B2-13c Auto Flight SR

Student Resource

Subject B2-13c Autoflight

Copyright © 2008 Aviation Australia All rights reserved. No part of this document may be reproduced, transferred, sold, or otherwise disposed of, without the written permission of Aviation Australia.

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AA Form TO-19 B2-13c Autoflight

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CONTENTS Topic Definitions

ii

Study Resources

iii

Fundamentals

13.3.1

Command Signal Processing

13.3.2

Modes of Operation

13.3.3

Yaw Damping and Trim Control

13.3.4/6

System Interface

13.3.7/8

Autothrottle and Autoland

13.3.9/10

Helicopter Autoflight

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DEFINITIONS Define 

To describe the nature or basic qualities of.



To state the precise meaning of (a word or sense of a word).

State 

Specify in words or writing.



To set forth in words; declare.

Identify 

To establish the identity of.

List 

Itemise.

Describe 

Represent in words enabling hearer or reader to form an idea of an object or process.



To tell the facts, details, or particulars of something verbally or in writing.

Explain 

Make known in detail.



Offer reason for cause and effect.


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STUDY RESOURCES E. H. J. Pallett, & Shawn Coyle (1993). Automatic Flight Control, (fourth edition): Blackwell Science, London. Jeppesen Sanderson, (1974). Avionics Fundamentals, United Airlines.


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INTRODUCTION The purpose of this subject is to allow you to gain knowledge in the operation Aircraft Automatic Flight Systems. On completion of the following topics you will be able to: Topic 13.3.1

Autoflight -Fundamentals Explain the fundamentals of automatic flight control including working principle and define current terminology.

Topic 13.3.2

Autoflight - Command Signal Processing Explain command signal processing.

Topic 13.3.3

Autoflight - Modes of Operation Explain the modes of operation in the following control channels:  Roll;  Pitch and  Yaw.

Topic 13.3.4

Autoflight Systems - Yaw Damping State the purpose of Yaw Dampers and explain their operation.

Topic 13.3.5

Helicopter Autoflight Systems State the purpose of Stability Augmentation System (SAS) in helicopters and explain its operation.

Topic 13.3.6

Autoflight Systems - Trim Control State the function of Automatic trim control and explain its operation.

Topic 13.3.7

Radio System Interface Explain the operation of Autopilot Navigation Aids interface.

Topic 13.3.8

Flight Management System Interface Explain the operation of Flight Management Systems (FMS) including the navigation database.

Topic 13.3.9

Autoflight Systems- Autothrottle State the purpose of Autothrottle systems and explain their operation.

Topic 13.3.10 Autoflight Systems- Autoland Explain the principles and categories of Automatic Landing Systems:  Modes of operation;  Approach,  Glideslope,  Land and  Go-Around.  System Monitors and Failure Conditions and  Downgrade and Upgrade Procedures.


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TOPIC 13.3.1 - FUNDAMENTALS Auto Pilot Purpose The aircraft autopilot controls the aircraft in vertical speed, attitude, and heading to reduce workload and fatigue on the flight crew and to provide improved flight comfort and stability. The basic principles of an autopilot are: 

Error sensing



Correction



Follow-up and command

Any autopilot follows the basic principles of error sensing, correction, follow-up and command, although different types of sensors, servos are used. The “inner loop” typically handles internal conditions – Attitude sensing, attitude changes in terms of error signals.

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The following are typical inner loop conditions; •

Attitude sensing,

•

attitude changes in terms of error signals,

•

processing of error signals,

•

conversion of error signals into movement of flight control surfaces.


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Roll – Inner Loop

Outer Loop Raw data inputs such as air speed, altitude, magnetic heading and interception of ground based radio aids for instance can also constitute “outer loop” control.


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Modes of Operation Aircraft operating in automatic flight mode are capable of maintaining set operating parameters depending on the stage of the flight. These can include: •

attitude hold

•

heading hold

•

turbulence

•

vertical speed hold

•

airspeed hold:

•

altitude hold

•

control wheel steering

•

navigation:

•

VOR

•

ILS

•

Aux or several other modes as selected by the pilots to suit the particular stage of the flight. A flight director system provides cues for the pilot to navigate and fly the aircraft, but a flight director cannot control the aircraft. Only the pilot or the autopilot system can fly the aircraft.


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Single axis – normally only roll axis – most basic concept – used in small light aircraft – heading hold and basic radio coupling capable Two axis – Typically around roll & pitch – typically heading hold & radio coupling capable – altitude and attitude hold capabilities – sometimes rudder has yaw damper incorporated – this is not considered a 3 axis system. Three axis – attitude control about all three axes – typically full AFCS system Multi axis autopilot – a system which controls an aircraft about the roll and pitch axis (two axis) or roll, pitch and yaw axis (three axis)

Authority Limits may be placed on control signals that are demanded to prevent excessive attitude changes or harsh maneuvering. It is necessary to monitor what is known as the authority of the AFCS. This means that limits must be placed on any control signals that are demanded to prevent excessive changes to the attitude of the aircraft, or to cause any harsh maneuvering. For example, within a typical roll channel incorporating a bank angle limiter and a roll rate limiter. The bank angle limiter limits any input signal to a value which is required by a particular mode of operation. The maximum bank angle limit is 25 degrees to 30 degrees, but can be changed automatically depending which mode is selected. For example, 20 degrees for VOR on course mode. The roll rate limiter limits the rate at which the aircraft changes its bank angle, by limiting the rate at which the servo motor turns. Typical roll rate limit outputs are 1 1/2 degrees per second to 7 degrees per second.

Capture Radio deviation signals are sensed by the AFCS, but are not used until capture occurs. The term capture means the point at which the radio deviation is used by the AFCS as a reference to fly the aircraft.

In the case of the ILS the capture point is determined by the vertical beam sensor in the pitch channel and the lateral beam sensor in the roll channel. The beam sensors are voltage level sensing circuits that satisfy certain switching functions and apply radio deviation to the signal chain.


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Couple Autopilot Principle of Operation To control autopilot steering the flight control computer can utilise many references. Heading, attitude and instrument landing system commands. The basis of operation to maintain a selected reference is typically conducted by selecting a reference and then having the flight control system generate corrective attitude changes whenever a deviation from the selected parameter is selected. For example, the pilot flies onto a heading and engages heading hold. The actual heading and desired heading signals are compared in an operational amplifier and any variation from the desired heading will produce a differential between the two signals (phase or amplitude) which will be applied to a servo amplifier to correct for the deviation. Similarly, a parameter can be selected on an autopilot control box, eg rate of climb. When auto pilot is selected the difference between the actual rate of climb and the selected rate of climb will produce an error signal (from an Op Amp) which will only be nulled when the aircraft is climbing at the same rate as selected. This principle is the basis of all automated flight management. Aircraft actual parameters are applied to Op Amps (or something similar) and are compared with desired parameters whenever automatic pilot is engaged. Whenever a differential between selected parameter and actual parameter is detected the aircraft attitude will be corrected to re-align. In a fully computerised system heading changes can be programmed in advance. Assume an aircraft is programmed to fly from Brisbane to Coffs Harbour on a heading of 180° and upon reaching Coffs Harbour heading is to change to 190° to then track toward Sydney. The parameters are typed into a Flight Management Computer (FMC), the aircraft takes off and heads for Coffs Harbour. When the inertial Reference System determines the aircraft is over Coffs Harbour a signal will trigger the change of heading required and the aircraft flight control system will respond and turn onto the new selected heading of 190° automatically. The maximum rate of turn permissible is typically programmed into the flight control computer so as not to throw the aircraft into violent manoeuvres.

Selecting a New Heading With heading hold engaged, when the pilot decides to select a new heading, he/she selects a new heading by turning the heading marker on the HSI dial. The aircraft is in position A flying on the original heading. The pilot selects a new heading by turning the heading marker on the HSI. The new heading is compared to the aircraft’s actual heading. The AFCS will immediately notice the difference and send an error signal through to the aileron servo actuator to deflect the ailerons and the aircraft rolls and turns onto the new heading. Once the aircraft has reached the new heading, the error signal is nulled and the aircraft returns to its straight and level attitude.


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Couple – Related to the mode of operation – it is the provision of raw data input to the AFCS relevant to a particular flight path


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Runway visual range (RVR). The range over which the pilot of an aircraft on the line of a runway can see the runway surface markings or the lights delineating the runway or identifying its centre line. Decision height (DH). A specified altitude or height in the precision approach at which a missed approach must be initiated if the required visual reference to continue the approach has not been established. Categories of precision approach and landing operations: (a) Category I. DH 200 feet and RVR 2,400 feet (with touchdown zone and centerline lighting, RVR 1,800 feet); (b) Category II. DH 100 feet and RVR 1,200 feet; his Category IIIA. No DH or DH below 100 feet and RVR not less than 700 feet; (d) Category IIIB. No DH or DH below 50 feet and RVR less than 700 feet but not less than 150 feet; and (e) Category IIIC. No DH and no RVR limitation.

NOTESpecial authorization and equipment required for Categories II and III.

Engagement Autopilot Engagement The basic principle of autopilots is to hold the aircraft in basic heading, pitch and roll channel attitude at the time of engagement. An autopilot system is designed so there will be a gradual transition when it is engaged, therefore if heading hold is engaged when aircraft is 90° from selected heading the aircraft will not immediately throw itself into a violent bank to capture the commanded heading. The aircraft will be limited in its rate of heading change to perhaps 3° per second, thereby taking 30 seconds or more to align to the commanded heading. Typically the rate of change of heading can also be selected by the pilot. The same gradual engagement is replicated for any autopilot function. Autopilot Control Panel provides for engagement for the range of autopilot options, eg Heading Hold, Roll stabilisation and Vertical speed hold all engaged simultaneously to control a climb to assigned altitude. Often autopilot cannot be engaged until preset conditions are met, eg roll stabilisation cannot be engaged until bank angle less than 10°, Autoland can only be engaged if Radar Altimeter system functioning, Radar altitude hold and barometric altitude hold cannot be engaged simultaneously.


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Autopilot is engaged by selecting the appropriate switches and buttons to select the autopilot functions desired. Autopilot will not engage if: •

Interlock circuit not complete

•

Circuit breakers not engaged

•

Switches in incorrect positions

Synchronisation The synchronisation of signals is normally performed as a pre-engage function. This is to remove any standing signals that may cause snatching of the controls when initially engaged, and allows the autopilot to take control in a smooth manner. The attitude sensing elements, such as the vertical gyro are continuously monitoring the aircraft’s attitude, and supplying signals to the servo motors. If the aircraft has been placed, as an example, in a nose down attitude prior to engagement, on engaging the autopilot would immediately receive an error signal from the vertical gyro. The computer would then signal the servo motors to move the control surfaces to provide a nose up attitude. It is therefore necessary to oppose the vertical gyro error signal and reduce it to zero before engaging the autopilot. This ensures that the system is synchronised with the attitude of the aircraft.

Amplification Any error signal derived from a detection device or command signal must be amplified, for they do not have the magnitude to operate a servo actuator. The type of aircraft and its handling characteristics will have a great bearing on the way it responds to control surface movement. For example, an empty aircraft will respond quite differently to the same one that has a full load of fuel and passengers on board. It is therefore necessary for the AFCS to have a system which will adjust the ratio of the output signal to the input signal to achieve the desired rate of response. The response of an aircraft to a command or detection signal will therefore be determined by the gain of the amplifier circuits in the flight control computer. The gain or transfer ratio of an amplifier is the ratio of the output to the input. To put it another way, the gain is equal to the transfer ratio, which is equal to output divided by the input. Turbulence An aircraft experiencing turbulent air conditions in flight will have varying degrees of load applied to its structure. In these circumstances, the pilot will adjust speed, power and use of controls to suit the prevailing conditions. If the aircraft encounters turbulent conditions while the AFCS is engaged, the AFCS will correct for any movement that takes place. However, this correction, in these conditions, could possibly lead to additional loads being applied to the structure due to the rate at which the AFCS responds. This can cause the AFCS response rate to get out of phase with the disturbance rate. The pilot would normally disengage the AFCS in these conditions, but in some AFCS there is a mode selection which reduces the gain of the pitch and roll channels thereby softening the response of the AFCS to the disturbance.(Turbulence Penetration)


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Gain Diagram

Adaptive Gain Control This term refers to the way in which the sensing of the handling characteristics in flight is carried out by a system which is classified as either: • semi-adaptive gain control • self adaptive gain control.

Semi Adaptive Gain Control This is a system of adjusting the gain of the amplifier by the use of external signals. It is called gain programming or gain scheduling. As an example, a radio deviation signal could be used to control the gain within the computer. Self adaptive gain control This is a system whereby it monitors its own operation and automatically adjusts its own gain for the best aircraft performance. The system is controlled by a model reference which will adjust the system to give the best response regardless of the flight conditions. Command signals are fed through the model to the system, and as the gain is increased, the error signal is amplified and fed to the control servo. The overall system is known as high loop gain, because maximum gain is needed to get the best response.

Wash Out Wash out is the automatic sensing and opposition of any standing signals existing from attitude signal transmitters. This enables engagement of the Autopilot system without any sudden control surface position change or ‘snatching’ of control surfaces.


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Artificial Feel Artificial feel is used to provide an artificial mechanical feedback between the control surfaces and the pilot. This feel is provided by a variety of methods depending on aircraft complexity and sophistication. The feedback is provided so the pilot can determine what aerodynamic loads are on the control surfaces. This is necessary as in some situations without this feedback the pilot could demand a control surface movement that exceeds the allowable aerodynamic loading.

Control Surface Actuation On small light aircraft the power to move the control surfaces is provided by the pilot. The control column is physically connected to the control surfaces by cables, and the pilot moves the control surfaces by repositioning the control column or rudder pedals. As the control column or rudder pedals are displaced the movement is mechanically transferred to the control surface, the aircraft attitude changes, and the pilot re-centres the control column or rudder pedals when the desired attitude or heading is achieved. Any autopilot function in this simple system is nothing more than the use of trim tabs to trim the aircraft to eliminate excessive drift and to relieve the pilot of the necessity to continually maintain a force on the control column or rudder pedals in order to maintain straight and level flight.

Controls When control column is moved backwards, the elevators are raised thereby decreasing lift of the horizontal stab so that the aircraft is displaced by a pitching moment about the lateral axis into a nose up or climbing attitude. Forward movement of the control column lowers the elevators to increase the lift of the horizontal stabiliser and so the pitching moment causes the aircraft to assume a nose-down or descending attitude. Pitch displacements are opposed by aerodynamic damping in pitch and by the longitudinal stability and as the response to elevator deflections is a steady change of attitude. Elevators are essentially displacement control devices. Control column and control wheel movements are independent of each other so that lateral and longitudinal displacements can be obtained separately or in combination. This concept of having separate control movements for each axes goes right through to the most complex automatic flight control systems with multi axis control.

Servomotors and Servo Actuators On larger aircraft it is physically impossible to move the control surfaces by muscles alone. All aircraft, with the exception of light aircraft or very old aircraft, will incorporate some form of power assistance (like power steering in a car) to move control surfaces. The power assistance is provided by actuators or servo’s and these devices can operate from either/or mechanical input (like your cars power steering) or electrical input. Once a flight control system is capable of repositioning control surfaces by use of electrical signals, these signals can be provided by a number of sensors to control the aircrafts flight path. Instead of the pilot detecting an uncommanded roll or heading change and moving the stick or pedals to counteract it, gyros, accelerometers and other sensor equipment can detect the uncommanded attitude changes far more accurately and then provide an electrical output to an actuator or servo. This is the basis of a fly-by-wire or automatic flight control system. The sensors detect uncommanded attitude changes and counter them. When the pilot moves the control column or rudder pedals (a commanded attitude change) an electrical signal from the stick or pedal transducers is transmitted to the electrically operated actuator and the control surface is deflected by pilot input to achieve an attitude change. In this electrically operated


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system, it is electrical signals not mechanical inputs which control the actuator or servo operation. Going one step further, the electrically operated flight control system can be programmed to fly a specific route, at a specific altitude and then the pilot is simply along for the ride. The avionic systems of the aircraft provides the flight control computer with inputs of heading, altitude, waypoints, etc and the flight control computer repositions the actuators with electrical signals to maintain the aircraft on the programmed flight path. This attribute in an automatic flight control system is called an autopilot.

Actuators and Servos The components used in an automatic flight control system (AFCS) to move the aircraft’s control surfaces are called servomotors, servo actuators, or by the name of the control surface or channel that it controls, for example rudder servo or pitch actuator. The signals received from the AFCS computer are electrical. Therefore the control of the actuators is electrical. The servo actuators convert these electrical signals into control surface movement by converting the electrical signal into mechanical motion which is usually done by torque motors or solenoid controlled valves (electro-hydraulic valves). The three main types of servomotors are: • electromechanical • electro-pneumatic • electro-hydraulic. Electromechanical and electro-pneumatic actuators or servos are more suited to smaller aircraft, and the typical installation in a modern commercial aircraft is an electro-hydraulic system. Even though some of the surface actuators may be electro-hydraulic, it is not uncommon for others to be electromechanical or electro-pneumatic, eg flap motors may be electrically driven and throttle boost actuators may be pneumatically driven when the remaining actuators may all be hydraulically driven (aileron, rudder, elevators).


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Servomotors may be connected in series or parallel with the AFCS. A series servomotor is one that moves the control surfaces without moving the pilot’s controls, whilst a parallel servomotor moves the control surfaces and the pilot’s controls. The most common actuator used on commercial aircraft flight control systems is the electrohydraulic actuator.

Pneumatic & Electric Servos Pallett Automat ic Flight Control pg 200 & 201 Simple servos found on light aircraft use vacuum sources like those which operate the gyroscopic instruments. Pneumatic pressure is obtained from either an engine driven pump or from a tapping at one of the engine compressor stages. The vacuum is directed to pneumatic servos that are mechanically connected to the normal mechanical flight control linkages. The pneumatic servo is an airtight housing which contains a moveable diaphragm. When vacuum is applied to the servo the diaphragm is displaced pulling on the cable to reposition the flight control surface. Two of these servos would be needed for each control surface, a “pull” and a “push” actuator. Another method of driving the control surface of light aircraft is by use of an electric motor. These servomotors may be powered by either AC or DC depending on the type of automatic flight control system used. An electric motor servo can use a reversible DC motor and reduction gearing to supply the force to move the flight control surface in both directions. Alternatively a constant direction motor can be used with magnetically switched clutches to engage a mechanism to apply force to a control cable. The servomotor consists of an electromagnetic clutch, gearbox and drive mechanism. It may also include an amplifier to amplify the command signal and a feedback system such as a potentiometer or tachogenerator. The constant drive type has the advantage that the inertia forces in starting and stopping the motor are eliminated so it can be engaged and disengaged more rapidly and precisely. When the AFCS computer sends a signal to the motor of the electric servomotor, it will drive the gear train and subsequently the control surface in the desired direction. At the same time, it drives a tacho generator to provide feedback to the computer for speed limiting and smoothing. A follow up synchro is also driven by the motor which will send a signal back to


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the computer indicating the actual position of the control surface. The synchro signal is of a phase opposite, but in proportion to the control surface displacement and will null the output signal of the computer when both signals are equal. Thus control surface movement will cease. Both the pneumatic and electric servos are only power assisting servos, with the flight control system still fundamentally powered by the pilots’ muscles. The pneumatic and electric flight control servos are limited to use in only light and simple aircraft. Neither of these examples are a fly-by-wire system. The actuators used in larger and more modern aircraft are typically electro-hydraulic.

Pneumatic & Electric Servos

Pneumatic Servo – when vacuum is applied the diaphragm pulls the cable repositioning the flight control surface Electric Motor Servo – can use a reversible DC motor or a constant direction motor with magnetically switched clutches Electric motor and Pneumatic (vacuum) powered servos normally only incorporated in light aircraft Neither of these are fly-by-wire systems – large modern aircraft typically use electro-hydraulic actuators which are entirely electrically controlled


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Electro-Pneumatic Servo Electro-pneumatic servo; consists of electro-magnetic valve, with dual poppet ports connected via pressure ports & orifices to two cylinders containing pistons sealed against pressure loss Valves controlled by electrical command signals from the auto pilot With no signal present both valves are open


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Electro Mechanical Servo This mechanism consists of a motor coupled to the flight control system via an electromagnetic clutch, a gear train and a sprocket and chain. Feedback is provided by a potentiometer in a DC motor, and a CX synchro and a tachogenerator to provide position and rate feedback signals for AC motors.


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Electrohydraulic Servo-Actuators Power Assisted Powered flight controls These are used in high performance aircraft and consist of two main types: • power assisted • power operated. The main difference in the two systems is the way in which the actuators are connected to the control surfaces Power assisted control In the power assisted system, the pilot’s control stick is connected to the control surface via a control lever. When the pilot pulls back on the stick to begin a climb, the control lever pivots about point X and commences moving the control surface up. At the same time, the control valve pistons are displaced allowing hydraulic fluid to flow to the left hand side of the actuating jack which is secured to the structure of the aircraft. The pressure exerted on the piston causes the whole servo unit and control lever to move to the left, and because of the greater control effort produced, the pilot is assisted in moving the control surface further.


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Electrohydraulic Servo-Actuators - Power Operated Power operated control In this system, the control column is connected to the control lever only, whilst the servo unit is directly connected to the control surface. The effort required by the pilot to move the control column is that needed to move the control lever and control valve piston. The power required to move the control surface is supplied solely by the servo unit’s hydraulic power. As there are no forces transmitted back to the control column, the pilot has no feel of the loads acting on the control surfaces, and a means of artificial feel must be introduced at a point between the control column and the connection to the servo unit control lever.


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Electro-Hydraulic Actuators The transfer valve is an electrically controlled hydraulic valve which operates a piston assembly called the autopilot actuator, which in turn operates the main control valve for the actuating cylinder. The movement of the actuator is monitored by the output of a linear voltage differential transducer (LVDT). This will provide the follow up signal back to the computer. Direct operation of the hydraulic power unit has two main advantages; one is the very low computer power output required and the other is that it is more sensitive and accurate, due to the absence of cable slack, stretch and drag.


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Manual Operation In manual operation, the control column moves the control quadrant, we will assume this relates to a back stick input to move the elevators. The pilot pulls the stick backwards to start a climb (moves the cockpit control to the left on the slide), the control cables will turn the control input quadrant which will move the upper end of the control valve actuator (long green arm) to the left. At this point before the surface begins to move the control surface actuator is hydraulically locked in position (shown still centred on the diagram, because this is before it begins to move) so the control valve actuator (long green arm) is anchored at the bottom. The end result of the top of the arm moving left is that the control valve will be displaced left. When the control valve moves left hydraulic supply pressure is ported to the left hand side of the control surface actuator, which will force the piston to the right.

The pressure applied to the left of the control surface actuator will force the piston to the right moving the control surface. This will move the bottom anchor point of the control valve actuator (long green arm) to the right, and this time the top of the arm is held stationary (pilot still has control column pulled back) so the control valve spool will be moved to the right, thus synchronising again and causing a hydraulic lock on either side of the control surface actuator piston, locking the control surface in the commanded position (whilst ever the pilot maintains the back stick input).


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A Summary of Actuator Operation to this Point When the pilot applied back stick, the control valve ported hydraulic pressure through to the control surface actuator, thus hydraulically driving the control surface to the commanded position. The pilot input is only moving the control valve, it is the hydraulic pressure which drives the control surface actuator. While the pilot retains the stick input the control surface will remain in the position it has been driven to (just like a pilot holding back stick on a cable system – the control surface will remain in the back stick position until the stick is recentred). If no hydraulic pressure were applied you can move the stick all you want, but the only effect it has is to move the control valve. The control surface will not move without hydraulic pressure applied.

The control surface will remain in the back stick position until the stick is recentred. When back stick is released, the same process as previously described occurs again. The bottom of the control valve actuator (long green arm) is locked in place because the control surface is initially still hydraulically locked in the extended position. With the bottom of the control valve actuator (long green arm) locked, the control valve will be displaced to the right, porting hydraulic pressure to the right side of the control surface actuator.

With the bottom of the control valve actuator (long green arm) locked, the control valve will be displaced to the right, porting hydraulic pressure to the right side of the control surface spool piston. The hydraulic pressure applied to the control surface spool piston will force the piston to the left and retract the control surface. This will also reposition the control valve to the left recentreing it and again hydraulically locking the control surface in the central position until the control column is again displaced.


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This description of flight control system operation still refers to a manually operated system. No electrical inputs have been described yet. As you can see actuator operation is dependant upon control valve position. If we can electrically drive the control valve, we can control the actuator with electrical signals alone.

On/Off Solenoid An ON/OFF solenoid is simply a hydraulic relay. With no electrical power applied hydraulic pressure is shut-off because the solenoid spring holds the seat against a seal, preventing pressure from being felt downstream. When power is applied the solenoid coil magnetises and unseats the valve (overpowers spring pressure) and permits hydraulic oil to flow. In the hydraulic actuator, the ON/OFF solenoid provides pressure to the transfer valve when autopilot is activated. Power to the ON/OFF solenoid is typically controlled through a series of monitors which detect any failures in the autopilot system. In the event that an autopilot failure is detected, the ON/OFF solenoid is de-energised, isolating autopilot inputs from the actuator.


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ON/OFF Solenoid Operation When the ON/OFF solenoid is energised, the transfer valve is provided with hydraulic pressure. Typically the pressure provided to the transfer valve nozzle will also be provided through the IN/OFF solenoid, but that is not shown here to reduce complexity of the diagram. When the transfer valve is provided with hydraulic pressure it is primed to convert the electrical inputs into hydraulic outputs which will drive the autopilot actuator – thus controlling the control valve and control surface actuator electrically instead of mechanically as previously described.

Transfer Valve Before looking at the operation of the actuator, we must understand the workings of the transfer valve, which transforms electrical signals from the computer into hydraulic pressure. A transfer valve is also often called a Electro-Hydraulic Valve (EHV), or a hydraulic servo can be driven by torque motors directly connected to the Autopilot Actuator spool. The torque motor style is a high current application though, so the transfer valve and EHV style of electrical interface to the hydraulic actuator are more common in modern flight control installations. On the right hand side there is a coil of windings around a C shaped core. If a signal is presented to this coil, it will move the permanent magnet armature up or down about its pivot. The computer outputs a DC signal and the polarity of the signal determines the direction of movement. Hydraulic fluid is fed into the unit through the feed pipe, passing through a Issue B: January 2008

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flexible tube which then divides across the pointed divider, just under the flexible tube. The feed pipe provides full hydraulic system pressure to the transfer valve nozzle, but it is supplied through only very narrow gauge plumbing because the work it has to perform to manipulate the spool valve is minimal, so a high rate of flow of pressurised hydraulic fluid is unnecessary. If there is no electrical signal to the coil, the flexible tube remains in the neutral position, due to spring loading (represented by the two black lines connecting the nozzle point and the spool valve piston assembly). In this position, the spool valve and feedback springs sense equal hydraulic pressure at both ends and take up the neutral position, closing off both control ports.

Transfer Valve Operation When the flight control computer sends a signal to the coil windings (either drawing the permanent magnet attached to the nozzle up or down) the permanent magnet will rotate Moving the nozzle in a direction dependant upon polarity of the input signal from the computer. In the first illustration this will cause a greater pressure to be directed to the top of the spool valve than at the bottom. This moves the spool valve down. The spool valve will continue to move down until the force of the feedback springs is sufficient to bring the flexible tube back almost to the neutral position. With the spool valve moved down, hydraulic supply pressure is ported out through the upper control port. This pressure is used to control the autopilot actuator spool valve, which will be explained next slide. With the spool valve down, hydraulic pressure is ported out the top control port, and the bottom control port is opened to the hydraulic return line. If the electrical signal is of reversed polarity, the spool valve will move up instead of down, porting hydraulic pressure through the bottom control port and opening the top control port to the hydraulic return line.


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Autopilot Actuator Operation On engaging the Automatic Flight Control System (AFCS), the ON/OFF solenoid opens and supplies hydraulic pressure to the transfer valve. When an AFCS command signal is supplied to the coil windings in the transfer valve, the spool valve is nozzle is displaced. The hydraulic pressure is applied to the right hand side of the autopilot actuator, causing it to move to the left. The control valve actuator (long green arm) pivots on the control surface actuator and moves the control valve to the left. This is the same movement as shown in the manual operation, except that the input to the control valve actuator is provided by the autopilot actuator, not the cockpit control. When the command signal is of the opposite polarity the transfer valve nozzle moves down forcing the spool valve up which ports hydraulic pressure to the left side of the autopilot actuator. When the autopilot actuator moves right the control valve is forced to the right providing pressure to the control surface actuator to drive the control surface. On the actuator illustrated on the slide the electrical inputs driving the transfer valve and the autopilot actuator reposition the control surface, and also reposition the control column in the cockpit. Any corrections made by the flight control computer will be felt at the control column.


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Other actuators are designed so that electrical inputs only move the control surface and have no effect on the cockpit controls. Damper signals are typical of this method of operation. When aircraft oscillations are detected by a gyro, it outputs a signal through the flight control computer to the actuator to counter the oscillations, but the rudder pedals or control column will not me moved. This design of the actuator is not that different from the type described here, but the differences will not be covered in this lesson. As the autopilot actuator moves to the left, the autopilot LVDT produces an electrical output which is sent back to the computer to null the command signal input.


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In Series Operation. Because the pilot has no feel of the aerodynamic loads acting on the control surfaces it is necessary to incorporate an “artificial feel” device at a point between the pilot’s controls and their connection to the servo-unit control lever.

This is commonly a “q” feel unit in which the feel force varies with the dynamic pressure of the air sensed by the Pitot/static system. Q=1/2ΡV2

It monitors hydraulic pressure and produces control forces dependent on the amount of control movement and forward speed of the aircraft. A “q” feel unit monitors hydraulic pressure and produces control forces dependent on the amount of control movement and forward speed of the aircraft. Issue B: January 2008

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In a “q” feel unit the feel force varies with the dynamic pressure of the air sensed by the pitot/static system. An artificial feel unit is connected at a point between the pilot’s controls and their connection to the servo-unit control lever – it restricts movement of the control column.


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Electrohydraulic Servomotors

Servomotors Duplex In some aircraft the possibility of a hard over or runaway condition resulting from automatic flight control malfunctions is prevented by utilising two independent control systems which displace control surfaces via duplex servomotors & differential Gearing. The pitch & roll servo motors are of equal authority and torque, and their outputs are summed by their respective differential gearing. The yaw servomotor is of a single type with torque limiting. When the commands from AFCS computer 1 & 2 to a servomotor are identical (normal operation) the motions of both motors within the servo are identical, so providing doubled authority to operate the appropriate control surface. If however a malfunction in one system occurs such that a hardover roll is commanded by that system, then it will turn the differential gear in the direction commanded. The other system however will (at the outset) detect the undesired attitude change and will command its motor to rotate the differential in the opposite direction with the net result that the deflection is prevented. Each motor is coupled electrically to a sensor known as a speed monitor, which in turn is connected to braking units. The purpose of the monitor is to identify a runaway motor, which it does within about 2 milliseconds, and then to apply a signal to the respective brake thus locking out ½ of the differential gear and enabling the good motor to drive the control surface through its half of the gearing. Since the servo power is halved, then any hard over risk in the remaining control system is reduced.


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Runaway conditions resulting from malfunctions are prevented by utilising two independent control systems which displace control surfaces via duplex servomotors. When commands from AFCS computer 1 & 2 to a servomotor are identical (normal operation) motions of both motors within servo are identical & complimentary Purpose of monitor is to identify a runaway motor (it does within ≈ 2 milliseconds) Applies signal to respective brake thus locking out ½ of differential gear Enabling good motor to drive control surface through its half of gearing No failsafe after one system fails – no monitor to counteract a failure of second motor or computer – additional redundancy typically incorporated – 3 or 4 computer channels Hydraulic servo-actuators also incorporated duplex operation


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Servomotors Duplex

Common command applied to all EHV’s from all 4 channels – all signals identical & complimentary. DP’s monitor EHV outputs – when all pressures balanced (normal operation) – no failure signal generated. Any detected computer fails (in a single channel) turns off the command signal – remaining signals still command actuator with no loss of efficiency: eg LVDT fail, cross channel mismatch wiring short or open cct. If DP sensor detects EHV failure – SOV controlling EHV turned OFF removing hyd pressure from EHV – pressure still provided to MCV & main ram – actuator continues to function on remaining good EHV with no efficiency loss. Bypass damper ports return pressure to MCV to prevent hydraulic lock with EHV deenergised If 2nd EHV fails – 2nd SOV turns OFF & actuator reverts to mechanical operation – input directly to MCV


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No command input – transfer valves in neutral position – pressures/flow rates P1 + P4 = P2 + P3 Failure Sensor centred & LVDT O/P normal Quadruple redundant duplex hydraulic servoactuator. Upper half of actuator – autopilot actuator – mechanical input applied to main control valve from command select mechanism if both EHV’s fail.

Command input & transfer valves respond correctly: P1 + P4 still equals P2 + P3. Failure Sensor centred & LVDT O/P normal If Transfer valve blocked or misaligned: P1 + P4 > P2 + P3. Failure Sensor piston driven from neutral – LVDT O/P to FCC indicates failure – SOV de-energised. Primary function of failure sensor is to detect fault & disengage EHV before uncommanded flight control input effected


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In duplex actuator, remaining non-failed half of actuator continues to function normally – driving control surface.


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Servomotors Duplex Quadruple redundant duplex hydraulic servo actuator Upper half of actuator – autopilot actuator – mechanical input applied to main control valve from command select mechanism if both EHV’s fail. Schematic diagram of hydraulic duplex servoactuator: Components: •

SOV’s

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EHV’s (Transfer Valves)

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DP’s (Failure Sensors)

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System Selector valve

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Mechanical input – mode selector valve

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MCV

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Bypass Damper

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CAS LVDT – Rate feedback

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RAM LVDT – position feedback


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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.


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TOPIC 13.3.2 - COMMAND SIGNAL PROCESSING


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B2-13.3.2 Command Signal Processing Page 1 of 82


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Two moving vane transducers can be mounted on the inner and outer gimbal rings of a vertical gyro to sense pitch and roll attitude changes. A square wave a.c power supply is applied to coils 1 and 2, the voltages being out-of phase. In level flight there is no output but with change of attitude there is relative movement between vane and coil assembly with the resultant output as a command signal to the computer


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The dynamic vertical sensor is a pendulum actuated synchro transmitter the axis of which is aligned with the fore and aft axis of the aircraft. The unit is oil damped and in a coordinated turn the pendulum aligns with the resultant of centrifugal and gravitational forces. At the dynamic vertical there is no signal output but if the aircraft is slipping or skidding the pendulum will be displaced from the vertical with a signal output.


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The rotor of a synchro is attached to the gimbal .

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A torsion bar restricts the motion in one axis.

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Rate gyros detect short term attitude changes, these signals are used to modify displacement inputs, e.g. rudder channel for turn co-ordination.

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They are also the primary sensor for Yaw Damping.


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The gyroscopic sensors used with autopilots are similar to the gyro instruments. These sensors detect aircraft pitch, roll, and yaw motions. Deviations in attitude or in the rate of change from a selected attitude are converted into electrical error signals and sent to the autopilot computer. On small aircraft, the autopilot sensors are frequently built in to the attitude and heading gyros. The latest types of autopilots use sensors that employ laser beams instead of a spinning gyro rotor. these laser sensors, called ring laser gyros, or RLGs. The RLG has two laser beams that travel in opposite directions around a triangular course. Sensitive detectors measure the Doppler shift or frequency change whenever the unit is rotated. One RLG is needed for each axis measured. RLGs are much more expensive than an actual gyro, but they do not precess, and they eliminate the moving parts that cause a conventional gyro to gradually wear out.


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A Hydro-mechanical system – used in DC3/4 A/c. Displacement signals are picked off the air driven gyros as small vacuum inputs to the balanced oil valves which provide hydraulic signals to the control surfaces.


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All automatic flight control systems (AFCS) are based on the closed loop servo system. Their input signals originate from either a command signal device or a gyro stabilisation signal (attitude gyro). The signal is amplified to provide an increase in signal strength to power the correction unit (servomotor) to move the aircraft control surface. The amplifier output is reduced to zero (null) by a follow up system (direct feedback) that cancels the input signal and the control surface movement stops. Since movement of the control surface causes aircraft response, the original input signal will be cancelled by the change in aircraft attitude. The control surface is now returned to neutral by the follow up (direct feedback) system providing the only input to the amplifier to drive the servomotor back to a null or zero signal from the follow up system. Basic AFCS loop consists of the detection, amplification, correction follow up and the aircraft response loop. Response loop. The function of the follow up signal is to cancel the input signal by applying an opposite signal proportional to the command required to bring the aircraft to the desired attitude. Position feedback provides control surface displacement proportional to the strength of the input signal. Rate feedback allows a control surface displacement to be applied at a rate that matches the rate of input signal application. It provides a damping effect on the control surface response, providing tighter control over the aircraft.


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Inner Loop The inner loop is the basis of any AFCS and is responsible for the basic attitudinal stability of the aircraft. Its function may be called a stability augmentation system (SAS) or a damper system. The operation of inner loop is by the sensing of aircraft attitude changes and the transmission of error signals, which is accomplished by the use of gyros, accelerometers and transducers.


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An example of a circuit as applied to a roll control channel is shown in the figure and is one which uses the characteristics of both an operating amplifier in the integrating function, and of servomechanism feedback. Assuming that prior to engagement of the AFCS, the aircraft is at some angle of roll there will be a corresponding roll attitude signal output from the roll sensing transducer of the verticalaxis gyroscope unit, and this is supplied to summing point 2. All other command inputs normally supplied to summing junction 1 are at zero, since the control system is not at this stage coupled to any mode. The roll attitude signal is therefore, an error signal which flows out from summing junction 2 to summing junction 3 via an amplifier and the roll displacement path, and also back to summing junction 1. Since the AP ENG and ROLL HOLD switches are both closed prior to engagement of the control system, the error signal is also applied to the input of the integrating amplifier to drive it until its output to summing junction 2 is equal to the roll attitude signal, thereby zeroing the output from junction 2. During the time that the foregoing synchronisation process is taking place, an error signal is also produced at the output of summing junction 3 as a result of summing the signal along the roll displacement path, with a roll rate signal developed from the roll attitude signal after passing it through a ‘rate taker’ circuit; the purpose of the rate signal is to provide a shortterm damping of servo-actuator operation. The resulting error signal passes to the aileron servo-actuator, via the summing and servo amplifiers, and so it is driven in the appropriate direction; the ailerons are not displaced of course, since with the control system not engaged the servomotor clutch is de-energised.


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At the same time that the motor operates, it drives a position feedback synchro (CX), the output of which is demodulated and fed to summing junction 3 to oppose the error signal which, as a result of the synchronisation process, becomes less than the feedback signal, leaving the latter as the sole means of driving the servomotor until the feedback signal itself is reduced to zero. An example of a circuit as applied to a roll control channel is shown in the figure and is one which utilises the characteristics of both an operating amplifier in the integrating function, and of servomechanism feedback. Assuming that prior to engagement of the AFCS, the aircraft is at some angle of roll there will be a corresponding roll attitude signal output from the roll sensing transducer of the verticalaxis gyroscope unit, and this is supplied to summing point 2. All other command inputs normally supplied to summing junction 1 are at zero, since the control system is not at this stage coupled to any mode. The roll attitude signal is therefore, an error signal which flows out from summing junction 2 to summing junction 3 via an amplifier and the roll displacement path, and also back to summing junction 1. The resulting error signal passes to the aileron servo-actuator, via the summing and servo amplifiers, and so it is driven in the appropriate direction; the ailerons are not displaced of course, since with the control system not engaged the servomotor clutch is de-energised. At the same time that the motor operates, it drives a position feedback synchro (CX), the output of which is demodulated and fed to summing junction 3 to oppose the error signal which, as a result of the synchronisation process, becomes less than the feedback signal, leaving the latter as the sole means of driving the servomotor until the feedback signal itself is reduced to zero. The servo-actuator also drives a tachogenerator which supplies a rate feedback signal to summing junction 4, the purpose of this signal being to synchronise the servomotor operation and to prevent any tendency for it to over shoot its nulled position. Thus, in the synchronised condition, the net signals at summing junctions 2 through 4 are zero, and since the servomotor is stopped at a position synchronised with the datum attitude detected by the vertical-axis gyroscope unit, it can be engaged with the aileron control system without snatching. On engagement, the :AP ENG and ROLL .HOLD switches are opened and so the integrator is isolated from the circuit. since the aircraft is still some angle of roll, the appropriate roll attitude signal is now predominant and from summing junction 3 is able to drive the synchronised servo actuator motor which thereby displaces the ailerons to restore a wingslevel attitude.


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The rates at which different types of aircraft respond to displacements of their flight control surfaces vary between types and their basic handling characteristics. In particular they vary with altitude, speed, aircraft load and configuration, and rate of maneuver. Thus, it is necessary to incorporate ‘gearing’ elements within flight control systems which will adapt them to aircraft and thereby reduce the effects which variations in flight parameters can have on handling characteristics. Similarly, in applying particular types of AFCS to individual aircraft control systems, it is necessary to provide facilities for altering the response of an automatic system to any given level of input signal, thereby obtaining a signal ratio best suited to the operation of the systems when working in combination.


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Such a ratio is known as ‘gain’ and may be considered as having a function analogous to the changing of gear ratios in a mechanical gearing system . The signal path from error to response is known as the system forward path, and the amplification from error to response, measured as amplification ratio, is the loop gain. Satisfactory closed-loop performance depends on determining a loop gain which compromises between long-term accuracy plus initial response, and acceptable settling time plus limited overshoot. These factors, in turn, require sufficient inherent damping in the load. Certain adjustments of command and feedback signals can be pre-set within amplifier and/or computer units in order to produce gain factors which establish a basic ‘match’ between an AFCS and aircraft characteristics. Adjustments are based on the variation of electrical resistance at appropriate sections of signal circuits, and as in several types of control system, this is accomplished by means of potentiometers located on a calibration panel that forms an integral part of an amplifier or computer unit.


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In addition to the foregoing adjustments, it is also necessary, particularly when a control system is operating in any of the outer loop control modes, for gain factors to be altered automatically to offset variations in handling characteristics resulting from changing flight conditions. This process is called ‘gain programming or scheduling’, and is part of a technique referred to as adaptive control. The diagram illustrates what is termed a self-adaptive control system and it is one which is capable of changing its parameters throughout an internal process of measurement, evaluation, and gain adjustments, without direct sensing of changing flight characteristics. The overall response of the system is optimized, irrespective of flight conditions, by means of an electrical analogue system referred to as a ‘model reference’. The model defines the optimum dynamic behaviour of the aircraft, based on selected response characteristics subject to any constraints that the airframe may impose, and the control system parameters are adjusted to match the response of the aircraft to that of the model.


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The program is accomplished by using signals from a low range radio altimeter which are supplied to a gain programmer control section in a vertical path module of the pitch control channel. After the GS mode is engaged plus 10 seconds, the GS deviation beam signal is modulated and amplified, and supplied as a pitch down command signal. Initially, the gain of the beam deviation amplifier is zero, but it then increases to, and is held at, 100% until the aircraft descends to 1,500 ft radio altitude.


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The altitude signal then produced by the radio altimeter develops a bias voltage which is applied to the GS beam deviation amplifier so that its gain is reduced as the aircraft descends (graph ‘A’). The gain programmer control section of the vertical path module also supplies a gain control signal to a lateral path module in the roll control channel in order that the gain of the LOC beam deviation amplifier may also be reduced. In this case, the reduction is gradual from 100% to 57%.


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In the event that there is an invalid signal from the radio altimeter, a time programme control is developed (see graph ‘B’). In respect of the GS, this programme starts 10 seconds after GS has been engaged, and results in an initial gain of 80%, decreasing to approximately 20% over a period of time of 120 seconds. The LOC beam deviation amplifier gain programme is initiated directly at GS engage, and after 120 seconds time period the gain is reduced to 57%.


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Under automatically controlled flight conditions, it is necessary to monitor what is generally termed the ‘authority’ of the control system notably in respect of the roll and pitch channels; in other words, limits must be placed on commanded control signals to prevent excessive attitude changes and harsh maneuvering. In the example of roll control channel shown in the figure, it will be noted that there are two limiting elements in the signal processing chain: a roll command rate limiter, and a roll command limiter. The rate limiter limits the rate of change of command signals to some selected value, e.g. 5 degrees/second, so as to ‘soften’ aileron displacements and prevent harsh rolling of the aircraft. Signal processing through the limiter network is such that it imposes a specific time constant on the roll command signal. The roll command limiter controls the roll angle authority of the control system, the limit which the circuit is actually capable of being dependent on the control mode selected. Limits are accurately pre-set, and when the control channel is operating in the appropriate mode they are controlled by a d.c. bias signal applied to limiting diodes within the limiter module.


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This system is provided in some AFCS to enable the pilot to maneuver the aircraft in pitch and roll through the AFCS by exerting normal pressure on his control wheel. On releasing the control wheel, the AFCS will hold aircraft at the newly established attitude.


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When power is applied to an AFCS, the attitude sensing elements are monitoring the aircraft’s attitude. The sensing signals are supplied to the appropriate AFCS channel and servo amplifier. Any output signal from the servo amplifier is also fed to the appropriate trim indicator. The purpose of the trim indicator is to provide an indication of signals being supplied to the servomotors, whether in the engaged or disengaged conditions, and also to indicate any out of trim conditions of the aircraft under normal operating conditions of the control system.


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Trim Sensor This consists of a fixed mounting plate with two adjustable contacts and a sliding bar with a common contact which is attached to pulleys that ride on the up and down elevator cables.

Trim Servo The servomotor is a geared DC reversible motor installed on a mounting bracket together with its amplifier. The output from the motor is transmitted to the elevator trim tab via a capstan connected to the trim tab control cables.

The elevator channel command signal to the elevator servomotor is used as the appropriate command signal to the pitch autotrim. Its direction is sensed by up trim and down trim sensor circuits, which operate relays that apply a signal to the trim servomotor to run it in the appropriate direction. The signal is pulse width modulated, the pulse duration being a function of airspeed.


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If an aircraft under auto-pilot control encounters turbulence it will sense the turbulence as disturbances to aircraft attitude, but in applying corrective control it is possible for additional loads to be imposed. In this mode the gain of pitch and roll signals is reduced gain, therefore softening flight control system response.


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Axes of an Aircraft An aircraft in flight is controlled within three stabilized planes. Movement within each plane is about an axis, rather than centred on an axis. All three axes pass through the centre of gravity (C of G). The three principal axes are: •

longitudinal (X axis) which runs nose to tail through the C of G

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lateral (Y axis) which runs parallel with a line from wing tip to wing tip and intersects the X axis at the C of G

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normal/vertical (Z axis) runs perpendicular to the other two axes intersecting them at the C of G.

Roll Movement around the longitudinal axis is called rolling, its control or stability is called the lateral stability, and is controlled by ailerons.

Pitch Movement around the lateral axis is called pitching, and the control or stability is called the longitudinal stability, and is controlled by elevators.

Yaw Movement around the vertical axis is called yawing. Control or stability is called the directional stability and is controlled by the rudder.


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Flight Control Surfaces In straight and level flight, a flight control surface can be considered as an extended aerofoil. When the control surface is deflected, the amount of lift produced by the extended aerofoil either increases or decreases depending on the direction of the control surface. Movement – If the control surface is deflected down the shape of the aerofoil is extended so the amount of lift produced increases. If the surface is deflected up the aerofoil shape is distorted and the amount of lift is decreased. Primary control surfaces consist of the following: ailerons/spoilers, elevators, rudder, stabilators, canards, elevons, ruddervators. Ailerons/spoilers – The roll attitude of an aircraft is controlled by ailerons, and are operated by sideways movement of the pilot’s control column. They are found on the trailing edge of the wing near the wing tip. Placing them near the wing tip gives greater airflow deflection, meaning only a small column movement will create a large amount of roll. They are connected in opposition to each other, so that when the left aileron is raised the right one is lowered. When the aileron is lowered lift is increased and the wing will rise. The raised aileron will assist the lowered one to roll the aircraft. Movement of the control column is instinctive. If you move the column to the left, the aircraft will bank to the left and vice a versa. Spoilers can be used as a substitute for ailerons, and operate on the same principle as ailerons, except they only extend upwards. When the spoiler of one wing is raised, the wing drops due to the loss of lift. Elevators – These control the longitudinal attitude of the aircraft, and are coupled together and are found on the trailing edge of the tail planes horizontal surface. They are operated by means of fore and aft movement of the pilot’s control column. Pulling the column back will cause the elevators to rise, which will decrease the amount of lift on the tail, which will drop, causing the nose to rise. Thus, the aircraft will climb. Pushing the column forward has the opposite effect, and the aircraft will drive.


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Rudder – This controls the direction of the aircraft in the same way as the rudder on a boat. The rudder is hinged at the rear of the vertical fin, and is operated by the pilot’s rudder pedals. Pushing the right pedal forward causes the aircraft to turn right. Stabilators – Some aircraft have dispensed with the elevators on the tail plane, and replaced them with a horizontal surface that moves in its entirety and is called a stabilator. It is mainly used in high performance fighter aircraft. Canards – Some aircraft have an additional set of wings set forward of the C of G instead of a tail plane. They are used to give additional lift to the main wings, and improve the handling of the aircraft at low and high speeds. Elevons – These are combined control surfaces that act as both elevators and ailerons on delta winged aircraft. Ruddervators – These are a combined control surface to operate in pitch and yaw, and fitted to aircraft that have a butterfly of V tail.


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Control Surface Actuation On small light aircraft the power to move the control surfaces is provided by the muscles in the pilots arms and legs. The control column is physically connected to the control surfaces by cables, and the pilot moves the control surfaces by repositioning the control column or rudder pedals. As the control column or rudder pedals are displaced the movement is mechanically transferred to the control surface, the aircraft attitude changes, and the pilot recentres the control column or rudder pedals when the desired attitude or heading is achieved. Any autopilot function in this simple system is nothing more than the use of trim tabs to trim the aircraft to eliminate excessive drift and to relieve the pilot of the necessity to continually maintain a force on the control column or rudder pedals in order to maintain straight and level flight.

Control Surface Actuation When control column is moved backwards, the elevators are raised thereby decreasing lift of the horizontal stab so that the aircraft is displaced by a pitching moment about the lateral axis into a nose up or climbing attitude. Forward movement of the control column lowers the elevators to increase the lift of the horizontal stabiliser and so the pitching moment causes the aircraft to assume a nose-down or descending attitude. Pitch displacements are opposed by aerodynamic damping in pitch and by the longitudinal stability and as the response to elevator deflections is a steady change of attitude. Elevators are essentially displacement control devices. Note on the illustration that control column and control wheel movements are independent of each other so that lateral and longitudinal displacements can be obtained separately or in combination.


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This concept of having separate control movements for each axes goes right through to the most complex automatic flight control systems with multi axis control


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Servomotors and Servo Actuators On larger aircraft it is physically impossible to move the control surfaces by muscles alone. All aircraft, with the exception of light aircraft or very old aircraft, will incorporate some form of power assistance (like power steering in a car) to move control surfaces. The power assistance is provided by actuators or servo’s and these devices can operate from either/or mechanical input (like your cars power steering) or electrical input. Once a flight control system is capable of repositioning control surfaces by use of electrical signals, these signals can be provided by a number of sensors to control the aircrafts flight path. Instead of the pilot detecting an uncommanded roll or heading change and moving the stick or pedals to counteract it, gyros, accelerometers and other sensor equipment can detect the uncommanded attitude changes far more accurately and then provide an electrical output to an actuator or servo. This is the basis of a fly-by-wire or automatic flight control system. The sensors detect uncommanded attitude changes and counter them. When the pilot moves the control column or rudder pedals (a commanded attitude change) an electrical signal from the stick or pedal transducers is transmitted to the electrically operated actuator and the control surface is deflected by pilot input to achieve an attitude change. In this electrically operated system, it is electrical signals not mechanical inputs which control the actuator or servo operation. Going one step further, the electrically operated flight control system can be programmed to fly a specific route, at a specific altitude and then the pilot is simply along for the ride. The avionic systems of the aircraft provide the flight control computer with inputs of heading, altitude, waypoints, etc and the flight control computer repositions the actuators with


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electrical signals to maintain the aircraft on the programmed flight path. This attribute in an automatic flight control system is called an autopilot.

Linear Variable Differential Transformers (LVDT) A linear variable differential transformer (LVDT) is an electromechanical device which translates straight line motion into a linear Alternating Current (AC) signal (proportional to amount of movement). Transformer theory explains how an alternating current passed through a coil induces a current into a coiled conductor in the vicinity. The expanding and contracting magnetic field in the primary coil induces a current into the secondary coil. If the magnetic flux is concentrated in an iron (ferrite) core, in lieu of just a hollow air gap the transformer is more efficient and a stronger signal is induced into the secondary winding. The strength of the signal induced into the secondary winding is therefore variable by inserting and removing the ferrite rod core. This is the basis of operation of an LVDT

LVDT Operation By incorporating two secondary coils (or a single coil with a centre-tap) whenever one end of the secondary is positive the other end will be negative. If the signals from each end of the coils are measured and compared to earth, the two signals will be of equal amplitude and frequency, but of opposite phase. If the two signals are combined the resultant will be zero volts because the two signals will cancel each other out. The two signals will only be of equal amplitude when the ferrite rod is in the centre of the secondary coil. If the rod is displaced in either direction, one of the secondary coils signals will be stronger than the other, and the resultant signal will be indicative of direction (the Issue B: January 2008

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phase indicates this) and amount of movement (amplitude is proportional to amount of movement). The AC signal produced by the LVDT can then be rectified and combined with the initial error signal applied to the transfer valve, nulling it out – this is called feedback.


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Force Transducers In an E & I bar force transducer, the principle of operation is virtually the same as for an LVDT. The AC input signal is applied to the centre winding on the E bar and the outer legs support the secondary windings. In the force transducer pictured, any input on the left hand end will move the I bar with respect to the E Bar because the outer case of the transducer is designed to expand and contract as force is varied. The magnetic relationship between the E & I bar will vary, producing output signals in the same manner as the LVDT.

In an E & I bar force transducer, the principle of operation is virtually the same as for an LVDT. The AC input signal is applied to the centre winding on the E bar and the outer legs support the secondary windings. In the force transducer pictured, any input on the left hand end will move the I bar with respect to the E Bar because the outer case of the transducer is designed to expand and contract as force is varied. The magnetic relationship between the E & I bar will vary, producing output signals in the same manner as the LVDT.


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Example of signal interface between sensor and control surface. Sensor electrical output can be from LVDT, E & I bar, Synchro, etc. Signal amplified and applied to servomotor. Feedback signal fed back to amplifier – nulling error signal (Op amp for example) In modern aircraft – error signal digitised within computer. Computer processor interprets attitude change requested by pilot and sends appropriate analogue signals to servoactuators to achieve required attitude change. Has advantage of gain scheduling signals before applying them to servoactuators, and computer can select appropriate control surface to move to achieve attitude change most efficiently – FCC software load. Although not new in concept, complete re-development of fly-by-wire systems has been necessary in recent years, as a means of controlling some highly sophisticated types of aircraft. Mechanical linkages to the control surfaces required would have been prohibitively complex, thus fly-by-wire systems where wires carry the electrical signals from the pilots controls, replaced mechanical linkages entirely.


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In a fly-by-wire system the control column will not be physically connected to the servo actuator, the control column and rudder output is only applied to an LVDT (or similar transducer) to convert the mechanical movement into and electrical signal proportional to the degree of movement. This electrical signal is then transmitted to the Flight Control Computer where the phase is detected to determine the direction in which the control column has moved. The amplitude of the signal represents the distance the control column has been displaced. Once phase detected the signal is rectified and amplified. The now DC signal is applied to the transfer valve to reposition the control surface as explained on the previous slide. When the autopilot Actuator displaces (due to the hydraulic force applied from the Transfer Valve) the autopilot LVDT will be displaced and the signal generated within the autopilot LVDT will be phase detected, amplified and rectified and applied to the servo amp, to oppose or null out the initial signal generated by the control column LVDT. When the feedback voltage nulls the initial “error” voltage there will be no servo amplifier output, so the transfer valve will recentre. The actuator will remain in the extended position until the control column is released, then the autopilot LVDT signal applied to the servo amp will drive the transfer valve, recentreing the autopilot actuator and control surface until the autopilot LVDT is again at its null (centred) when the servo amp will no longer have an input from either the control column LVDT or the autopilot LVDT This concept is called closed servo loop operation. The feedback signal opposing the initial input signal nulls it out, thus closing the loop of operation. When flight control surfaces are aligned the control column LVDT’s and autopilot LVDT’s must all be calibrated and aligned to the null position, this alignment is called rigging.


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Typically a flight control surface will have several flight control channels, so the amount of LVDT’s will be trebled. That is there will be three for the control column in the pitch axis, three in the roll axis and three connected to the rudder pedals. The Hornet aircraft has 4 flight control computer channels to operate 2 Ailerons, 2 Rudders 2 leading edge flaps, 2 trailing edge flaps and two stabilators, and has a total of approximately 100 LVDT’s within the flight control system for position feedback, rate feedback and error sensing monitors.

Position feedback servo loop – simplified using DC potentiometer theory. Initial DC signal from stick or gyro or accelerometer to amp. Amplified & drives Servo. Servo drives load, and also repositions feedback potentiometer. Feedback potentiometer voltage opposes initial input voltage – nulling input signal Rate feedback tacho generator signal – highest amplitude at greatest RPM – highest signal output when servo motor and hence control surface is moving at highest speed. Highest speed of surface provides the greatest strength feedback signal – damping


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Position feedback servo loop – simplified using DC potentiometer theory. Initial DC signal from stick or gyro or accelerometer to amp. Amplified & drives Servo. Servo drives load, and also repositions feedback potentiometer. Feedback potentiometer voltage opposes initial input voltage – nulling input signal Rate feedback tacho generator signal – highest amplitude at greatest RPM – highest signal output when servo motor and hence control surface is moving at highest speed.


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Comparison of signals to achieve position and rate feedback to achieve fast response & critical damping


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Automatic Flight Control System Elements The basic AFCS consists of the following major elements: Detecting element This consists of attitude gyros, rate gyros, accelerometers, pitot static systems and air temperature probes. They detect the movement of the aircraft about its three flight axes and the rate of that movement. It is considered to be the internal controls or inner loop of the system. Command element This element consists of the pilot’s control panel and manually operated controls, that include control column, rudder pedals, navigation systems and radar. It is considered to be the external controls or outer loop of the system. Computer amplifier This is the brain of the system. It computes, amplifies and processes the signals from the detecting and command elements and directs the output element to respond to the pilot’s and/or system’s requirements. Output element This element consists of the units which move the control surfaces of the aircraft in response to the computer demands. These servo units, as they are known, can be electric motors, electromagnetic solenoid valves controlling hydraulic actuators. Fly by wire This system refers to the replacement of mechanical linkages by wires that carry electrical signals from the pilot’s control stick to the servomotors. The movement and forces created by the pilot on his controls are measured by electrical transducers. The signals are amplified and then sent to the various hydraulic actuator units which are directly connected to the control surfaces. When an aircraft is able to be mechanically controlled (through electrohydraulic power controlled actuators) and electrically controlled (fly-by-wire), when engaging the AFCS the transition to electric fly-by-wire must be smooth and without violent control surface deflections. The AFCS engagement controlling circuitry will align to the current control surface positions (and will incorporate any current control column or rudder inputs). With no command inputs (stick or pedals) the AFCS will only steer to straight and level flight gradually (non straight and level flight detected by sensors, which all develop error signals) without violent or instantaneous maneuvering. Issue B: January 2008

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These units are used to measure the aircraft’s rate of turn. If the nose of the aircraft is deflected, the gyro will sense this, and signals will be sent to the rudder to move the aircraft back to the correct heading. The rate gyro output shown below the aircraft represents the 400 Hz synchro signal developed during an aircraft’s flight path. It also shows the rate gyro’s DC voltage output. When the aircraft is flying in a straight line, there is not signal output from the synchro, but when the aircraft turns, there is a signal with a constant amplitude.

On an aircraft is which is constantly changing direction, and as the rate of turn is constantly changing so is the output from the rate gyro synchro. After the signal has been demodulated and filtered, as seen in the DC graph, we have a DC voltage that is constantly changing value and polarity with each aircraft turn. The DC values are at their highest when the rate of turn is greatest. This signal is the one used by the autopilot to eliminate the condition known as dutch roll. Pallet Automatic Flight Controls Pg 25


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All automatic flight control systems (AFCS) are based on the closed loop servo system. Their input signals originate from either a command signal device or a gyro stabilization signal (attitude gyro). The signal is amplified to provide an increase in signal strength to power the correction unit (servomotor) to move the aircraft control surface. The amplifier output is reduced to zero (null) by a follow up system (direct feedback) that cancels the input signal and the control surface movement stops. Since movement of the control surface causes aircraft response, the original input signal will be cancelled by the change in aircraft attitude. The control surface is now returned to neutral by the follow up (direct feedback) system providing the only input to the amplifier to drive the servomotor back to a null or zero signal from the follow up system. The aircraft drift has therefore been corrected by the sensor output and when the aircraft is returned to the selected attitude by the control surface displacement, the control surface is returned to the neutral position. Basic AFCS loop This consists of the detection, amplification, correction follow up and the aircraft response loop. The function of the follow up signal is to cancel the input signal by applying an opposite signal proportional to the command required to bring the aircraft to the desired attitude. Position feedback Provides control surface displacement proportional to the strength of the input signal. Rate feedback This allows a control surface displacement to be applied at a rate that matches the rate of input signal application. It provides a damping effect on the control surface response, providing tighter control over the aircraft.


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Accelerometers Inertial Navigation Systems simply measure the magnitude of aircraft acceleration, and then use that information to compute velocity and distance traveled. The graphic on the slide shows a simple pendulum. You can visualize what happens to the pendulum if it is attached to a vehicle which starts to move forward, the pendulum swings towards the rear of the vehicle. This characteristic is explained by Newton’s First Law of Motion, which states: “A body at rest tends to remain at rest and a body in motion tends to remain in motion in a straight line unless forced to change its state by an external force”. Newton’s second law of motion states: “The acceleration of a body is directly proportional to the force causing it and inversely proportional to the mass of the body.” Newton’s second law applies to the pendulum in the vehicle on the slide and we can use the pendulum to measure acceleration. The magnitude of acceleration can be measured because the distance moved by the pendulum is proportional to the applied force. That is, the greater the acceleration, the further the pendulum will move rearwards. A device used to measure acceleration is called an accelerometer. Note that an accelerometer will measure acceleration only – that may seem like an obvious statement but it is important to realize that once the vehicle is in a steady state of motion (or rest) the pendulum will hang vertically. For example, let us assume our vehicle accelerates from rest to 100 km/hr, then cruises at that speed for a short period before stopping. When the vehicle is accelerating the pendulum will swing backwards. However, during cruise at 100 km/hr the pendulum will hang vertically as velocity is constant and acceleration is zero. When the brakes are applied the pendulum swings forward and the harder the brakes are applied the further forward it will swing. Finally, when the vehicle is stationary the pendulum hangs vertical again. Accelerometers are sensitive only along a single axis. This axis of maximum sensitivity is known as the sensitive axis of the accelerometer. To measure N-S & E-W accelerations, two accelerometers are required. An INS measures acceleration in longitudinal and Lateral axes and plots aircraft movement with respect to aircraft velocity (speed and direction or heading). Because the functionality of an INS basically depends on gravity as the constant, it is unaffected by wind and atmospheric conditions like a pitot/static system. An INS can very accurately detect any induced accelerations, eg drift and cross wind so can calculate groundspeed, heading and time (the 3 requirements for a dead reckoning system) with exacting accuracy, providing a reliable dead reckoning navigation reference.


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Accelerometers The accelerometer is basically a pendulous device. When the aircraft accelerates, the pendulum, due to inertia, swings off its null position. A signal pick off device will tell how far the pendulum has moved. There are many different types of accelerometers. Figure on slide shows the progression from a simple pendulum to an E and I pick off device. The E and I pick off device accelerometer is an with its armature spring loaded to a null position by two leaf springs. The output windings are connected in phase opposition to each other so that when the armature is in a null position, the resultant output signal is null.

Accelerometers Movement to one side causes the signal of one phase to dominate the other. The illustration on the slide shows the movements and subsequent outputs of the accelerometer. Accelerometer A is indicating that is it accelerating to the left and the armature, due to inertia, is lagging to the right. Because of the position of the armature relative to the windings, the output on the right will be greater. Accelerometer B is accelerating to the right, so obviously the output on the left will be greater. Sometimes two accelerometers are used in the one system. In this configuration, one accelerometer acts as a reference and the other is used to detect accelerations in another location. The illustration on the slide shows a pair of accelerometers connected in such a way, that if both experience movement in the same direction, the output signal will be a null. If they experience different accelerations, then there will be output, which is used by the autopilot computer. With no acceleration detected by the accelerometer, the output signals from each of the capacitive pickoffs are equal in amplitude but opposite in phase. Therefore, the output from the acceleration restoring amplifier will be zero and the output from the accelerometer will be


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zero. During accelerations the pendulum, due to its inertia, will begin to swing away from the null position. The greater the acceleration the further it will tend to swing.


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The movement of the pendulum will be detected by the pickoffs causing an imbalance in the output signals from the capacitive devices which are applied to the accelerometer restoring amplifier (ARA). The output signal from the restoring amplifier is applied to the rebalance torquers which generate a torque that will return and hold the mass at its null position. The current required by the rebalance torquers to maintain the pendulum at the null position is proportional to the amount of acceleration that the accelerometer is experiencing. This is very useful, we now have an electrical signal that is proportional to acceleration. In summary, when no acceleration is present, the inertial mass will be centred. The signals generated by the pickoffs will be equal in amplitude but opposite in phase. Therefore, the output from the restoring amplifier will be zero. When acceleration is present the following takes place: Pendulum tends to move away from centre pick-offs generate an imbalance signal which is applied to the amplifier output from the amplifier is applied to the rebalance torquers. Torquers generate a torque which holds the mass at the null position signal from the amplifier is proportional acceleration.


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Air Data Sensor (this is simply ADC info regurgitated) In their earliest form ADC’s where mechanical computers filled with gear trains, cams, aneroid sensors, temperature sensors, etc and were extremely sensitive and delicate components. Parameters were sensed and mechanically converted to electrical signals (potentiometers, etc) for output to other systems or for display. Modern ADC’s are fully electronically computerised. Although the sensing elements are still mechanical type devices, the transducers which convert the mechanical inputs to electrical signals are much more sensitive and do not load down the mechanical sensor, so are also more accurate. Having mostly slid state components ADC’s are very reliable components, rarely suffering unserviceabilities. ADC’s typically receive inputs from Pitot/static system, AOA, OAT (or TOT, Impact Temp, etc) ADC’s also typically receive inputs providing information on flap position, undercarriage position, etc to compensate AOA and pitot/static measurements for position error. The ADC computes air data information from the above inputs to help create a clear picture of flight characteristics. All parameters are measured, converted to electrical signals then digitised for transmission to FMC’s or for display on a VDU screen or HUD. The computer processor performs all these functions, thereby replacing many of the old style analogue instruments such as the ASI, Altimeter, VSI, OAT gauge, AOA gauge. The outputs can be used as inputs to a Flight Control Computer (autopilot), or thrust management computer to automatically maintain an airspeed or altitude, or to fly the aircraft to a selected altitude. ADC’s (as do all other computer systems) constantly perform Self-monitoring tests, often called Built-in-test (BIT) to monitor the serviceability of sensors and processing circuitry., On the indicator pictured, the following displays are available. Analog moving tape displays: IAS: 20-350KTS, Altitude: –1,000 to 32,000 Ft, IVSI: –9,000 to +9000 FPM, TAS, computed Density altitude, OAT: –50° to +50° C, MPH, KTS or KPH, Inches Hg or mm Hg, Mode C altitude encoding, Altitude warnings, Mach number Speed warnings.


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Computer Technology in Aircraft The pilot moving the control stick in the cockpit generates control inputs. The inputs are electronically sent to the flight control computer which senses what the pilot wants to do – bank, dive or climb. The flight control computer sends its signals to the aircraft’s flight control surfaces, rudder, ailerons, elevators or stabilizers, which move to perform the desired manoeuvre. The surfaces selected to perform the manoeuvre are dictated by the software program. During aircraft design and test flying of prototypes, phases of flight are tested within the flight envelope and the software program to manage the flight control system is designed for efficiency and flight safety. Eg selecting a spoiler or aileron to achieve a roll will depend on the rate of roll requested by the pilot (degree of stick deflection), airspeed, AOA and OAT to name but a few. Amount of deflection will also vary greatly with the same stick input depending on the phase of flight. Considering a fighter aircraft, a full surface deflection at low speed will result in a high G turn. The same surface deflection at high speed would likely tear the control surface off, or severely overstress the aircraft. With a CPU managing the flight control surfaces all factors effecting control of flight can be considered before the electrical signal is transmitted to the control surface actuator to deflect the surface the desired amount. It is the flight control computer software load which selects the most efficient method to achieve the attitude change requested by the pilot. Another added benefit is that the flight control surfaces can be easily controlled by other electronic avionic equipment, for example to fly an approach down an ILS localiser and glideslope; autopilot functions of attitude hold, attitude hold, heading hold, etc; coupling the flight control system to the navigation system to fly a pre-programmed course and track.


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In fighter aircraft flight control computers can recover from a stall or out-of-control (flat spin) by configuring flight control surfaces such that there is the greatest possibility of the aircraft recovering to controlled flight. The surface positions would be as determined by test pilots during the flight testing of prototypes. In large passenger aircraft the FCS system is programmed to avoid entering a phase of flight where a stall or out of control condition is a possibility, eg stick pushers, stall warning and stall avoidance systems.

Fly-By-Wire Benefits The Airbus fly-by-wire system offers significant tangible benefits in terms of greater safety, through unconstrained control input freedom within the flight envelope, and protection against exceedance of operating limits, stalling, overspeeding or overstressing the aircraft outside the envelope. To these can be added windshear protection, reduced pilot workload, lower costs and improved aircraft performance. It makes a major contribution to reducing maintenance costs by eliminating much of the complex mechanical system of cables, pulleys and associated gear which need post-maintenance rigging work and checks.

Sidesticks Because the pilot does not have to physically force the control column to deflect the control surfaces with his/her muscular effort, a simple little computer joystick (sidestick) can be incorporated to control the aircraft, with electronic outputs applied to a flight control computer. Replacement of old-generation control columns by modern sidesticks has the benefits of an unobstructed view of the instrument panel and a slide-out working table in front of each pilot.


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Damping Systems When a flight control system can be managed by electrical signals, the output of sensor systems can be harnessed to automatically correct for any variations from the intended attitude. This can take the form of damper systems to counter porpiosing, Dutch-roll or turbulence induced motion where gyro or accelerometer outputs are used to automatically correct for any non-pilot induced attitude changes (any variations not initiated by control column or rudder input). Consider an aircraft flying in a straight and level flight attitude with the AFCS engaged. If a sudden gust of wind should move the aircraft, the attitude gyros would sense the movement and send a signal to the computer. The computer will process this information and send a signal to the servomotors to move the appropriate control surface to bring the aircraft back to its original attitude. The servomotors will send a feedback signal to the computer, telling it that the control surface has been displaced.

Note When the aircraft is returned to its original attitude (before the gust of wind hit), this will null the original error signal produced by the gyro. This is called aerodynamic feedback. Where the attitude change of the aircraft initially produced an error signal detected by the gyro, the response by the flight control system to counter the drift will null out the initial gyro error signal. Damper correction will not be felt back through the control inputs (control column or rudder pedals). The correction will be applied to the servoactuator to drive a control spool which is not mechanically connected back to the pilots controls, so the control surface will correct for the sensed variation in attitude keeping the aircraft in straight and level flight, but the pilot will not feel any of the correction inputs in the cockpit.


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Autopilot Engagement The basic principles of autopilots are to hold the aircraft in basic heading, pitch and roll channel attitude at the time of engagement. An autopilot system is designed so that there will be a gradual transition when it is engaged. if heading hold is engaged when aircraft is 90° from selected heading the aircraft will not immediately throw itself into a violent bank to capture the commanded heading. The aircraft will be limited in its rate of heading change to perhaps 3° per second, thereby taking 30 seconds or more to align to the commanded heading. Typically the rate of change of heading can also be selected by the pilot. The same gradual engagement is replicated for any autopilot function. Autopilot Control Panel provides for engagement for the range of autopilot options, eg Heading Hold, Roll stabilisation and Vertical speed hold all engaged simultaneously to control a climb to assigned altitude. Often autopilot cannot be engaged until preset conditions are met, eg roll stabilisation cannot be engaged until bank angle less than 10°, Autoland can only be engaged if Radar Altimeter system functioning, Radar altitude hold and barometric altitude hold cannot be engaged simultaneously, etc. Autopilot is engaged by selecting the appropriate switches and buttons to select the autopilot functions desired.

Autopilot Principle of Operation To control autopilot steering the flight control computer can utilise many references. Heading, attitude, instrument landing system commands, etc. The basis of operation to maintain a selected reference is typically conducted by selecting a reference and then having the flight control system generate corrective attitude changes whenever a deviation from the selected parameter is selected. For example, the pilot flies onto a heading and engages heading hold. The actual heading and desired heading signals are compared in an operational amplifier and any variation from the desired heading will produce a differential between the two signals (phase or amplitude) which will be applied to a servo amplifier to correct for the deviation. Similarly, a parameter can be selected on an autopilot control box, eg rate of climb and when auto pilot is selected the difference between the actual rate of climb and the selected rate of climb will produce an error signal (from an Op Amp) which will only be nulled when the aircraft is climbing at the same rate as selected. This principle is the basis of all automated flight management. Aircraft actual parameters are applied to Op Amps (or something similar) and are compared with desired parameters whenever automatic pilot is engaged. Whenever a differential between selected parameter and actual parameter is detected the aircraft attitude will be corrected to re-align. In a fully computerised system heading changes can be programmed in advance. Assume an aircraft is programmed to fly from Brisbane to Coffs Harbour on a heading of 180° and upon reaching Coffs Harbour heading is to change to 190° to then track toward Sydney. The parameters are typed into a Flight Management Computer (FMC), the aircraft takes off and heads for Coffs Harbour. When the inertial Reference System determines the aircraft is over Coffs Harbour a signal will trigger the change of heading required and the aircraft flight control system will respond and turn onto the new selected heading of 190° automatically. The maximum rate of turn permissible is typically programmed into the flight control computer so as not to throw the aircraft into violent manoeuvres.


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Selecting a New Heading With heading hold engaged, when the pilot decides to select a new heading, he/she selects a new heading by turning the heading marker on the HSI dial. The aircraft is in position A, flying on the original heading. The pilot selects a new heading by turning the heading marker on the HSI. The new heading is compared to the aircraft’s actual heading. The AFCS will immediately notice the difference and send an error signal through to the aileron servo actuator to deflect the ailerons and the aircraft Rolls and turns onto the new heading. Once the aircraft has reached the new heading, the error signal is nulled and the aircraft returns to its straight and level attitude.


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Autopilot Coupling ILS Localiser element The transmitter is located at the far end of the runway. To direct an aircraft onto the centreline of the runway, the transmitter radiates azimuth guidance signals to the left and right of the centreline. The signal transmitted to the left has a 90 Hz signal superimposed on it, and a 150 Hz signal is superimposed on the signal transmitted to the right. The two transmissions overlap along the centreline. When an aircraft is approaching, the ILS receiver receives both signals at equal strength. This is indicated on the indicator. If the aircraft deviates to the left of the centreline, the strength of the 90 Hz signal will be greater than that of the 150 Hz signal. Both signals pass through a comparator circuit, (Op Amp as described last slide) which then produces an output, causing the vertical bar of the indicator to deflect to the right. This tells the pilot to fly right to intercept the centerline. The same occurs if the aircraft deviates to the right, except that the 150 Hz signal becomes the stronger. Glidepath element The transmitter is located at the threshold of the runway. The transmitter radiates a pattern similar to that of the localiser, but they provide vertical guidance above and below the decent path at an angle of 2.5° to 3°. When the aircraft is approaching along this path, 90 Hz and 150 Hz are received at the same strength. The same conditions occur as with the localiser, except the indicator will indicate up and down deviations. By following the displayed commands, a pilot is able to carry out an ILS approach to an airport runway. In order to carry out an approach under automatic control, it is necessary for the AFCS to be coupled to the ILS. The signals from the ILS are purely command signals varying in amplitude with displacement from the beam centres, but they have no directional properties and cannot take into account the heading of the aircraft. It is therefore necessary Issue B: January 2008

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for the pilot to align the aircraft with the runway in heading (this function can be carried out automatically on modern computerised aircraft). Upon intercepting the localiser beams the deviation pointer of the ILS indicator will display a fly left command. The pilot flies the aircraft onto the appropriate heading, and now the ILS system can automatically control the aircrafts approach to the runway threshold. The localiser mode is selected so that the AFCS roll channel will respond to the result of the beam signal and heading error signal. When the signals are in balance, the aircraft will fly straight and level on the intercept heading. As the aircraft enters the normal width of the beam, the signal is reduced and the runway heading signal causes the aircraft to turn towards the centre of the beam until both beam and runway heading signal are in balance. Any deviation from the runway heading or localiser beam will produce an error signal and the deviation will be corrected so the aircraft will follow the centre of the localiser beam right to the runway threshold. The AFCS will also be receiving signals from the glide path transmitter, and these are fed to the pitch channel. The glideslope signals will keep the aircraft on the optimum 3° glideslope. Any deviation (so the 90 or 150Hz signals are imbalanced) will produce an error signal, and when applied to the AFCS system to deviation will be automatically corrected.


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Autoland –During cruise and initial stages of approach to land, the control system operates as a single channel system, controlling the aircraft about its pitch and roll axis and providing the appropriate flight director commands. As multichannel operation is required for an automatic landing, at a certain stage of the approach, the other two channels are armed by a switch on the flight control panel. This will also arm the localiser and glideslope modes. Both the off line channels are continually supplied with the relevant outer loop control signals and operate on a comparative basis the whole time. Altitude information essential for vertical guidance to touchdown is always provided by signals from a radio altimeter which becomes effective as soon as the aircraft’s altitude is within the altimeter ’s operating range. When the aircraft has descended to 1500 feet radio altitude, the localiser and glideslope beams are captured and the armed off line control channels are then automatically engaged. The localiser and glideslope beam signals control the aircraft about the roll and pitch axis so that any deviations are automatically corrected to maintain alignment with the runway. At the same time, the autoland status displays LAND 2 or LAND 3 on the indicator and computerised control of flare is also armed. At a radio altitude of 330 feet, the aircraft’s horizontal stabliser is automatically repositioned to begin trimming the aircraft to a nose-up attitude. The elevators are also deflected to counter the trim and to provide pitch control in the trimmed attitude. When the landing gear is 45 feet above the ground (gear altitude), the flare mode is automatically engaged. The gear altitude is based upon radio altitude, pitch attitude, and the known distance between the landing gear, the fuselage and the radio altimeter antenna. The flare mode takes over pitch attitude control from the glideslope, and generates a pitch command to bring the aircraft on a 2 feet/second descent path. At the same time, a throttle retard command signal is supplied to the auto throttle system to reduce engine speed. Prior to touchdown and about 5 foot gear altitude, the flare mode is disengaged and there is transition to the touchdown and roll out mode. At about 1 foot gear altitude, the pitch attitude of the aircraft is decreased to 2 degrees, and at touchdown, a command signal is supplied to the elevators to lower the aircraft’s nose and so bring the nose wheel in contact with the runway and hold it there during roll out. The AFCS remains in control until disengaged by the pilot.


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Control Wheel Steering (CWS) Mode CWS is an operating mode for the autopilot in addition to the command operating mode. The command mode is the normal autopilot mode where the pilot does not touch the controls because autopilot is flying the aircraft. In CWS mode the controls are moved by the pilot in normal flight and the force applied to the controls is measured and used as an input signal to the autopilot (Flight Control) computers. In effect the human pilot is flying the aircraft but the autopilot is helping to move the control surfaces. The transducer on the control wheel which converts the input force into an electrical signal is typically an LVDT or similar sensor. This system is provided in some AFCS to enable the pilot to manoeuvre his aircraft in pitch and roll through the AFCS by exerting normal pressure on his control wheel. On releasing the control wheel, the AFCS will hold the aircraft at the newly established attitude. The pitch and roll forces that the pilot applies are sensed by transducers which will create output signals proportional to the forces and are supplied to the pitch and roll channels of the AFCS. In some cases, limits may be imposed, for example. if a roll angle is less than 5°, the wings levelling will automatically occur and the control system will hold the aircraft on an established heading one second after the wing are level. Prior to and during capture phases of radio navigation modes, the pilot can use the CWS to override the AFCS. This means the pilot will then always have control and the aircraft does not have to follow a pre-programmed flight path.


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Autopilot Override The autopilot can be overpowered at any time by the pilot if he moves his cockpit control with enough force. A typical overpowering force would be 25 to 35 pounds of turning force on the control wheel, or 40 to 50 pounds of force on the control column. If the pilot overpowers the autopilot, in the preceding example, he would be pushing the autopilot actuator back toward the right. In doing so he would increase the hydraulic pressure in the pressurized side of the actuator enough to open the top relief valve (just below and to the right of the auto-pilot actuator). The relief valve would then dump the excess pressure into the return line. The operating pressure of the relief valve determines the amount of force required to overpower the autopilot. The actual overpowering operation is considerably more complicated than indicated here, but this is the basic principle involved. It is actually arranged so that if the autopilot is over-powered, a portion of the autopilot actuator moves the LVDT center slug hard over, developing a high LVDT signal. The autopilot LVDT and the control surface LVDT are so arranged that, in normal operation of the actuator by the autopilot, these two signals are equal. But, if the autopilot has been overpowered, the autopilot LVDT signal becomes very high, and it is then quite easy to detect electrically that the autopilot has been overpowered, resulting in the flight control computer disengaging the autopilot to permit manual control inputs commanded by the pilot.


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Typical Autoflight Components Flight Control Computer This is the brain of the system. It computes, amplifies and processes the signals from the detecting and command elements and directs the output element to respond to the pilot’s and/or system’s requirements. Trim indicator The trim indicator carries out the above functions by monitoring the outputs of the servo amplifiers and producing deflections of the pointers from this zero datum marks in response to the supplied signals. The illustration on the slide depicts a three channel trim indicator showing all pointers aligned with the zero datum marks. The pointers which symbolise the flight control surfaces are actuated by DC milliammeters and are deflected each time servo amplifier command signals are applied to the servomotors. When the servo commands are satisfied, the signals are balanced out and the pointers return to their zero trimmed positions. If a servo amplifier produces a continuous correction signal to keep the attitude of the aircraft, the appropriate pointer will be continuously deflected showing that the aircraft is out of trim. Electrohydraulic Servo Actuators already described in detail AFCS and Autopilot Sensors AFCS sensors are gyro’s accelerometers and components considered to be in the inner loop. Autopilot sensors are all typically outer loop components and systems, eg navigation systems, landing systems, attitude sensor (typically INS), Air Data Computer, Radar Altimeter, etc Autopilot Disengage Switch typically located on control column. Provides readily accessible disengagement of autopilot as required.


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Flight Director Display Flight director information is displayed on Attitude Indicator and can be used for Monitoring the Autopilot and for providing visual commands for the pilot to follow so as to achieve selected profile of flight path as programmed (flight director cannot fly the aircraft – only pilot or autopilot can do this). Heading bug and course lines can be set in the HSI and coupled to the autopilot system to fly a pre-programmed course. A flight director uses sensors and computers but does not command servos to correct for deviations from flight path, it only displays the deviation. Autopilot Control Panel provides for engagement for the range of autopilot options, eg Heading Hold, Roll stabilisation and Vertical speed hold all engaged simultaneously to control a climb to assigned altitude. Autopilot functions can be adjusted utilising latest installations, eg dial in airspeed, altitude, heading, etc. In older style autopilots the aircraft would only assume the altitude, speed, etc at the time of engagement. In modern systems, the aircraft can be flown just by inputs to the autopilot control panel. Typically located on uppermost part of instrument panel, in the centre. The mode control panel of a modern A/C provides the point at which the pilot programmes the A/P-F/D into which mode it is to operate. Also this information is passed to the flight director so as to give the pilot a visual presentation and allow monitoring


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Flight Management System The Boeing 767 FMS has the capability of automatically controlling the aircraft from just after takeoff through to roll out after landing at the destination airport. The human pilot must then taxi to the appropriate gate. This system may not be used on any or all flights, but it has the capability to perform as described. The FMC’s provide the following functions: Flight Planning – the entire flight can be programmed into the computer using a cockpit keyboard (FMC Control Display Select Unit) Performance Management - the system can provide optimum profiles for climb, cruise, descent and holding patterns. A minimum cost flight can be flown automatically by using optimum climb and cruise settings. Navigation Calculations – The FMC can calculate great circle routes, climb and descent profiles etc. Auto Tune of VOR & DME – The FMC can automatically tune the radios to the correct station frequencies required for the air route flown (data stored on navigation software which is upgraded monthly).The FMC is in effect the master computer which integrates the functions of air data computers, inertial reference units, navigation computers, EICAS or ECAM computers and thrust management computers. Flight Control Computers – are the autopilot computers and there are typically three of them. Each computer is entirely independent of the others in that it has its own dedicated inputs and provides its own outputs. All computers function in unison during normal operation, although their outputs are compared against each other so that a malfunction resulting in an erroneous signal will be recognised immediately, and the malfunctioning computer (or sensor) will be isolated from the flight control system. Thrust Management Computer – automatically sets the proper thrust level for the engines. The electrical output from the TMC commands a servo which moves the throttle linkage to set the appropriate level of engine power as calculated by the TMC. The system includes engine mounted sensors measuring important engine operating parameters (TIT, N1, EPR, Fuel Flow, RPM, etc). Engine parameters are monitored to prevent exceeding any engine operating limitation. Autothrottle can be used to maintain a given climb rate, indicated airspeed, Mach number or Descent rate. The autothrottle supports autoland capability and will maintain optimum AOA and airspeed for approach to the runway threshold and will automatically close the throttles just prior to landing to ensure a smooth touchdown. The autopilot and autothrottle can be engaged separately or in unison depending on what automatic parameters the flight crew wish to utilise.


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Define the difference between an AFCS and the Autopilot to eliminate confusion between the two. A flight director system cannot fly the aircraft, it only provides directions for the pilot to follow. FLY-BY-WIRE BENEFITS The Airbus fly-by-wire system offers significant tangible benefits in terms of greater safety, through unconstrained control input freedom within the flight envelope, and protection against exceedance of operating limits, stalling, overspeeding or overstressing the aircraft outside the envelope. To these can be added windshear protection, reduced pilot workload, lower costs and improved aircraft performance. It makes a major contribution to reducing maintenance costs by eliminating much of the complex mechanical system of cables, pulleys and associated gear which need post-maintenance rigging work and checks. SIDESTICKS Because the pilot does not have to physically force the control column to deflect the control surfaces with his/her muscular effort, a simple little computer joystick (sidestick) can be incorporated to control the aircraft, with electronic outputs applied to a flight control computer. Replacement of old-generation control columns by modern sidesticks has the benefits of an unobstructed view of the instrument panel and a slide-out working table in front of each pilot. There will be an inbuilt delay time that must be satisfied before you can engage the system. This allows for the warming up of amplifiers, correct speed and erection of gyros. Issue B: January 2008

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Engagement will also depend on the selection and coupling of other systems into the AFCS, the capture, hold and lock in of various radio navigation aids. It is provided as a safety measure to ensure a smooth transition from manual to automatic flight.


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TOPIC 13.3.3 - MODES OF OPERATION

The vertical gyro provides roll displacement signals The accelerometer is rate damping The manual turn control enables selection of a banked turn Control from roll computer to aileron power unit with aileron position feedback


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Basic stabilisation mode is selected whenever the auto-pilot is engaged without outer loop mode selection. It can be of two types: 1. Attitude hold (aircraft maintains roll attitude at engagement) or 2. Wings level ( if banked aircraft will roll out to level flight).


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Heading Hold: When selected on the Mode Control Panel (MCP) this locks the aircraft heading to the current magnetic heading as sensed by the Directional Gyro (DG).


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Modes of Operation

The Turn Command Knob must be in the centre detent position before the autopilot can be engaged. Turn Command is initiated by pilot operation of the MCP turn knob. The aircraft will disengage from heading and attitude hold modes and roll to the angle commanded. Yaw displacement during the turn is sensed by a rate gyro or pendulous accelerometer and appropriate rudder compensation is applied. When the displacement as measured by the Vertical Gyro satisfies the Turn Knob input the aircraft will maintain this bank angle until another command is executed.


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During manual turn versine data crossfeeds from roll channel to pitch channel to compensate for the loss of lift in roll.


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VOR mode is an enroute radio mode, when selected the aircraft will maintain a path along a radio beam to or from the tuned VOR station. The VOR stations transmit on the 108 to 118 MHz band of frequencies. The ILS [Instrument landing system] comprises LOC [Localiser] and GS [Glide slope]. LOC mode is coupling of the autopilot to a radio signal for lateral guidance to the airport runway. The Localiser stations transmit on the 108 to112MHz band of frequencies The GS [Glide slope] operates in the 330 to 334MHz band with each GS station paired with a LOC station.


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Pilot reports that when a roll command is initiated by Turn Knob, aircraft continues to roll past selected bank angle. Reason: Loss of aileron position feedback signal.


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The patterned area represents a ten second period after capture where the autopilot initiates a 700 feet per minute descent


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Pitch Command is initiated by input from the A/P MCP Pitch Control Wheel. The aircraft establishes a pitch attitude when the VG signal matches the Pitch Wheel position and control surface position.


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Vertical speed mode when selected on the MCP locks the A/C to a vertical speed as selected on the MCP. The A/C is commanded to pitch up or down to satisfy the signal generated by the Air data Computer baro-rate sensor. Vertical speed mode when selected on the MCP locks the A/C to a vertical speed as selected on the MCP. The A/C is commanded to pitch up or down to satisfy the signal generated by the Air data Computer baro-rate sensor. Airspeed hold when selected on the MCP locks the A/C to the speed at selection as sensed by the Air Data Computer. At high altitudes Mach hold is used the air data computer measures a composite signal of airspeed and altitude. This then becomes the Mach reference speed for this mode.


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Altitude hold mode is selected on the MCP, this locks the aircraft to the altitude at selection. Any deviation is sensed by the baro-altitude capsule in the Air Data computer (ADC) and sent to the autopilot computer for processing as a pitch command output to maintain altitude. Chaser motor holds null until Alt Hold engaged – then Alt error is signal.


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Airspeed hold when selected on the MCP locks the A/C to the speed at selection as sensed by the Air Data Computer.


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At high altitudes Mach hold is used the air data computer measures a composite signal of airspeed and altitude. This then becomes the Mach reference speed for this mode. As the Mach No. increases the stab trims more nose up.


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The Mach Trim System automatically senses increases of speed above a specific Mach number and, through servo coupling, automatically re-adjusts the position of the horizontal stabiliser, maintaining the pitch trim of the aircraft. There will be two methods of Mach trim described. The first method uses a change in the aircraft's centre of gravity. The second method trims the horizontal stabiliser.


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The mach trim system provides nose up commands to the horizontal stabiliser to compensate for normal pitch down tendency that occurs with swept-wing aircraft at high mach numbers. This tendency is called mach tuck or tuck-under and is apparent above speeds 0.7M.


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The purpose of the yaw damper system is to improve the airplane’s directional stability and ride quality. The yaw damper system commands small rudder movements as required to ensure correct turn coordination, to correct for Dutch roll and to suppress body structural modal oscillations. It uses sensor inputs from the inertial reference system (IRS), air data computers (ADC) and dedicated accelerometers for computing the commands.


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TOPIC 13.3.4/6 - YAW DAMPING AND TRIM CONTROL


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Pallett Automatic Flight Control Pg 214 All aircraft are subject to dutch rolling, particularly those having a swept wing configuration. First paragraph details why some aircraft do not require yaw dampers to counter dutch roll. Some aircraft designed such that they exhibit stable characteristics & return to straight & level flight following dutch roll motion without need for manual or automatic correction For some aircraft however, natural damping of dutch roll tendency is dependant not only on size of vertical stab & rudder, but also on aircrafts speed, with damping being more responsive at higher speeds than at lower speeds. In these cases corrective action must be taken. Corrective action requires displacement of rudder in order to assist vertical stab in its stabilising function and this action is referred to as yaw damping. Usual for these aircraft to incorporate two axis control system with control about the third axis being provided by a yaw damper system. System is designed independently so it can be operated independently of the Automatic control system so that in the event the aircraft is flown manually the dutch roll tendencies can still be counteracted. The system may be switched on by selecting the damper position on the main engage switch on the AFCS panel or by selecting a separately located switch. Primary component is yaw damper rate gyro, coupled to rudder servoactuator or electromechanical servo. Logic circuits filter, integrate, synchronise & demodulate the servo amplifier signal. Servo amplifier output signal supplied to transfer valve of rudder servoactuator. This unit differs from the standard servoactuator in that it has an additional actuator (yaw damper) & does not include the automatic control system engage mechanism. Summing lever combines pilot inputs using rudder pedals and yaw damper rate gyro inputs Issue B: January 2008

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Whenever aircraft yaws, rate gyro senses motion and provides output to electronic circuitry coupling gyro to servoactuator. Rate gyro output is filtered so only those oscillations at a frequency corresponding to aircrafts dutch rolling tendency are permitted to pass, this eliminates the rate gyro from countering normal turns. Signal modified (gain scheduled) by airspeed to reduce signal as airspeed increases – reduces degree of rudder deflection. Transfer valve in actuator responds to yaw damper signal, driving actuator to respond to rate gyro detected deflection, and counter it. LVDT produces feedback signal to cancel rate gyro input when actuator has moved required amount. Feed back signal also supplied to display element to indicate direction of deflection of yaw damper actuator. When yaw damper oscillation has been countered, LVDT output has no rate gyro signal to oppose, so causes transfer valve to drive back to neutral position. Rudder pedals are not displaced during this process – the system is of a series connected type. Rate gyro output is filtered so only those oscillations at a frequency corresponding to aircrafts dutch rolling tendency are permitted to pass, this eliminates the rate gyro from countering normal turns. Signal modified (gain scheduled) by airspeed to reduce signal as airspeed increases – reduces degree of rudder deflection. Transfer valve in actuator responds to yaw damper signal, driving actuator to respond to rate gyro detected deflection, and counter it. LVDT produces feedback signal to cancel rate gyro input when actuator has moved required amount. Feed back signal also supplied to display element to indicate direction of deflection of yaw damper actuator. When yaw damper oscillation has been countered, LVDT output has no rate gyro signal to oppose, so causes transfer valve to drive back to neutral position.

Rudder pedals are not displaced during this process – the system is of a series connected type. When the servo actuator is used in a series operation, it is principally used as a rudder or yaw damper actuator. The rudder and yaw damper actuator differs from the typical hydraulic servoactuator discussed in previous lessons, in that when it operates it does not move the rudder pedals.


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Yaw Damping and Trim Control


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When the servo actuator is used in a series operation, it is principally used as a rudder or yaw damper actuator. The rudder and yaw damper actuator differs from the typical hydraulic servoactuator discussed in previous lessons, in that when it operates it does not move the rudder pedals. The yaw damper actuator and control valve are both shown with caging springs to the right, and the white sections of the springs’ assemblies are moveable towards each other. If no force is applied, they remain against the stops in the body of the actuator, positioning the control valve or yaw damper in the neutral position. Moving the rudder pedals rotates the heavy long arm about its pivot shaft at the bottom, and at the lower end of the shaft, a short lever moves the summing lever. The yaw damper actuator holds still because the caging spring maintains it in the neutral position. This will result in the summing lever pivoting at its upper ball joint. As it pivots around the upper ball joint, it pushes the control valve to one side or the other, allowing pressure to be present at one side of the main actuator. This will cause rudder movement and follow-up action without the movement of the yaw damper actuator. A signal from the yaw damper to the transfer valve causes the upper black spool in the transfer valve to move to one side of the other. This puts pressure to one side of the yaw damper actuator piston assembly. When the piston assembly moves, it moves the top of the summing lever which pivots at the ball joint in the middle. The summing lever then moves the main control valve causing the rudder to move. It can be seen that this action moves the rudder, but not the rudder pedals, and the actuator is known as operating in series.


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Balancing of aerodynamic forces & moments and establishing desired flight attitudes are continuous processes &, as we have already observed, are governed by the degree of inherent stability of an aircraft & by the maneuvering capability afforded by the primary flight controls system., In flight, however, control must be exercised over changes in weight & C of G locations which occur as a result of consumption of fuel, disposition of passengers and cargo and flight under asymmetric conditions, etc in addition the attitude changes resulting from lowering the flaps must be controlled.

Although the required control could be maintained by repositioning the relevant primary flight control surfaces, varying degrees of physical effort on the part of the pilot would be needed to keep the control surfaces in specifically displaced positions. It is usual therefore to provide secondary control system which can be separately adjusted so that it will displace the primary control surfaces thereby reducing the effects of aerodynamic loads on the primary control system, and so relieving the pilot of undue physical effort. The operation of such a system is referred to as trimming, and some typical methods by which it is operated are described in this lesson.


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Although the required control could be maintained by repositioning the relevant primary flight control surfaces, varying degrees of physical effort on the part of the pilot would be needed to keep the control surfaces in specifically displaced positions. It is usual therefore to provide secondary control system which can be separately adjusted so that it will displace the primary control surfaces thereby reducing the effects of aerodynamic loads on the primary control system, and so relieving the pilot of undue physical effort. The operation of such a system is referred to as trimming, and some typical methods by which it is operated are described in this lesson.


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Pallett Automatic Flight Control A flying tail is one which controls both maneuvering in the pitching plane, and trimming, by means of a variable incidence stabiliser. Elevators may also be provided but can only be operated by direct movement of the stabiliser itself, thereby supplementing its control functions rather than serving as an independent pitch maneuvering control surface. Stabiliser incidence is varied either by appropriate movements of the control column or by rotation of a pitch trim wheel. In each case, movements are finally transmitted via a common differential unit & a hyd servomotor to the actuator jack.

Pallett Automatic Flight Control In manually controlled flight, trimming is usually effected about the three axes, under automatically controlled conditions it is generally confined to control about the pitch axis only. In most cases it is accomplished by a separate trim servo motor coupled to the elevator trim tab system and operated in parallel with the elevator servo motor. For aircraft which vary horizontal stabiliser angle of incidence, a separate trim servo motor is coupled to the stabiliser servo motor. Illustration on slide is control system for trimming horizontal stab. Trim servo is 3 phase dual speed dual winding motor which operates in parallel with elevator power control unit.


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Operation – Pallett Automatic Flight Control

Pallett Automatic Flight Controls In aircraft which are capable of flying at high subsonic speeds and supersonic, larger than normal rearward movements of the wing centre of pressure occur & in consequence larger nose down pitching moments are produced, the attitude change generally being referred to as ‘tuck under’, or ’transonic tuck’.


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Pallett Automatic Flight Controls Attitude change is corrected & trimmed by designing aircraft so that they have the essential stability characteristics and trimming method, eg variable incidence horizontal stabiliser. However at certain mach numbers, compressibility effects arise which make counteracting nose-up pitching moment produced by trimming the horizontal stabiliser to a negative angle of attack position, less effective as aircraft speed increases. Under manually controlled flight conditions, this would necessitate the pilot having to make prolonged trim changes and to hold higher forces on the control column when displacing the elevators relative to a specific trimmed condition. It is therefore normal practice to incorporate what is called a Mach Trim system. Which automatically senses increases in speed above the appropriate datum mach number and, by means of servo coupling, automatically re-adjusts the position of the horizontal stabiliser thereby maintaining the pitch trim of the aircraft.


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Pallett Automatic Flight Controls Read about the Concorde method of achieving Mach Trim on pg 48 During transition to supersonic speed – nose down moment is counteracted by transferring fuel from front trim tank to a rear trim tank moving C of G rearwards thereby removing the force couple introduced by the movement of the centre of pressure


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Pallett Automatic Flight Control Flaps can be lowered to increase the camber of wing sections. In many cases they can also be extended so that the wing area is increased. Most common form is the fowler flap. Lowered and extended by hydraulic actuators or on smaller aircraft by electric actuators When flaps are lowered, AOA is reduced. As a result of increased wing camber there is a lower stalling angle for each increase in flap angle. The change in pressure distribution & rearward movement of the increased lift force also produces a nose down pitching movement which is generally opposed by an increase in the downwash at the tail of the aircraft, this being an added effect of lowering the flaps. Although lift is increased by lowering flaps, drag also incr4eases progressively but under approach & landing conditions this can be used as an advantage, and for each type of aircraft flaps are lowered to corresponding angular positions. At the approach position, the increase of lift enables a lower approach speed to be made & since stalling speed is decreased the aircraft can touchdown at a lower speed. For landing the flaps are fully lowered & a higher drag permits a steeper approach without speed becoming excessive, in other words the flaps have a braking effect during approach & during the period an aircraft floats before touch-down appropriate. By incorporating slots between the flaps, the air flowing over the entire surface of the wing (& flap) is kept smooth (refer to landing position illustration). At each point where the air flows over the leading edge of the extended flap, the flow is laminar, whereas if there was no slot the air would have broken away & become turbulent by that stage.


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Pallett Automatic Flight Control When flaps are operated there is a change in pitch attitude. It is usual therefore to incorporate a flap compensation circuit. Flap compensation circuits vary between types of automatic control system, but in general a circuit is activated by passing a DC signal through a switch controlled by the flap system and feeding it to the automatic pitch trim actuator, in a number of circuits the signal is also fed through a timer circuit to correspond to the specific ‘flaps in motion’ time period.


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Pallett Automatic Flight Control When flaps are operated there is a change in pitch attitude. It is usual therefore to incorporate a flap compensation circuit. Flap compensation circuits vary between types of automatic control system, but in general a circuit is activated by passing a DC signal through a switch controlled by the flap system and feeding it to the automatic pitch trim actuator, in a number of circuits the signal is also fed through a timer circuit to correspond to the specific ‘flaps in motion’ time period. On circuit flap compensation accomplished by a switch activated when flaps selected to down position, this causes a speed change relay to de-energise and thereby connect the 115 VAC supply to the trim servomotor via its fast speed winding. When flaps selected up the motor is powered through the low speed windings to remove the trim input.


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TOPIC 13.3.5 - HELICOPTER AUTOFLIGHT

Fixed and Rotary Wing Compared There is a lot of difference between fixed wing AFCS and rotary wing AFCS. Firstly, there are no ailerons or elevators to move, and instead of a rudder we have a tail rotor. On the helicopter, the controls have to operate the pitch change mechanism of both the main and tail rotor. Instead of a signal being sent to a servo actuator to move a control surface, in the helicopter, a signal is sent to a linear actuator, which shortens or lengthens the control linkages. The helicopter has another axis, collective, to consider. It means the AFCS has to be able to control the aircraft in vertical flight and in the hover.


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AFCS Components The components that make up an AFCS will vary depending on how complex the manufacturer, or indeed customer, requires it to be. However, a basic system will contain the following, and there are at least two systems operating to ensure safety.

Vertical Gyro This provides attitude signals in pitch and roll.

Directional Gyro This provides heading information to the autopilot, flight director, and other navigation systems.

Air Data Computer This provides altitude and airspeed information derived from pitot and static pressure sources.

Radar Altimeter This supplies absolute altitude above terrain to the flight control system, and radar altimeter indicator.

Flight Control Computer This processes information provided by the actual attitude of the helicopter as compared to the commanded attitude. This is required by the selected flight mode to produce autopilot pitch, roll, yaw, and collective control outputs. It also provides flight director pitch, and roll, steering command outputs.

Autopilot Static and dynamic stabilisation must also exist after a single system failure, which means that a dual autopilot system is needed. There is another advantage in having a dual system which is the fact that the failure response is softened if a hard over failure occurs, which allows the pilot more time to act than on a single system. The basic autopilot controls the helicopter by signals from an external source such as altitude, airspeed, etc. The number of modes incorporated in the system depends on the type of helicopter and the manufacturer. Most will have altitude signals from a radio/radar altimeter fed into the autopilot height hold, along with airspeed signals from an air data computer. A difficult task for the pilot is to hold the helicopter on heading by use of the rudder pedals. So by preventing yaw rates from developing, the pilot can reduce the workload involved in maintaining heading. The only difference between the yaw axis and the other axes are the sensors that are used, and the means of allowing manoeuvring. The sensors used will include some sort of compass system, rate gyros, and/or attitude gyros. To allow the pilot to change the heading, various methods are used. One way is to have a force threshold on the pedals which the pilot has to overcome; another is to have switches on the pedals that the pilot has to press. Yet another is to use the force trim switch on the cyclic stick to remove the pedal position signal from the system.

Automatic Flight Control System (AFCS) Most of the modern day helicopters now have advanced flight control systems fitted. These systems not only have an autopilot to control the aircraft, but have a flight director system, usually integrated with the autopilot, as well. Not only will the system reduce the pilot’s work load, but the system can be programmed to fly a specific flight path. The AFCS will combine autopilot and flight director functions for stabilisation and automatic flight path control. It will have four axis control, pitch, roll, yaw, and collective, and provide automatic approach to the hover and automatic climb out With Doppler radar and long range navigation systems fitted, fully automatic approaches to designated areas, terminating in a hover can be made. The Issue B: January 2008

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AFCS provides pitch, roll, and yaw stabilisation and collective power control. The system is made up of dual vertical gyros, dual flight control computers, and dual control actuators in each axis. Control position transducers connected to the cyclic stick, anti torque pedals, and collective stick provide position information to the computers. Commands for flight path guidance can also be coupled into the computers.

Autopilot Controller This is used to engage various autopilot flight control functions, to carry out preflight tests, and to couple and de-couple flight director commands.

AFCS Indicator This panel displays the position of the pitch, roll, or yaw actuators in relation to their centre of travel.

Rate Gyro This provides a rate of turn signal to the autopilot and the EADI.

Accelerometer This provides a voltage output proportional to aircraft lateral acceleration.

Series Actuator This extends or retracts the control linkages on command of the flight computer to alter the pitch change mechanism.

Artificial Feel and Trim Actuator This provides artificial feel and control linkage trim facility.

Primary Servo Actuator This is used to assist the pilot in moving the swashplate.

Pedal Damper Trim Actuator This damps yaw rates and provides a trim facility.

Navigation Receiver This contains VOR, ILS, and marker beacon receivers.


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Stability Augmentation Systems (SAS) Conventional helicopters are inherently unstable, particularly in pitch and roll, and can be tiring to fly for long periods. To overcome this, simple rate damping systems can be utilized and these will prevent any rates of pitch or roll from developing. However, these can also cause problems; such problems were overcome with the development. of Stability Augmentation Systems (SAS), which provide corrective control inputs proportional not only to the rate of change of attitude, but also to the deviation from a datum attitude. The rate gyroscope is still the basic sensor, but the signals from it are integrated, using a 'leaky integrator', to simulate an attitude signal that can be used as a datum. With the system in the normal state (i.e. pilot not commanding a manoeuvre), the rate gyro sends a signal directly to the computer; and also sends information through a leaky integrator that produces a 'pseudo' attitude by integrating the rate signal. The present attitude signal is compared to that which existed 20-30 seconds previously (hence the reason for the term 'leaky' - it does not have long-term memory). If the flight path is disturbed, for example by a gust of wind, the rate signal produced by the gyro is used to stop rate, and the leaky integrator produces an 'attitude' that is different from the attitude that existed 20-30s previously. Both error signals are sent to the computer, and this results in a correction signal being sent to the actuator. These signals stop the angular rate and then return the helicopter to the datum attitude. The 'leak' in the integration means that the pseudo attitude will disappear after a short time and the system will regard the continually updated new attitude as the datum. Simply stated, the system will try to maintain the datum to which the helicopter is trimmed, and it will attempt to return to this datum if it is disturbed. Such a system can overcome the problem of the nose dropping during a turn by feeding a derived bank angle signal to the pitch channel, and also by comparing the pitch attitude to the 'datum' pitch attitude.


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Due to the errors in integrating very small signals from the rate gyro, such as in a steady condition with no gusts, the system will drift off the desired condition. Such systems can only be regarded as providing a limited duration 'attitude' hold. The ability of an SAS to provide a long-term attitude hold will largely be determined by: 

the sensitivity of the rate gyro, and



the time constant of the integrator.

SAS Mode This mode is used during low and slow manoeuvring where the pilot may be making continual attitude changes such as preparing to land. The helicopters response is determined by the AFCS which will provide the pilot with a greater control to enhance the handling qualities of the aircraft. As this is a hands on control mode, the AFCS may be operated with the force trim on or off.

Stability Augmentation System (SAS) This system was developed to provide control inputs not only from a rate of change but also a deviation from the desired attitude and SAS is usually incorporated in the controls of pitch roll and yaw. The rate gyro is still the basic sensor, but the signals to the computer are also integrated to simulate an attitude signal. The SAS provides rate stabilisation, and may be referred to as a damper system, that is stabilising the helicopter against outside disturbances. If you look at the example of what happens in a gust of wind, as used previously, the rate signal produced by the rate gyro not only goes to the computer but also to an integrator, see Figure 4.2 The integrator will produce a pseudo attitude signal which the computer receives as an error signal and a correction signal will be sent to the actuator to return the helicopter to its original datum. The system will try to maintain the attitude that the helicopter was trimmed to, and it will try to return it to this attitude if it is disturbed. The SAS will provide short term stabilisation, and its efficiency will be determined by the sensitivity of the rate gyro and the time constant of the integrator.


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The sensors are electrically oriented so that their outputs are nulled when the cyclic stick is at its reference position corresponding to a level flight attitude. Displacements of the stick from this position produce output voltages from the sensors which are proportional to the amount of stick displacement, and are phase related to the direction of displacement. Both the rate damping and the pseudo attitude hold will interpret any movement of the flight controls by the pilot as a disturbance, and will try to return the aircraft to datum. Therefore, a means must be provided to allow the pilot to carry out a maneuver. The control stick is fitted with a stick position sensor, which, when the stick is moved, will allow a rate sensing switch to operate, disconnecting the rate gyro signal from the integrator. The rate gyro will continue to supply the computer with an error signal which means that rate damping will still occur, but there will be no attempt by the system to return the aircraft to its original attitude. When the pilot returns the stick to the detent position, the rate sensing switch will close, and the rate gyro signal will again be supplied to the integrator. In other words, the SAS will hold the helicopter in datum until the pilot moves the stick out of datum.


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Attitude Stabilisation A single pilot aircraft is required to exhibit longitudinal static stability and dynamic stability. If the aircraft fails to meet the minimum stability requirements, artificial stabilisation must be introduced in the form of an autopilot system. These systems can range from the simple pilot assist type to the very complex integrated systems. Helicopters are very unstable machines, and are extremely tiring to fly for long periods of time. Basic rate damping systems are fitted to overcome this problem, and these systems form the basis of all autopilot systems. It must be clearly understood that rate damping will not maintain a specific attitude. An example of rate damping, is to imagine a helicopter flying a straight and level course when it encounters a gust of wind causing it to roll to the left. The rate sensor will detect the movement after it occurs and send a signal to the computer, which will then command the actuator to move the pitch change mechanism. This will stop the roll rate, but the pilot will have to return the helicopter to its original attitude. Figure 4.1 shows a simple roll channel circuit. As can be seen, any deviation of the helicopter from its flight path, will be felt by the rate gyro, whose output is then computed, amplified, and then sent to the servo amplifier. The signal is then sent to the servo actuator to alter the pitch of the rotor to bring the helicopter back to the selected flight path. As the attitude of the aircraft changes, the aerodynamic feedback will be felt at the rate gyro whose output signal will reduce, thus reducing the signal to the computer. At the same time, the mechanical feedback from the servo actuator will be reducing the output from the servo amplifier.

Attitude Stabilisation (ATT) This forms the basis of the most complex autopilot systems and makes use of attitude gyros and parallel actuators. The system will still need a means of rapid damping which is provided by a series actuator and an SAS type circuit. Figure 4.3 shows a simple attitude hold circuit. By using a vertical gyro, an accurate attitude signal is available, which means that a commanded attitude will be held for a long time with great accuracy. The requirement to disconnect the system while manoeuvring still exists, and is done in much the same way as in an SAS system.


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TOPIC 13.3.7/8 - SYSTEM INTERFACE

VHF Omni Range (VOR) NDB navigation is not necessarily easy. First, there is a course to be flown. Second, the aircraft may be on-course or slightly (or hugely) off-course. Thirdly, the heading of the aircraft may be different from track, to accommodate a crosswind. During an instrument approach, a pilot has to continuously determine all this information, and also calculate and fly corrections. Determining which heading to hold to accommodate a crosswind is an empirical matter. One guesses a heading and determines drift (range of divergence of track from course) and then corrects - first twice as much, to get back onto course, and then when back on course, enough to follow course. All this is quite tricky and one needs to be in practice. This is crucial when flying an instrument approach, since strict adherence to course and altitude restrictions are the only things that guarantee that the aircraft flies clear of obstacles. Anybody who has flown an NDB instrument approach to an airport runway knows how labor-intensive it is. One has to achieve course-following using the above procedure, correcting for probably-changing crosswinds as one descends in altitude, especially in non-level terrain, very accurately and all inside of 2 or 3 minutes. ADF systems provide the pilot with an indication of the aircrafts relative bearing to that station. The only means the pilot has to determine his position using the ADF system is to plot the bearings of two different stations on a navigation chart and triangulate the aircrafts location at a point where the two lines intersect. This method of establishing current position is effective, but can be very cumbersome. This problem was addressed in 1945 with the introduction of VHF OmniRange (VOR) navigation stations. VOR is much more sophisticated and has definite advantages over ADF Issue B: January 2008

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navigation. VORs are radio beacons that transmit a signal which contains precise azimuth information, so that upon reception of the signal, an aircraft can tell precisely what bearing with respect to magnetic north the station is from the aircraft. ADF’s are still in wide use especially for aircraft flying to and from airports that are not equipped with VOR facilities. VOR has been the standard radio navigation system for cross country flying around the world for many years. The major advantages of the VOR system are: Provides an infinite number of radials or course indications Reduces the amount of indication errors from adverse atmospheric conditions Accurately provides directional information VOR operates in the VHF range 108 to 118 MHz. VOR reception is strictly line-of-sight. This limits the useable range at low altitudes or over mountainous terrain VOR is a VHF navigational aid for short and medium range flight distances, which is used mainly along airways and in the airport terminal control areas. A VOR station emits position lines, rather like a lighthouse, referring to magnetic north.

VOR Operation When the appropriate VOR frequency is entered into a navigation radio, the VOR indicator connected to that radio is used to find where the aircraft is relative to the VOR station. The RMI or HSI pointer will indicate bearing to or from the VOR ground station regardless of aircraft heading.


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The system provides flight crews with the ability to navigate accurately along a planned route using VOR transmitters as waypoints. In effect, it provides the ability to follow a desired path in the air.


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Introduction The VHF Omni-directional Radio Range, the abbreviations for which are 'VOR' and 'Omni', enables a pilot to determine the direction of his aircraft from any position to or from a VOR beacon, and, if necessary, track to or from the beacon on a selected bearing. VOR is a Very High Frequency (VHF) navigation aid. Because it is a VHF aid, its ground to air range is limited to 'line of sight' reception which is typical of VHF transmission. The range achieved is dependent, therefore, on the sitting of the VOR beacon with relation to surrounding terrain, and on the height at which the aircraft is flying. As a VHF navigation aid, the VOR is static-free, and the information given by it is displayed visually on easily read and interpreted cockpit instruments. An infinite number of bearings can be obtained and they may be visualised as radiating from the beacon like spokes from the hub of a wheel. However, for practical purposes, the number of bearings can be considered to be limited to 360, one degree apart, and these 360 bearings are known as radials.

Operational Use The VOR enables a pilot to select, identify, and locate a line of position from a particular VOR beacon. The following information can be obtained: 

the magnetic bearing of the aircraft from the VOR beacon;

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the magnetic bearing from the aircraft to the VOR beacon

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the position of the aircraft, i.e. port or starboard, of a selected radial

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when the aircraft is closing and when it is flying along a selected radial; and

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when the aircraft passes over the VOR beacon.


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Automatic VOR This term is give when the 30 Hz Reference signal is phase shifted automatically to measure the phase angle between the Variable 30 Hz (V) and the Reference 30 Hz (R) signal. The pilot need do no more than switch on and tune the receiver to an in range station in order to obtain a continuous indication of the bearing to the VOR station on the Radio Magnetic Indicator (RMI).


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Manual VOR This requires the pilot to switch on and tune the receiver to the station, and then select a particular radial on which he wants to position his aircraft. If the course selected in the course indicator does not correspond with the VOR bearing set by the Omni Bearing Selector (OBS), the deviation indicator will provide the appropriate left or right indications (as well as TOFROM indication).


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Cone of Confusion Over the VOR transmitter, there is a cone-shaped area called the Cone of Confusion (also known as; cone of silence or cone of ambiguity) in which the signals emanating from the VOR station have a spurious, unusable quality. In this area, the TO/FROM indicator is likely to fluctuate severely, accompanied by equally distorted presentations on the deviation indicator needle. At such a situation, the signals would be ignored until they had settled down again. The period of disturbance would depend for its length on the height at which the aircraft was passing over the station. Due to its cone-shaped property, its effect would be felt around an angle of about 40 degrees above the station. An aircraft flying over it at 1200 feet would traverse a linear distance of 800 ft within the cone, while an aircraft flying at 12,000 ft would suffer for a distance of 8,000 ft.


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VOR Principle of Operation Bearing to a lighthouse can be gauged by the following method. Time how long it takes for the light to complete a full revolution. This can be done by timing from when the beam is pointing directly at you, until it completes a full rotation and points back at you again. Lighthouses are often fitted with a red light on the top of them which flashes when the beam is pointing directly North. So if the beam hits you exactly when the red light flashes, you are directly North of the lighthouse. If a full rotation takes 10 seconds and the beam hits you 5 seconds after the red light flashes, you are directly south of the lighthouse. By knowing the time taken for a rotation and timing how long it takes from North for the vector to reach you, you can calculate your angular displacement from the lighthouse with respect to North. VOR functions by the same principle, with the reference signal replacing the Red light, and a varying “vector signal” providing sufficient information to the aircraft VOR system to calculate angular displacement from the VOR transmitter, with respect to North. The VOR station transmits two signals, one is constant in all directions, and the other varies the phase relative to the first signal. For example, at magnetic North both signals are in phase, but at magnetic south the variable lags the fixed signal by 180 degrees. The principle behind VOR systems then, is comparison of the fixed and reference signals. The VOR receiver senses the phase difference between the two frequencies and the difference identifies 360 different directions or "radials" from the VOR. The system does not provide distance information. The name Omni Range leads one to believe that Range to a station is provided. The inclusion of the word RANGE is an unfortunate choice of title, because a VOR transmitter provides only bearing to a station, not range.


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VOR Radiation Pattern Reference Phase Signal The reference phase signal includes an RF carrier and a 9960-Hz subcarrier. The subcarrier is frequency modulated between 10,440 Hz and 9480 Hz at a rate of 30 Hz. The reference phase signal is radiated from the centre antenna of a 5-element antenna array. This signal is omni-directional and has a constant phase throughout 360 degrees of azimuth. When voice communication or station identification is transmitted, the FM subcarrier of the reference phase signal is amplitude modulated. The identification signal (in Morse code) is a 2 or 3 letter word repeated three times in a 30 second period with a modulation tone of 1020 Hz (+ or - 50 Hz). The frequency range of the voice modulation is limited to 300-3000 Hz.


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Variable Phase Signal The variable phase signal is a pure RF carrier. This signal is radiated by the diagonal antennas of the 5-element array. The field strength pattern is a Figure of Eight configuration that is rotated through 360 degrees of azimuth at a rate of 1800 r/min (30r/s). The phase of the signal varies directly with azimuth; this is referred to as space modulation.

Composite Signal The reference phase signal and the variable phase signal are in phase when the positive lobe of the Figure eight pattern is aligned with magnetic north. However, as the positive lobe of the variable phase signal is rotated through 360 degrees, the phase angle between the reference phase signal and the variable phase signal increases. When the variable phase signal is aligned other than magnetic north, the positive voltage maxima of the variable phase signal occurs at a given time later than the positive voltage maxima of the reference signal. The phase displacement in the 2-voltage maxima at any point in azimuth designates the bearing at that point. The difference in phase can be measured and indicated by instrumentation in the aircraft. Therefore, a radial can be selected, and the aircraft's flight path can be related to that selected radial. The Figure of Eight pattern reinforces the omni-directional field strength pattern on the inphase side and weakens the field strength on the out-of-phase side. The resultant field strength pattern is cardioid revolving at a rate of 30 r/s. As a result of the space modulation of the rotating cardioid, an additional 30-Hz signal is apparent at the receiver input. The receiver circuit detects what appears to be a 30-Hz FM signal with a 30-Hz AM component. The phase of the two 30-Hz signals are compared, and the signal output is proportional to the phase relationship of the two 30 Hz signals. The phasing of the reference and variable signals is adjusted so they are in phase at magnetic north. The frequency of the 9960-Hz subcarrier is frequency modulated between 9480 and 10,440 Hz at a 30-Hz rate and is synchronised with the rotation, of the dipole antenna. When the frequency of the subcarrier is 10,440 Hz, the positive lobe of


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the dipole antenna is pointed toward magnetic north. The subcarrier frequency decreases to 9480 Hz as the dipole antenna rotates clockwise to 180 degrees (south). At 90 and 270 degrees from north (0 degree), the subcarrier frequency is 9960Hz. A maximum positive audio modulating voltage produces 10,440 Hz, and a maximum negative voltage produces 9480 Hz. Thus the output of at FM detector will be positive when the frequency is 10,440 Hz and negative when the frequency is 9480 Hz. The frequency of this voltage will be equal to the rate the subcarrier frequency is varying, which is 30 Hz.


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The Horizontal Situation Indicator One system for VOR presentation is the HSI. This type of display is for manual VOR indication. The HSI allows us to select, using the course select knob to select the desired radial to fly along. The course select may be part of the indicator or mounted remotely from the indicator on a mode control panel. The calibration of the HSI will correspond to the magnetic bearings relative to magnetic north. The display has a needle indicating the selected course on a rotating remote magnetic compass card. The left-right needle indicates the aircraft's position in relation to the radial selected with the course select knob. If the left-right needle is centered the aircraft is on the radial. If the needle is off centre, the aircraft is off the selected radial. The to-from indicator shows which side of the VOR station the aircraft is positioned. With the left-right needle centred, the aircraft can be anywhere on the line drawn through the VOR station on the radial selected. The information from the to-from indicator also allows the pilot to determine which direction to fly the aircraft. When the to-from indicator shows "TO" and the aircraft is flown on the course indicated by the HSI, it will be tracking towards the station. Conversely, when the indicator shows "FROM" and the aircraft is flown on the same course as indicated by the HSI as above, it will track away from the VOR station. There is another very important function of the left-right needle. Not only does this needle indicate when the aircraft is off the selected radial, but it indicates how far off that radial. This needle, by the direction of its deflection, also tells us in which direction to fly to get back on Issue B: January 2008

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the selected radial. If the aircraft is moving towards the VOR and the to-from indicator indicates "TO", the aircraft is flown towards the needle to get back on course. That is, assuming that the HSI is set to the desired radial and that the left-right needle is centred and now deflects to the left, the selected radial is to the left, and the aircraft must be turned towards the left to get back on the radial. The amount we are off course depends on the amount of needle deflection, and on the calibration of the needle. Typically, a full scale deflection indicates being 10 degrees off the selected course. A full scale deflection, is from centre to one side. If the needle is half way between centre and one side it will indicate 5 degrees. The importance of the course width should be clear. A very sensitive needle would move too much for an insignificant off course amount; an insensitive needle would not give a warning to take corrective action until the aircraft was way off course. Suppose the aircraft was 30 miles from the station, which is a perfectly reasonable and usable distance for the VOR system, and is 1 degree off course. The aircraft would be half a mile off course (see below). At 30 miles, 3 degrees off course would be approx. 1.5 miles. (This approximation applies up to a few degrees). At 60 miles 1 ° off course at 60 miles the off course distance approximates to 3 miles.

Automatic Omni Indicator and the Radio Magnetic Indicator (RMI) The RMI is more than an omni presentation. The first part of the RMI is an automatic means for indicating the omni bearing. There is no to-from indicator, no omni bearing selector, and no left-right needle. The needle on the automatic omni simply points to the "TO" the bearing to the station.


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Instrument Landing System (ILS) Prior to World War II, aircraft approaches and landings in poor weather conditions were often high risk endeavors. Early civilian air navigation systems provided only basic lateral position information, and 1930s airline pilots flying in low visibility conditions had to take a heading off a radio navigation station and then use speed, time and distance calculations to figure out when they should see the runway at their destination. Safe descents were dependent on the accuracy of their calculations and all final approaches had to be visual, because there was no instrument guidance system that could direct a pilot on a safe, precise descent profile clear of terrain. An instrument landing system (ILS) provides pilots with commands which enable a nonvisual approach to a runway. The obvious advantage of such a system is that pilots can descend even in the worst weather conditions. In an aircraft equipped with ILS, outside visibility is not crucial until moments before touchdown. This lesson will the following aspects of an ILS: •

ground based radio transmitting equipment

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airborne receiving equipment

•

instruments pilots use to navigate along the required flight path Although the functions of ILS can be replaced by other systems (such as microwave landing systems), it is likely that ILS will remain as the predominant world wide landing aid for the next few decades.


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The Instrument Landing System is in every sense a precision approach system. With modern equipment it can guide an aircraft right down to the runway—zero Decision-Height and zero visibility. An ILS system provides adds glide-slope, or elevation information, localiser information and distance to go to the runway threshold.

The ILS Components In order to guide an aircraft from several kilometers out, to a runway threshold, a system providing lateral and vertical guidance is used. The lateral guidance portion is called the localiser and the vertical guidance part is called the glideslope. In addition to the localiser and glideslope signals there are marker transmitters which provide pilots with an indication of their approximate position at intervals along the approach path. This total system using ground based and airborne equipment is called the instrument landing system (ILS). An aircraft flying ILS follows two signals: a localiser for lateral guidance (VHF); and a glide slope for vertical guidance (UHF). When the ILS receiver is tuned to a localiser frequency a second receiver, the glide-slope receiver, is automatically tuned to its proper frequency. The pairing is automatic. There's more to an ILS than the localiser & glide slope signals. FAA categorises the components this way: Guidance information: the localizer and glide slope. Range information: the outer marker (OM) and the middle marker (MM) beacons. Visual information: approach lights, touchdown and centerline lights, runway lights. The indicator for an ILS system uses a horizontal needle and a vertical needle. When the appropriate ILS frequency is entered, the horizontal needle indicates where the aircraft is in relation to the glide slope. If the needle is above the center mark on the dial, the aircraft is below the glide slope. If the needle is below the center mark on the dial, the aircraft is above the glide slope.


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Descriptions of the ILS Components An aircraft on an instrument landing approach has a cockpit with computerized instrument landing equipment that receives and interprets signals being from strategically placed stations on the ground near the runway. This system includes a "Localiser" beam with a width from 3° to 6°. It also uses a second beam called a "glide slope" beam that gives vertical information to the pilot. The glide slope is usually 3° wide with a height of 1.4°. Three marker beacons (outer, middle and inner) are located in front of the landing runway and indicate their distances from the runway threshold. The Outer Marker (OM) is about 8 km from the runway. The Middle Marker (MM) is located about 1 km from the landing threshold, and the Inner Marker (IM) is located between the middle marker and the runway threshold where the landing aircraft would be 100 feet above the runway (about 150 meters). Because of Australia’s good weather conditions relative to the rest of the world, most airports do not have an inner marker installation. The theory here is that even under really bad weather conditions, pilots will usually have the runway in sight by the time they reach the inner marker point.


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An ILS is a fixed installation useful for only one runway, accordingly there may be more than one ILS system at an airport (ie one for each runway). The ground-based equipment consists of: • a localiser (LOC) transmitter and antenna located at the far end of the runway • a glideslope (GS) transmitter and antenna located near the runway threshold • an outer marker antenna located approximately 8 kilometres from the runway threshold • a middle marker antenna located approximately 1 kilometre from the runway threshold • an inner marker beacon antenna located approximately 150 metres from the runway threshold.

Glideslope The Glideslope is the signal that provides vertical guidance to the aircraft during the ILS approach. The standard glide-slope path is 3° downhill to the end of the runway. The glideslope antenna is capable of operating between 328 and 335 MHz, (UHF) with the lowest assigned frequency at 329.15 MHz. The glideslope signal is radiated to produce two intersecting lobes, one above the other. The upper lobe is modulated by a 90 Hz signal, the lower lobe by a 150 Hz signal. When the aircraft is on the centre line, the two audio signals are equal. This occurs at approximately 3° above the horizontal. This line of equal modulation defines the glideslope approach path. If the aircraft is too high the 90 Hz signal will predominate and if it is too low the 150 Hz signal will predominate. How these signals effect the aircraft instrumentation will be explained later in this lesson. The glide path projection angle is about 3° above horizontal and it intersects the inner marker at 100ft, the MM at about 200 feet and the OM at about 1,400 feet above the runway elevation. The glide slope is normally usable to a distance of 10 NM.


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Localiser The localiser signal provides azimuth, or lateral, information to guide the aircraft to the centerline of the runway. It is similar to a VOR signal except that it provides radial information for only a single course; the runway heading. The localiser operates between 108 and 112 MHz, with the lowest assigned frequency at 108.1 MHz. The localiser signal is similar to the glideslope except that the 90 and 150 Hz signal radiated lobes are side by side and directed along the centre of the extended line of the runway. The line of equal modulation defines the centre line of the runway approach path. If the aircraft is left of the centreline the 90 Hz signal will predominate. If it is right of the centreline the 150 Hz signal will predominate. How sensitive is the Localiser? Near the Outer Marker, a one-dot deviation puts you about 500 ft. from the centerline. Near the Middle Marker, one dot means you're off course by 150 ft. The localiser signal also carries the audio identification.

Specifics of the Localiser The localiser antenna is located at the far end of the runway. The localiser signal is normally usable 18 NM from the field. The Morse code Identification of the localiser consists of a three-letter identifier preceded by the letter I. Localiser and glideslope frequencies are paired so that selecting the localiser frequency automatically selects the glideslope frequency.


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This table shows the pairing between localiser and glideslope frequencies. When the localiser frequency is selected on the VOR/ILS controller, the paired glideslope frequency is automatically selected.

Audio Audio, which consists of the station ident (at 1020) Hz and voice information from the station (between 300-3000 Hz), is supplied to the aircraft audio integration system. Filters in the audio integration system can be selected to filter out neither, either or both signals to be heard by the pilot. 100 mW nominal output at 30 percent modulation.


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Flag Outputs Two high-level warning signals (super flag) and one low-level warning signal should be provided by both localizer and glidepath receivers. The high-level flag characteristic is 28 VDC for valid status with current capabilities; 25 mA for AFCS warning; 250 mA for instrument warnings. The low-level flag should provide a voltage of between 300 and 900 mV into up to five parallel 1000 Ω loads.

Monitoring Warning signals when: no RF, either 90 or 150 Hz missing, total depth of modulation of composite 90/150 Hz signal is less than 28 per cent, etc.

Deviation Outputs Localizer: high-level 2 V for 0.155 d.d.m, low level 150 mV for 0155 d.d.m. Dual outputs in parallel for AFCS. Output characteristics should not vary for loads between 200 Ω and no load. When 90 Hz predominates the `hot' side of all deviation outputs should be positive with respect to the `common' side; in this case 'fly-left' is given. Glideslope: similar to localizer but high and low-level outputs are 2 V and 150 mV respectively for 0175 d.d.m.


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Autopilot Localises and glideslope signals are also fed to the autopilot system, if one is fitted. The localiser deviation will be used to supply the appropriate demand signal to the roll (aileron) and yaw (rudder) channels. The pitch (elevator) channel will respond to glideslope. The signals supplied to the autopilot is the same low voltage signals supplied to the instrument. Dual outputs in parallel for AFCS.

Figure above shows what would happen to an airplane, controlled by an autopilot, while flying a VOR or localizer course if a crosswind were to come up, and no crosswind correction capability were included in the autopilot computer. (Autopilots always have a crosswind correction capability). In position “A”, the airplane is on the radio beam and on the selected heading (deviation needle centered and course select cursor at the lubber line). In position “B”, the airplane has been moved off the radio beam by the crosswind (dotted pattern). The heading has not yet changed. In position “C”, the autopilot has begun to change heading back toward the beam, In position “D”, the autopilot has achieved a stable “standoff’ condition (defined later). Under airplane “A” is an HSI showing the pilot’s indication of the fact that the airplane is on the chosen heading of 900, and centered on the radio beam.


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Under airplane “B” the HSI shows the airplane heading has not changed yet, but the aircraft has been moved off the beam center. The two controlling signals into the autopilot roll channel are radio deviation and course select error. If the radio beam is to the left of the airplane, the deviation signal in the roll computer calls for a left turn, The vector chart below the HSI (airplane “B”) shows that the deviation signal in the roll computer is calling for a left turn, and there is no course select error signal. Under airplane “C” the HSI shows that a course select error of 5° has been developed as a result of the autopilot turning the airplane toward the beam center. The radio deviation signal has increased to one dot. The chart below the HSI shows that the radio deviation signal, calling for left turn, still exceeds the course select error signal, calling for a right turn. Under airplane “D”, where steady “standoff’ conditions have been achieved with a steady crosswind, the two controlling signals are “standing each other off’. The HSI shows a course select error of 10°. The deviation needle shows a deviation error of 11/2 dots. The chart below the HSI shows the deviation signal, calling for a left turn, balanced by the course select error signal, calling for a right turn. Consequently, the roll channel does not have an output to the ailerons, and the airplane wings are level. This situation is described by saying that heading error is “standing off’ radio deviation. The heading error signal (course select error) is “required” by the crosswind. In order for the airplane to follow the radio course, it must head into the wind; its heading must be different from its direction of travel, or ground path. This condition could only develop in an autopilot without crosswind correction (typically accomplished with an integrator circuit). Understanding what would happen without the integrator correction for crosswind will make it easier to understand the function of an integrator circuit in providing crosswind correction.


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Crosswind Correction by Integrator Figure above illustrates how an integrator would function to correct for the crosswind standoff condition developed in the previous figure. Assume that an integrator is switched into the roll computer after the stabilized standoff condition was built up. The time of switching in the integrator is illustrated in the figure with airplane “A”. The vector chart below the airplane shows the deviation signal calling for left turn balanced by the course select error signal calling for right turn. The input to the integrator is the course select error signal. Over a period of time, usually on the order of 100 seconds, the output of the integrator will build up to equal its input. The sense (phase or polarity) of the input will be inverted by the integrator. In this case, the heading error input calling for a right turn will be inverted by the integrator whose output will call for a left turn. Airplane “B” shows that the integrator output has begun to build up. Since the integrator output is now cancelling part of the required heading error signal, a smaller deviation signal cancels the remainder for a stable condition. Therefore, the airplane has moved closer to the beam center. Airplane “C” shows a later time when the integrator output has further built up, cancelling more of the required heading error signal, leaving less deviation necessary for a stable condition. Airplane “D” shows a later time when the integrator output has been built up even more, leaving only a small amount of deviation necessary to cancel the remainder of the required heading error.


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Airplane “E” shows that the output of the integrator now equals its input, and therefore complete cancellation of the required heading error is being accomplished with the integrator output; no deviation error is needed to achieve a stable condition. The airplane is now following the radio beam center, with the required heading error cancelled by the output of the integrator. As long as airplane speed and wind conditions remain constant, this situation will prevail. In normal autopilot, the integrator action would begin as soon as any heading error develops, and the airplane would not have been moved as far off course as we have indicated. The crosswind correction function could have been accomplished with an amplifier, noninverting integrator, with the input signal being the deviation signal. In that case, some small amount of residual deviation (determined by amplification factor) would be required to hold the necessary output. Integrators have other uses in autopilot, and their operation can often be observed in a parked airplane when a mechanic is checking the autopilot. If the mechanic does not understand what the autopilot is doing, he may think something is wrong. For example, the control wheel may mysteriously creep all the way to the right or left, or the control column may mysteriously creep forward or aft.


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VOR/LOC The operation of the roll channel is the same for both the VOR and LOC modes of operation. Here we will discuss the roll channel operation of the VOR mode only. In the VOR mode of operation the pilot selects the VOR frequency. The system works on the same principle as the heading mode except that the reference signal is now received from a radio navigation aid instead of the aircraft directional gyro. When the aircraft detects the signals from the radio navigation beacon, these signals are sent to the roll channel, so the aircraft can start to turn onto the beam, as shown in Figure.

Block Diagram Operation Figure is a simplified diagram of the autopilot roll channel for the VOR/LOC mode of operation. The radio navigation signal (Nav Deviation) is fed to the roll channel where it is combined with the aircraft course select error signal. The resultant signal is sent to the autopilot bank angle limiter and roll rate limiter. This limits the amount and rate of roll permitted from the autopilot. The signal is combined with the roll attitude signal and sent to the aileron servo actuator to turn the aircraft. Once the aircraft has "captured" the beam, it will continue to follow this heading until it reaches a pre-determined position over the VOR beacon. Once over the beacon, the aircraft could start detecting the convergence of other VOR radials which may confuse the aircraft flight control system. To prevent this, an over station sensor circuit is incorporated which cuts off the roll channel for a pre-determined period.


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Once this period has elapsed, the aircraft then detects the VOR signal corresponding to the outbound radial. This is switched in, allowing the aircraft to continue on that radial.


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Glideslope The glideslope mode of operation is similar to the VOR/LOC mode of operation of the roll channel in that it requires the aircraft to capture a radio navigation aid beam for the aircraft to follow. Instead of rolling the aircraft, the pitch attitude is adjusted and the aircraft follows the glideslope beam to the airfield to land the aircraft. There are two different operations for the pitch computer in this mode; 

Capture



Flare

Glideslope Capture Figure shows the method of capturing a glideslope beam from above and below. The mode engaged prior to capture of the glideslope will be disengaged when the glideslope capture occurs. The sensing circuit in the pitch computer which initiates glideslope capture mode is called the "Vertical Beam Sensor" (VBS). The autopilot follows the glideslope using a preset descent rate for example some passenger airliners use 700 feet per minute.


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Block Diagram Operation Figure is a simplified block diagram of the pitch channel in glideslope mode. The glideslope receiver generates an error signal which represents the position of the aircraft relative to the centre of the glideslope beam. This signal is sent to the pitch computer. The pitch computer then modifies the error signal and sends it to the control valve of the elevator servo-actuator. The servo-actuator then drives the elevators to the commanded position. When the aircraft is positioned in the centre of the glideslope beam, the error signal will be nulled, the elevators will return to the streamlined position and the aircraft will fly down the glideslope beam.

Glideslope Flare "Flare" is the term used to describe the touchdown manoeuvre. A signal from the RADAR altimeter system is used to initiate the flare operation of the pitch computer. This signal is normally initiated at 50 feet.


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FLIGHT MANAGEMENT SYSTEM INTERFACE A typical FMS has the capability of automatically controlling the aircraft from just after takeoff through to roll out on the runway after landing at the destination. The pilot must take-over to turn off the runway and taxi to the gate. Not all flights use a flight management system to its fullest capacity, but the autopilot and flight director will be used for some portion of each flight under normal circumstances


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Flight Management Computers In the 1950s and 1960s, aviation electronics, referred to as avionics, were simple standalone systems. The navigation, communications, flight controls, and displays consisted of analog systems. Often these systems were composed of multiple boxes, or subsystems, connected to form a single system. Various boxes within a system were connected with point-to-point (analogue) wiring. The signals mainly consisted of analog voltages, synchroresolver signals, and switch contacts. The location of these boxes within the aircraft was a function of operator need, available space, and the aircraft weight and balance constraints. As more and more systems were added, the cockpits became more crowded, the wiring more complex, and the overall weight of the aircraft increased. By the late 1960s and early 1970s, it became necessary to share information between the various systems to reduce the number of black boxes required by each system. A single sensor, for example that provided heading and rate information, could provide that data to the navigation system, the weapons system, the flight control system, and pilots display system. However, the avionics technology was still basically analog, and while sharing sensors did produce a reduction in the overall number of black boxes, the connecting signals became a “rat's nest” of wires and connectors. Moreover, functions or systems that were added later became an integration nightmare, as additional connections of a particular signal could have potential system impacts. Additionally, as the system used point-to-point wiring, the system that was the source of the signal typically had to be modified to provide the additional hardware to output to the newly added subsystem (additional amplifiers, or output multipliers boxes). As such, inter-system connections had to be kept to the bare minimum.


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By the late 1970s, with the advent of digital technology, digital computers had made their way into avionics systems and subsystems. As time and technology progressed, the avionics systems became more digital. And with the advent of the microprocessor, things really took off. The Flight Management Computer typically performs as the data management computer and is responsible for directing the flow of data on the data bus. FMC is master computer which integrates functions of: 

Navigation Computer

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Flight Control Computer

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Thrust Management Computer

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Air Data Computer

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Maintenance (EICAS or ECAM) Computer


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Flight Management System The Flight Management System provides automatic navigation, guidance, flight planning, navigation map display, flight information, and flight performance optimization. The system makes use of information from several external aircraft systems coupled with algorithms programmed in the FMC. The outputs from the FMS are fed to the main computer where this information can be used in a variety of ways including feeding into the aircraft autopilot for automatic guidance.


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FMS Major Components The major components of a FMS are shown in the block diagram above. The major systems include the Flight Management Computer (FMC), the Flight Management System Computer, two FMS Control Display Units (CDU) and the I/O Concentrator. The FMS Chassis and I/O Concentrator are standalone computers containing PC processors running operating systems.

Flight Management Computers Provide a number of advanced features and functions which were not found on earlier autopilot systems. Some of the functions of the FMC’s are: Flight Planning – the entire flight can be programmed into the computer using a cockpit keyboard, cruise altitudes, waypoints, overfly points, ascent and descent points, etc. Performance Management – the system can provide optimum profiles for climb, cruise descent and holding patterns. A minimum cost flight can be flown automatically by using optimum climb settings, cruise settings, etc. Navigation Calculations – the FMC can calculate great circle routes, climb and descents profiles, etc. Auto tune of VOR & DME – FMC can automatically tune radios to the correct and appropriate station frequencies for navigation aids as well as tower approach frequencies. Autothrottle Speed Commands – these are displayed on the EADI as fast/slow indications. The FMC is in effect the master computer which integrates the functions of the inertial navigation unit (or laser ring gyro sensors), flight control computers, thrust management computers, air data computers, navigation sensors and EICAS computers. The Flight Management System (FMS), in conjunction with other interfacing equipment in the aircraft, forms an integrated, full-flight regime control and information system which provides Issue B: January 2008

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automatic navigation, guidance, map display, and in-flight performance optimization. It reduces cockpit workload by eliminating many routine tasks and computations normally performed by the flight crew. The system operates continuously at all times if properly initialised.


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Flight Management System (FMS) Operation The primary function of the FMS, when coupled to the flight guidance system (autopilot), the auto-throttle and the EFIS systems, is to guide the aircraft in both the lateral (navigation) and the vertical (performance) axis planes. A total of up to some 24 inputs can be interfaced with the FMS. The FMS system has several major functions which it carries out simultaneously to maintain control of the aircraft. •

Data Verification Function. The purpose of this function is to receive and transmit digital data to check the validity of data from the aircrafts systems which having inputs to the FMC.

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Navigation Function. This function uses the navigational database in the FMC and the output from the Inertial Reference Units (IRU’s) for computing the aircrafts exact position, velocity and other performance parameters.

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Guidance Function. This function uses the output data from the navigation function to guide the aircraft against a preset flight plan that has been entered into the FMS by the flight crew. The flight plan is usually selected by the crew before departure, but it can be changed during flight if required.

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Vertical Profile Function. Using performance data and the data output form the navigation function, the FMC compares actual to selected altitude, altitude rate and then through the Flight Control Computer (FCC) and the auto-throttle, adjusts the pitch of the aircraft so as to maintain the preselected vertical position of the aircraft.

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EFIS Function. The FMC provides data to the EFIS system for the purpose of display. The EFIS is used to display such data as radio navigation aids, aircraft way points and some system failure messages. The outputs from all of these functions are then used to guide the aircraft to a predetermined flight in accordance with the operating parameters of the aircraft and the operating airline of the aircraft.


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FMS Interfaces In a normal aircraft installation, the FMC will directly connect to several basic aircraft systems such as Air Data Computers, Inertial Reference Systems, Flight Control Computers and other systems via multiple ARINC 429 serial data bus interfaces. The FMS Computer provides all of the FMS sensor inputs.

Discrete Input There are as many as 128 discrete input discrete channels available. The Digital Input Board is designed to read a voltage from a variety of devices. The signals may originate from electronic switching circuits, standard logic circuits, mechanical switch contacts, relay contacts, or numerous other sources. The inputs can be configured for current sinking or voltage sourcing input signals.

Discrete Output One card is used to provide up to 128 output discrete channels. The board is designed for a variety of applications such as relay drivers, lamp drivers, solenoid drivers, stepper motor drivers, LED drivers, and fiber-optic drivers.

ARINC 429 Interface Two boards handle ARINC 429 serial I/O. The ARINC 429 module is an intelligent I/O device with its own on-board processor and memory buffer. Each board can accommodate 16 receive channels and 8 non-multiplexed transmit channels. The transmit and receive channels can be programmed as high speed (100 KBs) or low speed (12.5 KBs) in channel pairs. The vast majority of the data that flows in and out of the FMC goes through the ARINC 429 channels. Issue B: January 2008

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Audio Output Another addition can be a programmable Audio Alert card. This card allows the triggering of programmed voice messages in response to logical events determined in the applications software. Different messages may be recorded and loaded into non-volatile memory using an external PC. The voice messages are fed into the aircraft inter-phone system for distribution throughout the airplane.

CPU Transition A PC Transition card takes the last VME card slot. This card works in conjunction with and is connected to the CPU card. It provides the interface between standard Ethernet, parallel port, and the two serial port connectors and the host board connector. The Ethernet and parallel ports contain passive circuitry only. The I/O controllers reside in the host CPU board. The serial ports contain active circuitry to provide multiplexing and buffering functions.


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Flight Management System To complete its functions, the FMC has several data bases. These databases include: •

Navigational Database – This database contains information on navigational radio aids, navigational way points, airline flight plans and flight routes.

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Performance Database – This database stores the engine installation data, aircraft aerodynamic performance and atmospheric models.

Database The Navigation Data Base contains all of the navaid, airport, airway, approach, and departure information required to operate the aircraft in the national airspace system. The FMC has 1 MB of nonvolatile bubble memory reserved for this purpose. The ARINC 424 compatible database is updated on a 28- day cycle and the memory is large enough to hold two update cycles. The non-volatile memory also holds the FMC operations program, performance data for the aircraft model and engine configuration, and specific customer option data. Data is loaded into the FMC from floppy disks using a standard ARINC 615 compatible data loader, which can be connected via a loader connector on the rear of the FMS.


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Navigation The navigation function of the FMC generates aircraft position, velocity, and altitude data for use in the FMC and other external aircraft systems. Navigation algorithms combine data from the Navigation Data Base, position and velocity data from the three IRS systems, range and bearing data from the DME and VOR receivers, and altitude and airspeed data from the Air Data Computers. The combined data is used to generate geodetic aircraft position, velocity, wind vector, altitude and flight path angle, local earth radius, and other data. The FMC has algorithms for the automatic navigation-aid selection based on aircraft position, desired track, and data from the Navigation Data Base and provides tuning outputs for the DME/VOR receivers. The FMC can make use of high accuracy GPS position data.

Navigation Accuracy: The FMS automatically selects and tunes VHR Omni- Range (VOR) and Distance Measuring Equipment (DME) in order to constantly update the position and speed of the aircraft. This information is used in conjunction with the Inertial Reference System (IRS) and Global Positioning System (GPS) to ensure accuracy in all phases of flight. Issue B: January 2008

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The FMS will primarily attempt to utilise GPS position information, then combine range information corrected for slant range from two separate DME locations, and finally position from three Inertial Reference Units (IRUs). If no usable VOR/DME information is available, the FMS will monitor aircraft position based on IRS/GPS data only, until the aircraft is determined to be in a location where DME/VOR information is once again available for position and velocity cross checking. The FMS navigation management system will also compute and provide true and magnetic track information, drift angle, magnetic variation for the current aircraft location and vertical flight path information. The FMC automatically determines which VOR/DME combinations will yield the best result given their position relative to the aircraft.

Flight Planning The FMC allows for the input of complex 2-D and 3-D flight plans via either of the two CDU's. The pilot or operator can select from thousands of airports, runways, waypoints, navaids, airways, approach and departure procedures defined in the Navigation Data Base to build a flight plan. Additionally, user defined waypoints can be entered based on latitude and longitude or distance and bearing from a pre-defined waypoint. Once entered, flight plans may be modified by adding, changing, or deleting any of its elements. The flight plans can be further modified with altitude and speed constraints, holding patterns, path offsets, etc. Two complete flight plans can be separately stored in the FMC and activated via CDU commands.

Guidance The guidance function defines the two or three dimensional flight path to be flown and the necessary computations and outputs for aircraft control referenced to the defined flight path. The flight path may be defined in the lateral and/or vertical planes from flight plan and performance data entered into the FMC from the CDU and augmented with navigation and performance databases. The lateral flight plan is normally defined by a string of navigation waypoints. The lateral path between waypoints may be any of 16 navigation leg types required to fly all published ATC procedures. The vertical and thrust guidance functions are closely integrated with performance management functions for optimum three-dimensional path guidance. The FMC computed guidance output parameters include desired track, track angle error, bearing and distance to the active waypoint, lateral and vertical path deviation, and speed error. These guidance outputs can be fed into the aircraft autopilot via the I/O Concentrator and Pilot Select Panel (PSP).


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Guidance Management: Two dimensional flight path management is available along an FMC programmed flight path in either the vertical navigation mode (VNAV) or lateral navigation mode (LNAV). Both of these modes are selected by using the LNAV/VNAV buttons on the Autopilot Mode Control Panel (MCP). When used in conjunction with one another, the FMS is capable of providing fully integrated three dimensional flight path management along the FMC defined flight path. The LNAV guidance function issues steering commands to the AFDS in order to keep the aircraft navigating correctly along the programmed route of flight. Deviations from the center of the desired flight track are corrected using intercept procedures and flight track adjustments. Normal lateral flight path deviation should not exceed 0.1nm in most phases of flight. In all phases of an LNAV managed flight, the FMS will monitor cross track error, which is defined as the lateral distance separating the aircraft from it’s desired path of flight. Roll and steering commands are provided to the AFDS Flight Control Computers in order to correct the cross track error. The FMS is capable of providing a great circle Direct-To track to any point programmed into the FMC/CDU displayed flight path. The VNAV guidance function controls the aircraft along the vertical flight path as defined by the FMC/CDU entered flight path and the aircraft’s performance limitations. VNAV takes position data from the navigation system and compares it to the vertical profile as defined in the FMC/CDU entered flight plan. The vertical navigation function then provides pitch and thrust commands to the AFDS in order to intercept and maintain the defined vertical profile for the current phase of flight. For vertical performance modes where vertical speed is unconstrained (most climbs) the VNAV system will provide pitch and thrust commands to the AFDS so as to maintain the most efficient climb based on the current thrust mode selected. This results in the most economically beneficial climb gradient, not necessarily the most rapid climb gradient. VNAV uses essentially two basic pitch control modes to manage the vertical flight profile: speed or rate of climb/descent. When speed the controlled factor the AFDS autothrottle will be given a target thrust setting by the vertical navigation function, and the elevator will be used to control speed, resulting in a variable rate of climb or descent based upon conditions. When vertical speed is the controlled factor, the AFDS will issue commands to the elevator for vertical speed control, and the AFDS will adjust the autothrottle to maintain speed, resulting in a fixed rate of climb/descent and variable speed based upon conditions.


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The lateral guidance function provides guidance between waypoints of the active flight plan route. The crew receives this guidance via the cockpit displays and can elect to couple the FMS to the autopilot if the aircraft is so equipped. The vertical guidance function provides control of the vehicle along the vertical route of the flight plan. Vertical guidance generates the pitch, speed, altitude, and thrust targets that are again displayed on the cockpit displays and can be coupled to the autopilot and auto-throttle equipment if the aircraft is so equipped. The time navigation function provides the calculated times of arrival at various waypoints along the flight plan. The crew can select a required time of arrival (RTA) at a waypoint and the FMS will generate the appropriate commands to achieve that time on station. The crew can specify the performance mode (RTA, fuel economy, maximum speed) and the FMS provides the proper commands to achieve that performance. The estimated times of arrival are displayed to the crew via the cockpit displays. Performance targets are translated into flight director cues and can be coupled with the autopilot/auto-throttle systems if so equipped.


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In target reference Vertical Navigation (VNAV) mechanizations, the FMS utilizes FCS vertical modes for Flight Level Change, Altitude Hold, Vertical Speed, Glide Path, or Pitch. The FMS is responsible for managing the vertical modes and providing target references to the FCS. Glide Path modes often require the FMS to supply a pitch command and Altitude Hold mode requires the FMF to supply an altitude reference. The remaining vertical modes, Flight Level Change and Vertical Speed, use the speed reference supplied by the FMS. To integrate the FMS with a Flight Control System (FCS), the data required to steer the aircraft to the desired flight path could be based on a command reference, target reference or a combination of command and target reference data. Typically, lateral guidance from an FMS is provided in terms of a roll command. Vertical guidance often uses a combination of pitch command and/or various target references, such as airspeed/mach, altitude, or vertical speed, for vertical guidance.

Performance The performance function manages aircraft flight profile by computing display data that assists the pilot in making decisions and commands that cause the aircraft to be controlled according to the selected flight plan. Algorithms in the FMC-PIP compute speed, thrust setting, and vertical guidance commands to meet the selected mode objectives, but are subject to aircraft performance limits and flight plan constraints. Prediction data is computed such as distance, time of arrival, altitude, speed, and fuel at future points on the active flight plan. Other performance calculations include warning messages if future flight plan constraints cannot be met, optimum and maximum cruise altitudes, time and fuel limits, climb and descent limits, and engine out drift down. Performance Management: The FMS is capable of managing nearly all aspects of aircraft performance so as to optimize precision and economy of flight. The FMS is only capable of providing such information if the gross weight, cost index target altitude and a route have been entered into the FMC/CDU by the crew.. Vertical Navigation can only be accomplished if the performance initialization page is complete. The performance model uses input from fuel flow, engine data, altitude, gross weight of the aircraft, flaps, airspeed, Mach, temperature, vertical speed, acceleration and location within a programmed flight plan to determine the optimum performance for the aircraft at any given moment. The performance management modeling used by the FMS attempts to provide a least cost performance solution for all phases of flight, including climb, cruise and descent. The default cruise performance management setting is ECON, or economy cruise. The airplane and engine data models are used to provide an optimum vertical profile for the selected performance mode, even if ECON has been overridden by the crew. During the climb, an optimum Mach speed target and a corresponding thrust target are computed by the FMS, with the speed target transmitted to the vertical guidance function of the autoflight director system. The AFDS will then generate commands to the elevator in order to maintain the correct pitch for the required speed. Thrust setting commands are delivered to the autothrottle servos by the FMS, and used in conjunction with the pitch setting commands to maintain the optimum speed and climb as directed by the FMS. During cruise, an optimum Mach setting is computed and thrust setting commands are delivered to the autothrottle. During descent, a vertical path is computed based on the flight plan entered into the FMC/CDU. The FMS will evaluate expected wind conditions, aircraft speed, altitude, position relative to the planned end-of descent point and any intermediate altitude or speed constraints between the aircraft and the end-of-descent point. This information will be passed to the AFDS for pitch based speed and vertical speed control and the autothrottles for vertical speed and thrust management. In ideal conditions, an idle thrust optimum descent profile is flown, however in many cases thrust and pitch will be varied to account for wind conditions or to ensure proper tracking of the vertical descent profile.


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Control Display Units (CDU’s) The Control Display Units (CDU’s) are mounted on the centre console in the cockpit. There are usually two fitted to an aircraft, there being one CDU fitted per FMC fitted to the aircraft. They provide control of the FMC’s and allow access to FMC fault data. System tests are started from the CDU’s, and test results are displayed on the CDU screens. Printing of the CDU display can also be carried out by using the on board printer in the aircraft centre pedestal. The CDU has two distinct areas on its display. •

The first area is the display area. This part of the CDU is used to display information from the FMC and also display information which has been entered into the FMC. Located next to each line of information on the CDU display is a touch key which is used to select the information located next to (if the option is available).

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The second area is the key pad area of the CDU. This area is used to enter in information into the CDU, change CDU displays and change FMC programming. This area has a full alpha numeric touch pad for data entry

Control Display Units (CDU’s) The CDU can also be used for the accessing of maintenance data and for the testing of aircraft systems. This information can be stored either in the FMC or in the Central Maintenance Computer (CMC). The maintenance pages of the CDU show maintenance related data that is accessible on the ground only. These pages are only available when the aircraft is on the ground. When the relevant system is accessed through the CDU, a warning message is usually displayed as well as a test message. This is a result of the fact that most tests require the system to be disconnected from normal operation. Output from these pages can be sent to the onboard printer, if a hard copy is required. In most FMC systems the available maintenance pages include, but are not limited to the following items: •

Navigation data downloading and cross-loading. This page is used to update the navigational database every 28 days. The database is update every 28 days to update for airways, radio navigational aid and other airline route change information that may have occurred.

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Performance. The purpose of this page is used by maintenance personnel to update and change the performance data of the aircraft. This data is used by the FMC to calculate flight plans and aircraft performance.

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IRS. This page shows an estimate of the aircrafts position for each Inertial Reference Unit (IRU).

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BITE. Built in Test Equipment (BITE). This page allows maintenance personnel to access different systems on the aircraft and perform set performance tests.

CDU Interface Normally, the CDU’s communicate directly to the FMC via ARINC 429 interface connections. The I/O Concentrator provides all of the required ARINC 429 interface connections to the two CDU’s. The I/O Concentrator communicates with the translation software in the Computer software.


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Flight Management Computer The Flight Management Computer (FMC) is the heart of the Flight Management System. The FMC, in conjunction with other aircraft systems, forms an integrated, full-flight regime control and information system, which provides automatic navigation, guidance, map display, and inflight performance optimization. It is designed to reduce cockpit workload during each flight phase by eliminating routine tasks and computations normally performed by the flight crew. When coupled to the autopilot, flight director, and auto-throttle, the system can provide automatic guidance through integrated commands for controlling roll, pitch, and engine thrust. The FMC/FMS provides flight planning, navigation, database storage, navigation display, guidance, and performance optimization.

Flight Management Computers Can be coupled to the autopilot, flight director, and auto throttle to provide guidance through integrated commands for controlling roll, pitch and engine thrust and provides course guidance for RNP/RNAV (Radio Navigation Point / Area Navigation) operations. All aircraft systems which are controlled by a computer also run built-in-tests on all associated sensors and processing circuitry to confirm system serviceability, and to electronically report equipment failures to initiate maintenance action.

Flight Management Computer (FMC) Operation Usually two FMC units are fitted to most aircraft, with some having three fitted. Each FMC is usually located to a crew position (left and right). All FMC’s receive the same input data from the aircraft systems and data entered into any CDU on the aircraft. In all cases the FMC is connected to its own CDU. Data entered into a CDU is sent to all the FMC’s fitted in the aircraft, where each FMC will action it independently from each other. The output from the FMC is sent to only the one CDU normally. This provides system redundancy in case one FMC fails and the other still has all the current aircraft data to control the operation of the aircraft. In case of FMC failure, the CDU can be selected to another FMC. This allows each flight crew member to still enter data and information into the operative FMC. If this option was not available, then the flight crew member whose FMC had failed, would then be unable to adjust the aircraft flight planning. Switching from one FMC to the other is usually controlled through the use of an alternate or source select switch located on the forward aircraft instrument panel. There is one for each of the crew member on the aircraft. In some modern aircraft the output from the CDU is also able to be switched from one CDU to another. The FMC interfaces with many areas of the aircraft to aid in its primary functions. These include such areas as the fuel quantity system, flight control system, air / ground data link systems, radio navigational systems, the inertial reference system, the thrust management system and the aircraft EICAS/ECAM and EFIS systems. The FMCS system is used by the flight crew to reduce the crew workload when planning, undertaking, monitoring and executing a flight plan. The major functions of the FMCS are: •

Flight Plan map display

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Automatic navigation/radio tuning

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Thrust management

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Lateral guidance

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Vertical guidance


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Map Display The FMC generates both dynamic map and map background data outputs which normally go to the EFIS on an ARINC 429 high-speed link. Dynamic map data is normally related to the airplane motion with respect to the flight plan and includes such data as airplane track, ground speed, time and distance to go, computed winds and vertical/lateral deviations from the active flight path. The computed background map data includes the location of waypoints, navaids, obstacles, and airports within the EFIS field of view. In the map mode, dynamic map data must be updated to reflect aircraft motion and is computed at 10 Hz. The background data is normally slow changing and is updated every three seconds.

Thrust Management: The FMS thrust management function is capable of performing autothrottle control law calculations based upon commands from the navigation function, as well as direct crew input from the FMC, manual adjustment of throttle position, or AFDS autothrottle commands. The autothrottle control law function provides automatic N1 equalization in all modes of flight, as well as thrust limit protection and N1 thrust requirement calculations to maintain MCP or AFDS speed and thrust settings. Autothrottle modes can be selected or overridden by the crew as required.


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TOPIC 13.3.9/10 - AUTOTHROTTLE AND AUTOLAND

Automatic Landing During this operation, the autopilot system must be at its most accurate. Limits must be imposed on the system, so that in the case of a runaway, the recovery rate is fast enough to avoid an accident situation. Two requirements govern the safety devices fitted into the system. They must: • limit the effect of a runaway such that a safe recovery can be made by the pilot • allow sufficient authority to the control system so that the required flight path may be followed even though the aircraft encounters disturbances such as crosswinds, etc. The AFCS must be capable of achieving the following during automatic landing. They must: • warn of a passive failure • complete the manoeuvre following an active or passive failure. To overcome these requirements, it was decided to use the concept of system redundancy. This is to utilise multiple systems operating in such a manner that a single failure within a system will have an insignificant effect on the aircraft during its approach and landing. The following is a list of terminology used in the description of system redundancy.

Failsafe (fail soft/fail passive) This describes the ability of a system to withstand a failure without producing excessive deviations from the flight path.


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Fail-operational (fail active) This is a system in which one failure can occur, but leaves the overall system still functioning, without causing degradation of performance outside the limits required for automatic landing.

Simplex (single or non-redundant) The term simplex defines a single automatic control system and its appropriate number of sub channels.

Multiplex This term applies to a system comprising of two or more independent simplex systems used collectively, so that in the event of a failure of a system, the remaining system is capable of performing the controlling function.

Duplex system A duplex system comprises of two complete systems of channels which are interconnected and which together provide continuous control.

Triplex system This is a fail-operational system of three complete systems or channels which are interconnected and which together provide continuous control. In the event of a failure of one of the systems or channels, that system or channel is overcome by the other two and is automatically engaged. Control is therefore continued in duplex.

Duplicate-monitored This refers to a system comprising two systems in parallel and with separate power supplies. The components of both are designed to be either self-monitoring or to have their outputs checked by parallel comparator circuits. Only one system is engaged at a time, the other being in the follow up mode. In the event of a failure, the system is automatically changed to the other.

Dual dual This is used by some manufacturers to define a twin fail-operational control system having twin passive monitoring systems. It should not be considered synonymous to a duplex system, since the control systems may or may not be active simultaneously. In the event of a monitor detecting a failure in its associated system, the second system with its monitor is switched in.


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Monitoring This term defines the process of comparing between two or more outputs, or between an output and datum. It can also be used as a limiter, to cause a system to disconnect when an output exceeds the prescribed limit.

Comparator This operates on data supplied from comparative stages in two or more similar systems.

Equaliser This device adjusts the subsystem in multiplex systems to remove differences between subsystem outputs that may arise other than fault conditions.


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Automatic Landing Sequence The figure shown is the indicator used for an automatic landing approach based on a system using triple digital flight control computer channels that allow for redundancy to operate in the fail operational and fail passive conditions previously described. Depending upon the number of channels that are armed and engaged, the system performs what is termed a land 2 status or land 3 status autoland. Land 2 signifies there is dual redundancy of engaged flight control computers sensors and servos (fail passive), while land 3 signifies triple redundancy of power sources, engaged flight control computers, sensors, and servos (fail operational). Each status is displayed on an autoland indicator.

Figure shows a typical automatic approach. During cruise and initial stages of approach to land, the control system operates as a single channel system, controlling the aircraft about its pitch and roll axis and providing the appropriate flight director commands. As multichannel operation is required for an automatic landing, at a certain stage of the approach, the other two channels are armed by a switch on the flight control panel. This will also arm the localiser and glideslope modes. Both the off line channels are continually supplied with the relevant outer loop control signals and operate on a comparative basis the whole time. Altitude information essential for vertical guidance to touchdown is always provided by signals from a radio altimeter which becomes effective as soon as the aircraft’s altitude is within the altimeter’s operating range.


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When the aircraft has descended to 1500 feet radio altitude, the localiser and glideslope beams are captured and the armed off line control channels are then automatically engaged. The localiser and glideslope beam signals control the aircraft about the roll and pitch axis so that any deviations are automatically corrected to maintain alignment with the runway. At the same time, the autoland status displays LAND 2 or LAND 3 on the indicator and computerised control of flare is also armed. At a radio altitude of 330 feet, the aircraft’s horizontal stabiliser is automatically repositioned to begin trimming the aircraft to a nose-up attitude. The elevators are also deflected to counter the trim and to provide pitch control in the trimmed attitude. When the landing gear is 45 feet above the ground (gear altitude), the flare mode is automatically engaged.


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The gear altitude is based upon radio altitude, pitch attitude, and the known distance between the landing gear, the fuselage and the radio altimeter antenna. The flare mode takes over pitch attitude control from the glideslope, and generates a pitch command to bring the aircraft on a 2 feet/second descent path. At the same time, a throttle retard command signal is supplied to the auto throttle system to reduce engine speed. Prior to touchdown and about 5 foot gear altitude, the flare mode is disengaged and there is transition to the touchdown and roll out mode. At about 1 foot gear altitude, the pitch attitude of the aircraft is decreased to 2 degrees, and at touchdown, a command signal is supplied to the elevators to lower the aircraft’s nose and so bring the nose wheel in contact with the runway and hold it there during roll out. The AFCS remains in control until disengaged by the pilot.


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Auto Throttle System This system is computer controlled and controls the thrust of the aircraft’s engines within specific design parameters. It is designed to operate in conjunction with the AFCS to maintain an aircraft’s speed and vertical path. When the AFCS mode is controlling the vertical path of an aircraft, the auto throttle maintains airspeed through thrust control.


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Take Off Mode This mode is initiated prior to take off by engaging the flight management system. The FME computer provides the engine rotational speed (NI) limits for each flight profile and an NI target speed. The limits and target speeds are displayed on the NI speed indicators. Arming of the auto throttle system is done by moving the engage switch on the control panel to the arm position. Engagement of the system with the servo-actuators controlling the throttles is done by pressing lever mounted switches designated as take off/go around (TOGA). The servo-actuators then advance the throttles at a particular rate to the position to obtain the correct NI value to attain the aircraft’s take off speed.


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Speed Control Mode This mode can be selected through the mode control panel (MCP) of the AFCS. The speed target set and displayed on the MCP is referred to as the MCP SPD. If vertical navigation (V NAV) control is selected for AFCS operation, the speed target is provided by the flight management computer and is referred to as FMC SPD. The auto throttle system is switched into this speed mode when an aircraft approaches a selected altitude under V NAV control, and will remain in this mode during altitude hold. Airspeed/Mach feedback signals are provided by the air data computer. When the aircraft begins to descend under V NAV control, the auto throttle system retards the throttles to idle. When the AFCS captures the glideslope beam, the VNAV mode is disengaged and the auto throttle system switches to the MCP SPD mode.

Auto Throttle System This system is computer controlled and controls the thrust of the aircraft’s engines within specific design parameters. It is designed to operate in conjunction with the AFCS to maintain an aircraft’s speed and vertical path. When the AFCS mode is controlling the vertical path of an aircraft, the auto throttle maintains airspeed through thrust control. Autothrottle can be utilised during approach to maintain the correct approach attitude and speed. Used in conjunction with the AFCS, the autothrottles can maintain an optimum AOA in the final approach phase before flaring and touchdown on the runway threshold. Typically engaged using Autopilot control panel. Disengagement switch typically on throttle, in illustration disengage is on outer side of each throttle lever (indicated by labels on throttles).


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Thrust Management Computer (TMC) Purpose is to automatically set the proper thrust level for the engines. The output servo motor moves the throttle linkage to set the level of engine power calculated by the TMC. Servos can be electrically powered, or powered by pressurised air, fuel or hydraulics. The system includes sensors on engines which monitor the important engine operating parameters. The monitoring of engine parameters is used to prevent exceeding any engine operating limitation for RPM, EPR, EGT, etc. The autothrottle system can be used to maintain a given climb rate, indicated airspeed, Mach number, or descent rate. The 767 has autoland capabilities and the autothrottle system will automatically close the throttles just prior to landing so that a smooth touchdown can be made. The TMC also provides a minimum speed protection which will maintain a safe margin above stall speed for the particular flight configuration


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Part 66 Subject

Selection of a desired speed value is made on the AFCS mode control panel. Approach gain of the auto throttle system is determined either by glideslope capture or by radio altitude and flap position. Approach gain provides high gain setting for more precise speed control and reduced throttle motion during changes of flap position. During an approach in turbulent conditions, the gain tends to cause the system to be high on speed and the degree of overspeed depends on the magnitude of turbulence. During the landing flare maneuver, the retard rate of thrust reduction is adjusted so that throttle angle is reduced to idle in 6 seconds. Retard occurs at 27 feet of radio altitude during an automatic or manual landing. If it is not initiated by radio altitude, it can also occur 1.5 seconds after an automatic flare. When the aircraft lands and the ground mode is sensed, the throttles are moved aft at 8 degrees a second to remove any residual displacement above the idle position. The autothrottle system is disengaged after 2 seconds.


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B2-13c Auto Flight SR

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