buckle so when the cell is complete, positive plates are completely enclosed by negative plates, keeping buckling to a minimum. The electrolyte consists of two constituents, sulphuric acid (H2SO4) and water (H2O), which are mixed in such proportions that the relative density (RD) is generally about 1.25 to 1.27 for a charged cell.


Page 3-3


3.1.4 Chemical Action During discharge of the cell, i.e. when an external circuit is completed between the positive and negative plates, electrons are transferred through the circuit from lead (negative plates) to lead peroxide (positive plates), the net result of the chemical reaction is that lead sulphate (PbSO4) forms on both plates. At the same time molecules of water are formed, thus weakening the electrolyte solution. The cell is therefore considered discharged when both plates are covered with lead sulphate and the electrolyte has become weaker. Figure 3 shows a lead acid battery in charged and discharged states.

E L E CT R O N S

Pb

O2

Pb

L E AD

L E AD P E R O X ID E

H2

SO4

S U L P H U R IC A C I D

B AT TE R Y IN A C H AR G E D S T AT E

Pb

Pb

L E AD S U LP H A TE

Pb

O2 Pb

SO4

SO4

L E AD S U LP H A TE

H2

SO4

D IL U T E D S U L P H U R I C A C ID

B AT TE R Y IN A D IS C H AR G E D S T AT E

Lead Acid – Charged/Discharged Figure 3

Page 3-4




The cell may be recharged by connecting the positive and negative plates, respectively, to the positive and negative terminals of a D.C. source of a slightly higher voltage than the cell. All foregoing reactions are then reversed; the lead sulphate on the positive plate being restored to lead peroxide, the negative plate restored to lead, and the electrolyte restored to its original relative density. 3.2 LEAD/ACID BATTERIES The battery in figure 4 is made up of two blocks, each containing six cells of 2 volts per cell, connected in series. Hence, each block delivers 6 x 2 volts = 12 volts. Since there are 2 blocks of 12 volts in series, Battery Voltage = 24V.

Lead/Acid Battery Figure 4


Page 3-5


3.2.1 Cell Characteristics The chemical action of a lead acid battery is shown in Table 1.

Battery Type

State of charge

Positive Plate

Negative Plate

Electrolyte

Lead-Acid

Charged

PbO2 (Lead Dioxide)

Pb (Lead)

H2SO4 Concentrated Sulphuric Acid

Lead-Acid

Discharged

PbSO4 (Lead Sulphate)

PbSO4 (Lead Sulphate)

H2SO4 Weak Sulphuric Acid

Table 1 Fully charged cell voltage

=

2.2 volts (Approx). 2.0 volts (Nominal).

Discharged cell voltage

=

1.8 volts.

Relative Density (RD) of electrolyte: Charged

=

1.25 – 1.27.

Discharged

=

1.150.

Note: The solution becomes weaker on discharge and that the SG figures may vary and manufacturer's instructions should be referred to.

Page 3-6


JAR 66 CATEGORY B1 CONVERSION COURSE

MODULE 11.6

MODULE 11 AEROPLANE AERODYNAMICS, STRUCTURES & SYSTEMS

ELECTRICAL POWER

3.2.2 Capacity of Batteries The capacity of a battery, or the total amount of energy available, depends upon the size and number of plates. The capacity rating is measured in AmperesHours and is based on the maximum current, in amps, which it will deliver for a known time period, until it is discharged to a permissible minimum voltage of each cell. The time taken to discharge is called the “Discharge rate” and the rated capacity of the battery is the product of this rate and duration of discharge (in hours). Thus a battery which discharges 5A for 5 hours is rated at 25 Amperes-Hours capacity. Figure 5 shows the discharge rate for a 24V lead-acid battery rated at 20 Amperes/Hour.

LEAD-ACID - RATED AT 20 AMPERE/HOUR

T E R M I N A L V O L T A G E

24V

21.6V

DISCHARGED

20 AMP DISCHARGE

½

1

10 AMP DISCHARGE

1½

2

TIME IN HOURS

Discharge Rate – 20 Ampere/Hour Battery Figure 5


Page 3-7


3.2.3 State of Charge All batteries display certain indications of their state of charge, and these are of practical help in maintaining operating conditions. When a lead-acid battery is in the fully charged condition each cell displays three distinct indications: 1. Terminal voltage reaches its maximum value and remains steady. 2. Relative density of the electrolyte ceases to rise and remains steady. 3. The plates gas freely. The relative density is the sole reliable guide to the electrical condition of the cell of a battery which is neither fully charged nor yet completely discharged. If the relative density is midway between normal maximum (1.25 – 1.27) and minimum (1.150), then the cell is approximately half discharged. Figure 6 shows a Hydrometer used to check the RD of lead-acid batteries. CHARGED = 1.260 RUBBER BULB

DISCHARGED = 1.150

SYRINGE

1.100 1.150 1.100 FLOAT

1.250 1.300 1.350 1.400

RUBBER TUBE

Hydrometer Figure 6

Page 3-8




3.3 BATTERY CHARGING (WORKSHOP) 3.3.1 Preparation for Charge The procedure is as follows: 1. Unscrew the vent plugs but leave in the vent holes. This allows the cell to gas freely during the charge. 2. Adjust the level of electrolyte, if required, to the level specified in the manufacturer's instructions by adding distilled water. The plates must always be covered, do not over-fill. Record amount of distilled water added in ccs. Note: Most batteries have a perforated strip above the plates to protect against foreign objects. The level is measured from this strip. 3. Connect to charging board. 3.3.2 Charging The Battery The charging rate (current) is the value specified in the manufacturer's instructions. A typical figure is 3.5 amps. The larger the ampere-hour (AH), the higher the current required to charge a battery. The charge must be monitored at frequent intervals to: 1. Adjust the charging current, as cell voltage will increase during the charge. 2. Ensure electrolyte remains above plates and cells are gassing. Adjust the level by adding distilled water. Record the quantity added; if the battery is always requiring distilled water, it must be rejected. 3. Monitor electrolyte temperature; stop the charge if the value rises above that specified by the manufacturer (approximately 60°C) until the temperature drops (to approximately 12°C). 4. Record the terminal voltage to determine when the battery is fully charged. 5. Record the RD. This will indicate when the battery is fully charged. Do not forget that if you want to find the state of charge by measuring the RD, it is relative to a temperature of 15°C and a correction must be made.


Page 3-9


The manufacturer's instructions give RD figures at a standard temperature, therefore a correction must be made to the actual RD reading to bring it back to the standard temperature reading. The standard temperature is 15°C (60°F). Correction figures are: 1. For every 4°C above 15C° add 0.003 to RD. 2. For every 4°C below 15°C subtract 0.003 from RD. (For 60°F use 0.001 and 2.5°F in the same manner) Without this correction we will not know the state of the cell. After level adjustment, the electrolyte is mixed by the cell gassing while still on charge, therefore, any RD reading taken immediately after adding distilled water will be incorrect. 3.3.3 Completion of Charge Completion of charge will be indicated as follows: 1. Constant terminal voltage, with charging current flowing, for three hours, 2. Constant RD and within the manufacturer's limit (after temperature correction), 3. Cells gassing freely. If all three conditions are met, the battery is fully charged and charging should cease. The final 'on charge' voltage will vary on different batteries. It is normally between 30 and 32.4 volts (2.5V to 2.7V per cell). The RD is approximately 1.260 at 15°C. Following a charge, the voltage immediately falls. If it falls below 28.5 volts the battery must be rejected for service. Over a longer period the voltage will fall below 28.5 volts. Note that gently rocking the battery disperses any gas retained in the electrolyte. Check electrolyte level one hour after battery is removed from charge. If distilled water needs to be added, re-connect to the charging board and adjust while on a low charge and gassing. NB: Gassing aids mixing. Finally, record date and state of charge.

Page 3-10



MODULE 11.6


ELECTRICAL POWER

Figure 7 shows the charge curve for a lead-acid battery.

T E R M I N A L

FOLLOWING CHARGE BATTERY TERMINAL VOLTAGE MUST NOT FALL BELOW 28.5V FINAL “ON CHARGE” VOLTAGE

30V 28V 24V

V O L T A G E

CONSTANT CHARGE CURRENT

TIME Charge Curve – Lead-Acid Battery Figure 7 3.3.4 Capacity Test Reasons for a Capacity test are as follows: 1. After initial charge. 2. Routine maintenance at specified periods: e.g. 3 months. 3. If the capacity of the battery is in doubt. This test is to determine whether the battery will be able to carry out its function as an emergency power source on the aircraft. To measure capacity, a fully charged battery is discharged at the battery rating, whilst the time to discharge is recorded, i.e. a 30AH battery at the one hour rate is discharged at 30 amps.


Page 3-11


3.3.5 Fully Discharged Condition There are two conditions to check: 1. First cell to reach 1.8 volts, or, if cell voltage cannot be read (block construction), read battery voltage. A 24V battery reads 21.6V (12 x 1.8), OR, 2. First cell to reach discharge level of RD. After a capacity test, the battery should be re-charged as described earlier. 3.4 NICKEL/CADMIUM (NI/CD) CELL The Ni/Cd cell is one of three possible alkaline cells. The three are: 1. Nickel Cadmium (Ni/Cd). 2. Nickel Iron (Ni/Fe). 3. Silver Zinc. Of the three, the Ni/Cd cell has become that preferred for use in aircraft batteries. Figure 8 shows the construction of a Ni/Cd battery cell. CELLOPHANE NYLON NYLON

PLATE

NICKEL MESH

Ni/Cd Battery Cell Construction Figure 8

Page 3-12




The case is made of a plastic/nylon material, which allows for slight expansion of the cell when fully charged. It acts as an insulator between cells and is impervious to electrolyte. The electrolyte in a Ni/Cd cell is an alkaline: Potassium Hydroxide, and may be topped up with distilled or de-ionized water. The relative density is normally between 1.240 and 1.300 depending on the manufacturer's instructions. The plates are made from wire screens sintered with nickel powder. They are impregnated with the active plate material. 1. Positive Plate - Nickel. 2. Negative Plate - Cadmium.

3.4.1 Chemical Action During charging, the negative plates lose oxygen and become metallic cadmium. The positive plates are brought to a higher state of oxidation by the charging current until both materials are completely converted; i.e. all the oxygen is driven out of the negative plates and only cadmium remains, the positive plates pick up the oxygen to form nickel oxides. The cell emits gas towards the end of the charging process, and during overcharging; the gas being caused by decomposition of the water component of the electrolyte into hydrogen at the negative plates and oxygen at the positive plates. A slight amount of gassing is necessary to completely charge the cell and so it therefore loses a certain amount of water. The reverse chemical action takes place during discharging, the negative plates gradually gaining back the oxygen as the positive plates lose it. Due to this interchange there is no gassing on a normal discharge. In this way, the chemical energy, and the electrolyte is absorbed by the plates to a point where it is not visible from the top of the cell. The electrolyte does not play an active part in the chemical reaction; it is used only to provide a path for current flow.


Page 3-13


The chemical reaction of a nickel-cadmium cell is summarized in Table 2.

Battery Type

NickelCadmium

State of charge

Positive Plate

Negative Plate

Electrolyte

Charged

Ni2O2 & Ni2O3 (Nickel Oxides)

Cd (Cadmium)

KOH (Potassium hydroxide) unaffected by state of charge

Ni(OH)2 (Nickel Hydroxide)

Cd(OH)2 (Cadmium Hydroxide)

KOH (Potassium hydroxide) unaffected by state of charge

NickelCadmium Discharged

Table 2 Fully charged cell voltage

=

1.55 volts (Approx). 1.2 volts (Nominal).

Discharged cell voltage

=

1.1 volts.

Relative Density (RD) of electrolyte: 1.24 – 1.30.

Page 3-14



MODULE 11.6


ELECTRICAL POWER

3.4.2 Nickel Cadmium Batteries

BATTERY CASE

Ni/Cad CELL

VENT PIPE

71°C THERMOSTAT (RED TOP)

57°C THERMOSTAT (BLACK TOP)

NUT

THERMOSTAT CONNECTOR

O-RING

LID ASSEMBLY WASHER

PINS C - D RED 71°C

PINS A - B BLACK TOP 57°C

HOLD DOWN BAR

NON-RETURN VALVE

STRAP

RUBBER GASKET

Figure 9 shows the construction of a Ni/Cd battery. Note the inclusion of a thermostat for warning of a 'thermal runaway'.

Ni/Cd Battery Construction Figure 9 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-15


3.4.3 Thermal Runaway Thermal runaway, perhaps more appropriately termed overcharge runaway, is a condition of overcharge instability. It occurs in the later part of the charge cycle. In a normal charge cycle, the heat generated by the charging current is dissipated within the battery and its temperature does not rise appreciably. As the Ni/Cad cell reaches its charged state, higher gassing takes place. All power sources, including batteries, have 'internal resistance'. If the cell temperature were allowed to rise higher, the internal resistance and the terminal voltage would fall. If the internal resistance falls, the charging current will increase, which in turn causes more heat. This chain reaction builds up rapidly and leads to the destruction of the gas barrier, then the cell and finally a fire or even an explosion. So thermal runaway takes place very rapidly and is a danger to aircraft. 3.4.4 Causes 1. Some of the causes of thermal runaway are: 2. Aircraft battery location, poor ventilation. 3. Higher than normal charging current. 4. Frequent or lengthy engine starts. (Electric starter). 5. Loose cell connection. 6. Low electrolyte. 7. Damaged gas barrier. 8. Unbalanced cells. Although Ni/Cd batteries are more susceptible to thermal runaway, the process can occur in Lead/Acid batteries.

Page 3-16




3.5 NI/CD BATTERY CHARGING 3.5.1 Cell Caps All batteries give off gas during charging. The cell caps of a Lead/Acid battery are open and the cell can vent at all times. In a Ni/Cd cell however, the cap is 'semi-open'. It is fitted with a non-return valve to allow gas to vent but not allow air to enter. This is because carbon dioxide in the atmosphere contaminates the electrolyte and reduces its RD. Semi-open caps are susceptible to being blocked by potassium crystals and cell gassing increases during charge. To prevent damage therefore, the caps are removed during the charge and may be cleaned with warm water and then rinsed in de-mineralized or distilled water. 3.5.2 Voltmeter Voltages are critical in the servicing of Ni/Cd batteries. The voltages we are required to measure are to two decimal places (1.24V, 1.55V, and 0.04V). To achieve this accuracy a Digital Voltmeter must be used. 3.5.3 Cell Shorting Links If a single cell in a battery discharges to 0.0 volts and current is still flowing, the cell will have a reversed charge and reversed polarity. This reversal, if continued, can damage the cell. A preventative measure is to short the cell when it is discharged, or nearly discharged, with a shorting link. The clip connectors must be firm. Sometimes a 1 ohm 2 watt resistor is used to compensate for internal resistance lost in shorting out the cell.


Page 3-17


3.5.4 Battery Characteristics The discharge and charge characteristics of a Ni/Cd battery are shown at Figure 10 and Figure 11 respectively. The example is from a 24V battery rated at 36 AH at the one hour rate and 100% capacity.

C E L L V O L T 1.0 V A G E

DISCHARGED 72 amps

½h

36 amps

1h

18 amps

1½h

2h

TIME

Ni/Cad Discharge Curve Figure 10 Note: 1. The voltage falls on initial discharge then remains almost constant. Finally it falls more rapidly. 2. If the discharge current is more than the 36 amps, the voltage falls more sharply and we do not get a 100%, while if the discharge current is less than 36 amps, we obtain more than the battery rated output.

Page 3-18



C E L L V O L T A G E


IF THE CHARGE CONTINUES EXCESSIVE GASSING OCCURS AND THE ELECTROLYTE RUNS OUT

VOLTAGE INITIALLY RISES THEN SETTLES DOWN TO A STEADY RISE

AT THE SECOND RISE THE CELLS START TO GAS

TIME

Ni/Cd Charge Curve Figure 11 3.5.5 Battery Characteristics The charge curve at Figure 9 shows: 1. The voltage initially rises then settles down to a steady rise. This is followed by a second rise where it levels off at the fully charged condition. 2. At the second rise in voltage, the cells start to gas and at the fully charged state, the gassing becomes livelier. If the charge is continued, excessive gassing takes place and the electrolyte will flow out of the cell. This loss of electrolyte can lead to overheating and thermal runaway. 3.5.6 Preparation For Charge Assuming no defects on the battery and satisfactory physical condition, its state of charge must be determined and the balance of cells confirmed. This is done by discharging the battery. The results from this discharge will determine the next step we must take to electrically recondition the battery. The battery is discharged at a constant current stated in the manufacturer's instructions. An example is a 36 AH battery which is discharged at 30 amps. MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-19


During this discharge, the time is recorded in hours, from the start of the discharge and at each stage. The individual cell voltages are also recorded. The battery voltage and each cell voltage are monitored periodically to ensure the discharge remains constant. Initially, the first cell to reach 1.0 volt is looked for then the discharge is continued until the battery voltage is equal to an average of 1.0 volt per cell. A 24V Ni/Cd battery has 20 cells. At 1.0 volt per cell, battery voltage will equal 20 volts. By discharging the battery to its minimum capacity and timing the discharge, the capacity of the battery has been ascertained. Also, by discharging to a known point, the amount of charge required is found and the risk of overcharging is minimized. 3.5.7 Constant Current Charging Constant current charging is always used in the battery workshop. Its advantage over constant potential (voltage) charging is that constant current charging maintains cell balance and capacity. Note: Always follow the manufacturer's instructions for charging. 3.5.8 Charging Rate Consider a battery rated at 24V, 40AH at the 1hr rate. The charge current is expressed in multiples of 'C' amps. Therefore, when we say the charge rate is: 0.1C = 4 amps

0.5C = 20 amps, etc.

3.5.9 Action Prior To Charge 1.

Battery cover removed, (we require to measure the cell voltages).

2.

Vent caps released but not removed from vent.

3.

Check electrolyte is above the plates. If below the plates, high temperature and damaged gas barrier will be caused. Note: this is not looking for a set level.

4.

Remove cell shorting links.

Page 3-20




3.6 METHODS OF CONSTANT CHARGING Two basic methods of charging can be used. 3.6.1 Method 1 Charge at 0.1C amps until battery voltage reaches an average of 1.5 volts per cell (20 cells = 30 volts), then continue the charge for a further four hours. The time for this charge should be between fourteen and fifteen hours. Figure 12 shows charge method 1.

T E R M I N A L

CHARGE AT 0.1C AMPS UNTIL BATTERY VOLTAGE REACHES AN AVERAGE OF 1.5 PER CELL (20 CELLS = 30 VOLTS)

30V V O L T A G E

CONTINUES THE CHARGE FOR A FURTHER 4 HOURS TOTAL TIME FOR CHARGE 14 - 15 HOURS

4 HOURS

10 - 11 HOURS

TIME

Charge Method 1 Figure 12


Page 3-21


3.6.2 Method 2 Charge at 0.5C amps for two hours. Battery voltage should have reached an average of 1.55 volts per cell (20 cells = 31 volts). If it is not up to 31 volts, then charge for a further half-hour at this rate. Continue the charge at 0.1C amps for a further four hours. Figure 13 shows charge method 2.

T E R M I N A L

BATTERY VOLTAGE SHOULD HAVE REACHED AN AVERAGE OF 1.55 V PER CELL (20 CELLS - 31 VOLTS)

31V

V O L T A G E

IF NOT UPTO 31 VOLTS THEN CHARGE FOR A FURTHER HALF HOUR ONLY AT 0.5 C AMPS

CONTINUE THE CHARGE AT 0.1 C AMPS FOR A FURTHER FOUR HOURS CHARGE AT 0.5C AMPS

2 HOURS

0.5 HOURS

4 HOURS

TIME

Charge Method 2 Figure 13

Page 3-22




3.6.3 Completion Of Charge During the last thirty minutes of charge, the electrolyte level should be adjusted. Remember that the level rises during charge and the cells will be gassing more freely at this stage. At completion of charge the cells' voltages should be between 1.5 and 1.7 volts. When the charge is complete, tighten and torque-load the cell vent caps. Refit the battery cover. 3.7 BATTERY TESTING 3.7.1 Capacity Test Start the capacity test with a fully charged battery. This means carrying out the initial discharge and then charging the battery. After charge, stand for fifteen to twenty four hours. Discharge at 1.0C amps 36 AH = 36 amps and record the time. Stop the discharge when battery voltage reaches an average of 1.0 volts per cell. The minimum acceptable capacity is 80% of the stated capacity, or the authorized capacity. 3.7.2 Capacity Test Note: As in the initial discharge, where we looked for the first cell to 1.0 volt, we can do the same in the capacity test. It can indicate an out of balance cell, if that cell voltage falls faster than the rest. The capacity test is carried out at the periods stated in the maintenance manual. This is normally no longer than three months. If the battery fails to achieve 80% capacity, it is not immediately rejected. A cell balancing procedure is carried out to restore the capacity and then a further capacity test is carried out.


Page 3-23


3.7.3 Capacity Recycling This helps to prevent premature damage and failure by cycling the battery, discharging it and then charging it. 1.

Discharge at the required rate, 1.0C amps or less. As each cell approaches, or is zero volts, connect a shorting link across the cell.

2.

When all cells are discharged, stand the battery, with shorting links on, for sixteen to twenty four hours.

3.

The shorting links should then be removed and the battery charged at the recommended recycling charge rate for twenty four hours.

4.

After the first five minutes of charge, any cell over 1.5 volts requires the addition of water, distilled or de-mineralized.

5.

After the first ten minutes of charge, cell voltages must be between 1.20 and 1.55 volts. If any cell is below or above these voltages, it must be rejected and replaced.

6.

After twenty hours of charge, record each cell voltage and adjust the electrolyte level.

After twenty four hours of charge, record each cell voltage. These readings must not be below the twenty hour readiness. Any cell below by more than 0.04 volts must be rejected and replaced. 3.7.4 Deep Discharge This is a term sometimes used. It is when the battery is discharged and all the available capacity is removed.

Page 3-24




3.7.5 Cell Balancing This is similar to the capacity recycling but uses a normal charge, not the long twenty four-hour charge. This is used if the battery fails to give 80% capacity, or the cells are out of balance. 1.

Discharge the battery at 1.0C amps until battery voltage falls to 20 volts and then stop the discharge.

2.

Monitor cell voltages prior to the battery reaching the 20 volts. If any cell falls to zero volts, or goes reverse, then stop the discharge. Zero voltage indicates a shorted cell and reverse indicates a weak cell.

3.

Continue the discharge at 0.1C amps.

4.

Record the time each cell falls to 1.0 volts.

5.

As each cell approaches, or is zero volts, connect shorting link

6.

When all cells are discharged, stand battery, with shorting links on, for sixteen hours.

7.

Charge battery.

Repeat the 'cell balancing'. The time that each cell takes to fall to 1.0 volt should have improved: i.e. longer time recorded. Reject any weak or shorted cells. Note: As already discussed, a battery must be able to deliver at least 80% of its rated capacity to be suitable for aircraft use. Ni/Cd batteries are discharged then charged before a capacity test to determine their efficiency. With certain types of cell, this figure of 80% does not apply. These cells and their rated capacity are as follows: After discharge – Charge – Capacity Test carried out, the capacity of the cells must be at least:

White Ni/Cd cell -

85%

Blue Ni/Cd cell

100%


-

Page 3-25


3.7.6 Voltage Recovery Check The purpose of this is to detect high resistance connections and short circuits inside the cells. 1.

Deep discharge the battery, stand for sixteen to seventeen hours with shorting links on.

2.

Remove shorting links.

3.

Stand the battery for a further twenty four hours, without the shorting links on.

4.

Measure the cell voltages, they should have recovered to above 1.08 volts. Below this voltage indicates a high internal resistance, or an open circuit inside the cell.

3.7.7 Storage Ni/Cd batteries should be stored in a clean, dry, well ventilated room and separate from L/A batteries. 3.7.8 Ready For Service Store in a charged condition. Ni/Cd batteries will self discharge if left standing, therefore a trickle charge is required, (approximately 1 mA per AH). 3.7.9 Long Term Deep discharge and store with main terminals shorted. 3.7.10 Facts And Figures Batteries, L/A or Ni/Cd, proved to have less than 80% capacity should be rejected for aircraft use. During emergency use, e.g. main power failure, batteries must be able to sustain essential services for at least 30 mins.

Page 3-26



MODULE 11.6


ELECTRICAL POWER

3.7.11 Aircraft Charging Systems Aircraft battery chargers generally charge in a constant current mode, or constant voltage mode. Note: The figures below refer to a Ni/Cd battery installation. 3.7.12 Constant Current Mode Supplies a constant current, (an example is 38 amps but this value depends upon the capacity size of the battery fitted to the aircraft). As the battery starts to charge, its terminal voltage will rise at a predetermined value (approximately 31 to 32 volts, giving 1.55v per cell). The constant current mode is switched off and the constant voltage mode switched on. On some aircraft, the battery temperature is monitored and the voltage cut off point of the constant current mode is reduced with a corresponding rise in battery temperature. 3.7.13 Constant Voltage Mode In this mode, the charger holds a constant voltage, or TR mode (as in TRU), at 27.5v or 28v. As the battery terminal voltage is higher than the charger voltage, no charging current will flow. Figures 14 and 13 illustrate typical battery charger charging cycles.

CHARGER VOLTAGE

31.5 V

CONSTANT CURRENT MODE

CONSTANT VOLTAGE MODE

27.5

POWER FIRST APPLIED

TIME

Battery Charging Cycle Figure 14


Page 3-27


Note: When electrical power is first applied to the charger, it goes into the constant current mode and the battery is being charged. CHARGER VOLTAGE BATTERY VOLTAGE

V O L T A G E

INITIAL CONSTANT MODE

CONSTANT VOLTAGE MODE

CONSTANT CURRENT MODE

TIME

Four Charging Cycles Figure 15 It can be seen that the initial constant current mode is independent of battery voltage. The second constant current mode is slightly shorter; as the battery becomes charged, the charge cycles get shorter. These cycles are monitored when the battery charging current starts to flow in the constant voltage mode. After the four charging cycles, the voltage is held at 27.5v, unless the power supply is interrupted, in which case it starts all over again. 3.7.14 Charger Isolation There are certain times when we want to switch the charger off automatically and have no DC output. These are: 1.

If the battery temperature is rising to a dangerous level (towards thermal runaway),

2.

If the battery to be used for excessively high discharge currents, for example engine starting, that would pull the battery voltage down.

On some aircraft, flight deck indication of battery over-temperature is provided for the pilot who takes the appropriate action. Page 3-28




3.8 TYPICAL AIRCRAFT BATTERY SYSTEM Figure 16 shows the circuit arrangement for a battery system from a turboprop aircraft. The circuit serves as a general guide to the methods adopted. Four batteries, in parallel are directly connected to a battery busbar which, in the event of an emergency, supplies power for a limited period to essential services (Radio, Fire Warning Systems, Navigation Systems etc.) Direct connections are made to ensure that the battery power is available to the busbar at all times. The batteries also require to be connected to ensure that they are maintained in a charged condition. In figure 14, this accomplished by connecting the batteries to the main D.C. busbar via a battery relay, power selector switch and a reverse current circuit breaker. Under normal operating conditions of the D.C. supply system, the power selector switch is set to the “Battery” position (This is normally termed “Flight” on modern aircraft). With the switch in this position, current will flow from the batteries through the battery relay coil, the switch and then to ground via the reverse current circuit breaker contacts. The battery relay coil energises connecting the batteries to the main D.C. busbar via the reverse relay coil and its second set of contacts. The aircraft‟s generators supply the main D.C. busbar and so the batteries will also be supplied with charging current form this source. Under emergency conditions (e.g. generator/busbar failure) the batteries must be isolated from the main busbar since their total capacity is not sufficient to keep all aircraft services in operation. The power selector switch must therefore be put to the “OFF” position, thus de-energising the battery relay. The batteries then supply the essential services for the time period pre-calculated on the basis of the battery capacity and current consumption of the essential services. The reverse current coil will reverse its polarity when the battery current flows up to the DC Main busbar (in the event of a failure of the Main DC busbar). This reverse in polarity will cause the reverse current circuit breaker to open, thus isolating the battery from the Main DC busbar.


Page 3-29

Page 3-30

EXT

EXT PWR

OFF

MAIN D.C. BUS-BAR

BATT

REVERSE CURRENT (C/B)

TO GENERATOR SYSTEMS AND ALL D.C. SERVICES

MAIN BUS

OFF

BATTERIES

BATTERY RELAY

VOLTMETER

BATTERY BUS

BATTERY BUS-BAR

TO ESSENTIAL SERVICES



Typical Battery System Figure 16



MODULE 11.6


ELECTRICAL POWER

3.8.1 Parallel/Series Batteries

BATTERY SWITCHING RELAY

BATTERY 2

EXT

2a 3a

2b 3b

1a

1b

BATTERY 1

BATT EXT

OFF

BATTERY RELAY 1

BATTERY BUSBAR

BATT

TO ENGINE STARTING SYSTEM

OFF

BATTERY RELAY 2

The battery system on certain types of turboprop aircraft are designed so that the batteries may be switched from a parallel configuration to a series configuration for the purpose of engine starting from battery power. The circuit arrangement of this type of system using two 24-volt Nickel-cadmium batteries is shown in simplified form in Figure 17.

Parallel/Series Battery System Figure 17


Page 3-31


With reference to figure 15; under normal parallel operating conditions, battery 1 is connected to the battery busbar via its own battery relay and also contacts 1a – 1b of a battery switching relay. Battery 2 is directly connected to the busbar via its relay. When its necessary to use the batteries for starting an engine (i.e. an internal start). Both batteries are connected to the battery busbar in the normal way and 24V supply is fed to the engine starting system via the battery busbar. Closing the starter switch energises the corresponding starter relay, and at the same time the 24V supply is fed via the starting circuit, to the coil of the battery switching relay energising it. Contacts 1a – 1b are now opened to interrupt the direct connection between battery 1 and the busbar. Contacts 3 a – 3b are also opened to interupt the grounded side of the battery 2. Since contacts 2 a – 2b of the switching relay are closed they connect both batteries in series so that 48V is supplied to the busbar and the starter motor. After the engine has started and reached self-sustaining speed, the starter relay automatically de-energises and the battery switching relay coil circuit is interrupted to return the batteries to their normal parallel configuration. Note: The power selector switches are left in the “Battery” position so that when the generators are switched onto the battery busbar, charging current will flow to the batteries. 3.8.2 Aircraft Battery Charger Units In most modern transport aircraft, the battery system incorporates a separate unit for maintaining the batteries in a state of charge. These units also provide some method of sensing the temperature of the batteries during the charging cycle and will automatically isolate the charging unit whenever an over temperature is sensed. The circuits of “On-Board” charger units vary between aircraft types. The following explains the operation of the charger unit fitted to the McDonnell Douglas DC-10.

Page 3-32


A 115V 3 A.C. B MAIN C

3 REVERSE CURRENT C/B

DIRECT FROM BATTERIES


D.C.

A3

B2

A2

B1

PWR SUPPLY INTERUPT

TEMP SENSOR CURRENT

TEMP CONTROL REF VOLTS

B2

D.C. FROM T.R.U.

TEMP SENSOR

LOGIC CCT

CHARGING CURRENT

PWR SUPPLY MONITOR

SCR SW CCT

SCR’S

A3 SENSING RELAYS

A2

B1

TEMPERATURE CUT-OFF

REGULATOR PWR SUPPLY

D.C.

BATTERY SWITCH

BATT

BATTERY SWITCH

D.C. FROM T.R.U.

B1

A1

D.C. RELAY

B2

OFF

A2

A.C. RELAY

TO BATT BUS

1  A.C.


TEMP SENSOR

TRANSFORMER RECTIFIER UNIT

OFF

ON

EMERGENCY POWER SWITCH


MODULE 11.6

ELECTRICAL POWER

Figure 16 shows the charger circuit for the DC-10.

DC – 10 Charger Unit Circuit Figure 16

Page 3-33


3.8.3 DC-10 Charger Unit On the DC-10, the D.C. systems operate from a 28V supply, this is achieved by connecting two 14V batteries in series. Under normal conditions the D.C. busbar is fed via two “Transformer Rectifiers Units” (TRU) from the main A.C. generating system. If the normal D.C. power is not available, then the batteries are automatically connected to the D.C. busbar via the “Charger/Battery” relay and “Sensing” realys. When the batteries are supplying the D.C. busbars the charger is isolated form the batteries. 3.8.4 Operation When power is available from the main generating system, D.C. is supplied to the battery busbar from the TRUs, which also feed the coils of the sensing relays. With these relays energised, the circuit through contacts A2 – A3 is interrupted, while the contacts B1 – B2 are made. The battery switch, which controls the operation of the Charger/Battery relay, is closed to the ”Batt” position when the main electrical; power is available, and the emergency power switch is closed in the “OFF” position. The charger/Battery relay is of the dual type, one relay being A.C. operated and the other D.C. operated. The A.C. relay coil is supplied with power from 1 phase of the main three-phase supply to the battery charger and is energised via contacts B1 – B2 of the sensing relays, the battery switch and the emergency switch. When energised, the contact A1 – A2 close to connect the D.C. positive output from the battery charger to the batteries, thus supplying them with charging current. In the event of a main power failure, the battery charger will become inoperative, the A.C. charger relay will de-energise to the centre off position, and the two sensing relays will also de-energise, thereby opening contacts B1 – B2 and closing contacts A2 – A3. The closing of contacts A2 – A3 now permits a positive supply to flow direct from the battery to the coil of the D.C. battery relay which on being energised also actuates the A.C. relay thereby closing contacts B1 – B2 which connect the batteries directly to the battery busbar. The function of the battery relay contacts is to connect a supply from the battery busbar to the relays of an emergency warning light circuit. 3.8.5 Charging Unit The charging unit converts the main three-phase supply 115/200volts A.C. into controlled D.C. output of a constant current and voltage. This is achieved using a transformer and a full-wave rectifying bridge circuit made up of silicon rectifiers and silicon controlled rectifiers SCRs. The charging current is limited to approximately 65A, this is controlled via current/voltage monitoring circuits and temperature sensing elements within the batteries. The output if these circuits are fed to a logic circuit which in turn controls the operation of the SCRs, thus controlling the charging current.

Page 3-34



MODULE 11.6


ELECTRICAL POWER

3.8.6 Boeing 373 Charger System

BATTERY CHARGER

BATTERY CHARGER TRANSFER RELAY

BUSBAR GRND SERV

GROUND SERVICE RELAY

EXTERNAL A.C. BUSBAR

APU START INTERLOCK RELAY

TRANSFER RELAY

NORM

No 1 GENERATOR BUSBAR

ALT

No 1 TRANSFER BUSBAR

TRANSFER SIGNAL 28V D.C. (No 2 GENERATOR CONTROL UNIT)

No 2 MAIN BUSBAR

No 2 GENERATOR BUSBAR

The charger operates from the aircraft‟s 115V 3-phase A.C. power supplied from a “Ground Service” busbar, which in turn is normally powered by the No 1 generator system, or the external power source. This ensures that the aircraft‟s battery is maintained in a state of charge both in flight and on the ground. Figure 17 shows the charger‟s A.C. input circuit.

Battery Charger A.C. Input circuit Figure 17 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-35


BATTERY BUSBAR “HOT” BATTERY BUABAR

BATTERY THERMAL SWITCH BATTERY CHARGER

MODE CONTROL RELAY 115V 400Hz 3

TRU

HIGH

LOW

FROM MAIN TRU

BATTERY TRANSFER RELAY BATTERY BUSBAR RELAY

EXT POWER RELAY

METERING SHUNT

EXT A.C. BUSBAR

BATTERY SWITCH

115V 400Hz EXT POWER BUSBAR

In flight the A.C. supply is routed to the charger throughh the relaxed contacts of a battery charger transfer relay and an APU start interlock relay. This interlock ensures the charger is inhibited when the APU is starting. Figure 18 shows the battery charger control circuit.

Battery Charger Control Circuit Figure 18 Page 3-36




The D.C. supply for battery charging is obtained from a TRU within the charger unit, which will maintain the battery cell voltage levels in two modes of operation: 1.

High Charge.

2.

Low Charge.

Under normal operating conditions of the aircraft‟s power generation system, the charging level is in the high mode since the mode switch is energised by a rectified output through the batteries thermal switch, and the relaxed contacts of both the battery bus relay and the external power select relay. The charger operates firstly in the high mode, providing an unregulated supply to the battery until the battery voltage rises above that of the charger. The charger current then falls to zero until the battery voltage falls below that of the charger, at which time the charger provides the battery with a pulsed charge and the process is repeated. This pulsing continues until the control circuits within the charger change the operation to the low mode, approximately 2 minutes after pulse charging commenced. In the event that the number 1 generator supply fails, there will be a loss of A.C. power to the ground service busbar, and therefore, to the battery charger. However, with the number 2 generator still on line, a transfer signal from the number 2 generator control unit is automatically supplied to the coil of the battery charger transfer relay, it‟s contacts change over to the connect the charger to the number 2 A.C. supply. Thus the battery charger‟s operation is not interrupted. The APU start interlock relay is connected in parallel with a relay in the starting circuit of the APU, and is only energised during initial stage of starting the APU engine. This prevents the starter motor from drawing part of its heavy starting current through the battery charger. The interlock relay releases automatically when the APU engine reaches 35% rev/min. In addition to the control relay within the battery charger, there are three other ways in which the charging mode can be controlled, each of them fulfilling a protective role by interupting the ground circuit to the mode control relay and so establishing a low mode of charge. They are: 1.

Opening of the battery thermal switch in the event of the battery temperature exceeding 46ºC.

2.

The loss of D.C. power from the designated TRU, causing the battery transfer relay to relax and the battery bus relay to energise.

3.

Energising the fuelling panel power select relay when external A.C. power is connected to the aircraft


Page 3-37


3.9 DC POWER GENERATION The majority of today's aircraft are equipped with ac generation systems. However, dc generation systems are still in use and this section gives an overview of these. 3.9.1 DC Generator Figure 19 shows the construction of a D.C. generator.

GENERATOR END HOUSING

FIELD WINDING

FRAME DRIVE SHAFT

LAMINATED ARMATURE

POLE PIECES

D.C. Generator Figure 19

Page 3-38




3.9.2 Fixed Winding Arrangement Figure 20 shows the arrangement of the fixed windings of a basic four-pole machine suitable for use as a self-excited generator.

COMMUTATOR POLE

TERMINAL BOX YOKE

Z A A1 Z1 FIELD WINDING

BRUSH

Fixed Winding arrangement Figure 20 The fixed portion of the armature circuit consists of four brushes, the links connecting together brushes of like polarity and the cables connecting the linked brushes to the terminals A and A1. The four field coils are of a high resistance and connected in series to form the filed windings. They are wound and connected in such a way as to produce alternate North and South polarities. The ends of the windings are brought out to the terminals indicated Z and Z1.


Page 3-39


3.10 VOLTAGE REGULATION The efficient operation of an aircraft‟s electrical equipment requiring D.C. depends on the fundamental requirement that the generator voltage at the distribution busbar system be maintained constant under all conditions of load and at varying speeds, within the limits of a prescribed range. It is necessary, therefore, to provide a device that will regulate the output voltage of a generator at the designed value and within a specified tolerance. There are a number of factors which, either separately or in combination, affect the output voltage of a D.C. generator, and of these the one which can be most effectively be controlled is the “Field Circuit” current, which in turn controls the flux density. 3.10.1 Vibrating Contact Type Regulator Figure 21 shows a vibrating contact regulator circuit.

TO DISTRIBUTION SYSTEM

VOLTAGE REGULATOR

CURRENT REGULATOR

SHUNT WINDING

SERIES WINDING

GEN

RESISTOR

Vibrating Contact Regulator Circuit Figure 21

Page 3-40




The regulator consists of two armatures (1 for current regulation, 1 for voltage regulation). The voltage regulator consists of windings assembled on a common core. The shunt winding consists of many turns of fine gauge wire and is connected in series with the current regulator winding and in parallel with the generator. The series winding consists of a few turns of heavy gauge wire and is connected in series with the generator‟s shunt field winding. When the generator is operating, the contacts of both regulators are closed so that a positive supply flows through the generator field winding providing the necessary excitation for raising the generator output. At the same time current passes through the shunt winding of the voltage regulator, which will increase the electromagnetic field. As soon as the generator output voltage reaches the preadjusted regulator setting, the electromagnetic field becomes strong enough to oppose the tension of the armature spring and opens the contacts. The circuit in the series winding is opened causing the field to collapse, at the same time the supply to the generator field passes through the resistor, reducing the field current which will cause the generator output to reduce. This reduced output in turn reduces the electromagnetic field strength of the regulator causing the spring tension to close the contact to restore the generator output voltage to its regulated value. The operation is then repeated to maintain the correct voltage output. The frequency of operation dependant on the electrical load carried by the generator, typically between 50 to 200 times a second. 3.10.2 Carbon Pile Voltage Regulator The carbon pile voltage regulator derives its name from the fact that the regulating element (variable resistance) consists of a stack or pile of carbon disks. The disks are contained in a ceramic tube with a carbon or metal contact plug at each end. At one end of the pile, a number of radially arranged leaf springs exert pressure against the contact plug, thus keeping the disks pressed firmly together. For as long as the disks are compressed, the resistance of the pile is very low. If the pressure on the carbon pile is reduced, the resistance increases. By placing an electromagnetic in a position where it will release the spring pressure on the disks as the voltage rises above a predetermined value, a stable and efficient voltage regulator is obtained. The carbon pile regulator is connected in a generator system in the field circuit and an electromagnet to control the resistance. The carbon pile is in series with the generator field and voltage coil is shunted across the generator output. A small manually operated rheostat is connected in series with the voltage coil to provide a limited amount of adjustment. This is necessary where two or more generators are connected in parallel to the same electrical system.


Page 3-41


REGULATED SETTING

LOAD

Figure 22 shows a carbon pile voltage regulator circuit.

VOLTAGE COIL

PILE PRESSURE

CARBON PILE

VOLTAGE OUTPUT

GEN

RPM

MAX

F I E L D

PILE RESISTANCE VOLTAGE OUTPUT

Carbon Pile Voltage Regulator Figure 22 Page 3-42



MODULE 11.6


ELECTRICAL POWER

3.10.3 Transistorised Voltage Regulation Vibrating contact and carbon pile regulators are mainly used on light aircraft: larger aircraft have transistorized voltage regulators (See figure 23). These regulators use a “Zener Diode” to regulate the field current. The two key points to understand with respect to the operation of the transistorized voltage regulator are the “Zener” diode operation and the control transistor. The zener diode can be compared to a relief valve that opens at a given pressure in a hydraulic system. When the zener diode conducts current, it causes the control transistor to switch “ON” which in turn causes the power transistor to switch “OFF”. With the power transistor off, the current flow to the generator field winding is zero. Once the generator output starts to fall, the zener diode will close, switching “OFF” the control transistor which will cause the power transistor to switch “ON” again, thus restoring the current flow to the generator field winding. The operation is then repeated to maintain the correct voltage output. The frequency of operation dependant on the electrical load carried by the generator, typically between 50 to 200 times a second. TR1 R1

POWER

D1 TR2

R2

D2 C1

CONTROL

R3

R4 C2

ZD1

R5

R6

DC GEN

FIELD

TO LOAD

Transistorized Voltage Regulator Figure 23


Page 3-43


3.11 REVERSE CURRENT CUT-OUT RELAY In every system in which the generator is used to charge a battery as well as supply operating power, an automatic means must be provided for disconnecting the generator from the battery when the generator‟s voltage is lower than the battery voltage. If this is not done, the battery will discharge through the generator and may burn out the armature. To prevent this occurring, a “Reverse Current Cutout Relay” is used. Figure 24 shows a reverse current cutout relay circuit.

A+ F+

LOAD +

CURRENT COIL

VOLTAGE REGULATOR

VOLTAGE COIL

F– A–

Reverse Current Cutout Relay Circuit Figure 24 The voltage coil and current coil are both wound on the same soft-iron core. The voltage coil has many turns of fine wire and is connected in parallel with the generator output. The current coil consists of a few turns of large wire connected is series with the generator, thus it carries the entire load current of the generator.

Page 3-44




A pair of heavy contact points is placed where it will be controlled by the magnetic field of the soft iron core. When the generator is not operating, these contacts are held open by the spring. When the generator is operating, and its voltage value is slightly higher than the battery, the voltage coil in the relay magnetizes the soft iron core sufficiently to overcome the spring tension. The magnetic field closes the contact points and connects the generator to the load. As long as the generator voltage remains higher than the battery voltage, the current flow through the current coil will be in a direction that aids the voltage coil to keep the points closed. This means that the field of the current coil will be in the same direction as the magnetic field of the voltage coil and the two will strengthen each other. When the engine turning the generator slows down or stops, the generator voltage will decrease and fall below that of the battery. In this case the battery voltage will cause current to start to flow toward the generator through the relay current coil. When this happens, the current flow will be in a direction that creates a field opposing the field of the relay. This results in the weakening of the total field of the relay, and the contact points are opened by the spring, thus disconnecting the generator and the battery. The tension of the spring controlling the contact points should be adjusted so that the points close at approximately 13.5 V in a 12 V system or 26.6 – 27 V in a 24 V system. 3.12 CURRENT LIMITER In some generator systems a device is installed that will reduce the generator voltage whenever the maximum safe load is exceeded. This device is called a “Current Limiter”. It is designed to protect the generator from loads that will cause it to overheat and eventually burn the insulation and windings. The current limiter operates on a principle similar to that of the vibrator type voltage regulator. Instead of having a voltage coil to regulate the resistance in the field circuit of the generator, the current limiter has a current coil connected in series with the generator load circuit. Figure 25 shows the circuit of a current limiter.


Page 3-45


CURRENT COIL

TO REVERSE CURRENT LIMITER

VOLTAGE REGULATOR

CONTACT POINTS

RESISTOR

Current Limiter Circuit Figure 25 When the load current becomes excessive, the current coil magnetizes the iron core sufficiently to open the contact points, which adds the resistor to the generator field circuit. This causes the generator voltage to decrease, with a corresponding decrease in generator current. Since the magnetism produced by the current limiter coil is proportional to the current flowing through it, the decrease in generator load current also weakens the magnetic field of the current coil and thus permits the contact points to close. This removes the resistor from the generator field circuit and allows the voltage to rise again. If an excessive load remains connected to the generator, the contacts of the current limiter will continue to vibrate, thus holding the current output at or below the minimum safe limit. The contact points are normally set to open when the current flow is 10% above the rated capacity of the generator.

Page 3-46




3.13 THREE UNIT CONTROL PANEL A three unit control panel consists of a: 1.

Voltage Regulator.

2.

Current Limiter.

3.

Reverse Current Cutout Relay.

This combination will provide for both voltage regulation and protection from excessive loads. It has proved very successful for the control of 12 and 24-volt generator systems. Figure 26 shows the circuit for a three-unit generator control panel.

VOLTAGE REGULATOR

CURRENT LIMITER

REVERSE CURRENT LIMITER

A+

F+

A–

Three Unit Generator Control Panel Figure 26


Page 3-47


Figure 27 shows a typical three unit Generator Control Panel.

Typical Three Unit Generator Control Panel Figure 27

Page 3-48




3.14 PARALLEL & LOAD SHARING In a multi-engine aircraft, it is generally desirable that both generators driven by each engine should operate in parallel thereby ensuring that in the event of an engine or generator failure, there is no interruption of the primary power supply. Parallel operation requires generators carry equal shares of the system load, and so their output voltages must be as near equal as possible under all operating conditions. Generators are provided with voltage regulators which independently control the generator‟s voltage output, but as variations in output and electrical loads can occur, it is essential to provide additional regulation circuits, having the function of maintaining balanced outputs and load sharing. The method most commonly used for this purpose is that which utilises a “LoadEqualising” circuit to control generator output via their voltage regulators. This is shown in figure 28.

BUSBAR

LINE CONTACTORS

CP

GEN No 1

FC

VC

CP

VC

EC

FC

EC EQUALISING RELAYS

INTERPOLE COIL

INTERPOLE COIL

Load Sharing (Carbon Pile Regulation) Figure 28 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-49

GEN No 2


Both generators are each feeding 150 amps to a common bus bar. The current is fed to the loads and the 150 amps complete the circuit to each generator through its earth connection. It can be seen that there is a voltage drop across (usually) an interpole of each generator and since the load currents are equal, the voltage drops are also equal. The same potential occurs at each end of the equalizing loop and so no current flows in the equalizing coils. When there is an imbalance of currents between the generators, the voltage drop across the interpole of the generator supplying the largest current is greater than the voltage drop in the other generator. The end of the equalizing loop with the largest voltage drop will be driven more negative, causing equalizing current to flow from the lightly loaded generator to the overloaded generator. Current flow in this direction causes one regulator to increase the output of the associated (lightly loaded) generator. The other regulator causes the associated (over loaded) generator to reduce its output. Figure 29 shows the operation of the equalising circuit when the No 2 generator‟s output is higher than that of the No 1 generator. 300 AMPS BUSBAR

100 AMPS

LINE CONTACTORS

CP

GEN No 1

FC

200 AMPS

VC

CP

VC

EC

FC

GEN No 2

EC EQUALISING RELAYS

100 AMPS

- 0.17V

- 0.34V

Load Sharing – Unbalanced Loads Figure 29 Page 3-50


200 AMPS



FIELD EXCITATION CURRENT

GEN No 1

No 1 VOLTAGE REGULATOR

BUSBAR

PARALLEL RELAY UNIT

GEN No 2

No 2 VOLTAGE REGULATOR

FIELD EXCITATION CURRENT

Paralleling load sharing can also be controlled utilising vibrating contact voltage regulation. Figure 30 shows a load sharing circuit using vibrating contact voltage regulators.

Load Sharing (Vibrating Contact Regulation) Figure 30 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-51


For this method of load sharing the circuit comprises an additional coil “Eq” in the voltage regulation section “A” of each regulator and a paralleling relay unit. When both generators are in operation and supplying the correct regulated outputs voltages, the contacts in the voltage and current (“B”) regulation sections of each regulator are closed. The contacts of the reverse current relays “C” are also closed thereby connecting both generators to the busbar. The outputs from each generator are also supplied to the coils of the paralleling relay unit so the contacts of its relays are closed. The paralleling relays and the equalising coils form the paralleling (equalising) circuit between both generators. Under load-sharing conditions, the current flowing through the coils “Eq” is in the same direction as that of the voltage coils of the voltage regulating sections of each regulator, but in equal and opposite directions at the contacts of the paralleling relay unit. If the voltage output of the No 1 generator should rise, there will be a greater voltage input to the voltage regulating section of the number 1 voltage regulator compared with that of the corresponding section of the number 2 regulator. There will be an unbalanced flow of current through the equalising circuit such that the increase of current through coil “Eq” of the number 1 voltage regulator will now assist the electromagnetic effect of voltage coil “D” causing the relay contacts to open. This connects the resistance into the field circuit of the number 1 generator, reducing its excitation current and its voltage output. Because of the unbalanced condition, the increased current in the equalising circuit will also flow across the paralleling relay unit contacts to the coil “Eq” in the number 2 voltage regulator so that it opposes the electromagnetic effect of its associated coil “D”. In a paralleled generator system utilising solid sate regulation, any unbalanced conditions are detected and adjusted by interconnecting the regulators via two additional paralleling transistors, one in each regulator.

Page 3-52



MODULE 11.6


ELECTRICAL POWER

3.15 AC POWER GENERATION AC generators are used as the primary source of electric power in almost all transport category aircraft. The AC system supplies most of the electrical power required for the aircraft. 3.15.1 Brushless Generators Brushless generators were developed for the purpose of eliminating some of the problems of generators that employ slip-rings and brushes to carry exciter current to the rotating field. The advantages of a brushless generators are: 1. Lower maintenance cost, since there is no brush or slip ring wear. 2. High stability and consistency of output, because variations of resistance and conductivity at the brushes and slip rings are eliminated. 3. Better performance at high altitudes, because arcing at the brushes is eliminated. The brushless generator‟s operation is to use electomagnetic induction to transfer current from the stationary components of the generator to the rotating components and use a three-phase Star connected armature. Figure 31 shows a schematic of a brushless generator. TO AC BUSBAR SYSTEM

TO GCU

T1

T2

T3

FROM GCU

PMG

PERMANENT MAGNET GENERATOR

RECTIFIER

EXCITER GENERATOR

MAIN FIELD

MAIN GENERATOR

Brushless Generator Figure 31 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-53

N


The permanent magnet, which is connected to the rotor, is used to induce an alternating current into the stationary PMG three-phase armature winding. The Generator Control Unit (GCU) rectifies the AC armature current to DC voltage, which is applied to the exciter filed winding. The exciter field induces an AC into the exciter armature. The exciter armature is connected to the rotating rectifier, which changes the AC to DC for the main generator field winding. The main field induces an AC voltage into the main generator stator fields. The stator fields are induced with 115V phase voltage giving the 200V between two phases. 3.15.2 Constant Speed Drive (CSD) Unit When an AC generator system is used as the primary source of power, numerous consumer services are dependent on a “Constant-Frequency”. A constant frequency is inherent in an AC system only if the generator is driven at a constant speed. The aircraft‟s engines cannot be relied on to do this directly and, if a generator is connected directly to the accessory drive of an engine the output frequency will vary with engine speed. Some form of conversion equipment is therefore required and the type most commonly used utilizes a transmission device interposed between the engine and the generator, and which incorporates a variable ratio drive mechanism. Such a mechanism is referred to as a “Constant-Speed Drive” (CSD) unit. Figure 32 shows the basic arrangement of a CSD unit. CHARGE OIL

VARIABLE DISPLACEMENT

FIXED DISPLACEMENT

HYDRAULIC UNIT CONTROL CYLINDER

DIFFERENTIAL UNIT OUTPUT TO GENERATOR

INPUT FROM ENGINE

TO OIL PUMPS AND GOVERNOR GOVERNOR

Constant-Speed Drive Unit Figure 32 Page 3-54




The power used to drive the generator is controlled and transmitted through the combined effects of the units. Oil for the system operation is supplied from a reservoir via charge pumps within the unit, and a governor. 3.15.3 Variable Displacement Unit The variable displacement unit consists of a cylinder block; reciprocating pistons and a variable angle wobble or swash plate, the latter being connected to the piston of a control cylinder. Oil to this control cylinder is supplied from the governor. This unit is driven directly by the input gear and the differential planet gear carrier shaft, so that its cylinder block always rotates (relative to the port plate and wobble plate) at a speed proportional to the input gear speed and always in the same direction. 3.15.4 Control Cylinder When the control cylinder moves the wobble plate to some angular position, the pistons within the cylinder block are moved in and out as the block rotates, and so the charge oil is compressed to a high pressure and then “ported” to the fixed displacement unit. Under these conditions the variable displacement unit functions as a hydraulic pump. 3.15.5 Governor The supply of charge oil to the unit‟s control valve is controlled by a governor valve which is spring biased, flyweight operated and driven by the out gear driving the generator. It therefore responds to changes in transmission output speed. 3.15.6 Fixed Displacement Unit The fixed displacement unit is similar to the variable displacement unit, except that its wobble plate, which has an inclined face, is fixed and has no connection to the control cylinder. When oil is pumped to the fixed displacement unit by the variable displacement unit, it functions as a hydraulic motor and the volume of oil pumped to it determines its direction of rotation and speed. 3.15.7 Differential Gear Unit The differential gear unit consists of a carrier shaft carrying two meshing (1:1 ratio) planet gears, and a gear at each end; one meshing with the input gear and the other with the gear which drives the variable displacement unit cylinder block. The carrier shaft always rotates in the same direction and at a speed which via the input , varies with engine speed. Surrounding the carrier shaft are two separate “housings”, and since they have internal ring gears meshing with the planet gears, then they can rotated differently. Each housing also has an external ring gear; one (input ring gear) meshing with the fixed displacement unit gear, and the other (output ring gear) meshing with the output gear drive to the generator.


Page 3-55


3.15.8 CSD Operation With the CSD in operation, the output ring gear “housing” serves as the continuous drive transmission link between the engine and the generator. Since the input ring gear “housing” is geared to the fixed displacement hydraulic unit, then depending on the direction of rotation of this unit , the “housing” can rotate in the same direction as, or opposite to, that of the carrier shaft and the output ring gear “housing”. In this way, speed is either added to, or subtracted from the engine input speed, and through the gear ratio (2:1) between the ring gears and the carrier shaft planet gears, the output ring gear “housing” rotational speed will be appropriately adjusted to maintain constant governor speed. 3.15.9 Underdrive Phase If the input speed supplied to the transmission exceeds that required to produce the required output speed, the governor, in sensing the speed difference will cause oil to flow away from the control cylinder. In this condition, the transmission is said to be operating in the “Underdrive” phase. Figure 33 shows the CSD operating in Underdrive Phase. VARIABLE UNIT

FIXED UNIT CONTROL CYLINDER

CARRIER SHAFT

D

T EN ER F IF

IAL

OUTPUT RING GEAR

TO GENERATOR

OUTPUT GEAR TO GOVERNOR INPUT GEAR

INPUT RING GEAR

TO PUMPS

FROM ENGINE

Underdrive Phase Figure 33 Page 3-56




The control cylinder changes the angular of the variable displacement unit‟s wobble plate so that the oil is pumped at high pressure to the fixed displacement unit causing it to rotate in the same direction as that of the variable displacement unit. This rotation is transmitted to the input ring gear ”housing”, of the differential unit, so that it will rotate in the same direction as the output ring gear “housing”, and the carrier shaft. Because the input ring gear “housing” is now rotating in the same direction as the carrier shaft then the speed of the freely rotating planet gear meshing with the housing will be reduced. The speed of the second planet gear will also be reduced in direct ratio thereby reducing the speed of the output ring gear “housing”. This hydromechanical process of speed subtraction continues until the required generator drive speed is attained. 3.15.10 Overdrive Phase When the output speed supplied to the transmission is lower than that required to produce the required output speed, the governor causes charge oil to be supplied to the control cylinder. In this condition, the transmission is said to be operating in the “Overdrive” phase. Figure 34 shows the CSD operating in Overdrive Phase.

Overdrive Phase Figure 34


Page 3-57


The control cylinder changes the angular of the variable displacement unit‟s wobble plate so that the oil is pumped at high pressure to the fixed displacement unit causing it to rotate in the opposite direction as that of the variable displacement unit. This rotation is transmitted to the input ring gear ”housing”, of the differential unit, so that it will rotate in the opposite direction as the output ring gear “housing”, and the carrier shaft. Because the input ring gear “housing” is now rotating in the opposite direction as the carrier shaft then the speed of the freely rotating planet gear meshing with the housing will be increased. The speed of the second planet gear will also be increased in direct ratio thereby increasing the speed of the output ring gear “housing”. This hydromechanical process of speed addition continues until the required generator drive speed is attained. Figure 35 shows a typical CSD-Generator Unit.

BRUSHLESS GENERATOR

OIL LEVEL SIGHT GLASS

CONSTANT SPEED DRIVE UNIT

CSD- Generator Unit Figure 35 In multi CSD generator systems the control of the drives is important in order that real electrical loads are evenly distributed between generators. This will be covered at a later point in these notes.

Page 3-58




3.15.11 CSD Disconnection The disconnection of the CSD transmission system following a malfunction may be accomplished mechanically via handles located in the flight crew compartment utilising electro-mechanical principles. Figure 36 shows a typical CSD disconnection system.

RESET SPRING

DOG TOOTH CLUTCH SEPARATION POINT

INPUT SPLINE SHAFT

INPUT SHAFT

THREADED PAWL

SOLENOID

INPUT GEAR

PAWL SPRING

HANDLE SPRING

TRANSMISSION CASE

CSD Disconnection System Figure 36


Page 3-59


The drive from the engine is transmitted to the CSD via a dog-tooth clutch, and disconnect is initially activated by a solenoid controlled from the flight deck. When the solenoid is energised, a spring-loaded pawl moved into contact with the threads of the input shaft, which serves as a screw causing the input shaft to move away from the input spline shaft (driven by the engine) therby seperating the driving dogs of the clutch. Resetting of the disconnect mechanism can only be accomplished on the ground following shut-down of the appropriate engine. This is accomplished by pulling out the reset handle to withdraw the threaded pawl from the input shaft, and allowing the reset spring on the shaft to re-engage the clutch. At the same time, and with the solenoid de-energised, the solenoid nose pin snaps into position in the slot of the pawl. 3.16 INTEGRATED DRIVE GENERATOR (IDG) The IDG is a state-of-the-art means of producing A.C. electrical power. It contains both generator and the CSD unit in one unit. This concept helps to reduce both the weight and the size of the traditional two-unit system. The CSD contains a hydraulic trim unit and a differential assembly, which converts the variable engine rpm to a generator input speed of 12,000 rpm. Figure 37 shows a typical IDG as used on Boeing 757 aircraft.

OUTPUT 115 V 400 Hz

GENERATOR

CONSTANT DRIVE UNIT

Integrated Drive Generator (IDG) Figure 37 This type of unit is capable of producing 90 kVA continuously, 112.5 kVA for a 5minute overload, and 150 kVA for a 5-second overload. The output voltage is 115V A.C. at 400 Hz. Page 3-60




3.17 FIELD EXCITATION The production of a desired output by any type of generator requires a magnetic field to provide excitation of the windings. In D.C. generators, this is achieved in a fairly straightforward manner by residual magnetism in the electromagnet system and by the build up of current through the field windings. The field current is controlled by a voltage regulator system. The excitation of A.C. generators involves the use of more complex circuits, the arrangement of which are essentially varied to suit the particular type of generator and its controlling system. However, they all have one common feature, i.e. the supply of D.C. to the field windings to maintain the desired A.C. output. 3.18 VOLTAGE REGULATION Regulation of the output of a constant-frequency generator system is based on the principle of controlling the field excitation. In installations requiring a multiarrangement of constant frequency generators, additional circuitry is required to control output under load-sharing or paralleling operating conditions; this also involves control of the field excitation. With reference to figure 37, the regulation circuit is comprised of three main sections: 1. Error Detection. 2. Pre-Amplification. 3. Power Amplification. 3.18.1 Operation The function of the error detector circuit is to monitor the generator‟s output voltage, compare it with a fixed reference voltage and to transmit any error to the pre-amplifier. It is made up of a three-phase bridge rectifier connected to the generator output, and a bridge circuit of which two arms contain gas-filled regulator tubes, the two arms containing resistors. The inherent characteristic of the tubes are such that they maintain an essentially constant voltage drop across their connections for a wide range of current through them and for this reason they establish the reference voltage against which the output voltage is continuously compared. The output side of the reference bridge is connected to an “error” control winding of the pre-amplifier and then from this amplifier to a “signal” control winding of the power amplifier. Both the amplifiers are three-phase magnetic amplifiers. The output of the power amplifier is supplied to the shunt windings of the generator‟s A.C. exciter stator.


Page 3-61


GENERATOR

PMG ERROR DETECTOR

REFERENCE

A

SENSING

R1

R2

B

RV1

PRE-AMP

ERROR CONTROL WINDING

SIGNAL CONTROL WINDING

POWER AMP

A.C. EXCITER

EXCITER FIELD

Figure 38 shows the voltage regulation circuit for an AC Brushless Generator.

AC Brushless Generator Voltage Regulation Figure 38 Page 3-62




The output side of the reference bridge is connected to an “error” control winding of the pre-amplifier and then from this amplifier to a “signal” control winding of the power amplifier. Both the amplifiers are three-phase magnetic amplifiers. The output of the power amplifier is supplied to the shunt windings of the generator‟s A.C. exciter stator. The output of the error bridge rectifier is a D.C. voltage slightly lower than the average of the three A.C. line voltages: This voltage may be adjusted via a variable resistor (RV1) to bring the regulator system to a balance condition for any nominal line voltage. A balanced condition of the reference bridge circuit is obtained when the voltage applied across the bridge (points A & B) is exactly twice that of the voltage drop across the two tubes. Since under this condition, the voltage drop across R1 & R2 will equal the drop across each tube, then no current will flow in the output circuit to the error control winding of the pre-amplifier. If the A.C. line voltage should go above or below the fixed value, the voltage drops across R1 & R2 will differ causing an unbalance of the bridge circuit and a flow of current to the “error” control winding of the pre-amplifier. The direction and magnitude of the current flow will depend on whether the error in line voltage, is above (positive error signal) or below (negative error signal) the balanced nominal value. The output from the pre-amplifier to the power amplifier will either be positive or negative. For a positive error, the exciter current will be decreased and for a negative output it will be increased.


Page 3-63


3.19 VARIABLE-SPEED CONSTANT-FREQUENCY POWER SYSTEMS In an effort to simplify and improve the production of ac power for aircraft and to get away from the need for hydro-mechanical constant-speed drives, a number of systems have been devised for producing 400 Hz three-phase electric power through electronic circuitry. This has been made possible by great advances in solid-state technology developed in recent years. Variable-speed constant-frequency systems are referred to as VSCFR, VASCOF, CFG (constant-frequency generator), and ECEPS (electronic convertor, electric power supply). The systems employ a generator whose variable speed and variable-frequency power would not be suitable for power needs in aircraft system but the variable-frequency power is converted to constant-frequency power by means of solid-state circuitry, and this makes the power suitable for aircraft use. The drawing at Figure 38 is a block diagram showing the principal elements of an ECEPS system. The brushless ac generator is similar to those described in module 3 but since it is driven directly by the engine, its speed and output frequency will vary as engine speed varies. The variable three-phase power is fed to the full wave crystal-diode rectifier, where it is converted to direct current and filtered. This direct current is fed to the conversion circuitry, where it is chopped into square-wave outputs that are separated and summed up to produce three-phase 400 Hz alternating current. Variable-speed constant-frequency systems can be designed with separate components, or as an integrated unit. The generator and static inverter are mounted as a unit on the engine. The GCU, which contains the voltage regulator, is mounted in the aircraft.

Page 3-64




NEUTRAL

B PHASE

C PHASE 3 PHASE AC 400 HZ

VOLTAGE REGULATOR

BRUSHLESS AC GENERATOR

FULL-WAVE CRYSTAL-DIODE RECTIFIER & FILTER

3 PHASE AC VARIABLE FREQUENCY

FILTERED DC

CONVERSION CIRCUIT

A PHASE

VOLTAGE SENSING

Figure 39 shows the basic elements of a Variable-Speed, Constant-Frequency Electric Power System

Variable-Speed, Constant-Frequency Electric Power System Figure 39 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-65


3.20 FREQUENCY-WILD SYSTEMS A frequency-wild system is one in which the frequency of its generator voltage output is permitted to vary with the rotational speed of the generator. Although such frequency variations are not suitable for the direct operation of all types of a.c. consumer equipment, the output can (after constant voltage regulation) be applied directly to resistive circuits such as electrical de-icing systems. It can also be transformed and rectified to provide medium – low voltage DC, which in turn could be fed into a static inverter to produce a frequency controlled AC supply. 3.20.1 Frequency-Wild Generator Construction The construction of a typical generator used for the supply of heating current to a turbo-propeller engine de-icing system is shown in Figure 40.

Frequency-Wild generator – Construction Figure 40 Page 3-66




3.20.2 Operation The Frequency-Wild Generator has a three-phase output of 22kVA at 208 volts and supplies full load at this voltage through a frequency range of 280 to 400 Hz. Below 280 Hz the field current is limited and the output relatively reduced. The generator consists of two major assemblies: 1. Fixed stator assembly in which the current is induced. 2. Rotating assembly (rotor). 3.20.3 Stator Assembly The stator assembly is made up of a high permeability laminations and is clamped in a main housing by an end frame having an integral flange for mounting the generator at the corresponding drive outlet of an engine-driven accessory gear-box. The stator winding is star connected, the star or neutral point being made by linking three ends of the winding and connecting it to ground. The other three ends of the winding are brought out to a three-way output terminal box mounted on the end frame of the generator. Three small current transformers are fitted into the terminal box and form part of a protection system. 3.20.4 Rotor Assembly The rotor assembly has six salient poles of laminated construction; their seriesconnected field windings terminate at two slip rings secured at one end of the rotor shaft. Three brushes are equal-spaced on each slip ring and contained within a brush-gear housing which also forms a bearing support for the rotor. The brushes are electrically connected to d.c. input terminals housed in an excitation terminal box mounted above the brush-gear housing. 3.20.5 Generator Cooling The generator is cooled by ram air passing into the main housing via an inlet spout at the slip ring end, the air escaping from the main housing through ventilation slots at the drive end. An air collector ring encloses the slots and is connected to a vent through which the cooling air is finally discharged. Provision is made for the installation of a thermally operated switch to ensure there is overheat protection.


Page 3-67


R O T O R

STATOR

COMPOUNDING RECTIFIER

EXCITATION RECTIFIER

SIGNAL

COMPOUNDING TRANSFORMER

VOLTAGE REGULATOR

DISCONNECT

TO 208V AC BUSBAR AIRCRAFT DE-ICING SYSTEM

28V DC BUSBAR

START SWITCH

Figure 40 shows the circuit for the frequency wild generator shown in figure 41.

Frequency Wild Generator – Schematic Circuit Figure 41 Page 3-68




3.20.6 Frequency Wild Generator Excitation Referring to figure 40, excitation is provided by DC from the aircraft‟s main busbar and by rectified AC. The principle components and sections of the control system associated with excitation are: 1. Control Switch. 2. Voltage Regulation Section. 3. Field Excitation Rectifier. 4. Current Compounding Section. (Consisting of threephase current transformers and rectifier). The primary windings of the compounding transformer are in series with the three phases of the generator and the secondary windings in series with the compounding rectifier. When the control switch is in the “Start” position, DC from the main busbar is supplied to the slip rings and the windings of the generator rotor; thus, with the generator running, a rotating magnetic field is set up to induce an alternating output to the stator. This output is tapped to feed magnetic amplifier type of voltage regulator, which supplies a sensing current signal to the excitation rectifier. When this signal reaches a pre-determined “Off-Load” value, the rectified AC through the rotor winding is sufficient for the generator to become self-excited and independent of the main busbar supply which is then disconnected. The maximum excitation current for wide-speed-range high-output generators is quite high, and the variation in excitation current necessary to control the output under varying “load” conditions is such that the action of the voltage regulator must be supplemented by some other medium of variable excitation current. The compounding transformer and rectifier provide this, and are connected in such a manner as to ensure the excitation current is proportional to load current.


Page 3-69


3.21 THREE PHASE GENERATOR The output terminals of a three-phase generator are marked to show the phase sequence, and these terminals are connected to busbars, which are also identified correspondingly. Figure 42 shows a basic three-phase A.C. generator.

PHASE A STATOR

N

ROTOR

PHASE B PHASE C

N

N

Basic Three-Phase Generator Figure 42

Page 3-70



MODULE 11.6


ELECTRICAL POWER

3.21.1 Interconnection Of Phases Each phase of a three-phase generator may be brought out to separate terminals and used to supply separate groups of consumer services (see figure 38). This is an arrangement rarely encounted in practice since pairs of “Line” wires would be required for each phase and would involve uneconomic use of cable. The phases are therefore interconnected by two different methods, these are: 1. Star connection. 2. Delta connection. 3.21.2 Star Connection The “Star” connection is commonly used in generators. One end of each phase is connected to the “Neutral” point, while the opposite ends of the windings are connected to three separate lines. With this arrangement, two-phase windings are connected between each pair of lines. A star connection arrangement is shown in figure 43.

A A PHASE VOLTAGE

A1

NEUTRAL

A

LINE VOLTAGE = 3 x PHASE

B PHASE VOLTAGE

C

B

B

C1 B1 LINE VOLTAGE =  3 x PHASE C PHASE VOLTAGE

LINE VOLTAGE =  3 x PHASE

C PHASE VOLTAGE = 115V LINE VOLTAGE = 200V Star Connection Figure 43


Page 3-71


3.21.3 Delta Connection In this configuration, the windings are connected in series to form a closed “Mesh” and the lines being connected at the junction points. As only one phase winding is connected between each pair of lines then, in the delta method, phase voltage is always equal to line voltage. Figure 44 shows the Delta method of connection.

LINE VOLTAGE

C

A1

C1

A

E AS GE PH TA L VO

A VO PH L T AS AG E E

L CURRENT = 3 PHASE CURRENT

C

LINE VOLTAGE


B1

B

B PHASE VOLTAGE

LINE VOLTAGE


PHASE CURRENT = 100A LINE CURRENT = 173A

Delta Connection Figure 44

Page 3-72




3.22 BUSBARS In most types of aircraft, the output from the generating sources are coupled to one or more low impedance conductors referred to as “Busbars”. These are usually situated in junction boxes or distribution panels located at central points within the aircraft, and they provide a convenient means for connecting positive supplies to the various consumer circuits. Busbars vary in form dependent on the methods to be adopted in meeting the electrical power requirements of a particular aircraft type. In a very simple system a busbar can take the form of a strip of interlinked terminals while in more complex systems main busbars are thick metal (usually copper) strips or rods to which input and output supplies are connected. The strips or rods are insulated from the main structure and are normally provided with some form of protective covering. 3.22.1 Busbar Systems The function of a distribution system is prmarily a simple one, but it is complicated by having to meet additional requirments which concern a power source, or a power consumer system operating either separately or collectively, under abnormal conditions. The requiremnets and abnormal conditions may be considered in relation to three main areas, which are as follows. 1.

Power-consuming equipment must not be deprived of power in the event of power source failure unless total power demands exceeds the available supply.

2.

Faults on the distribution system (i.e. fault currents, grounding of busbars) should have the minimum effect on the system function, and should constitute minimum possible fire risk.

3.

Power-consuming equipment faults must not endanger the supply of power to other equipment.

These requirements are met in a combined manner by paralleling generators where appropriate, by providing adequate circuit protection devices, and by arranging for faulted generators to be isolated from the distribution system. Most distribution systems are so arranged that they may be fed from a number of different power sources. In adopting this arrangement it is usual to categorise all consumer services into their order of importance. The categories are: 1.

Vital Services (Hot Battery Busbar).

2.

Essential Services.

3.

Non-Essential Services.


Page 3-73


Vital Services; are those which would be required after an emergency wheels-up landing, e.g. emergency lighting and crash switch operation of fire extinguishers. These services are connected directly to the battery. Essential Services; are those required to ensure safe flight in an in-flight emergency situation. They are connected to D.C. and A.C. busbars, as appropriate, and in such a way that they can always be supplied from a generator or from batteries. Non-Essential Services; are those which can be isolated in an in-flight emergency for load shedding purposes, and are connected to D.C. and A.C. busbars, as appropriate, supplied from a generator. Figure 45 shows a typical two D.C. generator distribution system.

GEN No 1

No 2 INVERTER

No 3 INVERTER

GEN No 2

NON-ESSENTIAL A.C. CONSUMERS No 2 BUSBAR

No 1 BUSBAR

NON-ESSENTIAL D.C. CONSUMERS


BATTERY BUSBAR

VITAL D.C. CONSUMERS

No 1 INVERTER

ESSENTIAL A.C. CONSUMERS

ESSENTIAL BUSBAR

ESSENTIAL D.C. CONSUMERS

Two D.C. Generator Distribution System Figure 45

Page 3-74




In figure 44, the power supplies are 28v D.C. from two engine driven generators operating in parallel, 115v A.C. 400 Hz A.C. from rotary inverters, and 28v D.C. from the batteries. Each generator has its own busbar to which are connected the “Non-essential” consumer services. Both busbars are in turn connected to a single busbar, which supplies power to the “Essential” consumer services. With both generators operating, all consumer services are supplied with power. The essential busbar is also connected to the battery busbar ensuring that the batteries are maintained in the charged condition. In the event that one generator should fail it is automatically isolated from its respective busbar and all busbar loads are taken over by the operative generator. Should both generators fail, the non-essential services are no longer provided with power, the batteries automatically supply power to the essential busbar to supply the essential services (A.C. essential services via the No inverter). The batteries will maintain the essential busbar for a period calculated on consumer load requirements and the battery states of charge. 3.22.2 Split Bus-Bar A.C. Generation System The generators supply three-phase power through separate channels, to the main busbar and these in turn supply the non-essential consumer loads and Transformer rectifier Units (TRUs). The essential A.C. loads are supplied from the essential busbar, which under normal operating conditions is connected via a change over relay to the No 1 main busbar. The main busbars are normally isolated from each other (i.e. the generators are not paralleled). If however one of the generators should fail, the busbars are automatically inter-connected by the energising of the “Bus-Tie” breaker, thus maintaining supplies to all A.C. consumers and both TRUs.


Page 3-75

Page 3-76

VITAL D.C. CONSUMERS

BATTERY D.C. BUSBAR

BATTERY RELAY

ESSENTIAL A.C. CONSUMERS

ESSENTIAL A.C. BUSBAR

NON-ESSENTIAL A.C. CONSUMERS CHANGE OVER REALY

ESSENTIAL D.C. CONSUMERS

ESSENTIAL D.C. BUSBAR

STATIC INVERTER

No 1 A.C. BUSBAR

GEN No 1

No 1 TRU

ISOLATION RELAY

BUS TIE BREAKER

No 2 TRU


NON-ESSENTIAL D.C. BUSBAR

NON-ESSENTIAL A.C. CONSUMERS

No 2 A.C. BUSBAR

GEN No 2

EXTERNAL POWER RELAY



Figure 46 shows a schematic of an A.C. Split Busbar system.

A.C. Split Busbar System Figure 46




The supply of D.C. is derived from independent TRUs and the batteries. The No 1 TRU supplies essential loads and the No 2 TRU supplies non-essential loads connected to the main D.C. busbar. Both the main and essential D.C. busbars are automatically interconnected by an isolation relay. In the event that one generator should fail it is automatically isolated from its respective busbar and all busbar loads are taken over by the operative generator. If, for any reason, the power supplied from both generators should fail the nonessential services will be isolated and the change over relay between No 1 main busbar, and the essential A.C. busbar will automatically de-energize and connect the essential A.C. busbar to an emergency static inverter. The main D.C. busbar is isolated from the essential D.C. busbar by the isolation relay de-energizing. The essential D.C. busbar deriving its D.C. supply directly from the battery busbar to maintain the operation of the essential D.C. and A.C. supplies. External power can also be connected to the whole system to supply both A.C. and D.C. to the system.


Page 3-77


3.22.3 Bus-Bar Supply Priority The supply of the automatic ac bus transfer system and the dc bus transfer system are given below at Figures 47.

P R I O R IT Y

1

GEN 1

GEN 1

ESS AC BUS

GEN 2

EXT PWR

2

EXT PWR

GEN 2

E M E R IN V

EXT PWR

A C B US 2

3

GEN 3

GEN 3

-

GEN 3

-

4

GEN 2

EXT PWR

-

GEN 1

-

AC BUS 1

ESS A C B US

E M E R AC B US

AC BUS 2

A C G RN D S E R V B US

* M A N U A L O P E R A T IO N

P R I O R IT Y

1

T RU 1

B AT 1

T RU 3

T RU 3

B AT 2

T RU 2

D C EX T PW R

2

T RU 2 *

-

DC BUS 1

DC BUS 1

-

T RU 1 *

GR N SER V T RU

3

-

-

-

B AT TE R IE S

-

-

-

DC BUS 1

B AT B US 1

ESS D C B US

EM ER DC BUS

DC BUS 2

D C G RN D S E R V B US

B AT B US 2

Bus-bar Supply Priority Figure 47 Page 3-78



MODULE 11.6


ELECTRICAL POWER

GCB 1

EMERG AC BUS

AC Busbar System Figure 48


AC GRND SERV

GSBC 1 GSBC 2 EBTR

EMERGENCY INVERTER

GEN 1

AC BUS 1

BTB 1

GPTC

GEN 3

EBTC

APTC

APC

EPC

ESS AC BUS

EXT PWR

BTB 2

GEN 2

AC BUS 2

GCB 2

Figure 48 shows the AC Busbar system, which would satisfy the requirements of AC priority as shows in Figure 47.

Page 3-79


AC GND SERV BUS

2

BATT 1

DC Busbar System Figure 49 Page 3-80


DC EXT PWR

DC EBTC BIC 1 BATT BUS 1

BATTERY CHARGER

ESS DC BUS

DC EPR

1

DC BTC DC BUS 1 AC BUS 1

AC BUS 1

TRU 1

DC GRND SERV BUS

BPC

EMERG DC BUS

1 2

TRU 3

ESS AC BUS

DC EBPC

TRU 2

DC BUS 2

GND SERV TRU

BIC 2

BATT 2

BATT BUS 2

AC BUS 2

AC BUS 2

BATTERY CHARGER

Figure 49 shows the DC Busbar system, which would satisfy the requirements of DC priority as shows in Figure 47.


MODULE 11.6


ELECTRICAL POWER

3.22.4 Parallel Electrical System In a parallel electrical system, all A.C. generators are connected to one distribution bus called the “Tie-Bus”. This type of system maintains equal load sharing for three or more generators. Since the generators are connected in parallel to a common bus, all generator voltages, frequencies and their phase sequence must be within very strict limits to ensure proper system operation. Figure 50 shows a simplified schematic of a parallel electrical system.

GPU

APU XPC 2 APB

TIE BUS

BTB 1 AC BUS 1

BTB 2

AC BUS 3

AC BUS 2

GCB 1

GCB 2

GEN 1

BTB 3

AC BUS 4

GCB 3

GEN 2

BTB 4

GCB 4

GEN 3

GEN 4

Parallel Electrical System Figure 50


Page 3-81


3.22.5 Split Parallel Electrical System A split parallel electrical system allows for the flexibility in load distribution but maintains isolation between systems when required. When closed, the split system breaker connects all generators together, thus paralleling the system. When open, the split system breaker isolates the right and left systems, thus creating a more flexible parallel system. Figure 51 shows a split parallel system for a four engine aircraft.

GPU

GPU

APU XPC 2

XPC 2 APB SSB TIE BUS

BTB 1 AC BUS 1

TIE BUS

BTB 2

AC BUS 3

AC BUS 2

GCB 1

GEN 1

BTB 3

GCB 2

GEN 2

BTB 4 AC BUS 4

GCB 3

GEN 3

GCB 4

GEN 4

Split Parallel Electrical System Figure 51

Page 3-82




3.23 GENERATOR CONTROL UNITS (GCU) Aircraft electrical power control systems include functions such as: 1. Voltage Regulation. 2. Current Limiting. 3. Protection for out-of-tolerance Voltages. 4. Protection for out-of-tolerance Frequencies. 5. Crew Alerting The major component used to perform these functions is called the “Generator Control Unit” (GCU). The GCU regulates generator output by sensing the aircraft‟s system voltage and comparing it with a reference signal. The voltage regulator then sends an adjusted current flow to the exciter field of the main generator. This in turn controls the main generator‟s output voltage. Protection circuitry monitors various electrical system parameters including over voltage and over current conditions, frequency, phase sequence, and current differentials. If a fault occurs, the protection circuitry then operates corresponding electric relays in order to isolate the defective components. In the case of a generator system failure, the GCU senses partial loss of electrical power and automatically sends the appropriate signal to the “Bus Power Control Unit” (BPCU). In this event the BPCU will automatically isolate any defective generator and reconnect the load bus to another power source. 3.23.1 Power Distribution System Control On modern aircraft employing a parallel, or split-bus system, a centralized means of controlling the power distribution between individual load busses is essential. If a generator fails or a bus shorts to ground, the appropriate bus ties and generator circuit breakers must be set to the correct position. In the event of a system overload, the control unit must reduce the electrical load to an acceptable level. This is called “Load Shedding”. In load shedding, the aircraft‟s galley power is normally the first nonessential load to be disconnected. The control unit must automatically reconnect any essential loads to an operable bus. This power manipulation must take place within a fraction of a second to ensure an uninterrupted flight. To achieve this, modern aircraft employ a solid state “Bus Power Control Unit” (BPCU).


Page 3-83


The BPCU also receives data from the Generator Control Unit GCU, the Ground Power Control Unit (GPCU) and various bus ties and circuit breakers of the system. It also receives input information concerning system loads from “Load Controllers”. Load controllers are electric circuits that sense real system current and provide control signals for the generator‟s CSDU rpm governor. The CSDU output rpm in turn affects the generator output frequency. Load controllers receive their input signals from current transformers. Figure 52 shows current transformers as fitted to the Boeing 737 aircraft.

GEN 1 LINE CURRENT

APU LINE CURRENT

GEN 2 LINE CURRENT

TO PROTECTION CIRCUITS

Current Transformers Figure 52

Page 3-84




3.23.2 Current Transformers Current transformers consist of three inductive pickup coils that provide current sensing signals. The main power leads carrying the three phase A.C. supply from each generator are routed through the corresponding holes in the current transformer. As the A.C. travels through the cable, the corresponding magnetic field induces a voltage into the current transformer. The electrical signals from the current transformer, in conjunction with the GCU and BPCU, are used to control protection circuitry and supply signals to load meters on the overhead panel in the flightdeck. Figure 53 shows the operation of a current transformer.

Current transformers Figure 53


Page 3-85


BTB BTB

GTB

LEFT IDG

LEFT GCU

RIGHT IDG

GTB

RIGHT GCU

BTB

BPCU

APU GEN

APB

APU GEN GCU

Figure 54 shows a schematic of the Power Distribution Control System (PDCS).

Power Distribution Control System (PDCS) Figure 54 Page 3-86




3.23.3 Generator Control & Protection Each GCU contains the following items: 1. A Field Power Supply (TR unit) which converts 3-phase ac power from the generator to a rippled dc voltage for the generator exciter. 2. A Control dc Power Supply (TR unit) which converts 3-phase ac power from the generator to 28 volt dc power for the generator switch and the protection circuits. 3. A transistor voltage regulator, which controls dc power returning to the field power supply from the generator exciter. 4. A double-coil magnetic latching relay called the Generator Control Relay (GCR) which connects the output of the field power supply to the generator exciter. 5. Transistorised protection circuits for: Over-voltage (OV)

Over-current (OC)

Under-voltage (UV)

Differential Current Protection (DP)

Over-frequency (OF)

Under-Frequency (UF)

The GCR can be tripped by three manual actions and automatically by the detection of five faults: Manual 1.

Generator switch off. (Normally left in the ON position).

2.

CSD disconnect switch activated (where applicable).

3.

Fire handle pulled.

Automatic (Figures quoted are typical) 1.

Over-voltage (130 ± 3 volts).

2.

Under-voltage (100 ± 3 volts).

3.

Over-current (170 amps).

4.

Differential Protection (20 amp difference).

5.

Over Frequency.


Page 3-87


6.

Over-voltage Unit

The unit is designed to protect electrical systems of 200 volts, 3-phase, 400 Hz from over voltage faults. The unit will operate, to protect the system, when the voltage rises to (typically) above 220 volts and in effect disconnects the generator from its loads. Figure 55 shows the circuit for the Over-voltage Unit. VOLTAGE NORMAL

TR1

DE-ENERGISED

OVERVOLT RELAY

TR2

D1

D3

D9

D5 R1

D2

D4

LOW VOLTS

D6 HIGH VOLTS

ZD4 NO BREAKDOWN

D8

R4

T1

RV1 ZD1 ZD2 ZD3 C1

ZD5

OFF

T2 ON

T3 OFF

D7 R5

R2 R3

C2

Over-voltage Unit Figure 55 The supply to the unit is via two open-delta connected transformers TR1 and TR2 via the full-wave Rectifier Bridge, across which is connected the potential divider network R1, RV1 and R2. Under normal conditions, the voltage developed across RV1 and R2 is not sufficient to break down Zener diodes ZD1, ZD2 and ZD3. Transistor T1 is therefore not conducting but transistor T2 is, due to the high potential on its base. This means that T3 is not conducting and the over voltage relay is not energised.

Page 3-88




When the supply voltage exceeds 220 volts, ZD1, ZD2 and ZD3 break down and (after a time delay afforded by R4, R5 and C2) the voltage across R3 is able to energise transistor T1. Thus T1 conducts, shutting off T2, which causes T3 to conduct and the relay to be energised. C1 is smoothing capacitor. D7 enables C2 to discharge through R3 for operation on subsequent over-voltage faults. Overall temperature compensation for transistor T1 is effected through Zener diodes ZD1, ZD2 and ZD3, and by Zener diodes ZD4 and ZD5 for transistors T2 and T3 respectively. The increased gain of the transistors due, to a rise in ambient temperature, is compensated for by the decreased output of the Zener diodes. Diode D9 suppresses the peak inverse voltage, which would be applied to transistor T3 from the relay coil when the transistor is shut off. The length of the time will vary considerably, depending on the value of the overvoltage and on the rate at which the voltage is rising. The following figures are given as a rough guide: at 220 volts the unit will operate in “less than 10 seconds”; at 225 volts it will operate in 0.7 to 1.6 seconds; at 250 volts it will operate in 0.35 to 0.65 seconds. From this it can be said that the time delay is Inversely proportional to the value of the over-voltage (i.e. the higher the overvoltage the shorter the time delay). 3.23.4 Under-voltage & Reverse Phase Sequence Unit This unit is designed to protect electrical systems of 200 volts, 3-phase, 400 Hz from under-voltage faults. It will operate, to protect the system, when the voltage falls to (typically) 173 ± 2 volts and in effect disconnects the generator from its load. It also affords protection against incorrect phase rotation in the event of the generator being incorrectly connected. The supply to the unit is applied to two open-delta connected transformers TR1 and TR2. Each transformer has two secondary windings S1 and S2. The S1 secondary windings supply the phase sequence circuit, whilst the S2 windings supply the transistor amplifier and voltage sensing circuits.


Page 3-89


Figure 56 shows the Under-voltage & Reverse Phase Sequencer Unit circuit.

PART OF GCB CLOSE “AND”

S1

TR1

TR2

S2

S2

UNDERVOLT RELAY

D4

S1

D1

D3

D5

D2

D4

D6

R6

T1 D2

C1

D1

ZD1

T2

ZD2 D3

R1 R 4

C2

T H R

R 5

R 7

C4

C3

Under-voltage & Reverse Phase Sequencer Unit Figure 56 Under normal operating conditions the under volt relay is energised, i.e.: transistors T1 and T2 are switched “on”. This is accomplished using a two-input “AND” gate circuit formed by D1, R6 and D3, R5, thermistor „THR‟ and R4. When D1 and D3 are “blocked” by the phase sequence circuit and by the voltage sensing circuit respectively, the transistors are switched “ON” and the relay is energised. Should either diode be “unblocked”, current will flow through it and through R6, resulting in the relay being de-energised.

Page 3-90



MODULE 11.6


ELECTRICAL POWER

3.23.5 Time Delay Activation The time delay is to prevent tripping of the GCR due to transient under-voltages, or to allow the CSD to slow down to a UF (Under-frequency) condition on engine shutdown and inhibit tripping of the GCR. The GB can be tripped by a fault tripping the GCR, or an over-frequency condition: (430 ± 5Hz). Faults of under-frequency or over-frequency on their own do not trip the GCR. Figure 57 shows the time delay circuit. NO R E G U L AT E D DC

B R E A KD O W N

F RO M G C U PO W E R SU PPL IES

Z D1

R1

D1 NO OU TPU T (G C U N O T R I P )

C1

+V E W H E N G E N

NO C H AR G E

O U T PU T N O R MA L

Q1

ON R2

B R E A KD O W N R E G U L AT E D DC

(A F T E R C 1 C H A R G E D )

F RO M G C U PO W E R SU PPL IES

Z D1

R1

D1 OU TPU T (G C U T R IP )

0V

WH EN GEN

O U T P U T U ND E R V O LT A G E

C1

C H AR G E

O R IN C O O R E C T P H A S E S E Q U E N CE

Q1

OFF R2

Time Delay Circuit Figure 57 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-91


3.23.6 Abnormal Frequency Protection Both the over-frequency (OF) and under-frequency (UF) detectors are transistorised voltage sensitive electronic circuits within the GCU, normally sensing one phase at the generator breaker terminal. At a level greater than 430 ± 5Hz the OF detector will send a signal to trip the GCB. At a level of less than 365 ± 5Hz, the UF detector will send a signal to trip the GCB, to inhibit the power ready circuit involved in closing the GCB and to prevent an under-voltage (UV) signal from tripping the GCR during engine shut down. Figure 58 shows the Abnormal Frequency Protection circuit.

PWR READY INHIBIT UV ON ENG SHUTDOWN

GEN O/P

UF DETECTOR 365 Hz ± 5Hz

0.5 - 1 SEC

OF DETECTOR 430 Hz ± 5Hz

Abnormal Frequency Protection circuit Figure 58

.

Page 3-92


GB TRIP


MODULE 11.6


ELECTRICAL POWER

3.23.7 Differential Current Protection The purpose of a “Differential Current Protection” system is to detect a shortcircuited feeder line or generator busbar which would result in a very high current demand on the generator, and possibly result in an electrical fire. The difference between the current leaving the generator and the current arriving at the busbar is called a “Differential Fault” or a “Feeder Fault”. In an A.C. system, current comparisons are made phase for phase, by two three-phase current transformers, one on the neutral side of the generator (Ground DPCT) and the other on the up stream side of the busbar (Load DPCT). Figure 59 shows the arrangement and principle of a system as applied to a single-phase line. BUSBAR

LOAD DPCT

GROUND DPCT

GEN

GENERATOR CONTROL RELAY

I

I-IF

FAULT IF LOAD

(I - I F) + (IF) = I

I-IF DP DETECTOR IN GENERATOR CONTROL UNIT

AIRCRAFT STRUCTURE

Differential Current Protection Figure 59


Page 3-93


If the current from the generator is I, and the fault current between the generator and busbar equals If, then the net current will flow through the aircraft‟s structure and back to the generator through the ground DPCT. The remainder of the current I - If , will flow through the load DPCT, the load, the aircraft structure, and then back to the generator via the ground DPCT. Thus, the ground DPCT will detect the generator‟s total current (I - If) + (If) which is equal to I, and the load DPCT will detect I - If. The difference in current (i.e. the fault current) between the two current transformers on the phase line is sensed to be greater than the specified limit (20 – 30A are typical values) a protector circuit within the GCU will trip the generator control relay. Figure 60 shows a 3-Phase Differential Protection circuit. GCB

CT1

CT4

CT2

CT5 CT6

CT3

GEN

R 1

R 3

R 5 D1

R7

D2 D3 R 2

R 4

R 6

C1

R V 1 ZD1

D4

3-Phase Differential Protection Circuit Figure 60

Page 3-94


TO TRIP GCR & GCB


MODULE 11.6


ELECTRICAL POWER

3.23.8 Over-Current Protection If a very heavy, potentially damaging load is placed on the generator, it is possible to form three simultaneous faults: over-current (OC), under-frequency (UF) and under-voltage (UV). If the underlying cause is the OC, it in itself will cause the voltage to drop and possibly the generator shaft to slow down. In this case, the OC detector will send out „OC LOCKOUT‟ signals to inhibit UF trip of the GCB or UV trip of the GCR. The OC fault takes priority in this case and will trip the GCR itself. Figure 61 shows an Over-current Protection circuit.

GCB

CT1 CT2 CT3

GEN

TO TRIP GCR & GCB

R 1

R 3

R 5 D1

R7

D2 D3 R 2

R 4

R 6

C1

R V 1 ZD1

Over-Current Protection Circuit Figure 61


Page 3-95


3.23.9 GCU Operation The primary purpose of the Generator Control Unit (GCU) is to connect the output of the field power supply to the generator field. This is achieved using a “Generator Control relay” (GCR). The GCR can only be closed by momentarily placing the generator control switch to „ON‟. The GCB can only be closed by placing the generator switch momentarily to „ON‟, energising the power ready relay (GCR closed, not UV, not UF) and tripping other breakers necessary to prevent paralleling. The GCR can be tripped by three manual actions and automatically by the detection of five faults. Manual actions: 1. Generator Switch to „OFF‟. 2. CSD disconnect switch activated. 3. Fire handle pulled. Automatic Trip: 1. OV (220V ± 6V). 2. UV (173V ± 6V). 3. OC (170A B737). 4. DP (20A Typical for Boeing). 5. OF (430 ± 5 Hz). The GCR having been tripped, or an UF (365 ± 5 Hz) condition will trip the GCB. Note: A fault of Under Frequency will not trip the GCR. Figure 62 shows a block schematic of the GCU.

Page 3-96



CSD DISC GEN OFF FIRE HANDLE

CUURENT DETECTOR OC - DP

FREQUENCY DETECTOR UF/OF

VOLTAGE DETECTOR (OV - UV)

PS

CONTROL DC PWR

CLOSE GCR

UF

UV

OV/UV

OF

UF

GCR TRIP

GB TRIP CIRCUIT

TRIP

GCB

CLOSE

TO GEN FIELD

TRIP & INTERLOCK CCTS


LINE CCTS

GEN AC 3 PHASE

BATTERY BACK UP

FIELD POWER SUPPLY

PART OF POWER READY RELAY


MODULE 11.6

ELECTRICAL POWER

Generator Control Unit (GCU) Figure 62

Page 3-97


3.24 TRANSFORMERS A transformer is a device for converting A.C. at one frequency and voltage to an A.C. at the same frequency but at another voltage level. It consists of three main parts: 1. An Iron Core: provides a circuit of low reluctance for an alternating magnetic field. 2. A Primary Winding: connected to the main power source. 3. A Secondary Winding: which receives electrical energy by mutual induction from the primary winding and delivers it to the secondary circuit. There are two classes of transformers: 1. Voltage or Power transformers. 2. Current Transformers. 3.24.1 Voltage transformers Voltage transformers are connected so that the primary windings are in parallel with the supply voltage, in the current transformers, the primary windings are connected in series with the supply voltage. These transformers may be single phase or three-phase devices. Transformers for three-phase circuits can be connected in one of several combinations of star and delta connections depending on the requirements for the transformer. When the star connection is used in a three-phase transformer for the operation of three-phase equipment, the transformer may be connected as a three-phase system. If a single phase load has to be powered from a three-phase supply it is sometimes difficult to keep them balanced, itr is therefore essential to provide a neutral wire so that connections of the loads may be made between this and any one of the three-phase lines. Figure 63 & 64 show transformer connections.

Page 3-98



MODULE 11.6


STAR CONNECTION THREE-WIRE

ELECTRICAL POWER

STAR CONNECTION FOUR-WIRE

Transformer Three-Phase To Three-Phase Supplies Figure 63

STAR-WOUND PRIMARY

DELTA-WOUND PRIMARY

STAR - DELTA CONNECTION

Transformer Three-Phase To Single-Phase Supply Figure 64 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-99


3.24.2 Transformer Ratings Transformers are usually rated in Volt/Amperes or Kilovolt/Amperes. The difference between the output terminal voltages at full load and no-load, with a constant input voltage is called the regulation of the transformer. As in the case of an A.C. generator, regulation is expressed as a percentage of the full load voltage, and depend not only on actual losses (i.e. eddy current, magnetic leakage and hysteresis losses), but also on the power factor of the load. Thus, an inductive load, i.e. on having a lagging power factor, will give rise to a high percentage regulation, while with a capacitive load, i.e. one having a leading power factor, the regulation may be a negative quality giving a higher output voltage on full load then on no-load. Changes in power supply frequency, or the connection of a transformers supply whose frequency differs from that for which the transformer was designed, has a noticeable effect on its operation. This is due to the fact that the resistance of the primary windings is so low that they may be considered to be a purely inductive circuit. If the frequency is reduced at a constant value of voltage, then the current will rise. The increased current will in turn bring the transformer core nearer to magnetic saturation and this decreases the effective value of inductance leading to still larger current. Thus, if a transformer is used at a frequency lower that that for which it was designed, there is a risk of excessive heat generation at the primary windings and subsequent burn out. On the other hand, a transformer designed for low frequency can be used with higher frequencies, since in this case the primary current will be reduced.

Page 3-100



MODULE 11.6


ELECTRICAL POWER

3.24.3 Transformer Rectifier Units (TRU) Transformer-rectifier units (TRU) are a combination of a static transformer and rectifiers, and are used in some A.C. systems as secondary supply units, and also as the main conversion units in aircraft having rectified A.C. power systems.

D4 D2

D6

D3 3 PHASE A.C. SUPPLY

D1

D5

D.C. LOAD

Figure 64 shows the basic principle of operation of a Transformer-rectifier.

Transformer-Rectifier Operation Figure 64 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-101


Figure 65 shows a TRU designed to operate on a regulated three-phase input of 200V at a frequency of 400Hz and to provide a continuous output of 110A at approximately 26V.

RECTIFIER SECTION TRANSFORMER SECTION

AMMETER SHUNT TERMINALS

A.C. INPUT TERMAINLS

TEMPERATURE WARNING SYSTEM TERMINAL

D.C. OUTPUT TEMINALS

Transformer-rectifier Unit Figure 65

Page 3-102




METER

RECTIFIER UNIT A.C. I/P

STAR

STAR

TRANSFORMER UNIT

DELTA

D.C. O/P

WARNING SYSTEM

Figure 66 shows the schematic circuit for the TRU shown in figure 65.

TRU Schematic Circuit Figure 66 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-103


The unit consists of a transformer and two three-phase bridge rectifier assemblies mounted in separate sections of the casing. The transformer has a conventional star-wound primary winding and secondary windings wound in both star and delta configurations. Each secondary winding is connected to individual bridge rectifier assemblies made up of six silicon diodes, and connected in parallel. An ammeter shunt (dropping 50mV at 100A) is connected in the output side of the rectifier to enable current taken from the main D.C. output terminals to be measured at ammeter auxiliary terminals. These terminals together with all others associated with input and output circuits are grouped on a panel at one end of the unit (see figure 60). Cooling for the unit is by convection through gauze covered ventilation panels and in order to give warning of over-heating conditions, thermal switches are provided at the transformer and rectifier assemblies, and are connected to independent warning lights. These switches are supplied with D.C. from an external D.C. source (normally on of the busbars) and their contacts close when temperature conditions at their respective locations rise to approximately 150ºC and 200ºC.

Page 3-104




3.25 ROTARY INVERTER Used to produce 26V or 115V 400Hz from a D.C. source. It consists of a D.C. motor driving an A.C. generator, and since many of the systems, which are to be operated from it, are dependent on constant voltage and frequency, the A.C. supply must be regulated accordingly. Figure 67 shows a Rotary Inverter circuit. 28V D.C.

INVERTER CONTROL BREAKER

115V 400Hz

MOTOR

GENERATOR

Rotary Inverter Circuit Figure 67 When the inverter is switched on, D.C. is supplied to the motor armature and shunt filed winding, and also to the excitation field winding of the generator. The motor will start to drive the generator, which in turn produces a three-phase A.C. output at 115V. In order to control the voltage at this level, the D.C. supply is passed through a resistor in series with the generator field. This resistor is preset to give the required excitation current at the regulated D.C. system voltage level. Since the frequency of the generator output is dependent of the speed of rotation of the motor, this requires some form of control. This is achieved by using another pre-set resistor , which is connected in series with the motor shunt filed to provide sufficient excitation current to run the motor and generator at the speed necessary to produce a 400Hz output.


Page 3-105


Figure 68 shows a schematic of another type of rotary inverter, and although it is only found on older types of transport aircraft, it shows an example of variation in application of principles.

28V D.C.

INVERTER CONTROL BREAKER

CARBON PILE

115V 400Hz

SHUNT FIELD

SERIES FIELD

RECTIFIER REGULATOR FIELD

Rotary inverter Circuit Figure 68 The motor and generator share a common armature and filed system, and control of voltage and frequency is based on the carbon pile regulator system. The D.C. section of the machine is of the four-pole compound wound type, the D.C. being supplied to the armature winding, series and shunt field-windings. The A.C. section corresponds to a star-wound generator, the winding being located in slots of the armature and beneath the D.C. windings. The A.C. winding is connected to a triple slip ring and brush-gear assembly at at the opposite end to the commutator. When the inverter is in operation, a threephase output is induced in a rotating winding and not a fixed stator winding as in the case of a conventional A.C. generator. The A.C. output is rectified and supplied to the voltage coil of the regulator which varies the pile resistance in the usual manner, this in turn, varying the current flow through the common field system to keep both the voltage and frequency of the A.C. output within limits.

Page 3-106



MODULE 11.6


ELECTRICAL POWER

3.25.1 Static Inverter

REGULATOR NOTCH CONTROL

TURN ON DELAY

PULSE SHAPER

400Hz SQUAREWAVE GENERATOR

28V D.C.

FILTER NETWORK

NOTCH TIME

POWER DRIVER

CONSTANT CURRENT GENERATOR

NOTCH TIME

OUTPUT STAGE

CURRENT SENSOR

VOLTAGE SENSOR

115V 400Hz A.C.

ODD HARMONIC FILTER

These inverters perform the same conversion function as the rotary machines, but by means of solid state circuit principles. They are employed in a number of types of aircraft and in some cases providing the normal source of A.C. They are more commonly used in supplying emergency sources of A.C. to certain essential A.C. systems when a failure of the normal source of A.C. has occurred. Figure 69 shows a block schematic of a static inverter.

Static Inverter Block Schematic Figure 69 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-107


The D.C. is supplied to transistorised circuits of a filter network, a pulse shaper, a constant current generator, power driver stage and the output stage. After any variations in the input have been filtered or smoothed out, D.C. is supplied to a square wave generator which provides the first stage conversion of the D.C. into square-wave form A.C. and also establishes the required operating frequency of 400Hz. This output is then supplied to a pulse shaper circuit, which controls the pulse width of the signal and changes its waveform before passing it onto the power driver stage. The D.C. required for pulse shaper operation is supplied via a turnon delay circuit. This is to cause the pulse shaper to delay its output to the power driver stage until the voltage is stabilised. The power driver supplies a pulse-width modulated symmetrical output to control the output stage, the signal having a square-wave form. The power driver also shorts itself out each time the voltage falls to zero, i.e. during “notch-time”. The output stage also produces a square-wave output but of variable pulse width. This output is finally fed to a filter circuit, which reduces the total odd harmonics to produce a sine wave output at the voltage and frequency required for operating the systems connected to the inverter. As in the case of other types of generators, the output of a static inverter must be maintained within certain limits. In the example shown, this is achieved by means of voltage and current sensors, both of which control the notch time of the pulse width shaper output via a regulator and notch control circuit.

Page 3-108



MODULE 11.6


ELECTRICAL POWER

3.26 BOEING 737 ELECTRICAL SYSTEM

TRU 3

DC BUS 1 GRND SERV BUS

TRU 1 GRND SERV RELAY

26V A.C. BUS 1

TRANSFER RELAY 1 TR 1

MAIN BUS 1

115 V GEN BUS 1 EXT PWR BUS

TRANS BUS 1

BTB 1 GCB 1

GEN 1

TRU 3 DISCONNECT DC BUS 2

TRU 2

APU GCB EPC EXT POWER

BATTERY BUS

26V A.C. BUS 2 TRANS BUS 2

TRANSFER RELAY 2

MAIN BUS 2

GEN BUS 2 115 V

BTB 2

APU GEN

GCB 2

TR 2

GEN 2

The Boeing 737 aircraft‟s electrical system operates using a split busbar system and is shown in Figure 70

Boeing 737 Electrical System – Schematic Figure 70 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-109


3.26.1 Controls & Indications The controls for the electrical system on the Boeing 737 300 are on the 'ELECTRIC' section of the overhead panel. There are 4 panels controlling the electrical system. Figure 71 shows the P5-13 panel.

Boeing 737 P5-13 Electrical Panel Figure 71 Page 3-110




3.26.2 Boeing 737 P5-13 Electrical Panel D.C. AMMETER – Indicates current of a source selected by the D.C. meter selector – for TR 1, 2, 3 AND BAT. D.C. VOLTMETER – Indicates voltage of a source selected by the D.C. meter selector (all positions). D.C. METER SELECTOR – Selects the D.C. source for the D.C. voltmeter indications. TEST – Used by maintenance personnel. Connects the voltmeter to the power systems test module for selection of additional reading points. BATTERY SWITCH – (guarded in “ON”). „OFF‟ no power to battery bus unless STANDBY POWER Switch (P5-5) is in the „BAT‟ Position. FREQUENCY METER – Indicates frequency of the source selected by the A.C. meter selector. RESIDUAL VOLTS SWITCH – Push to read the residual voltage of the selected generator (with associated generator control relay tripped). A.C. VOLTMETER – Indicates (130 volt scale) voltage of the source selected. When the residual volts switch is pushed, the 30 volt scale reads residual voltage of the selected generator. A.C. METERS SELECTOR – Selects the A.C. source for A.C. voltmeter and frequency meter indications. TEST – Used by maintenance personnel. Connects the voltmeter and frequency meter to the power systems test module for selection of additional reading points. GALLEY POWER SWITCH – “ON” provides electrical power to the galleys when both generator buses are energized.


Page 3-111


Figure 72 shows the P5-13 panel.

Boeing 737 P5-5 panel Figure 72 3.26.3 Boeing 737 P5-5 Electrical Panel STANDBY POWER OFF LIGHT – Illuminated amber when standby buses are not powered. GENERATOR DRIVE LOW OIL PRESSURE LIGHT – Illuminated amber when No 1 or 2 generator CSD oil pressure is below minimum operating limit. GENERATOR DRIVE HIGH OIL TEMPERATURE LIGHTS – Illuminated amber when No 1 or 2 generator CSD oil temperature exceeds limits. GENERATOR DRIVE OIL TEMPERATURE INDICATORS – No 1 or 2 “RISE” scale displays the temperature rise within the CSD. “IN” scale displays temperature of oil entering the CSD. Page 3-112




ELECTRICAL STANDBY POWER SWITCH – (Guarded in “AUTO”) AUTO - Normal operation switch position. With loss of all A.C. power in the air, the battery is automatically connected to supply the D.C. standby bus power for essential D.C. equipment and to the inverter to supply A.C. standby bus power for essential A.C. equipment. To prevent unnecessary battery drain on the ground, the standby busses are not powered by the battery. OFF – Turns off the power to standby power busses. BAT – Battery supplies power to the battery bus. The D.C. standby bus and turns on the inverter to supply the A.C. standby bus. GENERATOR DRIVE DISCONNECT SWITCHES – (Guarded in the connected position). Disengages the CSD from No 1 or No 2 generator. GENERATOR DRIVE TEMPERATURE SWITCH – Selects CSD oil temperature displayed on indicators (“N” or “RISE”).


Page 3-113


Figure 73 shows the P5-4 panel.

Boeing 737 P5-4 panel Figure 73 GROUND POWER AVAILABLE LIGHT – Illuminated blue when ground power is connected. TRANSFER BUS OFF LIGHTS – illuminated amber when the No 1, or 2 generator transfer bus, is not powered. BUS OFF LIGHTS – Illuminated amber when the No 1, or 2 generator bus, is not powered. GENERATOR OFF BUS LIGHTS – Illuminated blue when the generator is not supplying the generator bus (generator is “OFF” the bus). APU GENERATOR OFF BUS LIGHT – Illuminated blue when the APU is running but the generator is not supplying a generator bus. GENERATOR SWITCHES – For No 1 or 2 generator, three-position switch, momentary ON – OFF and spring loaded to centre, neutral position. Page 3-114




“ON” – Disconnects either external power or APU generator power from the generator bus and connects engine generator to the generator bus. “OFF” – Disconnects generator from the generator bus. APU GENERATOR BUS SWITCHES - Three-position switch, momentary ON – OFF and spring loaded to centre. “ON” – Connects APU generator to No 1 (or No 2) generator bus. “OFF” – Disconnects APU generator from No 1 (or No 2) generator bus. NOTE; Both APU generator switches work the same on the ground. Only one bus can be powered by the APU generator in the air. GENERATOR A.C. AMMETER – Displays engine generator No 1 or 2 load in amperes. GROUND POWER SWITCH – Three-position switch, momentary ON – OFF and spring loaded to centre position. “ON” – If ground power is available, the engine or APU generators are tripped off the generator busses and ground power is supplied to the generator busses. “OFF” – Ground power tripped off the generator busses. BUS TRANSFER SWITCH – (Guarded in AUTO). “AUTO” – Transfer bus automatically transfers to opposite generator power source if one becomes inoperative. “OFF” – Isolates D.C. systems (de-energizes TR3 disconnect relay) and deactivates automatic transfer feature.


Page 3-115


Figure 74 shows the APU control panel

APU Control Panel Figure 74

Page 3-116




The controls for the APU are located on the forward overhead panel and consist of the following: APU SWITCH – Three position switch “OFF” momentary “START” and spring loaded to “ON”. “START” – Start circuit armed. Momentarily holding in “START” and allowing return to “ON”, fuel valve opens, air inlet door opens and the start sequence is initiated. “ON” – Operating position after start. “OFF” – Shuts down the APU. LOW OIL QUANTITY LIGHT – Illuminated blue when oil tank fluid level is low. LOW OIL PRESSURE LIGHT – Illuminated amber when oil pressure is low. APU automatically shuts down. HIGH OIL TEMPERATURE LIGHT – Illuminated amber when APU oil temperature exceeds maximum allowable. APU automatically shuts down. OVERSPEED LIGHT – Illuminated amber when APU turbine speed exceeds allowable rpm. APU automatically shuts down. A.C. AMMETER – Displays APU generator load current. EXHAUST GAS TEMPERATURE INDICATOR – Displays APU exhaust gas temperature in ºC.


Page 3-117


3.26.4 B737 Electrical Power Distribution The primary power source is non-paralleled 115/200V 3 phase 400Hz A.C. from two 40kVA generators. A source of A.C. power can be supplied from another 40kVA generator riven by an Auxiliary Power Unit (APU). Power can also be supplied form an external A.C. power unit. All D.C. is supplied via three TRUs. The four power sources are connected to the busbars by six 3-phase breakers and two transfer relays, which are energised and de-energised according to the switching selections, made on the system control panels. An interlocking circuit between breakers and switches is also provided to enable proper sequencing of breaker and overall system operation. A source of power switched onto or entering the system always takes priority and so will automatically disconnect any existing power source. The switches on the control panel (P5-4) are of a “Momentary Select” type in that following a selection, they are returned to a neutral position by spring loading. The bus transfer switch is retained in the “Auto” position by a guard cover to provide a path for signals controlling the “Normal” and “Alternate” positions of the transfer relays. In the “Off” position the transfer relays are prevented from being energised to the “Alternate” position so that the two generating systems are completely isolated from each other. The indicating lights on the P5-4 panel are as follows: Ground Power Available (Blue) aircraft.

When external power is plugged into the

Transfer Bus Off (Amber) relay is de-energised.

-

When either the normal coil or a transfer

Bus Off (Amber) open.

-

If both the respective GCB and BTB are

Gen Bus Off (Blue)

-

If the respective GCB is open.

APU Gen Bus Off (Blue)

-

If APU engine is running and over 95% rev/min, but there is no power from the generator.

The ammeters indicate the load current of both main generators.

Page 3-118




3.26.5 Operation When external power is connected to the aircraft and is switched on, the external power contactor closes and energises both bus-tie breakers (BTB) to connect power to the whole busbar system. The connection between the generator busbars and the transfer bus busbars is made via the transfer relays which are energised to the “Normal” position by the BTB‟s. Once the number one engine is started and the generator switch (P5-4) is selected to “ON”, the BTB 1 will open and GCB 1 will closes, ensuring that all the number one system‟s power is delivered from the number one generator. At this time however the external power is still providing power for the number 2 system. When the number two engine has been started and its generator switched on, BTB 2 trips open, GCB 2 closes to connect the generator to the number two system busbars, and the external power contactor also trips open. If it is only necessary for the services connected to the ground service busbar to be operated from external power, this may be achieved by leaving the ground power switch on P5-4 in the “OFF” position, and switching a separate ground service switch (forward attendant‟s panel) to “ON”. The switch energises a ground service relay the contacts of which change over a connection from the generator bus 1 to the external power busbar. The APU generator is connected to the entire busbar system via its own threephase breaker, this, in turn being energised by two APU generator switches on the P5-4 panel. Placing the left (No 1) switch to “ON” closes the APU generator breaker and also BTB 1, and with the right or (No 2) switch placed to “ON” the BTB 2 is closed. As in the case of connecting an external power supply, the transfer relays are energised to the “Normal” position by the BTB‟s. The normal in-flight configuration of the power distribution system is for each generator to supply its respective busbars through its own breakers (GCB1 & GCB 2). The generator switches (P5-4) then energise these breakers; the interlock circuits keep BTB‟s 1 & 2 in the open position, so that the generator systems are always kept entirely separate. GCB 1 and GCB 2 have a set of auxiliary contacts which in the closed position energise transfer relays to their “Normal” positions and so provide connections between generators and transfer busbars 1 and 2. As will be noted from the diagram, the transfer busbars supply TRU‟s 1 and 2 while TRU 3 is supplied direct from main busbar 2. In the event of a loss of power from one or the other generators, say the number 1 for example; GCB 1 will open thus isolating the corresponding busbars. When GCB 1opens, another set of auxiliary contacts within the breaker permit a D.C. signal to flow from the control unit of the number 2 generator via a bus transfer switch, the “alternate” coil of transfer relay 1. The contacts will change over so that power is supplied to the number 1transfer busbar from generator 2 generator, which is still supplying its busbars in the normal way. A similar transfer of power takes place in the event of loss of power from generator 2.


Page 3-119


Generator busbar 1 and main busbar 1 which carry non-essential loads, can not be supplies with power from generator 2 under the power loss example given on the previous page. If, however, power to these busbars is required, the APU may be started in flight and its number 1 switch (P5-4), momentary placed to the “ON” position, thereby closing the APU breaker and the BTB 1. At the same time, transfer relay 1 contacts would change over from “Alternate” to “Normal” so that the APU supplies the whole number 1 system. If a loss of power from the number 2 system should occur, it is not possible to connect it to the APU since its number 2 switch is electrically locked out during in flight operations. 3.26.6 Generator Feeder Lines Figure 75 shows the routing of the feeder lines from the main generators and the APU generators. At the wing/fuselage junction, the lines pass through sealed connectors into the underfloor area. All lines are then routed through an electrical/electronics compartment. Those from the main generators pass through sealed connectors into unpressurised nosewheel well to connect up with the generator breakers. TO LOAD BUSBARS P6 PANEL

APU GENERATOR BREAKER

LINE CURRENT TRANSFORMER ENGINE/WING DISCONNECT No 1 GENERATOR BREAKER

No 2 GENERATOR BREAKER

ENGINE/WING DISCONNECT

CSD/GENERATOR NUMBER 1

CSD/GENERATOR NUMBER 2

FROM APU GENERATOR

Generator feeder Lines – B737 Figure 75 Page 3-120



MODULE 11.6


ELECTRICAL POWER

The feeder lines from the APU generators are connected to its breaker located above floor level within a special compartment (P6) on the flight deck to the rear of the First Officer‟s position. This compartment contains most of the A.C. and D.C. busbars, bus-tie breakers, and voltage control and protection units for all three generators and an external power control unit. Figure 76 shows the location of the electrical/electronics compartment and the P6 compartment.

EXTERNAL POWER

GENERATOR BREAKERS

P6 PANEL

APU BREAKER

BUSBAR PROTECTION PANEL

GENERATOR & APU CONTROL UNITS

BUSTIE BREAKERS

TRUs BATTERY CHARGER

NOSE WHEEL BAY

EXTERNAL POWER CONTACTOR

INVERTER E1 RACK E3 RACK BATTERY

EXTERNAL D.C. RECEPTACLE E2 RACK ACCESS DOOR TO COMPARTMENT

Electrical/Electronic/P6 Compartments Figure 76 The feeder lines from the main generator breakers pass into this compartment to connect with the A.C. busbars. A circuit breaker panel is mounted on the front side of the compartment, thus this compartment is termed the “Load Control Centre” of the aircraft.


Page 3-121


Figure 77 shows the P6 panel.

APU GENERATOR CONTROL UNIT

BUS PROTECTION PANEL

G7 POWER SUPPLIES FOR TEST EQUIPMENT

G5

115 V A.C. 28V D.C.

No 2 GENERATOR CONTROL UNIT

G4 G3

GEN BUS GEN BUS No 2 No 1

No 1 GENERATOR CONTROL UNIT

P6 PANEL (BEHIND F/O)

P18 PANEL (BEHIND CAPT)

P6 Panel Figure 77

Page 3-122



MODULE 11.6


ELECTRICAL POWER

3.26.7 Boeing 737 D.C. Power The 28V D.C. system consists of: 1. Three 50-amp transformer-rectifier units (TRU). 2. A 36 Ampere/hour battery. 3. Battery Charger. The TRU‟S convert 115V A.C. to 28V D.C. and are identified as TRU 1, TRU 2 and TRU 3. The three TRU‟s are connected in such a way that the loss of any one unit will not result in the loss of a D.C. busbar. The relay between TRU 1 and TRU 3 is held closed by supplying D.C. signals from the generator control units via the bus transfer switch in its “Auto” position. Figure 78 shows the D.C. power distribution system.

115 V AC GRND SERV BUS

115 V AC MAIN BUS 2

ALTERNATE BATT CHRG TRANS RLY

115 V AC TRANSFER BUS 1


BATT OVERHEAT RLY

BATT CHRG

T/R 1

T/R 2

T/R 3

APU START INTERLOCK RLY

DC BUS 1

DC BUS 2

HOT BATT BUS ALTERNATE

NORMAL

OFF

BAT

BATTERY BUS CONTROL

ON STANDBY POWER

BATT BUS BATTERY BAT

OFF

AUTO

B737 D.C. Power Distribution Layout Figure 78


Page 3-123


3.27 B747 GENERATING SYSTEM

GEN 4 GEN 3 GEN 2 GEN 1

GCB 4 GCB 3 GCB 2 GCB 1

AC BUS 4 AC BUS 3 AC BUS 2 AC BUS 1

BTB 4

TIE BUS

BTB 1

GPU

XPC 2

APB

APU

BTB 2

SSB

BTB 3

TIE BUS

XPC 2

GPU

The B747 aircraft uses a further variation of the split busbar system. Its A.C. power generating system utilises a system of interlocking GCB‟s and BTB‟s, but in this case various combinations of generator operation are possible. Figure 79 shows a simplified diagram of the B747 A.C. power generation system.

B747 – A.C. Power Generating System Figure 79

Page 3-124




3.27.1 Operation If the GCB‟s only are closed, then each generator will only supply its respective load busbar; in other words, they are operated individually and unparalleled. The generators may, however, also be operated in parallel when the BTB‟s are closed to connect the load busbars to the Tie-Busbar. As will be noted from figure 77, this busbar is split into two parts by a split system breaker (SSB) which, in the open position allows the generators to operate in two parallel pairs. Closing of the SSB connects both parts of the Tie-Busbar so that all four generators can operate as a fully paralleled system. By means of the interlocking system between breakers and the manual and automatic sequencing by which they are controlled, any generator can supply power to any load busbar, and any combination of generators can be operated in parallel.


Page 3-125


3.28 LOAD SHARING When a load is placed on an ac busbar, the nature of the load will determine the power factor of the system. Any load current, whether leading or lagging, can be thought of as having two components: 1. One in phase with the voltage. 2. One in quadrature with the voltage. The component in phase with the voltage is termed the real load component and the quadrature component is termed the reactive load component. Figure 80 shows the In-phase and In-Quadrature components of load current.

CURRENT (REACTIVE)

VOLTAGE

CURRENT (REAL) CU

RR

EN

T(

AC

TU

AL )

POWER FACTOR =

CURRENT (REAL) CURRENT (ACTUAL)

In-Phase & In-Quadrature Components Of Load Current Figure 80

Page 3-126



MODULE 11.6


ELECTRICAL POWER

3.28.1 Real Load Division Division of real load among paralleled generators becomes necessary, because it is not possible to attain exactly identical speed governor settings on all four generator constant speed drives. Therefore, in a paralleled system, the generator, which has the highest speed governor setting, will carry more than its share of real load. The unbalance in real load among paralleled generators is detected by means of current transformers and a real load division loop, whereby signals proportional to the unbalance are supplied to control devices, which correct the torque on the generator rotors. Figure 81 shows balanced real loads whereas figure 82 shows unbalanced real loads.

TOTAL REAL LOAD

I1

(AC TU

AL )

GEN 1 I2 (REAL) I2 (REACTIVE)

I1 (REACTIVE)

TOTAL REACTIVE LOAD

I1 (REAL)

I2

(AC TU

AL )

GEN 2

GENERATORS SUPPLYING SAME LOAD AND SHARING LOAD EQUALLY POWER FACTORS EQUAL

Balanced Real Loads Figure 81


Page 3-127


TOTAL REAL LOAD

I1

(AC TU

AL )

GEN 1 I2 (REAL) I2 (REACTIVE)

I1 (REACTIVE)

TOTAL REACTIVE LOAD

I1 (REAL)

I2

(A CT U

AL )

GEN 2

GENERATOR 1 IS SUPPLYING MORE REAL LOAD THAN GENERATOR 2 GEN 1’s POWER FACTOR INCREASES GEN 2’s POWER FACTOR DECREASES

Un-balanced Real Loads Figure 82 The frequency of an isolated generator is determined by the initial setting of the basic speed governor on its associated constant speed drive. Since the ac generators are synchronous machines, two or more generators operating in parallel will be locked together with respect to frequency, whereby the frequency of the paralleled system is that of the generator which supplies the highest frequency. If the speed governor setting on one constant speed drive is higher than others in a parallel operating system, its associated generator will motor the generators with which it is paralleled. In this case, the generator with the higher speed governor setting rotates at the same speed as its constant speed drive output. Since each generator is mechanically coupled to its own constant speed drive through an overrunning clutch, the generators which are being motored rotate at a speed which is higher than their associated constant speed drive output. Therefore there is less transfer of energy from the constant speed drives to the generators, which are being motored.

Page 3-128



MODULE 11.6


ELECTRICAL POWER

Since the energy supplied to the motored generators originates from a generator with a higher speed governor setting, this generator carries more than its share of real load and the motored generators carry less than their share of real load. To equally divide real load among parallel generators, equal amounts of energy must be supplied in the form of torque on the generator rotors.

TO CSD GOVERNOR TO CSD GOVERNOR TO CSD GOVERNOR TO CSD GOVERNOR

MAGNETIC AMPLIFIER MAGNETIC AMPLIFIER MAGNETIC AMPLIFIER MAGNETIC AMPLIFIER

LOAD CONTROL LOAD CONTROL LOAD CONTROL LOAD CONTROL

GEN 1

ERROR DETECTOR

GEN 2

ERROR DETECTOR

GEN 3

ERROR DETECTOR

GEN 4

ERROR DETECTOR

Real Load sharing is controlled by adjustment of the Constant Speed Drive Unit. Figure 83 shows a real load sharing loop.

Real load Sharing Figure 83 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-129


3.28.2 Reactive Load Division Gen A and Gen B are two ac generators operating in parallel. If the excitation (field current) of an ac generator is altered, the output voltage will be altered also. Assume Gen B to be overexcited and thus receiving more field current than Gen A. The output voltage of Gen B will increase and the voltage difference between the generators will cause a circulating current to flow from Gen B to Gen A. Figure 84 shows reactive current circuit.

GEN A

115V

REACTIVE CIRCULATING 125V GEN B CURRENT

Reactive Current Circuit Figure 84 Because of the very inductive nature of the ac generator, this current will be lagging by approximately 90° to the generator‟s emf and so represents a reactive load. Although no real power circulates between the generators, the current causes the generators and their lines to overheat due to power developed within the resistance of the copper conductors. This loss restricts the real power output current of the generator and limits the amount of power or torque available to keep the generators synchronised. The excitation of the generators must therefore be kept equal, so that zero reactive currents flows between them.

Page 3-130



MODULE 11.6


ELECTRICAL POWER

Figure 85 shows balanced reactive loads whereas figure 84 shows unbalanced reactive loads. TOTAL REAL LOAD

TOTAL REACTIVE LOAD

I1 (REACTIVE)

I1 (REAL) I1

(AC TU

AL )

GEN 1

I2 (REACTIVE)

I2 (REAL) I2 (AC TU

AL )

GEN 2

GENERATORS SUPPLYING SAME LOAD AND SHARING LOAD EQUALLY POWER FACTORS EQUAL

Balanced Reactive Loads Figure 85 TOTAL REAL LOAD

I1

(A CT U

GEN 1

AL )

I2 (REACTIVE)

I1 (REACTIVE)

TOTAL REACTIVE LOAD

I1 (REAL)

I2 (REAL) I2 ( AC TU AL )

GEN 2

GENERATOR 1 IS SUPPLYING MORE REACTIVE LOAD THAN GENERATOR 2 GEN 1’s POWER FACTOR DECREASES GEN 2’s POWER FACTOR INCREASES

Un-balanced Reactive Loads Figure 86


Page 3-131


TO PRE-AMP AND GENERATOR SHUNT FIELD TO PRE-AMP AND GENERATOR SHUNT FIELD TO PRE-AMP AND GENERATOR SHUNT FIELD TO PRE-AMP AND GENERATOR SHUNT FIELD

ERROR DETECTOR ERROR DETECTOR ERROR DETECTOR ERROR DETECTOR

GEN 1

MUTUAL REACTOR

GEN 2

MUTUAL REACTOR

GEN 3

MUTUAL REACTOR

GEN 4

MUTUAL REACTOR

Reactive Load Sharing is controlled by adjustment of generator field. Figure 87 shows a reactive load sharing.

Reactive Load Sharing Figure 87 Page 3-132



MODULE 11.6


ELECTRICAL POWER

The mutual reactor is a device capable of sensing reactive load. Note that its primary winding (P) is fed from a current transformer at phase C. Its secondary winding (S) is fed directly from the same phase. The air gap in its core causes a phase shift of 90º between the primary current and the secondary voltage. Figure 88 shows the circuit of the mutual reactor operation.

GENERATOR OUTPUT

REACTIVE LOAD SHARE LOOP

MUTUAL REACTOR VOLTAGE REGULATOR

TO THE TIE BUS

Mutual Reactor Operation Figure 88


Page 3-133


Figures 89 – 92 show how reactive load division is calculated.

Ip

DUE TO THE INDUCTIVE CHARACTERISTICS OF THE MUTUL REACTOR PRIMARY CURRENT SETS UP SECONDARY VOLTAGE 90° TO IT

Vs

Figure 89

SECONDARY CURRENT, FED DIRECTLY FROM PHASE C IS AT 90° TO V SECONDARY AND THUS 180° TO THE PRIMARY CURRENT

Is

Ip

Vs Figure 90

Page 3-134



MODULE 11.6


ELECTRICAL POWER

WITH A POWER FACTOR OF 1 OR UNITY, THE VOLATGE FROM WITH A REACTIVE LOAD PHASE C WILL BE INTHE PHASE VOLTAGE FROM WITH THE SECONDARY CURRENT PHASE C WILL BE IN PHASE WITH THE SECONDARY VOLTAGE

Is

Ip

I PHASE C (SECONDARY)Vc V SECONDARY V PHASE C THE REGULATOR SEES THIS AS TOO HIGH A VOLTAGE AND REDUCES THE FIELD EXCITATION

I PRIMARY

Vs Figure 91

Is

Ip WITH A REACTIVE LOAD THE VOLTAGE FROM PHASE C WILL BE IN PHASE WITH THE SECONDARY VOLTAGE

Vc

THE REGULATOR SEES THIS AS TOO HIGH A VOLTAGE AND REDUCES THE FIELD EXCITATION

Vs

Figure 92


Page 3-135


3.29 EMERGENCY AC POWER GENERATION Emergency A.C. is available in case of all aircraft engine driven generators failing. Emergency A.C. generation is available from either: 1. Standby Generator. 2. Auxiliary Power Unit (APU). 3. Ram Air Turbine (RAT). 3.29.1 Standby Generator Some larger aircraft have standby generators. These generators are a variable speed, variable frequency type and operate when the engines are running. They connect to a back-up generator converter, which makes the standby generator output into a stable 115V, 400Hz. 3.29.2 Auxiliary Power Unit (APU) Boeing 737 Normally located in the aft end of the fuselage, behind the pressure bulkhead and below the horizontal stabilizer and is a single shaft gas turbine. The APU generator is normally identical to the engine generators and will supply 55kVA on the ground or 45kVA as an alternate electrical power source. Figure 93 shows the location of the APU (Boeing 737).

COOLING FAN

APU ENGINE

APU AIR INLET

EXHAUST DUCT COOLING AIR INLET EXHAUST MUFFLER COOLING AIR EXIT

ACCESS DOOR GENERATOR

Boeing 737 APU Location Figure 93 Page 3-136




3.29.3 Ram Air Turbine (RAT) The RAT is used on aircraft to supply an emergency source of hydraulic power, electrical power, or both, in the event of failures. The RAT is stowed in the fuselage, usually in the underbelly or in the lower side of a wing to body fairing. Should the main hydraulic system pressure fall to zero (i.e. all pumps failed), or the electrical generators should all fail, the RAT will automatically deploy into the airstream. Figure 94 shows the RAT from a Boeing 777 aircraft.

FWD PROPELLER/GOVERNER UNIT

INBD G E N E R A TO R O U T P UT

RAT GENERATOR

E L E CT R ICA L C O N NE C T O R S

HYDRAULIC PUMP M O U N TIN G H OL ES

G C U S IG N AL S & POW ER

IN P UT S H AF T

FW D

RAT Generator (Boeing 777) Figure 94


Page 3-137


The RAT consists of the following components: 1.

A variable pitch propeller,

2.

A hydraulic pump,

3.

An electric retract/deploy actuator,

4.

A speed sensing device,

5.

A teleflex lock cable,

6.

Up and down limit switches.

3.29.4 Emergency Pump The ram air turbine drives a hydraulic Pump, which directly powers flight controls and landing gear. The RAT supplies electric loads such as actuators and DC Power loads by driving a hydraulic motor generator. Some aircraft have sufficient battery reserves for minimal functions. 3.29.5 Emergency Generator The ram air turbine drives a generator when the aircraft makes use of electric motor pumps (EMP) for hydraulic functions. Flight controls and landing gear are operated by the RAT through the EMP, while AC and DC power loads are drawn directly off the RAT generator. 3.29.6 Generator And Pump The ram air turbine can drive both the generator and pump if the use of other subsystems for power conversion is not desired. Higher total power availability is achievable by managing pump, or generator loading, at the RAT according to system priorities and turbine capability. 3.29.7 Extended Twin Engined Operations (ETOPS) Certification for ETOPS operation must consider operation of the aircraft with partial loss of power due to one engine out. Additional consideration of loss of main generator or pumps is also a concern in this situation. Therefore ram air turbine operation throughout the envelope in all weather for 120 or 180-minute diversions can help optimize the number of backup subsystems aboard the aircraft.

Page 3-138



MODULE 11.6


ELECTRICAL POWER

3.30 EXTERNAL/GROUND POWER Electrical power is required for the ground operations of aircraft for „turn rounds‟, engine starting, lighting and so on. While the aircraft batteries are capable of supplying these services for a very limited period, they should be conserved for the important role of supplying power under emergency conditions. A separate circuit is therefore incorporated to allow for connection of external power supplies to aircraft. 3.31 DC EXTERNAL POWER Figure 95 shows a 3-pin external power receptacle.

3 PIN EXTERNAL POWER RECEPTACLE EARTH

EXTERNAL SUPPLY SOCKET POSITIVE D.C. 3 PIN PLUG POSITIVE D.C.

ACCESS DOOR

D.C. External Power Plug & Receptacle Figure 95


Page 3-139


A basic external power circuit incorporating a 3-pin socket is shown at Figure 96.

MAIN D.C. BUS

REVERSE CURRENT CIRCUIT BREAKER EXTERNAL POWER RELAY

BATTERY SYSTEM

EXT

BATT

POWER SELECTOR

D.C. External power Schematic Figure 96

Note that there are 2 large pins (carrying the main busbar load) and one small pin, which carries the supply to the coil of the external power relay. This means that if ground power is being applied and the external power plug is withdrawn, the small pin leaves the power receptacle first, thus breaking the supply to the external power relay, which opens its contacts. In this way, the main busbar load is no longer supplied and arcing is prevented as the 2 large pins leave the power receptacle.

Page 3-140



MODULE 11.6


ELECTRICAL POWER

3.31.1 External DC – Multiple Busbar System

MAIN MAIN

AUX

GROUND POWER PLUG

ON

OFF

GROUND SUPPLY MASTER SWITCH

AUX

GROUND SUPPLY CONTACTOR

MAGNETIC INDICATORS

No 1 BUS TIE CONTACT

AUX

MAIN

No 1 DC BUS

ESS DC BUS

No 2 DC BUS

No3 BUS TIE CONTACT

AUX

No 3 DC BUS

In some aircraft D.C. power is distributed from a multiple busbar system and it is necessary for certain services connected to each of the busbars to be operated when the aircraft is on the ground. This requires a more sophisticated arrangement of the external power supply system. Figure 97 shows a schematic of a multiple D.C. busbar system.

Multiple D.C. Busbar System – External Power supply Figure 97 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-141


Referring to Figure 97, it can be seen that in addition to the external supply relay or contactor, there are contactors for tying the busbars together. There are also magnetic indicators to indicate that all connections have been made. When the external power unit is connected to the aircraft and the master switch is selected to “ON”, it energises the external power supply contactor, thus closing its auxiliary and main sets of contacts. One set of auxiliary contacts complete a circuit to the magnetic indicator, which then indicates that an external supply is connected, and on (Indicator “C”). A second set complete circuits to coils in the No 1 and No 3 bus-tie contactors while a third and main heavy-duty set connect the supply direct to the “Vital” and No 2 busbars. When both bus-tie contactors are energized their main contacts connect the supply from the external supply contactor to their respective busbars. Indication that both busbars are also “Tied” to the ground power supply is provided by magnetic indicators “A” and “B” which are energized from the “Vital” busbar via the auxiliary contacts of the contactor.

Page 3-142




3.32 AC EXTERNAL POWER Figure 98 illustrates an ac external power receptacle. There are 4 large pins through which three phases and neutral are fed and 2 small pins, which are fed by dc for operating a ground power contactor. These small pins carry out the same function as the small pin in the dc receptacle in that they ensure the 3 phase and neutral load bearing pins are open circuited before the external power supply plug can be disconnected, thus preventing arcing.

EXTERNAL POWER READY LIGHT

SERVICE INTERPHONE CONNECTION

NOSE WHEEL WELL LIGHTS

A.C. PHASE “A” A.C. PHASE “B”

A.C. PHASE “C”

A.C. NEUTRAL D.C.

A.C. External Power Receptacle Figure 98


Page 3-143


Three phase 400 Hz 115V AC power is supplied to the aircraft through the external power receptacle. Before being connected to the aircraft systems, a Ground Power Control Unit (GPCU) monitors the supply. Although aircraft ground power systems vary from type to type the following parameters are generally monitored: 1. Over voltage 2. Phase sequence 3. Under voltage 4. Over current 5. Over frequency 6. Open phase 7. Under frequency 8. Phase imbalance When these parameters are within specified limits, then ground power is allowed to feed the aircraft systems.

Page 3-144



MODULE 11.6


ELECTRICAL POWER

3.32.1 A.C. External Power Circuit

28V DC FEEDBACK

RESISTOR

3

3

AC GROUND POWER PLUG LIMITING

SELECTOR SWITCH

FREQ VOLTS

EXT

CONTROL RELAY

TRIP CLOSE

GROUND POWER SWITCH

28V DC BUS BAR

1

GROUND POWER AVAILABLE

PHASE SEQUENCE PROTECTION

GROUND POWER TRU

1

3

TRIP

CLOSE

3

3 A.C. MAIN BUS BAR

GROUND POWER BREAKER RELAY

Figure 99 shows a typical external A.C. power supply circuit.

A.C. External Power Supply Circuit Figure 99


Page 3-145


When external A.C. power is coupled to the receptacle a three-phase supply is fed to the main contacts of the external power breaker, to an external power TRU and to a phase sequence protection unit. The TRU provides a 28v D.C. feedback supplies to a hold-in circuit of the ground power unit. If the phase sequence is correct the protection unit completes the circuit to the control relay coil, thus energising it. A single-phase supply is also fed to an amber light, which comes on to indicate that external power is coupled, and to a voltmeter and frequency meter via a selector switch. The circuit is controlled by an external power switch connected to a busbar supplied with 28v D.C. from the aircraft‟s battery system. When the switch is set to the “Close “position current flows across the main contacts of the energized control relay, to the “Close” coil of the external power breaker, thus energising it to connect the external supply to the three-phase A.C. main busbar. Selecting the “Trip” position on the external power switch disconnects the external power supply. This action connects a D.C. supply to the trip coil of the external power breaker, thus releasing its main and auxiliary contacts and isolating the external power from the A.C. main busbar. Figure 100 shows the A.C. external power receptacle and nose wheel bay control panel for the Boeing 737 aircraft.


FLIGHT

SERVICE

PILOT

NOSE WHEELWELL ON NORM

NOT IN USE

CALL LIGHT

Boeing 737 External Power Figure 100 Page 3-146



BATTERY

HOT BATT BUS

ALTERNATE

Boeing 737 External Power Schematic Figure 101

Page 3-147

OFF

AUTO

STANDBY POWER

BAT

NORMAL

DC BUS 2

BATT BUS

T/R 2


BATTERY BUS CONTROL

ALTERNATE

DC BUS 1

T/R 1


TO GEN 1 VIA EPC 1

A.C. EXT PWR

ON

BAT

OFF

T/R 3

115 V AC MAIN BUS 2

EPC2


D.C. EXT PWR

APU START INTERLOCK RLY

BATT CHRG

BATT OVERHEAT RLY

BATT CHRG TRANS RLY

115 V AC GRND SERV BUS

FROM GEN BUS 1

GRND SERV RELAY


MODULE 11.6

ELECTRICAL POWER

Figure 101 shows an external power schematic for the Boeing 737 aircraft.


3.33 B747 ELECTRICAL SYSTEM A detailed schematic diagram of the electrical system for a Boeing 747 commercial transport is shown in Figure 102. 3.33.1 Normal Operation There are many AC and DC buses throughout the Electrical Power Generation System (EPGS). Power from these buses is distributed to all the electrical loads required for flight and ground operations. The Bus Control Units (BCU), Generator Control Units (GCU), and the Auxiliary Generator Control Units (AGCU), receive inputs from the flight deck control panel, current transformers, and current sensors to monitor and control the contactors, breakers, and relays to provide distribution of both AC and DC power to the different buses. The 115 volt AC power from the four IDGs is applied to the four main ac buses; ac bus 1, 2, 3, and 4 through the GCBs. The output of the four IDGs is normally operated in parallel through the Bus Tie Breakers (BTB), sync bus, and the Split System breaker (SSB). For ground operations, provided external power is not available, the APU generators can power the main AC buses through the Auxiliary Power Breakers (APB), sync bus, and the BTBs. The APU generator can not be paralleled through the SSB. External power applied to the airplane supplies the main ac buses through the External Power Contactors (XPC), sync bus, and the BTBs. External power can not be paralleled through the SSB. If only one APU generator or one side of the external power is available, both the right and the left sync bus halves can be powered through the SSB. The 115 volt AC three phase power from ac bus 1, 2, 3, and 4 is supplied to the four TRUs, which supply the inputs of the main DC buses; DC bus 1, 2, 3, and 4. The main DC buses are normally tied together by the DC isolation relays (DCIR) and the DC tie bus. 3.33.2 Ground Handling and Ground Service Systems The AC ground handling bus is energised automatically through the ground handling relays (GHR), provided BCU No. 1 determines the power from either external power No. 1 or APU generator No. 1 is of good quality. External power has priority over auxiliary power if both are connected. The external power TRUs are supplied with 115 volt AC three-phase power from the AC ground handling bus. The 28 volt DC output of the external power TRUs will power the DC ground handling bus.

Page 3-148




The ground service bus receives 115 volt AC three phase power from either AC bus 1, external power No. 1, or APU generator No. 1 through the ground service select relays (GSSRs) and the ground service transfer relays (GSTRs), depending on the source selected by BCU No. 1. External power has priority when connected for ground operations. AC bus 1 powers the ground service bus for electrical loads necessary during normal flight operations. An autotransformer powered by the ground service bus, supplies 28 volt AC single phase power to the 28 volt AC ground service bus. An autotransformer powered by the ac bus 3, supplies 28 volt AC single phase power to the main 28 volt AC bus. The main and APU battery chargers receive power from the ground service bus. The output of the battery chargers power the main and APU hot battery buses, also maintaining the batteries at full charge condition. 3.33.3 Main Standby System During normal power operations, DC bus 3 powers the main and APU battery buses through the battery transfer relay. The AC standby bus is powered by AC bus 3 through the AC standby power transfer relay. If power is lost on AC and DC bus 3, the system switches to standby power, energising the main battery relay, and also de-energising the battery transfer relay and the AC standby power transfer relay. The main and APU battery buses will now be powered by the main and APU hot battery buses. The static inverter will also receive 28 volt DC power from the main battery hot bus. The 115 volt AC single phase power from the static inverter will power the AC standby bus. The captain's transfer bus receives power from AC bus 3 through the instrument bus voltage sense unit (IBVSU), and the first officer's transfer bus receives power from AC bus 2 through another IBVSU. If the voltage is lost on AC bus 3, the IBVSU will automatically switch to AC bus 1 to power the captain's transfer bus. If AC bus 2 has a power loss, the IBVSU will switch the first officer's transfer bus to AC bus 1.


Page 3-149


3.33.4 APU Standby Power System The primary flight displays, navigation displays, and the flight management computers normally receive 115 volt AC power from the captain's and first officer's transfer buses. With a loss of the captain's transfer bus, the APU static inverter will energise and supply power to the left flight display, left navigation display, and the left flight management computer. A power loss of the first officer's transfer bus will result in the loss of the right primary flight display and right navigation display. Power can be restored by the APU static inverter by switching the EFIS power display switch to F/O position. If both the captain's and the first officer's bus have a power loss, then only the captain's or the first officer's displays may be powered at one time. The APU static inverter receives 24 volt DC power input from the APU inverter transfer relay when the system switches to standby power.

Page 3-150


EXT PWR 1


GHR

GHR

AC GHB 2

AC GHB 1

GSTR

GSTR

DC GHB

TRU

APU BATT

APU BATT CHRG

MAIN BATT

MAIN BATT CHRG

AC GSB UTIL

AC GSB

DC GHB

TRU

DC BUS 2

AC BUS 2

DC BUS 1

IDG 1

BTB 2

TRU

DCIR

GCB 1

LEFT SYNCH BUS

TRU

AC BUS 1

28V AC XFMR

BTB 1

APB 1

IBVSU

IDG 2

APU HOT BATT BUS

MAIN HOT BATT BUS

APU INV XFR RLY

DCIR

FO TRNS BUS

GCB 2

MAIN BATT RLY

FO XFR RLY

APU BATT XFR RLY

APU START RELAY

APU STATIC INV

DC BUS 3

TRU

AC BUS 3

DCIR

GCB 3

APU BATT BUS BATT XFR RLY

PRI FLT DISP - R NAV DISP - R

FMC

IDG 3

BTB 4

RIGHT SYNCH BUS

PRI FLT DISP - L NAV DISP - L

MAIN BATT BUS

CAPT XFR RLY

BTB 3

CPT TRNS BUS

IBVSU

DC BUS TIES

SSB

STATIC INV

28V AC XFMR

DC BUS 4

TRU

AC BUS 4

STBY PWR XFR RLY

DCIR

GCB 4

APB 2

XPC 2

AC STBY BUS

28 AC MAIN BUS

IDG 4

APU GEN 2

EXT PWR 2


28V AC GSB

GSSR

GSSR

XPC 1

APU GEN 1


MODULE 11.6

ELECTRICAL POWER

Boeing 747 Electrical System Figure 102

Page 3-151


3.33.5 Electrical System Control Module Switches are provided on the electrical system control module for manually selecting and monitoring the operating status of the functions selected by the operator. Figure 103 shows the Electrical System Power Module.

Electrical System Power Module Figure 103 

Generator Control Switches (Gen Cont)

The generator control switch is a latched, alternate action switch. When depressed to the latched (ON) position, the electrical power generating system is directed to place the 115 volt AC power from the integrated drive generator onto the AC bus. The release (OFF) position removes the integrated drive generator voltage from the ac bus.

Page 3-152






Generator Drive Disconnect Switches (Drive Disc)

The generator drive disconnect switch is a protected momentary action switch that controls mechanical disconnection of the integrated drive generator from the engine gearbox. 

Bus Tie Breaker Switches (Bus Tie)

The bus tie breaker switch is a latched, alternate action switch. When depressed to the latched (AUTO) position, automatic paralleling of the ac buses occurs. The release (ISLN) position isolates the selected bus from the other AC buses. 

APU Generator Control Switches (APU GEN)

The APU generator switch is a momentary switch. Actuation alternately allows APU generator power to be applied or removed from the AC buses. 

External Power Control Switches (EXT PWR)

The EXT PWR switch is a momentary switch. Actuation alternately allows external power to be applied or removed from the AC buses.


Page 3-153


3.33.6 Electrical Synoptic EICAS Display The electrical synoptic display presents graphic views and parameters for generator control, generator drive disconnect status, bus tie, split system breaker, main bus, electrical power flow, and galley and utility bus status. Illuminated combinations of green coloured flow segments within the bus outlines indicate electrical power flow. The electrical synoptic display is selected via the EICAS display select panel located on the glareshield. Figure 104 shows the EICAS Display Select panel and the EICAS synoptic display for the electrical system.

LOWER EICAS DISPLAY Collins

EICAS DISPLAY SELECT PANEL

EICAS Display Select Panel & Synoptic Display Figure 104

Page 3-154






Split System Breaker

The Split System Breaker (SSB) status is indicated pictorially in the open or closed positions. 

Bus Tie

Bus tie breaker status is indicated by square box symbology. If bus tie breaker 14 is open, an amber coloured ISLN message is displayed within a box. If bus tie breaker 1-4 is closed, two white coloured vertical lines overlay the bus outlines and are enclosed within a box. For invalid bus tie data, the respective bus tie symbol is replaced by a low intensity white box. 

Main Bus

A coloured message and outline indicate main bus status. If ac bus 1-4 is on, the bus number message and outline are coloured white. If ac bus 1-4 is off, the bus number message and outline are coloured amber. For invalid main bus data, the outline changes to low intensity white. 

Galley and Utility Bus

Galley and utility bus status is indicated by the display of an associated message located next to the appropriate main bus indicator. If main bus 1-4 is on and the galley bus 1-4 relay is closed, the GALLEY message is coloured green; otherwise, the message is coloured amber. If main bus 1-4 is on and the utility bus 1-4 relay is closed, the UTILITY message is coloured green; otherwise, the message is coloured amber. If galley or utility bus data becomes invalid, the associated message changes to low intensity white. 

Generator Control

Generator control status is indicated by square box symbology. If generator control breaker 1-4 is open, an amber coloured message OFF is displayed within an amber box. If generator control breaker 1-4 is closed, two white coloured vertical lines overlay the bus outline and are enclosed within a white box. For invalid generator control data, the associated symbol is replaced by a low intensity white box. 

Generator Drive

Generator drive disconnect status is displayed by a message enclosed in a coloured box. If a drive overtemperature or low oil pressure condition exists, then an amber DRIVE TEMP/PRESS message is displayed within an amber coloured box. For normal conditions, a white coloured box is displayed. For invalid data, the box changes to low intensity white.


Page 3-155


3.33.7 DC Distribution There are four main transformer rectifier units (TRUs), each TRU powers an associated main 28 volt DC load bus. In addition, under normal operation, the battery buses are supplied from DC bus No. 3. The TRUs are normally operated in parallel through the DC isolation relays (DCIRs), 75 amp thermal circuit breakers, and a DC tie bus. The TRUs are identical and are interchangeable. 

Main Transformer Rectifier Unit

The main TRUs are static devices which convert 3-phase nominal 115/200 volt, 400 Hz ac input power into unregulated 28 volts dc output power for the airplane main dc system. Each TRU employs a transformer with a star connected primary and paralleled star-delta secondary connections. The 6-phase output from the secondaries is connected to a 6-phase full wave rectifier bridge and filtered to produce the desired output. Each TRU has an output voltage of 29 volts at no load and approximately 27 volts at rated load. The four main TRUs are rated for a continuous output load of 75 amps. 

DC Bus Isolation Relay

The DC isolation relay parallels the four main DC buses when de-energized during normal operation. When energized, they isolate their associated main buses. The four DCIRs operate in conjunction with the BTBs and the bus tie switches. 

DC Current Sensor (DCCS)

The DCCS is used in six locations in the airplane DC system to operate in conjunction with the batteries, TRUs, BCUs, and GCUs to provide the required current sensing. They are rated for continuous operation at 180 amps DC. The DCCSs are used for the following functions: 1. Provide TRU current sensing to the GCUs and BCUs for indication on the EICAS maintenance page. 2. Provide monitoring of TRUs for proper operation and relaying this information to the GCUs and BCUs which provide failed TRU indication on EICAS and for triple channel autoland operation. 3. Provide battery current indication on the EICAS maintenance page and battery discharging indication at an advisory level on EICAS via the BCUs.

Page 3-156



PFD - L ND - L

PMG - L

PWR TEST APU BAT RELAY

CAPT XPR RELAY

EPC 1

PFD - R ND - R

BCU 1

AGCU 1

APU GEN 1

APB 1

GCU 1

GCB 1

CAPT XPR RELAY

IDG 1

AC BUS 1

BTB 1

AC 1 DIST

APU BAT XPR RELAY

DC BUS 1

TRU

LEFT SYNC BUS

EFIS SW

FO

CAPT

APU STATIC INVERTER

DC BUS 2

TRU

AC 2 DIST

APU INV XPR RELAY

CAPT XPR

IBVSU

GCU 2

GCB 2

FO XPR

IBVSU

IDG 2

AC BUS 2

BTB 2

SSB

OFF BAT SW

ON

APU BAT HOT BUS

DC BUS 3

TRU

AC 3 DIST

IDG 3

VOLTAGE SENSE RELAY

DC BUS 4

TRU

GCU 3

AC 4 DIST

SINGLE TRU RELAY

MN BAT HOT BUS

AC BUS 3

GCB 3

BTB 3

TO BCU

BAT XPR RELAY

MN BAT RELAY

IDG 4

EPC 2

STBY PWR SW

BAT

OFF AUTO

STBY PWR XPR RELAY

APU GEN 2

AGCU 2

APB 2

BCU 2

STATIC INVERTER

AC BUS 4

GCB 1

BTB 4

GCU 4

RIGHT SYNC BUS

AC STBY

EXT PWR No 2


APU BAT BUS

EXT PWR No 1


MODULE 11.6

ELECTRICAL POWER

Figure 105 shows a schematic of the total electrical system.

B747 Electrical System Schematic Figure 105

Page 3-157


3.34 CIRCUIT PROTECTION A common cause of circuit failure is called a “Short Circuit”. A short circuit exists when an accidental contact between conductors allows the current to return to the source through a short, low-resistance path as shown in Figure 106.

SHORT

HIGH CURRENT FLOW

LOAD

A Short Circuit Figure 106 If the current flow caused by a short circuit at some section of a cable is left unchecked, the heat generated in the cable will continue to increase until something gives way. A portion of the cable may melt, thereby opening the circuit so that the only damage done would be to the cable involved. However, there is a probability that much greater damage would result; heat could char and burn the cable insulation and that of other cables within the loom, and so causing more short circuits and setting the stage for an electrical fire. This failure is prevented by making sure that all insulation on the wires is in good condition and strong enough to withstand the voltage of the power source. Furthermore, all wiring should be properly secured with insulating clamps or other devices so that they cannot rub against any structure and wear through the insulation. To further protect the circuits the installation of protective devices, such as “Fuses” and “Circuit Breakers” are used.

Page 3-158




3.34.1 Fuses A fuse is a thermal device designed primarily to protect the cables of a circuit against the flow of short-circuits and overloads currents. In its basic form, a fuse consists of a low melting point fusible element or link, enclosed in a glass or ceramic casing. This casing not only protects the element, but also localizes any flash, which may occur when “Fusing”. The link or fusible element is made of either: 1. Lead. 2. Lead/Tin. 3. Tin/Bismuth. Or some other low melting temperature alloy. When the current flowing through a fuse exceeds the capacity of the fuse, the metal strip melts and breaks the circuit. The strip must have low resistance, and yet it must melt at a comparatively low temperature. When the strip melts, it should not give off any vapor or gas that will serve as a good conductor, because this would create an arc between the melted ends of the strip. The metal or alloy used must be of a type that reduces the tendency towards arcing. Fuses are generally enclosed in glass or some other heat-resistant insulating material to prevent an arc from causing damage to electrical equipment or other parts of the aircraft. Fuses in aircraft are classified as: 1. Cartridge Type. 2. Plug-in Type. 3. Clip Type. All types are easily inspected, removed and replaced. 3.34.2 Current Limiters Current Limiters are essentially a “Slow-blow” fuse. That is, when the circuit becomes overloaded, there is a short delay before the metal links melts and disconnects the circuit. This is because the link is made of copper, which has a higher melting point than the alloys used in other types of fuse. The current limiter will carry more than its rated capacity and will also carry a heavy overload for a short time. They are designed to be used in heavy-power circuits where loads may occur of such a short duration that they will not damage the circuit or equipment. The capacity of a current limiter for any circuit is so selected that the current limiter will always interrupt the circuit before an overload has had time to cause damage.


Page 3-159


Figure 107 shows some typical aircraft fuses and current limiters.

LIGHT DUTY FUSES

FUSE HOLDER

FU S IB LE E LE M E N T

TE R M IN A LS

HEAVY DUTY FUSES

CURRENT LIMITER

Typical Aircraft fuses & current limiters Figure 107

Page 3-160




3.35 CIRCUIT BREAKERS Circuit breakers, unlike fuses or current limiters, isolate faulted circuits and equipments by means of a mechanical trip device actuated by a bi-metallic element through which the current passes to a switch unit. Figure 108 shows two types of circuit breakers found on aircraft.

TYPICAL CIRCUIT BREAKER

CIRCUIT BREAKER WITH A “MANUAL TRIP” BUTTON

Circuit Breakers Figure 108 They are used for the protection of cables and components and, since they can be reset after clearance of a fault, they avoid some of the replacement problems associated with fuses and current limiters. Furthermore, close tolerance trip time characteristics are possible because the manufacturer, to suit the current ratings of the element, may adjust the linkage between the bi-metallic element and the trip mechanism.


Page 3-161


The mechanism is of the “Trip-Free” type; i.e. it will not allow the contacts of the switch unit to be held closed while fault current exists in the circuit. The design and construction of circuit breakers varies, but in general they consist of three main assemblies: 1. A Bi-metallic element. 2. A contact type switch unit. 3. A mechanical latching mechanism. A push-pull button is also provided for manual resetting after thermal tripping has occurred, and for manual tripping when required to switch off the supply to a circuit or system. The construction and operation of a circuit breaker is shown in figure 109. PUSH-PULL BUTTON

CONTROL SPRING TRIPS MAIN CONTACT

MAIN CONTACT

CONTROL SPRING

LATCH MECHANISM OPERATES THERMAL ELEMENT

THERMAL ELEMENT OPENS LATCH MECHANISM LOAD

SUPPLY

CLOSED CONDITION

TRIPPED CONDITION

Circuit Breaker Operation Figure 109

Page 3-162




In the closed position; current passes through the switch unit contacts and the thermal element, which in thus carries the full current supplied to the load being protected. At normal current values, heat is produced in the thermal element, but is radiated away fairly quickly, and after an initial rise in temperature remains constant. If the current should exceed the normal operating value due to a short circuit, the temperature of the thermal element starts to rise and becomes distorted. This distortion will eventually become enough to release the latch mechanism, allowing the control spring to open the main contact, thus isolating the load from the main supply. At the same time the push-pull button extends, exposing a white band to indicate that the circuit breaker has operated. After the circuit breaker has tripped, the distorted thermal element starts to cool down and reverts to its original state. Once the fault causing the trip has been rectified, the circuit can then be reset using the push-pull button. In a three-phase a.c. circuit, triple-pole circuit breakers are used, and their mechanisms are so arranged that in the event of a fault current in any one or all three of the phases, all three poles will trip simultaneously. Similar tripping will take place should an unbalanced phase condition develop as a result of a phase becoming “Open-circuited”. The three trip mechanisms actuate a common pushpull button. 3.36 REVERSE CURRENT CUT-OUT RELAY A reverse current cut-out relay is used principally in a D.C. generating system either as a separate unit or as part of a voltage regulator (see section 3.11). These circuit breakers are designed to protect power supply systems and associated circuits against fault currents of a magnitude greater than those at which cut-outs normally operate. Furthermore, they are designed to remain in a “Locked-out” condition to ensure complete isolation of a circuit until a fault has been cleared.


Page 3-163


Figure 110 shows an example of a Reverse Current CB.

RUBBER SHROUDED SETTING HANDLE

MANUAL TRIP BUTTON

TERMINAL BLOCK

MAIN TERMINAL (BUSBAR)

MAIN TERMINAL (GENERATOR)

Reverse Current CB Figure 110 It consists of a magnetic unit, the filed strength and direction of which are controlled by a single-turn coil connected between the generator‟s positive output and the busbar via a main contact assembly. An auxiliary contact assembly is also provided for connection in series with the shunt-field winding of the generator. Opening both contact assemblies is controlled by a latching mechanism actuated by the magnet unit under heavy reverse current conditions.

Page 3-164



MODULE 11.6


ELECTRICAL POWER

CONSUMERS

GEN

VOLTAGE REGULATOR

CUT-OUT

LINE CONTACTOR

REVERSE CURRENT CIRCUIT BREAKER

Figure 111 shows the circuit arrangement for a Reverse Current CB.

Reverse Current CB Circuit Figure 111


Page 3-165


3.36.1 Operation Under normal current flow closing of the relay energises the line contactor, the heavy-duty contacts, which connect the generator output to the busbar via the coil, and main contacts of the normally closed reversed current circuit breaker. The magnetic field set up by the current flow assists that of the magnetic unit, thus maintaining the breaker contacts in the closed position. The generator shunt filed circuit is supplied via the auxiliary contacts. When the generator is being shut down, or a failure of its output occurs, the reverse current resulting from the drop in output to a value below that of the battery flows through the circuit. The cut-out relay will operate and de-energise the line contactor, which takes the generator off line. Under these conditions, the reverse current circuit breaker will remain closed, since the current magnitude is much lower than that at which it will operate. If the cut-out or line contactor failed to open, then the reverse current would continue to flow towards the generator, and in addition to its motoring effect on the generator, it would also reverse the generator field polarity. The reverse current passing through the circuit breaker coil would continue to increase, thus its magnetic field strength would also increase until the latch mechanism opened. This would isolate both the main and auxiliary contacts of the circuit breaker. Note; the breaker must be reset after the circuit fault has been cleared.

Page 3-166



MODULE 11.6


ELECTRICAL POWER

3.37 OVERVOLTAGE PROTECTION Overvoltage is a condition which could arise in a generating system in the event of a fault in the filed excitation circuit, i.e. internal grounding of the filed windings or an open-circuit in the voltage regulator sensing lines. Devices are therefore necessary to protect consumer equipment against voltages higher than those at which they are normally designed to operate.

GEN

VOLTAGE REGULATOR

VOLTAGE SENSING COIL

OVER VOLTAGE RELAY

SENSING COIL

SHUNT FIELD

LINE CONNECTOR

D.C. BUSBAR

The methods used vary between aircraft systems and also on whether they supply D.C. or A.C. Figure 112 shows an overvoltage relay method of protection for a D.C. system.

Overvoltage Relay Figure 112 MOD 11 BOOK 2 PART 3 ISSUE 7 - 01/02/11

Page 3-167


3.37.1 Operation The relay consists of a number of contacts connected in all essential circuits of the generator system, and mechanically coupled to a latching mechanism. The mechanism is electromagnetically controlled by a sensing coil and armature assembly, the coil being connected in the generator‟s shunt field circuit and in series with a resistor, the resistance of which decreases and the current through it is increased. Under normal regulated voltage conditions, the sensing coil circuit resistance is high enough to prevent generator shunt-field current from releasing the relay latch mechanism, and so the contacts remain closed and the generator remains connected to the busbar. If an open circuit occurs in the regulator voltage coil sensing line, shunt field current will increase. Because of the inverse characteristics of the relay sensing coil resistor, the electromagnetic filed set up by the coil causes the latch mechanism to release all the relay contacts to the open position, thereby isolating the system from the busbar. After the fault has been cleared, the contacts are reset by depressing the push button.

Page 3-168



MODULE 11.6


ELECTRICAL POWER

3.38 SOLID STATE OVERVOLT PROTECTION Figure 113 shows a solid state overvoltage protection system

TO BUSBAR

GEN

GB TRIP RELAY INVERSE TIME DELAY

S1

OVERVOLTAGE DETECTOR

GENERATOR CONTROL RELAY S2

GENERATOR FIELD

O.V. LIGHT RELAY

28V DC

28V DC

FIELD SUPPLY FROM VOLTAGE REGULATOR

Overvoltgage Protection – (Solid State) Figure 113 The detector uses solid state circuit elements which sense all three phases of the generator output, and is set to operate at a level greater than 130  3 volts. An overvoltage condition is an excitation type fault probably resulting from a loss of sensing to, or control of, the voltage regulator such that excessive field excitation of a generator is provided. The signal resulting from an overvoltage is supplied through an inverse time delay to two solid state swicthes. When switch S1 is made it completes a circuit through the coil of the generator control relay, one contact of which opens to interrupt the generator excitation filed circuit. The other contact closes and completes a circuit to the generator breaker trip relay, this in turn, de-energises the generator breaker to disconnect the generator from the busbar. The making of switch S2 energises the light relay causing it to illuminate the annunciator light.


Page 3-169


The purpose of the inverse time delay is to prevent nuisance tripping under transient conditions. Figure 114 shows the principle of operation of the inverse time delay.

V O L T A G E

160V

GB

TR

IP

145V

130V

0.1 SECS

0.55 SECS

1 SEC

TIME

Inverse Time Delay Operation Figure 114

Page 3-170




3.38.1 List of Abbreviations A.

ABTR CBX APU AC EBTR ACR AIR APC APPR APTC ASC ASCR ASR

autoland bus transfer relay control bus transfer relay AC emergency bus transfer relay autoland control relay autoland inverter relay auxiliary power contactor auxiliary power pilot relay auxiliary power transfer contactor APU start contactor APU start control relay APU start relay

B.

BCR BIC BPC BPO WR BTC BTCR

battery charger relay battery isolate contactor battery power contactor battery power only warning relay bus tie contactor bus tie control relay In GCU 1/ 2/ 3

D.

DC BTC DC EBPC DC EBTC DC EPR DC GR

DC but tie contactor DC emergency bus power contactor DC essential bus transfer contactor DC external power relay DC galley relay

E.

EBTC EIC EPC EPIR EPPR EPRR

essential bus transfer contactor emergency inverter contactor external power contactor external power interlock relay external power pilot relay external power ready relay

GC CR GCU GHBC GHR GLC GLPPR GPCU GPPR GPTC GSBC

galley contactor generator control relay generator control unit ground handling bus contactor ground handling relay generator line contactor galley power pilot relay ground power control unit generator power pilot relay generator power transfer contactor ground service bus contactor

H.

HSOR

high speed oven relay

In equipment panel

I.

IDG

integrated drive generator

On engines

G.


In GPCU In GPCU

In GCU 1/ 2/ 3 In flight compartment

In equipment panel In flight compartment

Page 3-171


O....OIR....

off indicator relay

In relay box 1/ 2

P.

PMG POR PRR

permanent magnet generator point of reference power ready relay

In IDG 1/ 2 and Gen 3

TFR TPR TRU

TRU fault relay TRU power relay transformer rectifier unit 1/ 2/ 3

T.

Note:

Page 3-172

In GCU 1/ 2/ 3

In forward avionics bay

All relays are located in the electrical power centre unless mentioned.



PART 4 CONTENTS 4

LIGHTS ......................................................................................... 4-1 4.1 4.2 4.3 4.4 4.5 4.6 4.7

4.8 4.9 4.10 4.11 4.12

4.13 4.14 4.15 4.16 4.17

EXTERNAL LIGHTING .................................................................... 4-1 INTERNAL LIGHTING ..................................................................... 4-1 BOEING 737 EXTERNAL LIGHTING ................................................. 4-2 NAVIGATION LIGHTS ..................................................................... 4-3 LANDING LIGHTS .......................................................................... 4-4 4.5.1 Retracting Landing Lights.............................................. 4-5 RUNWAY TURN-OFF AND TAXI LIGHTS .......................................... 4-6 ANTI-COLLISION LIGHTS ............................................................... 4-6 4.7.1 Strobe Light Operation .................................................. 4-7 4.7.2 Strobe Light Safety........................................................ 4-7 4.7.3 Rotating Beam Anti-Collision Lights .............................. 4-8 WING ILLUMINATION LIGHTS ......................................................... 4-10 EMERGENCY ESCAPE SLIDE LIGHTS.............................................. 4-10 EXTERNAL LIGHT CONTROL PANEL ............................................... 4-11 CARGO & SERVICE LIGHTING ........................................................ 4-12 FLIGHT COMPARTMENT LIGHTS ..................................................... 4-14 4.12.1 Pillar & Bridge Lighting .................................................. 4-16 4.12.2 Wedge Lighting ............................................................. 4-16 4.12.3 Master Caution/Failure Lights ....................................... 4-17 PASSENGER COMPARTMENT LIGHTS ............................................. 4-18 4.13.1 Passenger Service Unit (PSU) ...................................... 4-19 PASSENGER READING LIGHTS ...................................................... 4-20 ATTENDANT CALL SYSTEM ........................................................... 4-21 EMERGENCY EXIT LIGHTING.......................................................... 4-22 4.16.1 Emergency Lighting Operation ...................................... 4-24 SELF ILLUMINATING SIGNS............................................................ 4-27


Page 1



Page 2



4

LIGHTS

Lighting is important in the safe operation of aircraft and aircraft systems and falls into two main groups: 4.1 EXTERNAL LIGHTING 1. Navigation Lights to mark the extremities of an aircraft and give position reference. 2. Flashing lights to mark the position of an aircraft. 3. Landing and taxiing lights for forward and lateral illumination. 4. Lights to illuminate wings for ice inspection. 5. Illumination to assist in the evacuation of passengers and crew in the event of an emergency landing. 4.2 INTERNAL LIGHTING 1. Lights to illuminate consoles/control panels. 2. Lights for passenger compartments and information signs. 3. Warning lights to indicate system-operating condition. 4. Emergency lighting.


Page 4-1


4.3 BOEING 737 EXTERNAL LIGHTING Figure 1 shows the layout of the Boeing 737 aircraft‟s external lighting.

110° 140°

110°

140°

Boeing 737 External Lighting Figure 1

Page 4-2



4.4 NAVIGATION LIGHTS All aircraft in flight or moving on the ground during the hours of darkness must display the following lights. A GREEN light, at or near the starboard wing tip so that it is visible in the horizontal plane from a point directly ahead through an arc of 110° to starboard. A RED light, at or near the port wing tip so that it is visible in the horizontal plane from a point directly ahead through an arc of 110° to starboard. A WHITE light visible from the rear of the aircraft in the horizontal plane through an arc of 140°. Figure 2 shows the layout of the Boeing 737 aircraft‟s navigation/position lights.

WING TIP NAVIGATION LIGHTS

TAIL CONE STROBE LIGHT

WING TIP REAR POSITION LIGHT

Navigation/Position Lights Figure 2


Page 4-3


4.5 LANDING LIGHTS These lamps provide illumination for aircraft landing and taxiing in conditions of night or poor visibility. The term „Landing and Taxi‟ lamp also covers such equipment as „flare-out lights‟ and „runway turn-off lights‟. The Boeing 737 aircraft has a total of 4 landing lights. Two are the fixed type and are located on the wing leading edges near the fuselage. Two are the retractable type and are located in the outboard flap track fairing. Note: When the retractable landing lights are extended the lights shine forward, regardless of the flap position. The outboard lights provide good visibility under adverse weather conditions and minimise the effect of reflected light into the flight deck. Figure 3 shows landing and turn-off lights Boeing 737 aircraft.

FIXED LANDING & TURN OFF LIGHTS

RETRACTABLE LANDING LIGHT

Landing & Turn-off Lights Figure 3

Page 4-4



4.5.1 Retracting Landing Lights These lamps are extended by means of an actuator via a slipping clutch or shear links. This ensures lamp retraction in the event of it failing to do so at high speed, for whatever reason. The lamp unit is generally of the sealed beam type. Figure 4 shows the typical arrangement of an extending landing lamp. The lamp is shown fully retracted.

E

EXTEND

D A

LIGHT

F

115V 400Hz

M B RETRACT

C G

Retracting Landing Light Circuit Figure 4 When an „extend‟ selection is made, the motor „M‟ is supply through contacts „D‟. As the motor runs, cams „A‟ „B‟ „C‟ rotate. „A‟ rotates to open contacts „D‟ and close contacts „E‟ when the lamp is fully extended. Cam „B‟ closes contacts „F‟ to enable a supply via closed contacts „E‟ to the lamp. Contacts „G‟ close as cam „C‟ rotates to arm the circuit for a „retract‟ selection.


Page 4-5


4.6 RUNWAY TURN-OFF AND TAXI LIGHTS Runway turn-off lights in each wing root inboard of the inboard landing lights. They are aimed ahead and to the side of the aircraft to illuminate taxiway turnoffs. A nose gear taxi light is mounted on the inner cylinder of the nose gear shock so that it turns with the nose gear. 4.7 ANTI-COLLISION LIGHTS Anti-Collision Lights are intended as „attention-getters‟ to warn of the presence of an aircraft and identify its position. Anti-collision lights are mounted on the top and bottom of the fuselage, aft of the wing leading edge. Each anti-collision light is a strobe light covered in a red lens. Note: In addition to the anti-collision lights, some aircraft are fitted with white strobe lights on each wing tip and tail and act as aircraft position indicators. Figure 5 shows the lower anti-collision light (strobe), Boeing 737 aircraft.

Anti-Collision (Strobe) light Figure 5

Page 4-6



4.7.1 Strobe Light Operation The strobe light works on the principle of a capacitor-discharge flash tube. The capacitor converts an input power of 28V dc or 115V ac into a high dc output of around 450V. This discharge occurs between two electrodes in a neon-filled tube and this in turn produces a high intensity flash of light at a rate of approximately 60 flashes per minute. 4.7.2 Strobe Light Safety WARNING: 1. Do not handle the unit for at least 5 minutes after power is removed. 2. Never touch a new flash tube with bare hands. 3. Damage to the eyes may result from looking directly into high intensity light. Figure 6 shows the circuit for Anti-Collision Lighting 115V 400Hz GND SRV

AC BUS 1

OFF

ANTI-COLL CONTROL PANEL

WARNING LIGHT PWR

ANTI-COLLISION TOP NOISE FILTER

TIMING CCT

VOLTAGE DOUBLER CCT

VOLTAGE DOUBLER CCT

STORAGE CCT

STORAGE CCT

ANTI-COLLISION BOTTOM

TRIGGER CCT

Anti-Collision Lighting Circuit Figure 6


Page 4-7


4.7.3 Rotating Beam Anti-Collision Lights There are two types of rotating beam anti-collision light: 1. Rotating Reflector. 2. Rotating Lamp. Figure 7 shows a rotating reflector beacon type anti-collision light.

SPREAD BEAM

NARROW BEAM

Rotating Reflector Anti-Collision Light Figure 7 In this type of light a motor drives a rotating reflector which reflects light from one lamp. The speed of rotation is (typically) 40-45 rpm giving a flashing frequency of 80-90 Hz/min. The reflector has one half flat to emit a narrow high intensity beam, while the other half is curved to increase the vertical spread of the light beam to 30° above and below the horizontal.

Page 4-8



Figure 8 shows a rotating lamp unit type of anti-collision light.

Rotating Lamp Anti-Collision Light Figure 8 This type of lamp employs two filaments mounted in tandem; each pivoted on its own axis. One half of each lamp consists of a reflector and a motor rotates the two lamps through 180°. Since the lamps are set in opposite directions, 180° to each other, then the effect is of a continuously rotating light beam.


Page 4-9


4.8 WING ILLUMINATION LIGHTS Two wing illumination lights are provided for scanning the wings and engines in flight for ice detection purposes. Also used on the ground to illuminate the immediate area. The lights are flush mounted, one on each side of the fuselage forward of the wing leading edge and just above the cabin floor level. 4.9 EMERGENCY ESCAPE SLIDE LIGHTS Four exterior lights illuminate the escape slide areas for the forward and aft entry and forward and aft service doors. An additional four lights illuminate the overwing escape doors and the areas just aft of the wing trailing edge. Figure 9 shows emergency lights from the Boeing 737 aircraft.

OVERWING ESCAPE LIGHT

DOOR ESCAPE SLIDE LIGHT

Emergency Lights Figure 9

Page 4-10



4.10 EXTERNAL LIGHT CONTROL PANEL The control panel for the external lights is normally located on the lower overhead panel. Figure 10 shows the external light control panel for the Boeing 737 aircraft.

RUNWAY TURNOFF

LANDING E X T E N D

RETRACT

ENGINE START

STROBE OFF

TAXI L

L ON R OUTBOARD

APU

OFF

L ON R INBOARD

OFF

R

OFF

OFF

GRD OFF CONT FLT

ON

ON

BOTH IGN L

GRD OFF CONT IGN FLT R

POSITION ON BAT

ANTI COLLISION

WING

WHEEL WELL

OFF

OFF

OFF

ON

ON

ON

OFF ON

ON

ON

START

RUNWAY TURNOFF

LANDING E X T E N D

RETRACT

L ON R OUTBOARD

OFF

L

TAXI

R OFF

OFF

ON

ON

L ON R INBOARD

STROBE OFF

ANTI WHEEL WING WELL POSITION COLLISION ON BAT

OFF

OFF

OFF

ON

ON

ON

OFF ON

ON

External Light Control Panel Figure 10


Page 4-11


4.11 CARGO & SERVICE LIGHTING The purpose of the cargo and service compartment lights is to provide the necessary illumination for cargo handling and for performing all service activities. Dome lights and floodlights are used to provide illumination in the cargo compartments, wheel wells and servicing compartments. The lights can operate from internal power or from an external power source. Light switches are provided within the compartments the lighting serves. Figure 11 shows the distribution of cargo/service lighting for the Boeing 737 aircraft.

Cargo/Service Lighting Figure 11

Page 4-12



Figure 12 shows a simplified diagram of the compartment lights circuit.

OFF

ELEC RACK

LOWER NOSE

ON OFF

E/E

ON

A/C BAY

OFF

RIGHT A/C ON

28V GRND SERV BUS

OFF

LEFT A/C ON

OFF

AFT ACCESS APU BAY

ON

ACCESSORY BAYS

OFF

TAIL CONE

ON

Compartment Lights Circuit Figure 12


Page 4-13


4.12 FLIGHT COMPARTMENT LIGHTS The purpose of the flight compartment lights is to provide illumination for the flight compartment, its instruments, controls and other equipment so that the flight crewmembers can perform their jobs. Dome lights supply general illumination for the cabin. The lightshield provides background lighting for the pilots. Each instrument and instrument panel has integral lighting. The control stand is illuminated from an overhead floodlight. Floodlights illuminate circuit breaker panels. There are also lights for the standby compass and for map lighting. Figure 13 shows the position of the flight compartment light control panels.

FLOOD

DOME WHITE

PANEL BRIGHT

BRIGHT

DIM OFF

OFF

OFF

BRIGHT CIRCUIT BREAKER BRIGHT

OFF

PANEL BRIGHT

OFF

LIGHTS TEST BRT DIM

MAP

PANEL

BRIGHT

BRIGHT

BACKGROUND

MAP

BRIGHT

BRIGHT

AFDS FLOOD BRIGHT

BRIGHT

OFF OFF

PANEL

OFF

OFF OFF

OFF

Flight Compartment Light Control Panels Figure 13

Page 4-14



Figure 14 shows the rear flight compartment and circuit breaker panel illumination.

Rear Flight Compartment & Circuit Breaker Panel Lamps Figure 14


Page 4-15


4.12.1 Pillar & Bridge Lighting The pillar light contains a miniature center contact filament lamp (pea lamp). A single cable carries the supply, while a ground tag completes the circuit for the lamp. An aperture has a filter through which the light is distributed. These lights can be used as single items or in a bridge configuration. 4.12.2 Wedge Lighting Wedge lighting uses two wedges of glass, inner wedge “A” and outer wedge “B”. Light is introduced to wedge “A” by a lamp, some light penetrates directly to the instrument dial while some is trapped within wedge “A” to be distributed down the dial. Light escaping into wedge “B” is reflected down the wedge but is prevented from further escape by none reflective black paint. In this way, the light is retained and illuminates the dial. Figure 15 shows Pillar/Bridge and Wedge type lighting. APERTURE

PILLAR LAMPS

BRIDGE LIGHTING GROUND CONNECTION

LAMP

ELECTRICAL SUPPLY

INNER WEDGE “A”

PILLAR LIGHTING

PILLAR & BRIDGE LIGHTING

INSTRUMENT DIAL OUTER WEDGE “B”

BLACK PAINT

WEDGE LIGHTING

Pillar/Bridge & Wedge Integral Lighting Figure 15

Page 4-16



4.12.3 Master Caution/Failure Lights The master caution/failure system informs the flight crew that a system fault annunciator has illuminated on the forward overhead, aft overhead or fire control panels. The system receives inputs from various fault annunciators to illuminate two master caution lights and one of twelve sections of the master caution annunciators. Both annunciators have a “Push to Cancel”, Push to Recall” function. Figure 16 shows the Master Caution lights for the First Officer and captains position.

ANTI-ICE

ENG

HYD

OVERHEAD

DOORS

AIR COND

MASTER CAUTION

FIRE WARN

PUSH TO RESET

PUSH TO RESET

FIRST OFFICER‟S PANEL

FIRE WARN

MASTER CAUTION

PUSH TO RESET

PUSH TO RESET

FLT CONT

ELEC

IRS

APU

FUEL

OVHT/DET

CAPTIAN‟S PANEL

Master Caution/Failure Panel (B737) Figure 16


Page 4-17


4.13 PASSENGER COMPARTMENT LIGHTS The passenger compartment is illuminated by ceiling and window lights. Entry and threshold lights provide additional lighting for the doorways. Other cabin lighting systems include lavatory lights, reading lights, passenger information signs and attendant call system. Figure 17 shows the layout for passenger lighting.

CEILING LIGHTS DEFUSER PANELS MID AISLE SIGN

EXIT

PASSENGER LIGHT PANEL WINDOW LIGHTS

CEILING LIGHTS

EMERGENCY EXIT DIRECTION INDICATORS

Passenger Compartment Lighting Figure 17

Page 4-18



4.13.1 Passenger Service Unit (PSU) The passenger service unit contains: 1.

Reading Lighting.

2.

Fasten Seat Belt & No Smoking signs.

3.

Attendant Call button.

4.

Air conditioning fans.

5.

Life vest storage.

6.

Oxygen mask storage.

Figure 18 shows the Passenger Service Unit.

PUSH INSIDE PUSH LI

FEVEST

LIFE VEST STORAGE

NO SMOKIN

G

FASTEN SE AT BELT

READING LIGHTS

READING LIGHT SWITCHES

AUDIO SPEAKER OXY MASK STORAGE

Passenger Service Unit (PSU) Figure 18


Page 4-19


4.14 PASSENGER READING LIGHTS Passenger Panel and Fwd Cabin Attendant Panel circuits are shown at Figure 19.

DC BUS 1

DC BUS 2

AC BUS 1

AC BUS 2

16 S DELAY

RH OTHER PASSENGER LTS

LH FLT DECK SW

ATT CALL

N/S LIGHTS

F/S LIGHTS

READ LT TOUCH CONTROL READ LT

CALL

READ

Passenger Panel and Fwd Cabin Attendant Panel Circuits Figure 19 All passengers have a reading light with a touch-control button on a Passenger Panel to switch the light on or off. There is a „Reading Reset‟ switch on the Attendant Panel to switch off all reading lights in one action. When operated, the reading lights supply is broken for 16 seconds only. When the supply is restored, the lights remain off but are ready for „ON‟ selection. A „Read Lights Test‟ switch on a maintenance and test panel enables all reading lights to be switched on for inspection. Note:

Page 4-20

N/S Lights F/S Lights S/R Read Lt

= = =

No Smoking Lights Fasten Seat Belt Lights Set/Reset Reading Lights



4.15 ATTENDANT CALL SYSTEM Figure 20 shows an Attendant Call System circuit.

DC BUS 2

CALL

READ

ATT CALL READ LT TOUCH CONTROL READ LT

N/S LIGHTS

F/S LIGHTS TO/FROM OTHER PASS PANELS

PASS TO ATT PA SYSTEM

LAV TO ATT

ATT TO ATT

ATT CALL LAVATORY

PILOT TO ATT ATT TO PILOT

CREW CALL SYSTEM

Attendant Call System Circuit Figure 20 AN ATTENDANT CAN BE CALLED FROM EACH PASSENGER STATION THROUGH THE „ATT. CALL‟ TOUCH BUTTON. THIS ACTION BRINGS ON A LIGHT IN THE „AREA CALL LIGHTS AFT‟ AND „FWD‟ STATIONS. A CHIME ON THE P/A SYSTEM ALSO ALERTS THE ACTION. RESET IS ACHIEVED BY A SECOND TOUCH AT THE „ATT CALL‟ SWITCH.

Operation of a switch in either lavatory will call an attendant in similar fashion and from which lavatory the call was made is identified at the „Area Call Lights Aft‟ and „Fwd‟ stations.


Page 4-21


4.16 EMERGENCY EXIT LIGHTING Emergency lights automatically illuminate exit signs and egress paths when normal lighting system power is lost. They are powered by battery packs and are located in the flight and passenger compartments. The system also includes the external lights used to illuminate the escape routes form the doors and overwings hatches. Floor proximity lighting provides visual; guidance for cabin evacuation when all sources of cabin lighting above four feet is obscured by smoke. They are positioned on the left-hand side of the aisle and have an illuminated arrow spaced every 40 inches to indicate the direction to the nearest exit. Figure 21 shows the position of the emergency exit signs and floor proximity lighting.

FLOOR TRACK LIGHTING

EXIT SIGNS

Emergency Exit Signs & Floor Proximity Lights Figure 21

Page 4-22



Figure 22 shows details of the floor proximity lighting. LENS COVER

LAMP ASSEMBLY

TRACK COVER

FLOOR TRACK

Floor Proximity Lighting Figure 22 Figure 23 shows over-wing escape hatch exit lights.

EXIT

EXIT INDICATOR

Over-wing Escape Hatch Lights Figure 23


Page 4-23


4.16.1 Emergency Lighting Operation The emergency lights (6V) are fed from battery packs, which under normal conditions are trickle charged from the aircraft main electrical system. The 6V lights operate as for the following conditions. A three-position switch located on the flight deck controls the system. The three positions are: 1.

OFF.

2.

ARMED.

3.

ON.

ARMED Is the normal in-flight position of the switch. A warning light, „NOT ARMED‟ is displayed if the switch is in the „OFF‟ or „ON‟ position during flight. ON The battery packs are not charged. 6V lights „ON‟ even though main electrical power available and normal lights are also „ON‟. OFF The battery packs are charged. 6V lights remain „OFF‟ even in the event off main electrical power loss.

Page 4-24



AC BUS 2 OFF IND

AC BUS 1 OFF IND

GSBC 1

28V DC

STANDBY LIGHTS

DC BUS 1

CABIN/ ENTRANCE

28V DC

EXIT SIGNS PASS DOOR LT ESCAPE HATCH LT CEILING LT

TOILETS

EMER POWER

EMER DC BUS

DN R LG UP

EXIT LTS

ON EMER LIGHTS

UP

BATTERY POWER SUPPLY UNIT (BPSU)

NORM

C

C

A

Emergency Lighting System Figure 24

Page 4-25

B

B

SCR 1

ARMED

EMERGENCY LIGHTS

A

Q1

OFF

K

K

ON



Figure 24 shows an Emergency Lighting system circuit.


The following explains the operation of the circuit in Figure 24 with the selector switch in the ARMED position. The battery power supply units supply electrical power to the emergency lights (6Volt) and the inverters of the floor proximity emergency lighting when the transistor Q 1 is on (conducting). Transistor Q1 comes on when the gate in the supply unit has: 1.Logic “zero” at the inverting input and; 2.Logic “one” at the non-inverting input To have logic “one” at the non-inverting input, the Silicon Controlled Rectifier SCR 1 must be on. This occurs when there is a positive voltage on pin B (ARM). As long as 28 Volt is on pin A (CHARGE and HOLD-OFF), the internal batteries are charged and logic “one” at the inverting input of the gate keeps the transistor off. When the voltage is removed from pin A and B, there is a logic “zero” at the inverting input and a logic “one” from the internal batteries at the non-inverting input. This turns the transistor on and the emergency lights and floor proximity emergency lighting comes on. The only way to switch off the emergency lights is to put 28 Volt on pin C (DISARM). This reverse biases the Silicon Controlled Rectifier SCR 1. Note: The voltage at pin B, necessary to “arm” the emergency lights, can also be supplied by the battery power supply unit itself, through pin K (6 Volt) and the attendant switch or EMER LIGHT rotary selector in the ON position. WARNING Minimum use of the battery packs (testing etc) must be made. The battery packs take up to 20 hours to recharge.

Page 4-26



4.17 SELF ILLUMINATING SIGNS The only possible hazard attendant upon the use of such signs is that due to inhalation or absorption into the body of gas released in the event of breakage of the glass envelope. Tritium gas is mildly radioactive; therefore, the signs should be handled carefully to avoid breakage. Should breakage occur, the aircraft should be evacuated and all doors left open to allow maximum ventilation. Disposal of broken signs are subject to the Radioactive Substances Act 1960 and the Radioactive Substances (Luminous Articles) Exemption Order 1962 and should, therefore, be returned to the manufacturer for disposal. All selfilluminating signs should be checked for luminosity level on initial fitting and at periods specified in the relevant maintenance schedule. Such signs usually have a scrap life of 5 years and should then be returned to the manufacturer for disposal.


Page 4-27



Page 4-28



PART FIVE CONTENTS 5

ON BOARD MAINTENANCE SYSTEMS ..................................... 5-1 5.1

5.2 5.3

5.4

5.5 5.6

5.7

MULTI FUNCTION COMPUTER SYSTEM (MFC) ................................ 5-1 5.1.1 Function ........................................................................ 5-1 5.1.2 Maintenance Panel ....................................................... 5-3 5.1.3 Built-In Test Equipment (BITE) ...................................... 5-6 5.1.4 Operation ...................................................................... 5-10 DATA LOADING ............................................................................ 5-11 5.2.1 Navigation Data Base ................................................... 5-12 STRUCTURE MONITORING ............................................................. 5-13 5.3.1 Low Cycle Fatigue......................................................... 5-14 5.3.2 Health & Usage Monitoring (Hum) ................................. 5-14 5.3.3 Structural Monitoring ..................................................... 5-15 CENTRAL MAINTENANCE COMPUTING SYSTEM (CMCS) ................. 5-17 5.4.1 Flight Deck Effect (FDE)................................................ 5-18 5.4.2 Maintenance Access Terminal (MAT)............................ 5-18 PORTABLE MAINTENANCE ACCESS DEVICE (PMAT) ...................... 5-21 AIRPLANE CONDITION MONITORING SYSTEM (ACMS) .................... 5-22 5.6.1 Airplane Condition Monitoring Function (ACMF) ........... 5-24 5.6.2 Quick Access Recorder (QAR) ...................................... 5-25 AIRPLANE INFORMATION MANAGEMENT SYSTEM (AIMS) ............... 5-26 5.7.1 Flight Compartment Printing System ............................. 5-28


Page 1



Page 2



5

ON BOARD MAINTENANCE SYSTEMS

On board maintenance systems enable the engineer to confirm faults and in some cases go straight to the defective item, thus saving time and money in the maintenance of aircraft. There are many different on board maintenance systems in use on modern aircraft, ranging from a simple magnetic indicator on an LRU, to complex systems that allow engineers to connect laptop computers to down load system parameters and fault data. 5.1 MULTI FUNCTION COMPUTER SYSTEM (MFC) In flight monitoring and ground test capabilities are provided by the MFC system (as fitted to the ATR 72). It consists of two independent computers MFC1 and MFC2. The use of these two computers has meant the removal of a total of 9 redundant LRUs. Each computer includes two independent modules, Module A & B. Each Module receives signals from all the various systems and system controls. They also include a self-test capability so that each module can be tested to ensure it is operating correctly. 5.1.1 Function After processing the input information, the unit will output to the various systems to: 1. Monitor, control and authorise operation of the aircraft systems. 2. Manage system failures and flight envelope anomalies and command triggering of associated warning in the "Crew Alerting System" (CAS). 3. Provide readout of BITE memory via a maintenance panel on the flight deck, giving information of any system failures.


Page 5-1


Figure 1 shows a simplified block diagram of the MFC system.

FAULT ACTIVATE

FAULT ACTIVATE

MFC 1

MFC 1A STATUS

MFC 1B STATUS

INPUTS

INPUTS

MFC 1A

MFC 1B

OUTPUTS

PRIMARY SECONDARY

OUTPUTS

ELECTICAL POWER

ELECTICAL POWER

PRIMARY SECONDARY

MFC Block Schematic Diagram Figure 1

Page 5-2



5.1.2 Maintenance Panel The ATR 72 maintenance panel (located right-hand console), enables the operator to identify faults on the system using a rotary switch and a failure display. The control panel (located on the overhead panel) allows the switching on and fault monitoring of the MFC system. Figure 2 shows the MFC Maintenance and control panels.

MFC 1A

1B

2A

2B

FAULT

FAULT

FAULT

FAULT

OFF

OFF

OFF

OFF

MFC CONTROL PANEL (OVERHEAD) BITE ADV DISPLAY

8

4

2

F F

MFC

1

DATA BUS

F F

BITE LOADED

NORM FLT WOW & L/G

ERS MFC

DOORS

3

BOOTS

PTA/ERS MISC

2

MAG IND TEST

NAV 1

BRK FLT CTL

MFC MAINTENANCE PANEL (OVERHEAD)

MFC Maintenance & Control Panels Figure 2


Page 5-3


The Maintenance panel has the following functions: Bite Loaded Indicator - Indicates when a fault has been recorded by the maintenance system. System Selector Switch - Normally placed in the NORM FLT position. During maintenance operations, enables the various systems to be selected and the relevant failure codes displayed. Bite Advisory Display - Indicates, through illuminated lights, the code of the failure recorded. Combination of illumination of these lights enables up to 14 failures per system to be coded. PTA/ERS push-button - PTA function (push to advance) enables recorded failures on selected system to be run. At the end of the selected system test FFFF is displayed. It also acts as an "Erase" function; this will clear current faults from the system. They will be stored in the systems non-volatile memory. Test push-button - Used to check operation of the "BITE LOADED" magnetic indicator. Data Bus connector - Enables the connection of the Maintenance Test Set system to be connected. This enables the down load of all faults onto a Notebook type computer.

Page 5-4



The failure codes are all listed in the aircraft maintenance manual. Table 1 shows an example of the code/failure relationship.

SYSTEM: WOW/L/G CODE 1 2 3 4 5 6 7 8 9 A B C D E

8

4

2 F F

F F F F F F F F F F F F

1 F F F

F F

F F

F F F F F F

F

DEFINITION Right Main Gear Prime DnLk Prox Switch Fail Nose Gear Prime DnLk Prox Switch Fail Left Main Gear Prime DnLk Prox Switch Fail Right Main Gear Sec DnLk Prox Switch Fail Nose Gear Sec DnLk Prox Switch Fail Left Main Gear Sec DnLk Prox Switch Fail Left Main Gear WOW 1 Prox Switch Fail Nose Gear WOW 1 Prox Switch Fail Right Main Gear WOW 1 Prox Switch Fail Left Main Gear WOW 2 Prox Switch Fail Nose Gear WOW 2 Prox Switch Fail Right Main Gear WOW 2 Prox Switch Fail

F F F

F

End of list for selected system

Failure Codes - De-icing Boots System Table 1


Page 5-5


5.1.3 Built-In Test Equipment (BITE) Large aircraft often incorporate "built-in Test Equipment" (BITE) systems to monitor and detect faults in a variety of aircraft systems. Before BITE systems, faults finding often required the connection of special “Test Equipment” then lengthy tests to establish where the fault lay. Then the rectification by replacing the required Line Replacement Unit (LRU) followed by a functional test to confirm the system serviceability, and finally, the removal of the test equipment. The use of BITE systems reduces the time-spent fault finding and thus eliminates the need for specialist test equipment. The BITE continuously tests the various systems and stores all fault information to be recalled later, either by the flight crew or a maintenance team. Once the appropriate repair has been made, the BITE system can then be used to reset the system for operation. Most BITE systems are capable of isolating system faults with at least 95% probability of success on the first attempt. The introduction of digital systems on the aircraft has made BITE systems possible. Discrete digital signals are used as the code language for BITE systems. The BITE system interprets the various combinations of digital signals to determine a system's status. If an incorrect input value is detected, the BITE system records the fault and displays the information upon request. This information may be by illuminating a number of Light Emitting Diodes (LED's), or, as with modern systems, a display on a CRT or TV display. A complex BITE system is capable of testing thousands of input parameters from several different systems. Most BITE systems perform two types of test programs: 1. Operational Test 2. Maintenance test Normal operational checks start with initialisation upon switch on of system power supplies.

Page 5-6



Figure 3 shows the BITE flow sequence.

POWER POWER UP UP RESET RESET

PROTECTION PROTECTION

INITIALIZE INITIALIZE

CONTROL CONTROL

INPUT INPUT

OUTPUT OUTPUT

OPERATIONAL OPERATIONAL BITE BITE

BITE Flow Diagram Figure 3


Page 5-7


The operational BITE program is designed to check: 1. Input signals. 2. Protection circuitry. 3. Control circuitry. 4. Output signals. 5. Operational BITE circuitry. During normal system operation, the BITE monitors a "Watchdog" signal initiated by the BITE program. This watchdog routine detects any hardware failure or excessive signal distortion, which may create an operational fault. If the BITE program detects either of these conditions, it automatically provides isolation of the necessary component, initiates warnings and records the fault in a Nonvolatile memory. The maintenance program of the BITE is entered into only when the aircraft is on the ground and the "Maintenance Test" routine is requested. On aircraft fitted with Flight Management System FMS, a more complex BITE system is provided. In the Boeing 737, the FMS BITE provides fast and accurate diagnosis of the main FMS components.

Page 5-8



Figure 4 shows the Boeing 737 FMS Bite System.

Boeing FMS BITE System Figure 4


Page 5-9


5.1.4 Operation Self-contained In-flight monitoring and ground test capabilities are provided for the main FMS components. Each major FMS component contains comprehensive tests for itself, its sensor inputs, and other interfaces. In-flight data is automatically stored for analysis on the ground through the BITE system. BITE is controlled via the FMS Control Display Unit, CDU. The FMS display will display (in plain English), system status for all systems under test. The operator simply selects from a menu of test options and inputs interactive responses via the CDU. BITE runs the test and provides corrective action diagnostics. The system is designed for line maintenance fault isolation to a single line replacement unit (LRU), within minutes. The BITE system will also carry out system verification; to check interfaces after corrective maintenance action.

Page 5-10



5.2 DATA LOADING Navigation information required by the aircraft systems is loaded using "Data Loaders". These loaders are capable of downloading thousands of bytes of information into the required system in a matter of seconds. The validity of the current data loaded into an aircraft can be checked using the FMS CDU, which will show the current version, loaded into it. Figure 5 shows a Data Loader as fitted to the Boeing 737

DISK STORAGE

DISK STORAGE

429 BUS INTERFACE

POWER PROG

CHNG

COMP

RDY

XFER

R/W

FAIL

SPARE FUSE

PROG CHNG COMP RDY XFER R/W FAIL

DATA TRANSFER IN PROGRESS DATA CHANGE IS REQUIRED DATA TRANSFER IS COMPLETE UNIT READY FOR OPERATION DATA TRANSFER FAILURE UNABLE TO ACCESS DISK DATA SYSTEM TEST FAILURE

LINE FUSE ON/OFF

Boeing 737 Data Loader Figure 5


Page 5-11


5.2.1 Navigation Data Base The Navigation database (NDB) contains data that describes the environment in which the aircraft operates. The type of information loaded includes: 1. Approaches. 2. Country Name. 3. Waypoints. 4. Airports. 5. Runways. 6. Marker Beacons. 7. Holding Patterns.

This information is used by the Flight Management Computer (FMC), to create flight plans that define the aircraft route from origin to destination. The source data and the NDB are updated on a 28-day cycle that it corresponds to the normal revision cycle for navigation charts. Each update disk contains the data for the current cycle and the next one. This arrangement provides the user with greater flexibility since it is not necessary to load a new disk on a specific day. Each PCMCIA card contains 8 megabytes of storage.

Page 5-12



5.3 STRUCTURE MONITORING Structural (Health), monitoring and usage monitoring have evolved over the years to improve the methods of monitoring critical aircraft components. Structural monitoring was first applied to the monitoring of aircraft engines. This was for two primary reasons: 1. To prevent engine damage and possible hazard to the aircraft following a catastrophic failure. 2. The detection of failures before any real damage has occurred. The engine was monitored for: 1. Engine Speeds. 2. Engine Temperatures. 3. Engine Pressures. 4. Engine Vibration. Figure 6 shows Engine monitoring set up.

ENGINE MONITORING RECORDER

OFF AIRCRAFT ANALYSIS

Engine Structural Monitoring Figure 6


Page 5-13


The first British application of engine health monitoring was on BEA Trident aircraft. Initially simple discrete inputs giving the engine parameters were monitored and recorded on a suitable on-board recorder. This data was then removed from the aircraft for the necessary analysis using ground-based equipment. 5.3.1 Low Cycle Fatigue The availability of cheap microprocessors in the early 1970's enabled a further development to be embodied. This has allowed more precise measurements and calculations to be made of which "Low Cycle Fatigue" is a typical example. The Low Cycle Fatigue Counter (LCFC), receives inputs from the engine for such parameters as engine speed (NL and NH) of compressors and turbines. It then processes this information to calculate engine damage cycles. These damage cycles are not related to actual damage but more a measure of the component life being consumed by these critical items. 5.3.2 Health & Usage Monitoring (Hum) Typical parameters monitored by modern HUM systems are: 1. Engine Speed. 2. Engine Temperature. 3. Engine Pressure. 4. Engine Torque. 5. Accelerations. 6. Vibration Levels. 7. Aircraft Stress.

The extension of HUMS is extended to the monitoring of gearboxes and transmission trains on helicopters where the continued operation of the power train is essential to airworthiness.

Page 5-14



A typical HUMS is shown at Figure 7. The engine and other health parameters are conditioned and converted into suitable digital format for use by the microprocessor. After the necessary calculations and algorithms have been executed the data is stored in non-volatile memory until conclusion of the flight. The data is then extracted by means of a suitable "Data Transfer Unit" (DTU).

PROCESSOR

MONITORED PARAMETERS

INPUT SIGNAL CONDITIONING

HUMS ALGORTHIMS

DATA BUS INTERFACE

NON-VOLATILE MEMORY

HUMS

DATA TRANSFER UNIT

OFF AIRCRAFT ANALYSIS

HUMS Figure 7 5.3.3 Structural Monitoring When an aircraft comes into service, the manufacturer will have calculated its life as the number of cycles (take-off - flight - landing) it will achieve. This is normally in the region of 10s of thousands of cycles, with say an average flight time. The aircraft’s life in hours is calculated by taking an average flight time, this could be 1½ hours. Therefore to calculate the aircraft’s life in hours simply multiply the average flight time by the number of cycles, e.g. 1½ hours X 60,000 cycles = 90,000 hours life.


Page 5-15


Manufacturers will also test a sample airframe to destruction, simulating the effects of flight, 24 hours a day until the airframe fails. Structural testing could also be carried out on actual aircraft in service. Strain gauges, positioned at various points on the airframe, measure the structural stress on the airframe. This information is gathered by an on board computer for analysis after every flight. With the requirements for modern aircraft having flight data recorders, these can also be used to monitor the aircraft’s structure and thus identify any faults before they cause catastrophic failure. Figure 8 shows the FDR system.

FDR ARINC 429 ANALOGUE

AIRCRAFT SYSTEMS

ANALOGUE DISCRETES

ARINC 573

ARINC 629

FAULT MONITORING

DFDAU AIMS

FDR System Figure 8

Page 5-16



5.4 CENTRAL MAINTENANCE COMPUTING SYSTEM (CMCS) The CMCS supports both line and extended maintenance functions through menu selections on the Maintenance Access Terminal (MAT) or Portable Maintenance Access Terminal (PMAT). Other menu selections support special maintenance functions, on-line help and report production. Figure 9 shows the location of the MAT.

MAT KEYBOARD

MAT KEYBOARD SLOT

MAINTENANCE ACCESS TERMINAL (MAT) FLIGHT COMPARTMENT REAR RIGHT SIDEWALL

Maintenance Access Terminal (MAT) Figure 9


Page 5-17


The CMCS is used for: 1. Monitoring the aeroplane’s systems for faults. 2. Processing fault information. 3. Supplying maintenance messages. 4. Monitoring flight deck effects (FDE). Maintenance messages give the engineers detailed fault information to help in troubleshooting. The Aeroplane Condition Monitoring System (ACMS) monitors for any system faults, if a fault is detected, a maintenance message is sent to the CMCS. The CMCS processes the data and shows a maintenance message so the maintenance crew can examine it and find a corrective action. 5.4.1 Flight Deck Effect (FDE) FDE inform the flight and ground crews of the conditions relating to the safe operation of the aircraft. The ground crew must find the cause of an FDE to find the corrective action. The FDE data is used along with the aircraft’s maintenance manuals to isolate the fault. The ACMS monitor conditions related to the loss of a system or function. If a condition exists that requires repair or deferral, the ACMS sends FDE data to the AIMS Primary Display System (PDS). The PDS will show the FDE on the MAT and PMAT. 5.4.2 Maintenance Access Terminal (MAT) The MAT has a display screen and controls for selecting and viewing fault data. A keyboard is also provided (stored when not in use) which allows certain entries and controls displayed data. The MAT also has a cursor control device, which has a power supply module that receives 115V ac via the “MAINT ACCESS TERMINAL” circuit breaker located on the overhead panel. This PSM then distributes power for the remainder of the MAT. The cursor control device contains the following controls: 1. Track Ball. 2. Selection Keys. 3. Brightness Control.

Page 5-18



Figure 10 shows the MAT and cursor control device.

MAT DUAL DISK DRIVE

MAT DISPLAY

MAT CURSOR CONTROL DEVICE

SELECTION KEYS (3) TRACK BALL

POWER SUPPLY MODULE

BRIGHTNESS CONTROL

CURSOR CONTROL DEVICE

MAT & Cursor Control Device Figure 10


Page 5-19


Figure 11 shows the MAT display showing FDE data.

LINE MAINTENACE

EXTENDED MAINTENANCE

OTHER FUNCTIONS

HELP

N77701 TBC1234 KBFI/KMWH LEG STATRT WAS 1753Z 07 JUL 00 THIS DATA IS FROM LEFT CMCF

INBOUND FLIGHT DECK EFFECTS Select text of Maintenance Message, then select the MAINTENANCE MESSAGE DATA button to get more data.

MAINTENANCE MESSAGE DATA

Flight Deck Effects recorded during the present leg

FDE: F/D ZONE TEMP CTRL

STATUS

Fault Code : 216 011 00

FDE: CAPT RA FLAG

Maintenance Message: 34-42011 Approach

NOT ACTIVE 1948z 07JUL00

PFD FLAG

Fault Code : 343 311 31

REPORT

ACTIVE 1948z 07JUL00

ACTIVE 1941z 07JUL00

Radio Altimeter Transceiver (left) has an internal fault.

GO BACK

ERASE FAULT

MAT Displayed Data Figure 11

Page 5-20



5.5 PORTABLE MAINTENANCE ACCESS DEVICE (PMAT) The PMAT is stored within the electronics bay and has the same functions as the MAT. There is a PMAT terminal receptacle located on the MAT position. There are also four other PMAT receptacles located throughout the aircraft. These are located: 1. Electronics Bay. 2. Nose Gear. 3. Right Main Gear Bay. 4. Stabilizer Bay.

Figure 12 shows a PMAT and receptacle.

PMAT

SELECTION SWITCHES

POWER SWITCH

CURSOR CONTROL

PMAT RECEPTACLE

LCD DISPLAY

KEYBOARD DISK DRIVE

Portable Maintenance Access terminal (PMAT) Figure 12


Page 5-21


5.6 AIRPLANE CONDITION MONITORING SYSTEM (ACMS) The ACMS (Boeing 777) collects monitors and records data from the aircraft’s system. The data collected by the system is used to produce reports. These reports are used to: 1. Analyse airplane performance. 2. Analyse trends. 3. Report significant events. 4. Troubleshoot faults. Figure 13 shows the layout of the Boeing 777 ACMS.

AIRPLANE CONDITION MONITORING SYSTEM (ACMS)

ACMS REPORTS ACMS REPORTS ACMS XXXX REPORTS XX X XX XXXXXXX XXXXX XXXX XX X XX XXXXXXX XXXXX XXXX XX XXXXXXX XXXXX XXXX XX XX X XX XXXXXXX XXXXX XXXX XX XXXXXXX XXXXX X X XXXXXXXXXXXXX XX XXXX XX XXXXXXX XXXXX XXXXXXX XXXXXXXXXX XXXXXXXXXXXXX XX X XXXXX XXXXXXXXXX X XXXXXXXX XXXXXXXXXXXXX XXXXXXXX XXXX XXXX XXXXXXXXXX XXXXXXX XXXXXXX XXXXXXXX XXXX XXXX XXX XX XXXXXXXXX XXXXXXX XXXXXXXX XXXX XXXXXXX XXX XX XXXXXXXXX XXXXXXX XXXXXXXXXX XX XXXXXXXXX XXXXXXXXXXXXXX X X X X XXXXXXX XXXXXXX X X X X XXXXXXXXXXXXXX XXXXXXXX X X X XXXXXXX XXXXXX XXXXXXXXXXXXXX XXXXXXX XXXXXX XXXXXXXXXXXXXX XXXXXXXXXXXX XXXXXXX XXXXXX XXXXXXXXXXXX XXXXXXX

ACMF

XXXXXXXXXXXX XXXXXXX

PDF CMCF

QAR

AIMS FMCF

DCMF

TMCF

FDCF

TA DA

DFDAF

Boeing 777 ACMS Figure 13

Page 5-22



The ACMS receives data from the Airplane Conditioning Monitoring Function (ACMF) which is located in the left-hand AIMS cabinet. The ACMS is accessed through formats on the Maintenance Access Terminal (MAT), Portable Maintenance Access Terminal (PMAT) or the side displays on the flight deck. The ACMS can if required be programmed by the user airline to carry out custom features. Figure 14 shows the general arrangement of ACMS.

RH DISPLAY LH DISPLAY

QAR

MAT

FLIGHT COMPARTMENT PRINTER

PMAT

A I R C R A F T

FLIGHT CONTROL ARINC 629 BUS (3)

SDU

VHF TX/RX

SYSTEMS ARINC 629 BUS (4) ARINC 429 ANALOG DISCRETES

LEFT HAND AIMS CABINET

ACMS (Boeing 777) Figure 14


Page 5-23


5.6.1 Airplane Condition Monitoring Function (ACMF) The ACMF is a combination of standard and custom software. The custom software is set to the following functions: 1. Report Format. 2. Report Content. 3. Triggers. Triggers are logic equations that detect conditions and cause data to be recorded, e.g. engine exceedances. The ACMF sends data to the following units: 1. Quick Access Recorder (QAR). 2. Maintenance Access Terminal (MAT). 3. Portable Maintenance Access Terminal (PMAT). 4. MAT or PMAT disk drives (to record data onto diskette). 5. Flight deck Side Displays (SD). 6. Data Communication Management Function (DCMF). Note: The DCMF is used to send data to the airline base while the aircraft is airborne via either the VHF communication or Satellite communication system. The ACMS collects data to record and sends reports to many output devices. The MAT and PMATs allows the user to see the ACMS data and control the function of the ACMS. Aircraft systems send data into the AIMS cabinet input/output modules on: 1. Flight Control ARINC 629 Buses. 2. System ARINC 629 Buses. 3. ARINC 429 Buses. 4. Analogue Inputs. 5. Discrete Inputs.

Page 5-24



5.6.2 Quick Access Recorder (QAR) The QAR records data sent from the ACMF onto a 3.5 inch 128 MB optical disk and holds 41 hours of data. A spare disk is located within the unit should the active disk become full. Figure 15 shows a QAR and optical disk.

PRESS SPARE DISK POWER ON

OPTICAL QAR

DISPLAY DISPLAY

PENNY & GILES

FAIL

LOW CAPACITY

MAINTENANCE

EJECT

MADE IN U.K.

OPTICAL DISK CARTRIDGE

QUICK ACCESS RECORDER

Quick Access Recorder (QAR) Figure 15 The optical disk has a magnetic surface with an infrared laser optically tracking the disk. Data from the ACMF (Core Processing Module, CPM) is received by the QARs CPU. The CPU does a self-test to check the validity of the data and then sends control information to the memory device.


Page 5-25


The QRA memory device contains two memories: 1. Flash memory (non-volatile). 2. Formatter memory. The flash memory holds configuration data, system data and identification files and sends this data to the formatter. The formatter arranges the received data, then sends it to the cartridge drive circuits. The cartridge drive circuits control the position of the laser tracking recording head. They also write data on and read data from the optical disk. The front keyboard is used to read information from the optical disk and to run functional tests. The CPU also sends data to the 16 bit LCD displays. These displays show: 1. Stored data. 2. QAR menus. 3. Test results. 4. Messages. The QAR sends data and status to the CPM/COMM in the left AIMS cabinet. The ACMF monitors the data and status. 5.7 AIRPLANE INFORMATION MANAGEMENT SYSTEM (AIMS) The AIMS collects and calculates large quantities of data and manages this data for several integrated aircraft systems. The AIMS has software functions that do all the calculations for each aircraft system. The AIMS has two cabinets, which do the calculations for these systems. Each cabinet contains: 1. Cabinet Chassis. 2. Four input/output Modules (IOM). 3. Four Core Processor Modules (CPM). The IOM and CPM are in the cabinet chassis, which has a backplane data bus and a backplane power bus to distribute data and power to the IOMs and CPMs.

Page 5-26



The IOMs transfer data between the software functions in the AIMS CPMs and external sources. The CPMs supply the software/hardware to do the calculations. There are four types of CPMs: 1. CPM/COMM – Core Processor Module/Communication. 2. CPM/ACMF - Core Processor Module/Aircraft Condition Monitoring Function. 3. CPM/B - Core Processor Module/Basic. 4. CPM/GG - Core Processor Module/Graphics Generator. Figure 16 shows the AIMS system (Boeing777).

AIRCRAFT CONDITION MONITORING SYSTEM (ACMS)

FLIGHT DATA RECORDER SYSTEM (FDRS)

FLIGHT MANAGEMENT COMPUTING SYSTEM (FMCS)

PRIMARY DISPLAY SYSTEM (PDS)

CENTRAL MAINTENCE COMPUTING SYSTEM (CMCS)

AIMS LEFT-HAND CABINET AIMS RIGHT-HAND CABINET

THRUST MANAGEMENT COMPUTING SYSTEM (TMCS)

DATA COMMUNICATION MANAGEMENT SYSTEM (DCMS)

AIMS System Figure 16


Page 5-27


5.7.1 Flight Compartment Printing System The flight compartment printer supplies high-speed hard copies of text for the following systems: 1. Primary Display System (PDS). 2. Airplane Condition Monitoring System (ACMS). 3. Central Maintenance Computing System (CMCS). The flight compartment printer receives data from the print driver partition of the Data Communication Management Function (DCMF). The DCMF is located within the AIMS. The DCMF prioritises data sent to the printer in the following order: 1. Flight Deck Communication Function (FDCF) of the DCMS. 2. Central Maintenance Computing Function (CMCF) of the CMCF. 3. Airplane Condition Monitoring Function (ACMF) of the ACMS. 4. Multi Function Display (MFD). The printer can print at 300 dots per inch (DPI). It uses a roll of paper, which is 125 feet long and is A4 European Air standard paper. The printer contains all mechanical components and electronics necessary for printer operation. The mechanical components include: 1. Printer head. 2. Rollers to move paper. 3. Motor and drive system. The electronic components include: 1. Power supply module. 2. Processor board. 3. Controller board. 4. Interconnection board

Page 5-28



Figure 17 shows the flight compartment printer.

FAIL

PAPER

CUT

SLEW

RESET

TEST

TOP VIEW

SIDE VIEW

Flight Compartment Printer Figure 17 Controller Board – Receives brightness controls from dimmer controls that drive the lights on the front panel. Processing Board – Processes all inputs for the left AIMS cabinet and changes the data signals to control the thermal printer. Interconnection Board – Controls the flow of data between the processor board and the controller board and the mechanical devices that print three paper.


Page 5-29



Page 5-30


MODULE 11 COMPLETE

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