CONTENTS

Page

Regulations 1 Flying control systems 4 Primary flying controls 8 Control system components 14 Tabs 28 Balancing of controls 37 Ailerons 43

1' Tailplanes-, Fnreplanesn --

Spoilers I I

Flaps I Slats/ slots vortex generators wing fenbes I

I Saw tooth leading edge I Fixed spoilers .

Canard configuration Delta a g e d aircraft

1 1

Butterfly tailplarie Jntegrd @st-locks . -

Control position indication Flying control rigging Rigging equipment

REGULATIONS

This section might be more easily understood if you read (and understand) the rest of this book first and read the book 'PFCUs and Autopilots'. Even if you do read it now it is worth a second read after completion of the two books.

Requirements are published to cover all aspects of airframe design in EASA Certification Specifications CS 25 (large aircraft) and EASA CS 23 (srnaI1 aircraft). Here we will concentrate on the control systems of large aircraft.

Control Surfaces

M u s t meet the airborne and ground gust loads specified for the airframe. Must be fitted with control stops and hinges and must have a factor of safety of 6.67 of the ultimate bearing strength.

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For the lateral, longitudinal, ditectional cbntrol of the control systems - the support structure mqst have a factor the control surface maximum rkment lbad.

I_ - --__ \ \

The sy;stLm must operate easili, smoothljr &d be positive control $tops fitted and should bd design&@ 40 prevent objects. t I I . ,

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I t shoulcl-be-so designed as to:mirilmjse-the possibility of incorrect assembly and must be capable of continued operation if:

I) A single failure of a hydraulic, mechanical or structural element occurs.

2) Dual electrical or hydraulic systems fail. 3) R jam occurs in the system. 4) The system experiences a run-away powered flight control.

With the system operating at 80% specific limit load and on 100% powered load there should be:

a) No jamming. b) No excessive deflection. c) No excessive friction.

Struc turd deflection should not affect the system adversely.

Minimum and maximum forces are specified for the pilot's input and minimum forces specified when pilots are operating in opposition on duaI control systems,

moodull IA-2

Rotary Control Joints

Must have a factor of safety of 3.3 with a push/pull rod system and a factor of safety of 2.0 with a cable system.

Cable System

Cables used on aileron, elevator and rudder must not be smaller than 0.125" (3.1 7mm) diameter.

Tensions must be kept reasonably constant.

Pulleys must be fitted with guards to prevent cable displacement or fouling.

A cable m u s t not change direction more than 3" after passing through a fairlead.

I - --

~~ekified-m.rts of the cable system must have access for inspection. I

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I / , I

Gus t Lo'cks ' I ' I I ' I I I - -

Must bd , , fitted to a specificatioi andprovision made that they:

I not possible. -. -. -

Trim Systems

M u s t meet minimum input force standards with loadings specified for:

a) Trirntabs. b) Balance tabs, c) Servo tabs.

They must operate in the correct sense and be designed to prevent abrupt changes of aircraft trim. Flight deck indications must be provided and the tab must be irreversible unless it is a balance tab.

Stability Augmentation

(This is Active Stability as fitted to some systems of some aircraft). Can be de- activated or overridden by the pilot without affecting safe control of the aircraft. It must be provided with a failure warning system.

Configuration (Config) Warning

An aural warning is provided if, during the take-off run, the aircraft is not correctly configured. The configuration to include:

1 Flap position, 2 ) Slat position. 3 ) Spoiler position. 4) Wheel brake configuration. 5) Tailplane position (tailplane/stabiliser not in the green area).

The warning to be cancelled by either:

a) Changing the incorrectly selected system. b) Abandoning take-off. c) Aircraft rotation. d) Pilot de-activation. -. - - 1 -.

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Wing Flaps and High Lift ~evicds 1

Minimum loads are speciiied wi(h factois Lf for take- landing conditions. I I I -.. - Lift ani drag devices must be skl ctabldby 9 they a f ~ designed t o be used on1 the ground *here provision n h s t be made to prevent inadverteht flight operatidn.

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Flight deck'indication must be-pro~ided'with warnings of asynirnetrical - operation of a symmetrical system reg flgps and slats).

Wing flaps must have a synchronising system strong enough to prevent asymmetric operation occurring with one side completely jammed and the other side under full operating power.

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FLYING CONTROL SYSTEMS

In order to allow the pilot to control the aircraft in the air flying control systems are fitted. These consist of moveable surfaces fitted to the trailing edges of the fin (rudder), tailpIane (elevators) and mainplanes (ailerons). These are called Primary Flying Controls and are connected via a control system to the pilot's controls in the cockpit or flight deck. The controls must be instinctive and work in the correct sense.

Other flying controls include; tabs, moveable tailplanes, spoilers and rarely, moveable outer wings similar to moveable tailplanes.

Note. In some books tailplanes are called stabilisers.

Instinctive

The,flying controls are said to-be instinctive. This means that when the pilot pushes the control column forwadthe aircraft dives or pitches nose down; when bd/she pulls the control column back the aircraft climbs or pitches nose up. ~ h k n the pilot moves the control column, or hand wheel left the aircraft rolls to fhe left; when he/she &oves the control column, or hand wheel right the aircraft rolls to the right. $hen the pilot pushes his/her left foot forward on the rudder bar the aircraft tuds to -the.left; when the right foot is pushed fonvaid the aircraft turns right. a

I

Sense:

The control surfaces should move 5 the correct relationship to the control column or rudder bar. This is called correct SENSE (see Table 1).

Elevators

Hinged t o the trailing edge of the tailplane and connected to the control column. Movement of the elevator gives longitudinal control about the lateral axis. May be interconnected with the tailplane.

Rudder

Hinged to the trailing edge of the fin and connected to the rudder bar. Movement of the rudder gives directional control about the vertical axis.

Ailerons

Hinged surfaces on the trailing edges of the mainplanes. Movement of the ailerons gives lateral control about the longitudinal axis.

On many civil aircraft are interconnected with the spoilers. May also be split into inboard and outboard ailerons with the inboard only operating at high speed.

CONTROL COLUMN

Arrows show the pilot puling the control column backwards and the elevator moving up to cause the alrcran to climb.

CONTROL CABLES PIVOT

PIVOT,' ' \

4 I .- -. -.. - -

I ' -

/-\ - \ , I I< - L,

ELEVATOR - -.

I ,

i i I I I '

I j , 8 ! : ! I

COGRPIT CONT;ROE 1 CONTROL SURFACE M ~ V E M E M T ~ V E M E N ~ u -

Control col~imn pulled Elevator hove& up,

I Control column pushed 1 Elevator moves down.

I I I 1 EFFECT i j

~irflod hits] the control surfat -mid produces a downward force, this is transferred to the tailplane causing the taiI to go dawn and the nose to go up. Similar to the above but in

a down force, which pushes the wing down. The down

forward. Control column or handwheel moved to the left.

I I I going aileron experiences an I I I I upward force therefore 1

Ailerons. Left one up. Right one down.

the opposite direction. The up going aileron (on t h e down going wing) experiences

I to the right. I one down. - I the opposite direction. 1 Control column or handwheel

- - Rudder pedals. Right foot I Rudder to the right. 1 The airflow pushes on t h e 1

Ailerons. Ri&t one up. Left

I forward. I t h e opposite direction. 1

pushing the right wing up.

Similar to the above but in

forward.

Rudder pedals. Left foot

- 5 - rnoodull l A-6

Rudder to the left.

rudder producing a force to the left, this pushes the tail to the left and the nose of t he aircraft to the right. Similar to the above but in

Arrows show pilot's right faot forward and rudder movement to the right and

RUDDER BAR aircraft flying right. CONTROL RUDDER

1 CABLES -

<

PIVOT 4 1

1 1 \ S + PIVOT

Fig. 2 A SIMPLE RUDDER CONTROL SYSTEM SHOWING CORRECT SENSE

Arrows show conhol column movement to the left with the leil (port) alleron moving up and the right (starboard) alleron moving down and the

CONTROL atrcrafl banking keft. RIGHT COLUMN AILERON - 1

1- -

C O N T ~ O L ~ CABLES .

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Fig. 3 A SIMPLEAICERQN CONTROL SYSTEM SHOWING CORRECT SENSE

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A Typical Large ,Commercial ~ircraft - Introduction , - .

The following is a brief description of the flying controls and lift augmentation/ drag systems of a typical large aircraft. For more detailed information you should refer to other sections within this book and refer to t h e book in this series 'PFCUs and Autopilots'. Of course, for specific information on an aircraft's control system you should refer to the aircraft's AMM.

Elevator

Hydraulically powered and controlled by a dual cable control system from the flight deck. The system may have all er any of the following: artificial feel, autopilot servo, stj ck shaker/ stick push, disconnect detents and cable tension regulators.

Tail plane

Usually trimable using a hydraulic or electric motor with standard trim and Mach trim inputs. May have elevator inputs.

Rudder

Hydraulically powered and controlled by a cable system from the flight deck. Usually fitted with yaw dampers, artificial feel, cable tension regulators, autopilot servos and may have speed related range limiting devices.

Ailerons

May be organised to droop for take-off and landing (called flaperons [flaps and ailerons] on the B777) to provide more lift and are interconnected with the spoilers (asymmetric operation) to provide better roll control. Are powered hydraulically a d operated by cables from the flight deck. Will usually have autopilot input, artificial feel and cable tension regulators.

All the above controls wiIl normally have automatic provision to give indication on the flight deck of their position and warning systems in the event of major component failure (Powered Flying Con.@ol Units PFCUs etc)., . - --

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Spoilers I I ,

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~ ~ d r a u l i k a l l ~ powered t o opera the flight or ground mode( ~esi&ekl to create drhg and dump lift when, operated: .._ _-

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Slats , ,

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Fitted to thii leading edge of the rnainpl[mes, are usually hy&hlically powered and symmZScZlly operated to create lift and increase the 'stdifig angle.

Leading Edge or Kruegar Flaps

Often used at locations inboard of the inboard engines and may be pneumatically powered. Symmetrically operated to give the same affect as slats.

Flaps

Usually of the Fowler variable area type. Are symmetrically operated to increase lift (and increase drag).

RUDDER

STARBOARD (RIGHT1 AILERON

OUYBOARD FLAP

w LWMWG EDGE FLAPS i

GROUND SPOILER

FLIGHT SPOILERS

-1 GRWWD SPOILER

i i I PRIMARY FLIGHT CONTROLS

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control Surface f ieory - -.- -

The control surface is said to work by a combination of two theories - the mean camber line theory and the force theory.

a) M e a n Camber Line Theory

When the control surface moves it alters the Mean Camber Line of the main surface to which it is attached. This alters the lift on that surPace and it is caused to move up or down in the airflow.

NEW MEAN CAMBER UNE WITH CONTROL SURFACE DEFLECTED EXTRA LIFT FORCE

CONTROL SURFACE

AEROFOlt 1

CONTROL SURFACE

2

Fig. 5 MEAN CAMBER LINE THEORY

- 8 - moodull IA-9

b) Force Theory

When the surface moves into the airflow it experiences an aerodynamic loading, if it is held in that position then a component of this load (thc vertical component) is transmitted to the main swface of the aircraft which will move in that direction. The horizontal component is drag.

D W G FORCE EFFECTIVE CONTROL \ FORCE ON AIRCRAFT

\TOTAL

Fig. 6 FORCE THEORY USING VECTORS

1 \ ' Movemerit of the pilot's control$ i i transmit'fed to the contrdl surfaces by a system of rods or cables, or a conibinatioh df both. Chains dvir sprockets are

I also used. The system must tranqpit the/'c&trol surface lo<di,ng back to the . .

I I pilot (on hon-powered systems) I 'I . -/' ,.

I 1- . . 1 ----- On some aircraft, such as the 320 and $77, transmission o:f cont/rol signals 4 1 to the (pdwered) control surfaces l s via a cbquterised fly-bb-+ire system. The flight deck contrbi movement istrwsducq'd ihto an electricdl dimd, sent to a computer, digitiskd, processed h a the resuit is an analoguk $&trical signal sent to s i g d . a hydraulically p d ~ k ~ d " F ~ ~ ~ . - to move the aohtrbl surface. More

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of this latcr in module 1 1.

On some aircraft a fly-by-light system is used (also computerised).

These systems are covered in t he book PFCUs and Autopilots.

There are two basic types of mechanical systems:

a) Rod System

Light alloy push-pull rods, supported by idling links, bell cranks, roller bearings or graphite impregnated bushes, form a simple rigid link system which is free from backlash. Changes in direction of the control run are obtained by the use of bell crank levers or torque tubes. The rods provide both a push and pull input. Not often used as a complete system, but push / pull rods are used in cabIe systems. Figure 7 shows an example a rod system - it's not a civil aircraft but it does show that the system is used on some aircraft.

ARTIFICIAL PEEL

COMf ROL

FEEL UNIT

! -. -- I --" - A- - --------- ,---- - --- -- ma- - - r-~g . '7 r ; ~ r ~ p ur. A ruanj ruhb KUU ~ ~ W K E ~ Y L

(ENGLISH ELECTREC LIGHTNING) ; !

b) ,Cable System (figures 8, 9 & 10) I i ~bn-corrodible extra fle$blehgh tensile steel cables; tensioned to ieQrninate lag, farm an efcective continuous loop over pulleys and !quadrants so as to provide1 a pull iri either direction. Each length of cable ]has end fiitings swaged in position, some of which are drilled and tapped 'to' accommodate turnbuckles used t o tension the cable sun.

- - - - - . . - .

On long straight runs tie rods may replace cables. The cables are supported on pulleys and fair-leads and pulleys are used to change the direction of the cable run. These pulleys and fair-leads must be kept clean to reduce system static friction. The cable systcms are usually duplicated with 'port and starboard' systems interconnected via 'disconnect' rods that will disconnect should one side jam.

Cable systems are used on most aircraft.

For more details on cables the reader is referred to J A R (EASA) modules 6and 7.

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W e s 1. The system I r nrm-pcwered servo tab operated. 2, rhsarmwsshwfhe movement c4 the systam Iar a bank

to the IM lcorrecl sense), mth each servo lab movlng In the opposlte dlrectlon lo 11% respect& conlrol surface.

8. Study the syalem and nMe Ihe rorlowlng' (a) CaMe lenslm icglualors. (b] Auto pllot servo <onel. (c) The dctenl strut 1 d ) ~ h e balance cables. {e) Ttwdtsconnwt devke. (f) 7 he Catlm pressure reals.

AILERON SERVO

DISCONNECT

CABCETEHSWN

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Fig. 8 EXAMP~E OF A ' \ C ~ L E OPERAT D: AILERON CONTROL S*~SITEM - BAe 14'6 ,

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Figure8showsthe aileron system of the--BAe 146. The ailerons kre operated via servo tabs (see later sections in this book). The system is cable operated and shows a commanded role to the left (left servo tab down - right aileron up). [If you are not sure about this read on, and after studying the section on tabs you may care to re-read this section again.1

The cables are routed from cable tension regulators via pulleys, pressure seals, gearing and push /pull rods to the ailerons. Because the ailerons are senro tab operated there is no need to fit powered flying control units - and therefore no need to have artificial feel.

There is one autopilot servo.

There is an interconnect or balance cable ensuring that when one aileron goes up the other goes down.

There is a detent strut fitted between the pilot's and co-pilot's controls so that if one side jams the other side can still be moved. Control is not so effective but there is still some lateral control.

PRIMARY STOPS

DATUM a ARTIF lClAL P E L SWllNQ STRUT

TRIM EeARBOR

1 DAMPERS

hte- ma fdluwlng testumm: 1. Thcwmrrable raw d mwsment umh a stepped

z;G=p:G :& :, a =:. * 2 : ~ *k:>-: Increases so the range aI rmnrarnerrt ~CCFESSFS.

2. The prrrnary and secondary stops. 3. 7lw1 arrows ind~calrng the mrrrct senSC*rlM Ihe

rtghr Toot forward the n ~ l l d ~ r mows right 4. TIIP d~rpllcated yaw dnrnpar*. 6. Thc rlupHcatcd PFCUs,

CONNECTKmS

Fig. 9 RUDDER CABLE SYSTEM - W e 14'6

Figure 9 shows the rudder system for the same aircraft - again cable operated using cable tension regulators, pulleys and thc cables passing through pressure bulkhead seals.

The system is power operated using duplicated PFCUs and duplicate yaw dampers.

Note the primary and secondary stops with a mechanism (Q pot) to reduce the range of movement as the speed increases.

Figure 10 shows the elevator cable system as fitted to a Boeing aircraft.

Note the triplication of the PFCUs and autopilot servos. Thesc arc all fed from three: separate hydraulic systems and can work individually if necessary - but with reduced authority.

Take a few moments and study the drawing - make sure you understand how it works.

Note: 1. Three PFCUs per elevator with a separate

hydraulic supply f o r each. CONTROL COLUMN 2. The overload device at each PFCU input. OVERRIDE 3. The control column override.

STALL WARNING f 4. The stick nudger operated by the stall

MODULE warning module.

ANGLE OF ATTACK 5. The stick shakers also operated by the

FLAPISL4T POSITION stall warning rnodulc. 8. T h m N P servos each powered by a

separate hydraulic system. 7. The artificial feel controlled by the feel

AUTOMATIC FLIGHT computer.

CONTROL COMPUT ER (8 ) ) I 1

I l l 1 - 1 W-H- DYNAMIC Q PRESSURE lNPUT (2) STAWLER POSITION IYPCrr (2)

A W I L O T PITCH CONTROL SERVO (3)

WWER CONTROL ACTUATORS (5)

W V E IMERCONNECT CABLE

AFT QUADRANT OVERRIDE SnCK NUDGER

SLAVE CABLE OVERRIDE MECH#JIISM

.<#..--"-'-'

. . . .. -

Most aircraft have a combination system of pushlpull rods and cables. The push/pulI rods are used for the shorter runs (under the flight deck, in the tailplane etc) and cables are used for the longer mns - down the fuselage, along the wings etc.

Advantages of Cable Systems

The overall advantages of a cable system over a rod system are that they have a better strength/weight ratio and are less expensive.

However, the structure of an aircraft is continually changing its length due to temperature changes and since the expansion rate of aluminium alloy (a = 23 x 10-6) in the structure of the aircraft is nearly twice that of the steel (a .;: 15 x 10- 6) in the cables, cable tensions can vary considerably. (a = coefficient of linear expansion).

QUESTION: Explain what would happen to the tension of n cable system running dong the fuseIage without any automatic adjustment when the aircraft increases altitude? (10 mins)

ANSWER: As the aircraft climbs so the ambient temperature reduces (as low as -56°C if it climbs high enough). This will cause the fuselage (and the rest of the structure) to contract in length quicker than the cables, so the pulley hinge points get closer and the cable tension reduces.

To overcome this problem requires cither very high tensions at ground level settings (as was used on some older aircraft) or the use of a tension regulator in the cable system.

Most aircraft systems are now regulated by means of Cable Tension Regulators, which give a retativeIy constant system tension irrespective of changes in temperature resulting in much lower rigged tensions (more of this later). - -..- -

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i I I CONTR~L'SYSTEM COMPONENTS i

These' include: I .. .

- - -

Cables. Push Jpul.1 rods. BeIlcranks. : I

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Torque tubes. Turnbuclcles. - .

Cable connectors. Fairleads. Bulkhead seals. Pulleys. Cable tension regulators. Quadrants . Powered flying control units. Artificial feel units. Position traducers. Position indicators. Warning systems. Autopilot servos. Computers. Stick shakers. Stick nudgers.

This book will not cover d l the items in the list as some of them are in module 7 and in the books in this series on Avionic Systems m d Powered Flying Control Systems - but together d l are covered.

Flight deck controls usually consist of a control column or control wheel and rudder pedals. Where there are twin controls (pilot and co-pilot) they are interconnected with a spring link Jdetent or automatic disconnect link so that if one side jams the other can be operated.

On small aircraft disconnect/ spring links are not usually fitted.

On larger aircraft the control column is fitted with a 'stick shaker'. This is a small electrical motor driving an out-of-balance wheel. When the aircraft approaches the stalling angle the motor is operated to cause the stick to shake and warn the pilot (with aural warnings as well). The warning signal comes from an angle of attack vane fitted on the side of the fuselage.

On some aircraft a stick nudger may be fitted to push the control column forward when the aircraft approaches the stalling angle. The pilot can overcome this if he / she wishes.

The rudder pedals are fitted with._an_adju s tment mechanism-to-allow for ~ d j i i s t m ~ n t tn s11it the 1 ~ 3 Im@h nf ench inrlivirl~~al pilnt. I-.

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The handwheel may have contrbls fitted t b ii such as: I , , ,

I . I , ,

a) ' Parking brake. , I , , . I , , I I

b) Auto pilot switch . ., , i , -'

c) Intercom switch 1 \ 1 , ...,'

',, ',, d) Elevator/tailplane tfih switch1 i ! , i !

i I I ! , !

, , i i ! , i . .. . -

I \ control. Co~umn -, - . , I

. , - -. . .. . . . . -- -- L . .-

Usually fitted with a chain, sprocket and cable system to transmit aileron control inputs to the system and is connected under the floor to the elevator control system.

On some aircraft (eg the A320) the control column has been replaced by a side stick fitted to the left hand console (for the captain) and the right hand console for the co-pilot. This is connected to the fly-by-wire transducers.

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- 15 - moodull 1A-16

AUf 0 PILOT CUTOUT & GO AROUND BUTTON

AUTO PILOTIFLIGHT DIRECTOR SYNC BUTT

INTERCOM SWlTC

OP SPROCKET

TORQUE TUBE

I// , ''AILERON CABLES

ELEVATOR CABLES

Fig. $1 CONTROL COLUMN - - .

Rudder 'Pedals I

I

May be bf the simple htdder bar' type or each pedal supported on separate lever'mechanisms. The foot pedals are adjustable to cater for different leg lengths and &e usually fitted S t h wheeI brake control foot motors.

RUDDER BAR ADJUSTER

SECONDARY CONTROL STOPS

CABLES TO / RUDDERCONTROL SYSTEM

Fig. 12 RUDDER BAR

Detent or Disconnect Strut (Torque Limiter on the L- 10 1 1)

Fitted on some aircraft where duplicate controls run from the flight deck to the control surfaces. It is fitted between the pilot's and co-piIotYs controls so that should a jam occur on one system the other system can be operated normally - after the detent strut has "broken out'. These break-out struts vary in design but the following description is typical.

Figure 13 shows the detent s t r u t from the BAe 146 aileron control system. I t is located under the flight deck floor and connects the pilot's control column to the co-pilot's control column. Effectively it is a rigid link as the rollers are forced into the detent grove by the action of the spring collet.

SLIDER

CaMTROC COLUMN SmocnET LEVER

,

Fig. 13 THE DETENT STRUT OF THE BAe 146

If one side of the control system becomes jammed and the other side is operated then [at a break out load of 35 lb force) the link 'breaks out'. This allows one side of the system to be operated even though the other side is jammed.

During this operation the inner shaft is allowed to slide in and out of the outer shaft as the rollers have been forced out of their detent positions.

The break out of the strut operates a microswitch on the strut that releases the disconnect device on the aileron balance /interconnecting cable circuit, allowing movement of either aileron.

Stick Shaker

This may be initiated by a:

--a)._ _ Leading edge stall warning vane. ' b ) Rotating angle of-attack probe.

I c ) Trailing angle of attack vane. 1 I

Thesei day operate the stick sh&r and Stick pusher directly v i a a micro switch (A), or v ia the DADC (~istal Air Data Computer) (b) and (c).

' i -.

The stick shaker is an electric mbtor driving an out-of-balance wheel attached to the, cpntrol column, or closejby on the system. When it is switched on, just before the stalling angle (12" to 1,4" with a "clean" wing), the out-of-balance

I wheel1 causes the control column: to shake, warning the pilot of an impending stall. -

. . - -- - -. . . - - . --

Stick Pusher

If the pilot ignores the stick shaker and the angle of attack increases still further, then the stick pusher system is activated.

This system operates a device, which is connected to the elevator control system. This gives a positive push to the control column and causes the aircraft to pitch nose down when the system is activated. It c a n be overridden by the pilot, if necessary, by the operation of a switch, which on the pneumatic system releases the pressure in the supply line to the jack. Even if this f d s the system is so arranged that the pilot can manually overcome the force of the jack, if necessary, by pulling on the control column. When not supplied with pressure the jack moves freely when the controls are moved.

The pneumatic system i s supplied with air pressure from a tapping on the jet engine via the pneumatic system (typically 40psi).

- 18- rnooduli 1A-19

The system usually has the following inputs:

t Airspeed switch. Increases speed of operation with reduced airspeed.

* Stick shaker relay, which receives the signal from the angle of attack indicator probe.

v Flap and slat position. The deployment of these will affect the stalling angle.

Figure 14 shows the stall warning system for the BAe 146. Note the following inputs /outputs:

* Weight on wheels (squat) switch. * Flap position. * Test. * Power supplies. * Airspeed. *

- Fail. , , , .... ,

I I

, 8

. --, ,

I ' A& AIRFLOW SmlSOR VANE POSlTlON

. .p-.2 -SIGNAL---- - - - SUMMlNG UNIT - QOA SIGNAL

FAIL I------

AIRSPEED TRANSDUCER ONE CHANNEL SHOWN THERE ARE TWO CHANNaS PER SYSTEM

AIRSPEED TRANSDUCER

AOA AIRFLOW SENSOR VANE POSITION

SIGNAL VANE EXCITATION WARM SUMMlNG AOA SIGNAL

I

FLAP INPUT

Fig. 14 STALL WARNING CIRCUIT - BAe 146

- 19 - rnoodulll A-20

Control Stops

Usualtlly adjustable and fitted to the front and rear of each system and will control the range of movement of the system. Primary control stops are fitted to the control surface end of the system while secondary control stops are fitted to the cockpit or flight-deck end of the system. On some aircraft the range of movement is progressively reduced as the speed of the aircraft is increased - the BAe 146 rudder system for example. This is achieved by automatic moveable control stops or Iimiters that are controlled by either a computer (having airspeed data) or from an airspeed module (see figure 9).

Fig. 15 CONTROL STOPS I --

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i chains knd Sprockets I

cables hay go around a pulley or be connected to the pulley end fittings. The cable may terninate at a chain fitting - usually a turnbuckle - and the chain passed around a sprocket. This provides a positive drive to the sprocket. Chains may be of the 'non-reversible type', which means that they are so designed that they cannot be put on the sprocket the wrong way round.

1 %' EKwULL CHAIN

NSlON ROD TYPE

NN ECTED TO

Fig. 16 SPROCKET & CHAW DETAIL

Wsh/ Pull Rod Support

Push/pull rods may be supported on idling links or various types of bearings. Remember to keep the bearings clean and dry at all times. To change the direction of the run a bell crank lever, torque tube or pulley is used.

PUSHIPULL ROD END DETAIL PUSHIPULL

INSPECTION HOLE

RoLLER RACE (9) GRAPHWE LOCKNUT IMPREGNATED BUSH ADJUSTABLE END FITTING

Fig. 17 PUSW/PULL'RODS-=-SUPPORTS & END-FI'l"rXNGS -, . -

, ' , , ', \,

I , I , ,

! '

I 9 - - --'

, .

I . ..

I - -' LA..

.. . .,

Fig, 18 BELL C W K LEVER

SUPPORT - AIRCRAFT STRUCTURE , STRUCTURE

BBlNG STRIP

SUPPORT CLIP

Fig. 19 FAIRLEADS

Cable Support

Cables can be supported by pulleys and special quadrants where they can change angular direction - without limit. Where little or no change in direction is required various types of fairleads c a n be used. Fairleads are usually made of composite material and must not be lubricated.

Pulleys

Made from fibre, plastic or metal and are used to give a more abrupt change of direction of the cable run. Guards are usualIy fitted to retain the cable on the pulley and often a cover to keep out unwanted small items, which might foul their operation.

SUPPORT BRACKET

PULLEY ' SUPPCF??

.,K, fE \

. -

- -. . . . . . . . Fig. 20 CABLE PULLEYS & GUARD PINS

QUADRANT PIVOT

AIRCRAFT

CABLE I CABLE 2 ATTACHMENT ATTACHMENT /

QUADRANT

Fig. 21 CABLE QUADRANTS

- 22 -

moodul 13 A-23

Turnbuckles

These may vary in design and commonly may be of the Barrel Rod type or the Tension Rod type. I n general they all have a left hand thread at one end and a right hand thread at the other. When the centre part is rotated - holding the two cable ends to prevent them rotating - then the cable tension will either be increase or decreased. It is important that, after adjustment and prior to wire locking that the threads are in safeq.

For the barrel type turnbuckle that means that all the threads must be buried in the barrel. For the tension rod type the threads must be screwed in deep enough into the fork ends so that a piece of locking wire will not pass through the inspection hole. The wire should be the same size as the inspection hole and should not come out the other side.

BARREL LOCKING WlRE

FORK END

I I , ' I I

~ i g . 22 BARREL ROD.+J!PE TURNBUCKLE I I

LO~KING RE TYPE I i

Fig. 23 BARREL ROD TYPE TURNBUCKLE - LOCKING CLIP TYPE

THREAD RIGHT HAND THREAD INSPECTION HOLE

LEFT HAND THREAD

\ TENSION ROD /

FORK END /

LOCKNUTS LOCKING WIRE

Fig. 24 TENSION ROD TYPE TURNBUCKLE

Note the various locking methods (figures 22, 23 and 24). Some barrel rod types are locked with locking wire in a figure-of-eight fashion, others have a special Iocking clip. The tension rod type is locked with lock-nuts and locking wire.

Cable Connectors

These are fitted to some cable systems at positions where the cables need to be disconnected. Each half of the connector is keyed in such a way that it can onIy be fitted back to its mating half. The connectors allow for quick cable disconnect and re-connect without the possibility of connecting two wrong cables together. They usually do not provide for any cable tension adjustment.

Torque Tubes

Used to change linear motion into rotary motion - or vice-versa. .--

. .

cable\ Tension Regulators : The dajority of modern aircrdt now use cable-operated systems for their flying contrdd. This is due, in a large part; to the development of an efficient Cablc Tensiori Regulator. 1 7

: I

I

Cable tension regulators are mechanical devices and can be made in many configurations, for example, quadrants, bell crank levers, pulleys etc . For the purpose of a brief description, we will consider the quadrant type.

- - - - -.

QUADRANT \ TEMPERATURE SCALE

Should be withln 5 degees of actual temperature - but check AMM.

Fig. 25 CABLE TENSION REGULATOR

This consists of a pair of spring-loaded quadrants with a pointer scale for recording the cable tensions. The cables are inserted through slots in the recessed ends of the V grooved quadrants and the cable ends are secured at the cable anchorages in each quadrant.

When the cables are tightened equally (as with the fuselage getting Ionger as the aircraft descends) the quadrants rotate about the centre shaft and the links pull the crosshead freely along the locking shaft, compressing the springs and, ie effect tensioning the cables - or at least keeping the tensions correct.

When the aircraft climbs the fuselage contracts and the cables tend to slacken, but the springs react against the crosshead and push the crosshead back along the shaft, thus tightening the cables - or at least keeping them at the correct tension.

- I 8 ! - i I CABLE TENSIONS: KEPT CONSTANT AS 1 ,

- I ' CABI$ T E N S ~ N S KEPT CONST~NT AS

FUSELAGE TEMFERATURE DECREASES ' - FUSEIAGE GETS WARMER &'INCREASES 'a FUSELAGE GETS SHORTER IN. LENGTH I - - -~

NO PlLOf 1NPUT - CROSSHEAD MOVES FREELY ON LOCKING SHAFF WlTH SPRINGS MAINTAINING CABLE TENSIONS

PILOT INPUT - CROSSHEAD TILTS AND LOCKS ON LOCKING SHAFT & THE uNrr BEHAVES AS A PULLEY

fig. 26 CABLE TENSION REGULATOR - OPERATION

When a control load is applied by the pilot only one link will tend to move, tilting the crosshead on its locking shaft (by a very small amount) and locking it to the shaft, preventing movement of one quadrant relative to the other with the whole system now acting as a pulley. Both quadrants are, therefore, locked together and operate as a solid pulley until the control load is released.

Each tension regulator incorporates a scale and pointer, which provides a visual tension indication. When rigging a regulated cable system therefore, a tensiometer is not required, the cables being tensioned until the correct reading is obtained on the regulator scde. The correct reading depends on the ambient temperature and must be obtained from a special graph provided for each regulator in the aircraft.

Pressure Bulkheads

On pressurised aircraft the control run will have to pass out of the pressurised area to the un-pressurised side of the cabin or pressure hull.

- I -- -

mi \ - , ' I L - - L - - - - - A --Ll-- "- I nus p vsrl/ pull rods, it~~-+ LUULB a l l ~ L ~ V L L ~ illlist p a ~ ~ t:ii-~iigIi seals :G

preveht #undue air leakage. The pressure bulkhead seal. must allow freedom of moveLent, be sclf-aligning, require little or no maintenance - and provide a good air seal.

# I

I

: I I I

I TOBULKMEAD

I . PRESSURE SIDE I I

- - - . - - -. - -

Fig, 27 BELLOWS TYPE SEAL

One such device consists of a rubber or polymer bellows, which moves with the control. This arrangement is used with control systems using twin cabIes (one up and one down), as the cabin pressure acting on the bellows causes a load on the control which must be balanced by an equal and opposite load on the other cable.

This type of seal is self aIigning and completely airtight but does impose a load into the system - particularly on the pulley bearings o n the pressure side of the bulkhead.

Fig. 28 GLAND TYPE SEAL

Alternative methods for control rods or cables include some form of gland assembly. Several types have been developed all of which rely on packing rings or silicon rubber composite to provide the airtight joint. Remember, they should be kept clean and not lubricated.

Improved-sealing is affected if-the movement through the bulkhead is rotary rather than linear (figure 29). ~ h e - ~ p i ~ a l &rrangernent ill6stred-consists of a rubber seal, clipped to and rot&img with the control tube. ~ k r pressure acting on the splayed outer end of the rhbber seal !forms an airtight j int. i

I I . P I '

The seBJ ;slides on a smooth rubbing plate fixed to the bulkhead. - - .

PRESSURE

Fig. 29 ROTATING SEAL

Powered Controls

The control surfaces of many modern aircraft are subjected to high aerodynamic loads due to the airspeed and/or the size of the control surface. These loads are often greater than the pilot can comfortably overcome and the system must be powered.

I t is usual to power the control surfaces hydraulically, using hydraulic pressure from the aircraft's hydraulic system to operate a jack the control vaIve of which is moved by the pilot via t h e control system.

However, to ensure that the surface moves only when, and as far as the pilot wants it to, a feedback from the jack to the control valve must be incorporated. (Negative feed back) :).

The basic jack and control valve are incorporated in one unit called a Powered Flying Control Unit (PFCU or PCUJ, and may contain the autopilot servo.

The usual type of PFCU feeds off an aircraft hydraulic system but there are other types which are self-contained hydraulic systems requiring only electrical power to drive their pump motors,

For more information on PFCUs see the book in this series entitled Powered Flylng Controls and Autopilots.

-

FCC: S-J-Gtc;ns I

when' a control system is fully Powered the pilot loses all sense of feel, since the work is done hydraulically. As feel is essential for the pilot to fly the aircraft properl-$ under adequate contrbl, it must be provided artificially.

I , I .. .-

The simplest form, a spring bok in the control run, supplies a constant feel force iriespective, of air load variations, and 'is therefore not completely satisfactory - but it is chcap. i

-. . Another system measures dynamic pressbre (q) and therefore surface loading, and varies the feel force accordingly. Feel force thereforc increases as speed increases and vice versa and decreases with altitude and vice versa. I t is known as a 'Q Feel System'. (Again, see the book in this series 'PFCUs and Autopilots 7 .

TABS

These are ancillay surfaces attached to the rear of the primary flying control surfaces.

A control surface may have several different types of tab fitted to it and in some cases more than one function may be built into one tab. Tabs can be fitted to non-powered systems and tabs (balance & anti-balance) may be fitted t o same powered flying control systems,

In general, tabs are designed so that if moved in one direction they produce an aerodynamic force, which causes the main control surface to be moved in the opposite direction. Different tabs, however, have different functions.

Fixed Tab

This is adjustable only on the ground by the maintenance engineer. I t may be fitted to non-powered controls and is used to correct for inherent flying faults. I t may be an actual tab as shown in figure 30 or may be a metal strip riveted to the trailing edge of the control surface. The tab, as shown, is adjusted by removing the flxing plate, repositioning the tab and refitting the plate. The metal strip type tab is bent into a new position usually using a special bending tool.

PUSHIPULL CONTROL SURFACE CONTROL ROO \

/ LOCKING PLATE

TAB

I i 1 . >

, I ~ i k . 30 F T X E ~ TAB I

, ; 1 I '

I . ,

The tab is moved in the opposite direction to that which we I 1 surfack to move, eg to correct for an inhereht nose down flying a t t i t~de the

e~evatdr needs to be raised whidh means thdtab on the elevbt* i s adjusted do ~ndai-ds . i 1 i 1

, i The pilot will report any tendency for the ,aircraft to fly in a-p&cular attitude. The criiineeiiiill consult the ~ M M *-d;if not stated in thk AM^, work out which way the tab is to be moved to correct the fault. The tab is repositioned and on the next flight the pilot checks on the flying characteristics of the aircraft - if necessary the tab will have to be adjusted again.

Controllable Trim Tab

This is moved by the pilot during flight to allow the aircraft to be trimmed to fly straight and level, although it is sometimes used to trim the aircraft into a climb or a descent path.

The tab may be manually operated using cables, chains and screwjacks, or it may be electrically Aperated with an electric actuator controlled from the flight deck. Fitted to non-powered controls.

QUESTION: Why should the pilot need to trim the aircraft into straight and level flight during ff ight? (5 rnins)

ANSWER: The trim of the aircraft might change due to fuel usage, or one engine (multi-engined aircraft) shut down. To save fatiguc on the pilot the aircraft can be trimmed into an attitude where the pilot has to put little or no input into the system to keep his aircraft flying on the correct course and altitude.

PUSHIPULL ROD TO FUGHT DECK FLYINS; CONTROLS

A

=

v \ PUSHIPULL ROD TO FLIGHT - DECK TRIMMING CONTROLS

Fig. 31 CONTROLLABLE TRIM TAB

E3alarice Tab I I ! I

This askists the pilot to move his/ her contr'ols on a non-powered system, or relieve (he load on a powered system.. It is automatic in operation.

: I ! * I I I I---

The tab1 is fixed by a rod to the, niainplane, tailplane or fm. When the control surfakcis maved.by the pilot the tab is caused to move in the opposite direc~ion. The airflow hitting the tab willcause a force to be created in the direction that the control surface is being moved.

~ h i s f ~ s < ( f ) (Ghilst small) is 3TS6me distance (d) from the control surface hinge line - thus a turning moment is creatcd (f x d) which is significant enough to assist the pilot to move his / her controls.

CHED TO MAIN SURFACE May be adjusted by engineer

Fig* 32 BALANCE TAB

QUESTION: Could the balance tab also be used to function as another type of tab and if so what? (5 rnins)

ANSWER: If the attachment rod is adjustable on the ground (which it usually is) then the tab can be used as a fixed tab as well as a balance tab. If the length of the attachment rod can be adjusted from the flight deck in the air then the tab can perfom the dual function of a controllable trimming tab and a balance tab.

Anti-Balance Tab

To make the controls more effective and to give the pilot more feel an anti- balance tab may be fitted.

I t is similar to a balance tab except that the linkage is so connected that the tab moves in the same direction as the control surface - but further. It makes the control surface more effective by giving the control surface itself a curved mean camber line.

MAIN CONTROL SURFACE

I -- - Not connected to any control system. > I

, 1 ---- $ Y

1 1 : i \ 1- I

I ' I \ I

P u s H l P ~ u ROD TO F L ~ H T , ' - . DECK FLYING CONTROLS I-

F&. 33 SE ~ V O TAB I . ' 1

I I I I

I I ! I ,

I t is interesting - to note that the a h ~ i ~ b d m c e tab fitted to thk Adder of the Canadian de Havilland Dash S&i the same chord length as t l i k rudder itself - presumably making for very effective directional control.

Servo Tab

This is similar to the bdance tab in principle but it is operated directly by the pilot. The control surface is not connected to the control system in any way but is free to move in any direction. Movement of the piIot's primary control moves the tab, aerodynamic pressure on the tab will cause a turning moment on the control surface, which will move in the opposite direction. Control, however, is still instinctive (control column forward - aircraft descends etc) . Fitted to non-powered controls.

QUESTION: Which way would the tabs move on the ailerons to cause the aircraft to roll to the right? (5 rnins) (Hint - in your mind move the primary control surface first - the tabs move in the opposite direction).

moodull l A-32

ANSWER: The sight hand tab will move down causing the right hand aileron to move up and the right hand wing to move down. The left hand tab will move up and its aileron will move down.

This type of tab works very well at reasonably high speed (all the primasy flight controls on the Bristol Britannia are servo tabbed), but at low speed the system has problems. When the tab moves into low airspeed it produces littJe force and consequently poor control surface response - at high speed response is good. To overcome this problem Spring Tabs were invented.

PIVOT LINK

- Ce

-

SPRING- - - J Any,direct control surface movement

; PILOT'S INPUT I '

from the flight deck is wla the spring only. I I

F&. 34 SPRING TAB

i .:

This is similar t o the servo tab but it only operates at the higher airspeeds. At the lower airspeeds the pilot operates the control surfaccs as normal.

.... . - . - . . . .

The control linkage is connected directly to the tab with a connectioi~ to the control surface via a torsion bar [shown as a spring in the drawing for case of explanation).

At low airspeeds the loading on the control surface i s insufficient to overcome the pressure of the spring/ torsion bar and the movement of the control system moves the control surface directly through the spring/ torsion bar. At high airspeeds the aerodynamic loading on the control surface is sufficient to overcome the force of the spring/torsien bar and the link moves. Movement of the link compresses or extends the springJtorsion bar and moves the tab= The tab in turn moves the contra1 surface - acting as a servo tab.

CONTROLLABLE TRIMMING TAB SYSTEM

These are operated by the pilot independentIy of the main controls and are not fitted to powered flying controls. They are used to trim the aircraft to a particular flight attitude, eg to trim it to fly straight and level or trim it to descend.

They are operated from the flight deck and the system may be: mechanical using cables and pulleys or i t rnay be operated electrically (a switch - often on the control wheel - operating an electric actuator at the tab end).

The cockpit controls are designed so thcir operation are 'instinctive' - ie handwheel fonvard - nose trimmed down.

The tab may be combined with say a balance tab and may be fitted to the rudder, elevator and, usually, one of the ailerons.

+ CHAIN

CABLE \ FUGHT DECK HAND-WHEEL

- -

Direction .of. Movement I .A__-

TAB

I

I

The control run of the controllable trimming tabs is usually complicated and because of the screwjack or other similar device, their operation may not be readiIy understood. Therefore, it is advisable to re-check that the movement of the cockpit control does result in t h e correct movement of the tab.

Elevator Trimming Tab

These are usually operated by a handwheel mounted in the vertical plane fore and aft, so that when the top of the handwheel is moved fornard, that is, wheel wound forward, the nose of the aircraft goes down and vice versa, Markings on or near the handwheel, such as Wose up' and 'Nose downJ indicate the direction in which to turn the handwheel. [Note the mechanical trim wheel on the centre pedestal of the Airbus A320).

TABLE 2 ELEVATOR TAB MOVEMENT

TAB

HANDWHEEL MOVEMENT

I a , -,

.- -

SUBSEQUENT PRIMARY CONTROL

SURFACE MOVEMErn

UP

AIRCRAFT FLYING FAULT

Nose heavy reward

AFFECT IN FUGHT

Fig. 36 MOVE~ENTOF ELEVATOR TRIM TAB I , I

I , '

PILOT'S ACTION

Control wheel

Down Tail heavy

Ailerdn ;Trimming Tab

TAB MOVEMENT

Down

This is- usually 'operated by a-hadwheel mounted vertically .yon a fore and aft spindle. As the ailerons arc interconnected, a controllable trimming tab may be fitted to one aileron onIy. Markings on or near the handwheel, such as, To correct for port wing low' and To correct for starboard wing low' indicate the direction in which to turn the handwheel.

Control wheel

TABLE 3 AILERON TAB MOVEMENT

UP

TAB MOVEMENT (RIGHT AILERONJ

AIRCRAFT FLYING FAULT

Left wing low

PRIMARY CONTROL SURFACE

MOVEMENT - Right up

PILOT'S ACTION

Right wing low Right kfr down

Wheel to the right

- 34 -

moodull 1 A-35

Down

Wheel to the left UP

HANDWHEEL MOVEMENT

AILERON 1 MOVEMENT

Fig. 37 MOVEMENT OF AILERON TRIM TAB (FITTED TO THE LEFT WING)

Rudder Trimming Tab

These are usually operated by a handwheel mounted horizontally, though some aircraft may have the handwheel mounted vertically on a fore and aft spindle. Markings on or near the handwheel,-such as 'Correct for VTW to port' and 'Correct for yaw to starboard' indicatethe direction in which to1:Ui-n t he handwheel. I I ', \

I I ' I

I I l 1 I ' I

TABLE 4 RUDDER-TAB MOVEMENT ' , - --- 8 .

HANDWHEEL MOVEMENT

I i 1 I --_-

AFFECT IN FLIGHT

A I R C V FLYING FAULT

I I '

Yaws left

Fig. 38 MOVEMENT OF RUDDER TRIM TAB

PILOT'S ACTION ' I

I

Wheel ctockwisc--- Yaws ri g ht Wheel anti-clochse - g . -

- Ri ht I k f t -

:TAB MOVEMENT I I PRIMARY CONTROL

I

4

, Left , - , MOVEMENT

\ , ~ i ~ h ?

SERVO TAB

/ TRIM TAB

'I-. TRIMWHEEL ,, ( MOVEMEN-

1 '

Fig. 39 AILERON TRIM SYSTEM OF THE BAe 146

I ' I

~i 39 shows a typical system for operating the aileron trim tabs. Study the drawihg and note how it works. ,

Electrically Operated Trim Tabs

Most of the electric him systems are an extension of the manual system with provision to allow for manuaI trimming in the event of electric supply failure or electric actuator failure. The actuator i s usually reversible and incorporates an overload clutch that will slip in an emergency.

Cockpit control is usually through 'thumb' switches that return to the centre off position when finger pressure is released. Pushing the switch one way will cause the tab to move, say, down, while pushing it the other way will cause the tab to move up. Again the system is instinctive. On some aircraft the three axes or trim control are incorporated into one switch, eg switch forward - nose down, switch tilted to the left - aircraft rolls left, switch turned to the left - aircraft yaws left.

Operation of the switch will cause a voltage to be supplied to the motor. The motor will operate a reduction gearing or a screwjack to move the tab direct.

On smaller aircraft a dc motor of the permanent magnet type is used with reversal being achieved by reversing the current in the m a t u r e . On larger aircraft the motor is a split field motor.

BALANCING OF CONTROL SURFACES

Control surfaces are usually mass balanced and aerodynamically balanced. M a s s balance is used to reduce the possibility of flutter and aerodynamic balance is used to assist the piIot to move the controls.

QUESTION: Canyouexplainwhatismeantbythe term 'Elutter'? (10mins)

ANSWER: Like all things aeronautical, 'flutter' is a cornpIex subject and it comes in many forms. In general, however, it can be considered as a form of vibration which is induced by aerodynamic forces and is a function of the 'stiffness' of the structure and the control system and the flying control surfaces. I n its mildest

-

I form it may no! be noticed, ,or may show upxs-increased wear in control surface b~afings-. In its more severe form it can cause the aircraft to disinteg?? in fright with explosive fdrce. Aerodynamic

I pressure variations can cads4 the structure/ b ontrol sqistern to flex in a syrnpathLtic mode - .if this is sever the structlre can

I vibrate violently and -even'disintegrate. I ! I - , F- - -- ,

, , ? \,

M a s s Balance 'I , \ I I : I I

I I

I I

i Flutter can be reduced, or even prevented, by the mass bal c g of the control surfaces. -- / I

I . . -- L i n --- - -

During the design stage the centre of gravit-y of the control surface (chord wise and span wise) is calculated to be within certain limits.

This is usually achieved by the addition of carefully calculated weights (to be technically correct - masses) placed forward of the hinge line. This is donc on control surfaces whether they are powered or not, and the C of G must be within a certain range in plan view and in end elevation, ie in two pIanes.

MASS BALANCE HINGE LINE WEIGHT

Fig. 40 MASS BALANCE WEIGHT - CHORD WISE LOCATION

TRAILING EDGE \ C OF O RANGE u

n - MASS BALANCE -

I WEIGHTS

3 -Fy-m- -- -HINGE LINE

LEADING EDGE / HINGE

Fig. 41 C of G RANGE IN PLAN VIEW

The range of the C of G in end elevation may be such as to make the control surface nose or tail heavy or either depending on the actual C of G position.

r--- -- ; I.. ._ >.

- , , I :

I I

I I ' Fig. 42 C of G R A N ~ E - NOSE' HEAVY CONTROL SURFACE I

I Fig. 43 C of G'RANGE - NOSE OR TAIL HEAVY -

Fig, 44 C of G RANGE - TAIL HEAVY

Of course, the designer will decide where the C of G is to be and a range is given to allow for minor repairs and paint finishes to be carried out without the need to continuously adjust the mass balance weights.

The mass balance and total weight of the control surface will be checked in accordance with the AMM where the equipment will be specified and the calculations to be used will be shown.

QUESTION: Can you specify 2 or 3 occasions when control surfaces should be weighed/ mass balance checked? (1 0 mins)

ANSWER: 1. As laid down in the maintenance schedule. 2. On fitment of a new or repIacement surface. 3. After repair of modification to the control surface. 4. After reported flutter. 5. After aircraft/ control surface re-spray.

The check may be done on the aircraft but the surface may have to be removed and the check carried out on a bench (bench check).

The check may involve balancing the surface on its hinge line (using locally made special brackets, if on a bench check) m d adding test weights to the leading or trailing edge. The weights are added to bring the control surface horizontal. Calculations are then carried out, using this information on how much weight has to be added or removed to the mass balance weight. The mass balance weight is adjusted and the control surface re-tested and the test weights removed.

After adjudtment, the new b a h c e irlf0~mation (together with- the total wcight of the 'surface) is recorded in the aifEiaft logbook. There may 4lsG be a record plate on the control surface whicli will netdiamending. Thc ko'ntrol surface may also be balanced in the spanwise plane - chkck the manual!

I ; ' i I I

! ! I ' I

I I-- QUESTION: What are themassb35ficew,eightsmade of rfndhow can they

be adjusted? (15 mins) I

i I ! I I I ANSWER: I Mass balance should be dense (heavy) and cdn b e madc of:

* Steel (density 7800 kg/ 1-123). - *Lead(densitywl-1-300kgJm3). I

* Depleted Uranium (DU) (Densiw =I9000 kglrn3). Most of it phased out now because of health concerns.

* Tungsten (density - 19300 kg/rn3). Replacing DU. Expensive.

Note: In general the more dense a material the less space it requires in the structure for a given amount of mass.

Adjustments will be carried out in accordance with the AMM, but may involve the use of the following:

* Addition J removal of steel washers/ bolts. * Addition/ removal of mass balance weights. * Addition / removal of steel shims. * Re-location of mass balance weights - fore 8t aft. * Changing weights for different sizes. * Machining of weights - only if allowed in the AMM, and

remember DO MOT cut or damage depleted uranium.

Aerodynamic Balance

The force, which is needed to move a flying control system in flight, depends on: air density; aircraft velocity; control surface size and angle of deflection.

This can be calculated by the equation:

where F - force in Newtons -

P = air density in kg per m3 V = velociiy in m/ s2

S - - area in m2. This is related to control surface size and amount of deflection.

(NOTE: The symbol 'ockmeans 'is proportional to')

On large/fast aircrart his force may become too much for the pilot to overcome so the coitrols are powered m d / W aerodynamically - balanced.

Aerodydarnic balance can be dchieved by: !

;(a) Balancetabs. , (b) Servo tabs. - -

' I I (G) Spring tabs. I (d) Pressure balance. ' I (ej Horn balance. I I

( Inset hinges. I

g ) Balance panels. -. - . - -

Having already dealt with the tabs let us have a look at the rest,

Pressure Balance - With this system part of the control surface, in the lorm of a beak, extends fornard of the hinge line into an enclosed area within the main structure of the wing, tailplane or fin (figure 45). When the pilot moves the control surface, say up, the air pressure increases above the surface and decreases below it. This differential pressure is felt across the beak thus assisting the pilot to move his/ her controls.

Sometimes the gap between the beak and the aircraft structure is sealed by a flexible apron (Westland/ Irving Type) or hinged flap (balance panel system).

blank

SHROUD - CONTROL SURFACE MOVEMENT

/ LOW PRESSURE

\ BEAK

SMALL GAP

Fig. 45 PRESSURE BALANCE CONTROL

Horn Balance - This is where part of the control surface extends forward of the hinge line (figure 46). When the surface Is moved in one direction the horn moves in the other direction, but out into the airflow, thus it experiences an aerodynamic force, which helps the pilot to move his/her controls. The horn

I -- - - --.

=zy a!sc hzuee t h e =ass b d ~ q c e ~ r z i g ~ : : I I I: I

A problek sometimes experiended with h4A balances is thdt &f 'snatci'. When I I the control surface is moved from, the inlipe;position, as thelhorn prdtrbdes

into the airflow so it experiences a sudden .- force causing the/ cbnt-rol to snatch. i r--- - x x h

I

I %

ELEVATOR

Fig. 46 HORN BALANCE

Inset Hinge - On this control surface the complete leading edge extends forward of the hinge line and the effect is similar to that of the horn balance. Again this area houses the mass balance weights. Snatch may also be a problem.

Fig. 47 INSET HINGE

-

Balancc Panel - Similar to the pressure balance control. The balance pancl, is housed in a balance bay forward of the control surface (fippre 48). When the mntro! surfclre is mnv~ld R pressure-difference is felt either side of thc control surfabe which is allowed to pass :through the gap between control surface and shroud and act on the balancd panel. This action assists control surface movement.

~ i g u r k 48 shows the elevator balance panel of the Boeing 737-400 I . I

INSET HINGE INSET HINGE

- HINGE LINE -I

Fig. 48 BALANCE PANEL - EXAMPLE 737-400

I -e

- C - - . ' - ,- -

- - - -

AILERONS

When operated, all control surfaces produce some drag and with most of them this is not a major problem. Ailerons, however, are different. They can cause a problem called adverse yaw.

When the ailerons are moved the down going aileron tends to produce more drag than the up going one. In a turn the down going aileron is on the up going wing. This wing is on the outside of the turn. If the down going aileron produces too much drag then it may tend to turn the aircraft in the wrong direction. To counteract this problem the aircraft may be fitted with:

* Frise Ailerons

* Differential Ailerons

* Aileron Upfloat

Frise Ailqrons 1 I

I I

These are designed so that the up' aileron)(oh the inner win$ o f the turn) produces more drag than the down goingiode, thus the airciqt is helped to lurn in the correct direction. ~ d e ~ l e r o n has a low set hingd sb that when it is moved :up the leading edge of the GlEETn protrudes into the lairflow-and creates drag. W h e n the aileron is moveti down it produces less dm$. I

I I

\ HINGE LINE

DFZAGCAUSEDBY AILERON NOSE

Fig, 49 THE FRTSE AILERON

Differential Ailerons

The aileron control system is designed so that the up going aileron moves through a greater range of movement than the down going one. Thus the aileron on the inside of the turn produces as much, if not more, drag than the one on the outside of the turn - thus preventing adverse yaw.

45OANOULAR RANGE OF MOVEMENT UP & DOWN

CAB ACTUAL LINEAR RANGE OF MOVEMENT

I t PUSHIPULL ROD TO AILERON

Fig. 50 DIFFERENTXAL AILERONS

With reference to figure 50 and assuming a 45" angular range of movement of the pulley, it can be seen that the linear range of the push/puIl rod is greater when it moves up than when it moves down. Thus the up going aileron (lower wing) ;loyes through a greater range - and produces more drag than the down going ;one.

1 , I I ,

C A U ~ O N . Some people get confvsed with the word "differential" thinking that it means that the ailerons move in opposite directions. All ailerons always move in; opposite directions bqt dqferential - ailerons have a different range of movement - up and down. j I I

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Aileroh 'Upfloat '. .

On some-smaller aircraft the -ailerons may be rigged into their "neutral' position with a certain amount of 'upfloat'. I n other words the neutral position of both ailerons is set above the trailing edge of the wing (refer to the AMM). This will mean that the up going aileron will move even higher into the airflow - with an increase in drag and the down going aileron will not move so far into the airflow - producing less drag.

QUESTION: This last method is an inexpensive way of counteracting adverse yaw, but it does have one disadvantage. Can you think what it is? (2 mins).

ANSWER: You might have thought of several disadvantages, but one that springs to mind is the continuous drag penalty. When flying straight and level both ailerons are high in the airflow and creating drag - not a good idea.

Another disadvantage is that they are both creating a slight amount of 'negative lift', ((On most large aircraft both ailerons are usually set down a small amount when the aircraft is in the landing or take-off configuration. This increases the camber of the mean camber line of that part of the wing and increases lift. They still work in opposite directions of course.

Enhanced Roll Control

On many larger aircraft the aileron system is interconnected with the spoilers to give better roll control. The spoilers are operated asymmetrically in conjunction with the up going ailerons to increase drag (and reduce lift) on the down going wing. The operation of the spoilers may be related to speed and/or range of aileron movement.

Fig. 5 1 ASYMMETRIC SPOILER CONTROL - EXAMPLE

In some aircraft the operation of the asymmetric spoilers is by a direct link between the aileron system and the spoiler selectors /actuators. In other systems the spoilers are operated via a control modute/cornputer (figure 5 1).

With reference to figure 51 - note the inputs to the control module - roll - speedbrake lever - feedback signal.

WARNING

I. Spoilers operate quickly and can cause serious injury.

2. They also have a fail-safe system which means that they will close automatically if either hydraulic pressure or electrical power is removed. (On some aircraft they will float upwards in flight if selected and the hydraulic power has failed. This is caused by the reduced pressure above the wing).

3. With the hydraulic system pressurised and the aircraft an the ground the spoilers will operate automatically if reverse thrust is selected.

Enhanced Lift Facility 7 - ... - .

I

Zln some of t he larger rnudcr r i +i ciaft"-i-bot:i afieroils will aut~rnzi icdy- szt ink the ' d r ~ $ ~ ' position for take-of4 and landing. This enhances the lift characteristics for that part of khk wing because (in effect) the ailerons are acting similar to flaps - whilst still allowing the pilot roll control via the ailerons.

I 1 I - - - 8 8

~ i gu r t 52 shows the location ok a typical droop actuator - it also shows the spring feel unit and the electritally operated trim system. As with most powexed controls t he trimming of the system is usually achieved by setting thc sys teh to. a.'riew neutral' - exckpt with many elevator systems where thc tail&ane is used as the trimming device.

Some aircraft such as delta wing aircraft (Concorde) are fitted with a set of control sudaces at the trailing edge of the wing. Having no tailplane these surfaces must do the job of elevators and ailerons - hcnce the term elevons.

When the control column is pulled back all control surfaces rise (and vice versa).

When the control handwheel is moved to the left - the left hand elevons rise and the right ones fall (and vice versa).

When the pilot puts both roll and pitch inputs in simultaneously the system 'mixes the two signals'to give a combination of both, eg aircraft climbing and banking to the left - control column back and to the left - all eIevons up but those on the left move up further than those on the right.

Flaperons

Combine the function of a flap and an aileron. Fitted to- the Boeing 777 (inboard aileron). Similar to droop ailerons.

ROLL CONTROL TRANSDUC

-TRIM SCREWJACK

Fig. 52 TYPKCAL AILERON ELECTRIC TRIM SYSTEM

TAILPLANES

In some manuals called stabilisers or horizontal stabilisers - the vertical stabiliser being the fin. The term stabilator is sometimes used for a slab tailplane.

The tailplane is designed to give the aircraft longitudinal stability about the lateral axis (module 8), but may be used for pitch control as well.

Tailplanes may be:

* Fixed - with elevators. On small aircraft.

k Variable Incidence PI). Also fitted with elevators. The tailplane may be powered (electric or hydraulic) on some aircraft or manually operated (screwjacks) on small aircraft.

x All Flying or Slab. Used as the primary flying control surface and therefore has no elevators. Used in place of elevators and often used in place of ailerons on fighter aircraft (Taile~ons). Will act together as elevators and differentially as ailerons - or a combination of both.

Variable Incidence Tailplane

Controllable trim tabs become less effective at higher speeds and totally ineffective on fully powcr-operated systems. A variable incidence tailplane overcomes these problems. "Phe complete ttarlplane pivots about a main hlnge bearing and is usually moved by an electric or hydraulic actuator.

An instinctive switch in the flighi deck enables the pilot to increase or decrease the tailplane incidence. The VI:tailplaule is more effective than a trim tab and prodyes less drag. i

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On small aircraft the tailplane ksmovkd Aanually.

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All Flying -or- Slab Tailplane

This is similar to the VI tailplanc except that it does not have an elevator. The tailplane is operated directly by the fore and aft movement of the control column and on large aircraft it is fully powered. On small aircraft it is manually operated and may be fitted with a trim tab. Trimming of the fully powered tailplane is by setting t h e tailplane to a new neutral - similar to ordinary powered control systems.

Advantages of this tailplane are :

(a) Less drag for the same control effect. (b) More rigid and less liable to flutter, (c) Simpler than an elevator and tab system. (d) More effective control.

blank

FRONT SPAR

,mi \ SPAR TAILPLANE HINGE

Fig. 53 VARIABLE INCIDENCE TAILPLANE

T A I L P ~ E PIVOT

ACTUATOR Manually operated on small sitcrak powered on large airwaft

Fig. 54 ALL-FLYING TAILPLANE

STOP NUT

BALL HI /' -. - MAIN ACTUATOR I MOTOR

\

TORQUETUBE

CABLES TO '1 1 MANUAL FLIGHT DECK TRIM WHEELS :

\\ I NUT I. hhachl trim R auto pllal Inputs are via I Ihe lrlm scrvomotar.

2. The ritain actuator operates the system vla a blutch, gear train a gear box.

39dndal opcralran 1s vla a hand wheel

I In !he flight deckwhich disconnecls - * the metn actuator.

/ CABLE DRUM 8 DISCONNECT SWITCH

AlTACHM ENT TO AIRFRAME

TRIM SERVOMOTOR (AUTOMATIC CONTROL SYSTEM]

Fig. 55 TYPICAL VI TAILPLANE TRIM SYSTEM

QUESTION: As a general knowledge question, can you think why tailerons are not usually fitted to civil airliners? (5 rnins).

ANSWER: The twisting force would be too high on the fuselage because of its length and high moment of inertia due to the engines being placed out on the wings (for most civil aircraft).

FOREPLANES

Not common, particularly an civil aircraft. These are similar to tailplanes in effect but are fitted to the front of the aircraft. Their main difference is that they can be designed to help prevent the aircraft from stalling. As the aircraft approaches the stall the foreplane stalls first and lowers the nose so preventing the mainplanes stalling. They also make the aircraft more manoeuvrable,

Foreplanes can be VI with elevators or can be slab tailplanes.

SPOILERS

These may not be considered as primary flying controls but on some aircraft they are connected to the aileron system and as such are part of the primary flying control system. They are normally situated on top of the mainplanes forward of the flaps. They may carry out more than one function but are generally classified as: Symmetrical; Differential; Ground Effect and Gust Alleviation. When extended they-dump Iift and create drag.

I An aidaft may have 6 or 7 spoilers per wing and they may have collective and individual functions. They a r e pbwer operated.

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Thc sbdilers may be operated by direct mechanical connection to the flight deck (fibre 57) or may be operated-via a computer (figure 58).

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I .

~~rnrhebical ' I Spoilers -

~ ~ ~ r a t e d - - s ~ m m e t r i c a ~ l y in flight to reduce the lift/drag ratio. This will incrcase the rate of descent and reduce speed. Sometimes used on automatic landing approach runs.

Differential Spoilers

Used in conjunction with the ailerons to give improved lateral control. When the aileron is moved up: (a) passed a certain angle, and/or (b) the aircraft is flying within a certain speed range, t h e spoilers extend on that wing. This creates drag and dumps lift, hence increasing t he desired turning effect,

Figure 56 shows the general layout of the control surfaces including the spoilers and figure 57 shows the spoiler arrangement of the BAe 146. The spoiler push/ pull rod operating system is connected to the aileron cable control system at the first quadrant in the mainplane (cable quadrant). The push/ pull rod connects the cable quadrant to the Spoiler Cam Box. From the spoiler cam box the pilot's input is feed to the servo valve of the spoiler hydraulic actuator - via a spring strut.

.*

During downward movement of the aileron the cam foIZower in the cam box moves in the non-effective portion of the cam track and the spoiler is not selected. After the first (approximatcly) 5" of cockpit handwheel movement to raise the aileron, the spoiler is selected to give a non-linear movement in relation to ailcron movement.

Ground Effect Spoilers

These extend automatically on landing to dump lift and increase drag. In terms of getting the aircraft ta 'sit' firmly on the runway on landing they play an important role together with the shock absorbers (oleos).

They operate when the aircraft is configured for landing with the weight switch* operated on the landing gear; the throttles are in their correct (usually idle) position; the spoiler selector in the 'arm' position; wheels are rotating (picked up by the anti-skid transducer); bogie rotation micro switch operated; md u l t showing close to ground etc. Not all of these may be applicable to all aircraft so check the AMM of the aircraft comcerned-. . - -----

r - , - 8 i

* The weight switch may b{ called (amdngst other things): 1 1 I I , I,

Weight On Wheels switch(W0W) - .~iibus. Ground/Air sensor - ~oeing:-- - .

. - squat switch. I ,

Weight switch. \ i \

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IUBOWO AILERONS

Fig. 56 CONTROL SURFACES - LOCATION

Fig. 57 ROLL CONTROL SPOILERS - BAe 146 I

Example - A300 Ground Effect Spoilers -.

Speed brakes and roll spoilers &re used when landing as ground effect spoilers. Deflection angles being 50" for all surfaces.

They automatically extend when:

* They are selected. ~r The aircraft is on the ground.

The ground effect spoilers are selected when the two following conditions are fulfilled:

* Speed brake control lever pulled upwards (when in t he RET position) or thrust reverser selected on one engine,

and * both throttle levers in the idle position,

- 52 - moodull 1 A-53

The 'aircraft on ground signal' is sent when the following conditions are fulfilled:

* Two main landing gear aft landing wheels speed is greater than 70kt. or

+? For landing:

1. Boogie beam rotation. 2. Shock absorber compressed (signal sent 3 seconds after

touchdown). 3. Radio altitude lower than 5ft (1.5rn).

. . . . - . - -

GROUND EFFECT SPOLlER9

Fig. 58 A320 GROUND EFFECT SPOILER LOGIC CIRCUIT

Automatic extension is achieved for an aborted take-off only when two main landing gear aft wheels speed is higher than 70kt.

Ground effect spoilers will remain extended during bounces due to the ground conditions logic circuit and as long as both throttles are in the idle position and pre-selection order fulfilled.

Ground effect spoiler retraction is achieved:

* Either by pressing the SPEED BRAKE control lever down (pre-selection cancelled). or

* By pulling one throttle lever out of the idle position.

QUESTION: With reference to figure 58. Can you work through the logic gates to check that the above text confirms the wiring diagram logic. Remember an AND gate must have ALL the inputs positive for there to be an output. An OR gate will give an output if ANY ONE input is positive.

ANSWER: If you have problems contact your tutor.

Gust Alleviation Spoilers - - . -. - .

-. . I-

i

These, are fitted to the A3 20 [and, 6thEf aircraitj and operate automadcaiiy to relievk in-flight gust loads. They hre a form; of active stability and give a more stable +d comfortable flight ahd reducestructure fatigue.

! When the aircraft is disturbed labout the longitudinal axis, gyros sense the movement and send a signal to d computer. If the pilot has not commanded this Aokcrnent the computer will know this (all the pilot's dontGl inputs being sent to bne or more computer$ on the A320).

I I b

The c&hputek will send a sign& to ahydraulic control valve to extend the spohers-on the .up going wing--thus dumping lift - preventing the upswing of the wing and helping to keep the aircraft level.

This is a form of Active Stability as it relies on the use of computers and wros and not on the aerodynamic design of the aircraft.

HIGH LIFT DEVICES

Strictly these are not classed as Primary Flying Controls but are included in this section because they are moveable surfaces attached to the leading and trailing edges of the mainplane. When extended they increase lift and drag. Trailing edge devices are called flaps and leading edge devices are usually called slats, slots and sometimes leading edge flaps.

FLAPS

When extended they increase the camber of the mean camber line of the aerofoil. Thus they increase the amount of lift produced at any given airspeed. Many flaps also extend rearward when they are lowered which also increases the effective wing chord length and effective wing area.

Any protrusion into the airflow causes drag and flaps are no exception. In most cases it is an unwanted by-product of their use, but on some occasions the drag produced can be useful in slowing the speed of the aircraft.

I t is important that port (left) and starboard (right) flaps operate together (s~~mrnetricdly). To this end they are connected together mechanically on most aircraft, though on a few they may be inter-connected hydrauIically.

QUESTION: What would happen if the flaps moved asymmetrically in flight? . -

In other words the f lapcd~ one side of the Aircraft moved and . -

I the flaps on the dtEG siae d?d not. (5 mins) . I

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ANSWER: ' I

If one side flaps were to rnokqinto the airflowmore t h k ;the other side, there (prduld be more lift created oh &at wingithan on

I the other. This w9ul-d c-aGse the aircraft t o roll. i'he roll might be I

1 - significant enough (depefiiing on the amoun6 of flap akyrnme try) to be uncorrectablel by the pilbt - unless he/ s'hd c a n get both

' I flaps up before td td aircrafti control loss. (An Lifcrait [BEA Elizabethan] crashed at ~onddn LHR carryinel +ow horses because of a 1inkage.failu-e in one side of the:flap system. The aircraft rolled violentlyr cmshed into a line bf-parked aircraft killing all the crew and the horses and writing off several other passenger aircraft.

The flaps are operated:

(a) Manually- Being connected by rods and Ievers to a handle in the cockpit, similar to a car handbrake (light aircraft).

(b) By electric actuators driving a common drive shaft. (c) By a hydraulic jack or jacks (split and plain flaps usually) - with

mechanical interconnection. (d) By hydraulic motors - the Fowler type flap - driving a common

driveshaft.

A simple hydraulic jack or actuator connected to a split or plain flap will be capable of lowering or raising the flap, with the port and starboard flaps being connected together by a mechanical linkage to prevent asymmetric operation.

When the flaps are power operated a feed-back system is used to cancel the selection signal once the flaps have reached their selected position.

Where Fowler type flaps are used (most large aircraft) t he rangc of movement is such as to rcquire the use of a large jack arrangement, this wouId be too heavy. In these cases it is common to operate the flaps using a drive shaft system driven by an hydraulic motor. The motor might be in the centre of the aircraft (or there might be more than one motor) with a drive shaft running along the rear of the port and starboard mainplane rear spars. At each flap location the lateral drive is converted into a longitudinal drive by a gear box arrangement. The flaps arc moved by a rotating screwjack arrangement (ball screwjack), which moves them back and down dong guide tracks, which are covered by fairings when the flaps are retracted.

RIGKT WNG 0mmIvE s TORQUE LrPElfER

m R I K L aSSYMmRY TORQUE LIMITER 1 /BUAKE

Fig. 59 FLAP OPERATING SYSTEM - BAe 146

Should asymmetry occur then detectors will operate a warning on the flight dcck and automaticdly stop the operation of the flaps.

Figure 59 shows the flap operating system for the BAe 146. Note the common drive shaft; duplicate drive chains to the lower gearboxes and the asymmetry brakes.

For more detailed information on thc operating systems (seIectoss, hydraulic systems etc) you should refer to the book in this series entitled Hydraulics.

FLAP OPERATION - GENERAL

Take-off and Landing

A 'flap-less' take-off may be used on some aircraft such as the A300. This is employed when the runway length exceeds 2000m (allowing for certain weight and weather restrictions).

On somc aircraft there is a 'long runway' flap position and a 'short runway' flap position.

For landing there may be 2 positions - one for approach and one for landing.

Flight Deck SeIector

On older aircraft this is marked in degrees. On newer aircraft i t is marked LANDING APPROACH, FINAL APPROACH, etc.

1 - ,-- -

. . ' . I \ ! I - - -

Flight Deck Indication i I '

Flap posiiion will be indicated e h e r on gccl"ockwork' gauge L bessyn s$stem, moving cbil, synchro system etc!oi-shomMbn a CRT screen. 1 1

I -. ... I

I ,

: i '>>, \ 1, \ ' 1 . Asymmetric ~rotection I I a ,

I ! 1 I Any flap kyrnmetry (leading edie or trailing edge) will indude a violent roll. Asymmetry-is -preven ted hy having .the-port and starboard '$laps. (leading cdge or trailing edge) mechanically connected (cables, push/puIl rods, drive shafts etc) . On some older system they may be hydraulically interconnected. This means that both port and starboard flaps will move together.

Should the flap operating system suffer a mechanical breakdown then a safeguard device is fitted to warn the pilot and stop the flaps moving.

Asymmetric detectors are usudly fitted to the outboard ends of the drivc system (one port, one starboard) and monitor the revoIutions/rate of movement of that side of the system. Their signals are sent to a comparator unit.

Should this show a discrepancy between the port and starboard flaps (outside a specified tolerance) then the operating system is shut down and the pilot warned. (Refer to the book in this series entitled Hydraulics).

Flap Load Relief

Should the flaps be lowered at excessive airspeeds then flap structural damage will almost certainly result. The flaps will be damaged, the mechanism may also suffer and any resulting debris may hit the fuselage, tailplane and may be ingested into any rear-mounted engines.

Also the aircraft may suffer from lateral asymmetry as well as loss of flaps for landing. This will result in a high-speed lading. Altogether a most unhappy state of affairs.

With a simple hydraulic jack operated hinged type flap, provision for 'blaw- back' can be incorporated into the hydraulic system. This is in the form of a pressure relief valve in t h e flap hydraulic down line - called a 'blow-back\alvve. This will allow the flaps to be blown back by the airflow (aerodynamic pressure) if left down after take-off. If the flaps are lowered during flight (at speed) then the relief system will prevent the flaps from going down too far and sustaining damage.

t I I 1 I

SELECTOR VALVEI

1 CONTROL UNIT

AIRSPEED UNIT

PILOT'S WARNING

Fig. 60 FLAP LOAD RELIEF SYSTEM

1 FLAP MAIN TROLLEY

MAIN TROLLEY T

FLAP FULLY UP

FORWARD TROLLEY TRACK \

TAKE-OFF POSITION - LONG RUNWAYS

i

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LANDING POSITION

Fig. 6 1 FOWLER TYPE FLAP - POSITIONS

With screwjack operated flaps (Fowler type) the aerodynamic loads on the flap are not transmitted to the flap operating motors so the above solution will no t work.

For screwjack operated flap systems the airspeed i s sensed by the aircraft's Pitot system. This data is sent to the Digital Air Data Computer (DADC). I t may also be sent to a flap load relief unit (airspeed unit).

If the flaps are selected down (or left down after take-off) and the airspeed is high then an electric load relief actuator is signalled to operate from a computer with air data sent to it from the DADC. (The signal may come from an airspeed unit dedicated to the flap load relief system - refer figure 60). This will change the geometry of the linkage between the flight-deck selector handle and the flap selector valve. This will cause the flap selectar valve to move to the retract position.

The flaps will retract (not necessarily fully up) and the pilot will get a warning. The flight-deck selector handle usually stays in the position selected.

TYPES OF FLAP

P-. Fitted to simpler smaller aircraft and gives about a 50% increase in lift for that section of the wing. Decreases t he stalling angle to 12" and moves the centre of lift rearwards so producing a nose down pitching moment. (Remember the normal clean wing stalling angle is 15".) When lowered fully produc&sa large amount of dr@. The - complete rear section of the wing moves down ;oh a simple hinge system. Sometimes called a camber flap.

m 8

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F&. 62 PLAIN FLAP

Split Flm-Giqes a 60% increase in lift with a stalling angle of 1 4 O . The lower rear section of the wing moves down. Gives a large amount of drag when fully down and produces a nose down pitching moment.

The Zap flap is similar to the Split flap but the flap moves partly rearward during lowering. The effects are the same but with a 90' increase in lift and a stalling angle of 13".

Fig. 63 SPLIT FLAP

Slotted Flap. Gives a 65% increase in lift with an increase in the stalling angle to 16" (which is good). The flap moves down and forms a slot between it and the wing.

- 60 -

moodull l A-61

This allows some air through the slot from the bottom of the wing to the top of the flap keeping the top side clean from eddy currents (boundary layer cuntrol), making the flap more efficient and producing less drag than with t h e previous flaps.

The double slotted flap is similar to the slotted flap except that there are two slots in front of the flap. Gives 70% increase in lift with a stalling angle increased to 18". Sometimes a triple slotted flap is used.

Fig. 64 SLOTTED FLAP

Fowler Flap. Produces 90% increase in lift (which is nearly double the amount of lift for that part of the wing) with a stalling angle of 15". The flap moves

I down- and back to effectively increase- the-wing area while producing a more cambered mean camber line. Fro+uces-a nose down pitching rnornenr.

I I I, I I

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' I I I

\ I

1 I I I

Cl I !

-2 1 i [- , /--TRACKSYSTEM I -. ---.-

6 \,

I / Fig. 65 FOWnER FLAP I I

I , , , ; I I I I

Double slotted Fowler Flap. 1 ~ 6 % increase Ji'n lift with an inkrease of stalling angle to 20". The double slot alI~w~s'~ii--f~orn under the fla+-tors&ecp the top surface clean of any turbulence. Treble slotted flaps sometimes used.

The hinged Fowler flap produces even bct-ter results.

TRACK SYSTEM

Fig. 66 DOUBLE SLOTTED FOWLER FLAP - /

TRACK SYSTEM 0(

Fig. 67 HINGED FOWLER FLAP

Krue~er Leading E d ~ e Flap. A leading edge flap which when hinged forward increases lift by 50% and stalling angle to 2 5 O , As is moves the centre of lift fonvard it causes a nose up pitching moment. Often fitted to the inboard sections of wings of large commercial aircraft.

SLIDES DOWN & FORWARD OR HINGES DOWN FROM UNDERNEATH

Fig. 68 KRUEGER FLAP

SLATS/ SLOTS

On some aircraft they are held in a Pxed position on the leading edge of the aerofoil. On most aircraft they are moveable. The moveable slats are intercmqected to prevent asymmetric -operation of the port and starboard sectidns. i f asymrnerric operation was iu uccur irl lligiit ik i i a uident rij:: would ehsue - as would happen if the flaps were to operate asymmetrically. Detectors are fitted to stop thel system if this were to happen.

! I

Most slats/ slots will move the centre of lift forward and produce a nose up , I . pitchibg moment. . I

TYPES OF SLATS/SLOTS I

-

Slotted. Win%-This is a fixed slot .&-the wing leading from the underside to the top side just aft of the leading edge. At high angles of attack air from under the wing rushes through the slot and sweeps the top of the wing clear of any turbulent airflow (boundary layer control). Increases lift by 40% and increases stalling angk to 20'. Somc extra drag at high speeds.

SLOT CUT THROUGH WlNG /

Fig. 69 SLOTTED WING

Fixed Slat. This is similar in operation to the slotted wing and fitted t o some small aircraft. Lift increases by 50% and stalling angle increased to 20". Produces some drag.

- 62 - moodull 1 A-63

Fig. 70 FIXED SLAT

Moveable Slat. This may be automatic in operation or operated manually from the cockpit, or operated electrically, or hydraulically. In automatic operation the slat is spring loaded in the closed position. At high angles of attack the negative pressure is felt by the slat causing it to pull out on a system of levers fr'orn the wing. The slat then directs the airflow over the top of the wing to sweep it clean of any turbulence.

When the angle of attack is decreased the negative pressure on the slat is insufficicnt to hold it out and the springs will pull it back flush fitting to the wing, - - - i--- - - -_ - -

. - - -. , I

The spring operated slat is normqly restiicied to (some) s m ~ l l aircraft. ' , I

I ! I I I I ' When operated the increase in kd will be 'about 60% with an increase ib angle of attack to about 22". I I -.- _

I I r-

I I ' I - I

I I SL~TM~VESFO~WARDONAPA~?OGRAPHLINKAGE I I I/; I /

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f-

**'- ! 1 :

S 1 ; -- - - - - .- -- --- I _ _ i

Fig. 71 MOVEABLE SLAT

On large aircraft the leading edge slats are selected out for take-off and landing along with the flaps with the pitching moment being neutdised.

They may be wound out using hydraulic motors and are fitted with asymmetric detectors that stop the movement immediately should asymmetric operation be detected.

The flaps and slats are usually operated together to increase lift (for that part of the wing) by up to 120% and increase the stalling angle t o around 30".

DRlVE ARM FULLY EXTENDED

P L

IN TRANSIT

FOLDING HOSE RETRACTED PIISITION

Fig. 72 KRUEGER LEADING EDGE FLAP

MECHANISM

Fig. 73 TYPICAL SLAT OPERATING SYSTEM

VARTATIONS IN AERODYNAMIC DESIGN

Vortex Generators

These are small flat metal plates fitted to some aircraft t o mix high energy air into the (sluggish) boundary layer ta give it more energy, They are usually arranged on the top surface of the wing at a small angle to the relative airflow, Usually they are all angled at the same angle but some aircraft may have each alternate plates angled in the opposite direction.

They stick up higher than the boundary layer (about 2 in. 15Qrnrnl) and cause the high energy free stream airflow to became turbulent and mix with thc low energg boundary layer - thus giving it energy and making it become more effective. They create form drag but reduce skin friction drag. They also tend to wcaken the shock waves and hence reduce shock drag.

I RELATIVE AIRFLOW

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Wing Feqccs I I I

I I These are flat metal plates (up tb !2 in [360h] high) and ftt+di parallel to thc free stream flow. They are fitted tn-help-prevent spanwise mbvernent of the air on swept-wing aircraft and may-bcfound -in front of control surfaces to increase their effectiveness.

WING FENCE 1 R

AILERON - Fig. 75 WING FENCE

- 65 - moodull 1A-66

Saw Toath Leading Edge

Fitted to the leading edge of some swept wing aircraft. Sometimes called a Dog Tooth, it brings the centre of pressure forward on the outer part of thc wing so helping to prevent the tip of the lower wing 'digging in' during a turn. I t also encourages t he boundary layer to move in the direction of the free stream flow. I t also helps prevent airflow separation at the tip - which is always a problem with highly swept wings.

SAW TOOTH

... . , , - . . .- . . - . .- 1

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, I , ,

: I Fig 76 SAW TOOTH OR DOG TOOTH i i . ,

Fig. 77 STALL WEDGE OR PUIED SPOILER

Stall Wedge or Fixed Spoiler

During the stall if the outer part of the wing stalled at the same time as (or sooner than) the rest of the wing then there would be turbulent air over the ailerons and the pilot would loose Iateral control. During a stall this can be dangerous as a spin might result.

To prevent the outer part of the wing from stalling first, f&d spoilers are fitted to the inboard leading edges of the wing to cause that part to stall before the outer part of the wing - thus the pilot will still have some lateral control even though the aircraft is in a stall.

Ca.nard Aircraft

Sometimes called a tail-first aircraft as the tailplane (stabiliser) is fitted in front of the mainplane and called a foreplane. The first powered flight was a canard configured aircraft and some modern high performance fighters are also built

-- this way. -.

I I - - . --

I The foreplane acts similar to a donventio~al; tailplane in providing longitudinal stability and control. They can as a slab!foreplane (no elivAtors) - i f both move up (increased angle of attqck) then tl$ aircraft climbs 1 yith thk control column pulled back - and vice-vefsa.-Moved differentially t&y provide roll control,. I-= -' -- .

i 1 ; I \* ', ! - - - - 2

I 1 \ ' I If they re fitted with elevators, \&n, with control coludn pulled back the foreplane elevators move down (with a tailblane they move up): foreplane lift increases and the aircraft climb&. 'It the .control column is phshed forward then the reve-r-se happens. - . - A + , i '1

I t has many disadvantages including poor aerodynamic stability, but one advantage is that it may make the aircraft difficult to stall as the forepIane stalls before the mainplane, automatically putting the nose of the aircraft down before total lose of control occurs.

Delta Winged Aircraft

These may be divided into two categories, those with tailplanes and those without.

Those with tailplanes have conventiona1 controls like other aircraft - flaps, ailerons, eIevators, rudder etc. Those without a tailplane are different - they have elevons.

Cancorde is typical of a delta-winged aircraft without a tailplane. It has one set of control surfaces at the rear of the mainplane that do the job of both the elevators and ailerons. They are called elevons.

When the pilot pulls the control column back bath elevons move up - and the aircraft climbs (and vice-versa) . When he/ she moves the handwheel to the right the right-hand elevon moves up and the Ieft one moves down - and the aircraft rolls right (and vice-versa) .

If he/she pulls the control column back and moves the control wheel to the right then both elevons move u p but the right-hand one moves up further than the left-hand one - the aircraft will climb and roll to the right.

The pitch and roll conh-01s are put through a mixing unit, which sums (mechanically) the two inputs to give the requircd control surface response. On a fly-bywire/ fly- by-light aircraft this function would be performed by a computes. Figure 78 shows the principle of hew the mixing unit works.

ROLL INPW

BELL CRANK A

BELL CRANK &

PUSHIPULL RODS

SUMMING LINKS

Fig, 78 SUMMING LINK - ELEVON SYSTEM

When the pilot puts in a soll command input, bell crank A ratatcs about pivot X I and causes bell crank B to pivot about X2. This will causc A 1 to movc down or up with A2 moving in thc opposite direction - moving one elevon in one direction and the other in the opposite direction. The summing links pivoting about B 1 and B2 respectively.

When a pitch push /pull command is put in, the torque shaft is caused to rotatc and move both 13 1 and B2 push/ pull rods in the same direction. (Each push/ pull red connection to the torque shaft by means of a lever).

This movement will cause both elevons to move u p or down together, with the summing links pivoting about A l and A2 respectively.

If the pilot rnovcs both the pitch and soll controls together then both inputs will be summed by the linkage to produce the required control surface movements.

Butterfly or Vee Tailplane

Some (usually smaller) aircraft are designed with a tailplane/ stabiliser with a very high dihedral angle (dose to 4 5 O ) . This means that it can double as a tailplane and as a fin for stability purposes and the control surfaces can double as an elevator and a rudder (ruddervator).

Fig. 79 BUTTERFLY TAILPLANE . --- I

Figure 79 shows the butterfly tailplane o f ) ~ e French built ~aiobin. When the

P I ruddefiators are both moved up (lor dowd) tbgether they ac ab an elevator.

When one moves up and the ot&er-one.&es down then t h dffect id similar to a rudder! I -- i 1 -. ., ,,

~yrnm&&cal movement of the *ddemator,s (far pitch contrdl) is caused by fore and aft movemerit' of the control ~o lumn. AsJImmetric opera$oh (for yaw control) is caused by movement' of the rudder bar.

I -__ -- i 1

I r \ 7

- - .- I-. 1 when the pilot puts a pitch &-djaw~o&mand in (control column and rudder bar), then summing links similar to those already described will ensurc that both ruddervators will be displaced in the same direction but one will move further than the other - producing both a yaw and pitch change of the aircraft.

Its advantage is that there is less profile drag (there is no fin) and production costs are reduced - lor the same reason.

Of course, the aerodynamic efficiency in terms of stability and control is not as good as a conventional tailplane and fin assembly - but you can't have it both ways.

INTEGRAL GUST LOCKS

Some aircraft are fitted with a flying control locking system so that the controls can be locked for parking/ picketing/ mooring the aircraft.

- 69 - moodull l A-70

Usually operated by a handle fitted in the flight deck and connected to moveable locking pins by a cable system. When operated the pins are pushed into the locked position by passing through holes in pulleys/quadrants which lock the control system and prevent its movement due to wind laads.

The system is so designed that take-off is impossible (by regulation) with the locks still in - sometimes by fitting the handle in such a position that the throttles cannot be moved forward unless it is released.

CONTROL POSITION INDICATING SYSTEMS

O n large aircraft the position of the control surfaces are indicated to t he pilot by an indicator gauge or a display on a CRT (Cathode Ray Tube). The transducers can he a variable resistor (potentiometer) connected t o the flying control surface - or linkage close to it. AS thc surface moves the transducer sends a dc voltage to thc gauging system dependant on the surface position. This voltage can be used to move a moving coil instmrnent or a dc ratiometer typC'iiis-tfnrnent, or it c a n besent to a-computer where itlis converted to a digitai signai and sent to a symb6i-gkiieraior io show a position dispiay url iile CRT. , I ,

For more information on instddentation you a re advised to read the book in this sbries entitled ~nstrument$. -- -

I - .

j I 1

I I ! I I I FLYING CONTROL RIGGING I t 1 I ' 1 :

The fliihg. contrbl systems havk be rigged from time to time, ie set up and checked so-that they carry out their function correctly.

QUESTION: When would the controls be rigged? Try and think of at least 4 occasions. (5 mins) .

ANSWER: 1. At the manufacturer's. 2. When stated in the AMM. 3. When a component in thc system is changed. 4. After a heavy landing or flight through turbulent air, 5. When any adjustment is carried out to the system. 6. If the pilot reported a fault with the system - eg lack of

range, high static friction etc.

The actual process of controI system rigging will vary from aircraft t o aircraft. It wilI vary an whether the controIs arc manual or powered. It wiIl also vary nn whether the system is operated by cables, push/ pull rods or by fly-by-wire.

I t is most important therefore to refer to the AMM for the actual procedure to be carried out.

The following paragraphs give a general outline concentrating on a manually operated cable system. Some variations are included but most systems rigging would be similar. The process is not too dissimilar to that used on trimming systems.

The general procedure is:

Refer to the AMM. Configure the aircraft for test. Set the control system t o neutral. Check cable tensions. Carry out sense check. Carry out freedom of movement check. Carry out range of movement check. Carry out any special checks. Carry out a duplicate inspection.

Do r elate-this process to your ,ewn-aircraft. - .. y

.- - +

-. r

I

The following paragraphs exparid ,on theXrn+n headings above: and are a I , m

general approach. , , ' ) 1 I

f / 1 I i I 1

1. Refer to the Aircraft Maintenance M k a l -With some tasks &is may not be the first thing to do -withsome control rig& i fahibst certainly $hould come first. I t may the *aft to be jackea-and in rigging position -it may not. i I

I ' I ! 2 . Configure the system. ~ h k AMM --. . .. wili - sbecify, for powefed controls, that

hydrauLic..and electric power I---- is -- to be 'on and certain ;systefns to be on - such as the Air Data Computer, Flight Management Computer etc. Pitot static systems might have to be pressurised. I t will list the equipment to be used.

Check flying control systems and associated systems for completeness and se~ceabi l i ty .

On large aircraft place warning notices that controls are being moved and check that servicing personnel are not working on or close to control systems. Remember, when powered controls are moved under power they can cause serious injury if anyone gets in t he way.

3. Set the Control Systems to Neutral - This may mean slackening the control cables and usually requires mechanical locks/ pins / devices to be placed:

(a) On the pilots control. (b) At the control surface end. (c) At intermediate links/ pulleys in the system.

Check that each component in the system is in neutral:

(a) PFCUsandartificidfeel(iffitted). (b) Flight deck indicators (if fitted). (c) Trimming systems. (d) Tabs and indicators. (e) Chains and cables are equally placed around sprocket

wheels/ pulleys. (f) Control surfaces align with trailing edge.

Note. O n some aircraft there may be upfloat or downfloat specified - check the AMM.

4. Check Cable Tensions - This may be carried out using a tensiometer (see following paragraphs) and adjusting the turnbuckles in the system. A n alternative method is to read the tensions from the cable tension regulator while adjusting the turnbuckles. In this case reference must be made to the ambient temperature and a graph relating temperature to

'-He tension regulator -. readhg. Ensure that aircraft has settled to local .- --

j ambient: temperarure. I I I

Visudly inspect the lay of kll cables that they are correct and not fouling anything. Check that d l Iturnbuckles/ adjusters are in safely. Remove all :neutral setting pins/ devices- and re-check tensions and neutral settings.

-- . .. -

I

/ ~ o t e . Item 3 would not apply to a push/ pull rod system-- but push rod lengths may be adjusted.

' I

I

5. Carry Out Sense Check - This will require electr id and hydraulic power on-a-powesed system. Tl?e fight deck controls are moved and a check is carried out at the control surface end to check that they move in the correct sense. Check flight deck indicators at the same time.

Remember on some large aircraft the spoilers may move asymmetrically when the ailerons are moved. On other aircraft the tailplane may move in response to elevator movement. So when checking primary flying control movement do check related systems operation. Check correct sense under autopilot command.

6. Carry Out Freedom of Movement Check - In general this requires the control system to be pulled through its complete range of movement using, say, a spring balance attached to the control column/rudder pedals. The force required to operate the controls should not exceed that value laid down in the AMM. If it does the system must be given a complete visual examination and the cause ascertained and rectified.

If manual reversion is provided on a powered system, check the system in manual as well as in power. Check the AMM on spring balance readings related to artificial feel inputs.

7. Carry Out Range of Movement Check - In general the piIot's controls are moved to their full range in both directions and the range of movement of the control surface is measured. It may be measured linearly using a rule or angularly using an inclinometer. If measurements are incorrect then range of movement may be altered using (usually) the primary control stops.

If appropriate, check controls in power and manual and in autopilot mode.

Check correct indications on the flight deck.

Lock all system points where previous adjustments have been carried out.

8. Special Checks - The manual will specify the checks to be carried out on all the equipment fitted to the controls. The aircraft may have to be configured so as to assimilate certain conditions, and checks carried out

- .. .- -- c.1 tfie f ~ ! ! ~ ~ . ~ , ~ ~ w a 9 -- e gp$cpriz?g: , . -.

',

! , \ j I : , * Stick shaker. ' \ I ! ',I ',

, , * i Stick push. I : 1 ; ; :

m ,

: / * 1 , , . ~ e t e n t / discohrject syste&s. I '- - I

i .

* Artificial feel. : 1 1.. . * r-- - Autopilot seqo land system operation. -

. * , i , , Tab systems, trim and ~ d ~ h trim. * ;Yaw dampers. 1 ' I , I

, m i * I I *. .. Alternative p@er suppies/ alternative operation.

Emergency sthid-by-syst&s. , . I I

8 .. .- .. . . .* .. .. . , System coniputerJ-s- op4ration. i'--' !

* Flight deck indications/ warnings.

9. Carry Out a Duplicate Inspection - All the parts of a flying control system are generally classed as VITAL POINTS (as defined in BCARs section A A5-3) and if disturbed will require a duplicate inspection. Duplicate inspections are required by BCARs section A A6-2, which defines the following:

(a) Control System - A system by which the flight path, attitude or propulsive force of an aircraft is changed, including the flight, engine and propeller controls, the related systems controls and the associated operating mechanisms.

(b) Duplicate Inspection - An inspection first made and certified by one qualified person and subsequently made and certified by a second qualified person.

NOTES

I. A duplicate inspection may be limited to that part of the system which has been disturbed.

2. A duplicate inspection must be carried out as soon as possible after the first inspection and before the aircraft flies.

3. If the system is disturbed during or after the duplicate inspection then the part disturbed shall be subject to another duplicate set of inspections.

4. All the work done must be recorded together with part numbers / serial numbers of components replaced. A CRS must be signed and entered/attached to the aircraft log book.

EQUIPMENT

TENSIOMmERS I

-" I - - -- .-. r nese are used z'or checking Lh? tenslurls oi cabit=s irl a i r crdi, cvrlir oi sy sic~rls,

including engine controls. They are normally used on unregulated systems only. I

here are several types available and the type to be used may, o r may not be, specified in the AMM.

I I - - - I ;

, I : I . ,

SME i yke I

1 I

The SME en sib meter is supplied. in various marks to suit different sizes of cable-. Each tensiometer will.take fmo sizes of cable - with two scales and the size being marked an each scale.

Note. I t is important that the correct mark of tensiometer is used otherwise inaccuracies will result.

Instructions for Use

(a) Fit the instrument where there is a clear run or cable. (b) Pull the pointer over to its stop. (c) Pass the cable under the right hand fixed pulley, then over the centre

pulley then under the left hand flxed pulley. (d) Ensure that the tensiometer hangs freely. (e) Run the tensiometer back and for the along the cable a few inches then

tap the cable until the reading settles down. The tension is indicated in Tbs on the appropriate scale.

5 CVVT SCALE

20 C W SCALE I

PIVOT

I I I I

Fig. 80 s B m ~ E 8 m i m s I o M m m i 1 Y I 8 ,

I - . _ I I '

I

Check \for Accuracy I I

I I I i '

' I I i i I I Before use, the position of the spring anchoq,age pin should ibo checked to see that it has not moved outside it5 &raved circle. ~f movernent!has occurred the pin should be restored t o its ~ r i~ ind -~os i r ibn . The tensio6eter dhould be checked at regular intervals by the manufacturer or at an approved standards room.

The Pacific T5 Type

This tensiometer is suitable far various sizes of cables using tables to convert the reading into tension vdues depending on cable size and using the correct size riser.

Instructions for Use

(a) From the chart supplied find the correct riser to use for the size of cable. Fit the riser.

(b) Check that brake is off. Move trigger away from case - this Iowers the riser.

(c) Placethecableunderthetwosectorsandovertheriser.

[d) Close the trigger - this raises the riser.

(e) Note the reading on the gauge. (If t h e reading is to be held operate the brake lever).

( f ) Open the trigger and remove the tensiometer.

(g) Convert the dial reading to lbs tension by reference to a calibration chart.

SECTOR BRAKE

\ LEVER SECTOR RISER

SCALE L Fig. 81 FITTING THE T5 TENSIOMETER TO A CABLE

Note. Each tensiometer has its own calibration chart. Make sure the calibration chart bears the same serial number as the tensiometer.

INCLINOMETERS

Used for checking the angular range of movement of control surfaces.

These are made by various manufacturers and may have a range of only 10 degrees or so or may have a range of 90 degrees plus. Accuracies range from lo to 1 minute (60 minutes = 1 degree). They may be mechanical or electronic in operation and the mechanical ones use a spirit level as the reference.

For more information on measuring angles refer to JARlEASA module 7 series of books in this series.

CONTROL SURFACE RIGGING BOARDS

These are supplied by most aircraft manufacturers and fit onto the mainplane / tailplanel s tructure and, using a scale, will indicate the range of movement of the control surface/control component.

May be fitted with red warning flags to ensure their removal after use.

RIGGING JIGS

Often supplied to fit onto the ~ontrols in the cockpit to rig them into neutral. The jig is attached to the cantto1 coIumn and the rudder pedals and a part of the cockpit structure.

RIGGING PENS

Si-~ppli~rl h$ tbe r n a n ~ ~ f ~ c t i ~ r e r m h e fi_ft?d.. into r n m p n n e t l t ~ 3 ~ a s -q~-]a-rlrar?ts, pulleys and idling links to l o ~ k ' + t r n ~ 3 o , f h ~ neutral warning flags to cnsure their repmvaI aft& hse.

- 77 -

rnoodull l A-78

CONTENTS

Page

The atmosphere The ICAO standard atmosphere

Low speed flight Definitions - 1 Aerodynamic lift

Definitions - 2 Lift augmentation

Aerodynamic drag ' -ThGF -f~-gr f ~ r r p q -- -

I

[ - Manoeuvres I

The axis Stability I I Dynamic stability I

Flutter - - -- ..

V ~ i a t i o n s in aircraft dcsignl- - -- High sbecd flight I I

Transonic speed L

I

~ u ~ e r s o n i c ) speed I .

i I

Wing plan forms -

Kinetic "heating - Stability and control at supersonic speed Jet engine intakes

THE ATMOSPHERE

The parameters of air (pressure, density and temperature) vary considerably both with height and geographical location around the world. The air is made up of approximately 2 1% oxygen (02) and 78% nitrogen (N) by volume, with the remaining 1% being made up from other gases. The ratios of the gases vary little with height although the moisture content drops with increase in altitude.

Because of these variations and to allow standardisation and calibration of instruments and engine performance figures etc, a Standard Atmosphere has been devised.

This allows engines to be test run in almost any ambient conditions and the performance figures adjusted to standard atmospheric conditions - allowing the performance of one engine to directly compared to that of another.

Pitot-static operated instruments can be calibrated using the standard atmosphere and they can be set for flight using t he same parameters leg QNE on the altimeter).

. . - . , -- - - I -. -. . I-_ , -

THEICA~STANDARDATMOSPHORE

I 1 I

~t has bekn shown that the maih hariableb (pressure, tempeiat,ure and density) of the standard atmosphere relate well-with actual average <dues observed at

8 - - about latitude 40' N. This stand&717ttmosphere is regardedas a reference basis fhr 'certain parameters in free air (excluding those dependent on water

I I vapaur) :). # I I I I ' I

1 I I ' I I '

The standard-st he TCAO standard-atmosphere states that: ~ h k air is assumed to be dry. The-pressure at sea-level is 10 13.25rnb (rnillibw). The temperature is 15°C and the temperature lapse rate is 1.98"C per lOOOf t (feet) up to a height of 36,000ft where the ternperature will remain constant at -56.S°C to 65,800ft. The value of "g" (gravity) is given a uniform value of 9.8lm Jsec at sea level.

For heights above 65,800ft the ICAO law states that the temperature lapse rate is approximately +0.303"C per 1 000ft to -44.6"C at 105,000ft.

Effectively that means that the ternperature falls with altitude at a rate of about 2°C per 1000ft from 15°C at sea level to 36,000ft where it holds almost stoady at -56°C until about 36,000ft where the temperature starts to rise.

Pressure. Measured in Pa or psi or mb). Starts at 10 13mb (14.7psi) at sea level and falls at a non-linear rate with altitude. Losing most of its value at the lower altitudes so that at 18,00Oft., for example, the pressure is halved to 506mb.

These pressure readings are absolute pressure readings. If an ordinary pressure gauge is open t o atmosphere it will read zero.

If it is used to check a tyre pressure of, say 30psi, it will read 30psi, but the pressure in the tyre is in fact 30psi above atmospheric so the absolute pressure in the @re is 30 + 14.7 = 44.7psi. The tyre pressure as measured by the gauge is called gauge pressure. I ts absolute pressure would be gauge pressure plus atmospheric pressure = 44.7psi absolute.

Density. Defined as mass per unit volume (kmJm3). Starts at 1.2kg per cubic metre at sea level, and falls at a similar rate to pressure. I t s rate of change is non-linear which means the graph is a curve and the amount by which it drops changes with height.

At a given height density can change depending on the temperature and the relative humidity (RH) . It the temperature drops density will increase and if t h e RH increases the density will decrease. In the standard atmosphere the drop in pressure with altitude offsets any tendency for the density to increase because of the d r ~ p in temperature.

Temperature. Starts at 15°C at sea lcvel and falls at a rate of about 2°C (1.987 actual) per 1000ft to 36,OOOK 1 1 km) . This is called the lapse rate which is !kAe& f i $his altitr-~de (the gra:fih i s Y straight lin ej. T t remlain ~ steady . a t 9 hmlt minus 56°C to 65,000ft where ,it tp rise. I

I I

' I

~ u r n i d i b . The Relative ~ u m i d i t y (RH) fails with altitude. T ~ S is us<aily taken 1 ill hold' af a as a $eicentage of the total rn&hurn humidity that the air w

partidular temperature. (For air hiditioning purposes watkr ispray is added to I-- the a+ {ntering the cabin at altitude to'c&nter , the effect ok thedry ambient air

- low N). I

I

RH is the amoint of moisture that is in a volume (m3) of aii compared to the rnqimtirn ahount it will hold (ie! when i t i s saturated) - at 'thht, temperature.

-. -.. -.

The higher the humidity the less dense the air and as density is a function of lift so lift decreases with an increase in humidity. This means that with some airfields located in humid climates, large/ heavy aircraft may wait until nightfall when the RH drops before taking off.

Absolute Humidity (AH) is the mount of moisture in a cubic metre of air {at a specific temperature) in grams per cubic metre. A similar measure to R3-I but gives the value as a specific amount, eg 2.7g Jm? Not such a helpful parameter as RH.

When air is continuously cooled there comes a point when a temperature is reached which causes any moisture present to condense out - this is called the Dew Point. When dealing with breathing 0 2 , if it is considered contaminated its moisture content can be checked by passing it at low pressure across a mirror in a hygrometer. The mirror's temperature is gradually lowered and when the dew point is reached dew forms on the mirror and an electrical connection across the mirror causes a needle to move.

moodull 1 A-81

At this point the temperature of the mirror is noted (Dew Point temperature) and from tables the moisture content of t he oxygen can be established. This value is compared to data supplied by the oxygen manufacturer (BOC for example). There are also electronic instruments available for determining the Dew Point of a gas.

Vapour trails or Contrails from high flying alrcraft are caused by moisture in the atmosphere.

Most vapour trails come from the efflux of jet engines due to the condensation of the moisture in the efflux as the hat gasses cool at altitude. When there are no vapour trails the aircraft is flying through air with very little moisture in and/or it is flying at a low enough altitude so that rapid cooling of the efflux does not occur.

Vapaur trails from wing-tips are caused by the condensation of the moisture in the air as it looses pressure (and cools) by spilling over the wing-tip from the high pressure side on the bottom of the wing to the top low pressure side of the wing. In some conditions they can be seen coming from the ends of flaps.

-8 - - -- --- .. . m L

' I 'I j ! I L ~ W SPEED BLIGHT

I i I ! I '

I j DEFINITIONS

The follo&ing definitions should y-~emekbered: I I 8 8 ! I

\ \

~irspekd; The spebd of the aircfaft througb @e air. Not usu$ly the same as ground speed, fhf example: If tfie Faircraft ii's flying through the1 air at 140kts with a tail-wid of 30kts then i& ground -.- - - spied will be 140 4 310 = 170kts. If the aircraft turnsxound and fliesrinto_the wind then its ground speed . . is 140 - 30 = 1 1 Okts. Hence the reason why aircraft always land into wind - the actual landing speed is reduced by the amount of head wind.

NOTE. The knot (kt) is 1 nautical mile per hour and 1 nautical mile - 1.15 statute miles. 1 kt = 1.15mph = 1.85km/ h = 0.514rn/ s

Incompressible. Flow below sonic speed (the speed of sound - 762rnph at sea level) is assumed to be incompressible. Not strictly true but close enough for most practical purposes. At supersonic speeds the air is compressible. While supersonic speed is considered as Mach 1 and above, compressibility effect starts to make itself felt at about M0.7. BernouIIi (how air flows around an object) assumes incompressible flow and therefore only applies to subsonic speeds.

If the aircraft was to fly at speeds approaching the speed of sound (MCRIT and above) then shock waves would cause a large inarea= in drag (as well as buffeting etc). This process starts in the transonic speed range.

Laminar Flow. Fluid flow in which the streamlines maintain a uniform parallel separation with no turbulence. Shown as parallel straight lines on a flow diagram. Generally considered to be a good condition which reduces drag.

Streamline. A n imaginary line marking the path of a particle of fluid from one point to another especially in laminar flow. Usually shown as a line with an arrow indicating direction of flow.

Turbulent Flow. Random motion of fluid with unpredictable fluctuations and vortices. There are no streamlines present. Will cause considerable drag.

AERODYNAMIC LIFT

The Venturi Effect

When air passes through a tube which contracts to a throat, it can be shown by a simple experiment (Bernoulli) that the air pressure (called the Static Pressure) drops at the throat - where the air velocity is at its fastest.

8 --- 7-

If werey;esent t he flow by drawing ~ ~ e a k l i n e s of the flo% ofGFrthiflhrough such a ventuti we see that thc streahlines md, f4rced together wh&e the'speed is

1 greatesd and the pressure is lowest. ! / l 1 I ! I I I I

I ' 'I / I

I , 1 :

, I STREAMLINES I I

I I 4 '

I

I AIRFLOW DlRECTlON

Fig. 1 AIRFLOW THROUGH A VENTURI

If the two sides of the venturi are free to move they will move together as the pressure drops (a simple experiment can show this to be true). If we reverse the venturi sides and we put the low pressure areas on the outside of the shape then we have the makings of an aerofoil (figure 2) .

The Aerofoil

The wings of aircraft are of an aerofoil shape, as indeed are the tailplane (stabiliser) and fin. Other aerodynamic components are also aerefoil shaped to include propellers, flaps, slats, control surfaces, aerials (antenna) etc.

HH;H VELOCCPl AND

LOW PRESSURE AREAS

# Fig. 2 SYMMETRICAL AEROFOIL

Figure 2 shows a symmetrical aerofoil. Not common for aircraft that fly below the speed of sound, but some aircraft are in Iact fitted with these. The wings arc attached to the fuselage at a small positive angle (angle of incidence) so that they will create lift, but the more usual aerofoil for low speed aircraft is asymmetric with a well rounded leading edge and a straighter or slightly concave bottom surface (figure 3).

PI.*- Ilt: - "f"ii seGtion of tlie fiii qf z i ~iyidart is ilS-uaiiy s,, rlllXlci~ic.i;ii- i ill a Wcii rounded leading edge. The chord liiie isset' in-line with the fuselage longituaihd datum line, but some single engined propeller driven aircraft may havc the fin set at a small angle, on the fuselage to help offset the effect of thc rotating Slipstream from the prdpeller.

I

Some tailplancs have an aerofoil sektiCfi which give 'negative9 lift. In .other words the most cambered surface/ is on thk bottom with the flitter surfacc on the top: vsed to help balance the four forcks)acting on the aircraft. More of this later. . : I

I . ' . - ' I.. ' ,

The top-surfaceof an asymmetric-aerofoil for the wings usually&as a good convex tap camber with the bottom surface being nearly flat, or in some cases having a slight concave surface (figure 3).

NEQATWE ' 1 1 1 I . PRESSURE

AIRFLOW - Ttm Unes am drswn BO ~ c 1 o r Ilnaa and ars oblalcd by fitting manoma!wbrbss !a each pdni within 91s aeroloil model In a wlnd hmW.

POSITIVE PRESSURE

Fig. 3 PRESSURE DISTRIBUTION AROUND AN AEROFOIL 43 I b l

This type of aerofoil will produce a negative pressure on the top surface and a positive pressure on the bottom surface, thus creating lift. M o s t of the lift is created from the top surface with about a third being created from the bottom. The Centre of Lift of the aerofoil is about 1 J3rd chord distance from the leading edge.

The total air reaction can be organised into its component vectors - lift and drag. The lift vector always acts at right angles to the airflow (or free-stream flow) and the drag vector always acts at right angles to the lift vector and in line with the airflow (figure 4).

Fig. 4 ~ I F T A N D ' D ~ A G VECTORS i I <. 1 . I --

T WUNG I

EDGE WAKE

Fig. 5 AIRFLOW AROUND AN AEROFOIL

Figure 4 shows how the lift and drag vectors are summed to give the Total Air Reaction and figure 5 shows the airflow around the aerofoil.

When an object moves through the air it sends out pressure waves forward into the airstream. These pressure waves warn the oncoming air of the object's approach, and as the air gets near to the object so the air will start to move out of its way. This will cause Upwash in front of the aerafoil with most of the un- coming air moving upwards. The air over the top surface will speed up (Bernoulli) and the air under the bottom surface will slow down. This means a drop in pressure on the top surface and an increase in pressure under the aerofoil.

As the air leaves at the trailing edge so will move down. This i s called Downwash.

I t can be shown that the lift produced by an aerofoil is related to:

* I t s shape and angle of attack (CL) * The air density (p) * The air velocity (V) * The pIan wing area (A)

Lift is calculated by:

where

CL This is found by experimentation and is related mainly to the aerofoil shape. A higher Ct means greater lift. I t is just a number

- - - which is called a coefficient.

p , is ihe air density j i:Zk~,rnz at sea ieveij. The iower the-aititucie of 1 the aircraft the gr&ef i~fie~dknsity and the myre thP lift is

1 I created - dl other pkrimeters, bking fmed. ! I

V I s the air velocity. With an increase in speed more lift is cre:ated. In I I fact the lift is related to the s@are of speed. S o if the speed is

doublcd the lift is irideased fobr-fold. I ; A Isrelatedtothewi~gp~~-~e~,Ingeneralthegreaterthewing area the greater the1 lift. In some. equations this pay be designated

' ' , ( I I I . I I

I I

, 8

I t I I

The equation,%pV2 is sometime; ddled-the dynamic quatioh h d .. is given the designation- q: That is q = %p~ar - - - -

q is used in many other calculations including those related to Pitot static instruments and the calculation of drag - as we shall see later.

Some Mere Definitions:

Angle of Attack [AoA). This is the angle between the chord line of an aerofoil (mainplane, tailplane or fin) and the free-stream flow. In various manuals (pilot's notes in particular) it is called the Alpha angle (a). If a symmetrical aerafoil is given a positive angle of attack the speed of the airflow over the top surface increases - which produces suction, and the speed of the airflow under the aerofoil decreases - and the pressure increases. Hence lift is produced.

For an asymmetric aerofoil lift can be generated at zero angle of attack and ' kven at small negative angles of attack. ex

- 7 -

moodul I I A-86

ANGLE OF A n m (THEW)

Fig. 6 ANGLE OF A V A C K

This means that for a given airspeed lift increases with an increase in angle of attack - but only up to the stalling angle, usually 1 5" for a 'clean' wing (flaps up, slats in, spoilers in etc) . This increased angle will also increase upwash and downwash - and drag of course. More of this later.

Angle of Incidence. The angle the chord line makes with the longitudinal datum line of the aircraft. Fixed for most aircraft wings but variable for many tailplanes. The angle of incidence may change from root to tip on the wings. If it increases from root to tip it is said to Wash-in, if it decreases it is said to Wash-out.

- - A - - .-. .

I .. cent& bf pressure. All the pr&~siif~diffeAnces between the tap-and bottom surfabes of the aerofoil c a n be bdded together to produce tLe i ~ o t a l &ir: Reaction which chn be considered to act a: a p i n t c,hled the ~ e n t r c of ~ressur& (C of P).

I I ' lf 1 1 A s the gngle of attack increasds and fie&essure distribution changes, the positibn of the C of P moves f o k d ~ d s u a l l y reaching a pqin<about I/- chord length from the leading edge at the stal1ing:angle. After the stalling angle has

I been p+ssed it mbves rapidly b-ack to about mid-chord position. I

I I I

chordline. A'stdaight imaginary line joinini the centre of curvature of the leading edgeof an aerofoil section to the trailing edge. I _. 1

Downwash. An area behind the trailing edge of an aerofoil where the airflow tends to movc downwards.

Fineness Ratio, The ratio between the maximum depth of the aerofoil and the chord length. Thin' wings havc a high fineness ratio. 'Fat' wings have a low fineness ratio.

Mean Aerodynamic Chord. Similar te the Mean Chord. I t is the chord 01 an imaginary wing of constant aerofoil section producing the same forces (lift and drag) as those produced by an actual wing.

Mean Camber Line. A n imaginary line drawn from the centre of curvature of the leading edge to the trailing edge of an aerofoil, but equidistant from the top and bottom surfaces. This i s the same as the chord Line on a symmetrical aerofoil but will be curved on an asymmetric aerofoil and not the same as the chord line.

TURBULENT FLOW

LONG WAKE

Fig. 8 AEROFOIL STALLING

There is a Transition Point on the top of the aerofoil towards the rear where the airflow changes from laminar to turbulent flow, and as t h e angle of attack increases so this point will move forward.

At approqirnately 1 5" angle qf- attac-k- (with a clean wing), !the airflow can no longer r6ain laminar on the'-topsurface' *of the acmfoil h d it will break away and &dome turbulent. This will destroy tGy lift in this are4 +d the drag will increase sharply. At this paintthe aerafd,l is said ta stall. The acraf+ilis no longer 4ble to support the aircrakt weight (&though there is still some lift) m d itwilllldoseheightrapidly. 1 - -- . '

I

I I ! .. ! I

I i i-- i L - -

The Artraft will descend with thk altipeteiwinding back q ~ i f i k l ~ ; The pilot can do ndthkng to prevent this andl rdcovery itivolves pushing t$e throttles forward (more phcr) , &shing the contrcil column fhrward (putting the nose down to get m'ork - airspeed and reduce the angle of attack). - -

I

-- ' - .- - L A . - - - - . . - - - --<

When the normal airflow pt te rn around thc wing is re-c Atiblished the pilot pulls back on the control column and raises the nose of the aircraft. This process losses a lot of height and stall tests are carried out with sufficient altitude to allow recovery.

Stalling is usually accompanied by buffeting (due to the turbulence) and sometimes loss of control and possible engine pxoblems. Lose of control can be caused by turbulent air passing over the ailerons and sometimes the elevators (the turbulent air coming from the wings).

With rear mounted jet engines the turbulent air from the wings can sometimes cause the engines to stall (turbulent air in the intakes causing the compressor blade to stall). With stalled engines and ineffective elevators the aircraft is in very serious trouble. This is called a Super Stall with the aircraft falling in a nearly flat attitude with the pilot having few recovery options.

I

Increasing the angle of attack continuously (without increasing power) until the aerofoil stalls is not the only way an aerofoil can be stalled. I t can be stalled by gradually reducing the fonvard airspeed whilst holding the aircraft in straight and level flight.

A s the speed (Indicated Air Speed IAS) deceases so the aircraft will fly slower and to maintain height the angle of attack will gradually have to be increased. At some point an IAS will be reached where the wing is a t the stalling angle of attack, the wing can no longer support the weight of the aircraft and the wing stalls. This is sometimes called the stalling speed and is important when calculating the landing speed of an aircraft. In general the lower the stalling speed the better.

It is interesting to note that the IAS at stall is the same at all altitudes.

Note that on most aircraft there are stall warning devices fitted in the flight deck which include oral and visual warnings, devices that shake the control column (stick shaker) and stick pushers (to put the nose of the aircraft down, but the pilot can overcome this if helshe wishes) - all to warn of an approaching a stall. More of this later in the module.

Remember, the angle of attack is the angle between the chord line and the relative airflow. Aircraft with powerful engines can climb a t almost any angle relative to the horizontal, but the angle of attack must always be lower than 15" for the wing to be in a n un-stalled condition. --- -

- --

i 1 Wing Loading is the weight (mass) of the &?raft divided by thk gross wing area. Gross wing area is taken as the total wing area in plad vkw including any part of the fuselage in-between the wings. The mass of the aircraft is taken as the mass at that instant. Aircraft wytth a low wing loading (in general, light

-- I -- aircraft with large wing areas - gliders for example) have a lpw~r_stalling speed which k e a n s lower landing speed's. I

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Aircraft with high wing loading,swch as fighters have high l$nding speeds and high stalling speeds. - I --

- - - - - - --

The Lift Curve

Graph 1 shows the lift curve for an aerofoil. At about -4" (off the graph) there is no lift and as the angle of attack (AoA) increases so the lift increases until a t about 15" the wing stalls and the total lift reduces - even if the AoA is further increased.

The graph shows that some lift is created at 0' AoA (for most asymmetric aerofoils) with the lift coefficient starting at about 0.02 and peaking a t 1.2.

At the point where the wing stalls the wing losses lift rapidly and the aircraft looses height. Buffeting might occur due to turbulent air from the wings hitting other parts of the aircraft such as the tailplane and control surfaces. One wing might stall before the other making the aircraft roll whilst losing height.

5 L

LIFT COEFFICIENT

ANGLE OF ATTACK "

GRAPH I - ANGLE OF ATTACK (a) AGAINST LIFT COEFFICIENT (CL)

The pilot may have difficulty in controlling the aircraft. If the outer part of the wingstalls before the inner p r t - -. - then - - - - - the -turbulent air from the stalled area of fie!wing&odd alieci tiie diCi-Oiis. mi---- lrlc FI.LUL -I-& W V U ~ U 1-1 I--- lU3L - ~ i i 1u.u L"LLLLU.L ,---CLZI.J~~:-- U U I L L ~ ~ CL- L L I ~

,- -. -.

stall @d this could lead to thd &craft apbroaching a spin p~casd t state of affairs. ! I i i I

I I '

i 1 1

To hdpi to prevent this a fixed barldl' 'stall spoiler' is fit+d to the leading edge bf 'the inboard part of thelw~ngGi sbme aircraft. This Lauses the inboard part df ihe wing to stall before itfie outet\p&-t during the s td~l~rocess , so helpi+g the pilot ,to maintain roll: control &bing the stall. i

, I I I I 1 , I I I ' I

This means that when the air=)+ stal1s.h is the inner that stalls and the outer sections &ill have lamina air-flowing over them and-the ailerons are still eff& tive-, - - - --A

- . I ". . . - __-<- . ..

On swept wing aircraft, if the outer wing (towards the wing tip) stalls before the inner wing then this lose of lift may cause a nose up pitching moment. This is because this part of the wing may be behind the aircraft centre of gravity on the longitudinal axis and this lift element would be causing a nose down couple. With it removed (as in an outer wing stall) the nose may pitch up momentarily.

LIJV AUGMENTATION

Lift is increased for landing and take-off by the use of leading edge slats and trailing edge flaps. The whole idea is to give the aerofoil a more curved mean camber line. Birds can do this by changing the shape of their wings and the feathers slide over each other similar to platelets. It is not really a practical preposition to try and change the shape of a metal wing (although experiments have been carried out). So the best way, so far, is to effectively "bend" the front and rear part of the aerofoil down - to fit leading and trailing edge high lift devices.

Lift can be increased in flight by:

* Increasing the airspeed - throttles/ thrust levers. j: Increasing the angle of attack - elevators, all-flying tailplanes, all-

flying foreplanes etc. * Increasing the effective camber of the mean camber line - flaps and

leading edge devices. * Increasing the stalling angle - some flaps and leading edge devices. * Increasing the effective wing area - Fowler type flaps. * Using the ailerons as flaps. Using the ailerons in a combined

roll/flap mode (called flaperons). Setting both port and starboard ailerons partly down for take-off and landing, whilst the pilot still has differential movement for roll control. Set automatically when aircraft configured for landing/ take-off,

Lift augment~tion generally refers to leading and trailing edge devices, which means flaps, slats etc.

Y T T T ' A 'CieiCrl w i n g --

1 - - -

Figure 9 shows a section of a 'clFan7 wing, every thing rnovkable and attached to the wing is either in or up - landing ge&, flaps, slats, spoileks etc. Notice the Mean Camber Line (MCL). If we ,can make the MCL more cambered or make it longer (bigger wing area) then lift will b? increased.

I : - - ~ I

MEAN CAMBER LINE (MCL)

Fig. 9 A 'CLEAN' WING

Figure 10 shows that the Effective MCL (EMCL) becomes more cambered when flaps are lowered, which means lift is increased - by about 60% in fact. The stalling angle is not affected significantly. The drawing shows a split flap, but the same is true for all types of flap with the Fowler type producing the best EMCL.

- - - _

Fig. 10 SPLIT FLAP

- 13 - rnoodull l A-92

Figure 11 shows a Fowler type flap which not only goes down when selected but also travels rearward (on tracks). This means that the wing area is increased as well as the camber of the MCL. The increase in lift for this type of flap is up to 90% - that means, nearly double the lift for that section of the wing.

The double slotted Fowler uses an additional small slat type aerofoil in front of the main flap to direct air over the top of the main flap to help to prevent it from stalling - similar to a slat. This small slat allows higher pressure air from the bottom of the flap to wash over the top surface washing any eddies and stalling currents away. This means that this type of flap produces up to 100% more lift with an increase in the stalling angle to 20°.

Fig. ~ O E R TYPE FLAP [ : - -- I r - - - -

, , - , . ,

I ,

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-- ' : ,

, . . - - - - Fig. 12 D ~ ~ L E - . S L O T T E D FOWLER FLAP.

Some Fowlcr flaps have an additional hinged surface attached to the trailing edge that moves down when the flaps are lowered, similar to an additional small plane flap - thus increasing t he camber of the mean camber line still further and increasing the lift.

Figure 13 shows a Krueger flap. I t is type of leading edge flap that is hinged forward to increase the camber of the EMCL, and i t also increases the wing area a little.

%.-**

Fig. 13 KRUEGER FLAP

Popular on the inboard sections of mainplanes of many large airliners, it increases the lift by about 50% with an increase in the stalling angle to 25".

The slat in figure 14 increases the EMCL slightly as well as the wing area (slightly) but its main advantage is that it controls the boundary layer on top of the wing. At high angles of attack the air flows through the gap and is directed along the top of the wing. This airflow 'sweeps' any turbulent air away and holds the stall off until about 22". The lift increase is about 60%.

AIRFLOW SWEEPS TOP OF AEROFOIL CLEAR OF TURBULENCE

Fig. 14 THE SLAT

- - , I

On some &all aircraft the slats may-be f i e d (rare); on other smal1,aircraft they may be automatic - sucked out a t high angles of attack against a spring. On large aircraft they are powe(ed and selected by the pilot, often when the flaps are selected via a single levei- in the flight deck. I

1 I I

They a r e fitted to most large aircrgt with the most cornrnoncdnfigtu-ation being -

Krueger flaps on the inboard sehions of the wings and slat$ on the outboard sections. They are all selected dhen flaps are selected (by the pilot).

1 1 I 1 I

On some older aikcraft the wing day be pf the Slotted type (rare). The wing has a slot initconnecting the bottom surgce P- to the top surface. This allows air to pass through from the bottom-to the top. The action is iden6Ci.l to that of a fixed slat in that, a t high AoA, the air passes through and is allowed to sweep the top surface clean of turbulence thus putting off the stall to a much higher angle.

On some experimental aircraft Blown Flaps are used. This entails inbuilt compressors or tapings from the jet engine and ducting to blow air downwards from the aerofoil trailing edge creating, in effect, a flap.

Combination of Slats and Flaps

When slats and flaps are deployed together, which they usually are, the increase in lift is about 120% (well over double) and the stalling angle increased to about 30" (again, double the normal stalling angle).

Symmetrical Operation of High Lift Devices

All trailing edge and leading edge lift devices must operate symmetrically (ie the flaps/ slats on the right wing move at the same speed and the same distance as the flaps/ slats on the left wing).

If a ~ y r n m e ~ c operation was to occur then a violent roll would ensue - which would be uncontrollable (it has happened, and it has caused fatalities).

The aircraft i s fitted with systems to prevent asymmetric operation such as a common drive system for port and starboard flaps/slats and an asymmetric detection system that will stop the systems immediately any asymmetry is detected (and warn the pilot).

Landing /Take-off Configuration

Modern aircraft have computers to monitor the aircraft's configuration during landing and take-off and if something is amiss, to give a warning on the flight deck- G s t A ,- exactly what is wr3iigXjfst3ms - that are monitorEd-far correct -- - confi&i.ation can include: -\ b , r--

\ \, I i I * nap/ slat positions. 1 . 1 I j 1 1 I 1 I * Spoiler positions. I 1 I 1 I I

8 * Auto aileron droo$. ' cl //

* Landing gear do& Fd-lGCked. I , I - - I . * 1 Auto-brake set. I 1 . I i 1 :

I * Engjne parameters set. r I I ' . * Erlgine bleed inhibit' (take-o f;, $mottles open).

I 1 I ' * ' -- -Auto-throttle set, I ! .- ,

-*- - -Weight switches: opesation.-- -- -

* Cabin doors shut (take-off). * ILS (instrument Ianding system) monitored (landing) ,j. .X Auto-land system monitored. * Some aircraft can automatically receive airfield barometric and

visibility data. * Radar a1 timeter. * Aircraft transponder (providing aircraft ident and altitude to ATC). + Ground mapping radar. * Autopilot set. * Auto-trim set.

I t is interesting to note that auto-trim (normally from the autopilot) operates in fast mode when flaps are lowered (aircraft near t he ground and speed is important) and operate in slow mode when flapslslats are in and housed.

AERODYNAMIC D W G

Any device moved into the airflow will cause drag and this applies to flaps and slats, so they are designed to produce as little drag as possible. But if flaps are lowered to a large angle then they can be used to reduce the speed of the aircraft, ie they are designed at this angle to produce drag.

So when the pilot wants a s much lift as possible with as little drag as possible then the flaps are only lowered a little. So for take-off the pilot moves the flap lever (shaped like a flap so it can be identified by feel if necessary) to a detent marked 'take-off' and the flaps lower part way.

When landing, the pilot moves the lever to the 'land' position detent which causes the flaps to move further down, in some cases nearly a t right angles to the airflow. At this position a large drag force is created which slows the aircraft and lift is also created which helps to reduce the stalling speed and hence the landing speed.

For small aircraft the flap lever is a handle (like a motor vehicle handbrake) that moves tine fiaps rnanuaiiy7C~ large airc~al'i tilt: Gap iever is a r l eieciricai selector and the flaps are powerFd hflraulically (being ele<triclally selected).

I I ! I

If flapsjslats are lowered a t excks~ive airsbeed then they would be damaged by the airflow (even torn away - which has h'appened) so devices k-e fitted to the aircraft to prevent them being low'ered if the aircraft is travelling too fast. - ----

I

' 1 On mode'rn large aircraft the selection of high lift devices is Loverned by a computer which,takes account of airspeed, altitude, whether the aircraft is in the air or on the ground and wdether the pilot has made a selection or not. If hejshe has not and the computerl,thinks that a selection should be made (at take-off for example) the computer--will give a configuration warning.

If the pilot selects the flaps out a t too high an airspeed then the computer will not make the selection and a warning is given.

Air Resistance (Drag)

The total drag on the aircraft is made up of Profile drag and Induced drag. Profile drag increases with increased airspeed and Induced drag reduces with increased airspeed.

The drag calculation is not too unlike that for lift and written as:

Drag = CD% pV2A or C D q A

- 17 -

rnoodull l A-96

where

C D This is found by experimentation and is related mainly to the aerofoil shape. I t is just a number which is called a coefficient. With a flat plate the coefficient is 1.0 and with a highly streamlined shape this is reduced t o about 0.02.

p Air density. V Air velocity. A Is related to the wing frontal area. Sometimes designated 'S'.

Profile Drag

Profile drag is associated with the whole aircraft moving through ~e air and increases as the square of speed. Induced drag is associated with the production of lift and occurs mainly at the wingtips. I t is opposite to profile drag in that it decreases with the square of speed.

ProfiIe drag includes several types of drag and one, Boundary Layer drag, has featnresbuilt into t h e design-of-the-airer,aft to help minimise-its effects.

A ' -

I I , \ ' I ---- -- This 0undar-y layer is viscoud ($tic@) dith low energy lev& and will *adversely affect fl9ing control surfaces, en*ne perforkance - if it gets into intakks, and thc pkrfomance of aerofoils. I t is disliedby designers and iS bled hiy from engine intakes, and on some $e&foils-voitex Generators are lfitted io hove the free-svtrtrkarn flow down into the' bmiidary layer in an attembt to-liven it up and

1 give i t more energy. I I / \ \!

' i I i I I I i ? 1 : I --

TOP SURFACE OF WING

Fig. 15 VORTEX GENERATORS - PLAN VIEW OF WING

The vortex generators are small pieces of metal and inch or two high (25 to 51rnrn) set at a small angle to the airflow. They are set in rows usually towards the front of the aerofoil on the top.

Some Airbus A340s are fitted with a riblet film to aerofoils and fuselage, which are microgrooves to help reduce skin friction caused by the boundary layer.

The boundary layer may be lmm thick or several millimetres thick. The thicker it is the worse it is. The thickness is affected by several parameters:

* The further the air has to pass down the surface of a body the thicker the boundary layer becomes.

* The rougher the surface the thicker the boundary layer. * If the boundary layer gets too thick then it can become turbulent -

increasing it's depth still further. * The slower the airflow the thicker the boundary layer - in general.

FREE-STREAM

BOUNDARY LAYER

Fig. 16 BOUNDARY LAYER

The layer problems can be reduced by: I -- -- I

* I Having smooth highly polished wings, fuselage, tail and flying

control surfaces - reducing the boundary layer !tMickness. I * Bleeding away the boundary layer through many small surface air

inlets on the wings using vacuum pressure - experimental aircraft I - usually.

* Giving it more energy by mixing free-stream airflow down into the boundary layer air using vortex generators.

-k Bleeding boundary layer air away from engine intakes that are close to the fuselage, or designing the intake so that it is away from the surface of the fuselage.

The ideal arrangement is to have no boundary layer a t all, but this is not possible, so the thinner, and more laminar it is, the better.

A s the air passes over the wing so it starts to get turbulent a t a point towards the trailing edge. This Transition Point is usually close to the trailing edge but will move forward as the angle of attack is increased and if the boundary layer gets too thick.

During the stall it moves well forward to produce a very turbulent region of air 1 on the top of the wing.

- 1 9 -

rnoodull l A-98

Induced Drag

So far the drag that we have been dealing with increases with the square of speed. The faster the aircraft flies the greater it becomes - significantly, ie double the speed and the profile drag increases four-fold. With Induced Drag it is the opposite - the faster the aircraft flies the less it becomes. I t is caused by the lift generated by the wings/ helicopter rotor blades.

NEGATIVE PRESSURE

POSITW PRESSURE

Fig. 17 INDUCED DRAG - AIR MOVEMENT OVER-THE WIMG'CXPS

A s you know the press- on top of the wing is low whilst the pressure underneath is high. This is true of both a wing and a helicopter rotor blade. As the wing separates the two areas of positive and negative essure, they cannot equalise - except at the wing/rotor blade tips (and along thp trailing edge).

At the tip the airunder the winglrotor blade "spills over" to movc into the area of low pressure on the top.

This creates- wing tip vortices -which- use energy which ulti-mately comes from the aircraft engines - and costs fuel (like all drag), As these vortices spill over the wing tip thc local air pressure drops and so does the temperature, and under some atmospheric conditions this causes the moisture to condense out and vapour trails are produced.

This tip movement of the air means that there is some spanwise movement of the air on the top and bottom surfaces of the wing. There is a slight movement: towards the tip on the underside and a slight movement away from the tip on the top side,

Induced drag is a penalty we pay for the production of lift, but there are ways of keeping it to a minimum. These include:

Speed. With an increase in speed induced drag is reduced - but some aircraft can't fly fast, such as gliders, so they are stuck with this one. And all aircraft

I# - have to fly,slow to take-off and land. .

The effect of speed means that the air does not have time to move spanwise to spill over the wing tip - it is 'pushed' too quickly chordwise over the trailing edge.

Reducing the AoA. Induced drag is a t its greatest a t high angles of attack. Again we can't always change the angle of attack just to reduce the induced drag. The effect of a large angle of attack is to increase the pressure difference between the top and bottom of the wing and increase the induced drag.

AIRFLOW OVER THE WlNG AIRFLOW UNDER THE WlNG

INCLINATION ANGLE INCREASES TOWARDS THE TIP ANGLES SHOWN EXAGGERATED

Fig. 18 AIP,FLC?~ CVEP,jT,TXEEP, T9E ;;TIF?G I

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Winglets. A winglet helps prevent the air from spilling over the wing tip. Some aircraft have a winglet fitted beneath the wing tip, others on top, and some on the top and the bottom (the A380 for-example). I -

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I -

Some aikraft - the Boeing 777 for example - don't have them a t all. So the case for them is not as clear-cut as it appears. It is interesting to dote that even wingleks will produce both profile and induced drag - and increase weight. Some aircraft use wing tip fuel tanks-and other attachments a t the tips to help

-- reduce-induced drag. --

LEADING EDGE FIXED

Fig. 19 WINGLETS & FIXED SPOILERS OF THE LEARJET

Yjgure 19 shows a good example. Note the rathel latge winglets compared to the size of the aircraft. It also shows leading edge Fixed Spoilers or Stall Strips.

Remember, with swept wing aircraft there is a tendency for the wing tips to stall first which is not a good idea as this means loss of lateral control (ailerons), so some aircraft are fitted with inboard leading edge spoilers. These are triangular shaped devices fitted to the leading edge so as stall is approached they cause that section of the wing to stall first, allowing the pilot to still have lateral control.

Increasing Aspect Ratio. Aspect ratio is defined as the number of times the average chord length divides into the wing-span. When a wing is designed it is made to withstand a certain "load per unit area" (total mass of the aircraft divided by the gross wing area in plan view). This is called Wing Loading.

HIGH ASPECT RAT10 WlNG I AlRCWT 1 TOTAL WlNG SPAN = I70 ff WING CHORD = I D A

\ WINGAREA=170xlO =1700sqfl ASPECT RATIO = 17030 3 17

LOWASPPCT - AIRCRAFT 2 TOTAL WING SPAN = ff -- .-

RATIO WING "- '-- WING CHORD = 7s 8R- WlNG AREA = 19.7 w 86 : 4700 sq A ASPECT R4TIO = 86G19.8 = 4.3

Fig. 20 ASPECT RATIO - WING PLAN FORM OF TWO . - AXRCWW'I' WITH THE SAME WING AREA

Wing loading will effect stalling speed (high wing loading = high stall speed), maximum aircraft speed, gliding distance, aircraft performance etc, so other parameters are also used when calculating the wing loading of an aircraft and hence wing area.

We are only concerned with induced drag here, and for a given wing area the aspect ratio can be changed by changing the wing span.

Figure 20 shows an example of two aircraft with the same wing area and hence wing loading (if we assume both aircraft are the same mass), but the aircraft with the higher aspect ratio wing has smaller wing tips and hence will have less induced drag because there is less wing tip for the air to flow over.

Ip general, aircraft that fly slowly, such as gliders, will have high induced drag and there-fore will have high aspect ratio wings to help keep this drag as low as possible. High speed aircraft will have low aspect ratio wings because their induced drag is not high (at high speed) - their main probIem being profile drag.

Wing Fences

As the air under the wing tends to spill out over the tip it produces a span-wise movement from root to tip. Similarly a s the air comes over the top of the wing tip a span-wise component is produced on the top from tip to root.

On swept wings there is a tendency for the air to move spanwise towards the tips.

To try to counter these problems some aircraft have wing fences fitted. These are strips of metal u p to a foot high (30cm) running parallel to the airflow usually fitted on the top of the wing running from the leading edge to about 2 / 3rds chord length.

Fitted to the top surface of some swept winged aircraft, to promote correct airflow in front of ailerons to ensure correct airflow direction over the control surface.

The PrGfihk Drag C-IIF!~ I

With reference to graph 2 below. A s the angle of attack increases so the drag increases - not as a straight line but as a curve, as would be expected with the squared belocity in the equation. As speed rises drags increasks - as the square of speed. -

0 ANGLE OF ATTACK

GRAPH 2 - ANGLE OF ATTACK AGAINST DRAG COEFFICIENT

The Lift/Drag Ratio Graph

Ideally what is needed is a wing that will produce high lift with as little drag penalty a s possible. In other words a high liftldrag ratio. Graph 3 shows how this varies with angle of attack.

> 5 F 5r The graph shows the curve produced. Note the points o f interesting a t about 4" and 15" AoA. The 4" angle is known as the Optimum Angle of Attack and the 15" angle is the Stalling Angle.

rnoodull l A-1 02

The optimum angel of attack produces the best LID ratio and so it is the optimum cruise angle. At this angle the wing produces the most lift for the least drag. I t is the angle the wing is fsed to the fuselage, or there-abouts, called the Angle of Incidence.

ANGLE OF AT'FACK'

GRAPH 3 - L/D RATIO AGAINST ANGLE OF ATTACK - - . -.

As the angle of attack of the wihi is increased from, say 0° , so the lift/drag ratio gets better (higher) until zit about 4" where it is at its highest (24: 1 in graph 4 :above). (24: 1 means Wat there is 24 times more lift than drag on this particular aerofoil) . , --

8

After th& optimum AoA, lift still increases as the angle of attack is increased but drag starts to rises faster than before, so the LJD ratio gets smaller (worse).

At the stall there is a sudden drop in lift with the drag continuing to rise. This means that the L/D ratio reduces significantly at the stall.

Many large aircraft fly with the fuselage at a slight positive AoA when in straight and level f l ight as this will also produce some lift.

If induced drag and profile drag are plotted against velocity on the one graph then the total drag is found. Where the total is at a minimum - this is the velocity where the total drag is least and the speed that will give the aircraft it's greatest range for a specific volume of fuel used. I t is more a theoretical concept than a practical one as t he profile drag element is so much more than the induced drag element that in general, the faster an aircraft flies the greater the total drag will be.

THE FOUR FORCES

The important forces acting on an aerofoil in flight are lift and drag. However, when considering the aircraft as a whole there are other forces to be taken into consideration.

The other two important forces are weight (or more correctly mass) and thrust - from the engines.

So the four forces are:

* Lift * Drag * Weight * Thrust

Lift

This force is provided mainly by the wing of a fmed wing aircraft and some lift by the fuselage, and on some aircraft a small amount is provided by the tailplane. It acts a t right angles to the free-stream flow through the centre of pressure.

For fixed wing aircraft i t is varied by changing speed, AoA and wing cQr,f;,m ir9 tinn

--

6"' """'̂ ' -

I

For an aircraft in straight and level flight the lift acts vertically through the Centre of Pressure (C of P) of the wing. For a straight wing the C of P for each wing secrion is in the same position relative to the longitudinal datum line. For a swe& wing aircraft (or an aircraft-with a delta wing) as w?ng sections are considered further away from the fiiselage so the C of P positionpis further to the rear relative to the longitudinal datum line. This means that, for a swept winged or delta winged aircraft the average C of P fore and kft position has to be calculated.

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This also-demonstrates why, if-the outer part of a swept-wing stalls, the aircraft will tend to pitch nose up.

When calculating the total lift produced, consideration has also to be given to any lift produced by the tailplane (positive or negative lift) and the fuselage.

Remember the lift always acts a t right angles to the free stream airflow. For an aircraft in a vertical climb (a fighter for example) any lift created would be horizontal. The aircraft would be supported by the thrust from the enginels alone. Any lift created from the wings would tend to move the aircraft horizontally and if the pilot wanted to fly the aircraft exactly vertically then he/she would have to push the stick forward a little to reduce the AoA to a negative angle so the net result of the lift from the wings was zero. (Remember, for an aircraft in a vertical climb the relative airflow is vertically downwards.)

- 25 -

rnoodull l A-1 04

Drag

This acts backwards at right angles to the lift and resists the forward motion of the aircraft. I t opposes thrust and acts through the aircraft's centre of drag.

Changed by varying wing configuration, aircraft velocity and, to some extent, AoA - for fixed wing aircraft.

Thrust

For most aircraft this is provided by accelerating a mass of air backwards either by a propeller or a jet engine. This produces thrust to propel the aircraft forward. I t acts through the centre line of the je t engine or the propeller spinner.

When more than one engine is fitted to an aircraft the sum of all the engine thrusts can be calculated to produce one thmst line to act through the centre of thmst.

. ---. .

Varied by changing the engine ~hrottlesettings/VP prop settings. 1 I I

weight (Mass) I S

. - -

This always acts vertically down~&&, unlike the other forkes that act relative to the; aircraft" attitude. All the mass of the aircraft is saidito act through the Cenke of Gravity (C of G) of the aircraft. I 1

In general terms is fxed for any one instant in flight. It cannot be varied by the pilot; but over-the long term the .mass- reduces because of fuel-usage . (For most large aircraft the pilot can dump fuel in an emergency, but this is not relevant here.)

Arrangement of the Four Forces

These are so arranged an the aircraft as to make it reasonably stable. In straight and level flight at constant speed with no turning moments the aircraft is said to be in equilibrium. This means:

THRUST = DRAG and

WEIGHT = LIFT

Each pair is equal and opposite. AIthough they are opposite in direction they are not usually opposite in position.

For straight and level flight the AoA is adjusted by the pilot to make the lift equal to the weight, if it is greater the aircraft will climb (in general terms). If it is less the aircraft will descend (in general terms). Tlle engine thrust is adjusted by the throttles to make it equal to the drag, if it is greater the aircraft will increase speed - if it is less the aircraft's speed will decrease.

The Ideal Arrangement

The drawing below shows the ideal arrangement of the forces. Not all aircraft are like this. For various reasons some aircraft have to have their forces in a less than a n ideal arrangement - seaplanes for example have a high thrust line well above the drag line - to keep the engines (and propellers) out of the water.

The ideal arrangement is where the Centre of Gravity is forward of the Centre of Pressure (Centre of Lift), which produces a nose down couple - and the thrust line is lower than the Centre of Drag, which produces a nose up couple. Each couple opposing the other and cancelling each other out.

NOSE

-

NOSE DOWN COUPLE

L---

C DRAG

/ I \

COUPLE c i f G I c-of L

-- - I WEIGHT - -

Fig. 2 1 THE IDEAL ARRANGEMENT OF THE FOUR FORCES

The arrangement is 'ideal' for the following reasons:

* The lift-weight nose down couple opposes the thrust-drag nose up couple. Ideally the couples should exactly balance each other, with many aircraft they do not.

* Should engine power be reduced or engine failure occur then the aircraft will automatically take up a natural nose-down gliding position (there being no opposing nose-up couple).

* The forward C of G makes the aircraft more stable longitudinally.

NOTES

1. The function of the tailplane is fwo fold: I t is to counteract any tendency for the aircraft to pitch in flight (longitudinal stabiliw); and it may be fixed at an angle to produce an additional downward (nose-up) correcting couple in flight. Many airliners have tail planes with asymmetric aerofoils producing down-lift and a tail dawn moment.

2. A couple is defined as two opposing forces not on the same centre line. The turning effect of a couple is called a moment.

3. With highly swept wings, if the outer section of the wing (which is behind the C of G) was to stall before the rest of the wing then the C of L for the whole aircraft would move forward causing a nose-up moment.

MANOEUVRES

An aircraft can take up any position in the air m d the four forces will a1 act in - . - - - - -

relationship to each other, but reinembe'r: - .. .

WEIGHT always acts vertically downwards through the C of G, LIFr always acts at right angles to the relative airflow (frcc-

stream flow) through the Centre of Lift (or Centre of Pressure).

THRUST always acts in line with the aircraft engine,%- DRAG always acts itz,line with I the airflow, at right anglcs to lift and

opposes thrust.

All the above is correct irrespective of the position of the aircraft relative t o the ground; - .

The Banked Turn

For the aircraft to carry out a turn it must produce a force towards the centre of the turn (centripetal force). This is true for any object to go round a corner.

When the aircraft banks t he lift force is placed at an angle (when viewed from the front or rear) by the pilot moving the control column to one side moving the ailerons.

The down going aileron will produce an upward force (higher wing) and the up- doing aileron will produce a downward force (Lower wing).

"NEW' LlFT VECTOR "OLD" LlFT VECTOR = W Kt--- ANGLE OF BANK = 8 l \ d ,

CENTRIPETAL FORCE = W TAN 0 (ie dependant on weight & angle of bank)

(theta)

I WEIGHT = W

Fig. 22 A BANKED TURN

This causes one wing to go down and the other to go up. The force can then be divided into its component parts as-follows:

Lift Component: Must be equal and 'opposite to weight if the aircraft is 1 1 to not lose or gain height (the pilot's VSI to remain a t

zero). 1 --

I -

I

Centripetal To be equal and opposite to the ceritnfugal force Component: created by the aircraft, and to provide the force

required to 'pull' the aircraft arouna the corner. -

- - If the centripetal ---- -- force is larger than the centrifugal force then the aircraft will go into-a tighter turn, if it is smaller the radius of turn will get greater.

Can you see from what we have discussed so far, that when the aircraft goes into a banked turn the pilot must increase the angle of attack to increase the lift so as to give a longer lift vector? In figure 22 the 'new' lift vector is longer in the bank than it was in straight and level flight - compare the 'new' vector with the 'old'.

For a correct turn therefore, the pilot must apply rudder (to help the aircraft turn), aileron (to bank the aircraft and move the lift vector to produce a centripetal component), and elevator (to increase the angle of attack to increase the lift vector), and apply more throttle.

The throttle has to be increased slightly to counter the extra drag created by the increased angle of attack.

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With a decrease in the radius of a turn (tighter turn) the angle of bank must be greater. This means that the 'g' loadings will be greater (for both people and the structure).

During banking the down going wing would be the most likely to stall because of the increased AoA, but only if the rate of banking was severe. Whilst in the banked turn, because of the increased angle of attack of both wings, the stall is more likely than when the aircraft is in level flight under the same conditions.

With autopilot engaged and commanded t o turn by the navigation computer, it will cause the ailerons to move commencing the bank, the rudder to move providing a correct turn without side-slip or skid and the elevator to move providing lift compensation to maintain correct height,

Side-slip is the aircraft slipping downwards into the turn because of too much bank or not enough rudder. Skidding is the aircraft sliding out of the turn because of not enough bank or too much rudder.

.. . .. . . - . Lead Factor

Load Factor is the ratio of lift to weight (more correctly mas's). With the aircraft flying straight and level the LoadFactor is one. With the aircraft in a banked turn the Load Factor is increased, for example:

ANGLE OF BANK g LOADING LOAD FA~TOR ... I

For an average 200 ton (= 200 metric tonnes) airliner these figures arc, respectively 440 tons/ tonnes and 600 tons/ tonnes.

Because, effectively, the aircraft mass increases during the turn so does the stalling speed.

The Climb

When the aircraft climbs dl the force vectors move with the aircraft - except the weight vector. With reference to figure 23, it can be seen that the weight vector W can be split into its component parts W, equal and opposite to L (lift) and W2 - a force to be added to the effect of drag D. This means that the engine thrust must be increased (as you would expect) from To to T, to give a total thrust vector equal to D + W2. This will keep the aircraft speed constant whilst in the climb.

It is interesting to note that the new thrust force T, can be split into its component parts also with T2 representing the element of engine power that is assisting the lift so that L1 + T2 = W.

Fig. 23 AIRCRAFT IN A CLIMB - -

-

This means that a component of the thrust vector is helping the lift vector and the actuAl generated lift can be reduced. So L1 is equal to Ti + the vertical vectoral component of L.

1

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1 - -

--

A similar analysis can be used when the aircraft is in a dive. -

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So it is interesting that less lift is needed when the aircraft is climbing than when it i:s in straight and level flight. Consider for a moment how much lift is required by a fighter in a vertical climb - none.

--

So the greater the angle of climb the less lift needed - but - the greater the angle of climb the more the engine power is required.

A Glide

With the engines shut down (or failed) the aircraft will go into a glide, naturally or by the pilot pushing the control column forward and lowering the elevators - putting the tail up and the nose down.

In generally the object of a glide is to get as much distance as possible - in order to make an airfield, or make land. The gliding distance can always be shortened, if the pilot thinks he/she is gliding too far, by pushing the control column forward, but increasing the gliding distance is another matter.

For the aircraft to glide it must move forward," so the nose is moved down to create a component of the weight vector to produce a force in the required direction.

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moodull l A-1 10

This component takes the place of the thrust vector. The 3 forces now availabIe are weight, drag and lift and they must all "ealance' to give a glide in an equilibrium condition. The resultant of the lift (L) and drag (D) vectors (figure 24) must equal the weight (W) vector.

L D RESULTANT

GLIDE PATH #

HORIZONTAL

Fig. 24 AIRCRAFT IN A GLIDE I

By a process of simple geometry if can be seen that the angle'of glide (a) is the same as the angle between themliftvector (Lj and the drag vector (D). I t is called the Glide Angle. By inspection it can be seen that if the LiftlDrag (LJD) ratio is higher ( h e r drag or higher lifti) then a is smaller, the glide slope is shallower and the 'gliding distance longer. So, the maximum gliding distance is governed by the LID ratio. The higher the ratio the longer the gliding distance.

- . .- - .

To some extent the LID ratio can be considered as a measure of the aerodynamic efficiency of the aerofoil. The higher the L/D ratio the higher the efficiency.

As seen from the graph on LJD ratios earlier, the best LJD ratio is about 4" angle of attack. If this angle is varied either up or down the L/D ratio gets worse. This means that there is only one glide angle that gives the best range.

A higher angle of attack will reduce the LJ D ratio and reduce the range, and a lower angle of attack will, of course, steepen the descent.

If the pilot is gliding for distance, to get to an airfield to make a forced landing for example, then only one angle will do. An angle of attack indicator is a distinct advantage. If one is not fitted to the aircraft then the pilot will have to rely on instruments to check on rate of descent (VSI); airspeed; ground speed etc.

Should the pilot wish to pull out of the glide (or pull out of a dive) then, provided there is sufficient airspeed, all he/she needs to do is to pull back on the control column to raise the elevators to cause the nose to come up. Provided airspeed is maintained above the stalling speed then the pull-up is only limited to the strength limits of the airframe.

If the pull-out is too severe then damage may be sustained with possible structural failure. In general the lighter the wing loading the shallower the glide angle.

Glide Ratio

This is the ratio of the horizontal distance travelled to the height lost. The higher the ratio the better.

A Dive

In general to dive a n aircraft the control column is pushed foomard, this lowers the elevators, raises the tail and lowers the nose. To be more effective the thrott lp are pushed forward to'increase engine power and the airspeed of the aircraft increases.

I

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Once thd correct angle of descent has been established the elhators can be moved d o r e towards the neutral faiTG3 position and the pilot watches the instruments. These will show the airspeed indicator (ASI) indicating a n increasing speed, the Mach number rising and the altimeter showing a reducing altitude. Two things the pilot must worry about - the aircraft must not go too-fast, or get too low. If speed gets too great then, on modern aircraft, Never Exceed-Speed warnings sound (and on some systems thelnose is automatically pulled up). If the aircraft gets too close to the ground the ground proximity warnings sound.

Control Surfaces

Figure 25 shows the control surfaces a s fitted to a civil airliner. The slats, leading edge flaps and flaps (trailing edge) are for increasing the liftldrag ratio of the wing, although trailing edge flaps are often lowered further on torch- down to produce drag.

The ailerons are used for roll control (to bank the aircraft) - also to assist in improving the L/D ratio of the wing during take-off by being drooped (automatically in take-off configuration). The spoilers are also used to assist in roll control when they are deployed asymmetrically (also automatically).

, i

The pitch of the aircraft is controlled by the elevators and the rudder is used during the turn. Tabs are fitted to the elevator and aileron systems.

OUTBOARD CIA?

-

Fig. 25 CONTROL SURFACES - BOEING i737

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The a c t i d take-off (and landing) brocedurewill valy from &refaft to aircraft. The pilot will consult the Pilot's Notes (for a small aircraft these will be a small booklkt,, for a large aircraft the Pilots Notes will be an A4 ring :binder with several hundred pages - also on CD) - for . that particular airbrait, but in general the following sequence will apply: , . . . .

1 . Permission obtained from Air Traffic ControI (ATC). 2. Engine/ s running. 3. Engine/ s and systems checks completed. 4. The aircraft is configured for take-off - flaps down (take-off

position) - leading edge devices deployed - spoilers in - doors closed and armed - altimeter set - navigation systems set - auto brake set etc.

5. At the end of the runway t he engine/s are set to take-off power (with correct propeller setting if fitted with a VP propeller) and EPR, JPT, RPM etc checked.

6. Brakes off, Take-off into wind - it helps to keep ground-speed down while keeping air-speed as high as possible.

7. The aircraft accelerates, airspeed increases and lift is generated (as the square of speed). The thrust vector is greater than the drag vector so the aircraft accelerates - but drag is also increasing.

8. When the speed is high enough the pilot pulls the control column back to raise the elevators - this rotates the aircraft about the main landing gear (rotation speed - VR) and increases the angle of attack - to say about 30" (with slats and flaps out).

9. This creates enough lift to overcome the weight vector and thc aircraft rises into the air.

10. At the correct speed the flaps and slats are selected u p and the landing gear also - if retractable.

11. The aircraft will continue to accelerate until the pilot closes the throttles or the speed builds to a value where the drag equals engine thrust. At this point the aircraft will maintain a constant speed.

Landing

Permission obtained from ATC. Aircraft configured for landing - landing gear down, flaps/slats down (approach/landing position), radio altimeter checked, barometric altimeter set to OFE. autobrake set, autoland set etc. ILS set and checked if using a n instrument lanping approach. Engine/propeller power setting reduced to allow approach. May be automatic if auto-throttle used. Aircraft slowing (drpg vector greater than thrust vector) and lift vector slightly smaller than weight vector - airc~aft descending. A s aircraft slows, drag vector gets smaller, but still greater than thrust. I I

Aircraft lands into wind to reduce ground-speed. Jus t before touch-down the aircraft 'flares-out' and lands a t a landing approach angle of about 3 to 5 degrees. On some aircraft, spoilers, reverse thmst, and auto-brakes operate automatically when the weight switch is made. Engine power reduced to idle. Brakes applied. Aircraft slows. Flaps, spoilers etc selected in. Ailerons to normal position. Taxi speed maintained and pilot manoeuvres to park the aircraft as requested by ATC.

THE AXES

To enable the various attitudes of the aircraft to be discussed in a technical manner axes are defined - these are imaginary straight lines all perpendicular to each other and all running through the Centre of Gravity.

Lateral Axis runs from wing tip to wing tip - or parallel to a line from wing tip to virJng tip. When the aircraft moves about is axis it is said to be pitching (climbing or diving) using the elevators (control column back, elevators up, tail down, nose up, aircraft climbs).

rnoodull lA-114

Longitudinal Axis runs as a centre line from nose to tail. When the aircraft moves about this axis it is said to be rolling - left or right wing down. To roll, the control column /control wheel is moved from left to right to move the ailerons. One aileron moves up the other moves down. To roll to the left the contro1 column /control wheel is moved to the left the left aileron moves up (pushing the left wing down) and the right aileron moves down (pushing the right wing up).

Normal Axis is at right angles to the other two, and in straight and level flight is vertical. Movement about this axis is called yawing - nose moving to port (left) or starboard (right). Pushing the left foot forward on the rudder bar causes the rudder to move left pushing the tail to the right and the nose to the left.

Notes

1. These axes are relative to the aircraft - when it moves they move. 2. For most manoeuvres all the axes wilI be involved. 3. When the controls are moved in the flight deck their movement is

said to be instinctive. 4. For rigging purposes the aircraft will have DdtumLihes. There is a

lateral datum line - not coo different from the lateyd axls, and a longitudinal daturri line - not too different from the longitudinal axis. These ~a tu rn l~ incs are specified in the AMM and may not necessarily pass tdrough the Centre of GraviIy.1 Instruments can be placed on the airfrAme (when on jacks in the hbgar) at specified points to ascertain! how level it is prior to setting other airframe angles.

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Fig. 26 THE THREE AXES

STABILITY

Stability can be defined as the ability of an object to return to it's original position after it has been disturbed. In general it is opposite to manoeuvrability. If an aircraft is ve y stable then it is not very manoeuvrable and vice-versa. Most fighters are not very stable whilst most civil airliners are.

In general terms, an object may:

* Be stable. It will move back to it's original position after disturbance. The upright pyramid in figure 27 for example. If it is tilted a little to one side and let go it will fall back onto its base.

A Be unstable. Once moved from it's present position an object will continue to move in the direction of the original disturbance. In other words the disturbance will get worse. The inverted pyramid for example, if it is balanced on its point and a force moves it to one side the movement will continue once the force is removed.

* Have neutral stability. Once disturbed an object will take u p it's new position but it will not move further or try to-return to it's

I original position. The-ball-for example - if disturbed it will roll a little and stop. It will not return nor will its position get worse.

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When i n flight the aircraft will be subject to local air disturbances which will try to deflect it from it's flight path. If the aircraft returns to it's original flight path without the aid of the pilot - t h e n it is said to be stabld, if i t does not then it is said to be unstable or have neutral stability. Stability can be achieved in 2 ways - actively or passively. I

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- -

- - DISTURBING DISTURBING FORCE FORCE DISTURBING --

STABLE UNSTABLE OBJECT WITH OBJECT OBJECT NEUTRAL STAt31LlTY

UPRIGHT INVERTED BALL PYRAMID PYRAMID

Fig. 27 STABILITY

STABILITY

ACTIVE PASSIVE

Fig. 28 TYPES OF STABILITY

Active Stability

The aircraft is flown back to its trimmed flight path automatically by the controls. The flying controls are powered (usually by hydraulics) and controlled by computers that note the aircraft" movement from laser wros. The computers compare the aircraft movement with the pilot's input, and intervene if an un-commanded movement occurs. Used mostly o n military aircraft, but also on some civil aircraft - for example the gust alleviation spoilers of the A320.

When a gust disturbs the aircraft about the longitudinal axis causing a roll the gyros pick the movement up and send the appropriate signal to the flying control computers. These note that the pilot did not command the movement so send a signal to operate the selector valve of the (gust alleviation) spoilers.

These are hydraulically powered and deploy asymmetrically on the up-going wing. This action will dump lift on the high wing and cause it to drop - putting the aircraft back to it's trimmed position.

A fully actively stable aircraft allows-the designer l e build-Ithe airframe strictly in accordance with engineering principles without any considkration far aerodinamic parameters. (Note the shape df some modern, adtively stable, military aircraft).

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passive Stability I -- -. i

I I .. -

i The aircraft flies itself back to its original 'path after being disturbed because of the aerodynamic design of the hirframe. This is the stability that we shalI considei here. I t is achieved by the design of the tailplane (stabilator), fin, and the wings-: - - - . . .. .- ..

Almost d l aircraft are designed to be passively stable so the pilot need takc little or no action to return the aircraft to it's original path after it has been disturbed - although he/ she may assist it by using the controls if he/ she wishes.

Although the stability of an aircraft involves all three axis - as they all interact, it is usual to consider stability in three separate forms:

* Lateral Stability - about the longitudinal axis. * Directional Stability - about the normal axis, * Longitudinal stability - about the lateral, axis.

Lateral Stability

For a high winged aircraft lateral stability is helped by the Pendulum Effect. This is produced by the high position of the C of L and t h e low position of the C of G.

rnoodul I I A-I 17

If the aircraft is disturbed about the longitudinal axis the movement of the Centre of Lift to one side of the Centre of Gravity will cause a correcting movement to help put the aircraft laterally level.

) WEIGHT

Fig. 29 PENDULUM EFFECT

Also, the down-going wing will experience an increase in AoA as the relative airflow is moving upwards from a position forward of, and below, the wing. This will produce an increase in lift on that wing.

- -

The up-going wing experiences a decrease in AoA (as the reladive airflow is moving down from a position forward of, and above, the wing), so experiences a decrease: in lift. All this helps to correct the un-commanded role.

This differential lift effect applies to all conventional fxed wing aircraft whether they have a high wing or a low wing, whether the wing is swept or straight.

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Dihedral Angle - - - -

For low winged aircraft (most civilian airliners) lateral stabilityis assisted by the Dihedral Angle of the mainplanes - the upward and outward inclination of the mainplanes away from the fuselage (measured against the lateral datum line).

If a gust of wind raises one wing the down-going wing effectively has an increase in angle of attack thereby increasing the lift of that wing, and the up- going wing will have a reduced angle of attack and a reduced lift force, the total effect being to help to restore the aircraft to its original flight path. Similar to the case discussed before but the dihedral angle enhances this effect.

If the differential lift forces do not correct the aircraft (which sometimes happens) the aircraft will stay in the banked attitude and a side-slip will occur. (Unless the pilot intervenes by applying opposite aileron - or automatic roll control spoilers are asymmetrically deployed - on the up-going wing.)

Stability from the dihedral angle applies to straight, 'swept and delta winged aircraft.

INCREASED LIFT

t REDUCED UFT

DIHEDRAL ANGLE

Fig. 30 DIHEDRAL ANGLE

Swept Wings

Swept wings include delta wings. They may be designed for high speed flight to help reduce M C ~ T but come into play during an un-commanded roll at any speed. If the roll remains uncorrected and the aircraft starts to sidc-slip then the low wing will meet the airflow at a more effective angle (in-plan view) than k\e ;$+ng t,k,'d= creaeng e-.reE-EErs !if? c~ ?filxl'r Wibg E E ~ !iff_ the high &ing - helping to further Correct the aircraft.

AIRFLOW 01RECTION ' \ a DECREASED EFFECTIVE WING PLAN

Fig. 31 SIDE-SLIP

Also in a side-slip, whether the wings are swept or not the high wing (on a low winged aircraft) will be in the Aerodynamic Shadow of the fuselage and expe~ence a reduction of lift because of the turbulent flow over that part of the wing.

- 40 -

rnoodull I A-1 19

f SIDE-SLIP

Fig. 32 AERODYNAMIC SHADOW (LOW WINGED AIRCRAFT)

So it is hoped that with dihedral angle and sweep-back and aerodynamic shadEw [a%-d in snrn-e cases penduhim stahilip! that lateral~stahility will he achieved. However, it is the most difficult stability to obtain out of the three and for the majority of aircraft pilot intervention/autornatic stability control systems still have to intervene if the aircraft rolls more than a few degrees.

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Directional Stability I I 1

This is assisted by the fin and rudder and the side area of the fuselage aft of the C e n t ~ e of Gravity - taken all together called the Effective Keel Surface. If the aircraft is caused to yaw then, like a weather-cock or weather vane on a church spire, the airflow will "blow9t back to it's original position.

Remember that when it yaws the aircraft will tend to fly in it's original direction for a short time due to it's momentum (Newton's first law) - thus for a short time the airflow will be acting on the side of the fuselage. This correcting moment is also assisted by the small sideways "lift'' produced by the fin.

EFFECTIVE KEEL SURFACE

Fig. 33 EFFECTIVE KEEL SURFACE

- 41 -

moodull l A-120

This correcting action may set up an oscillating motion which is corrected by fitting powered automatic yaw dampers to the rudder system.

Note. For most aircraft the fin is vertical and its chord line is parallel to the aircraft's longitudinal datum line. For some single engined propeller driven aircraft the chord line may be set at a small angle to the longitudinal datum line to tv to counter the effect of the swirling propeller slipstream.

Longitudinal Stability

This is normally associated with the tailplane or horizontal stabiliser. For many large aircraft the tailplane chordline is set at a small negative angle to the longitudinal datum line (or the tailplane has a 'reverse camber'). The angle between the two chordlines of the mainplane and the taiI-plane is called the Longitudinal Dihedral Angle.

The negative tailplane angle of incidence (producing negative lift) helps the stability of the aircrdt and also creates a downwards force on the tailplane to help balance the four forces. :, -

- - -

If a gusi of wind causes the noseof the aircraft to be deflect=& up or down then the tailplane will experience a qhange in ADA but the aircraft's momentum will keep the aircraft going in the oi-iginaI direction for a short time.

I - - I

This change in AoA will create a fdiEe on the tailplane to coirect the nose-up or nose-down condition eg:

1. Aircraft nose deflected down.' I 2. . Tailplane experience$ a negative AoA. 3, - .Negative angle of-attack produces a downward-lift force. 4. The downward force on the tail creates a correcting turning

moment about the C of G to raise the nose and restore the aircraft to it's original attitude.

fJ CORRECTING MOMENT

AIRCRAFT'S MOVEMENT - ! DOWN FORCE ON TAILPUNE

Fig. 34 LONGITUDINAL STABILITY

The mainplane will also experience a negative angle of attack at the same time, but the tailplane creates a large turning moment which is strong enough to correct the aircraft.

NOTE

We have considered stability of an aircraft as three separate items. In reality they are connected, and one will have an effect on the other. For example, consider an aircraft that is directionally very stable, but not very stable laterally - the following could happen:

The aircraft is caused to move right wing low (un-commanded role) Not being very stable laterally the wing stays down. The aircraft starts to side-slip - to the right. Still the wing stays down and the airflow is now being felt on the side of the fuselage and fin. The Effective Keel Surface comes into play and starts to turn the aircraft towards the direction of the low wing (to the right). This causes the high (left) wing to go faster - creating more lift. The high wing becomeseven higher; the side-slip gets worse and the process repeats itself. We now have a very unstable aircraft which is starting to go into a spin. If not corrected by the pilot it will enter a spin which may be such that recovery may -- not - be possible. I - -

An aircraft may be statically stable in that it will return to its trimmed position after being disturbed without any help from the pilot - as discussed above. However, it may be Dynamically unstable in that when returning to it's original position it overshoots that position. Having shot past this position its normal stability comes into play to return it back to its original position - only for it to overshoot again.

These oscillates about it's original flight path may continue for some time and even get worse.

An aircraft is said to be dynamically stable when it returns to it's original trimmed position without any overshoot. It's movement is said to be heavily damped. There are no oscillations and the aircraft returns steadily to it's original flight path.

If the aircraft returns to it's original trimmed position but overshoots it is said to be lightly damped. The overshoot causes it to move back passed it's original

, position and then return back again to try to-stabilise about the position it had before it was disturbed. These oscillations gradually decrease in amplitude until the aircraft regains it's original flight attitude.

rnoodull l A-122

AIRCRAFT FUGM PATH

HEAW DAMPING - DYNAMICALLY STABLE ORIGINAL FLIGHT P A M

1

SOME DAMPING -DYNAMICALLY STABLE. OSCILLATIONS REDUCE

NEGATIVELY ~ ~ P E D - DYNAMICALLY UNSTABLE - OSCILLATIONS GROW - .. .

Fig. 3 5 DYNAMIC' STABILITY

1

With Neutral Dynamic ~tabilid it%-& to be under-damppd; In this case the aircraft returns to it's original flight path and overshoots. 1g corrects this overshobt to return back to the: same overshoot position but on the other side. Thcse overshoots either side of'the original flight path do not decrease or increasc in amplitude but r e m h the same.

. .

If it i<D$n-Gnicdly Unstable (negatively damped) the situation--is similar to the above but the amplitude of the oscillations get worsc. The amplitudes get greater until there is a rncchanical failure or the pilot or some automatic stability system takes a hand in correcting the situation.

Static Stability / Dynamic Stability

All the above 'dynamic stability's' are related to an aircraft that is statically stable because in each case the aircraft is trying to return to its original trimmed attitude. Dynamic stability only comes in to play whcn the aircraft is moving back to it's trimmed position. So in general terms static stabilitg acts -

first and dynamic stability acts second.

The figure above shows the aircraft longitudinal stability, but the principle is the same for directional and lateral stability. Although with lateral stability some of the corrective farces are non-oscillatory. .

As we have already seen the tailplane or stabiliser provide the main input for longitudinal stability and the fin and effective keel surface for directional stability.

For lateral stability it is the effect of increasing lift on the down-going wing and reducing lift on the up-going wing that provides the initial correcting forces. This effect can be helped by having a high wing (pendulum effect) or, on a low wing, by using dihedral. These forces tend to be corrective and not oscillatory.

When a side-slip occurs, of course, sweep back will increase the efficiency of the low wing and aerodynamic shadow will also affect the high wing. Again this is primarily 'damping' but may set up an oscillating motion in some circumstances.

Dutch Roll

This involves movement about the longitudinal axis (roll) and movement about the normal axis (yaw).

-- -

If the aircraft is disturbed about the normal axis (yaw) and the fin is moving to one side (say right) of it's normal position the wing on that side of the aircraft is going faster than the wing on the other side. So the initial yaw to one side causes the wing on that side to lift and cause a rolling moment to the left. So as thefin moves to the right the right wing goes faster, lifts and the - aircraft rolls to the left.

I I I I ,

~ u r i n k this time the airflow is acting on the fin and effectivk keel surface to move the fin to the left, as it does so it moves the left wing faster so increasing it's lift. This increase in lift of the left-wing occurs at the same time as the dihedral effect is trying to correct the low left wing. It , therefore,lifts while the fin is moving to the left and this combination sets up an oscillation motion, a combination of roll and yaw, called Dutch Roll.

Dutch Roll usually has a low frequency (say l/z to 2Hz). Dihedral and sweepback tend to make Dutch Roll worse while anhedral (negative dihedral) tends to improve the situation.

Because aircraft use dihedral and sweepback to help with static stability other means, such as Yaw Dampers (powered units fitted to the rudder system), are used to counteract Dutch Roll.

Yaw Dampers

The pilot can control Dutch Roll by the use of the rudder but this would be t;iring and difficult so yaw dampers are fitted to post large commercial aircraft.

I

- 45 -

rnoodull l A-124

The damper is fitted in series in t he rudder flying control system. I t is automatic and controlled by computers with gyro inputs - one sensing degree of yaw and one sensing rate of yaw.

The damper is usually hydrau3icalIy powered and it's piston is fed fluid pressure at the correct times to input into the rudder contra1 system to move the rudder to correct the yawing movement so correcting Dutch Roll.

FLUTTER

Flutter is a vibration that is set up in the airframe. It is caused by aerodynamic loads acting directly on the airframe, or on the flying control surfaces which affect the airframe. The condition can also be caused by engine vibration (the Lockheed Electra suffered seriously from this form of flutter).

It is a vibration or cyclic movement of the airframe, part of the airframe, or flying control surfaces due to elasticity within the system/ structure which is started by aerodynamic loadings. This cyclic movement of the structure is

- -. - - - . called A& Elasticity when induced by aerodynamic force-s.

-.

The elastic behaviour of the structure may be complicated with both flexing and torsional movements being involved. I

The flutter may be slight with no ~b-Vious signs of deterioration to the airframe or equipment, and may be prededt fir-many years. I t may de sever-enough to cause major svtructural failure isn 'flight (similar to an explosksn).

John IXrry in his DH 1 10 at Famborough for example, which disintegrated in flight kilIing it's occupants and niany spectators on the ground.

. -.. -- . .

There are several types of flutter and all must be kept to a minimum. In it's mildest form it causes fatigue which will cause failure in the later life of the airframe. In it's worst form it can cause immediate structural failure.

Better understanding of the problem, and better design of the airframe, flying control surfaces and engines will all go towards reducing flutter to zero or within limits which are acceptable.

As far as the maintenance engineer is concerned, strict observance of the procedures laid down in the manuals is the answer - particularly when dealing with flying controls, airframe structures, and engine and propeller balancing.

Flutter (mass balance) is dealt with in more detail in the Flying Control Systems book.

VARIATIONS IN AIRCRAFT DESIGN

Some aircraft are fitted with a fore-plane in place of a tailplane. The fore-plane is fitted to the fuselage forward of the mainplanes and generally provides less longitudinal stability than a tailplane. In some cases it can actually make the aircraft unstable longitudinally (aircraft pitches u p - fore-plane has an increased AoA with increased lift which increases the pitching moment). A n aircraft with a fore-plane is called a Canard aircraft.

FOREPLANE

Fig. 36 CANARD AIRCRAFT I

I

I One advantage of canard configured aircraft is that they can be made stall- proof. If the fore-plane is set a t a positive angle on the fuselage slightly greater

I than t h e angle the mainplane is set a t (angle of incidence), then just before the mainplane reaches the stalling 'angle the fore-plane will stall and the nose of the air'craft will not go any higher and the mainplanes will not stall.

The elevators on the fore-plane are connected to the control column in the usual way except that they move opposite to those fitted to the tailplane, eg for the pilot to climb he/she pulls the control column back (instinctive control as before) this lowers the foreplane elevators causing an upward force on the foreplane, causing the nose to rise.

On many high performance canard fighter aircraft the fore-plane has no elevators but moves as a complete surface (all-flying fore-plane) to change the pitch of the aircraft. (Some conventional tailplanes are also all-flying tailplanes with no elevators.)

Swept Back Wings

These have the following advantages:

* Helps lateral stability once a side slip has started (as already 2 discussed). F-

- 47 -

rnoodull l A-126

* Increases the critical Mach number (MCRIT) and allows higher speeds to be obtained in the transonic speed range without the onset of compressibility effects. Mach (Ernst Mach Austrian physicist 1838 - 19 16) is a measure of the aircraft's speed in relation to the speed of sound and the higher the MCMT the better. Mach is not part of the module 8 syllabus.

Disadvantages include:

* Structurally more difficult to design than straight wings. * Tend to suffer from tip stall - if this happens the lose of lift at the

tip will normally cause a nose-up moment and possible loss of lateral conlol. The tendency to tip stall may be reduced by wash- out (reduction of the angle of incidence of the wing towards the tip) .

* Less efficient at creating lift than a straight wing.

Swept Forward Wings - . . . . - . . . -. . -

Aerodynamically these behave in a similar way to swept back wings, but structurally there is a problem. Any structure that is placed out in the airflow is stable if it is allowed to trail in the airflow (swept back wings, control surfaces etc) . I I

I - -

I

If thc structure is fitted so that it1 lies forward of its mounting then it is unstable, Example - a rudder could be designed to be fitted in front of the fin with its hinges on the rear of the rudder so it is facing forward of the fin.

I t would work the same as a conventional rudder, but, as soon as it is displaced more than a degree or two out of alignment with. the airflow the force of the airflow would cause it to swing completely round tearing it off its hinges and causing possible loss of control of the aircraft.

Forward swept wings have this problem, though less dramatic. Because of their sweep any movement caused by aerodynamic loading will cause the wing to twist. This twisting will cause the wing tip Angle of Incidence and hence Angle of Attack to increase (wash-in). Increasing the tendency for wing-tip stall - which is always a bad thing.

With swept-back wings the opposite is t rue. Any movement caused by aerodynamic loading is a stable condition and any bending normally causes wash-out. Remember wash-out is better as it is preferred that the roots of the wing stall first so lateral control is maintained during the stall.

Tlie Delta Wing

Has the same advantages as the swept wing but generally has a greater wing area so reducing the wing loading (aircraft's mass per unit area of wing). It is generally stronger.

Also has the same disadvantages, except for the structures problem. But as most delta winged aircraft have a high degree of sweepback so it has further disadvantages:

* Poor lift characteristics a t low speed. * High angles of attack required for take-off and landing. Note the

high nose landing gear on Concorde to allow for these high angles.

I?&. 37 TYPICAL TAILLESS DELTA WING AIRCRAFT CONCORDE I

May be fitted with or without a tailplane. If there is a tailplane then the aircraft will have the normal flying controls - elevator, rudder and ailerons.

If there is not a tailplane, Concorde for example, then the ailerons and elevators are combined and called elevons.

Elevons are fitted to the trailing edge of the delta wing. For climb and descent they both move up and down together. For roll control they operate a s ailerons - ie, in opposite directions. For a combination of roll and climb/descent the inputs are mixed. For example, if the control column is pulled back and to the left (climb and roll to the left) then both elevons move up but the left elevon moves further up than the right elevon.

- 49 -

rnoodull l A-128

HIGH SPEED FLIGHT

When studying the theory of flight, High Speed Flight (HSF) is considered to start at the onset of MCRIT. This speed varies with the local speed of sound which could be considered as 762mph at sea leveI under standard atmospheric conditions, and the design of the aircraft. At speeds close t o this compressibility problems start to make themselves felt. Up to this speed the study of theory of flight is considered to be Low Speed Flight.

In fact the definitions could be refined still further to:

* Low speed flight - subsonic flight - up to MGRIT. * Transonic flight - going through the sonic range from MCRIT (about

M = 0.7 - depending on conditions and aircraft type) to about M =

1.2. * Supersonic flight - all speeds above M = 1.2.

The Speed of Sound - .

A s HSF is related to the local speed-of sound it is important to have a good knowledge of how sound travel& through air. Sound is transmitted through air as a series of sound waves and the speed the waves travel at is related to the air temperature and is calculated from the formula:

I I - I I

where a - speed of sdwd K = a constant T = absolute tcmperature (Kelvin)

-" ..

The speed of sound in air at sea level at stp (standard temperature and pressure - stp) is 762 rnph - the temperature taken as 15°C. This reduces to 660 rnph at 36,000 feet (tropopause) where the temperature is takcn as -56°C. So with increasing height the speed of sound decreases. Converting these values to metric givcs:

762 mph - - 1229 km/h 660 rnph - - 1062 h / h 36,000 feet - - 11 km

Definitions

Subsonic Speed. The aircraft is travelling up to speeds of about M = 0.7 where all the air over the aircraft is subsonic.

Transonic Speed. The aircraft is travelling at speeds between about M = 0.7 and M = 1.2 where some of the air moving over the aircraft is supersonic and some is sub-sonic.

moodull l A-129

Supersonic Speed. The aircraft speed is generally above about M = 1.2 where all the air moving over the aircraft is supersonic.

The Sound Barrier

In the early days, during the period immediately prior to the introduction of the jet engine, aircraft had flown u p to the speed of sound (usually in a dive) and so many problems occurred, including a significant increase in drag, that it was thought that there was a 'barrier' to going any faster. It was called the Sound Barrier. Even today, going faster than the speed of sound is usually mentioned as going through the sound barrier.

Of course there is no barrier, but for aircraft not designed to go faster than the speed of s o ~ ~ l d there are considerable problems in passing through this speed range.

Mach Number (M) I

-

Named after Ernst Mach Austrian physicist 1838 - 19 16. The Mach number refers to the speed that an aircraft is travelling compared to the local speed of sound, ie

I

M = true airspeed --

local speed of sound I

The speed of sound decreases with height (up to 36,000ft - the tropopause). I

For example, if an aircraft is flying a t 700mph a t seal level (stp) its M number would be: - -- - -

If it is flying a t 700mph a t 36,000ft (stp), its Mach number is:

The airspeed is indicated to the pilot by an airspeed indicator (ASI) while the Mach number is indicated on a Machmeter (which gives an accurate reading a t all altitudes). This 'sums' the values of airspeed (corrected for density) and altitude.

From the above you can see that if the aircraft maintained a constant speed and climbed then, starting a t 700mph at sea level, it rxould eventually go through the sound barrier although its actual speed would not increase.

rnoodull l A-1 30

Similarly on descent (at a constant true airspeed) its Mach number would decrease.

Mcrit

I t is important that the pilot knows the speed of the aircraft in relation to the local speed of sound as at that speed he/ she is likely to encounter significant problems in t he control characteristics of the aircraft.

As the aircraft approaches the speed of sound so some parts of the airframe become supersonic before others. This supersonic flow will occur over the lwger cambers of the airframe such as the tops of the wings. This supersonic flow will cause shock waves and turbulence and may cause the aircraft to exhibit some or all of the following:

* Vibration and buffeting - sometimes severe. -k Stability problems. * Control ineffectiveness. * Control reversd. ' . - - -

* Inability ta control thc aircraft (particularly with manual controls) * Turbulence. I '

Turbulence wilI occur behind the shock wave and this can affect the tail plane which will cause buffeting, pitch . Control - . and stability problems.

Because of the high acrodynarnic loads on the control surfaces at high spced, -

the control surfaces can act like trimming tabs on the main aerofoil surface. If this is associated with lack of rigidity in the structure the wingltail plane/ fin can twist-about its torsional axis and produce the opposite aircraft response to that desired by the movement of the-cockpit controL-md control surfaces.

Fig. 38 AILERON REVERSAL

With reference to figure 38. When the aileron goes up the expected response is that the complete wing moves down. If, however, the wing is not torsionally stiff enough it will twist about its torsional axis, increasing its angle of incidence (and angle of attack) thus causing the whole wing t o go up - thus the aircraft waves in the opposite direction to that intended. N o t a good state of affairs.

In the 'early days' this was a problem with some aircraft. Today, however, as designers know more about structural stiffness and its relationship to aerodynamic loads, the design engineers can anticipate the difficulties and design the airframe stiff enough to withstand the twisting loads.

So it is important that the pilot knows his speed in relation to the local speed of sound so he/she can keep his/her speed down below Mcrit - the speed a t which the first supersonic flow occurs around the airframe. Hence the fitment of a Machmeter to all aircraft that are likely to fly near to their Mcrit.

Of course, this only applies to aircraft that are not designed for supersonic flight. Those that are have enough power to overcome the significant rise in drag and are designed to have high MCRIT values. Aerodynamically they are also designed to pass through the transonic range with little adverse affect.

TRANSONIC SPEED

When a6 aircraft passes through the air a t a speed lower than the speed of sound ,(66 1 knots stp a t sea level) it sends out pressure waves ahead of it that 'warn' the air of the aircraft's approach. Thus the air starts to move out of the way of the oncoming aircraft before the aircraft actually gets there.

I I

I

At speeds below the speed of sound the air behaves as if it is incompressible, but as the speed of sound is approached so the behaviour df the - --- air gradually chang7s1 At high speed the air wiil compress or expand as nelessary (refer table 4). I

I I

SPEED IN KNOTS COMPRESSIBILITY I - - ERROR %

80 0.5 260 4 440 11 520 16

TABLE 1

In very general terms it is considered that the air acts as if it is not compressible a t speeds below the speed of sound and is compressible a t speeds above the speed of sound - not strictly true, but as a rough guide it is okay.

A moving object will send out pressure waves in all directions a t the local speed of sound. These pressure waves warn the air that the object is coming and the air is prepared to move out of the way.

If the object is moving through the air at the speed of sound then the pressure waves being sent out ahead of it do not move f k a r d ahead of the object but build up in front to form a Bow Wave. The air in front now has no warning that the object is coming and hence there is sudden change in pressure and velocity as the object hits the air.

rnoodull l A-1 32

SOUND

wAm w STATIONARY OBJECT

SOUND WAVES STIU MOVING MRWARD MSTERTHAN THE S P W OF THE OBJECT OWEET MOVING FORWARD

SLOWER THAN THE SPEED OF SOUND

WE BEGlNNlNG OBJECT MOVING FORWARD OF A BOW WAVE AT THE SPEED OF SOUND

Fig. 39 SOUND WAVE PROPAGATION

Shock Waves - . . - -. - . .

As the spied of the airflow increases over an aerofoil or stre-lined shape so the bdundary layer of air starts tb break.away from the surface towards the rear. This may start at low airspeeds and the position at which it starts on the aerofoil is called the Transition Pbint. As 'the speed increases :so this point moves fonvard and the boundw -- layer -- gets thicker. I

1 i ' -- - - - - = - , -

LAMINAR FLOW TURBULENT WAKE

. . .. - -- . - . . . Fig, 40 TRANSITION POINT (M w 0.6)

A s the speed of the aircraft increases to about MO.75 an incipient shock wave farms. This occurs at the point of maximum camber (usually on the top of the wing and some points on the fuselage near the flight deck),

SUDDEN INCREASE IN PRESSURE. AIRSPEED

INClPf NT SHOGK WAVE DROPS TO Mcrit

AIRSPEED HIGHER T HAN M I DECRE ASlnlG PRESSURE

PRESSURE INCREASlNG TO THE TRAlUNG EDGE

Fig. 41 INCIPIENT SHOCK WAVE [M = 1.0)

It causes a sudden rise in pressure and density of the air and a drop in velocity. There is also a tendency for the turbulent wake to move forward and start a t the point where the shock wave attaches itself to the aerofoil.

The Shock Stall

The shock wave causes a sudden large increase in drag (by as much as a factor of 10)) and a loss of lift.

The change in the pressure distribution around a conventional aerofoil causes a nose down pitching moment of the aircraft and the turbulent airflow behind the shock wave causes severe buffeting - particularly if it hits the tailplane. The effects are similar to a stall and it is often called a shock stall. This condition is not confined to the aerofoils but can occur to any part on the aircraft.

The ordinary stall is often called a high incidence stall to distinguish it from a shock stall. This shock stall is often called 'tuck under7.

Shock Drag !

~enerall; considered to be made up from wave drag and boundary layer drag. I I

profile drag a t subsonic speed varies with the square law (drag = Cd % pV2S) but a t transonic speed the square-law breaks down. As can beseen from the graph the Cd rises rapidly during the transonic period, but decreases thereafter until it becomes steady at about M = 2. But the Cd is still a t least twice that a t which it was at subsonic speed. I

-- --

0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0

MACH NUMBER

GRAPH 4 - Cd AGAINST MACH NUMBER

Critical Mach Number (Mcrit)

As the speed of sound is approached at below M = 1 compressibility effects are felt - lass of lift; buffeting; loss of control etc. The speed at which this starts is called the Critical Mach Number (Mcrit). In general the higher the Mcrit for an aircraft the better.

To increase the Mcrit aircraft are designed with slim wings and a slim fuselage with no 'lumps' to produce local shock waves. Swept wings also help, and having a high tailplane tends to keep the tailplane out of the turbulent air from the wings.

The Machmeter will have a lubber line indicating MCWT and for slow speed aircraft the pilot will not fly the aircraft beyond this mark under any circumstances (on most modern aircraft high speed warnings will sound).

Mcrit is also specified in the Pilot's Notes.

Flying ~ h r o u ~ h the Transonic Sb&& Range

A s the aircraft becomes transonic so there is a considerable change in the longitudinal trim - usuaIly a nose down pitch, This is accompanied by buffeting; lack of effectiveness of t h e t r i u n g devices and a considerable increase in the force required to move the controls.

I

I

The aircraft may:

Snake - Yaw from side b side. Porpoise - Pitch uyj and down. Dutch Roll - A combination or roll and yaw.

The aircraft may become difficult to control, The control surfaces are usually behind the shock wave and in the turbulent area, therefore, they may not be very effective. As the speed increases so the shock wave moves back over the aerofoil and over the control surface - this makes them difficult to move.

In some cases control surface reversal occurs. The lack of rigidity in the structure tends to move the structure about its flexural centre line.

Better airframe design, slab tailplanes, powered flying controls etc, will help overcome some of these probIems.

Increasing the Mcrit of an Aircraft

If the fuselage and wings are kept as slim as possible this will increase the Mcrit. On a wing this is called thicknesslchord ratio - and a wing with a lower % t/c ratio has a higher Mcrit,

THICKNESSICHORD 6 RATIO (%)

0.8 0.85 0.9 0.95 1 .O

Mcrit

GRAPH 5 - MCRIT AGAINST T/C RATIO

If the wing is swept this will also increase Mcrit. This is because the shock wave is caused by that component of the airflow running parallel to the chord 1 : 1 T - - -1 -- +L- --,A+,, +L, ,,,,,,,L,-.-lr +L- l..:-L--- + L A nrT,,:+ r l r lc (IL v LUD - au LIIL g L L a L \ r l LLLL ~ W L L ~ U Q L A LILL 11lf;ll~l L I ~ L LVLLILL.

Fig. 42 COMPONENTS OF AIR VELOCITY ON A SWEPT WING

Another advantage of the swept wing is that it has a lower Cd (graph 5).

Sweepback also helps to keep the wing tips within the leading edge shock waves (shockcone) created by the nose of the aircraft.

Swept wings also help to main lateral stability about the longitudinal axis at low speed (low speed theory of flight).

- 57 -

rnoodull l A-136

0.6 1 .O 1.2 1.4 1.6 1.8 2.0

MACH NUMBER

GRAPH 6 - Cd AGAINST MACH NUMBER

The swept wing has, however, several disadvantages:

(a) I t is prone to tip stalling.- -.

( 1 ' ' Ct ;I;= is I3-;v' therefEire ?atl&ir=g U~C.LUU c.*~-PAc. C C L ~ h;dh A L A ~ ~ .

- -

,(c) Angel of attack is high for CL max. (d) Bending stresses are high so structurally more difficult to design.

Going Through Mach 1 (figure 43) - -

The first signs of the incipient khock wave okcurs at below speed of sound, and on a symmetrical aerofoil !he wave wiU appear on bothmtlie top and bottom surfaces. For wings with a more cambered top surface the first wave starts on the top. As the speed increases sd the shock wave gets stronger and tends to move towards the trailing edge. - -

(The shock wave is sometimes described as a pressure wave made up of d l the sound waves moving forward from that part of the wing and when they meet air coming back at the speed of sound they cannot travel any further l o w a r d so build up to form a pressure wave.)

At just above M = I a bow wave forms in front of the aerofoil and the tail wave becomes curved and attached to the trailing edge.

As the speed increases further so the bow wave attaches itself to the leading edge and the angle of both waves becomes more acute. A t each wave there is an increase in pressure, density and temperature, and a decrease in velocity.

All air is supersonic at about aircraft speed M = 1.2.

Nbte the rearward movement of the centre of lift or centre of pressure.

SUBSONIC FLOW ,CENTRE OF PRESURE (C of P)

M = 0.6 NO SHOCK WAVE FORMED YE1

SUBSONIC FLOW

INCIPIENT SHOCK WAVE

-+----- (ON BOTH SIDES OF THE

SUPERSONIC AEROFOIL AS IT IS

AIRFLOW SYMMETRICAL)

SUBSONIC FLOW

M = 0.8 C of P = STARTS TO MOVE REARWARDS

SUPERSONIC SUBSONIC FLOW AIRFLOW \ P = SUDDEN INCREASE

P = SUDDEN INCREASE V = DECREASE

SUPERSONIC / AIRFLOW

M = 1.0 SHOCK WAVE NOW MORE FULLY DEVELOPED AND MOVING REARWARCIS

I

AIRFLOW

I BOW WAVE STARTS TO I

FORM AND APPROACHES FROM THE FRONT THE INCIPIENT SHOCK WAVE

NOW MOVED BACK TO THE TRAILING :EDGE

--

I I

FULLY DEVELOPED BOW AND TAIL WAVE ATTACHED TO THE AEROFOIL ALL AIR SUPERSONIC

Fig. 43 SHOCK WAVE DEVELOPMENT - BI-CONVEX AEROFOIL

Area Rule

In a n attempt to keep the drag as low as possible during the transonic period the area rule may be applied to the design of the aircraft. This states that the total frontal cross sectional area of the aircraft (including wings, tailplane and engines) should increase gradually from the front of the aircraft to the middle, then reduce slowly to zero at the rear. Thus where additions are fitted to the fuselage, such as wings and tailplane, the fuselage should be wasted.

Fig. 44 AREA RULE

SUPERSONIC SPEED

The Mach Cone (figure 45) - .- .

If an aircraft moves at a velocity-V! which is greater than t he speed of sound. it will send out pressure waves in all directions continuously atl every point along its path. Thus it will send out pressure waves at, say, points A, B, C and D. These~waves will have travelled to a poht on the line Dl3 during the time the aircraft has travelled from A to, D.

VELOCITY bF AIRCRAFT (Vl - Fig, 45 MACH ANGLE (MACH CONE)

The angle sin o~ is equal to g = 1, V M

where a - - speed of sound V = Velocity

And this is called the Mach Angle.

Mach Angle = sin x = a = 1 = angle ADE V M

The faster the aircraft goes the more acute the angle becomes. At M = 1 K = 90.

The line DE is often called the Mach Line. Mach lines are developed from many points on the aircraft,

rnoodul I l A-1 39

If the speed of the air over the aircraft is constant then all the Mach lines will be parallel. If the airflow is accelerating then the Mach lines will diverge. And if the airflow is decelerating then the Mach lines will converge.

This gives some indication of how a shock wave is formed.

AIRFLOW CONSTANT AIRFLOW AIRFLOW \'ELOCIPI ACCELERATING DECELERATING

Fig. 46 MACH LINES

I

--

Supersonic Flow

In sub'sonic flow the air will anticipate objects in it's path ahd make changes gradually. It will also behave in accordance with Bernoulli's theorem (see the books in the LBP series on Science/Physics/Theory of ~ l i ~ ~ t ) . ' In supersonic flow things are different.

-

I 4

Supersonic flow may be divided into compressive flow and dxPansive flow. I

I ,

--

Compressive-Flow (figure 47)

Consider the supersonic airflow meeting the wedge angle of the leading edge of a mainplane. It will not anticipate the oncoming corner but will continue until it is forced to move by the wedge angle itself. At this point the air slows, and temperature, pressure and density increase. An oblique shock wave is formed.

This condition will occur whenever supersonic air meets:

* The leading edge of a wing - whether it be rounded or wedge shape.

* The nose of the fuselage. * Leading edges of tailplanes, fin, engine intakes etc. * The beginning of a contracting duct, as in engine intakes. * Any concave corner of the airframe.

- 61 -

moodull l A-140

SHOCK INCREASED PRESSURE INCREASED DENSITY INCREASED TEMPERATURE LOWER VELOCITY

AIRFLOW , P

WEDGE SHAPE LEADING EDGE

Fig. 47 COMPRESSTVE FLOW

Expansive Flow (figure 48)

When supersonic air passes over a convex surface/corner it is caused to move faster around the corner - othenvise a vacuum would form. This sudden change in direction causes the reverse to happen as happened with compressive flow. There is a reductiop in pressure, density and temperature, and an inrr~ase in velocity. At the same time Mach lines &e formed.

~e tween the old and new Mach lines the air follows a curveld path andthe second Mach line is at a more acute angl'c than thc first.

- - OLD NEW MACH MACH LINE. , LINE

SUPERSONIC AIRFLOW

.- -b LOWER PRESSURE LOWER DENSITY LOWER TEMPERATI

HER VELOCITY JRE

Fig. 48 EXPANSrVE FLOW

There may be many Mach Iines to a curved surface and the air flows through these (which are weak compared to shock waves) without sudden changes in direction or physical properties. The condition between these lines is called an expansion wave.

The angle the air can move through in any one expansion wave is small, but it can be moved through a large angle by passing through a succession of expansion waves, eg around a large curved surface.

moodull 1A-141

Supersonic Aerofoils

Figure 49 shows a typical high-speed symmetrical aerofoil a t a small angle of attack. Notice the shock waves where there is compressive flow and many Mach lines where there is expansive flow. There is also no upwash and no downwash.

MACH LINES \ -

\ - -

Fig. 49 SUPERSONIC FLOW OVER A SYMMETRICAL AERO'FOIL

I --

Curved surfaces are very good for-low speed flight but for supeTsonic flight straight surfaces offer a better solution. For example, a double-wedge aerofoil may be used.

1

Figure 50-shows a double wedge aerofoil a t zero angle of a tkck. The flow patterns are symmetrical with shock waves at the leading and trailing edges and expansion waves a t the point of maximum thickness. At the shock waves pressure, density and temperature increase (with a decrease in velocity) - and a t the expansion wave the reverse will occur.

AIRFLOW ____lt

\ COMPRESSION

\

SHOCK WAVE

MACH LINES 1 I -

Fig. 50 SUPERSONIC FLOW OVER A DOUBLE WEDGE AEROFOIL

- 63 -

r n o o d u l l l A-142

As the angle of attack increases so the top leading edge (L/E) shock wave becomes weaker, as does the bottom trailing edge (T/E) shock wave - this is because the wedge angle is becoming smaller.

At an angle of attack equal to half t he wedge angle the top L/E and bottom T / E shock waves disappear altogether. The other two shock waves get stronger because their wedge angles are greater.

Figure 51 shows the aerofoil at the best angle of attack for the best L/D ratio. (The best lift/drag ratio. The most lift for the least drag.)

MACH LINES SHOCK WAVE

- -

AIRFLOW - I

SHOCK WAVE \ \ MACH LINES

Fig. 51 S U ~ R S O N ~ C FLOW AT ANGLE GMNG--BEST LJD RATIO -.

If the angle of attack is increased still further the bow wave: bkcomes detached and expansion waves form on thk top L/E and bottom T/ E areas (figure 52).

Fig. 52 DOUBLE WEDGE AEROFOIL AT LARGE ANGLE OF ATTACK

Supersonic aerofoils may take various forms. The ideal would be a flat plate but this is impractical structurally. Bi-convex aerofoils and various 'straight' aerofoils may be used. Ideally they should have a low t /c (finesse) ratio (thickness/chord ratio); have straight lines and have no projections or bulges. The problem with all high speed aerofoils is that they must perform reasonably well a t low speed to allow the aircraft to take off, land and fly through the subsonic speed range. So most high speed aerofoils are a compromise - to give reasonable low speed characteristics with good high speed characteristics.

DOUBLE WEDGE X = 50%

AEROFOIL

61-CONVEX AEROFOIL t 1 = 10%

C -

- , I

I HEXANGONAL L = 10% AEROFOIL c

1 '

I

- -

Fig. 53 HIGH SPEED AEROFOILS

\ I i I

The x /c ratio (thickness/chord ratio figure 53) of a typical 9erofoil for high- speed flight can be between 40% and 60% for least drag. These values should not affect the C of P or the lift significantly. For low speed flight the best x/c ratio is between 30% and 40%.

The 'hexagonal' aerofoil is probably stronger and has a %lunterY leading edge compared to the double wedge. This has advantages when it comes to kinetic heating (heating of the aircraft's skin a t high speed).

A bi-convex aerofoil is better than the others at subsonic speeds and produces the same drag as a double wedge with an x/c ratio of 25%.

Lift/drag ratios for a conventional low speed aerofoil a t low speed can be u p to about 24, while the lift/drag ratio for a high speed aerofoil above M1.3 is about 12 (very poor).

The lift coefficient, while low, is similar for all the high-speed aerofoils and because of this it does mean high landing, take off and stalling speeds.

WING PLAN FORMS

At supersonic speed the wing plan form is very important.

Sweepback delays the shock sbll and increases the Mcrit. I t also means that the leading edge of the wing will be within the Mach Cone (if it is swept enough). This will put the leading edge behind the Mach Line and, if there are shock waves at the nose of the aircraft, within a region of air that is moving slower than the rest of the airflow (though it may still be supersonic).

To maintain this rule t h e faster the aircraft flies the more acute the sweepback must be - until a delta shape is formed.

The more sweepback there is the more structural problems there are - and of course there is tip stalling and lack of lateral control.

Delta wing plan forms have an advantage as regards strength/weight ratio because of their long chord length they can have quite a depth 01 wing and still have a - good t/c ratio. The depth of wing gives greater bending strength.

- .- . . - - . . - - -

Conventional swept wings tend, however, to create less drag than delta wings. With wings swept at more than say 55" an pdvantage is gdned in respect to lift and drag during the high-speed stall. The leading edge stall starts at the wing tips and works inboard - the bubble that is formed is thenswept back along the leading edge to form a trai$ng%dge vortex. This low pressme acts on the upper: leading edge and creates lift at the same time creates-a form of thrush (negative drag) - and because it is laminar causes little or no buffeting.

The striight wing has an advantkge when it comes to drag!at high Mach numbers. As can be seen from graph 7 the total amount of drag from a straight wing becomes less than a swept wing after about M 1.6. The wing tip will produce a Mach Line and that means that part of the wing will be within the cone of the Mach Angle (but only a small part).

DRAG

1 '12 1.4 1.6 t.8 2.0 MACH NUMBER

SWEPT WING

STRAIGHT WING

GRAPH 7 - DRAG AGAINST MACH NUMBER

Variable Geometry Wings

To try and overcome some of the problems associated with straight and swept wings a variable geometry wing plan form may be used. Thus for landing and take-off the wings are moved to the straight position while for high-speed flight the wings are moved to the swept position. This increases the structural problems and makes the aircraft technically more complex.

Typical Wing Plan Forms

Figure 54 shows some typical wing plan forms. (a) is the plan view of a YAK-40 with a maximum design cruising speed of M0.4 so is unlikely to have any problems as regards Mcrit. However, some high performance aircraft do have straight wings. These are usually short (low aspect ratio), placed well back on the fuselage and their tips are within the Mach cone created by the nose of the aircraft.

Figure 54 (b) shows a wing plane form of a typical airliner (B747). The 747 has a maximum cruise of M0.76 so there could he a possibility nf flying c l n s ~ tn its Mcrit. F'or many aircraft in this range the sweepback plays an important roll in giving thle aircraft stability about the longitudinal axis (lateral stability).

' I

Fig. 54 WING PLAN FORMS

' Figure 54 (c) shows the wing plan form of Coricorde. It is a delta wing aircraft without a tailplane. Some delta wing aircraft also have a tailplane (stabiliser) normally fitted in the form of a T tail.

moodull l A-146

Note the acute angle of the shock cone or Mach cone - with the wingtips designed to be within the cone. Concorde had a design top speed in the region of M2. Delta wings have a poor coefficient of lift (CL), particularly at low speed, so on landing and take-off they have a high angle of attack to obtain the required lift. For aircraft like Concorde this necessitates a long landing gear.

Figure 54 Id) shows the Mirage G fighter which as a top speed of M2 -5. The drawing shows the wings in the high speed position w i ~ the dashed profiles showing the extended position for landing and take-off. Remember, straight wings have a better coefficient of Lift (CL) than swept or delta wings particularly at slower speeds.

Note - the plans of the aircraft in figure 54 are not to scale.

KINETIC HEATING

A s aircraft move through the air at high-speed so heat is created. This comes from three sources - skin friction, air compression and shock waves.

- . - - .

Friction in any form. creates heat ana aerodynamic skin friction is no exception. Heat is always created this way when an aircrdt flies throdgh the air - at low speed skin friction is low and the heat created is very srnali. At high-speed the reverse is true.

- -

-

When air is compressed (cg the stagriation point at the leading edge of the aerofoil is compressed) its temperature rises - it is, afterall1 an adiabatic compression (ie does not lose o,r gain heat from another sovce) .

, ' I

I I

This fork of heating accounts for why the leading edges tend to get hotter than the rest of-the-aircraft. . .. .

TEMPERATURE "c

0 1 2 3 4

MACH NUMBER

GRAPH 8 - TEMPERATURE AGAINST MACH NUMBER (28,000ft)

Shock wave heating mostly tends to heat the air and has only a small effect on the aircraft skin temperature. So the faster the aircraft flies the more acute the problem of kinetic heating becomes - particularly skin friction heating and shock wave heating.

The formula:

gives the approximate rise in temperature in degrees Celsius where

t - - rise in temperature in degrees C and v - - speed in knots.

The rise in temperature can cause problems with:

(a) Structures. (b) Systems - fuel etc.

PVDTXT , n A ncaccongers. ' V I U " " U l L U lV! p"""""

(d)' Freight. I

The solubon for (b), (c), and (d) is insulation or some form qf cooling but (;I) is more difficult to solve.

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I I l 1

The stkucture cannot be allowed to heat up to temperature$ that-affect the mechdnical properties of the metals. So if the structure cadnot be artificially cooledor protected in some way, then limits to speed must Ibe imposed depending on the materials used; eg l l

Aluminium Alloy - about M 2 Titanium Alloy - about M 4 Stainless Steel - higher than titanium alloy Ceramics - higher than stainless steel

In the design of the aircraft, to reduce kinetic heating, it is better to have wave drag than boundary layer drag and to avoid all sharp corners. At supersonic speeds the boundary drag is relatively unimportant compared to its effect at low speed.

STABILITY &, CONTROL AT SUPERSONIC SPEED

Some of what has been said about stability and control in the transonic region applies to this area, particularly when it comes to control. Stability is complicated a t these speeds by the high inertia loads. Natural inherent stability provided by such things as effective keel surface and dihedral angle is not as effective. To increase the size of things like fins and wings to increase stability would only increase weight and drag (and increase the inertia loads).

rnoodull l A-148

The aircraft's inherent stability can be assisted automatically by such things as:

* Mach trim. * Auto pilot. * Active stability.

All these stability inputs coming via the flying control surfaces - or all-flying tailplane (slab tailplane) or all-flying aerofoils. All-flying aerofoils move similar to an all flying tailplane, they are rare.

Mach Trim

As we have seen there is a tendency for the aircraft to pitch nose down (Tuck under) when it moves through the transonic region. This is because the Centre of Pressure moves rearwards. A Mach Trim system is fitted to many high-speed aircraft to compensate for this by putting a signal into the longitudinal controll trim system during t he transonic period.

If Mach trim is via the tailplane it would cause the tailplane to decrease its angle of incidence (and hence its angle of attack) so putting a down-load on the tail and raising the nose.

I , . . . . . .

Auto pilot , ,

This fully automates the flying cqntroi system of the aircraft using computers and servos fitted into the flying control systems. In some aikraft Mach trim would be via t he autopilot system.

Active Stability

Most fmed wing aircraft are inherently stable [not so most helicopters). A fixed wing aircraft will return to its normal flight path without assistance by the pilot. This inherent stability is achieved by making the aircraft aerodynamically stable by the use of such things as Effective Keel Surface (Directional Stability); Dihedral Angle (Lateral Stability) and Tail Plane (Longitudinal Stability).

Somc modern aircraft - particularly military hardware - use Active Stability. This means that the aircraft is kept stable by active intervention of the controls. When the pilot moves the controls - usually via a computer system - the aircraft will respond.

If the aircraft is disturbed in flight laser gyros will sense the movement and inform the computer. If the pilot has not commanded the movement the computer assumes that t he aircraft has been disturbed aerodynamically which sends a signal to servos which move the powered control surfaces to return the aircraft to its normal flying altitude.

The advantages of active stability are:

* Quicker response times. Inherent stability is opposite to controllability. If the aircraft is very stable it is not very responsive to the controls. With active stability systems the aircraft can be made inherently unstable and be very aerobatic. Ideal for fighter type aircraft.

* Better structural strengthlweight ratios - the aircraft can be designed as a structure with little or no consideration for aerodynamic stability (note some of the "square" shapes of some of the more modern military aircraft).

A The aircraft can have varying levels of stability - just by the flick of a switch.

It is, of course, a more complex system with the safety of the aircraft depending on the correct qerat ion of the hardware and good software, and reliable supplies of electrical and hydraulic power. The systems are usually triplicated and may have up to 5 or 6 levels of redundancy with software designed by different software houses to hope the prevent the duplication of -- software errors.

Possibly the first use of active stability in civil aircraft was t h e ~ 3 2 0 gust alleviation system. If the aircraft is caused to roll by aerodynamic forces, the laser gyros pick this u p and serid a signal to the flying control computers (triplicated for each channel) who will know that the pilot has not commanded the manoeuvre (all his/ her flying control inputs go via the computers). The computer/s will send a signal tb the spoilers on the u p goirig wing; these spoilers kill deploy, destroying lift and bringing that wing down.

-- -- - -- -

Two features are worth noting about the design of the intake; (1) the bleed-off of the boundary layer if the intake is close to the fuselage or wing of the aircraft and (2) the requirement to slow the air down to a n acceptable velocity for entry into the compressor stages of the engine.

The Boundary Layer

As described earlier it is a layer of slow moving low energy air on the skin of the aircraft that gets thicker the further back it travels. The problem occurs a t all speeds and effectively the layer of air molecules next the aircraft's skin is almost stationary with respect to the aircraft. The next layer of molecules is moving slightly faster and the next layer after that is moving slightly faster still and so on. This state of affairs will continue for a centimetre or two until we get to the free-stream flow where all the air is travelling at the same speed. Should it become1 turbulent, as happens at the t ransi t i~n point then the boundary layer gets thicker and even more sluggish. This boundary layer is a problem.

moodull 1A-150

To try and get more energy into this layer vortex generators are fitted to some parts of the aircraft skin - normally the top side of the mainplanes near the leading edge. These are lines of small pieces of metal sticking up into the airflow by about an inch (25mrnj. They are at an angle to the airflow so they cause the free-stream flow to mix downwards into t he boundary layer to liven it up and give it more energy. May be fitted up-stream of the ailerons (or any other control surfaces) to improve their effectiveness.

If there is an engine intake close to the skin of the aircraft then this layer of air, if it got into the engine, would adversely effect its performance - so it is bled away from the engine through air ducts to the outside, or the intake is designed so that it is clear of the fuselage. For example, the rear engine intakes of the MD11 and TriStar and the Eagle fighter (figure 55) and Concorde [figure 56).

Where the intake is actually on the side of the structure then provision wiU be made within the intake to bleed off the boundary layer through ducts to the outside.

, 1 , \ BOUNDARY I

LAYER BLEED

Compressor Entry Speed

Fixed configuration circular intakes work well for aircraft travelling at subsonic speeds. As the aircraft approaches supersonic speed, however, shock waves will form in the intake and engine performance will be considerably reduced.

At speeds up to about M 1.4 the shock waves do not have much effect on the pressure recovery of the intake, as down-stream of the normal shock wave the velocity is always subsonic (during the transonic speed range of the aircraft).

As the aircraft Mach number increases so the pressure recovery behind the shock wave drops away and methods have to be found to restore the performance of the intake.

For supersonic flight, intakes have to be designed to create shock waves when required. This means that intakes must be variable geometry intakes and mechanical means must be found to change the shape of the intake to produce different pattern shock waves at different airspeeds.

BOUNDARY LAYER BLE

TAKE-OFF

SECONDARY AIR VALVE

SUPERSONIC

SHOCK PATTERN DIFFUSER

RAMP ASSEMBLY

SHUT-DOWN I

I I

Fig. 56 SUPERSONIC INTAKE - CONCORDE I

I I I

For circular intakes this was achieved by having a moving centre cone within the intake. Most modern intakes are of the 'square' type, which,allows easy adjustment-of square panels within the intake to control the pattern of shock waves.

The variable geometry intake is designed to supply the engine with the correct quantity of air at all times and to reduce the velocity to a subsonic value a t the compressor inlet.

Typical of the square section intake is that shown in figure 56. The intakc has a movable ramp assembly, an auxiliary door, a secondary air valve and a subsonic diffuser. The moveable ramp assembly creates the shock waves which will reflect within the intake to reduce the airspeed to subsonic values even when the aircraft is well within the supersonic region.

In supersonic flight the ramp assembly is lowered to focus the pattern of shock waves formed on the intake lip to obtain subsonic flow at the throat. Further compression and reduction of air velocity is obtained in the subsonic diffuser.

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rnoodull l A-1 52

Changes in engine aidow demand caused by varying ambient conditions are met by spilling excess air over the intake lip. This ensures maximum intake efficiency and good enginel airspeed matching.

Wi th engine shutdown in flight the ramp is lowered as far as possible and the auxiliary door opened to dump excess air overboard, and the secondary air valve is open. This reduces the chanccs of instability in the enginc and reduces drag.

During take-off and subsonic flight the engine requires maximum mass airflow, so the ramp assembly is fully raised, the auxiliary door open and the secondary air valve shut.

LBP Dec 03 - Feb 04 - Apnl04

Addendums module 11A books STRUCTURES pending amendment action in response to student feedback after taking the CAA examinations.

*** If there is partial failure of the passenger emergency lighting system the aircraft can be dispatched provided the passenger compliment is reduced to that number that can be carried in that part of the aircraft with a serviceable emergency lighting system. Check the MEL for the specific aircraft. The maximum number of emergency lights that can be out is 25% (www2. faa.gov/ certification/ aircraft).

*** A similar regulation applies to inoperable passenger exits. The passenger compliment is reduced and passengers are not seated near that exit. Again the MEL is consxlted.

*** Painting radomes. Use non-metallic acrylic lacquer or polyurethane. For neoprene coated surfaces the use of cellulose based paint is not recommended

i a s it is likely to attack the neoprene. Probably a module 7 question. I I ***

-

Alodizing is a n anti-corrosive treatment for aluminium alloys. The part is cleaned with a n acid or alkaline solution and rinsed with clean water. The part is then treated with Alodine solution (a propriety solution similar to Alochrome) which1r;sults in a hard greenish finish. After another watea wash the part is then treated with Deoxylyte (a proprim brand solution to leave the surface slightly acidic). Almost certainly a module 7 question.

I *** I

passedger seats may face in any direction. *** I

Aluminium alloy 2024 is an American specification and is used widely in airframe construction.

*** Acrylic windows have a larger coefficient of linier expansion than A1 alloy structure .

*** Kevlar is stored in moisture proof bags.

*** A heavy landing is likely to cause hogging (the fuselage bent so that it is high in the middle and low a t the nose and tail).

*****

NOTE: It is possible that some of the above statements may not be too meaningful when read out of context, so it is suggested that the appropriate book/ subject be read first then the information above be checked against that topic. -

(1) moodull l A-154