CAR 66 B1 Lic Module 11.5.1 Instrument System (ATA 31)

Module 11.5.1 Instrument System(ATA 31) and 11.18Onboard maintenance System (ATA 45) CAR 66 Basic Training Manual

MODULE 11.5.1

1


PAGE INTENTIONALLY BLANK

2


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: •

Pressure instruments

•

Gyroscopic instruments

•

Compasses

•

Mechanical indicators

•

Electronic instruments

PRESSURE INSTRUMENTS 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 remotelylocated 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. 3


Figure 11-5-1-1 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

Figure 11-5-1-1 - Aircraft Pressure Head

4


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. •

IAS = Pitot minus Static - (In knots).

• Mach No = Pitot - Static divided by Static. •

Baro Ht = Static - (In feet).

FUSELAGE

•Vertical Speed = Change in Static pressure - (X 1000ft/min).

STATIC VENT

Figure 11-5-1.2 shows typical aircraft static vent:

STATIC PIPE

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.

5

Figure 11.5.1.2 – Aircraft Static Vent


Figure 11.5.1.3 shows the location of probes and vents on a Boeing 737.

ALTERNATE PRESSURE HEAD CAPT

CAPT

STATIC VENTS

F/O

PRESSURE HEAD AUX 2

Figure 11.5.1.3 – Boeing 737 Air Data Probe and Vent Location

6


Pitot and static pressures are transmitted through seamless and corrosion-resistant 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. 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 11.5.1.4 shows a typical water drain valve.

ORANGE FLOAT INDICATOR

TRANSPARENT PLASTIC PIPE

DRAIN VALVE

BAYONET FITTING CAP

(SELF SEALING)

Figure 11.5.1.4 – Water Drain Valve 7


The three primary instruments in the Air Data System are: •

Altimeter (Baro Ht).

•Indicated Air Speed (IAS) Indicator. •

Vertical Speed Indicator.

The IAS is often combined to display Mach No as well as indicated airspeed as is referred to as the “Combined Speed Indicator”. Figure 11.5.1.5 shows the connection and equations for the primary Air Data instruments.

HEIGHT = S IAS = P- S VSI = CHANGE IN S

0

9

1

FEET

1 , 6 2 7

8

00 10

2

ALTIMETER 3

7 1 01 3 6

350

5

5

2 9.9 2 4

400

8 4

300

MACH

60 1 0 9

80

250

120 IAS KNOTS

240 220

200

IAS INDICATOR

140 180

PITOT/STATIC PRESSURE HEAD 2 11000FT PER MIN VSI

.5

Figure 11.5.1.5 – Air Data Instrumentation

4 6

UP

0DOWN 1

8

6

VERTICAL SPEED

.5

2

4

VSI INDICATOR


ALTIMETER The operates on the aneroid barometric principle, i.e. responds to changes in atmospheric pressure, and are calibrated to indicate these changes in terms of equivalent altitude values. Figure 11.5.1.6 shows a typical altimeter.

ALTIMETER POINTER 0 - 1,000 ft

9

ALTITUDE SCALE 0 - 1,000 ft

0

SBY

1013

8 X 100 ft 7

1

3 5 0 00

MILLIBAR SETTING

2

ALTITUDE SELECT BUG

3

5

ALTITUDE DIGITAL READOUT 0 - 99,000 ft

6

5

4

MB ALTITUDE SELECT KNOB

SET MILLIBAR KNOB

Figure 11.5.1.6 – Altimeter 9


The pressure sensing element consists of an aneroid capsule, which transmits deflections in response to pressure changes. They are contained in a sealed container that is evacuated to the static pressure. There is a mechanical linkage to a pointer, which indicates the aircraft’s height above sea level. There is a facility to set the correct pressure of the day in millibars so that the instrument displays the correct height. Figure 11.5.1.7 shows the simplified operation of the altimeter.

SEALED ANEROID CAPSULE

SEALED CASE

POINTER

MECANICAL LINKAGE STATIC VENT

Figure 11.5.1.7 – Simplified Altimeter operation 10


“Q” CODE SETTINGS FOR ALTIMETERS The setting of altimeters to the barometric pressures prevailing at various flight levels and airports is part of the flight operating techniques. It is essential for maintaining adequate separation between aircraft and for terrain clearance during take-off and landing. In order to make the settings, flight crews are dependant on observed meteorological data which is requested and transmitted from ATC and form part of the ICAO “Q” code of communication. There are three code letter groups commonly used in connection with altimeter setting procedures: •

QNH.

•

QFE.

• QNE. QNH: Setting the barometric pressure to make the altimeter read airport elevation above-sea level on landing and take-off. When used for landing and take-off, the setting is generally known as “Airport QNH”. Any value set is only valid in the immediate vicinity of the airport concerned. Since an altimeter with a QNH setting reads altitude above sea level, the setting is also useful in determining terrain clearance when an aircraft is enroute. Fir this purpose, the UK and surrounding seas are divided into fourteen “Altimeter Setting Regions”, each transmitting an hourly “Regional QNH” forecast. QFE: Setting the barometric pressure prevailing at an airport to make the altimeter read zero on landing at, or taking off from, that airport. The zero reading is regardless of the airport’s elevation above sea level. QNE: Also known as the “Standard Altimeter Setting (SAS)”. The barometric pressure is set to 1013.25 mb and is used for flights above a prescribed “Transmission Height” and has the advantage that with all aircraft using the same airspace and flying on the same altimeter setting, the requisite separation between aircraft can more readily be maintained. The transition altitude within the UK airspace is usually 3000 - 6000'. Figure 11.5.8 shows QNH, QFE and QNE definitions.

11


QNE FLIGHT LEVEL

QNH HEIGHT ABOVE SEA LEVEL

QFE HEIGHT ABOVE AIRFIELD

STANDARD SETTING 1013.25 MILLIBARS SEA LEVEL

Figure 11.5.1.8 – QNH, QFE and QNE Definitions

12


COMBINED SPEED INDICATOR This indicator is one, which combines the functions of both a conventional indicator and a Machmeter. Figure 11.5.1.9 shows a typical Combined Speed Indicator (CSI).

MACH No READOUT LIMIT SPEED POINTER (Vmo)

AIRSPEED POINTER

EXTERNAL INDEX MARKERS

350 300

400

60

1 8 4 0 MACH 9

80 120

250

COMMANDED AIRSPEED BUG

IAS KNOTS

240 220

200

140

180

AIRSPEED SCALE (KNOTS) COMMAND AIRSPEED BUG SET

Figure 11.5.1.9 – Combined Speed Indicator 13


The internal mechanism consists of two elements (pointer and fixed scale for IAS and a digital readout for Mach No). There is also a second pointer on the IAS scale, this is known as the “Velocity Maximum Operating (Vmo)”. It indicates the aircraft’s maximum safe operating speed over its operating altitude range. To set the desired speed for operation, the flight crew uses the command bug. This speed in turn is the datum speed for the Auto throttle or Fast/Slow speed indicator. The external index bugs are used to set various reference speeds (take-off, flap retract speeds etc.). Figure 11.5.1.10 shows a simplified IAS operation.

PITOT

STATIC

Figure 11.5.1.10 – IAS Operation

14


VERTICAL SPEED INDICATOR (VSI) These indicators (also known as Rate-of Speed indicators) are very sensitive differential pressure gauges, designed to indicate the rate of altitude change from variations in static pressure alone. Figure 11.5.1.11 shows a VSI.

RATE OF CLIMB/DIVE POINTER

1

RATE OF CLIMB SCALE 1,000 ft per sec

2

1000FT PER MIN

VSI

.5

4 6

UP

0DOWN VERTICAL SPEED

.5

1

2

6 4

MAX INDICATED 6,000 ft per sec

RATE OF DIVE SCALE 1,000 ft per sec

Figure 11.5.1.11 – Vertical Speed Indicator (VSI) 15


Since the rate at which the static pressure changes is involved in determining vertical speed, a time factor has to be incorporated as a pressure function. This is accomplished by using a special air-metering unit in the sensing system. Its purpose is to create a lag in static pressure across the system and so establish the required pressure difference. Figure 11.5.1.12 shows a simplified VSI operation.

METERING UNIT

STATIC VENT

CLIMB

0

DIVE

Figure 11.5.1.12 – VSI Operation

CAPSULE

16

MECHANICAL LINKAGE

POINTER AND SCALE


AIR DATA SYSTEMS The complexity of an Air Data System depends primarily upon the type and size of the aircraft, the number of locations at which primary air data are to be displayed, the type of instruments installed, and the number of other systems requiring air data inputs. Figure 11.5.1.13 shows a typical air data system for a large aircraft.

PRESSURE HEADS

UPPER

LOWER

F/O

VS

PC

PITOT

IAS

STATIC

Figure 11.5.1.13 – Air Data System MS 1

ALT

A/S 1

ADC 1

FLT REC

DIFF PRESS

PITOT

VS

STATIC

MS 2

ALT

IAS

A/S 2

CAPT UPPER PRESSURE HEADS

17

LOWER

ADC 2


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. 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 center (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: •Gyroscopic Inertia (Rigidity). •Precession.

18


RIGITITY The property which resists any force tending to change the plane of rotor rotation. It is dependent on:

Z FRAME

1.The mass of the rotor.

Y

X

2.The speed of rotation. PRECESSION

ROTOR

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:

OUTER RING

X

1.The strength and direction of the applied force.

Y

INNER RING

2.The angular velocity of the rotor. Figure 11.5.14 shows a gyroscope.

Z

Figure 11.5.14 - Gyroscope.

19


Figure 11.5.1.15 shows the characteristics of gyro rigidity.

A

B C

Figure 11-5-1.15 - Gyro Rigidity

20


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. DIRECTION OF ROTATION

This rigidity can be used in a number of gyro instruments including the directional gyro.

PRECESSION RATE = APPLIED FORCE 90º IN THE DIRECTION OF SPIN

PRECESSION If an external force is applied to a spinning gyro, its effect will be felt at 900 from the point of application, in the direction of gyro rotation. This is known as precession. It can be seen in Figure 11.5.1.16, 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.

SPIN AXIS 90º APPLIED FORCE

Figure 11.5.1.16 - Gyro Precession

21

DIRECTION OF PRECESSION


VERTICAL GYRO Figure 11.5.1.17 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.

FREE GYRO

Figure 11.5.1.17 - Behavior of a Vertical Gyro +

22

ONLY PERPENDICULAR TO THE EARTH’S SURFACE AT TWO POINTS


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 ten-minute period is small. Figure 11.5.1.18 shows a vertical gyro corrected to the local vertical.

CORRECTED TO LOCAL VERTICAL

Figure 11.5.1.18 – Corrected Vertical Gyro +

23


Instruments that use either the rigidity or the precession of gyros are: • Gyro Horizon Unit. •Attitude Director Indicator. • Standby Horizon Unit. • Direction Indicator. •Turn and Slip Indicator. • Turn Co-ordinator.

24


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 11.5.1.19 shows a displacement gyro and the two axis of displacement. Figure 11.5.1.19 – Displacement Gyro

ROLL

25

PITCH


Indications of attitude are presented by the relative positions of two elements, one symbolizing the aircraft itself, the other in the form of a bar stabilized by the gyroscope and symbolizing the natural horizon. Figure 11.5.1.20 shows a typical Gyro Horizon Unit.

AIRCRAFT SYMBOL

3 ROLL SCALE

6

6

SPERRY

HORIZON BAR

3

ROLL POINTER\

Figure 11.5.1.20 – Gyro Horizon Unit 26


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 longitudual axis (ZZ1). The element symbolizing the aircraft may either be rigidily fixed to the case, or it may be externally adjustable for setting a particular pitch trim reference. Figure 11.5.1.21 shows the construction of the Gyro Horizon unit.

X

Y

OUTER RING

ROTOR

Z1

SYMBOLIC AIRCRAFT

BALANCE WEIGHT

PIVOT POINT

Z Y1 ROLL POINTER & SCALE

27

X1

HORIZON BAR


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. 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 11.5.1.22 shows climb attitude operation.

X

Z

Figure 11.5.1.22 – Climb Attitude operation. Z

HORIZON BAR

28

X

1

1


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 11.5.1.23 shows roll attitude operation.

X

Y

Y

1

BANK TO PORT DATUM X

Figure 11.5.1.23 – Roll attitude operation 29

1


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. ATTITUDE DIRECTOR INDICATOR This unit performs the same functions as a Gyro Horizon unit; i.e. it establishes a stabilized reference about the pitch and roll axes of an aircraft. Instead, however, of providing attitude displays by direct means, it is designed to be operated via a synchro system, which produces and transmits attitude-related signals to the indicator. The synchro system includes a attitude reference source and a computer linked into the aircraft’s navigational system to produce flight director signals for the flight crew to follow to ensure the aircraft follows the required course. Figure 11.5.1.24 shows a typical Attitude Director Indicator (ADI)

FD F

1 2

T AT TEST

30

GSL

1

S

Figure 11.5.1.24 - Attitude Director Indicator (ADI)

2

RW

Y


STANDBY HORIZON UNIT Most aircraft currently in service use Flight Director systems, or more sophisticated electronic flight instrument systems, all of which comprise indicators displaying not only attitude data, but navigational data as well. In such aircraft, the role of the conventional gyro horizon is mainly used as a standby instrument located on the center instrument panel. It is used as a reference in the event of a failure that might occur in the attitude display systems. Figure 11.5.1.25 shows a Standby Horizon Unit (SHU). The gyro is powered by 115V; three phase ac supplied from a static inverter, which in turn is supplied by 28V from the battery busbar. In place of the stabilized horizon bar a stabilized attitude sphere is used as the reference. The upper element is coloured blue to display climb attitudes, and black/brown for descending attitudes. A pitch trim adjustment and fast erection facility is provided, both being controlled by a knob on the lower right-hand corner of the indictor. When the knob is rotated the aircraft symbol can be positioned through 5º, thereby establishing a variable pitch trim reference. Pulling the knob out and holding it actuates the fast-erection circuit.

ROLL SCALE AIRCRAFT SYMBOL

PITCH SCALE

20

20

20

20

POWER “OFF” FLAG

D

C

PITCH ERECTION/ TRIM KNOB

Figure 11.5.1.25 – Standby Horizon Unit

31


DIRECTION INDICATORS This indicator was the first gyroscopic instrument to be introduced as aheading 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.

HEADING SCALE

LUBBER LINE

180

170

Figure 11.5.1.26 shows a Directional Indicator

CAGING/SETTING KNOB

32


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 11.5.27 shows the construction of a directional gyro.

ROTOR ASSEMBLY

INNER GIMBAL RING

VERTICAL GIMBAL RING

COMPASS CARD

SYNCHRONISER RING

Figure 11.5.1.27 – Directional Gyro

CAGING/ SETTING KNOB

33


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: •Rate 1 - Turn Rate 180º per minute. •Rate 2 - Turn Rate 360º per minute. RATE OF TURN INDICATOR

•Rate 3 - Turn Rate 540º per minute. •Rate 4 - Turn Rate 720º per minute. Figure 11.5.1.28 shows a typical Turn & Slip indicator.

2 MIN Figure 11.5.1.28 – Turn & Slip Indicator

RATE OF TURN 2 MIN - 360º

34

SLIP/SLIDE INDICATOR


For the detection of rates of turn, a rate gyroscope is used and is arranged in the manner shown in figure 11.5.1.29. 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 INPUT operating position the rotor spin axis, due to AXIS 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.

FWD

Y1

X

Figure 11.5.1.29 – Rate Gyro Turn Indicator

F Y

P

35

X1


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. 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 utilizing a ball in a curved liquid-filled glass tube as shown in Figure 11.5.1.26. 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/over-banked), 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.

36


Figure 11.5.1.30 shows bank indication for various aircraft bank conditions.

CF

W

W

R

Figure 11.5.1.30 – bank Indications

CF

W

CF

R W

37

R


TURN COORDINATOR 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 11.5.1.31 shows a Turn coordinator indicator.

AIRCRAFT SYMBOL

TURN COORDINATION

RATE OF TURN

L

R 2 MIN NO PITCH INFORMATION

Figure 11.5.1.31 - Turn coordinator Indicator 38

TURN COORDINATOR


HORIZONTAL SITUATION INDICATOR This indicator derives its name from the fact that its display presents a pictorial plan of the aircraft’s situation in the horizontal plane in the form of its heading, VOR/LOC deviation and other data relating to navigation. Figure 11.5.1.32 shows a typical HSI. The aircraft symbol is fixed at the center of the instrument and displays the heading of the aircraft in relation to a rotating compass card and the VOR/LOC deviation bar (lateral bar). The selector knobs at the bottom corners of the instrument permit the setting of desired magnetic heading and VOR course.

3 4 5

5 9 2

MLS

GND SPEED

33

N

3

30

6

W

E

24

12 21

Figure 11.5.1.32 – Horizontal Situation Indicator

39

S

15


COMPASS SYSTEMS The compass has, since the earliest times, given information to travelers with regards to the direction to go. Mounting a compass on a moving object, whether it was a vehicle, a ship or an aircraft poses certain problems. This includes how to mount the compass without the, motion (maybe violent), upsetting the device.

GEOGRAPHICAL NORTH POLE MAGNETIC NORTH POLE

Another problem that besets compasses is the fact that they usually point to magnetic north, which slowly moves, and not true north, the difference between the two is something like 1,300-miles/2,000 km. This is of little concern if we are moving slowly, on a boat, in the vicinity of the equator, but vital in an aircraft flying what is known as a 'Trans-polar route' from say, New York to Tokyo. The effect this has on navigational charts is referred to as 'variation'.

0º E VARIATION 17.5º E VARIATION

Figure 11.5.1.33 shows the difference between True North and Magnetic North.

Figure 11.5.1.33 - True North & Magnetic North

40

11º W VARIATION


DIRECT READING COMPASS Direct-reading compasses have the following common principal features: • Magnet system housed in a bowl. • Liquid damping and liquid expansion compensation. Figure 11.5.1.34 shows a direct reading compass used as a standby compass in most aircraft.

HORIZONTAL (“B” & “C”) CORRECTORS

MOUNTING PLATE

FILLER PLUG

BELLOWS BOWL

Figure 11.5.1.34 – Standby Compass

STEM & BRACKET ASSEMBLY

41

MAGNET SYSTEM


The magnet system comprises an annular cobalt-steel magnet to which is attached a light-alloy card. The card is graduated in increments of 10º, and referenced against a lubber line fixed to the interior of the bowl. The system is pendulously suspended by an iridium-tipped pivot resting in a sapphire cup supported in a holder or stem. The bowl is of a plastic (diakon) and so moulded that it has a magnifying effect on the card and its graduations. It is filled with a silicone fluid to prevent the card oscillating or overshooting after changes of heading. The fluid also provides the system with a certain buoyancy, thereby reducing the weight on the pivot and so diminishing the effects of friction and wear. Changes in the volume of the fluid due to temperature changes, and their resulting effects on damping efficiency, are compensated by a bellows type of expansion device secured to the rear of the bowl. Compensation of the effects of deviation due to longitudinal and lateral components of aircraft magnetism is provided by permanent magnet coefficient “B” and “C” corrector assemblies secured to the compass mounting plate. A small lamp is also provided for illuminating the card. Figure 11.5.1.35 shows a complete standby compass indicator.

B

C

CO-EFFICIENT “B” ADJUSTMENT

21 CO-EFFICIENT “A” ADJUSTMENT

S

CO-EFFICIENT “C” ADJUSTMENT

15 12 LUBBER LINE

ELECTRICAL CONNECTION FOR LIGHTING

Figure 11.5.1.35 – Standby Compass

42


REMOTE READING COMPASS A remote reading compass, is basically one in which an element detects an aircraft’s heading with respect to the horizontal component of the earth’s magnetic field in terms of flux and induced changes in voltage. It then transmits these changes via a synchronous/servo system to a heading indicator. There are two types of remote reading compass systems: •The detector element monitors a directional gyro unit linked with a heading indicator. •The detector element operates in conjunction with the platform of an inertial navigation system (INS). DETECTOR UNIT (FLUX VALVE) The detector unit detects the effect of the earth’s magnetic field as an electromagnetically induced voltage and controls the heading indicator by means of a variable secondary output voltage signal. The construction of the element takes the form of a three-spoked wheel, slit through the rim between the spokes so that they, and their section of rim, act as three individual flux collectors. Figure 11.5.36 shows the construction of a flux valve.

LAMINATED COLLECTOR HORNS

A

A

AC POWER

B

EXCITER COIL

C SECONDARY PICK-OFF COILS

Figure 11.5.1.36 – Flux Valve Construction

43

B C


The paths taken by the earth’s magnetic field through the spokes for differnet headings is shown at Figure 11.5.1.37. The detector unit on its own is not very accurate by virtue of its limited pendulous suspension arrangement. Errors will occur as a result of its tilting under the influence of acceleration forces, e.g. during speed changes on a constant heading and during turns. It is necessary to incorporate within the system a means of monitoring the detector’s output. The horizontal directional gyro is used to give the system short-term accuracy with the detector unit providing long-term accuracy.

P A TH OF E A RT H’ S FI E LD A

A C

B

B

C

B

C

Figure 11.5.1.37 – Earth’s Flux path

A

C

B A

44


Figure 11.5.1.38 shows the arrangement of a remote reading gyro compass system.

115v 400 Hz

B

+

Figure 11.5.1.38 – Gyro Magnetic Compass System

_

6

E 12

21 VOR

S 15

A D F

_+

3

W 3 0 24

N 33

C

A D F

VOR

SLAVED

VOR/ADF

45

DG

SYNC


Figure 11.5.1.39 shows a schematic of a Gyro Magnetic Compass system.

Figure 11.5.1.39 - Gyro Magnetic Compass System Schematic

CT N

M

W

E

26V AC 400 Hz

S

TG

CX CT M DETECTOR UNIT

46

GYRO


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 AOA probe and the stall vane. The probe contains slots on the leading edge of the probe itself and, depending on the angle of attack; the air flowing through the different slots move a 'paddle' which indicates the AOA electrically in the cockpit. The stall vane is rather like a small weather vane mounted on the side of the aircraft. The vane follows the airflow, much like the weather vane, but indicating, not pitch angle, but the angle of the airflow relative to the aircraft centerline. i.e. the angle of attack. Figure 11.5.1.40 shows a vane type Angle of Attack transducer.

A330

ANGLE OF ATTACK

AIRCRAFT LONGITUDINAL AXIS

VANE ARM ANGLE OF ATTACK TRANSDUCER

∝ FLIGHT PATH

Figure 11./5.1.40 - Angle of Attack Transducer

AIRFLOW

47


STALL WARNING INDICATION To maintain lift at low airspeed, the angle of attack is increased. When this angle is above a critical angle, the aircraft wings will not produce enough lift to support the aircraft, which will begin to stall. Before this situation occurs, the aircraft will shake heavily, this being a natural alert to the pilot. If, however, the aircraft is configured for an approach (Wheels & Flaps down), the airspeed difference between the natural warning and the actual stall is very small, so an alert must be generated before the stall occurs. Modern performance aircraft use the output from an Angle of Attack probe, connected to a Stall Warning system. The stall warning system also has other sensor inputs (Flap, Slat positions). Once critical angle prior to actual stall is reached, the stall warning system initiates a "Audio warning" and operates a "Stick Shaker", which actually shakes the control column. Figure 11.5.2.40 shows simple stall warning system.

28V DC SUPPLY

the

ANGLE OF ATTACK

Figure 11.5.1.40 - Stall Warning System

>17.5º

M STICK SHAKER

48


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 11.5.1.41 shows a typical flight deck layout of an Airbus A320.

EFIS PFD

EFIS ND

ECAM ENGINE WARNINGS

ECAM SYSTEMS

Figure 11.5.1.41 - Flight Deck Electronic Instrumentation Layout

49

EFIS ND

EFIS PFD


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 11.5.1.42 shows the switching panel from Airbus A320.

Figure 11.5.1.42 – A320 EIS Switching Panel

ATT HDG CAPT 3

50

NORM

AIR DATA F/O 3

NORM CAPT 3

F/O 3

E/S DMC CAPT 3

NORM

ECAM / ND XFR F/O 3

CAPT

NORM

F/O


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 11.5.1.43 shows a block schematic of the EIS for the Airbus 320.

Figure 11.5.1.43 – Electronic Instrument System (EIS)

DISPLAY MANAGEMENT SYSTEM DMS No 1

51




ELECTRONIC FLIGHT INSTRUMENT STSTEM 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. ELECTRONIC ATTITUDE DIRECTOR INDICATOR (EADI) GENERAL The EADI displays traditional attitude information (Pitch & Roll) against a two-colour 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 w.r.t. 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.

52


Figure 11.5.1.44 shows a typical EADI display

ROLL SCALE

Honeywell

Figure 11.5.1.44 - Electronic Attitude Director Indicator (EADI) Display

ATTITUDE SPHERE LOC

HDG

ATT 2

F

S

PITCH SCALE

M .99 200 DH

INCLINOMETER

53

GS

20

20

10

10

10

10

20

20

M AP ENG

140 RA

AIRCRAFT SYMBOL


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).

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.

54


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.

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.

55


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-of-Turn display.

Vertical Navigation Display:

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.

56


ELECTRONIC HORIZONTAL SITUATION INDICATOR 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 dependant on the system fitted.

DISTANCE DISPLAY

Honeywell

Figure 11.5.1.45 shows an EHSI display.

NAV 1

CRS 315 +0

COURSE SELECT POINTER

H

6

24

3

WPT

2.1 NM

N

W

30

33

G

E

21

VOR 1

ADF 1

15

350

HEADING DIAL

12

S HDG

GSPD 130 KTS

COURSE OR AZIMUTH DEVIATION BAR

Figure 11.5.1.45 - Electronic Horizontal Situation Indicator (EHSI) Display

AIRCRAFT SYMBOL

57


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).

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:

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.

58


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 w.r.t. the lubber line represents drift angle left or right of the desired track. The drift angle bug w.r.t. 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.

Heading SYNC 59


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.

PART TIME EHSI DISPLAYS Vertical Navigation Display:

Glide Slope Deviation:

Bearing Pointer Source Annunciators:

Elapsed Time Annunciation:

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.

60


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 11.5.1.46 shows a Partial EHSI display (Compass Mode). Figure 11.5.1.46 - EHSI Partial Compass Mode Display Honeywell

DTRK

317

NAVAID

320 30

RANGE RINGS

FMS1 30 NM

33 N V

VOR 1

50 ADF 1

HDG

350

WIND VECTOR

61

25 15

GSPD 130 KTS


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.

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: 1.

Black -

No storm.

2.

Green -

Moderate storm.

3.

Yellow -

Less severe storm.

4.

Red

Severe storm.

-

62


Figure 11.5.1.47 shows an EHSI partial format with Weather Radar information.

Honeywell

DTRK

317

FMS1 30 NM

320 30

33 N

Figure 11.5.1.47 - EHSI Weather Radar Display

V VOR 1

50 ADF 1

HDG

350

GSPD

25

130 KTS

WIND VECTOR BLACK = NO STORM GREEN = MODERATE STORM YELLOW = LESS SEVERE STORM

63

RED = SERVERE STORM

WEATHER RADAR RETURN


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 11.5.1.48 shows an EHSI MAP mode display. Honeywell

Figure 11.5.1.48 - EHSI MAP Mode Display.

DTRK

317

WAYPOINT SYMBOL & IDENTIFIER

FMS1 30 NM

320 33

30 05

04 05

N V

VOR 1

50

03

ADF 1

HDG

350

GSPD

25

64

130 KTS


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.

TO/FROM ANNUNCIATOR

Honeywell

Figure 11.5.1.49 shows a typical composite display.

ATTITUDE SPHERE

120 NM HDG ILS

CRS FR ATT 2

20

20

F

10

10 HEADING DISPLAY

HEADING SELECT DISPLAY

10

010 M .99 200 DH

Figure 11.5.1.49 - EFIS Composite Display

10

000

S 33

00

03 DH

COURSE AZIMUTH BAR

65

M 140 RA

HEADING TAPE DISPLAY


DISPLAY 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.

DISPLAY SELECT BUTTONS

EFIS Display Controllers are shown at Figure 11.5.1.50.

FULL ARC CRS

GS TTG

WX

DIM

ET

DH

MAP BOT

SC CP

REV HDG

TOP

TEST RASTER DIM

DISPLAY CONTROLLER

SOURCE SELECT BUTTONS

Figure 11.5.1.50 - EFIS Display and Source Controllers NAV ADF 2 ADF 1

VLF

FMS

INS 1

INS 2

VOR 1 ADF 2

AUTO

HDG VOR 2

ADF 1

OFF

OFF

BRG

BRG

SOURCE SELECT CONTROLLER

66

ATT


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.

67


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: Rotation of the heading select knob allows the heading select bug to be rotated to the desired heading.

68


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.

69


The EFIS comprises the following units: Honeywell

1. Symbol Generator (SG).

GS

ATT 2 AOA F

20

20

10

10

10

10

G GS

WX

DIM

CRS

ET

TTG DH

SC

MAP

BOT

REV

CP

TOP

S CMD M .99 200DH

HDG

TEST

4. Remote Light Sensor.

VLF

FMS

INS 1

INS 2

ADF 2 ADF 1

ATT

HDG

I

140RA

Honeywell

VOR 2

CRS +0

OFF

BRG

BRG

NAV 1

345

ADF 1

OFF

DH

EFIS SG No 1

VOR 1 ADF 2

AUTO

20

N 33

ADF 1

E 1 2

VO R 1

HDG

NAV AID ILS/VOR

H 2.1 NM 3

6

Figure 11.5.1.51 shows the EFIS units and signal interface in block schematic form.

INERTIAL REF SYSTEM

NAV

20

21

3. Control Panel.

AIR DATA COMP

RAST ER DIM

W 3 24 0

ARC

S

15

2. Display units X 2 (EADI & EHSI).

FULL

GSPD

013

130 KTS

EFIS SG No 3

RAD ALT Honeywell

AOA F

20

10

10

10

10

G

FULL ARC

GS T TG

WX

DIM

CRS

FMS

ET

DH

MAP

BOT

SC CP

REV

T OP

20

20 DH

140RA

HDG

TEST RASTER DIM

EFIS SG No 2

AFCS

Honeywell INS 2

ADF 2 ADF 1

VOR 1 ADF 2

AUTO

ATT

HDG VOR 2

CRS +0

ADF 1 OFF

VO R 1

BRG

BRG

ADF 1

HDG

013

70

N 33

H 2.1 NM 3

E 1 2

OFF

NAV 1

345

30

INS 1

W 24

FMS

21

VLF

S

15

NAV

6

Figure 11.5.1.51 - EFIS Block Schematic

20

S CMD M .99 200DH

DME

GPWS

GS

ATT 2

WEATHER RADAR

GSPD

130 KTS


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. 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. 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: •Resistance Gauges. •Capacitance Quantity Indicators.

71


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. Figure 11.5.1.52 shows a simplified resistance gauge.

INDICATOR

N S

TANK RESISTOR

Figure 11.5.1.52 - Resistance Gauge

+ DC POWER

FUEL TANK

72


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. TANK UNIT Figure 11.5.1.53 shows a circuit of a capacitance quantity system.

EMPTY

REF C

Figure 11.5.1.53 - Capacitance Quantity Indicating System

FULL

2 - PHASE MOTOR

DISCRIMINATION STAGE

AMPLIFIER STAGE INDICATOR

73

REF PHASE

AMPLIFIER UNIT

IS

LOOP A

IB

LOOP B


FUEL FLOW INDICATOR As the name suggests, these indicators show the amount of fuel flowing into the engines. Fuel flow information can be represented as either LBS/HR, Gallons/HR or PSI. Some indicators will show both PSI and either LBS/HR or Gallons/HR. Figure 11.5.1.54 shows a fuel flow indicator.

PSI SCALE 2.5 PSI

Figure 11.5.1.54 - Fuel Flow Indicator

L

LBS/HR

195 PSI

45 55 65

FUEL FLOW

LBS/HR SCALE

50

75

100

80

R T.O.

170

95

LEFT ENGINE FUEL FLOW

150

RIGHT ENGINE FUEL FLOW

74


FUEL PRESSURE INDICATOR Some engines have a fuel pressure gauge that displays the pressure of the fuel supplied to the fuel control unit. Most display the pressure in pounds per square inch (psi) and provide indications to the pilot that the engine is receiving the fuel required for a given power setting. Figure 11.5.1.55 shows a fuel pressure gauge.

Figure 11.5.1.55 – Fuel Pressure Gauge There are two types of pressure gauge:

10 PSI

•Bourbon Tube type.

30

POINTER

FUEL PRESS

•Pressure Capsule type.

50

125 PSI 100

80

PSI SCALE

75


BOURBON TUBE FUEL PRESSURE INDICATOR 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 bourbon tube is connected to a capillary tube containing pressurized fuel. As the pressurized fuel enters the bourbon 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 11.5.1.56 shows a Bourbon type fuel pressure gauge and its operation.

POINTER STAFF

Figure 11.5.1.56 – Bourbon Tube Fuel Pressure Gauge

BOURBON TUBE

ANCHOR POINT GEARING

76


PRESSURE CAPSULE FUEL PRESSURE INDICATOR 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 pressurized fuel to the diaphragm. As the diaphragm becomes pressurized it expands, causing an indicator pointer to rotate. Figure 11.5.1.57 shows a pressure capsule type fuel pressure indicator.

DIAPHRAGM Figure 11.5.1.57 – Pressure Capsule Fuel Pressure Gauge

77


ENGINE RPM INDICATORS These instruments indicate the rotational speed of the engine. Low Pressure Compressor (N1), Intermediate Pressure Compressor (N2) and High Pressure Compressor (N3). Figure 11.5.1.58 a RPM gauge for N1 measurement.

PERCENTAGE SCALE POINTER

Figure 115.1.58 - N1 RPM Gauge DIGITAL REPRESENTATION OF POINTER

20 0

40

60

N1 % RPM 80 100

9 2 FIALURE FLAG

78

OPTIMUM TURBINE SPEED


The indicator use electromagnetic sensors (which contains a coil of wire that generates a magnetic field) to measure the RPM of the respective compressor blades. The sensor is mounted in the shroud around the fan so, when each fan blade passes the sensor, the magnetic field is interrupted. The frequency at which the fan blades cut across the field is measured by an electronic circuit and then transmitted to a RPM gauge in the cockpit. Figure 11.5.1.59 shows the operation of a N1 & N2 gauges.

Figure 11.5.1.59 – N1 & N2 Pressure Gauges Operation

PICK-UP COILS

ELECTROMAGNETIC SENSOR GEAR DRIVEN ROTOR

FAN BLADES

PERMANENT MAGNET

40

% RPM

80

20

N1 GAUGE

80

20 0

0

100 NI

79

60

60

40

N2 % RPM 100

N2 GAUGE


Engine Temperature Gauges Because turbine engines can be severely damaged by high temperature in the turbine sections, a means of measuring the temperature is required. Because of the high temperatures involved, this is carried out using thermocouples. There are a number of different terms and abbreviations used for the gas temperature in turbine engines, these are: •Turbine Inlet Temperature (TIT). •Inter Turbine Temperature (ITT). POINTER

•Turbine Outlet Temperature (TOT). TEMERATURE SCALE

•Engine Gas Temperature (EGT).

5 3

•Measured Gas Temperature (MGT).

EGT °C X 100

7

OVER-TEMP LIMIT POINTER

1

•Jet Pipe Temperature (JPT).

7 6 5

Figure 11.5.1.60 shows a typical EGT indicator

OVER-TEMPERATURE WARNING LIGHT

Figure 11.5.1.60 - EGT Indicator

80

9

DIGITAL READ-OUT


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 thermocouples. Figure 11.5.1.61 shows EGT indicator operation.

Figure 11.5.1.61 – EGT Indicator Operation

AMPLIFIER

8

9 EGT

3

7 C° X100

6

5

4

EGT INDICATOR

81

EGT THERMOCOUPLE PROBE

EGT THERMOCOUPLE HARNESS


ENGINE PRESSURE INDICATORS Engine pressure indicators provide indications of the thrust being produced by a turbojet or turbofan engine. Figure 11.5.1.62 shows an EPR indicator.

Figure 11.5.1.62 - EPR Indication System

EPR DIGITAL COUNTERS

EPR SCALE

0.8 FAILURE FLAG

POINTER

1.0 1.2

1 5 0 EPR

1.4 1.6

EPR REFERENCE DIGITAL COUNTERS

EPR REFERENCE BUG SET KNOB

82

1 5 0

EPR REFERENCE BUG


The EPR is the ratio of turbine discharge pressure to compressor inlet pressure. Pressure measurements are recorded by total pressure pickups, or EPR probes, installed in the engine inlet Pt2 section and at the exhaust Pt7 section. Once collected, the data is sent to a differential pressure transducer, which drives a cockpit EPR gauge. Figure 11.5.1.63 shows the operation of an EPR indicator.

Figure 11.5.1.63 – EPR Indicator Operation

Pt 7 PROBE

Pt 2 PROBE

1.0 0.3

EPR

1.2 1.4

EPR TRANSDUCER

83

1.6

1 5 0

EPR INDICATOR


Figure 11.5.1.64 shows the engine instrument grouping for a twin engine aircraft.

EPR

0.8

EPR

1.0

1.0 1.2 1.0 0.8

1.0 1.2

1 5 0 1.4

1.4

1.6

1.6

150

1 5 0

Figure 11.5.1.64 - Power plant instrument grouping

100

40 20 60 N1 80 % RPM 0

EGT

1

3 4

FF

2 1

9 2

5 EGT °C X 100 7

7 6 5

FF

5 EGT 3 7 °C X 100

9

9

5

X 1000

6 5 8

%RPM

100

9 2

3

EPR

1 5 0

20 40 60 N1 0 % RPM 80

%RPM

EPR

5 6

6

8

8

84

1

EGT

7 6 5

4 3

FF

X 1000

8 5 6

2 1

FF


ELECTRONIC INSTRUMENTS (ENGINE & AIRFRAME SYSTEMS) 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). 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. 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 colours are produced by the CRTs for displaying information. Table 11.5.1.1 shows the colours and description of there uses.

85


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 11.5.1.1

86


Figure 11.5.1.65 shows layout of the EICAS Displays.

Figure 11.5.1.65 - EICAS Primary and Secondary Display Formats CAUTION RESET CANCEL

0 SBY

1 1 01 3 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

87

L AUTO R

THRUST REF SET L

BOTH

R

MAX IND RESET


Figure 11.5.1.66 and 67 show display formats for primary and secondary displays.

CAUTION

TAT 15°c 10

CANCEL RECALL

6

0.0

10

2

Figure 11.5.1.66 – Primary EICAS Display

6

0.0 2

N1 0

0

EGT

VVVVVVV

88


Figure 11.5.1.67 – Secondary EICAS Display

50

50

OIL

PRESS

120

120

OIL

TEMP

18

18

OIL

N2 86

86 N3

4.4

4.4

QTY

N1

FAN

3.1

1.9 VIB

89

88

88.00

FF


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).

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. 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 11.5.1.68 shows an example of a status page. 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 sub-systems.

90


Figure 11.5.1.68 – EICAS Status Page

HYD QTY

L 0.99

C R 1.00 0.98

HYD PRESS

2975

3010 3000

APU

EGT 440

OXY PRESS

RPM 103

1750

RUD

AIL ELEV AIL

91

OIL 0.75

0.0

FF

0.0


DISPLAY SELECT PANEL To control the operation of the EICAS, a control panel is situated on the center pedestal. Figure 11.5.1.69 shows a typical EICAS control panel.

Figure 11.5.1.69 - EICAS Control Panel

COMPUTER

DISPLAY

BRT BRT

ENGINE

92

STATUS

EVENT RECORD

BAL

L AUTO R

L BOTH R

MAX IND RESET


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:

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.

93


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.

94


Figure 11.5.1.70 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

10 6

70.0

10

2

6

Figure 11.5.1.70 - Upper EICAS Display – Alert Messages

95

2

N1 999

775

EGT

VVVVVVV

110.0


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 11.5.1.71 shows a typical maintenance control panel.

Figure 11.5.1.71 - Maintenance Control Panel

PERFORMANCE AND AUXILLIARY POWER UNIT FORMATS ENVIRONMENTAL CONTROL SYSTEM AND MAINTENANCE MESSAGE FORMATS

SELECTS DATA FROM AUTO OR MANUAL EVENT IN MEMORY

ELECTRICAL AND HYDRAULIC SYSTEM FORMAT

EICAS MAINT DISPLAY SELECT

ECS

ELEC

PERF

MSG

HYD

APU

CONF MCDP

CONFIGURATION AND MAINTENANCE CONTROL/DISPLAY PANEL

ENG EXCD

ENGINE EXCEEDANCES

96

BITE TEST SWITCH FOR SELF-TEST ROUTINE

EVENT READ AUTO

MAN

REC

ERASE

TEST

RECORDS REAL-TIME DATA CURRENTLY DISPLAYED (IN MANUAL EVENT)

ERASES STORED DATA CURRENTLY DISPLAYED


ELECTRONIC CENTRALIZED AIRCRAFT MONITORING (ECAM) 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. 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 associated information in pictorial or synoptic format. Figure 11.5.1.72 shows the layout of ECAM displays.

350

400

8 4

30 0

MACH

60 1 0 9

80

250

120 IAS KNOTS

240 220

200

14 0 180

to 5

Figure 11.5.1.72 – ECAM Display Layout 5

LDG GEAR GRVTY EXTN RESET OFF DOWN

97


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 11.5.1.80. 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. Pre-flight, Takeoff, Climb, Cruise, Decent, Approach, and post landing. Figure 11.5.1.73 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

N1 %

10

5

6 50

80 1500

Figure 11.5.1.73 – ECAM Upper and Lower Display (Cruise Mode)

10 4 80

S

FLAP

KG

1530

OIL F

11.5

QT Y

80.2

FF KG/H

1500

ON ON F UL L F UL L

VIB 0.8

(N1) 0.9

VIB 1.2

(N2) 1.3

11.5 LDG ELEV AUTO

FULL CKPT 20

FWD 22

AFT 23

24

22

24

500FT

CAB V/S FT/MIN 250 CAB ALT FT 4150

L DG I N HIB IT A PU B L E ED

EC AM UPP ER D ISPLA Y

98

1530

A IR

N2 %

N O SM O K IN G : SE AT BE LT S: SP LR S: F LA PS :

F.USED FOB : 14000KG

5

EG T ºC

10 6 5. 0

TAT +19 ºC SAT +17 ºC

G.W. 60300 KG 23 H 56

C.G. 28.1 %

EC AM LOW ER D IS PLAY - CRU ISE


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 11.5.1.74 shows an advisory message following a Blue Hydraulic failure.

Figure 11.5.1.74 – ECAM Advisory Mode 10

5

87.0

650

ADVISORY MESSAGES

80 1500

65.0

N1 %

10

5

10

5

FOB : 14000KG 10

5

EGT ºC

480

N2 %

80.2

FF KG/H

1500

HYD B RSVR OVHT B SYS LO PR

FAILURE MESSAGES

99

1 FUEL TANK PUMP LH

S

FLAP

FULL

FLT CTL SPOILERS SLOW

F


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.

100


Figure 11.5.1.75 – 11.5.1.79 shows the 12-system pages and status page available.

Figure 11.5.1.75 - ECAM System Displays Note; These pages are displayed: Automatically due to an advisory or failure related to the system.Whenever called manually.

C ON D

TE M P ºC

C AB P R E S S

A LTN M O DE F AN

F AN

CKPT 20

FWD 22

24

22

C

H

C

LDG ELEV MAN 500FT

AP PS I

V/ S FT/M IN

2

8

AFT 23

0

0 4.1

24 H

C

DN

M AN

SY ST 1

H

1150 2

IN L ET

G.W. 60300 KG C.G. 28.1 %

AIR CO ND ITIO NIN G SY STEM PA GE

101

4150

SA FET Y

EX T RA C T

PA C K 1

23 H 56

SY ST 2

10 0

VE NT

H OT A IR

TAT +19 ºC SAT + 17 ºC

C A B ALT FT

UP


PA C K 2

G.W. 60300 KG 23 H 56

C.G. 28.1 %

PR ESSU RISA TION SYSTE M P AG E


Figure 11.5.1.76 - ECAM System Displays Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.

BAT 1 28 V 15 0A

ELE C DC

D C B AT

F/CT R

BAT 2 28 V 15 0A

1

GBY

DC 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

TR 2 28 V 15 0A

EX T PW R 11 6V 40 0H Z

GE N 2 26 % 11 6V 40 0H Z

C.G. 28.1 %

ELE CTRIC AL SY STEM PAG E

102

R AIL GB

PITCH TRIM G Y 3.2º UP

AC 2

G.W. 60300 KG 23 H 56

SP D BR K

L AIL BG

RUD GBY

L ELEV BG


R ELEV YB

G.W. 60300 KG 23 H 56

C.G. 28.1 %

FLIG HT CO NTR OL SY STEM PA GE


Figure 11.5.1.77 - ECAM System Displays

1550

F .U SE D 2

1550

FOB APU

Note; These pages are displayed: Automatically due to an advisory or failure related to the system. Whenever called manually.

FU E L KG

F .U SE D 1

H YD GR EE N

3000

LEFT

1 07 5 0


YE LL OW

5 60 0

1 07 5 0

3000

PSI

3000

5 50

G.W. 60300 KG 23 H 56

PSI

RIGHT

CTR

5 50

B LU E

2 87 5 0

C.G. 28.1 %

FUE L SYSTE M P AG E

103


G.W. 60300 KG 23 H 56

C.G. 28.1 %

HY DR AULIC SYSTE M P AG E



WHEEL

B LE E D

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.

20 ºC

24 ºC C

C

H RAM AIR

50 ºC

170 1

ºC REL

140

140

2

3

ºC REL

LO

2

GND APU LP

G.W. 60300 KG C.G. 28.1 %

LANDING GEAR/WHEEL/BRAKE SYSTEM PAGE

104

HI

4

AUTO BRK

23 H 56

LO

HI

140 1


H 230 ºC


HP

HP

LP

G.W. 60300 KG 23 H 56

C.G. 28.1 %

A IR BLEED SYSTEM PA GE



A RM

ARM

A VI O N IC

Automatically due to an advisory or failure related to the system. Whenever called manually.

A PU 26%

B LE ED 35 PSI

116 V 400 HZ

C A BIN

Note; These pages are displayed:

APU

O X Y 1 85 0 P S I

D OOR

F W D C O M PT C A RG O

ARM

EM ER EX IT

10

AR M

0

N %

80

FLAP OPEN

C A RG O B U LK C A BIN

A RM

5

AR M

3

Related flight phase. TAT +19 ºC SAT + 17 ºC

23 H 56

C.G. 28.1 %

DO O R/O XY SYS TEM PAG E

105


7

EG T ºC

580

23 H 56

C.G. 28.1 %

AP U SY STEM PAG E


CONTROL PANEL The layout of the control panel is shown in Figure 11.5.1.80.

DISPLAY ON & BRIGHTNESS CONTROL

Figure 11.5.1.80 - ECAM Control Panel

DISPLAY ON & BRIGHTNESS CONTROL

SGW SELECT SWITCHES

1

TOP DISPLAY

OFF MESSAGE CLEARANCE SWITCH

2

FAULT

FAULT

OFF

OFF

BOTTOM DISPLAY

OFF

STS

RCL

RECALL SWITCH

106

SGU

BRT

CLR

STATUS MESSAGE SWITCH

ECAM

ENG

HYD

AC

DC

BLEED

COND

PRESS

FUEL

APU

F/CTL

DOOR

WHEEL

SYSTEM SYNOPTIC DISPLAY SWITCHES

BRT


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 lefthand display

107

CAR 66 B1 Lic Module 11.5.1 Instrument System (ATA 31)

Recommend Documents