Auto Flight Part 1

PERTH COLLEGEBRAHAN BUILDING

CRIEFF ROADPERTH PH2 1NX

TEL: 01738 877105FAX: 01738 553369

AUTOFLIGHT – PART 1Part 66 – B2/014(a)

AIRSERVICETRAINING(ENGINEERING)LIMITED

A Subsidiary of Perth College

© Air Service Training (Engineering) Ltd

AERONAUTICAL ENGINEERING TRAINING NOTES

These training notes have been issued to you on the understanding that they are intended for your guidance, to enable you to assimilate classroom and workshop lessons and for self-study. Although every care has been taken to ensure that the training notes are current at the time of issue, no amendments will be forwarded to you once your training course is completed. It must be emphasised that these training notes do not in any way constitute an authorised document for use in aircraft maintenance.

All Rights Reserved

The copyright in these technical training notes remain the physical and intellectual property of Air Service Training (Engineering) Ltd, (AST). Copying, storing in hard copy or electronic format, transmission to third parties and use for teaching by establishments other than AST is forbidden, except with the written permission of the AST General Manager.

M HaufeTraining Manager February 2006

NOTES:

© Air Service Training (Engineering) LimitedPart 66 – B2014(a) Autoflight Part 1

CONTENTS PAGE

CHAPTER 1 : ERROR SIGNALS

SECTION 1 : Flight Controls 01.01.01

SECTION 2 : Attitude and Outer Loop 01.02.01

SECTION 3 : Sideslip Sensing 01.03.01

CHAPTER 2 : SIGNAL PROCESSINGSECTION 1 : Roll Axis 02.01.01

SECTION 2 :Integrators 02.02.01

SECTION 3 : Gain Programmes 02.03.01

SECTION 4 : Channel Data Interface 02.04.01

CHAPTER 3 : DEMAND SIGNALSSECTION 1 : Selectors/Indicators 03.01.01

SECTION 2 : Control and Limiting 03.02.01

SECTION 3 : Mode Compatibility 03.03.01

CHAPTER 4 : COMMAND SIGNALSSECTION 1 :Engage Interlocks 04.01.01

SECTION 2 : Stabiliser Trim 04.02.01

SECTION 3 : Standard Actuator 04.03.01

SECTION 4 : Clutches and Brakes 04.04.01

SECTION 5 : Hardover Protection 04.05.01

CHAPTER 5 : CONTROL AXESSECTION 1 : Autostabilisers 05.01.01

SECTION 2 :Damping 05.02.01

Issued February 2006 Contents 1

© Air Service Training (Engineering) LimitedAutoflight Part 1 Part 66 – B2014(a)

NOTES:

2 Contents Issued February 2006


CHAPTER 1 : ERROR SIGNALSSECTION 1 : FLIGHT CONTROLS

FLIGHT CONTROLS INTRODUCTION

The flight control system controls airplane attitude changes as necessary to perform a desired flight profile. and includes primary controls which directly effect airplane attitude and secondary controls which augment or change the effectiveness of primary controls.

Typical Primary Controls – seven moveable surfaces.

Roll or lateral control is by two inboard and two outboard aileron surfaces. Two elevator surfaces are used for primary pitch control. Yaw control is by one rudder surface.

Typical Secondary Controls – Twenty-nine moveable surfaces.

Twelve spoiler segments increase roll control and directly effect lift and drag. The moveable horizontal stabiliser increases pitch control.

Twelve leading edge slats and four trailing edge flaps are high lift devices which change the effectiveness of primary control surfaces.

Issued February 2006 Chapter 1 : Section 1 1


2 Chapter 1 : Section 1 Issued February 2006




FLIGHT CONTROL INPUTS

Flight Control Operation

All flight control surfaces are moved by hydraulic power control actuators (PCAs). PCAs are controlled by mechanical linkage with manual, autopilot and fly by wire attitude inputs.

Manual Controls

A system of cables, quadrants, rods, etc provide for manual control from the control wheel, control column and rudder pedals. Manual control of the ailerons is aided by lateral central control actuators (LCCAs) which drive the mechanical systems in the wings.

Manual inputs include artificial feel components to provide feel in the absence of aerodynamic feedback pressure.

Autopilot Controls

Autopilot computers provide electrical commands to control servos which for ailerons are the LCCAs. The servos provide the mechanical control of the PCAs. The required position feedback information is by electrical signal from the servos.

Attitude Controls

Automatic attitude control is provided for yaw damping. Electrical commands to the servos result in mechanical motion of the MCA (main control actuator) control linkage. Position feedback is also be electrical signal from the servos.





HYDRAULIC POWER SYSTEMS

Hydraulic System

The hydraulics provide power to accomplish main mechanical functions, one of which is to move flight control systems and aerodynamic surfaces, and is typically made up of three systems:

left

centre

right.

Each of the systems can provide power for primary flight controls.

The left (colour-coded red) and right (colour-coded green) systems each have two power sources; One engine driven pump (EDP) and one alternating current motor pump (ACMP). The ACMP operates on demand.

The centre system (colour-coded blue) includes four power sources. Two primary sources are alternating current motor pump (ACMP). One air-driven pump (ADP) operates as a demand source. The fourth source on some aircraft is an emergency ram air turbine (RAT) which deploys if both engines fail in-flight.

A hydraulic motor generator (HMG) is often located in the left wheel well and provides a limited backup source after loss of all generator electrical power. The generator is driven by a hydraulic motor supplied from the centre hydraulic system.



NOTES:



SECTION 2 : ATTITUDE AND OUTER LOOP

Signal Computations within the FCC are divided into two general types;

an outer loop and

an attitude inner loop

Signal integrity requirements and response to failure is governed by the type of computation involved.

LOOP FUNCTIONS

Outer loop

The outer loop computations use targets and sensors appropriate to the operating mode to command attitude changes to accomplish airplane control to a vertical speed, heading, etc. These commands are integrated to develop a desired attitude (pitch or roll) for the airplane.

Attitude loop

The attitude loop compares the desired attitude and the actual attitude and uses any error along with rate to develop a command for the surface.

INTEGRITY REQUIREMENTS AND FAILURE RESPONSE

Cruise outer loop

For cruise operation, outer loop signals need to be present from only one sensor to allow operation. Loss of the outer loop sensor signal in general results in a mode fail with reversion to attitude hold.

Attitude loop

At least two sensor sources must be available for the attitude loop signals for comparison. Loss of this level redundancy causes autopilot disengage.

Autoland requirements

When more than one channel is engaged in approach, at least two outer loop sensor sources are required. Loss of this level of redundancy causes autopilot disengage.



AFDS FCC – Attitude & Cruise Outer Loop



SECTION 3 : SIDESLIP SENSING

CONTROL

Displacing ailerons during a turn creates a difference between the lift and drag forces acting on each wing resulting in the aircraft yawing in a direction opposing the roll. The yaw (depending on bank angle and airspeed) produces SIDESLIP in a direction opposite to the yaw.

The aircraft also loses height in a turn, due to nose-down displacement. All these effects must be controlled to co-ordinate the turn.

Ordinary 3-axis displacement sensors will not detect the effects of sideslip therefore other sensors must be used.

One method uses 2 accelerometers, one mounted at the C of G position and the other mounted as far forward as practicable. The outputs of both accelerometers are summed giving a resultant signal in proportion to the yaw rate.

A turn command signal is now applied to the system roll control channel and starts aileron displacement for the turn. The changed attitude about the roll axis is sensed by the vertical gyro unit, who’s synchro detector provides a roll attitude change signal to be summed with the turn command signal. This ‘backs-off’ the turn command signal to only the amount required for the turn.



To overcome the opposing yawing motion, the rudder must be displaced in the direction of the turn and in order to achieve this a signal must be supplied to the yaw control channel and rudder actuator.

The signal is given by dividing the roll attitude change signal from the vertical gyro unit by true airspeed (TAS). TAS being obtained from a function generator being applied with airspeed and altitude signals from the ADC. The divider circuit output is therefore a signal representing the desired true yaw rate for a particular roll angle and airspeed.

Actual yaw rate and desired true yaw rate are now summed, any difference is gain programmed as a function of indicted airspeed and fed to the yaw control channel to reposition the rudder to provide further damping of the yawing motion rate. The gain is decreased as airspeed increases.

The yaw rate signals mentioned so far are essentially short-term. In the case of an extended period of turn, continuous co-ordination is achieved by routing the signal from the accelerometer positioned at the C of G via a ‘lag’ element and second gain programmer, thus producing a ‘steady-state’ rudder displacement.

In some aircraft types, airspeed switches are used to ‘lock-out’ computed yaw rate signals where such signals are not wanted if above a particular airspeed.



CHAPTER 2 : SIGNAL PROCESSINGSECTION 1 : ROLL AXIS

A/P ROLL AXIS CONTROL SIGNAL FLOW

Lateral Navigation

In LNAV mode, the FMC computes a desired roll angle to accomplish the desired lateral navigation. This command is received by the FCC as horizontal steering and processed for flight director commands and aileron commands.

F/D processing

The horizontal steering command is compared to roll angle to generate the F/D roll command. Roll angle is never washed out in LNAV operation.

Autopilot

With LNAV engaged, the horizontal steering command is compared to the roll command output of the roll attitude integrator. Any differences drive the roll command to be equal to the horizontal steering command. The drive is limited to limit the rate to 6° per second. The ailerons are driven to make roll angle equal to roll command.



NOTES:



SECTION 2 INTEGRATORS

Two types are used in automatic Flight Control systems, the most common of which is the integrating operational amplifier. It is used to provide a ‘ramped’ output whose final value is the value of the input signal. The time taken to reach this value is the integration time which is determined by the time constant of the input resistor and the feedback capacitor. The ‘delayed’ output obtained is brought about by the feedback being anti-phase to the output.

This type of integrator is typically used in a flying control command limiting circuit, providing a controlled command rate, or in channel signal processing for synchronizing or, if an aircraft is flying a VOR radial in crosswind conditions, where error signals due to recurring displacement must be ‘washed-out’.

The other type of integrator is an integrating servo-mechanism. A typical example is when an aircraft altitude increases very slightly, while in the cruise, over a long period of time due to fuelburn. In order to sense this a servomechanism is used with its output shaft driving the rotor of a synchro through a ‘low gear ratio’. The output is therefore a measure of the change in altitude and when it reaches a significant value, it can be used to drive the pitch channel to correct the altitude.



SUMMING POINTS

Summing Points in the signal processing network are typically resistor networks with a resistor for each voltage arriving at the junction. The voltage at the output will be the algebraic sum of the inputs.

Alternatively, a summing amplifier may be used, in which case the resistor network is located at the amplifier input and the output will be the algebraic sum of the inputs and the gain of the stage.



SECTION 3 : GAIN PROGRAMMES/BUFFERS

GAIN

The gain of an Automatic Flight Control System determines how a particular aircraft type will respond to the commands of a given autopilot system and therefore must be adjustable on installation or servicing. This can be achieved by using variable resistors in servo-mechanism feedback lines, or by using suitable ‘gearing pads’ which are fixed value resistors or capacitors which form the feedback of an amplifier stage.

Gain programming is where the gain of a system must be varied in ‘real-time’ to suit changing flight conditions.

Typical examples of GAIN PROGRAMMING are:

When turbulence or soft-ride mode is selected the system gain is reduced so that the autopilot will not respond violently to sudden disturbances caused by gusts or clear air turbulence.



During an ILS approach and subsequent to localiser capture, the ILS receiver captures the glide slope and the AFCS will command pitch down until, at 1500ft, which is the designed altitude for gain programme initiation, the gain will gradually be decreased from 100% to 57% as the aircraft proceeds down the glide slope and radio altitude decreases. If RAD ALT information becomes invalid during this period, a step-change in gain is initiated after 10 seconds of glide slope capture, reducing the gain to 80% and after 120 seconds a further step change reduces the gain to 57%.

If RAD ALT is lost 10 seconds after point of capture middle marker will re-establish glide slope gain signal (typical).

BUFFER AMPLIFIERS

Buffer Amplifiers are, by definition, amplifiers with unity gain, ie. the value of the output equals the value of the input. The benefit derived is of a clean signal. Buffers are normally used in long transmission lines which may be liable to crosstalk or induced stray signal from wires nearby. This type of buffer is normally found in digital transmission systems.

Buffer, amplifiers in analogue circuits are normally amplifiers of gain greater than unity, or of negative gain, where it is required to change the amplitude of a signal to match with various circuit components or system voltages.



SECTION 4 : CHANNEL DATA INTERFACE

FCC/CROSS – CHANNEL DATA INTERFACES

Digital data is transmitted cross-channel between FCCs for monitoring, signal selection, and synchronization.

The analog discrete logic for servo engage and detent trip are also sent cross-channel. They go to dedicated hardware to provide a redundant path for failed FCC disengagement.

Digital interfaces

The digital data busses utilize ARINC 429 high speed design. Cross-channel bus parameter labels generally are not identical to the respective source label. This allows a simplified, combined cross-channel receive function.

Analogue Discrete Interface

These discretes are either ground or 28 volt dc.

28 volt dc Ground

Servo ENGAGED NOT ENGAGED

Detent NOT TRIPPED TRIPPED





CHAPTER 3 : DEMAND SIGNALSSECTION 1 : SELECTORS / INDICATORS

SELECTORS

Typically, two types of mode selector exist and two discrete panels are employed:

The MODE SELECTOR panel

The AUTO PILOT CONTROL panel

In most applications they are mounted near to each other, often on the flight deck centre panel.

The mode selector panel has two rows of push-switches which light when the particular mode is selected ie. ‘APPR’ (Approach-ILS) ‘IAS’ (airspeed hold) or ‘HDG’ (Heading Hold). In systems which include a flight director, some of the modes will be purely F/D modes and some will be F/D & Auto-pilot modes (see Mode Compatibility in later chapter).

The auto pilot control panel provides switches to engage/disengage the auto pilot and yaw dampers if applicable. Normal turn control and pitch control switches are located here, in certain systems.





MODE INDICATORS

Mode indicators are sometimes found on the glareshield panel and usually take the form of a transparent strip behind which are lit coloured captions showing the selected mode and, perhaps more importantly, the STATUS of the AFCS in a particular mode. For example, with an AFCS operating in ILS (APPROACH) mode, subsequent to mode selection ‘APPR ARM’ would be displayed and simultaneously ‘G/S arm’ (glide slope) would be illuminated until the aircraft acquired the glide slope beam when only ‘APPR’ and ‘G/S' would be displayed, showing that the aircraft was established on the final approach to the runway.

For aircraft equipped to CAT 2 or CAT 3 configurations and during an approach, if any data from system sensors becomes invalid, making the control unreliable, a warning to this effect is displayed on the display panel.

On larger and more modern aircraft, the control and mode selection functions are together on the flight control panel which is typically located in the middle of the glareshield panel.



CONTROL AND DISPLAY UNITS

Control and Display units are common names for the typical flight deck installed panel allowing control by switching and display by lamp, LED 7-segment display, LCD, etc these panels are commonly used for INS, Global Nav or Area Nav systems. This terminology is extended to modern AFCS for maintenance information.



MAINTENANCE CONTROL & DISPLAY PANEL



MAINTENANCE CONTROL AND DISPLAY PANEL

A typical MCDP (Maintenance Control Display Panel) is located in the avionics bay and monitors input from all Flight Control Computers and multiple system sensors. Any fault is indicated on a Dot-matrix or 7-segment LED or LCD, or Magnetic Indicator (Doll’s eye) type display which assists maintenance. In some cases, the display is a binary fault code and during BITE testing (built-in test equipment) suitable messages are displayed to indicate the serviceability status of the system. In an aircraft such as the 757 (200 series) a plain word display is available which not only informs of status but indicates possible fault.

REMOTE MCDP OPERATION

On some aircraft maintenance is accomplished by one person through use of a remote MCDP control panel in the flight compartment. The remote MCDP control is a special test equipment. The hand-held, carry on panel can be connected to a plug on the flight deck. Access is by opening the panel. A cutout is provided for the connecting cord.

The remote panel’s switches provide for the switch functions on the front of the MCDP. Mode annunciations and alphanumeric messages are provided on the EICAS Display. The CONF/MCDP button must be selected on the EICAS maintenance panel. The airplane must be on the ground to view the EICAS maintenance pages.

TURN CONTROLLERS

The Turn Controller is located either on the Flight Control Panel or the Auto Pilot Control panel and provides a facility for the pilot to normally inject a turn command into the autopilot system. In some systems the control knob operates potentiometer or rotor of a synchro to generate the error signal. In its central position, the control knob engages a detent which operates, via a switch function, an interlock in the preselected heading signals from the compass system, or radio heading data. In the event of the turn controller being operated when in either of the coupled modes, ‘out-of-detent’ of ‘camlock’ warnings are usually given. Similarly, the autopilot may not be engaged with the turn controller displaced from its detent.

PITCH control can normally be achieved in a manner similar to the above except that instead of a turning knob, a ‘fore and aft’ movable switch provides a 2° pitch command in the appropriate direction each time the switch is deflected.





SECTION 2 : CONTROL AND LIMITING

AILERON FORCE TRANSDUCER AND LIMITER CWS

Purpose

Aileron Force Transducer – The purpose of the aileron force transducer is to convert mechanical force into an electrical signal to be used in the flight control computer.

Aileron Force Limiter – The purpose of the Force Limiter is to mechanically limit the amount of aileron command input while the autopilot is engaged.

Location

Aileron Force Limiter – The aileron force transducer is attached between the control drum and the bus drum of the captain’s control column.

Aileron Force Limiter – The aileron force limiter is attached to the lower end of the captain’s control column.

Physical description

Aileron Force Transducer – The force transducer is a cylindrical unit approximately 5 inches long with a diameter of 2.25 inches. Two electrical connectors provide excitation and command wire connections.

Aileron Force Limiter – The aileron force limiter has two sections which are the mechanical force limiting mechanism and a motor driven relay section. Two electrical connectors provide connections for clutch operation and motor and relay connections.





Operation

Aileron Force Transducer – The force transducer receives 26 volts ac 400Hz excitation for each of the transducers. The output represents the force applied on the control wheel. The force transducer is mounted between the bus drum and control drum, below the captain’s control column. The force transducer transmits the mechanical movement to the control drum to move the lateral controls when the A/P is not engaged. When the A/P is engaged, the force transducer develops the electrical signal to provide the control wheel steering signal for the A/P. The force transducer is constructed as a spring. The spring is expanded or contracted between the drums. An internal ‘E’ field transformer converts the mechanical movement to an electrical signal which is used in the A/P computations within FCCs. If the transducer fails mechanically, the bus drum contacts the control drum to maintain mechanical continuity.

Aileron Force Limiter – The aileron force limiter, limits control wheel movement during A/P operation. A clutch in the aileron force limiter connects it to the bus drum when the A/P is engaged. The limit is set by a spring and cam mechanism. The cam is driven from the control drum until the spring is contacted. The pilot may override the force limiter by compressing the spring by force on the control wheel. A motor assembly is used to provide stops of 17° with flaps up and 25° with flap not up. Micro switches within the motor assembly are used in conjunction with flap switches to assure correct stop positioning.

Maintenance practices

Aileron Force Limiter – The aileron force limiter is a line replaceable unit.

Aileron Force Transducer – The aileron force transducer is a line replaceable unit. Adjustment after installation is required. There are two adjustments on the transducer, coarse adjustment is made by turning the rod end, and fine adjustment is made with a knurled nut on the rod.



SECTION 3 : MODE COMPATABILITY AND COMMAND OVERRIDE

In many cases LATERAL autopilot/flight director modes will be compatible with VERTICAL modes, providing in certain cases that vertical mode is engaged first and vice versa. In other cases, modes will not be compatible.

Typical examples of mode compatibility

MODE 1 MODE 2 COMPATIBILITY

Heading Hold Heading Select No

Approach Go-around Yes, if appr first

Vert-speed Vert-nav No

Nav Vor No

Vert-speed Heading Yes

Obviously, two vertical modes or two lateral modes cannot be engaged simultaneously.



NOTES:



CHAPTER 4 : COMMAND SIGNAL OUTPUTSSECTION 1 : ENGAGE INTERLOCKS

INTRODUCTION

The autopilot engage interlock circuits in the FCC monitor the operation, power and components of the AFDS. When normal conditions exist, the circuits allow the autopilot to engage and to remain engaged. When the circuits detect certain non-normal conditions, the autopilot disengages or is prevented from engagement.

Autopilot engage

Autopilot engage starts by a push of the CMD button on the MCP. This makes the engage request.

In addition to proper input power (115v ac and 28v dc), these AFDS components must be valid to engage:

Flight control computer

Mode control panel

Aileron and elevator A/P servos

Also, data and conditions monitored on the cross-channel inputs must be proper to engage. The FCC supplies ARM and ENGAGE voltage to the aileron, elevator and rudder servos.

Manual disengage

The operation of either the A/P disengage switch or the disengage bar disengages the autopilot.



AFDS – AUTOPILOT ENGAGE INTERLOCKS



AFDS – Engage interlock inputs

Power

AC and DC power required to enable A/P engagement. The DC battery bus must be present to engage the autopilot. If engaged, loss of this BUS does not disengage the autopilot.

A/P engage request

A/P engage starts by a push of the engage button on the Automatic Flight Control System Mode Control Panel. The A/P engage logic in the Flight Control Computers receives the A/P engage request and MCP status and starts the engage process.

A/P disengage

The method for the pilot to disengage the autopilot is to use the disengage bar switches or the A/P disengage switches.

Cross channel data

The right and centre FCCs send engage status and cross channel data to the left FCC.

Autotrim valid

Autotrim valid from the stabiliser trim/rudder ratio module (SRM) is required to stay engaged if single channel.

AFDS – ENGAGE INTERLOCK INPUTS



AFDS – ENGAGE INTERLOCK OUTPUTS

Power

The servo LVDTs and surface LVDTs in the aileron, elevator and rudder servos receive 26v ac from the FCC.

A/P solenoid and LVDTs

Arm and engage discretes energise the solenoids in the aileron, elevator and rudder servos. The FCC monitors the servo and surface LVDTs for synchronization.

Cross channel data

The left FCC supplies servo engage status, detent trip status and cross-channel data to the right and the centre FCCs.

AFDS – ENGAGE INTERLOCK OUTPUTS



AFDS – ENGAGE INTERLOCK OPERATION



AFDS – ENGAGE INTERLOCK OPERATION

Power

DC power arms the contacts of four relays in the FCC. DC is also applied through the disengage bar switch to energise the arm and engage relays.

Elevator and aileron arm

The arm phase of autopilot engage starts by a push of the CMD pushbutton. The request goes through the data bus to the elevator/aileron arm logic. With normal conditions, the logic energises the elevator/aileron arm relay. This energises the arm solenoids in the aileron and elevator servos.

Elevator and aileron engage

At the end of the arm phase, engage status is initiated by the elevator and aileron engage logic. The elevator/aileron engage relay energises. This energises engage solenoids in the aileron and elevator servos.

Rudder arm

Rudder arm begins with multi-channel arm. The rudder arm logic energises the rudder arm relay. This energises the arm solenoid in the rudder servo.

Rudder engage

When multi-channel engage begins, the rudder engage logic energises the rudder engage relay. This energises the engage solenoid in the rudder servo.



SECTION 2 : STABILISER TRIM

MANUAL ELECTRIC TRIM SYSTEM

The stabiliser trim control switches, located on the outboard horn of each control wheels, provide up or down ARM and CONTROL signals to the two stabiliser trim and aileron lockout modules (SAM’s), located in the main equipment centre. The SAM’s provide up or down ARM and CONTROL signals to solenoids on the stabiliser trim control modules (STCMs), located above and forward of the stabiliser trim actuator in the stabiliser compartment. Control module ARM and CONTROL hydraulic valves function in series and allow hydraulic flow to the two hydraulic motors and two hydraulically released brakes on the stabiliser trim ballscrew actuator assembly, located forward of the stabiliser.

Hydraulic inputs from the elevator feel computers control the rate of stabiliser trim. The left and right limit switch and position transmitter modules, located below the stabiliser, provide position feedback through Flap/Stabiliser Position Modules (FSPMs) to the Flight Control Computers (FCC) and both SAMs. The position and limit switch modules also drive the flight compartment position indicators, and limit stabiliser travel by cam actuated microswitches which break the electrical path from the SAMs to the STCMs. Column operated cutout switches, located beneath the floor on each outboard side of the control column torque tube, also break the electrical path from the SAMs to the STCMs when the control columns are moved in a direction opposing stabiliser trim.



STABILISER TRIM BLOCK DIAGRAM



STABILISER TRIM

Autopilot Trim System

Three flight control computers (FCCs), located in the main equipment centre, provide autopilot input signals to the two SAMs to trim the stabiliser based on elevator out of neutral position. The operation of the SAM, the stabiliser trim control modules (STCMs), the stabiliser trim ballscrew actuator assembly and the limit switch and position transmitter modules is identical to the manual electric system.

The maintenance control and display panel (MCDP) tests the ability of each FCC (L, C, R) to command the SAMS to position the horizontal stabiliser. The MCDP also checks the FCC to SAM interfaces and stores flight faults associated with autotrim.

Mach trim system

Either SAM automatically trims the stabiliser according to airspeed signals from the air data computers (ADCs), located in the main equipment centre. Air/ground logic from the air/ground relays inhibit mach trim on the ground. Operation of the STCMs, the stabiliser trim ballscrew actuator assembly, and the limit switch and position transmitter modules is identical to the manual electric system described before.

Electric alternate manual trim system

The two electrical alternate trim switches provide up or down ‘ARM’ and ‘CONTROL’ signals directly to the dual coil solenoid valves. These signals energise one coil in the dual coil solenoid valves (the second coil is controlled by the trim signals from the SAMs). The valves operate as before, providing pressure to release the hydraulic brakes and drive the hydraulic motors to trim the stabiliser.

Stabiliser trim fault indication system

The SAMs control the logic associated with flight compartment amber annunciations on the P5 pilot’s overhead panel and displays on the engine indicating and crew alerting system (EICAS). The faceplate of each SAM has fault balls to record the status of various system LRUs.



AUTOMATIC STABILISER TRIM BLOCK DIAGRAM



AUTOMATIC STABILISER TRIM

Autotrim command

In cruise, autostabiliser trim is designed to keep the elevator within the neutral position of the elevator feel force unit. This is done to prevent pitch servo out-of-detent and to prevent large pitch transients if the autopilot is disconnected with the feel unit displaced from the neutral position.

During an approach when stab trim bias is applied, the stabiliser is positioned leading edge down causing the elevator to move in the opposite direction to maintain desired path. Trim bias leaves the airplane trimmed nose up in the event of total autopilot disconnect.

Command response monitor

The trim motion and direction detector verifies that stabiliser movement is consistent with the compound trim commands. If a movement/command anomaly exists, autotrim fault logic generates unscheduled trim or dead trim fault signals which activate autotrim failure annunciations and/or SAM/FCC select logic. Unscheduled trim faults are also detected in either SAM (Stabiliser Trim Aileron Lockout Module) where commands are generated to activate the Unscheduled Trim light on the P5 panel and annunciation on the EICAS display.

Command response faults are defined as follows:

DEAD TRIM – Up or down autotrim command with no up or down stabiliser movement.

UNSCHEDULED TRIM – Stabiliser movement with no autotrim command.

or

Stabiliser moves opposite direction to autotrim command.



NOTES:



SECTION 3 : STANDARD ACTUATORS

OPERATION

The accompanying diagram shows the schematic operation in a typical roll channel. When the aileron is not engaged with the power control unit, the solenoid valve (solenoid actuator) is de-energised, the select valve and engage mechanism are isolated from the hydraulic supply. The transfer valve and spool valve are also isolated by the select valve.

The cam spring of the engage mechanism is therefore relaxed (extended) releasing the A/P engage cam which allows the pilot to move the control wheel. The main control valve is centralized, thus permitting fluid to reach either side of the main actuator piston.

Remember that the actuator piston is anchored to the aircraft structure therefore when pressurized, it is the actuator body which moves and operates the aileron cable drum via a linkage.

The body (cylinder) of the main control valve is attached to the actuator body and will therefore move with it, with respect to its piston, which is attached to the control wheel via the input cranks.



When the roll channel is engaged, the solenoid actuator is energized allowing fluid under pressure to compress the spring of the engage mechanism. This action rotates the engage cam which locks the input cranks to the power control unit. The main control valve piston is now prevented from moving in relation to its body.

Hydraulic pressure is now admitted to the select valve which opens the hydraulic lines between the spool valve and the main actuator and also allows hydraulic pressure to the transfer valve.

Command signals from the roll servo-amplifier operate the deflection coils of the transfer valve which deflects the valve in the appropriate direction. This action causes a pressure differential across the spool valve which now moves allowing pressure to the appropriate side of the main actuator piston via the already open select valve.

As with normal control, the PCU will move bodily to operate the ailerons and because the inputs cranks are locked by the engage system, as the PCU moves, the pilot’s control wheel will also rotate. This is known as a parallel auto-flight system.

Finally, movement of the PCU also displaces the core of an LVDT (linear variable differential transformer) providing positional feedback signals which are required for reducing the command signal as the function is satisfied.



SECTION 4 : CLUTCH/BRAKES

Electro-clutches and electro-brakes are virtually the same equipment and, of course, are normally DC operated solenoids bringing together driving and driven plates clad in friction material. The main advantage of these devices is the prevention of control overrun when the signal is removed.

Torque limiting

Torque limiting of servo-motor drives can be accomplished in 3 ways:

By the insertion of a mechanical slipping clutch in series with the output drive (not shown).

By including a coil spring in the electro-clutch output shaft, which will compress when over-torqued and in doing so will allow rollers to override cams on the clutch body operating limit switches, to de-activate the solenoid. (figures shown).

Electrical torque limiting is commonly used with 2-phase servomotors. In this system a resistor is connected in series with the control phase and, additionally, torque-sensitive switches are situated in series with the reference phase. If the pre-determined torque is exceeded the reference phase will be open circuited (not shown).



NOTES:



SECTION 5 : HARDOVER PROTECTION

SENSING OF JAM AND HARDOVER PROTECTION

A potentially dangerous situation where a flying control servomotor is commanded to its ‘hardover’ position must be protected against. Typically, some systems use duplex servomotors.

The 2 servomotors are driven by duplex channel signals, which are also routed to a speed motor. If the signals are equal, the speed monitor output is zero, but if the speed of one servomotor is excessive, compared to the other, the speed monitor output operates the electro-brake on the faulty unit. The effect of this is to halve the servo power which reduces the risk of hardover. Since the torque on the remaining serviceable system is increase, its torque limiter will be activated which may cause eventual disconnect.

In systems which do not employ duplex servomotors, runaway (or hardover) is sensed by comparing analogue values of command signals so that any excessive values may cause an auto-pilot disconnect.

Some systems provide disconnect facilities which allow pilot override, such as a control wheel steering system where the operating system (springs and field-electric) detects excessive torque (normal application) and disengages the servo-clutches, at the same time giving indication of auto-pilot disconnect.



Control jam can be detected and compensated for basically in 2 ways:

MECHANICALLY For example, where 2 control columns are interconnected below floor panel level by a tube within a tube, separated by a spring loaded roller engaging a detent. If one control column jams (each column is connected to separate control runs), force applied to the other allows the spring loaded roller to override the detent and thus the other control column is free to control its own run.

ELECTRICALLY For example, if an aileron control run jammed, the pilot would apply force to rotate the handwheel; in this case a sheer pin would be severed, releasing the handwheel from the control run. Horseshoe spring and force switches now operate a solenoid releasing the control run to one aileron. A synchro transmitter in the handwheel hub row sends demand signals to the aileron servo-amp which now drives the aileron servo clutch of one aileron as a fly-by-wire system.



CAM OUT OPERATION

Dual channel

Mechanical operation

When both channels are hydraulically engaged (PITCH after G/S Capture ROLL after LOC APP) control surface position is given to the autopilot channel with the smaller demand signal. Each A/P actuator moves to satisfy its channel demand signal. When the A/P actuator of the channel with the smaller signal has moved sufficiently to satisfy its demand, it stops.

The A/P actuator on the channel with the larger signal keeps on moving, but its output side, driving through a force detent spring pot, is trying to overcome the combined force of the feel system and the force detent spring pot of the stationary actuator of the other channel. It is unable to do so, and its force detent spring pot collapses, this unlocks its fulcrum point, and the actuator output becomes ineffective. This is termed mechanical cam out.



EXAMPLE – LAND MODE

Dual A/P Cam-out, A/P with the smaller demand in control.

Both A/P Actuators move until (B) Actuator (smaller signal) satisfies its demand and stops.

‘A’ Actuator is unsatisfied and continues to move.

‘A’ Actuator moving but opposed by feel force plus force detent of the stopped ‘B’ actuator. Opposing force too great. Cam-out occurs.

Force detent breaks outs

Engage lock unlocks

Equalisation circuit in A/P synchronise out ‘A’ channel

Below 500ft wailer sounds.



PCU ACTUATOR CONTROL

A/P engaged lock (A) locks under hydraulic pressure so A/P actuator movement (B) goes to control run via the force detent spring unit (C).

If (C) breaks under cam-out conditions, the hydraulic pressure to the engaged lock (A) is relieved through the regulating valve (D) causing engaged lock (A) to unlock, A/P Actuator (B) position has no effect on control run.

The A/P LVDT (E) provides feedback to cancel A/P demand.



TRIPLE CHANNEL CAM OUT

With all three channels hydraulically engaged, (BOTH PITCH AND ROLL at LOC APP) control surface position is given to the channel with the medium signal (MID TERM VOTING), and the channels with the larger or smaller signals will be mechanically and electrically cammed out.

Cam out will be occurring continuously during approach due to small differences of channel command signals, but if a cam out on a channel reaches a certain magnitude and persists for a certain time, the cam out monitor will disconnect the affected channel.



ELECTRICAL ACTION

The A/P actuator has moved to satisfy its demand signal, but

It’s associated PCU has moved under control of the smaller demand signal of the other channel.

Position feedback signals are taken from the A/P actuator and the associated PCU ram and compared with the cam out monitor circuit.

If operation is normal, these two signals cancel out, but in the cammed out condition they are unbalanced. If the unbalance is of sufficient magnitude and persists for a set time, the cam out monitor will trigger and put on a steady RED light in the channel on which the cam out exists. Below 200ft R/A ht audio warning also sounds.



The circuit can be used to detect power failure or ‘Hardover’ ie. large error signals.

Assume the output of the ‘AND’ gate is keeping a relay energised and so allowing the A/Pilot to operate correctly.

Should a power failure occur the reference signal (28v or a typical value) would be lost. D1 would cease conducting, TR1 would ‘switch off’ and a non valid would cause the ‘AND’ gate to de-energise the relay.

Should a ‘Hardover’ occur then the two signals from the LVDTs (compared) would not be equal, creating an error signal of sufficient size to ‘breakdown D2’ and ‘switch on’ TR2 and a non valid would cause the ‘AND’ gate to de-energise the relay.

As a result of either actions, the A/Pilot could be disengaged, thereby protecting the system.



EQUALISING CIRCUIT

The unbalanced feedback signal in the cammed out channel is also fed back into the computer channel to reset it and reposition the A/P actuator.

Cam out is occurring continuously during an approach, due to small differences in channel command signals, this is normal operation, but a warning will only be given when the CAM OUT monitor threshold is reached.

NOTE: In the dual channel case, it should be borne in mind that the channel with the smaller signal may in fact be the faulty channel, so the appearance of a steady red warning light is purely an indication of ‘difference’ and identifies the channel with the larger signal.



NOTES:



CHAPTER 5 : CONTROL AXESSECTION 1 : AUTOSTABILISERS

The purpose of an automatic control system is to keep an aircraft on a correctly stabilized flight path by sensing and correcting any departures from the flight path.

Auto-stabilisers

An auto-stabiliser is an automatic device which improves the natural stability of an aircraft by operating control surfaces independently or as part of a pilot control system in such a way that the human pilot retains continuous control through his normal flying controls.

There is a tendency for modern aircraft to oscillate about their axis. The magnitude of this oscillation is quite small with the result that the human pilot tends to apply delayed over correction and thereby increases the problem.

Auto-stabilisers or dampers as they are sometimes known, are designed to detect and damp out oscillations and for this reason have no aircraft datums like an autopilot.

Before going on to explain a basic damper system, it must be understood what is meant by the term, phase advance.

When an aircraft is oscillating in this way, the purpose of applying a phase advanced signal is to maximize the damping of an aircraft’s response to its control signal.



Take the case of an aircraft which has been displaced in pitch along its flight path.

The natural longitudinal stability of the aircraft will tend to return it to its flight path. Its control system will also add to the natural stability. The aircraft will respond quickly to this deviation. There are two factors contributing to stabilizing the aircraft:

The control surface.

The natural stability.

When a control surface is being used for a correction in an aircraft with a low damping response, the inertial effect will cause it to overshoot.

To cancel this overshoot it is anticipated before the control surfaces reach the null and applied in the opposite sense to null out and dampen as the aircraft returns to its correct attitude.

In the example shown there is:

A rate signal which is proportional to the rate of attitude change.

A signal from the attitude reference unit (VG) which is proportional to the attitude displacement.

The resultant control signal is phase advanced with respect to the disturbances.



SECTION 2 : DAMPING

BASIC DAMPER

Let’s look at the basic damper shown with reference to auto-stabilisation. The disturbance would be sensed by the rate gyro and fed into a phase advance network (to increase damping) and then to the servomotor and control surface.

The servomotor, via a rate generator will supply a feedback signal in opposition.

The position feedback LVDT will provide a datum to centralize the control surface when acceleration is zero and to centralize the actuator prior to engagement. This basic damper would oppose deliberate manoeuvres.



ROLL DAMPER

Let’s update and apply it to a roll function as a turn compensated damper.

There are two inputs, the rate gyro and an aileron position transmitter.

As the aileron is moved to roll the aircraft the aileron position transmitter will output to the phase advance network and as the rate gyro senses the aircraft moving it will also output.

The aileron position signal opposes the rate gyro signal as the maneouvre is demanded.

When the desired roll attitude is acquired, phase advance output it zero, so the controls centralise and the damper will provide damping about that attitude.



Let’s look briefly at the yaw damper, as this is covered in more detail later.

This type of control crossfeed is used most with large transport aircraft, where the time constant needed for the wash-out network in the yaw damper is frequently unsuitable to achieve the required value of damping ratio for the closed loop system. The situation arises because the frequency of the dutch roll mode is very low, as are the frequencies of the structural bending modes, so that feedback of lateral acceleration cannot be used, otherwise there would be considerable coupling of the rigid body and structural motion. Furthermore, the relatively slow response of the rudder fitted to such large aircraft precludes the use of loop gain to suppress unwanted sideslip motion. Consequently the following technique is employed. The crossfeed signal is introduced into the summing junction of the yaw damper as if it were a command signal for some value of yaw rate which corresponds to zero sideslip angle.

The purpose is not only to provide damping but also to provide the signal to apply the correct amount of rudder to co-ordinate a turn.

The aileron position transmitter opposes the output from the rate gyro but drives the rudder to assist in the turn.

Sideslip

The sideslip monitor input can be included if the damper is used as part of the autopilot system.



RUDDER TO AILERON CROSSFEED

The efficiency of the auto-pilot in correction yaw disturbances is greatly improved by controlling the rudder and ailerons simultaneously. This is achieved by applying the yaw gyro output signal to both rudder and aileron channels.

This is known as rudder to aileron crossfeed and assists in co-coordinating turns made for small deviations from the desired heading.

Yawing disturbances are damped by the rudder but since the aircraft is steered by the ailerons, deviations from the desired heading are corrected by the ailerons banking the aircraft to turn it back on the original heading.


Documents

Auto Flight Part 1