Auto Pilot System

Introduction

The autopilot is the central component of every automatic flight system.

This auto-stabilization mechanism is designed to keep the aircraft in stable flight along one or more of its three axes:

Roll
Pitch
Yaw.

Initial autopilot were primarily created to ensure level flight by managing the aircraft's roll along its longitudinal axis, and similar systems are still utilized today in certain light aircraft. This type of autopilot is referred to as a single-axis or single-channel autopilot, offering lateral stabilization.

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

The single-axis lateral stabilizing autopilot uses a closed-loop control system to maintain aircraft roll by adjusting the ailerons through a servomotor. A rate gyro detects roll deviations and sends an error signal to a controller that compares this with the programmed 'w level' condition. The controller generates a correcting signal to adjust the ailerons, bringing the wings back to level. This system can only maintain wings level, lacking directional information, making it an inner loop control system.

SINGLE AXIS SYSTEM

The autopilot controller's 'wings level' reference is typically adjustable by the, allowing them to set the desired roll attitude using a knob.

The controller then directs the servo motor to adjust the ailerons and achieve the required bank angle. This autopilot loop maintains that angle until the pilot makes further adjustments.

Additionally, a second channel can provide automatic pitch control, maintaining a preselected attitude for level flight, climb, or descent. A two-axis autopilot, which includes both roll and pitch channels, is common, while larger aircraft may also feature a yaw channel for enhanced stability.

Some transport aircraft with autopilots use them primarily auto stabilization.

The autopilot receives external inputs to maintain a selected heading and altitude.

This allows easy integration of navigation system inputs to guide the autopilot along chosen flight path, referred to as outer loop controls.

AUTOPILOT INNER & OUTER CONTROL LOOPS

Here are some sources of outer loop control signals for autopilot systems in modern large transport aircraft:

Central Air Data Computer (CADC): Provides manometric for pitch control, enabling airspeed, altitude, vertical speed, and mach number selection and maintenance.
Magnetic Heading Reference System (MHRS): Supplies inputs for roll control to maintain the aircraft's magnetic heading.
Radio Navigation Aids: ADF, VOR, and ILS provide lateral guidance via roll control, while the ILS glideslope receiver aids vertical navigation during approach and.
Flight Management System (FMS): Supplies programmed lateral and vertical information for navigating along a predetermined flight path, incorporating refinements like pitch trim for fuel compensation.

Control Surface Actuation

Autopilot aircraft controls can operate via hydraulic actuators or servo-actuators, which are electrically driven devices that move control surfaces. Typically, the same servomotors respond to signals from either the autopilot or the pilot. In some systems, servomotors are in series with the pilot's controls, meaning pilot controls remain stationary when the autopilot moves the surfaces. In other systems, they are in parallel, allowing autopilot movements to be reflected in the pilot's controls.

Torque Limiters

Deflecting the control surfaces of a large aircraft can impose significant stress on the airframe, potentially exceeding design limits and causing serious damage. While manual flight control includes artificial 'feel' devices to warn pilots against excessive deflection, autopilot systems this feature. To prevent airframe overload, torque limiting devices are installed between the servo motor and control surface, slipping or disengaging if the deflection torque exceeds a preset limit. This mechanism also prevents-commanded servo motor operation by disengaging before significant surface deflection occurs.

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Autopilot engagement/Synchronisation

Before engaging the autopilot, several conditions must be met to ensure a safe transition from manual to automatic. The pilot must set the aircraft's trim to prevent sudden attitude changes, and all power supplies to the autopilot system must be operational Interlocks, including relays and switches, ensure that the autopilot can only be activated when all parameters are satisfactory, as they must all close in series for engagement.

Manual inputs

The simplest manual input to the autopilot is a rotary knob for adjusting roll or pitch attitude, still found on many aircraft's flight control panels.

However, in modern transport aircraft, roll and pitch inputs are more commonly made using the control yoke, through two methods:

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.

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.

These include

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

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.

N1 mode

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.

Level Change 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.

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.

Vertical Speed (V/S) 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.

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.

Speed Mode

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.

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

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.

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.

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.

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.

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.

AUTOLAND SEQUENCE

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

Operating Modes

The autothrottle system operates in one of three possible modes:

Take-Off,
Speed Control
Go-Around.

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 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.

Thrust Computation

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

Climb (CLB)
Economy (CON)
Cruise (CRZ)
Top-of-Descent (TOD)
Go-Around (GA)

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).

Yaw damper

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 trim tab pitch trim system.jpg

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