chapter 06 - flight controls

8/13/2019 Chapter 06 - Flight Controls

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WILJAM FLIGHT TRAINING

Chapter 6.

Flying Controls

Introduction

An aircraft can be rotated in flight about any one, or a combination of its three axes. Theseaxes act at right angles to each other, and all pass through the aircrafts centre of gravity (Fig. 6.1).

AXIS OF ROLL(LONGITUDINAL)

AXIS OF PITCH(LATERAL)

AXIS OF YAW(NORMAL)

FIG. 6.1

Movement about the lateral axis is known as pitch , movement about the longitudinal axis isknown as roll and movement about the normal axis is known as yaw . This is achieved via aprimary flying control system , which in its basic form consists of moveable control surfaceslinked by a series of cables and rods to controls in the cockpit (Fig. 6.2).

CONTROLWHEEL

RUDDER

PEDALS

AILERON

ELEVATOR

RUDDER

FIG. 6.2

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6-2

The primary control surf aces are the elevators , the ailerons and the rudder . These surfacesare hinged at the trailing edges of the main surfaces and are used to manoeuvre the aircraftabout its three axes, producing both primary and secondary effects.

Elevators

The primary effect of elevators is to provide pitch control about the lateral axis (Fig. 6.3).

Pushing the control column forwards causes the elevator to move downwards. This producesan aerodynamic force acting on the tailplane in an upward direction causing the aircraft to pitchnose-down. Pulling the control column rearwards has the reverse effect and causes the aircraftto pitch nose-up. The elevators produce no real secondary effect on an aircraft, althoughchanges in pitch attitude result in changes in angles of attack, and hence airspeed.

PITCH

LATERAL AXIS

ELEVATOR CONTROL

ELEVATOR UPNOSE UP PITCH

ELEVATOR DOWNNOSE DOWN PITCH

FIG. 6.3

The Stabilator

On some aircraft the tailplane and elevator are combined into one surface, known as astabilator , or an all-moving tailplane (Fig. 6.4).

AERODYNAMIC AERODYNAMICFORCE

FORCE

TAILPLANE

ELEVATORSTABILATOR

FIG. 6.4


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Forward movement of the control column will cause the leading edge of the stabilator to rise,thereby generating a force that causes the tail to rise, and the aircraft's nose to drop. Arearward movement of the control column will have a reverse effect.

The Rudder

The primary effect of the rudder is to provide yaw control about the normal axis (Fig. 6.5).

YAW

NORMAL AXIS

DEFLECTED

RUDDER CONTROL

RUDDER

LFIN

LFIN

FIG. 6.5

If the right rudder pedal is moved forwards the rudder will move to the right. In flight this willproduce an aerodynamic force on the fin and the aircraft will yaw to the right. If the left rudderpedal is moved forwards the reverse action will take place, and the aircraft will yaw to the left.The secondary effect of rudder is roll in the same direction as yaw (Fig. 6.6).

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6-4

YAWOUTER WING

MOVES FASTER MORE LIFT YAW

ROLL

FIG. 6.6

This occurs because the outer wing travels faster than the inner wing, thereby generating morelift.

Ai lerons

The primary effect of ailerons is to provide roll control about the longitudinal axis (Fig. 6.7).

If the control column is moved to the right the ailerons will move in opposite directions by equalamounts; the right aileron will deflect upwards, and the left aileron will deflect downward. Thiswill locally alters the shape of the wing where the ailerons are attached. In flight this willproduce a downward aerodynamic force on the right wing and an upward aerodynamic force onthe left wing, causing the aircraft to roll to the right. If the control column is moved to the left,the reverse effects occur.

ROLL

LONGITUDINAL

AXIS

AILERON RAISEDDECREASED LIFT

AILERON CONTROL

AILERON LOWEREDINCREASED LIFT

FIG. 6.7


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If the control wheel is turned the Elevons on one wing will rise, and on the other wingwill lower, as in the case of conventional ailerons. For example when the control wheelis turned to the right the right Elevons will rise and the left Elevons will lower, causingthe aircraft to roll to the right (Fig. 6.13).

ROLL

LEFT ELEVONS DOWNRIGHT ELEVONS UP

FIG. 6.13

The control systems are also interconnected so that the surfaces can be deflectedsimultaneously to produce combined pitching and rolling moments.

Ruddervators . Ruddervators are fitted on light aircraft having a Vee or Butterfly tail.They combine the function of the rudder and elevators (Fig. 6.14).

FIG. 6.14

RUDDERVATORS

The Ruddervators are operated using conventional control system inputs, by way of thecontrol column and rudder pedals. When functioning as elevators the Ruddervatorsmove in unison in the same direction by equal amounts. For example pulling the controlcolumn rearwards causes the Ruddervators to move up, and the aircraft to pitch nose-up (Fig. .15).

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Aerodynamic Balance

When the control surfaces are deflected, the product of the aerodynamic force acting throughthe centre of pressure of the surface and its distance from the hinge-line will produce anopposing moment (Fig. 6.17).

AERODYNAMICFORCE

FHINGE-LINE

HINGEMOMENT

HINGE MOMENT = FX

C OF P

X

FIG. 6.17

This is known as the hinge moment of the control surface and its magnitude determines theamount of effort required by the pilot to maintain its position i.e., stick force . Stick force alsodepends on the method by which the control column is linked to the control surface. The ratioof stick movement to control surface deflection is known as stick-gearing (Fig. 6.18).

STICK

CONTROL COLUMN

MOVEMENT CONTROL SURFACE DEFLECTION

STICK GEARING =STICK MOVEMENT

CONTROL SURFACE DEFLECTION

FIG. 6.18

If the stick forces are high some form of assistance will be needed to help move the controlsurface, and if the stick forces are too light the surface must be artificially loaded to increase theopposing moment. To achieve the necessary stick forces the control surfaces areaerodynamically balanced using one or more of the following methods:

Inset Hinge. With this method the hinge-line is placed inside the control surface. Thisreduces the length of the moment arm and therefore the size of the hinge moment, thusreducing the overall stick force (Fig. 6.19).

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AERODYNAMICFORCE

FHINGE-LINE

HINGEMOMENT

HINGE MOMENT = FX

C OF P

REDUCEDX

FIG. 6.19

The amount of inset is normally limited to 20 - 25% of the chord length to ensure thatthe centre of pressure does not move in front of the hinge-line at high deflection angles.

If the centre of pressure is allowed to move ahead of the hinge line the resulting hingemoment will no longer oppose the movement of the control surface, but will insteadassist it (Fig. 6.20).

X

F

HINGE-LINE

C OF P

ASSISTING

OVERBALANCE

OVERBALANCED CONTROLCP MOVES AHEAD OF HINGE LINEMOMENT

FIG. 6.20

This is known as control surface overbalance , and is detected as a decrease insteadof an increase in the progressive stick forces required to move the control surfacethrough a given deflection angle. In this condition the control surface will automaticallymove towards full deflection and to stop this the control input should be reversed. Thisis known as control reversibility .

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Tabs

Unlike the previous methods of aerodynamic balance tabs are small hinged surfaces formingpart of the primary control surface. In its basic form the pilot does not directly control the tab,but its deflection angle is changed automatically whenever the main control surface is moved.These tabs are used to partially balance the aerodynamic load acting on a control surface,thereby reducing the overall stick force.

Balance Tabs. These tabs are sometimes incorporated as part of the elevator onconventional tailplanes. They are mechanically linked to the tailplane by a linkage thatcauses them to move in the opposite direction to the control surface (Fig. 6.23).

For example if the control column is moved forwards the elevator will move downwardsand the balance tab will move upwards. The resultant aerodynamic force acting on thetab will produce a balancing moment, and reduce the overall st ick f orce .

MAIN AERODYNAMICFORCE

TO CONTROL COLUMN

X

Y

SMALL AERODYNAMICFORCE THAT REDUCES

HINGE MOMENT

FIG. 6.23

Anti -Balance Tabs . These tabs are fitted on aircraft, when it is necessary to increasethe stick force . This is because small movements of a control surface can producelarge aerodynamic loads, and may lead to over control (Fig. 6.24).

MAIN AERODYNAMICFORCE

X

YCONTROL

INPUT

SMALL AERODYNAMIC FORCE INCREASES

HINGE MOMENT

FIG. 6.24

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tab will move by way of the pivot point in the opposite direction to the control surface,providing the necessary aerodynamic assistance.

Mass Balance

During flight the main control surfaces may suffer from vibration . This condition is betterknown as flutter . It is caused by the combined effects of changes in the pressure distributionaround the control surface with changing angles of attack (aerodynamic forces) , and theforces due to the elastic nature of the aircraft structure (aeroelastic forces) itself. If theseforces become coincident, or act in phase with each other the resultant oscillations will quicklyincrease in amplitude, and if left unchecked may ultimately lead to structural failure.

To help eliminate flutter in flight all manually operated control surfaces are generally massbalanced. This is achieved by attaching weights forward of the hinge-line so that the centreof gravity acts through the hinge-line, thus altering the period of vibration and the liability to

flutter. These additional weights are usually stored internally along the leading edge of thecontrol surface, inside the horn balance, or on an arm attached to the surface (Fig. 6.26).

BALANCE WEIGHT

HINGE-LINE

ORIGINAL C OF G

NEW C OF GBALANCE WEIGHT

FIG. 6.26

Powered Flying Controls

Powered flying controls are used on most transport category aircraft to provide assistance inmoving the primary and secondary control surfaces against the large aerodynamic loads, whichmay exceed the physical capabilities of the pilot at high airspeeds.

The primary flying control surfaces are arranged in the same configuration as on lightaircraft, with ailerons , elevators and a rudder , although some differ because they areadditionally fitted with inboard ailerons (Fig. 6.27).

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AILERONS(ROLL)

ELEVATOR(PITCH)

RUDDER(YAW)

FIG. 6.27

The control surfaces are hydraulically activated and are powered from the aircrafts mainhydraulic systems. Due to the importance of the flying control systems the surfaces are alsonormally powered by at least tw o independent hydraulic systems (Fig. 6.28).

G

GG G

G Y

Y Y Y Y

Y

B

B B B

L AIL G G G

G G

YY

Y Y

Y Y

B

B B BR AIL

SPOILERS

G Y

TRIMMING HORIZONTALSTABILIZER ACTUATOR

HYDRAULIC B BLUE SYSTEM G GREEN SYSTEM Y YELLOW SYSTEM

LEADING + TRAILING EDGE HIGH LIFT DEVICES

L ELEV B G R ELEVY B

G Y B RUDDER

FIG. 6.28

Each system is primarily pressurised in flight by engine driven pumps, or alternatively byelectrically driven pumps, but for emergency purposes are normally backed up by either anelectrical pu mp , or by a Ram Air Turb ine (RAT) (Fig. 6.29).

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6-16

LEFT

ENGINE

ELEC MOTOR PUMP

ENGINE DRIVEN PUMP

RIGHT ENGINE

ELEC MOTOR PUMP

ENGINEDRIVENPUMP

ELEC

MOTORPUMP

ELEC

MOTORPUMP

RAM AIRTURB

FIG. 6.29

Powered Flying Control System l ayout and requirements

A basic powered flying control system comprises of the following components:-

A control input system

A power control unit

An artificial feel system

Irrespective of their design all powered flying control systems are regulated by the Joint Ai rworth iness Requirement s (JARs) , and must comply with the following standards:-

Sense. The aircraft must move in the direction signified by the control input, e.g. controlcolumn back, pitch nose-up.

Rigidity. The control system must be strong enough to withstand any operating loadswithout excessive distortion, e.g. airloads on the control surfaces (irreversibility).

Stability. The control surfaces must remain where selected by the pilot and must notbe affected by signals which are not self initiated, e.g. vibration and aerodynamic loads.

Sensitivity. There must be immediate response at the control surfaces to the pilotsinput signals.

Safety. Passengers, cargo and loose articles must safeguard the control systemagainst jamming, chafing, and interference. Guards must therefore be fitted whereappropriate to provide the necessary protection.

Fail-Safe. The control system must be duplicated or be capable of manual operation inthe event of hydraulic power failure.


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The Power Control Unit (PCU )

The power control unit is the main component in a power operated control system and providesall of the force necessary to move a control surface, with the pilot only having to supply a smallforce to operate a servo valve (Fig. 6.32).

CONTROLCOLUMN

RETURNFLUID OUT

HYDRAULICPRESSURE

IN RETURNFLUIDOUT SERVO VALVE

(NEUTRAL POSITION)

PORT

AIRCRAFTSTRUCTURE

JACK RAM PISTON JACK BODY

CONTROL SURFACE

PIVOT

FIG. 6.32

It consists of a jack ram/piston arrangement, which is fixed to the aircraft structure ,hydraulic fluid , inlet/outlet ports, and a jack body . These parts form a hydraulic actuator ,which is controlled by a servo (control) valve and is connected via a control run to the flightdeck controls. When the valve is displaced in either direction from its neutral position it allowshydraulic fluid under pressure to pass to one side of the piston, and opens a return path fromthe other side. For example a rearward movement of the control column will cause the servovalve to move to the left (Fig. 6.33).

CONTROLCOLUMN

RETURNFLUID OUT

HYDRAULICPRESSURE

IN RETURNFLUIDOUT

SERVO VALVE(DISPLACED FROM

AIRCRAFTSTRUCTURE

PISTON JACK BODY

CONTROL SURFACE

MOVEMENT OFJACK BODY

NEUTRAL POSITION)

PRESSURE FLUID

RETURN FLUID

PIVOT

FIG. 6.33

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Since the jack is fixed in position the resulting pressure differential across the piston will causethe jack body to move to the left, which in turn will deflect the control surface upwards via amechanical linkage. The body continues to move until it centralises itself on the servo valve, i.e.returning it to its neutral position (Fig. 6.34).

CONTROLCOLUMN

RETURNFLUID OUT

HYDRAULICPRESSURE

IN RETURNFLUIDOUT

SERVO VALVE

AIRCRAFTSTRUCTURE

PISTON JACK BODY

CONTROL SURFACE

(NEUTRAL POSITION)

TRAPPED (STATIC) FLUID

PIVOT

FIG. 6.34

At this point the hydraulic fluid will be trapped either side of the jack and will form a hydrauliclock . This in turn will maintain the control surface rigidly in its selected position, and it willcontinue to remain so , irrespective of the aerodynamic loads acting on it, until the servovalve is repositioned by further flight deck control inputs. This is alternatively known as anirreversible control system . Conversely if the control column is moved forward the sequenceof operations will be reversed, i.e. if the servo valve moves to the right, the jack body will moveto the right, and the control surface will deflect downwards. Some power control units alsooperate in response to electrical inputs from the Autopilot and Autostabilisation systems whenthey are engaged.

Ar ti fi cial Feel Syst ems

In a manually operated flying control system the aerodynamic loads acting on a controlsurface are fed directly back through the control runs to pr ovide stick f orce or feel on theflight deck controls . The loads thus vary depending on control surface deflection andairspeed. In the case of a power operated flying control system there is however no directlinkage between the control su rface and the fligh t deck cont rols . In fact the only force feltis that associated with the movement of a servo valve, and the effort provided by the pilottherefore bears no direct relationship to the actual loads acting on the control surface.

These loads are alternatively dissipated through the aircraft structure via the body of adedicated power control unit, thereby relieving the pilot of all control loads. Consequently, toprevent over-controlling and overstressing of an aircraft, some form of artificial feel isincorporated in the control system, so that the forces experienced are representative of a

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manually controlled aircraft. A suitable feel unit must therefore be capable of producing anopposing force that varies with airspeed and control surface deflection.

On transport category aircraft the requisite feel forces are provided by; spring or Pitot-static Qfeel units , or in some cases a combination of both. Artificial feel systems normally alsoincorporate a self centring mechanism , so that if the flight deck controls are released they willautomatically return to their neutral position, and will also centralise the control surface.

A Simple Spring Feel Uni t. This is the simplest form of artificial feel unit and isnormally fitted in the operating linkage between the flight deck controls and the powercontrol unit (Fig. 6.35).

CONTROL COLUMN

PIVOT POWER CONTROLUNIT

JACKRAM

SPRING FEEL UNIT

ELEVATOR

FIG. 6.35

It is designed so that any flight deck control movement is firstly made against springtension, so the larger the movement, the greater the opposing spring force. Forexample if the control column is moved rearward the left-hand side of the spring in thefeel unit will be compressed in proportion to the control column movement andsubsequent deflection of the control surface. When the control column is centralisedthe spring unit off-loads itself, thereby centralising the linkage and returning the flyingcontrol surface to its neutral position. This type of feel unit by itself may be adequate atlow airspeeds, but at higher airspeeds greater resistance to flight deck controlmovement is needed to prevent over-stressing the aircraft. This is because the amountof feel only varies in proportion to control surface deflection, and takes no account ofairspeed. On transport category aircraft this type of feel unit is only normally used byitself in aileron control systems.

Q Feel Units. These units like spring feel units are fitted in the operating linkagebetween the flight deck controls and the power control unit (Fig. 6.36)

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PITOT STATIC

TO POWERCONTROL UNIT

Q FEEL UNIT

AIRCRAFTSTRUCTURE

PIVOT

FIG. 6.36

A basic Q feel unit consists of a diaphragm with static pressure acting on one side andpitot pressure on the other, with the difference between the two being dynamicpressure . The unit is also arranged so that movement of the flight deck controls ineither direction deflects the diaphragm against pitot pressure (Fig. 6.37).

SAME LINEAR TRAVELIN SAME DIRECTION

FIG. 6.37

For example if the dynamic pressure increases due to an increase in forward airspeed(IAS), the forces required to move the flight deck controls would similarly increase.Conversely, a reduction in forward airspeed will cause the load on the flight deckcontrols to decrease. This system therefore ensures that the stick forces vary duringflight in proportion to varying loads acting on the control surfaces. These units howevertend to be very large, so the sensing pressures are alternatively used to operate apiston subjected to hydraulic pressure , thereby providing hydraulic Q feel (Fig. 6.38).

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6-23

Trimming Control Systems

The trimming control system is principally designed to reduce the stick forces or controlforces to zero . This allows an aircraft to maintain any yaw, pitch or roll attitude set by the pilotwithout further control input. On small aircraft this system comprises of moveable or fixed auxiliary surfaces. These surfaces are called trim tabs and are normally hinged at the trailingedge of the primary control surfaces (Fig. 6.40). Most light aircraft are fitted with elevator andrudder trim, but aileron trim is normally only fitted on more sophisticated types of aircraft.

RUDDER TRIM TAB

ELEVATOR TRIM TAB

AILERON TRIM TAB

FIG. 6.40

Principle of a Trim Tab

In operation a trim tab creates a hinge moment which exactly balances the hinge momentproduced by a control surface (Fig. 6.41).

b

a

F1

F2

TAB

CONTROL SURFACEF1 a F2= b

FIG. 6.41


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In this condition the control surface remains in its set position without any effort from the pilot,i.e. the control forces are zero.

Moveable Trim Tabs. These trim tabs are normally fitted on elevator and ruddercontrol surfaces. In each case the tabs are connected via a cable and gearing systemto trim wheels in the cockpit. For example consider the operation of an elevator trim tab(Fig. 6.42).

TRIM WHEEL

TRIM TAB

FLAP LEVERELEVATOR TRIM WHEEL

TRIM POSITION INDICATOR

DOWN UP

FIG. 6.42

In this system the trim wheel is mounted to give movement about a lateral axis and isrotated in the natural sense to give the required pitch trim change. A forwardmovement of the trim w heel will produce a nose-down tr im change , and vice versa(Fig. 6.43).

TRIM TAB

ELEVATOR

MOVEMENT OFTRIM WHEEL

MOVEMENT OFTRIM WHEEL

NOSE-DOWN TRIM NOSE-UP TRIM

FIG. 6.43

Notice that the trim tab moves in the opposite direction to the control surface. Inpractice if it is necessary to trim the aircraft in a given pitch attitude the elevator shouldfirstly be moved to produce the desired pitch. The force necessary to maintain this

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pitch is then eliminated by rotating the trim wheel in the same direction as the controlcolumn until the stick loading reduces to zero.

Note: Trim tab deflection reduces the maximum available elevator authority.

Rudder trim is provided in a similar manner to elevator trim, except in this instance thetrim wheel is mounted so that it rotates about a normal axis. For example to providenose right trim the trim wheel is rotated in the clockwise direction and vice versa(Fig. 6.44).

TRIM TAB

RUDDER

NOSE RIGHT TRIM NOSE LEFT TRIM

MOVEMENT OFRUDDER TRIMWHEEL

MOVEMENT OFRUDDER TRIMWHEEL

FIG. 6.44

On some light aircraft the trim tabs are moved electrically instead of mechanically. Inthis case the switch is normally spring-loaded towards the central off position and inorder to reduce the stick forces to zero the switch is operated in a natural sense. Whenthe switch is released it is designed to return to the Off position.

Note : irrespective of the positi oning method, the tab will remain in the same fixedposition relative to the control surface until it is necessary to re-trim the aircraftin a new attitude.

Fixed Trim Tabs. These tabs operate completely independently of the pilot and canonly be adjusted on the ground (Fig. 6.45).

TO CONTROL COLUMN

FIXED TAB ADJUSTABLE

ON THE GROUND

FIG. 6.45

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Their actual setting is determined by flight tests and when they are set to give noresultant stick forces, the trailing position of the control surfaces is governed by theactual deflection of the tab. On light aircraft this type of tab is only normally fitted onailerons to make wing level flight more easily achievable without having to maintain aconstant stick force, i.e., to correct for a wing-low tendency.

Combined Trim/anti-balance Tabs . On some aircraft the tab has a dual function ,and can operate as either a trim tab, or as an anti-balance tab, for example an all-moving tailplane. (Fig. 6.47).

ALL MOVING TAILPLANE

ANTI-BALANCE/TRIM TAB

PIVOT POINT

FIG. 6.47

When used to provide elevator trim the tab is positioned by way of a trim wheel untilzero stick force is achieved.

Trimming of Powered Flying Controls

On aircraft fitted with powered flying controls the position of the control surfaces is not affectedby aerodynamic forces, and is only altered by movement of the appropriate servo valve inresponse to flight deck control inputs. Thus in order to correct for any out of trim condition adevice is fitted which re-positions the neutral setting of the servo valve, and moves the controlsurface to a new neutral position. The flight deck controls similarly take up a new neutralposition in the direction of the required trim. These devices are fitted in the control-inputsystem, and consist of an electrical or mechanical linear actuator, controlled from the flight deck(Fig. 6.48).

SPRING STRUT

TRIM ACTUATOR

TO CONTROLSERVO VALVE

FIG. 6.48

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When a trim actuator is operated it alters the effective length of the input lever to the servovalve, thereby making a selection, and moving the control surface to a new neutral position. Aprotection device in the form of a spring strut is also fitted between the trim actuator and thecontrol input linkage, which normally operates as a fixed member, but should the servo valveseize a spring inside the unit compresses or extends to protect the valve from further damage.

Ai leron and Rudder Trim. On most transport category aircraft aileron and rudder trimis applied through the movement of electrically operated trim switches. These switchesare normally positioned on the centre pedestal and are spring loaded to t heir centraloff position (Fig. 6.49).

AILERON TRIM

6 4 2 0 2 4 6 AILERON TRIM INDICATOR

CONTROL COLUMN

AILERONTRIM

SWITCHES

RUDDER TRIM INDICATOR

RUDDER TRIM SWITCH

LEFT RIGHTWING WING

DOWN DOWN

AILERON

RUDDER TRIM15 10 5 0 5 10 15

NOSE LEFT UNITS NOSE RIGHT

NOSELEFT

NOSE

RIGHT R U D

D

E

R

FIG. 6.49

To provide aileron trim both switches are simultaneously moved in the samedirection to provide system integrity . The amounts of aileron and rudder trim appliedare conventionally displayed on dedicated trim indicators . The rudder trim indicator isnormally sighted on the centre pedestal, whilst aileron trim indicators are normallysighted on each control column.

Variable Incidence Horizontal Stabiliser. On most transport category aircraftvarying the angle of incidence of the horizontal stabiliser provides pitch trim .This has the same effect as that achieved by movement of the elevator, but isaerodynamically more efficient , particularly at high airspeeds, and if the required trimrange is considerable. The angle of incidence of the stabiliser is varied by an actuatorassembly which is positioned near the leading edge and normally operates inconjunction with the elevator, the combination of the two providing the least trim drag(Fig. 6.50).

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6-28

TAILPLANE

ACTUATOR ELEVATOR ACTUATORCONTROL INPUT

FIG. 6.50

The leading edge of the stabiliser is normally moved up or down in flight by anelectrically signalled hydraulic trim motor . This is in response to signals primarily

from the electrical trim switches that are situated on the control column, or alternativelyin response to electrical signals from the Autopilot pitch channel (Fig. 6.51)

AUTO PILOT

AP TRIM

COMPUTER 1

TRIMCOMPUTER 2

MANUALPITCHTRIM

ENGAGED M M

VALVE VALVE VALVE

MOTOR MOTORMOTOR B Y G

SCREWJACK

TRIMMABLE HORIZONTAL STABILISER

ELECTRICPITCH TRIM

SERVO-MOTORS

ELECTRICAL PITCH TRIM

LEGEND MECHANICAL LINKAGE ELECTRICAL LINKAGE

FIG. 6.51

Like the aileron trim system, pitch trim is provided by the simultaneous movement oftwo switches , located on each control wheel. This provides system integrity andprevents pitch trim runaway . The rate of trim varies and is controlled by trim controlmodules. The trim rate decreases with increasing IAS .

In the event of a malfunction in any of these methods pitch trim is alternatively providedby manually positioning the hydraulic servo valves via a series of control cables and


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Some aircraft are alternatively fitted with stabiliser trim levers , eg. Boeing 757. Theseare sited on the centre console, and like trim wheels pass manual pitch trim controlsignals to the pitch trim servo motors (Fig. 6.53).

M

M

STABILISERTRIM

MECHANISM

STABILISER

A/CNOSEUP

A/CNOSEUP

STABILISER TRIMLEVERS

ST A

B

TRIM

ELECTRIC PITCH TRIMSERVO MOTORS

SCREW JACK

FIG. 6.53

To vary the amount of pitch trim, both levers must be moved simultaneouslytogether in the same direction . This increases the overall system integrity, andprevents the possibility of pitch trim runaway . This system is also designed so thatsignals from the trim levers override all other trim inputs, and unlike trim wheels, they

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do not move in response to electrical trim inputs . On aircraft fitted with this facilityall stabiliser trim indications are alternatively displayed on two dedicated indicators ,which are normally sited on the centre pedestal (Fig. 6.54).

NOSENOSE

FLAP

0 2

4

6

8

10

12

14 O

FF A/C

NOSEUP

A/C

DN

S

T

A

B

T

R

I

M

STAB TRIM 02

4

6

8

10

12

14OFF A/C

NOSEUP

A/C

DN

ST

AB

TRI

M

STAB TRIM APLNOSEUP

APLNOSEUP

S T A B

T R I

M

STABILISER TRIM

LEVERS

STABILISER TRIM INDICATORS

FIG. 6.54

SpoilersSpoilers are flap type control surfaces, which are normally, located on the upper surface of thewing, just in front of the trailing edge flaps (Fig. 6.55).

PANELS SHOWNRAISED

FIG. 6.55

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These surfaces are individually hinged at their leading edges and are actuated by hydraulicpower supplied by dedicated power control units . As their name implies the main purposeof the surfaces is to disturb the smooth airflow over the top of the wing, thereby reducing the

wings lifting capability , and when fully extended considerably increasing aircraft drag .The spoilers are positioned so that aircraft pitch trim is not adversely affected by theirdeployment and the surfaces are extended in various sequences depending upon the mode offlight. They are used as:

Roll Spoilers. In this mode the spoilers operate asymmetrically in flight whenever thecontrol wheel is rotated to assist the ailerons in providing roll control , particularly athigh airspeeds. When this occurs the appropriate spoiler servo valves are signalled byway of the aileron operating system and the requisite surfaces are deflected upwards inproportion to the roll input . During a roll the spoilers on the lowering wing aredeflected upwards, so decreasing its overall lifting capability and increasing theaircrafts roll rate, whilst the spoilers on the opposite wing remain retracted (Fig. 6.56).

FIG. 6.56

54

3 21

UP GOING WING SPOILERSREMAIN RETRACTED

AILERON

SPOILERS RAISEDON LOWERING WINGPROPORTIONAL TOROLL INPUT

AILERON

SPOILERRETRACTED

The innermost spoiler on the lowering wing also remains retracted to prevent tail buffet,and degraded pitch control.

On aircraft fitted with two sets of ailerons, (inboard and outboard), as the airspeedincreases the aerodynamic loads on the ailerons tend to twist the wing at the tips,where it is more flexible. To overcome this tendency some aircraft use the technique oflocking the outboard ailerons in the faired or neutral position, and use an inboardaileron/spoiler combination above flap retraction speeds to provide the necessary rollcontrol (Fig. 6.57).

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6-33

INBOARD AILERON

OUTBOARD AILERON

FIG. 6.57

Flight Spoilers. In this mode the spoilers are operated symmetrically about theaircrafts longitudinal axis to slow an aircraft down in flight (speed brakes). Theirdeflection angle is determined from the flight deck in response to movements ofa speedbrake lever , which is typically located on the centre pedestal on transportcategory aircraft (Fig. 6.58).

SPEEDBRAKELEVER

FIG. 6.58


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In this mode the outboard spoilers usually remain retracted to prevent the aircraftpitching nose-up whilst the innermost spoilers are deflected by a lesser amount toprevent tail buffet. There are sometimes occasions in flight when both airbrake and rollcommands occur together. On these occasions both inputs are fed into a complex box,containing a mixture of levers, bell cranks and quadrants, called a spoiler mixer unit(Fig. 6.61).

DOWN

ARMED

FLIGHTDETENT

UP

SPEEDBRAKE

SPOILERMIXERUNIT

LEFT ROLLINPUT

SPEEDBRAKELEVER

SIGNALS TOSPOILER ACTUATORS

FIG. 6.61

This unit sums both inputs and gives a revised output, which in turn varies themovement of the spoilers during an aileron input depending upon the amount ofspeedbrake selected (Fig. 6.62).

RIGHT SPOILERSPARTIALLY RAISED

SPOILERSRETRACTED GREATER INPUTTO LEFT SPOILERS

FIG. 6.62


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Spoilers in this role can normally be used at any airspeed but at increasingly higherairspeeds they are forced down (blowback) progressively.

Ground Spoilers (lift dump). In this mode the spoiler panels on both wingsautomatically rise to their full extension after touchdown to increase an aircraftsrate of retardation (ground spoilers), when certain ground conditions are fulfilled (Fig.6.63).

123456789101112

FIG. 6.63

These conditions are typically:-

Speedbrake lever in the armed position.

Aircraft weight on the undercarriage (through the air/gound sensing system). All thrust levers in their idle positions.

Aircraft wheels ro tating (provides a time delay and ensures the aircraft is on theground).

As the spoilers deploy the speedbrake lever automatically moves to the up position inline with their movement (Fig. 6.64).

DOWN

ARMED

FLIGHTDETENT

UP

SPEED BRAKE

DOWN

ARMED

FLIGHTDETENT

UP

SPEEDBRAKE

FIG. 6.64

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The maximum deflection angles are greater in the ground mode than the flight mode.With the spoilers in their fully extended position approximately 80% of the wing/flap lift isdestroyed and the aerodynamic drag of the aircraft more than doubles. The

subsequent loss of lift causes the aircraft to fully settle on the main undercarriage andincreases its potential braking force. The flaps are also left in their landing configurationbecause of the drag benefits on deceleration. Should any of the thrust levers beadvanced the speedbrake lever automatically moves to the down position , and thespoilers retract.

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chapter 06 - flight controls

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