Easa Part 66 - Module 11.09 - Aerodynamics Structures and Systems

Page Mod 11.9 Issue 1

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MODULE 11.09

AERODYNAMICS, STRUCTURES AND

SYSTEMS

engineering

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Contents

1 PRIMARY FLIGHT CONTROLS ................................................... 1-6

1.1 AILERONS .................................................................................... 1-8

1.2 ELEVATORS ................................................................................. 1-8

1.3 RUDDERS .................................................................................... 1-9

1.4 SPOILERS .................................................................................... 1-10 1.4.1 Flight Spoilers ............................................................... 1-11 1.4.2 Ground Spoilers ............................................................ 1-11

2 TRIM CONTROLS ........................................................................ 2-1

2.1 FIXED AND ADJUSTABLE TRIM TABS ............................................... 2-1 2.1.1 Fixed Trim Tabs ............................................................ 2-1 2.1.2 Controllable Trim Tabs .................................................. 2-1 2.1.3 Servo Tabs.................................................................... 2-2 2.1.4 Balance Tabs ................................................................ 2-2 2.1.5 Anti-Balance Tabs ......................................................... 2-3 2.1.6 Spring Tabs ................................................................... 2-3

2.2 FULLY POWERED FLYING CONTROL TRIM SYSTEM ........................... 2-4 2.2.1 Typical Trim System ...................................................... 2-4 2.2.2 Rudder Trim System ..................................................... 2-4 2.2.3 Aileron Trim System ...................................................... 2-4 2.2.4 Tailplane Trim System .................................................. 2-5

3 ACTIVE LOAD CONTROLS ......................................................... 3-1

3.1 ACTIVE LOAD CONTROL ................................................................ 3-1

3.2 ACTIVE CONTROL TECHNOLOGY .................................................... 3-1 3.2.1 Advantages of Active Control Technology ..................... 3-3 3.2.2 Direct Lift Force ............................................................. 3-3 3.2.3 Direct Side Force .......................................................... 3-3

4 HIGH LIFT DEVICES .................................................................... 4-4

4.1 FLAPS .......................................................................................... 4-4

4.2 SLATS .......................................................................................... 4-6

4.3 DROOPED LEADING EDGES ........................................................... 4-7

4.4 KRUEGER FLAPS .......................................................................... 4-7

5 LIFT DUMP AND SPEED BRAKES ............................................. 5-1

5.1 LIFT DUMPERS ............................................................................. 5-1

5.2 SPEED BRAKES ............................................................................ 5-2

6 SYSTEM OPERATION ................................................................. 6-1

6.1 MANUAL OPERATION .................................................................... 6-1

6.2 POWERED FLIGHT CONTROLS (P.F.C.US) ..................................... 6-1 6.2.1 Proportionality ............................................................... 6-2


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6.2.2 Redundancy of hydraulic Supplies ................................ 6-2 6.2.3 Tandem PFCU .............................................................. 6-2 6.2.4 Dual Assembly PFCUs ................................................. 6-4 6.2.5 Duplicate/Triplicate PFCU's ........................................... 6-5 Duplicated Control Surfaces ....................................................... 6-6 6.2.7 Self Contained PFCU .................................................... 6-7 6.2.8 Input Systems ............................................................... 6-7 6.2.9 High Speed Primary Controls ........................................ 6-8

6.3 TRAILING EDGE FLAP CONTROLS .................................................. 6-9 6.3.1 Flap Control Utilising Linear Hydraulic Actuators ........... 6-9 6.3.2 General ......................................................................... 6-10 6.3.3 Hydraulic Power ............................................................ 6-11 6.3.4 Control Input Circuit ...................................................... 6-11 6.3.5 System Operation ......................................................... 6-12 6.3.6 Safety Aspects .............................................................. 6-13 6.3.7 Position Indication ......................................................... 6-14 6.3.8 Flap System - Hydraulic Motors and Torque Tube Drive 6-14 6.3.9 Maintenance of Flap Systems ....................................... 6-16

6.4 LEADING EDGE FLAP CONTROLS ................................................... 6-17 6.4.1 Leading Edge Flap Pneumatic drive Unit ...................... 6-20 6.4.2 Krueger Flap Drive Components ................................... 6-24

6.5 SPEED BRAKE/GROUND SPOILER CONTROL ................................... 6-25 6.5.1 Operation ...................................................................... 6-26

6.6 MECHANICAL & ELECTRICAL FLIGHT CONTROL SYSTEM .................. 6-29 6.6.1 Mechanical Controls ...................................................... 6-29 6.6.2 Electrical Flight Controls ................................................ 6-30

7 Q FEEL, YAW DAMPER, MACH TRIM, RUDDER LIMITER, GUST LOCKS 7-1

7.1 ARTIFICIAL FEEL ........................................................................... 7-1 7.1.1 Q Feel System Principles ............................................ 7-1 7.1.2 Mechanical Q Feel System .......................................... 7-1 7.1.3 Operation ...................................................................... 7-1 7.1.4 Hydraulic Q Feel System ............................................. 7-3 7.1.5 Mach Number Correction .............................................. 7-3 7.1.6 Operation ...................................................................... 7-3

7.2 YAW DAMPING ............................................................................. 7-5 7.2.1 Yaw Control .................................................................. 7-5

7.3 MACH TRIM .................................................................................. 7-6 7.3.1 Typical System .............................................................. 7-8 7.3.2 Controller ...................................................................... 7-8 7.3.3 Mach Trim Actuator ....................................................... 7-8 7.3.4 Operation ...................................................................... 7-8

7.4 RUDDER LIMITING ......................................................................... 7-10 7.4.1 Q Limiter ...................................................................... 7-10

7.5 GUST LOCKS ................................................................................ 7-10 7.5.1 Description .................................................................... 7-10 7.5.2 Controls locking mechanism (aileron and elevator) ....... 7-11


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7.5.3 Controls locking mechanism (rudder) ............................ 7-12 7.5.4 Power Supplies ............................................................. 7-14 7.5.5 Operation ...................................................................... 7-14

8 RIGGING AND BALANCING CONTROLS .................................. 8-1

8.1 RIGGING - INTRODUCTION ............................................................. 8-1

8.2 CHECKS BEFORE RIGGING ............................................................ 8-1

8.3 RIGGING PROCEDURE ................................................................... 8-2 8.3.1 Establishing the Neutral Setting .................................... 8-2 8.3.2 Rigging Pins .................................................................. 8-2 8.3.3 Control Surface Setting Gauges .................................... 8-4 8.3.4 Checking for Sense of Movement ................................. 8-4 8.3.5 Checking for Static and Running Friction ...................... 8-6

8.4 CHECKS AFTER RIGGING .............................................................. 8-6 8.4.1 Duplicate Checks .......................................................... 8-7

8.5 PRIMARY CONTROL SYSTEMS - EXAMPLE OF RIGGING .................... 8-7

8.6 RIGGING A TUBE-OPERATED CONTROL SYSTEM ............................. 8-8

8.7 RIGGING A POWERED FLYING CONTROL SYSTEM ........................... 8-9

8.8 RIGGING OF TRIMMING TAB SYSTEM .............................................. 8-11

9 STALL WARNING AND PROTECTION ....................................... 9-1

9.1 STALL WARNING SYSTEMS ............................................................ 9-1 9.1.1 Pneumatic Stall Warning System .................................. 9-1 Electric Stall Warning System ..................................................... 9-2

9.2 STALL PROTECTION SYSTEM ......................................................... 9-3 9.2.1 System Functions.......................................................... 9-3 9.2.2 Typical System Components ......................................... 9-3

9.3 ACTUAL STALL PROTECTION SYSTEM ............................................ 9-4 9.3.1 Incidence Probes .......................................................... 9-4 9.3.2 Nitrogen System ........................................................... 9-5 9.3.3 Automatic Ignition.......................................................... 9-6 9.3.4 Stall Warning ................................................................. 9-6 9.3.5 Stall Identification .......................................................... 9-7

10 FLY BY WIRE ............................................................................... 10-1

10.1 INTRODUCTION ............................................................................. 10-1

10.2 PRINCIPLES OF FBW .................................................................... 10-1

10.3 PRINCIPLES OF FBOW ................................................................. 10-1 10.3.1 Advantages of FBOW over FBW ................................... 10-1

10.4 OTHER INPUTS TO POWERED FLYING CONTROL UNIT ..................... 10-2

10.5 777 FLIGHT CONTROLS - INTRODUCTION........................................ 10-2 10.5.1 General ......................................................................... 10-2 10.5.2 777 Primary Flight Control System ................................ 10-2 10.5.3 High Lift Control System ................................................ 10-3 10.5.4 Benefits of the Fly-By-Wire System ............................... 10-3 10.5.5 Abbreviations and Acronyms ......................................... 10-3 10.5.6 Primary Flight Control System - Introduction ................. 10-5


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10.5.7 PFCS General Description ......................................... 10-6 10.5.8 Manual Operation.......................................................... 10-6 10.5.9 Autopilot Operation ....................................................... 10-7 10.5.10 PFCS Modes of Operation ............................................ 10-7 10.5.11 Flight Deck Controls ...................................................... 10-7 10.5.12 Main Equipment Centre................................................. 10-8 10.5.13 PFCS Flight Controls ARINC 629 BUS Interfaces ...... 10-8


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INTENTIONALLY BLANK


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1 PRIMARY FLIGHT CONTROLS

Aircraft theory of flight has already been discussed in Module 11.1. We shall now look at how the Aircraft are equipped with moveable aerofoil surfaces that provide control in flight. Controls are normally divided into Primary and Secondary controls. The primary flight controls are:

Ailerons

Elevators

Rudders

Spoilers

Because of the need of aircraft to operate over extremely wide speed ranges and weights, it is necessary to have other secondary or auxiliary controls. These consist of:

Trim controls

High Lift Devices

Speed Brakes and Lift Dump


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Note: There is some variation of opinion as to whether spoilers are considered to be primary controls. The JAR 66 syllabus includes them as primary controls, so that is how these notes will define them. Both types of controls are illustrated in the following diagram.

Typical Aircraft Flight Controls

Figure 1


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1.1 AILERONS

Ailerons are primary flight controls that provide lateral roll control of the aircraft. They control aircraft movement about the longitudinal axis. Ailerons are normally mounted on the trailing edge of the wing near to the wing tip.

Inboard and Outboard Ailerons

Figure 2

Some large turbine aircraft employ two sets of ailerons. One set are in the conventional position near the wing tip, the other set are in the mid-wing position or outboard of the flaps. At low speeds both sets of ailerons operate to give maximum control. At higher speeds hydraulic isolate valves will cut power to the outer ailerons so that only the inboard ailerons operate. If the outer ailerons are operated at high speeds, the stress on the wing tips may twist the leading edge of the wing downwards and produce aileron reversal.

1.2 ELEVATORS

Elevators are primary flight controls that control the movement of the aircraft about the lateral axis (pitch). Elevators are normally attached to hinges on the rear spar of the horizontal stabilizer. Fig 11.1 shows the typical location for elevators.


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1.3 RUDDERS

The rudder is the flight control surface that controls aircraft movement about the vertical or normal axis. Rudders for small aircraft are normally single structural units operated by a single control system. Rudders for larger transport aircraft vary in basic structural and operational design. They may comprise two or more operational segments, each controlled by different operating systems to provide a level of redundancy.

Rudder Figure 3


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1.4 SPOILERS

Spoilers are secondary control surfaces used to reduce or spoil the lift on a wing. They normally consist of multiple flat panels located on the upper surface of the wings. The diagram below shows the more common configuration.

Operation of Spoilers on a Typical Aircraft Figure 4

The spoilers lay flush with the upper surface of the wing and are hinged at the forward edge. When the spoilers are operated, the surface raises and reduces the lift. The spoilers may be used for different purposes.


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1.4.1 FLIGHT SPOILERS

Flight Spoilers are used in flight to reduce the amount of lift. If the pilot operates the controls left or right to roll the aircraft, the spoilers on the down-going wing move upward to aid rolling the aircraft. The movement of the spoilers is in proportion to the rate of roll required. On some aircraft, the spoilers are the primary flight control for rolling. If operating only as flight spoilers, only the surfaces on one wing will be raised at any one time. The flight spoilers are normally positioned outboard of the ground spoilers.

1.4.2 GROUND SPOILERS

Ground Spoilers are only used when the aircraft is on the ground. They operate with the flight spoilers to greatly reduce the lift on landing. The also reduce the drag after landing to slow down the aircraft. Ground spoilers will normally be deflected to their maximum position to give maximum drag on landing.


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2 TRIM CONTROLS

The majority of aircraft at some time during a flight, develop a tendency to deviate from a straight and level attitude. This may be caused by a fuel state change, a speed change, a change in position of the aircraft's load, or in flap and undercarriage positions. The pilot can counter this tendency by continuously applying a correcting force to the controls - an operation, which, if maintained for any length of time, would be both fatiguing and difficult to maintain. The tendency to deviate is therefore corrected by making minor trim adjustments to the control surfaces. Once an aircraft has been trimmed back to a 'balanced' flight condition, no further effort is required by the pilot until further deviation develops.

2.1 FIXED AND ADJUSTABLE TRIM TABS

2.1.1 FIXED TRIM TABS

A fixed trim tab is normally a piece of sheet metal attached to the trailing edge of a control surface. It is adjusted on the ground by bending to an appropriate position that give zero control forces when in the cruise. Finding the correct position is by trial and error.

2.1.2 CONTROLLABLE TRIM TABS

Controllable Trim Tab

Figure 5

A controllable trim tab is adjusted by mechanical means from the flight deck, usually with an indication of its position being displayed to the pilot. Most aircraft have trim on the pitch control and more advanced aircraft have trim on all three axes. Whilst the controls in the cockpit are by lever, switch etc., the actuation can be by mechanical, electrical or hydraulic means.


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2.1.3 SERVO TABS

Servo Tab Figure 6

Sometimes referred to as the flight tabs, the servo tabs are used primarily on large control surfaces, often found on larger, older aircraft. This tab is operated directly by the primary controls of the aircraft. In response to the pilot's input, only the tab moves. The force of the airflow on the servo tab then moves the primary control surface. This tab is used to reduce the effort required to move the controls on a large aircraft.

2.1.4 BALANCE TABS

Balance Tab

Figure 7

A balance tab is linked to the aircraft in such a manner that a movement of the main control surface will give an opposite movement to the tab. Thus the balance tab will help in moving the main surface, therefore reducing the effort required. This type of tab will normally be found fitted to aircraft where the controls are found to be rather heavy during initial flight-testing.


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2.1.5 ANTI-BALANCE TABS

These anti-balance tabs operate in the same way, mechanically, as balance tabs. The tab itself is connected to the operating mechanism so that it operate in the reverse way to the balance tab. The effect this has is to add a loading to the pilots pitch control, making it appear heavier. These tabs can often be found fitted to stabilators, which are very powerful and need extra feel to prevent the pilot over-stressing the airframe.

2.1.6 SPRING TABS

The spring tabs, like some servo tabs, are usually found on large aircraft that require considerable force to move a control surface. The purpose of the spring tab is to provide a boost, thereby aiding the movement of a control surface. Although similar to servo tabs, spring tabs are progressive in their operation so that there is little assistance at slow speeds but much assistance at high speeds.

Spring Tab

Figure 8


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2.2 FULLY POWERED FLYING CONTROL TRIM SYSTEM

As fully powered flying controls are irreversible, i.e. all loads (reactions) are fed via mountings to structure, trim tabs would be ineffective.

To overcome this, electric trim struts or actuators are used within the input system. These actuators commonly reposition the "null" position of a self centring spring device to hold the control input system in a new neutral position. Thus the main control surface will be held deflected and the aircraft trimmed.

2.2.1 TYPICAL TRIM SYSTEM

The following is a typical trim system as used on a fully powered flight control system.

2.2.2 RUDDER TRIM SYSTEM

In a typical rudder trim system for a powered system, trim commands from the trim switch causes an actuator to extend or retract, which rotates the feel and centring mechanism. This provides a new zero force pedal position corresponding to the trimmed rudder position. The trim switch is spring loaded to return to neutral. Both positive and negative elements of the circuit are switched to prevent a trim runaway should one set of switch contacts become short circuited. The trim indicator is driven electrically by a transmitter in the rudder trim actuator. The indicator shows up to 17 units of left or right trim. Each unit represents approximately one degree of rudder trim.

2.2.3 AILERON TRIM SYSTEM

In a typical aileron trim system for a powered system, trim commands from the trim switches causes the actuator to extend or retract, which repositions the feel and centring mechanism null detent. The trim switches must be operated simultaneously to provide an electrical input to the actuator, as both positive and negative elements of the circuit are switched to prevent a trim runaway should one set of switch contacts become short circuited. The available aileron trim provides 15 degrees aileron travel in both directions from neutral.


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2.2.4 TAILPLANE TRIM SYSTEM

For trimming the aircraft longitudinally (about the lateral axis) the elevators are not trimmed. Instead the angle of incidence of the whole tailplane is altered. Raising the leading edge of the tailplane will increase lift over the tailplane which imparts a nose-down attitude to the aircraft or vice versa.

This is done by mounting the forward end of the tailplane on a screw-jack. Depending on the system the screw-jack is rotated by two hydraulic or electric motors via a gearbox. Movement is induced by a lever in the flight deck which operates solenoid selector valves or an electric control circuit to operate the motors. Over-travel is prevented by micro-switch.

Reasons for fitting to transport aircraft:

1. All aircraft benefit from having as large a range of useable centre of gravity as possible. This gives flexibility in cargo loading and allows for fuel usage in a swept wing.

2. Aircraft benefit from a wide speed range. Very simply, when an aircraft is trimmed at a particular speed, a reduction in speed calls for "up" elevator and an increase in speed calls for "down" elevator. This would cause extra drag.

3. The need to compensate for centre of pressure changes due to slat/flap extension, gear extension.

4. To reduce trim drag to a minimum to give the optimum performance in cruise.


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Variable Incidence Tailplane Trim System

The tail-plane is pivoted at the rear of the centre section torsion box and attached to an actuator forward of the centre section. Operation of the actuator raises or lowers the leading edge of the tail-plane, altering the incidence angle.

Variable Incidence Tailplane Figure 9


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The actuator, comprises a re-circulating ball screw jack and nut assembly driven by two hydraulic motors with separate spur gear reduction trains.

Figure 10


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Friction brakes ensure that air loads cannot back-drive the actuator when the system is de-pressurised.

The actuator is signalled from one of three sources:

i) Auto-pilot servo

ii) Mach trim servo

iii) Trim hand-wheel operation.

A cable loop runs from the pedestal in the cockpit, under the cabin floor, and ends at a cable reduction gearing unit at the tailplane incidence actuator.

Hydraulic Power Supply

Each hydraulic motor is powered from a separate system. In the event of a single hydraulic system failure, a bypass valve permits that motor to "freewheel" when the system is de-pressurised.

Position Indication Systems

Geared indicator scales inboard of the cockpit hand-wheels present the demanded position of the tail-plane. This will be the actual tail-plane incidence with the hydraulic system(s) pressurized.

Actual tail-plane position is continuously displayed on the pilot's instrument panel, signalled by a position transmitter operated by the tail-plane.

External markings on the structure adjacent to the tail-plane give the approximate position of the tail-plane.

Tail-plane in Motion Warning

Some aircraft types have a tail-plane in motion warning system to alert the pilots of continuous motion of the tail-plane beyond a certain time period.


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3 ACTIVE LOAD CONTROLS

3.1 ACTIVE LOAD CONTROL

This system is a relatively new approach to civil aviation, although it has been in use for some time in military aircraft. It is a complex system that senses disturbances in the air that may cause both discomfort to passengers and crew, whilst causing extra unnecessary loading on the airframe.

The gusts that are about to hit the aircraft are sensed either by a tiny pair of vanes on either side of the nose or by accelerometers mounted inside the nose of the aircraft. These instantly send a signal, 'bump coming', to the flight control computers, which instantly send a correcting signal to the elevators that counter the bump and give a smoother ride.

The whole system requires the quick reactions of both the computers and the hydraulic jacks to be successful. If the aircraft senses a downdraft, the computers instantly signal just the correct amount of 'up elevator' to counteract the disturbance and leave the aircraft to fly smoothly on.

3.2 ACTIVE CONTROL TECHNOLOGY

Active Control Technology (ACT) can be defined as the use of a multivariable automatic flight control system to improve the manoeuvrability, dynamic flight characteristics and the structural dynamic properties of an aircraft by simultaneously driving an appropriate number of control surfaces and auxiliary force or moment generators in such a fashion that either the loads which the aircraft would have experienced as a result of motion without an ACT system are much reduced or the aircraft produces a degree of manoeuvrability beyond the capability of a conventional aircraft.

In essence ACT is the use of technology to make an aircraft and its control surfaces operate in an unconventional manner to effect high manoeuvrability or to reduce airframe stress.


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ACT is nothing new, it has been used on aircraft for many decades but it has increased in usage with the advent of flight control computers and fly-by-wire systems. The Tristar aircraft has a system installed that reduces the flight loads on the wings by partially deploying the spoilers. This changes the lift profile over the wing, bringing the lift closer to the wing root, which is much stronger (see next fig). This means that the wing can be lighter and the wing stresses will be reduced.

Figure 11

Numerous control surfaces, auxiliary force and moment generators can be added to make the aircraft operate unconventionally. Fighter aircraft and some executive jets may have a number of such devices fitted to make them more agile. These include:

Foreplanes which can only move together to give pitch control.

Canards, these differ from foreplanes as they can also move independently giving more response in roll.

Flaperons which are control surfaces that act as flaps and/or ailerons depending on the pilots selection. They have the ability to move both up and down independently for roll control, but can also move simultaneously for take off and landing.

Thrust vectoring, mainly used on combat aircraft, but the advantages gained with short take off and landing will mean that some form of vectoring system will be developed for commercial aircraft in the future.


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3.2.1 ADVANTAGES OF ACTIVE CONTROL TECHNOLOGY

The employment of Active Control Technology presents numerous advantages both for civil and military aircraft, namely:

The aircraft is more stable in flight

The aircraft are highly agile (military only)

A more comfortable flight for passengers

Reduced fatigue on the aircraft, therefore lighter construction can be utilised

Lighter construction gives better fuel consumption

Varying lift profiles means wings can be more streamlined (less drag)

It is impossible for the aircraft to be flown beyond its design limitations under normal conditions !

Conventional aircraft have four forces providing control and movement

Rolling moment

Pitching moment

Yawing moment

Thrust (Drag modulation)

The use of ACT can provide two more additional forces of control and movement:

Direct lift force

Direct side force

3.2.2 DIRECT LIFT FORCE

In order to change altitude a pilot must pitch the nose of the aircraft up, which may cause him to lose sight of his destination (the runway). Using ACT, the pilot can change altitude by causing the foreplanes and flaperons to operate together increasing the lift on the front and rear of the aircraft simultaneously. This is known as the direct lift force

3.2.3 DIRECT SIDE FORCE

The pilot, conventionally, must roll the aircraft to change its flight path in a sideways plane. ACT allows the aircraft to side step during normal flight by deploying the rudder and the canards together to pull the nose and tail of the aircraft across in the same direction. This is known as the direct side force.


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4 HIGH LIFT DEVICES

4.1 FLAPS

These devices have two primary aims, to provide extra lift during take-off and to provide greater lift as well as high drag during landing. The types of flap used on different aircraft depends on the type of aircraft, the method of aircraft operation and other variables. For example, a single engined light aircraft might only have some form of simple trailing edge flap, whilst a large airliner like the Boeing 777 has complex, triple slotted flaps.

Types of Flap System Figure 12

Flaps are fitted to most aircraft and are usually one of the types shown, together with the maximum increases of lift over the 'clean' configuration. As the complexity increases to improve performance, there is a proportional increase in weight, maintenance and cost.


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Whilst the term 'flaps' is used, it is taken as meaning trailing edge flaps, and the term 'leading edge flaps 'refers to those fitted to the leading edges of the wings of most large aircraft.

The methods of operation of flaps, are numerous. They can vary from simple, mechanical push rods or cables actuated, via a lever in the cockpit, by the pilot, to complicated, multiple flaps that are electrically selected on the flight deck and hydraulically or electrically powered.

Most flap systems have a number of positions, which can be selected at various times. As an example, five positions could be as follows;

00 - flaps up 250 - landing, first position

80 - take-off, first position 400 - landing, second position

150 - take-off, second position

These would all be selected by movement of a lever in the cockpit, which will have 'detents' at the various positions. This movement will, as can be seen in the illustration, be transferred to the control valve and on to the motor, which moves the actuators.

Flap Mechanism Figure 13


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Other high lift devices can be found on the leading edges of the wings and include slats, drooped leading edges and Krueger flaps. All of these devices are aimed at smoothing the airflow over the leading edges of the wings when they are at a high angle of attack, thereby maintaining, or increasing lift when the wing would normally be stalled.

4.2 SLATS

Slats are separate small aerofoils, which can be fixed or retractable. Their purpose is to control the air passing over the top of the wing at slow speeds. On larger aircraft, the retractable slats, have their extension interconnected with the trailing edge flaps.

This can be seen in the illustration, which not only shows the operation of the slats through three different positions, 'stowed', 'active' and 'open', but their association with the four positions of the trailing edge flaps.

Fixed slats are usually found on light aircraft, where the complications of weight, cost etc, can be balanced by the limitation of slightly higher drag than a 'clean' wing.

Leading and Trailing Edge Flap Settings

Figure 14


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4.3 DROOPED LEADING EDGES

Drooped leading edges are a different design, but are aiming at the same effect, that of smoothing the air over the top of the wing. They operate in much the same way as most high lift devices, by screw jack operation with the motive power for the jacks coming from the hydraulic system.

4.4 KRUEGER FLAPS

Krueger flaps are, again, a different design for the same effect. These are usually found fitted to the leading edges of the wing at the inboard sections where the effect of 'slats' or 'drooped leading edges' are not as efficient.

Figure 15

Krueger (left) and Drooped (right) Leading Edge Flaps


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5 LIFT DUMP AND SPEED BRAKES

5.1 LIFT DUMPERS

These devices are used to spoil lift from the wing after touchdown. This ensures that the aircraft's weight is fully on its landing gear, which enables the brakes to work at 100% for the full landing run. If this did not happen, the aircraft would tend to 'float' or bounce at touchdown, making the brakes inefficient and the risk of skidding much greater.

Lift dumpers are nearly always flat, rectangular panels, hinged at their leading edge and powered by hydraulics. They can usually be found on the top of the wing, and located about the maximum thickness, where their deployment would destroy the maximum lift from the wing.

To ensure that they deploy at the correct time and also without the need for the pilot to select them, at a very busy time, there is a simple system to deploy them automatically. A set of switches are fitted to the landing gear which 'make' and indicate weight-on-wheels to several systems, once the aircraft is completely on the ground. By giving the pilot a "lift dumper arming" button, he can arm the system, in flight, and know that it will deploy the lift dumpers at the correct time.


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5.2 SPEED BRAKES

The use of speed brakes is similar regardless of the aircraft type. If the aircraft is a sailplane it is so streamlined that it requires high drag when descending and landing in unprepared fields. A large 400 seat airliner needs to be able to follow Air Traffic Control instructions to descend and maintain certain speeds and a military jet fighter needs to have very high drag on approach, permitting the engines to accelerate quickly if the landing is aborted.

All types of speed brake use a variation of the same principle, to put panels of varying shapes into the airflow, to increase the drag. Some are able to modulate, (vary the amount of drag to suit the situation), whilst others are just 'IN' or 'OUT'. Some airliners use the same surfaces on the top of the wing to carry out more than one operation, such as speed brakes when in flight and needing drag; roll control to augment (or replace) ailerons; or as lift dumpers to be used after landing.

Light aircraft rarely need speed brakes because of their generally high drag designs. A reduction in power will produce a satisfactory slowing down of the aircraft. Streamlined sailplanes, however, usually have vertical panels that project from the wing, top and bottom, which produce large amounts of drag, enabling steep, slow and safe approaches when landing.

Military jets have a different need for drag, not only as mentioned during the approach to landing, but during combat and other operations where fast application of drag with a quick reduction in speed can have a life saving effect.

Speed Brake Installation Figure 16


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6 SYSTEM OPERATION

6.1 MANUAL OPERATION

6.2 POWERED FLIGHT CONTROLS (P.F.C.US)

In large modern aircraft that fly at high speeds, the air loads on the f'lying control surfaces far exceed the ability of the pilot to move them manually. To overcome this problem hydraulic pressure is used to move the control surfaces, a POWERED FLYING CONTROL UNIT or BOOSTER being used to convert hydraulic pressure into a force exerted on the control surface.

In its simplest form, a P.F.C.U. consists of a hydraulic jack, the body of which is fixed to the aircraft structure and the ram, via a linkage to the control surface.

To control the P.F.C.U. a servo valve (control valve) is mounted on the jack. The servo valve, which is connected to the pilot's controls by a system of cables and/or pushrods, called the input system, directs fluid to either side of the jack piston and directs the fluid from the other side to return. This flow of fluid will displace the jack ram and as this is connected to the control surface via an output system of pushrods or cables, the control surface is moved.

Figure. 17


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6.2.1 PROPORTIONALITY

To make the controls "proportional" (i.e. the degree of movement of the jack-ram and hence the control surface, should be proportional to the degree of movement of the pilot's controls), a "follow-up linkage" is used. This linkage connects the input system, through a series of levers to the output system in such a way that the movement of the output system (jack ram) tends to cancel the input once the desired position is reached and so output movement ceases. In effect the movement of the jack ram is always trying to re-centre the servo valve and stop fluid flow in the jack.

6.2.2 REDUNDANCY OF HYDRAULIC SUPPLIES

Hydraulically powered flight control units usually derive their hydraulic power from the aircraft hydraulic system. If a PFCU obtained hydraulic power from only one hydraulic supply, a failure of that hydraulic supply due to an engine shut down, loss of fluid due to a leak, or failure of a hydraulic pump. The result would be loss of powered control of the aircraft. The probability of hydraulic failure is too great to allow a system to rely on one hydraulic supply, so redundancy must be introduced into the flight control system.

As in the previous notes on hydraulic systems, modern large multi-engine aircraft, are arranged such that the engine driven pumps (and the other types of pumps) supply two or more independent hydraulic power supply systems.

The following are methods that use that arrangement of hydraulic redundancy to allow failure of one hydraulic supply and still maintain control of the aircraft.

6.2.3 TANDEM PFCU

These are similar to the arrangement shown. They consist of a single jack ram but with two pistons. These pistons are housed in two co-axial cylinders each of which receives pressure fluid from separate power supply circuits via their own duplicated servo valves. The servo valves, which are controlled by the same input system, are carefully set up in the overhaul workshop to ensure they work in unison. This prevents the two hydraulic pistons working against each other. With this arrangement a loss of one hydraulic supply will allow the relevant piston to "free stroke whilst the other piston operates the control surface.


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TANDEM ACTUATOR

Figure 18


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6.2.4 DUAL ASSEMBLY PFCUS

These are similar to the tandem arrangement but two piston rams are located in cylinders mounted side by side with the piston rams connected to a common output lever that transmits the movement to the control surface. The arrangement for the input system, the duplicated servo valves and hydraulic fluid supply are the same.

Dual Assembly PFCU FIGURE 19


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6.2.5 DUPLICATE/TRIPLICATE PFCU'S

In this arrangement each control surface is operated by two or three separate PFCU'S. For hydraulic redundancy, each PFCU is powered from separate hydraulic supply circuits. If one supply system should fail, or if one PFCU should malfunction the effected PFCU can be switched off. In this event a bypass valve within the PFCU will open interconnecting both sides of the jack ram. Therefore, as the pilot moves the input and operates the serviceable PFCU'S, the control surface will move and, "drag" the unserviceable PFCU ram with it. The open bypass valve will allow fluid to transfer from one side of the ram to the other as the PFCU "free strokes". Thus control will be maintained by the serviceable PFCU's driving the control surface, and a hydraulic lock in the unserviceable PFCU is prevented.

In this arrangement each control surface (rudder is shown in the diagram) is split into two or three independent sections. Each section is operated by its own PFCU. For hydraulic redundancy, each PFCU is powered from separate hydraulic supply circuits. If one supply system should fail, or if one PFCU should malfunction the effected PFCU can be switched of. In this event the PFCU and its control surface segment will be "blown back" to the neutral position by aerodynamic loads and held by a lock. Thus control will be maintained by the serviceable PFCU's driving their respective segments of control surface.

All PFCU's are controlled via a single input system to a common input lever connected to all PFCU servo valves. Therefore if one PFCU malfunctioned it could prevent the operation of the remaining serviceable PFCU'S. To prevent this the input to the servo valves from the common input lever is via compressible spring struts or spring boxes. In normal operation these spring struts/boxes resist compression and allow full control of all PFCU'S. If a PFCU is unserviceable, pilots input will compress the spring strut to that PFCU but the remaining spring struts/boxes will resist compression and operate the PFCU servo valves normally.


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6.2.6 DUPLICATED CONTROL SURFACES

Duplicated Control Surfaces Figure 20


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6.2.7 SELF CONTAINED PFCU

A self contained PFCU consists of a jack-ram powered by its own dedicated integrally mounted hydraulic generator" and hydraulic reservoir. The generator is a radial piston pump arrangement within a slip ring assembly. The slip ring position is control ' led by a servo valve piston arrangement. With the slip ring held concentric with the piston bank no movement of the pistons within the rotating piston bank is allowed and no fluid flow will result. If an input moves the slip ring the rotating bank of pistons will be allowed to "stroke" and a flow to the PFCU piston will occur and the PFCU ram will move. Movement of the slip ring in the opposite direction will cause fluid flow to the other side of the piston and the ram will move in the other direction. The piston bank is rotated by a drive from a 3 phase electric motor which derives its supply from the aircraft electrical system.

To maintain redundancy this type of PFCU will be duplicated and each may drive a duplicate and independent (split) control surface as above. As its source of power is electrical, it is independent of the aircrafts hydraulic system, therefore even with total hydraulic failure, control can still be maintained. On malfunction of a PFCU, or loss of electric power to that PFCU, it will lose hydraulic pressure and "blow back" to a neutral position where an integral lock will hold it. In this event further inputs to the servo valves are absorbed by spring-strut that allows unhindered operation of the remaining PFCU'S.

To give redundancy of electrical power supply, each PFCU in a "set" (i.e. rudder) gets its power supply from a different bus bar.

6.2.8 INPUT SYSTEMS

Generally the input system of the powered flying control system is mainly a cable system with the related quadrants, pulleys and fairleads with the connections to the control column and the PFCU input lever by push rods. To guard against loss of control due to cable breaks the cable system is duplicated. All duplicated runs are routed separately through the aircraft to avoid one incident damaging both control runs. The cable systems meet at a common input lever to the PFCU'S.


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Input Systems Figure 21

6.2.9 HIGH SPEED PRIMARY CONTROLS

Primary controls are designed to give adequate control in all flight phases. The flight phase at which the control surfaces are least effective is during low speeds (landing). This is because of the reduced aerodynamic effect with low speed. This means that the size and range of movement of each control surface must be sufficient to maintain sufficient control authority. With the control system designed to give efficient control at low speed, there may be a problem at high speed. This is that at high speeds the increased air-loads on the control surfaces will cause them to be too sensitive producing over control and possible loss of control or over-stressing of the airframe. To prevent this two systems may possibly be used.


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Geared Controls

In this system a single acting hydraulic jack may be fitted to an idler lever. The control rod is attached to this jack so that the radius of operation can be altered. Thus for a given angular movement of the idler lever, if the length of the jack is shortened, the linear movement of the control rod is reduced. This will maintain a constant range of movement at the pilots controls but reduce the range of movement of the control surface. Pressure at the jack is usually controlled by a pressure-modulating valve sensitive to a pressure transducer in the pitot system.

High Speed Control Surfaces (ailerons)

Normal, "low speed" ailerons are situated at the usual wing tips position to gain maximum authority due to the moment arm produced. But again at high speed their authority may be too great. In this system an additional set of "high speed" ailerons is also fitted at the wing root. Hydraulic isolate valves are incorporated in the control system such that at low speed the outer ailerons are functional, but at high speed, their hydraulic power is cut off and the high speed ailerons are powered to maintain roll control. The isolate valves are again controlled by pressure switches in the pitot system.

6.3 TRAILING EDGE FLAP CONTROLS

On small aircraft the flaps are operated using hydraulic jacks to operate a single flap on each mainplane. This arrangement is not suitable for use on larger aircraft due to the size of the airframe that requires that the flaps are manufactured and mounted in "segments" along the trailing edge.

6.3.1 FLAP CONTROL UTILISING LINEAR HYDRAULIC ACTUATORS

The following system that may be regarded as a simple system, similarly uses linear hydraulic actuators for an aircraft that has three flap segments on each mainplane each positioned by a separate hydraulic actuator.

Movement of each actuator is controlled by a servo valve (simiIar to that in a primary flight control unit). Control is by flap lever/quadrant on the centre console. This is connected to the actuator servo valves by a duplicated system of control cables and pushrods.


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6.3.2 GENERAL

The flap surfaces are operated through linkages by hydraulic actuators. The actuators respond simultaneously to the control-cable-relayed demands of a selector lever mounted on the flight compartment centre console.

The piston rod end of each actuator is structurally anchored; movement being confined to the unit body. A position control element (servo-valve) incorporated in the body is controlled by an attached operating lever that has limited travel on each side of the neutral position. The lever is moved towards or away from the anchored piston rod end to retract or extend the actuator. Each actuator incorporates internal restrictors that control the rate of response and an internal mechanical lock that engages when the flaps are fully up. The lock is hydraulically released when a down selection is made.

The control system consists of a duplicated input circuit, which through the medium of a spring strut, signals all six actuators. Beyond the spring strut the signal to the inner flap actuators is conveyed by a rod and lever system and to the mid and outer flap actuators by interconnected signalling cables.

The purpose of the spring strut is to "store" control lever movement due to the actuators' restricted rate of travel.

The adjacent ends of the mid and outer flap surfaces are connected by a link that allows sufficient free movement to accommodate normal variations of relative positions without the links being loaded. The links are incorporated as a safety feature and take effect to prevent an asymmetric flap condition.

The flap selector lever is afforded the following gated positions - 0, 5, 15 and 30.


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6.3.3 HYDRAULIC POWER

Hydraulic power for operation of the actuators is provided by main system pressure backed up by flap accumulator pressure, when the flight compartment selector lever is at any position other than fully up (0). The accumulator stored pressure is released to the flap system when a solenoid valve is energised open via a micro-switch operated by the selector lever. The 'back-up' pressure is introduced downstream of a non return valve in the main system pressure line; thus maintenance of a selected down position is assumed, for a limited period in the event of a main system failure.

Flap System Hydraulics Figure 22

6.3.4 CONTROL INPUT CIRCUIT

From the flap selector lever on the centre console, the duplicated input cables are routed aft through the roof structure to a position immediately aft of the rear spar. At this point, the cables are directed through the roof skin terminating with a double quadrant assembly. A double acting spring strut is connected between an output lever on the quadrant and a series of levers and control rods. These:

Operate the position control elements (servo valves) on the inner flap actuators and transmit actuator movement to the inner flap surfaces.

Provide an input to the left and right mid and out flap signalling circuits.


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The spring strut is incorporated to allow a selection to be made in one quick movement - the total input motion being absorbed by the spring strut and progressively released as all six actuators respond at their controlled rate of travel.

Each of the left and right mid and outer flap signalling circuits consists of a pulley drum from which cables are routed outboard to quadrant assemblies at the mid and outer flap positions. Output levers on these quadrants are linked by control rods to the position control element (servo-valve) operating levers in the appropriate actuator package assemblies.

The left and right pulley drums are interconnected by two tie-rods to ensure symmetrical operation of the left and right wing flaps.

Flap Control Input Circuit Figure 23

6.3.5 SYSTEM OPERATION

Immediately a selection is made the total input motion is absorbed by the spring strut and progressively released as all six actuators respond at their limited rate of travel. When the spring strut returns to its pre-selection settled length - the rod which connects to the position control element operating lever on each actuator arrests. The actuators will then marginally run on until their now restrained element operating levers reach neutral positions. This simultaneously creates a hydraulic lock at all six actuators and hence arrests the surfaces in alignment at the selected position.


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6.3.6 SAFETY ASPECTS

Two main safety requirements must be met.

1. One is that a control cable break will not mean loss of control of the flaps. System integrity is such that duplication of the input cables which allows for functioning in the event of loss of either circuit) will maintain control.

2. The other is that an asymmetric deployment of the flaps is prevented. An asymmetric condition could happen in several ways and the following mechanisms are designed to prevent these.

A. Controls jamming between an actuator and surface (input systems intact):-

Should this occur during a programmed selection, the input system of the relevant actuator will arrest and in consequence will stop signalling of the remaining actuators which will then run on marginally until their now restrained servo valve operating levers reach neutral positions - thus arresting all six surfaces in approximate alignment.

B. Mechanical failure between an actuator and surface (which will not impede surface movement):-

Should this occur at either of the inner flaps - the system will remain functional (full asymmetry between inner flaps can be adequately countered by aileron action).

Should this occur at a mid or outer flap - the link which interconnects the adjacent ends of these surfaces will take effect to allow full functioning of both surfaces from one actuator. Thus preventing an asymmetric condition that would be beyond the ailerons ability to counter.

C. Loss of signalling (cable break) to a mid or outer flap actuator.

Should loss of signalling to a mid or outer flap actuator occur and the 'free' actuator become hydraulically locked at any stage during a programmed operation - the interconnecting link will arrest the adjacent functional actuator and thus its intact signalling system. This will have the effect of simultaneously arresting the interconnected input circuits of the remaining actuators that then run on marginally, until their now restrained servo valve operating levers reach neutral positions - thus arresting all six surfaces in approximate alignment. The actuator arrested by the link will remain programmed to achieve intended travel in opposition to the locked adjacent surface. For this reason and to prevent excessive structural overloading - the actuators incorporate internal relief valves.

D. Loss of main system pressure


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Main system pressure is augmented by flap accumulator stored pressure via a solenoid valve when the FLAPS selector lever on the flight compartment centre console is at any position other than fully up (0'). The 'back up' pressure is introduced down-stream of a non return valve in the main system pressure line; thus maintenance of a selected down position is assured for a limited period in the event of a main system failure.

6.3.7 POSITION INDICATION

Flap position is indicated on a twin pointer scale calibrated to 0', 5', 15' and 30' settings. The flap position is signalled by two transmitters that are driven from the flap hinge arms via control rods.

6.3.8 FLAP SYSTEM - HYDRAULIC MOTORS AND TORQUE TUBE DRIVE

On large aircraft it is more common for the flaps to be driven by twin hydraulic motors, each motor deriving its hydraulic supply from a different hydraulic system.

Each motor is mounted on the same gearbox, such that drive from either or both motors will drive the gearbox.

The gearbox is commonly located in the main gear bay. The drive is transmitted to the flap surfaces by a system of torque tubes, gearboxes and screw-jacks. The screw-jacks drive trolley assemblies along flap tracks mounted to the wing structure via support units. The flap segments are mounted onto the trolleys.

System Description

The flap system of each side of the aircraft comprises of flap sections supported and moved by six support/operating units. (Flap Tracks) The flaps are manually controlled by a lever on the central console to UP (0), take off (20), approach (35) and landing (45) positions. This manual control operates independent electro/hydraulic systems A and B, employed simultaneously to power the drive unit (gearbox) and their supplies are drawn from the aircraft electric and hydraulic systems bearing the same suffix letter. Both systems normally operate together, but should a hydraulic system fail, or a fault develop which necessitates selection of ISOLATE on one system, the flaps travel only at half rate due to the design of the drive unit.

a. Drive Unit The drive unit, comprises a gearbox and selector drum assembly, powered by two hydraulic motors. It rotates a torque shaft system that operates screw-jack and trolley mechanisms at each support/operating unit.

The drive unit is mounted to the rear of the wing rear spar member in the left main landing gear bay. It is powered by two hydraulic motor/lock valve assemblies; one supplied from hydraulic system A and the other from system B.


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The motors drive a main shaft through a differential gear and a spur wheel reduction gearing. A gear driven selector drum operates micro-switches to arrest the flaps when they reach the selected position.

b. Flap Transmission System

Torque shafts extend outboard in each wing from either end of the drive unit main shaft. The sections of torque shaft couple via universal joints and serrated sleeve joints to bevel gearboxes and to intermediate bevel gearboxes. The bevel gearboxes and intermediate gearboxes are connected by serrated sleeve joints and universal joints to screw shaft assemblies located at each support unit. Flap trolleys fitted to each screw shaft engage via their rollers with trolley tracks fitted to the support units. These trolleys support the flap sections.

The flaps are hinged by pins to lugs on the flap trolleys. A torque link pivoted to each flap section carries a forward flap trolley, the rollers of which engage with the cam track on the support unit.

c. Hydraulic System

For redundancy the flaps are supplied by two independent hydraulic systems, which are identical. The following therefore describes one system only.

Hydraulic pressure is supplied to the flap selector valve via a flow control valve and isolating valve.

Movement of the flap selector lever energises the appropriate solenoid selector valve to allow pressurised fluid to pass to the hydraulic motor through the lock valve. Return fluid from the hydraulic motor passes through the lock valve and flap selector valve back to the main system. The flow control valve controls the rate at which the flaps move. A throttle valve slows down the flaps at all selected positions.

When the flaps reach the selected position the selector valve solenoid is de-energised, through the operation of the selector drum micro switches. The pressurised fluid is held at the selector valve and the two service lines from the lock valve are connected together and into return. The lock valve prevents the hydraulic motor from rotating.

d. Flap Control

Each separate flap operating hydraulic circuit is controlled by a separate 28 volt D.C. electrical system. Each supply is derived from a separate D.C. Bus Bar.

Each system is controlled by three micro-switches operated by control lever movement, these provide a circuit to the selector valve solenoids via six micro-switches operated by the drive unit selector drum.

Cams on the outer periphery of the selector drum operate one switch at both the normal, up and down limit positions of the flaps and two switches at the take-off (20') and approach (35') positions.


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e. Over-Run Protection If the micro-switches in the drive unit selector drum should malfunction there is a probability that structural damage may occur as the flap trolleys reach the end of travel on their screw jacks. If a malfunction should happen, a set of over-run micro-switches mounted on the flap support units, will be operated to interrupt the supply to the selector valve solenoid and prevent the trolleys bottoming on their screw jacks. These micro-switches are part of the complete control circuit and are operated by strikers on the flap support trolleys.

f. Asymmetry Protection

If a malfunction should occur in the flap transmission system causing one part to seize, great damage could occur as the drive system attempted to drive the flaps to their selected position. To prevent this, weak "fail safe" joints are incorporated in various torque tubes that are designed to fail under a certain load.

However, this will allow the damaged portion of the system to stop and the remainder to continue travelling, so producing an asymmetric flap condition. To prevent this an asymmetry protection circuit is incorporated in the control system.

This system uses an A.C. electrical supply and is controlled by four synchro's which are small devices mounted on and driven by screw shafts in board and out board of the flap systems. These are paired, and as they rotate send an alternating signal to an asymmetry control box. If the signals become out of phase with each other the over-travel/ asymmetry isolate relay will be energised to lock-out the system.

g. Position Indication

Flap position indication is provided by a d.c ratio meter indicating system comprising two transmitters, driven from the outboard end of the left and right torque shaft systems and dual indicators positioned on the centre in the flight deck.

6.3.9 MAINTENANCE OF FLAP SYSTEMS

Because of the exposed position of most flap system components regular lubrication of hinge bolts, screw jacks, trolleys etc is required. When carrying out this task all excess grease must be removed to prevent the accumulation of dirt or grit that may enter bearings etc.

Rigging

The flap operating system is a large complex system which will only work if all parts are in their correct relative positions at all times. To ensure this, whenever the system is disturbed by a maintenance task it must be checked or re-rigged. Provision is built into the system for this.


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6.4 LEADING EDGE FLAP CONTROLS

The following notes describe a typical leading edge flap control system (Boeing 747). There are two modes of operation, primary and alternate. The normal method of operation is by use of the primary mode and is initiated by use of the flap lever to a selected detent.

On the Boeing 747 there are 28 leading edge flaps, they in turn are divided in two categories, variable camber (22) off, Kruger (6) off.

Four pneumatic power units in each wing move the flaps up or down, (extend) or (retract). Air is supplied to the power units from ducts in the leading edge of the wing. The ducts also supply air to jets that spray the outboard drive units with hot air for anti icing. Each drive unit assembly has two motors, one pneumatic and one electrically powered. Torque developed by the drive units is supplied to rotary actuators. The actuators move the flaps.

Normal operation is achieved by operation of the flap control lever. Three rotary variable differential transducers (RVDTs) sense movement and signal the Flap Control Unit (FCU), which control the direction control motors. If pneumatic power is not available the FCU will switch to electric drive motor operation.

Figure 24


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L/E Flap System Components Figure 25


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Figure 26


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6.4.1 LEADING EDGE FLAP PNEUMATIC DRIVE UNIT

Purpose

Eight pneumatic drive units (PDU) power the LE flap system. Each power drive unit has both a pneumatic and an electric drive motor. The pneumatic motor is the primary drive source and is powered by the leading edge pneumatic manifold. The electric motor is an additional drive source for use when the pneumatic system is not available.

Leading Edge Flap Drive Unit- operation

a. Pneumatic Drive

The flap lever is used to command FCU operation. The flap control unit signal is passed to the directional control motor and the shutoff valve. Pneumatic pressure flows from the inlet duct through the alternate valve (normally open) to the shutoff valve. The shutoff valve (normally closed) opens to pressurise the regulator and the air Motor brake. Pneumatic pressure at the regulator opens the butterfly valve and regulates the pressure to the control valve. Pneumatic pressure at the air motor brake releases the brake. The direction and speed difference between the direction control motor and the output shaft follow-up gear is sensed by the differential. The differential uses the speed differences to position the control valve and maintain PDU speed. Travel limits are governed by the primary position controller. This translates the amount of distance that the nut travels. When the translating nut reaches its travel limit it stops the direction control motor rotation that, in turn, stops PDU operation.

b. Electric Drive

The signal to activate the electric drive motor closes the alternate solenoid valve. The electric motor brake then releases the electric motor drive. The pneumatic brake holds the sun gear of the planetary gearbox at the air motor output shaft. The electric motor drives the output shaft through the ring gear of the planetary gear reduction. When the translating nut in the alternate position controller reaches the end of its travel it opens the electric motor limit switches. The alternate controller position switches control the electric motor shutdown in both primary and alternate control modes.


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Figure 27

Operating Times

The approximate leading edge flap extension or retraction times are:

Pneumatic operation: 9 seconds

Electric operation: 90 seconds


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Electric Drive. Motor Control

Primary Mode: in the primary mode the FCU controls the LE flap operation. If the pneumatic drive motor is not available the FCU will select the electric drive motor. The alternate controller provides signals to the FCUs for control, monitoring, and indication functions.

Alternate Mode: in alternate mode the electric drive motor is the only method of moving the flaps. The alternate arming switch arms the system. Flap operation is commanded by using the rotary alternate control switch located on P-2 in the flight deck.


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Figure 28


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6.4.2 KRUEGER FLAP DRIVE COMPONENTS

Purpose/Location

The Krueger flaps modify the configuration of the inboard portion of the wing leading edge to increase low speed lift. There are three Krueger type LE flaps installed on each wing inboard of the inboard engines (flaps 11 through 16).

Figure. 29


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6.5 SPEED BRAKE/GROUND SPOILER CONTROL

Spoilers will normally be controlled by the pilot through the normal roll controls or by the automatic flight control system (auto-pilot). They may also be operated automatically as part of an automatic landing system. On a typical aircraft (Boeing 757) the spoilers are electrically controlled and hydraulically powered.

Figure 30


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6.5.1 OPERATION

Rotary variable differential transducers (RVDTs) convert control wheel inputs into electrical signals. Spoiler control modules receive the signals and command the Power Control Actuators (PCAs) to raise the spoilers. Placing the speed-brake lever in the UP position will raise all flight spoilers.

Figure 31


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Three RVDTs are grouped together in a can-like unit, mounted on the bottom of each control unit assembly. The RVDTs convert aileron control wheel rotation into a signal voltage proportional to the control wheel movement. The Spoiler Control Modules mix RVDT inputs with other inputs according to a programmed logic. Six SCMs control the 12 spoiler surfaces. Power Control Actuators operate the spoilers. Each spoiler has one PCA, powered by one of three hydraulic systems. Each PCA consists of a hydraulic actuator, an electro-hydraulic servo valve (EHSV) and a Rotary variable differential transformer (RVDT). The PCA extends or retracts as commanded to raise or lower the spoiler. The RVDT sends a feedback signal to the SCM proportional to the amount of surface deflection.

Electro-hydraulic Servo valves controls the flow of hydraulic fluid in the PVA in response to the SCM commands. The command operates a jet pipe that supplies hydraulic fluid to the EHSV control bobbin. The EHSV is spring loaded to the retract position, so the spoiler panel will retract if there is no command signal.


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Spoiler Electro Hydraulic Servo Valve Figure 32


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6.6 MECHANICAL & ELECTRICAL FLIGHT CONTROL SYSTEM

6.6.1 MECHANICAL CONTROLS

Most aircraft use conventional mechanical controls to move the flight controls. These will normally consist of cables, chains and control tubes. Many examples of this type of system have been described and illustrated previously. The ailerons and elevators on this type of system would normally be operated by a conventional control column and control wheel. Operation of this is instinctive to the pilot, the control wheel being rotated to the left to bank left and right to bank right. Pushing the control column forwards causes the aircraft to dive and pulling back causes the aircraft to climb. A typical control wheel and other cockpit controls is illustrated.

Figure 33


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JAR 66 CATEGORY B1

MODULE 11.09


SYSTEMS

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6.6.2 ELECTRICAL FLIGHT CONTROLS

Many modern aircraft use electrical inputs to the powered control units. This eliminates the need for mechanical controls and all of the chains, pulleys, fairleads and linkages associated with this type of system. This topic is covered in more detail in the Fly By Wire section, but the following paragraphs illustrate a typical Airbus system.

The electrical flight control computers are designed to ensure a high degree of safety. This is accomplished by using a high level of redundancy which consists of five EFCS computers installed in the aircraft, the use of dissimilar redundancy which consists of two types of computers with each being capable of achieving pitch and roll control along with other redundant features assuring aircraft control.

Each computer is also composed of one control unit and one monitoring unit. Control and monitoring software are different and the control and monitoring units are physically separated.

Monitoring

In each computer, one monitoring channel is associated to a control channel by use of self- monitored channels. Each computer is able to detect its own failures (microprocessor test, electrical power monitoring, input and output test). Input monitoring by comparison of signals of the same type, but sent by different sources, and checking of the signal coherence along with permanent cross talk between associated control and monitoring channels, consolidate and validate information received. This allows permanent monitoring of each channel by its associated one. Automatic test sequences can be performed on the ground when electric and hydraulic power is applied (no surface deflection during test).

Side-stick Controller The side-stick controllers are used for pitch and roll manual control and are shown below. The side-stick controllers are installed on the captains and first officer's forward lateral consoles. An adjustable arm-rest is fitted on each seat to facilitate the side-stick control. The side-stick controllers are electrically coupled.

Easa Part 66 - Module 11.09 - Aerodynamics Structures and Systems

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trim controls

fixed trim tabs

adjustable trim tabs

controllable trim tabs

typical trim system

rudder trim system

aileron trim system

tailplane trim system