Auto Pilot System
The core of every automatic flight system is the autopilot. This is an auto stabilisation system capable of maintaining the aircraft in stable flight about one or more of the aircraft’s three axis of roll, pitch and yaw. Early autopilot systems were only designed to maintain wings level flight by controlling the aircraft in roll, that is about its longitudinal axis, and such a system is still in use today in some light aircraft. This is known as a single axis, or single channel, autopilot providing lateral stabilisation.
Components of an AFCS
In general the components of any AFCS can be subdivided into three distinct groups as follows :
Sensors : These measure the relevant parameters and transmit the information in the computation group.
Computers : These convert the information from the sensors into the signals which are fed to the system output devices.
Output Devices : These convert the computed signals into a form which will result in the necessary aircraft control surface movements.
BASIC COMPONENTS OF AUTOPILOT
Inner Loop Control
The single axis lateral stabilising autopilot employs a closed loop control system in which the aircraft attitude in roll is controlled by operating the ailerons through a servomotor, or actuator. Any change in roll attitude is sensed by a rate gyro sensitive only to movement about the aircraft longitudinal axis. Such movement will cause the gyro to process and this precession will be picked off and transmitted as an error signal to a controller, which compares input signals with the ‘wings level’ message programmed in its memory.
The error signal will indicate the rate and direction of deviation from wings level and the controller will generate a correcting output signal of corresponding amplitude and rate of change.This signal is amplified and transmitted to the servomotor, which moves the ailerons in the appropriate direction to arrest the roll and restore the aircraft to wings level. The restoring movement is sensed by the roll gyro, with the result that the error signal received by the controller diminishes as the wings approach the level condition, until there is no error when the wings are once again level, by which time the servomotor has returned the ailerons to their neutral position.
It will be appreciated that a system such as that described is only capable of maintaining the wings level, because the only information the controller has is the programmed desired condition, the actual condition as sensed by the rate gyroscope and response feedback from the aileron servo motor.
It is incapable of maintaining a particular heading or course, since there is no directional information input to the system. Such a control loop is known as an inner loop, since it has no external references.
SINGLE AXIS SYSTEM
The ‘wings level’ reference programmed into the controller may be, and almost invariably is, adjustable by the pilot. By rotating a knob, for example, on a control panel the desired roll attitude can be biased to left or right and the controller will signal the servo motor to move the ailerons and roll the aircraft until the required bank angle is achieved.The autopilot loop will now maintain that bank angle until the reference attitude is re-adjusted.Thus the pilot can control aircraft heading through the autopilot system.
A second autopilot inner loop channel can be added, to provide automatic control of the aircraft about the pitch, or lateral, axis. Once again, this is purely an auto-stabilisation system that will maintain a preselected pitch attitude. The reference pitch attitude may be adjusted by the pilot, as before, in order that the autopilot maintains a selected aircraft attitude for level flight, climb or descent.
An autopilot system that has both roll and pitch channels is known as a two-axis autopilot and this is probably the most common form of auto stabilisation system. A third, yaw channel is often provided in larger transport aircraft, especially where stability about the yaw axis is a problem, as is the case with many swept-wing aircraft.
Outer Loop Control
Few transport aircraft fitted with autopilots use them just for auto stabilisation. Instead, the autopilot is supplied with inputs from external sources, which direct it to control the aircraft laterally to maintain a pre-selected heading or
course and to control it longitudinally to maintain a given altitude or vertical speed. Once this is done, it is a relatively straightforward progression to supply inputs, via the flight management system, from the aircraft’s navigation systems to direct the autopilot so that it flies the aircraft along a preselected lateral and vertical flight path. These external control inputs are known as outer loop controls.
AUTOPILOT INNER & OUTER CONTROL LOOPS
Listed below are some of the sources of outer loop control signals supplied to the autopilot systems of a typical modern large transport aircraft:
Central Air Data Computer (CADC) : Manometric data of airspeed, pressure altitude, vertical speed and mach number are supplied by the CADC for control of the aircraft in pitch to select and maintain a given airspeed (airspeed select and hold), select and maintain a given altitude (altitude select and hold) or maintain a given vertical speed or mach number (V/S or mach hold). These are all pitch modes.
Magnetic Heading Reference System (MHRS) : The aircraft magnetic heading is selected and maintained by control about the roll axis, using inputs from the MHRS. This is a roll mode, known as heading select and hold.
Radio navigation aids : The ADF, VOR and ILS localiser receivers provide signals for lateral guidance of the aircraft. In these modes of operation the autopilot uses control about the roll axis to achieve a required course and they are consequently roll modes. The ILS glideslope receiver provides signals for vertical navigation, through control about the pitch axis, during approach and landing phases. This is, of course, a pitch mode of operation.
Flight management system (FMS) : The lateral and vertical information programmed into the flight management computer, when supplied to the autopilot systems, enables them to navigate the aircraft vertically and laterally along a predetermined flight path. Through this source many refinements are possible, such as pitch trim to compensate for fuel consumption to name but one. In this mode of operation (LNAV and VNAV) the FMS will be coupled to the roll, pitch and (where fitted) yaw channels of the autopilot.
Control Surface Actuation
Autopilot operation of the aircraft flying controls may be by hydraulic power-operated actuators, as is usually the case in large aircraft, or by servomotors otherwise known as servo-actuators. These are electrically driven mechanical, hydraulic or pneumatic machines that move the control surfaces by linear or rotary motion. In most cases the same servomotors operate the control surfaces whether the signal controlling them is from the autopilot or from the human pilot’s controls.
In some aircraft control systems the servomotors are connected in series with the pilot’s controls. In this case, when the autopilot is engaged and the flying control surfaces are moved, there is no corresponding movement of the pilot’s controls on the flight deck.
Alternatively, the servomotors may be connected in parallel with the pilot’s controls and autopilot-initiated movement of the control surfaces will be mirrored by the flight deck controls
The stress loads imparted to the airframe when the control surfaces of a large aircraft are deflected can be enormous. Consequently, large or rapid deflections can conceivably induce structural loads beyond the design limitations of the aircraft, resulting in serious structural damage. In manual control of flight this danger is largely averted by artificial ‘feel’ devices, which warn the pilot if excessive control deflection is attempted. Since the autopilot system has no such ‘feel’ it is necessary to introduce safeguards into the control surface operation to prevent overloading of the airframe.
Torque limiting devices are inserted in the drive between servo motor and control surface, which will either slip or disengage if the torque required to achieve the attempted rate of deflection exceeds a preset limit. This device has the added benefit of preventing a servo motor runaway (un-commanded operation of the servomotor) from more than slight deflection of the associated control surface, before disengagement occurs. The torque limiter typically comprises a spring-loaded coupling and friction clutch.
Before the autopilot is engaged, and control of the aeroplane is transferred from manual to automatic, it is important that a number of conditions are satisfied to ensure that the changeover occurs without hazard to flight.
For example, the trim of the aircraft must be set by the pilot to avoid any possibility of sudden attitude change when the autopilot is engaged. Similarly, all power supplies to the autopilot system must be operational and a host of operational parameters must be met.
To ensure that it is impossible to engage the autopilot until all requirements are satisfied a system of inter- locks is interposed between the autopilot engage switch and its electrical supply. These interlocks take the form of relays and switches that only close when parameters are satisfactory. Since they are connected in series, they must all close before the autopilot can be engaged
Mention has been made of the simplest form of manual manoeuvring input to the autopilot, in the form of a rotary knob used by the pilot to bias the inner loops to change the roll or pitch attitude of the aircraft. This type of input control may still be found on the automatic flight system control panel of many aircraft types. In modern transport aircraft, however, it is much more common to apply roll and pitch manoeuvring inputs to the autopilot by means of the control yoke.
There are two methods of achieving this, known as
Control Wheel Steering (CWS)
Touch Control Steering (TCS)
Control Wheel Steering (CWS)
When the autopilot is engaged the pilots can override it, without disengaging, by applying normal manoeuvring force to the control wheel or column. Upon release of the control wheel the autopilot will hold the aircraft at its new attitude and in some cases, if the bank angle is less than 5°, roll the aircraft wings level and hold the new heading until a new automatic flight mode is set on the control panel.
Touch Control Steering (TCS)
With this system a thumb switch on the control column is depressed to disengage the autopilot whilst the pilot manoeuvres the aircraft. When the thumb switch is released the autopilot re-engages to hold the aircraft at its new attitude until an automatic flight mode is reselected.
Automatic Flight System
A typical automatic flight system for a passenger transport aircraft in the medium to large range is made up of a number of component systems.
the flight management system,
the flight director system and
an autothrottle system.
The role of the autothrottle system is to maintain selected EPR and N1 conditions at specific flight phases as directed by the flight management computer or as set by the pilots on the Autopilot Flight Director System (AFDS) mode control panel. The current automatic flight and autothrottle status is displayed on the EFIS ADI and HSI display screens.
The AFDS mode control panel is usually situated on the cockpit coaming beneath the windscreen and its function is to provide the pilots with control of the autopilot systems, the flight director, autothrottle settings and altitude alert settings. The design of the AFDS and its control panel will clearly vary according to the size and performance of the aircraft in question.
The Control Panel
The system uses two independent flight control computers that, in automatic flight, supply pitch and roll commands to the inner loops of the autopilot systems. In manual flight control the computers position the command bars on the captains and first officer’s ADI displays. Each pilot has a flight director selector switch; when switched on, the ADI command bars will appear in certain command modes; when switched off, the command bars will retract out of view. The various mode selector push button switches are depressed for selection and will illuminate to indicate mode selection. Depressing an illuminated switch will deselect that mode. The system will only accept a new mode selection provided that it does not conflict with the mode(s) currently in operation.
Engagement and disengagement of the autopilots is made with paddle switches, one for each autopilot. The paddles have three positions; OFF disengages the respective autopilot, labelled A and B, CWS engages the autopilot but control of flight is by operation of the control wheel and column and CMD is the position for full automatic flight control, enabling all the command modes and CWS operation as required.
In all flight phases other than approach (APP) only one autopilot may be engaged at any one time, but approach mode requires both autopilots to be engaged for a fully automatic landing. Command modes may only be armed or selected provided that at least one of the engage paddles is set to CMD and one or both flight directors are switched on. An armed mode is one that has been selected, but will only engage when certain parameters are met.
AUTOPILOT CONTROL PANEL
Autopilot Command Modes
The autopilot command modes of the system are as follows
Vertical navigation (VNAV) mode
When this selector is depressed the flight management computer commands the AFDS pitch control and autothrottle to follow the selected vertical flight profile programmed into the Flight Management System (FMS). The programmed climb and descent rates, cruise altitudes, speeds and height limitations will be followed through automatic selection of pitch attitude and thrust. With VNAV selected the EFIS ADI will display VNAV PTH or VNAV SPD, depending upon the phase of the planned flight and SPD, N1, RETARD or ARM for the current autothrottle mode.
Lateral navigation (LNAV) mode
Engagement of LNAV mode causes the flight management computer to command the AFDS roll control to intercept and track the lateral route programmed into the Flight Management System (FMS) from waypoint to waypoint. The programme includes all flight procedures such as SIDs, STARS and ILS approach. LNAV will only engage provided that there is a flight path programmed into the Flight Management Computer (FMC). It will automatically disengage if the planned track is not intercepted within certain criteria or if the HDG SEL push button is depressed.
With N1 selected the autothrottle system positions the thrust levers to maintain whatever limiting rpm is set on the Flight Management Computer (FMC) for the current phase of flight.
With this mode selected the autothrottle system positions the thrust levers to maintain the speed selected with the rotary speed select knob and displayed on the AFDS control panel. The autothrottle system will ensure that the selected speed is achieved without exceeding N1 limits and will equalise N1 on both engines provided that it can do so without exceeding 8° difference of thrust lever position.
Level Change (LVL CHG) mode
In this mode automatic control of pitch and thrust is co-ordinated for climb or descent to a preselected altitude at preselected airspeed. Before engaging LVL CHG a new altitude is selected with the rotary altitude select knob on the AFDS control panel and this is displayed digitally in the appropriate window on the panel.
Heading select (HDG SEL) mode
A selected heading is made by rotating the heading select knob on the AFDS control panel and is displayed digitally in the HDG window. Depressing the HDG SEL push button will send a roll command to the autopilot to intercept and hold the selected heading. The bank angle during the turn can be controlled with the bank angle select knob, which forms the outer perimeter of the heading select knob.
Approach (APP) mode
With approach mode selected the AFDS is armed to capture and hold the ILS localiser and glide-slope. Only when this mode is armed is it possible to engage both autopilots; at any other time moving one autopilot paddle to CMD will automatically disengage the other. To meet the requirements of a fail passive control system, both autopilots must be engaged for completion of a fully automatic landing sequence. In this mode the AFDS will command the autopilots through the ILS descent, landing flare, touchdown and roll-out phases.
Take-off/go-around (TO/GA) mode
The go-around mode is automatically armed when FLARE ARMED is annunciated on the flight mode annunciator and/or EFIS display. Depressing the TO/GA selector push button under these circumstances will engage go-around mode, where-upon the flight director will command a 15° pitch up attitude for a climb on present track to a radio altitude of 400 ft. The autothrottle system will simultaneously command the thrust levers to advance for go-around N1 rpm. Once 400 ft radio altitude has been passed, other pitch and roll modes may be engaged; prior to that both autopilots must be disengaged if pitch or roll attitude is to be changed.
Altitude hold (ALT HOLD) mode
Selection of altitude hold mode will either maintain the aircraft at the selected altitude or adjust the aircraft’s attitude until the selected altitude is attained, in either case by pitch commands. If a new altitude is selected with ALT HOLD engaged, the select push button will illuminate until the new altitude is reached. Alternatively, the new altitude may be selected first and then ALT HOLD engaged. With ALT HOLD engaged, LVL CHG, V/S and VNAV modes are inhibited.
Vertical speed (V/G) mode
In this mode the flight director provides pitch commands to maintain the selected rate of climb or descent and the autothrottle system adjusts the thrust levers to maintain the selected indicated airspeed. Engagement of V/S mode is annunciated on EFIS and/or the flight mode annunciator and the present vertical speed is displayed on the control panel, prefixed by + or - to indicate rate of climb or descent, respectively. The desired vertical speed is set by rotation of a thumbwheel on the mode control panel.
Automatic Landing (Autoland)
For an automatic flight control system to be capable of automatic landing it must meet certain criteria. It must contain a minimum of two independent autopilot systems and, in addition, it must satisfy the following safety requirements:
The response of the system must be such that there will be no deviation from the flight path in the event of external disturbance such as turbulence or windshear.
Control system faults must be indicated to the pilot as a warning or alert.
Control system failures must not cause the aircraft to deviate from the flight path.
The flight control system must have sufficient control authority to ensure accurate maintenance of the flight path.
The effect of a servomotor runaway must be limited, such that safe recovery by the pilot is not jeopardised.
The automatic flight control system must not prevent completion of the intended landing manoeuvre in the event of a system failure.
The above criteria are met by incorporating redundancy in the flight control system through duplication or triplication of the autopilot systems, so that a single failure within the system has a minimal effect on the overall aircraft performance during approach and landing. Depending upon the degree of redundancy, the autoland system is classified as being either a fail passive (fail soft) system or a fail operational (fail active) system.
Fail Passive System
An automatic flight system is considered to be fail passive if there is no significant deviation from the flight path, or out-of-trim condition, following a failure within the system, but the landing cannot be completed under automatic control. In simple terms it means that, if one of the autopilots fails, the other will disengage (since two are required for completion of an automatic landing), but there will be nothing to prevent the pilot completing the landing manually. It follows from this that an automatic flight control system incorporating two independent autopilots must be a fail passive system. Furthermore, a self-monitoring system is essential to ensure that both autopilots are in agreement at all times. These are the minimum requirements for the multiple type of control system necessary to meet autoland certification.
In the event of failure of either autopilot or the monitoring system during an automatic approach, the approach will continue on one autopilot, but automatic landing is no longer possible. The flight crew must take over manual control and revert to category 1 minima for landing, either continuing the landing or elecuting go-around procedures at decision height. The single autopilot will disengage automatically at about 350 ft radio altitude.
Fail Operational System
In order for a landing to be completed automatically, following a failure within the system, it follows that there must be at least three independent autopilots and two independent monitoring systems. A single failure in either of these will render the system fail passive, but it still has sufficient redundancy to meet the criteria for completion of an automatic landing. In a fail operational system all the autopilots and self-monitoring systems must be engaged for an automatic approach and landing.
The EFIS display indicates the number of engaged autopilots, with a caption reading LAND 3 indicating three autopilots engaged and a fail operational system, LAND 2 a fail passive system with two autopilots engaged and LAND 1 a passive failure with automatic disengagement pending and completion of the automatic landing impossible. At all other flight phases only one autopilot may be engaged at a time.
In the case of fail operational systems there is a specified alert height, determined by the performance of the aircraft and the automatic landing system. Failure of a redundant autopilot or monitoring system above this radio altitude will result in discontinuance of the automatic landing. If failure occurs below alert height the automatic landing is continued on the remaining autopilot, on the basis that manual reversion is more hazardous at this late phase (typically below decision height for manual completion of landing) than to continue in automatic control.
Automatic Landing Sequence
The radio altitudes for the events during the final stages of the approach to touchdown will vary according to aircraft size and performance, but the sequence is typical for most aircraft types is similar to the one discussed here.
During the descent from the cruise, approach mode is selected by depressing the APP pushbutton and this arms the off-line autopilots; the second in the case of a fail passive system and the second and third in the case of a fail operational system. At the same time the ILS glideslope and localiser channels become the armed pitch and roll modes.
The radio altimeter becomes effective at, typically, 2500 ft agl and provides all height measurements for the automatic flight control system from then until touchdown. At 1500 ft radio altitude, provided that the localiser and glideslope beams have been captured, the off-line autopilots engage and LAND 2 or LAND 3 is displayed on the autoland status annunciation, depending on the number of engaged systems.
The aircraft continues to be flown by one autopilot, with the remainder performing a comparative function, overseen by the monitoring system. If these sequences have been satisfactorily completed, FLARE mode becomes armed and the glideslope and localiser beams become the engaged pitch and roll command modes, maintaining the aircraft on the glidepath centre line.
When the aircraft has descended to 330 ft radio altitude, the AFCS commands a nose-up trim adjustment, with pitch control being maintained through the elevators. When the main landing gear is 45 ft above ground level, as measured by the radio altimeter and adjusted to take account of the height difference between the radio altimeter transceiver and the main gear, FLARE mode engages and provides pitch commands.
Roll commands are still from the localiser, to keep the aircraft on the centre line of the glidepath. The aircraft now follows a 2 ft per second descent path, rather than the glideslope beam, and the autothrottle system begins retarding the thrust levers to control airspeed for the touchdown.
First prior to touchdown, at about 5 ft gear altitude, flare mode disengages and touchdown and roll-out modes engage. At approximately 1 ft gear altitude the AFCS commands a decrease in pitch attitude to 2° nose-up and, at touchdown, the elevators are adjusted to lower the nose and bring the nose wheels into contact with the runway. Selection of reverse thrust by the pilot disengages the autothrottle system, but the AFCS remains in control of the roll-out until disengaged by the flight crew.
Automatic Thrust Control (Auto Throttle)
The autothrottle system receives its commands from an autothrottle computer, which is linked to the flight management and flight control computers and operates the thrust levers through servo-actuators. Its function is to control the thrust in terms of Engine Pressure Ratio (EPR), HP spool rpm (N1) or the aircraft’s flight speed. Its primary function is to operate in conjunction with the Automatic Flight Control System (AFCS) in its VNAV and approach modes, to attain a required airspeed and to maintain the programmed vertical flight path. The autothrottle system is armed by operation of a switch on the mode control panel of the automatic flight control system and is controlled through this panel during automatic flight.
AUTOTHROTTLE SIGNAL INTERFACING
The autothrottle system operates in one of three possible modes:
Before commencing the take-off the flight management system is engaged and its computer supplies the N1 limits for each stage of the flight profile, together with a selected, or ‘target’, N1 rpm. These values are displayed as markers on the N1 indicators of the engine displays. Switching the autothrottle engage switch on the mode control panel to ARM will arm the autothrottle system for take-off and this will be annunciated on the EFIS, or other, display.
The thrust lever servo-actuators are engaged by pressing switches mounted on the thrust levers, known as take-off/go-around switches. Once this has been done, the servo-actuators advance the thrust levers at a preset rate in order to reach the position for take-off N1 by the time a specific speed has been reached on the take-off roll.
For example, the advance rate for the thrust levers might be 15° per second to ensure all engines have reached take-off N1 before the aircraft has reached a speed of 60 knots. When this target speed has been exceeded by a preset amount, autothrottle movement of the thrust levers is interrupted by a speed detection circuit and the levers are held at their current position, a condition known as throttle hold (THR HOLD).
Should the speed detection circuit fail, a back-up system, activated by the main landing gear ‘squat’ microswitches, will operate to instate throttle hold shortly after the aircraft lifts off.
At a radio altitude of 400 ft the autothrottle system arms to control N1 for the vertical profile of the remainder of the flight and the automatic flight control system takes over control of the autothrottle system
Speed control mode
Speed control mode is selected through the mode control panel of the automatic flight control system, either by the pilot pressing the SPD push button switch or automatically if the system is in other than a speed mode (e.g. VNAV).
In either case, the autothrottle system will command the thrust lever actuators to adjust the levers until the IAS or mach No. selected has been reached and held. The autothrottle system controls airspeed/mach to maximum and minimum safe values, regardless of the selected airspeed/ mach, and it prevents the angle of attack (alpha angle) from exceeding a safe value. Minimum safe airspeed and maximum safe alpha are computed from data received from the flap position sensors and angle of attack sensors.
Under VNAV control mode the autothrottle system begins to retard the thrust levers at the top of descent at, typically, 2° per second until they either reach the idle stop or are arrested by pilot intervention. During retardation a RETARD annunciation appears, followed by ARM when the retarding movement ceases, to indicate that speed mode is armed.
At glideslope capture the AFCS mode changes from VNAV to approach (APP) and the autothrottle engaged mode changes to speed, with the displayed speed being that computed by the flight management system. The gain, or sensitivity, of the autothrottle system is increased for greater precision of speed control during the approach. During the flare manoeuvre the thrust levers are retarded at a rate computed to reach idle in 6 seconds. Immediately after the landing gear micro switches have indicated touchdown, further thrust lever retardation is initiated by the autothrottle system, until it automatically disengages 2 seconds after touchdown.
Depressing a take-off/go-around switch on one of the thrust levers with the autothrottle system engaged and the aircraft below 2000 ft radio altitude will initiate advance of the levers until they reach the position for reduced go- around thrust.
The caption GA will immediately appear on the ADI and the flight management system computer will calculate the full go-around thrust rating determined by the present aircraft all-up weight and the density altitude. A second depression of the thrust lever-mounted TO/GA switch will now advance the thrust levers to increase engine thrust to the full go- around thrust rating. The AFCS will generate the pitch-up and wings level commands necessary to establish the aircraft in the go-around climb-out.
Full Flight Regime Auto System (FFRATS)
This system performs all the functions described above and additionally provides engine overboost protection and selection of variable engine rating. The system monitors demands on the engines made by air conditioning system and anti-icing system air bleeds and adjusts the Engine Pressure Ratio (EPR) limits to suit. Most large modern passenger transport aircraft are fitted with an autothrottle system meeting the FFRATS specifications.
In order to achieve maximum fuel economy and to prolong engine life, advanced aircraft turbine engines utilise electronic engine control systems. Pull-authority electronic engine control systems receive data from the aircraft and engine systems to enable safe and efficient operation of the thrust management system over the entire operating range of the engines.
One aspect of such a system is computation of the optimum and maximum thrust requirement for every condition of flight. The computed total air temperature (TAT) and measured pressure altitude are used to compute the optimum and limiting Engine Pressure Ratio (EPR) for the current flight phase. EPR is the ratio of HP turbine exhaust pressure to LP compressor inlet pressure and has been found to be directly proportional to the thrust delivered by the engine.
The computed EPR for the current flight phase is presented on an indicator on the flight deck, which typically displays TAT, the current flight mode (e.g. take-off, climb, cruise, etc.) and the EPR limit for that mode. The actual EPR, with limit and target markers, continues to be indicated on the engine monitoring display (e.g. EICAS).
The flight mode for which EPR computation is required is selected by the pilot through an EPR limit control panel and this is fed to the EPR computer.
Typical modes for EPR limit computation are
When the automatic flight control system is in use, go-around EPR limit will automatically display as the glideslope is captured. Additionally, the panel may contain thrust rating selector switches, with which the pilot can command the computer to calculate the EPR for specific engine performance ratings. The system incorporates a test function for preflight testing. In the event of system failure or electrical power loss, a warning flag obscures the EPR limit indicator.
Flight Envelope Protection
Every aircraft design is tested mathematically and in flight to determine the limits of pitch, roll, yaw, angle of attack and ‘g’ force that the airframe can withstand in flight without suffering structural damage. These limits then form what is known as the flight envelope for that particular design, within which the aircraft can be safely operated. With a conventionally controlled aircraft it is clearly possible to exceed the limits of the flight envelope by applying excessive control movements.
As a means of eliminating the possibility of exceeding the limits of the flight envelope through human error, the fly-by-wire system of flight control has been developed. With such a system, the pilot’s control demands are transmitted to computers that are programmed to respond with signals to the appropriate flying control servo-actuators which will limit their rate of movement, thus ensuring that the aircraft response remains within the limits of the flight envelope.
In the Airbus series of aircraft, beginning with the A320, the fly-by-wire concept has been developed to the extent that the fly-by-wire computers have complete control over each of the flying control surfaces, in response to pilot demands from a small side-stick type of control. The response of the computerised system to pilot inputs must be the same as in a conventional direct control system, but the nature of the inputs is more complex because the pilot can demand, for example, a rate of pitch or roll instead of a simple control movement. This type of fly-by-wire system is known as an active control system.
Given that there is no provision for reversion to manual control in these aircraft, it is clearly vital that there must be a degree of redundancy in the fly- by-wire control system sufficient to sustain failure of a computer without degradation of aircraft control. This is achieved by employing a number of computers in an active control system, such that no single computer can command a control surface movement without being monitored by at least one other.
The A320 aircraft employs seven computers, connected by a data bus, to control the elevators, ailerons, horizontal stabiliser, spoilers and rudder. Two computers control the elevators, ailerons and the horizontal stabiliser and are known as the elevator/aileron computers (ELAC).
Three computers control the spoilers, elevators and horizontal stabiliser and are known as the spoiler/elevator computers (SEC).
It can be seen that control of the aircraft in pitch and roll is shared between the two computer systems so that a fault in one system will not adversely affect the aircraft control.
A third pair of computers controls the aircraft in yaw, known as the flight augmentation computers (FAC).
Swept-wing aircraft, to a greater or lesser extent, exhibit a tendency to develop an oscillatory motion in flight, following a disturbance, which is a combination of yawing and rolling and is known as ‘Dutch roll’.
In many cases the motion is damped out naturally by the ‘weather cocking’ effect of the vertical stabiliser and the aircraft quickly returns to steady flight.
However, swept-wing aircraft exhibit less natural damping because the yawing motion initiates rolling and, in some cases, the oscillations increase if unchecked, especially at lower flight speeds. The tendency can only be checked by deflection of the rudder and to achieve this manually throughout a long flight would place considerable strain on the pilot.
In aircraft that are susceptible to Dutch roll it is usual to install a yaw damping system that automatically applies rudder deflection to control the yawing tendency. The system comprises the third (yaw) alis of an autopilot system and can be operated in either automatic or manual flight control.
YAW DAMPER SYSTEM
Motion about the aircraft’s yaw axis is sensed by a rate gyroscope situated in a coupler unit and powered from the aircraft’s 115 V a.c. electrical system. Output signals from the yaw rate gyro are amplified and filtered to remove frequencies not associated with Dutch roll, and transmitted to an hydraulic transfer valve in the rudder power control unit (PCU). Movement of this valve directs hydraulic pressure to the yaw damper actuator.
The resultant movement of a piston in the yaw damper actuator operates a control valve in the main rudder actuator, which moves the rudder in the required direction to correct the yawing tendency sensed by the rate gyro. The yaw damper piston motion is sensed by a transducer, known as a linear voltage displacement transmitter (LVDT), and fed back to the gyro unit. when the actuator piston has moved by the amount demanded, this feedback of rudder position cancels the gyro output and rudder movement is arrested.
When the Dutch roll oscillations have ceased, the LVDT signal is integrated in the rate gyro coupler unit, to produce an output signal returning the rudder to its neutral, centralised position.
Yaw Damper Indicator
On many aircraft equipped with a yaw damping system the operation of the yaw damper is indicated on the EADI in conjunction with the rate of turn indicator. This receives a signal from the yaw rate gyro. Whenever the gyro precesses, the signal causes the rate of turn indicator to move away from its neutral position. Rudder movement is displayed on a control position indicator.
Pilot operation of the rudder is by direct linkage to the main rudder actuator and is therefore independent of the yaw damping system. Rudder movements commanded by the yaw damping system are not transmitted back to the rudder pedals.
Automatic Pitch Trim
In an aircraft equipped with a movable horizontal stabiliser (trimmable stabiliser) and elevator for pitch control, pitch trim is normally adjusted by first moving the elevators, followed by trimming the horizontal stabiliser until the elevator is returned to the neutral, centralised position.
The normal action of an autopilot system in compensating for an out of trim condition in pitch is to move the elevators until the condition is corrected.
The disadvantage of this system is that, once the elevators are deflected, the amount of remaining movement in that direction is limited and control authority in pitch is reduced. Furthermore, with the elevators deflected from their centralised position, drag is increased, with the obvious adverse effects upon fuel economy and, ultimately, range and endurance.
Consequently, it is not uncommon for aircraft with the stabiliser/elevator configuration to incorporate a system additional to the automatic flight system, which will automatically adjust the horizontal stabiliser until the elevators are restored to the neutral position. Such a system is known as an automatic stabiliser trim (AUTO STAB TPEM) system and it is usually engaged automatically with autopilot engagement. It is a requirement of autopilot engagement that the automatic stabiliser trim system must be operational.
The degree of elevator deflection necessary will depend on airspeed and the automatic stabiliser trim controls incorporate a feel unit which adjusts the trimming signal according to sensed dynamic pressure.
Pitch commands from the autopilot or from manual inputs are sent to the powered control unit, which deflects the elevators through a screwjack, either up or down depending upon the pitch attitude change required. The automatic trim system will then move the horizontal stabiliser, through the trim actuator and screwjack to apply the nose-up or nose-down trim adjustment initially required. As the stabiliser takes up its new position, its motion is mechanically transmitted to a feel and centring unit and a neutral shift sensor. The deflection of the stabiliser removes the need for elevator deflection and the neutral shift sensor sends a feedback signal to the elevator PCU, removing elevator deflection as stabiliser deflection increases, until the elevator and stabiliser are centralised.
AUTOMATIC PITCH TRIM UNIT
In aircraft with a filed horizontal stabiliser, the pitch trim is achieved by means of elevator trim tabs, which are deflected to assist the elevators, to relieve the aerodynamic loads and some of the drag created by elevator deflection.
Automatic pitch trim control is accomplished by means of a separate elevator trim tab servo-actuator coupled to the trim tabs and working in parallel with the elevator servo-actuator.
AUTOMATIC PITCH TRIM SYSTEM
A sliding bar on a mounting attached to the airframe is connected to a capstan, positioned between the elevator ‘up’ and ‘down’ control cables. The cables are lightly tensioned by pulleys so that, when the elevator is in the neutral (streamlined) position, the capstan and bar are centralised between the cables. An electrical contact attached to the sliding bar is supplied from the aircraft’s d.c. bus bar.
Adjustable contacts filed to the mounting are connected to the ‘up’ and ‘down’ field windings of a reversible d.c. motor, which is the trim tab servo-motor.
Let us suppose that the autopilot has demanded a nose-down pitch. The elevator actuator will deflect the elevator down, through the control cables, tensioning the down cable and relieving tension on the up cable. The difference in cable tension will force the sliding bar upward and electrical contact will be made with the trim tab ‘down’ line, supplying the ‘down’ field coil of the trim tab servo-motor and driving the tab down to reduce the aerodynamic force on the elevator.
As the load on the elevator decreases, the tension of the elevator cables will once again equalise and the sliding bar will return to the centralised position, cutting off supply to the trim tab servo- motor.
Automatic pitch trim systems normally include warnings and alerts in the event of system failure. These typically take the form of warning lights or captions and may include an aural alert should a runaway condition, resulting in excessive trim input, occur. In the case of the automatic stabiliser trim system, there is always a trim indicator on the flight deck