Easa Part 66 - Module 11.13 - Landing Gear

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JAR 66 CATEGORY B1 MODULE 11.13 LANDING GEAR

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4.1 WHEELS ... 4-22 4.2 TYRES ... 4-23 4.2.1. Tyre Construction ... 4-23 4.3 BRAKES ... 4-25 4.3.1 Energising Brakes ... 4-25 4.3.2 None Energising Brakes... 4-26 4.3.3 Expander Tube Brakes ... 4-26 4.3.4 Single Disc Brakes ... 4-26 4.3.5 Multi Disc Brakes ... 4-26 4.4 ANTI SKID SYSTEMS ... 4-28

4.4.1 Mechanical Anti-Skid Systems ... 4-28 4.4.2 Electronic Anti-Skid Systems ... 4-29 4.5 AUTOBRAKING ... 4-31 4.5.1 Selector Panel ... 4-31 4.5.2 Auto-Brake Control Unit ... 4-31 4.5.3 Auto Brake Solenoid Valve ... 4-32 4.5.4 System Operation ... 4-32 4.5.5 Auto Brake Termination... 4-32 4.6 STEERING ... 4-33 4.6.1 Steering Mechanisms ... 4-35 4.6.2 Nose Wheel Self Centreing ... 4-36

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JAR 66 CATEGORY B1 MODULE NO 11.13 LANDING GEAR

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INTENTIONALLY

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1 LANDING GEAR

1.1 INTRODUCTION

Landing gears can be divided into two main categories, fixed and retractable. All early aircraft had fixed landing gears, due to the lack of technical knowledge and the slow speed that aircraft flew. Later, as speeds increased and the technology to make the landing gears retract, became available, most of the faster aircraft were built with retractable landing gears.

Landing gears have two main functions, supporting the weight of the stationary aircraft and absorbing the loads during touch-down, the landing run and taxiing.

1.2 GENERAL

Generally, most early landing gears were of the format of two main wheels at the front and a little ahead of the Centre of Gravity, (C of G), supporting the majority of the weight of the aircraft, with a smaller wheel or skid at the rear end of the tail fuselage (figure 1). This gave ample clearance between the propeller and the ground during take-off and landing.

The disadvantage of this configuration is that the pilots view is restricted on take off, landing and taxiing, due to the nose up attitude. Also if the aircraft pitches nose down slightly whilst taking off or after landing, the C of G attempts to get ahead of the main wheels and may cause what is known as a ground loop.

Tail Wheel Type Undercarriage Figure 1

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JAR 66 CATEGORY B1 MODULE NO 11.13

LANDING GEAR

To overcome this problem, the tricycle configuration, which is used almost exclusively today, was developed (figure 2). This places the main landing gears a little behind the C of G, with a lighter supporting nose undercarriage at the forward end of the fuselage. It is possible, on larger commercial aircraft to have two, three or four main landing gear legs, depending on the maximum weight and size of the aircraft.

Nose Wheel Type Undercarriage Figure 2

When aircraft flew at low speeds parasite drag was a minor consideration and light weight and ruggedness were the prime requirements. When aircraft began to fly faster streamlined covers were installed over the wheels. Although they added weight they considerably reduced drag. When speeds became of major importance retractable landing gears were introduced.

It is unusual to find aircraft with fixed landing gear if the aircrafts speed is greater than 200 kts. The additional weight of the retraction system and its operating and maintenance costs must balance the reduction in drag and fuel savings. Some aircraft operating in a basic condition will have fixed landing gear to minimise the cost and maintenance requirements.

1.3 CONSTRUCTION

All landing gears have to be attached to strong points on either the fuselage or the wing structure, so that the landing loads can be absorbed and transferred safely to the aircraft structure.

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Smaller light aircraft use a steel leaf or tubular steel spring to act as an undercarriage (figure 3). One end is attached to a strong point on the airframe while located on the other end is the wheel and axle. The deflection of the spring tube on landing absorbs the landing loads and transmit them to the airframe. A properly conducted landing will not cause any undercarriage rebound.

Spring Tube Type Figure 3

Another simple method was to use elastic bungee cord encased in a loose weave cotton braid (Figure 4). The bungee cord is located on a series of support struts which support the wheel and axle. The bungee cord stretches on landing and transfers the landing forces into the airframe.

Bungee Cord Type Landing Gear Figure 4

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LANDING GEAR

Larger more modern aircraft, require more complex and heavier retractable systems (Figure 5). The larger the aircraft the larger the system. The components remain similar just the size and quantities change (Figure 6). Each landing gear unit is basically a wheeled shock absorber (oleo). A forged cylinder body is attached to the airframe on trunnions to allow it to pivot when lowered and raised. Articulated side stays are located between the cylinder body and airframe strong points to give the landing gear strength and rigidity and allow the landing gearleg to fold. Drag or bracing struts may also be fitted. These absorb the high acceleration loads during take off and deceleration loads during braking.

Landing Gear Leg With Bracing Struts Figure 5

The wheel and axle assembly (bogey) is attached to the piston end. A hinged torque (scissor) link is located between the axle yoke and the cylinder body. This allows the piston to move freely in and out of the cylinder but prevents the piston and wheel assembly from swivelling.

Two actuators are usually fitted. A main actuator attached to the cylinder body to raise and lower the gear and a downlock actuator located on the bracing strut which operates to cause a mechanical lock when lowered. It also unlocks the gear mechanism before raising.

B R A C I N G S T R U T M A I N O L E O S ID E S T A Y D O W N L O C K A C T U A T O R P IS T O N T R U N N IO N M A I N S U P P O R T F R A M E S M A I N A C T U A T O R D O W N L O C K L IN K A G E (T O G G L E L E V E R S )

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Oleo Type Landing Gear Figure 6

1.4 MULTIPLE AXLES AND WHEELS

To allow for maximum utilisation of aircraft when operating from different runways multi wheel landing gear is used. Typical configurations are shown in Figure 7.

Wheel Axle Configurations Figure 7 S ING LE T AN DE M D OU BL E B OG IE MAIN ACTUATOR DOWNLOCK ACTUATOR CYLINDER PISTON SCISSOR (TORQUE) LINK WHEEL BRACING STRUT

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LANDING GEAR

The advantages of using multi-wheel configurations are:

• They spread the landing loads over a larger area (footprint). • They are easier to stow as the wheel volume is reduced.

• They provide greater safety. As the loads are spread over several wheels a burst tyre is not so critical as the remaining wheels accept the extra loads. The main disadvantages are:

• There are more moving parts so they need more maintenance. • They are expensive to produce

• Due to the large footprint the turning circle is increased to prevent the tyres from crabbing and increasing wear.

1.5 SHOCK ABSORBING

To absorb shock the mechanical energy of the landing impact must be converted into some other form of energy. This is done on most larger modern aircraft by the use of air/nitrogen-oil or oleo shock absorbers. This uses the fluid flow from one chamber to another, free in one direction and restricted in the other to resist the 'rebound' that will naturally occur after touchdown. Mechanical energy is converted into heat energy in the absorber fluid.

Taxi shocks are cushioned by the air or nitrogen being compressed within the strut. Some light aircraft use rubber snubbers or oil damped coil springs to help damp out taxi shocks or any tendency to rebound on landing. Larger aircraft

1.5.1 Oleo Strut

The oleo strut is a piston within a cylinder and is charged with hydraulic oil and nitrogen. It is made up of 2 chambers that are separated by an orifice and a tapered piston metering pin which moves in and out of an orifice when the strut is compressed or extended.

When the aircraft takes off the combined weight of the wheel and the air/nitrogen pressure inside the strut fully extends the piston to the internal piston extension stop (Figure 8) the fluid to flow down into the hollow piston. This rate of extension has to be controlled to prevent damage to the oleo. The combination of the snubber tube and the snubber knob at the end of the metering pin controls the rate of extension.

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Oleo Strut Extended Figure 8

The holes in the snubber tube slow the rate of extension by controlling the rate at which the oil is allowed to flow into the piston. As the piston reaches its full extension the snubber knob enters the metering orifice and greatly restricts the fluid flow into the piston. This ensures that the extension is controlled up to when the strut reaches its internal extension stop.

Oleo Strut Compressed Figure 9 PISTON SNUBBER KNOB FLAPPER VALVE METERING PIN SNUBBER TUBE CYLINDER INNER CYLINDER METERING PIN SNUBBER KNOB FLAPPER VALVE (OPEN)

PISTON CYLINDER

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LANDING GEAR

When the aircraft lands the oil is forced into the upper chamber through the orifice into the snubber tube and into the inner cylinder through the flap (check) valve. The small end of the metering valve is in the orifice as the strut begins to compress and its tapered shape steadily decreases the area of the orifice as it compresses. The landing energy is absorbed by the oil as it is forced through the decreasing sized orifice and by the air/nitrogen which compresses as the oil is forced into the upper chamber.

The momentum of the aircraft landing compresses the strut to more than is required to support the weight of the aircraft and when maximum compression is reached the aircraft tries to rebound or bounce (Figure 9) back into the piston. This closes the flap valve and forces the fluid at a restricted rate to flow through holes in the snubber tube. This restriction of flow (damping) prevents the rapid extension of the strut which would otherwise cause the aircraft to bounce..

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2 EXTENSION AND RETRACTION SYSTEMS

As the speed of the aircraft becomes high enough that the parasite drag of the landing gear is greater than the induced drag caused by the added weight of the retracting system it becomes economically practical to retract the landing gear into the aircraft structure.

Raising or lowering of the undercarriage is carried out either hydraulically or pneumatically via a selector lever in the cockpit which is mechanically or electrically linked to a selector valve. When the selector valve is operated it directs the fluid to one side or the other of the piston.

The landing gear is uplocked and downlocked mechanically or hydraulically through the uplock boxes and the downlock toggle levers. Landing gear positions are sensed by proximity switches or microswitches and transmit these positions to the cockpit instrumentation via a control unit.

In the case of fluid or electrical failure, a mechanical emergency lowering system is available. An emergency handle located in the cockpit is operated and by a system of push-pull cables and gearboxes, the uplocks are released.

The landing gear selector valve or a freefall valve is also operated, which opens all extension and retraction lines to return. The landing gear is allowed to fall under gravity and aerodynamic forces but may be assisted by a spring or gas operated free fall assister.

Smaller light aircraft may use differing methods for operating the landing gear. Electric motors may drive actuators, a winding cable system, a simple operating lever with safety locks or a manual hydraulic jacking system may be used to raise or lower the landing gear.

Most modern aircraft use a hydraulic power pack. This is a self-contained system and was designed to be lightweight and easy to maintain. The pack contains the fluid reservoir, sight glass, pressure pump, filter, thermal relief valve, pressure relief valve, ground service and replenishment connections.

2.1.1 Extension System

When the selector lever is selected to GEAR DOWN a micro-switch on the lever is made which powers up the hydraulic pump, the hydraulic pressure is then fed to the uplock actuator valves to unlock the uplocks. Once operated, the uplock hooks remain mechanically open under spring pressure. Movement of the undercarriage legs break the uplock limit switches which indicates on the instrumentation panel that the landing gears are in transit.(red triangles) and that the undercarriage is unlocked.

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LANDING GEAR

The landing gear selector valve operates, and the down lines to the actuators and the return lines to the reservoir are opened. The fluid pressure flows through the selector valve to the actuators and extends the actuators. Once the main actuators are fully extended and the undercarriage legs have mechanically locked, excess pressure is bled back through the low pressure control valve to the reservoir.

When all 3 wheels are down and locked, proximity switches send signals to a control unit which turns the hydraulic pump off, closes the selector valve lines and sends signals to the instrument panel indicating that the undercarriage is locked down, (green triangles).

2.1.2 Retraction System

The retraction procedure is basically the opposite of the extension procedure. When the selector lever is selected GEAR UP a micro-switch on the lever is made which powers up the hydraulic pump, the hydraulic pressure is then fed to the downlock actuators to unlock the mechanical locks on the bracing struts. Its is also fed to the selector valve and opens the uplines to the main actuators and the return lines to the reservoir.

Movement of the undercarriage legs breaks the downlock proximity switches which send signals to the control unit which indicates on the instrumentation panel that the landing gears are in transit, (red triangles) and that the undercarriage is unlocked.

The fluid pressure flows through the selector valve to the main actuators and retracts the landing gear. The undercarriage legs on full retraction mechanically lock the uplocks. Once the main actuators are fully retracted and the undercarriage legs are locked up, excess pressure is bled back through the low pressure control valve to the reservoir. When all 3 wheels are up and locked, uplock limit switches send signals to a control unit which turns the hydraulic pump off, closes the selector valve lines and change the red triangles to black on the indicating panel.

If a red triangle remains on when the undercarriage is fully extended or retracted there is a fault in the system. A squat switch system and an electro-mechanical stop on the selector lever, will prevent the landing gear from being retracted when the aircraft is on the ground. The landing gear will not be able to be retracted until certain parameters are met. This is normally when all landing gear legs have fully extended after take off. This is sensed by proximity switches on each leg.

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2.2 SELECTOR VALVE

The selector valve on modern large aircraft will be normally operated by electrical solenoids signalled from micro-switches in the landing gear selector lever, but on some aircraft they may be mechanically operated. A spool valve in the selector valve is moved from a neutral position one way or the other allowing hydraulic pressure to one side of the main actuator piston, depending whether the landing gear is to be raised or lowered .

Normal operation of the selector valve can be overridden in case the landing gear has to be lowered in an emergency, if the landing gear fails to extend due to a system fault. The spool valve is moved mechanically by a system of rods, cables and levers to allow all lines to be opened to allow the free flow of hydraulic fluid around the system. This operation is normally inter-linked with the emergency mechanical opening of the uplocks.

A typical selector valve is shown in Figure 10

Selector Valve Figure 10

2.3 UPLOCK MECHANISM

On large modern aircraft when the landing gear is being retracted the uplocks will operate mechanically. A roller on the landing gear leg will locate and engage into the uplock hook. Limit switches will sense when the landing gear leg has engaged in the lock hook and will turn off the hydraulic pressure. The gear will then be held retracted in place purely mechanically. (Figure 11)

SOLENOID SOLENOID

MECHANICAL OVER-RIDE LINKAGE

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JAR 66 CATEGORY B1 MODULE NO 11.13 LANDING GEAR Locked Uplock Figure 11

Normal release of the uplock is by a hydraulically actuated valve. The supplied hydraulic pressure pushes a plunger against the lock lever which rotates about its pivot. This action allows the uplock hook to disengage under its own spring tension. The landing gear will then be extended hydraulically by the main actuator. (Figure 12)

Unlocked Uplock Figure 12

UNLOCK ACTUATOR VALVE

UPLOCK HOOK

LIMIT SWITCH LOCK LEVER ASSEMBLY

LANDING GEAR LEG ROLLER

UNLOCK ACTUATOR VALVE

UPLOCK HOOK

LIMIT SWITCH LOCK LEVER ASSEMBLY

LANDING GEAR LEG ROLLER PLUNGER

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2.4 DOWNLOCK MECHANISM

The downlock actuator can have either a single or double direction operation depending on the aircraft. A single direction operation would unlock the downlock mechanism (upper and lower toggles) prior to retraction, the leg relying on its own extension to provide the over centre lock. The double direction actuator will lock the downlock mechanism on extension and unlock it prior to retraction.

Once the landing gear has been fully extended and is sensed by a limit switch hydraulic pressure is directed to the downlock actuator which extends the actuator piston. The piston acts against a toggle lever which move both toggle levers to an over centre position. This over centreing of the toggle levers forms a mechanical lock which prevents the landing gear leg from collapsing. (Figure 13)

Linkage Downlocked Figure 13

Once the aircraft has landed and parked up, a red flagged safety pin is inserted through alignment holes in the toggle levers to prevent inadvertent collapse or retraction of the landing gear on the ground. This safety pin is removed before flight. DOWNLOCK ACTUATOR LOWER TOGGLE LEVER PROXIMITY SWITCH PROXIMITY SWITCH SIDE BRACE MAIN LEG UPPER TOGGLE LEVER CENTRE LINE

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LANDING GEAR

On selecting the landing gear up, the hydraulic pressure is directed initially to the downlock actuator and retracts the piston. As the piston retracts it moves the lower toggle overcoming the mechanical lock, moving both toggle levers from the over centre position to an under centre position, so that the landing gear can now fold. (Figure 14)

Linkage Unlocked

Figure 14

2.5 EMERGENCY LANDING GEAR OPERATION

The uplocks can be released manually if the actuator or hydraulic system fails. An emergency landing gear lever, operated from the cockpit will act on and rotate the hook locks, releasing the landing gear legs from the uplock hooks. The emergency mechanism lever will also operate a lever on the landing gear selector valve which will open all hydraulic lines to return. This allows the hydraulic fluid to free flow through the system, to allow the landing gear to extend.

Once the uplocks are released the landing gear legs will extend under gravity and aerodynamic forces. Spring or gas operated free fall assistors may be used to help the gear extend. The proximity and limit switches will operate as normal giving a cockpit indication of the gear in transit and down locked.

MAIN LEG PROXIMITY SWITCH DOWNLOCK ACTUATOR SIDE STAY UPPER TOGGLE LEVER LOWER TOGGLE LEVER PROXIMITY SWITCH CENTRE LINE

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Emergency Release Mechanism Figure 15

On aircraft fitted with hydraulically sequenced doors if the hydraulic system fails, the door jack is mechanically unlocked. This will also be carried out by a mechanical linkage connected to the cockpits emergency release mechanism (Figure 15)

2.6 LANDING GEAR DOORS SEQUENCING

To keep the aircraft as streamlined as possible and to reduce drag, the landing gear is normally retracted into bays within the aircraft structure. However some aircrafts landing gear do not fully retract into the structure and some access doors do not fully enclose the landing gear.

The bays have access doors which open and close in relation to the movement of the landing gear. Some doors are mechanically linked to the landing gear, by a system of connecting rods, bellcranks and links, whilst other doors open and close under operation from a hydraulic sequencing valve, signalled by micro-switches or proximity micro-switches via a control unit.

UNLOCK ACTUATOR VALVE UPLOCK HOOK

LIMIT SWITCH HOOK LINK ASSEMBLY

EMERGENCY OPERATING HANDLE

LEVER

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LANDING GEAR

To further reduce the drag some doors will close when the landing gear has been extended. The landing gear doors may have a manual unlocking mechanism to allow the door to be opened on the ground for access in carrying out maintenance tasks and inspections.

Anything that jeopardises the sequence can cause considerable damage to the aircraft structure and could lead to an unsafe landing condition. Door sequencing relies on the movement of valves operated by the doors and the movement of the legs. The sequencing valve can be therefore be either door operated or gear operated.

2.6.1 Door Operated Sequencing System

Only when the door is fully open is pressure allowed to flow to the main actuator. If the door is not fully open the main actuator remains isolated. Hydraulic pressure is initially fed to the landing gear door actuator which operates to open the door. When the door reaches its maximum travel it abuts against. and depresses a plunger. (Figure 16) The movement of the plunger unseats a valve in the sequence valve, which opens a gallery to allow fluid pressure to the main actuator and extends the landing gear down.

Sequence Valve – Door Shut Figure 16 PLUNGER PRESSURE IN TO DOOR ACTUATOR TO MAIN ACTUATOR VALVE SEAT

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Retraction of the landing gear is reversed. Pressure is fed to the main actuator which retracts the landing gear leg. When the landing gear leg is fully retracted it abuts against and depresses a sequence valve plunger. The movement of the plunger unseats a valve in the sequence valve, which opens a gallery to allow fluid pressure to the door actuator which closes the door. (Figure 17)

Sequence Valve – Door Open Figure 17

2.6.2 Gear Operated Sequencing System

The principle of operation is very similar to the door operated mechanism. The difference being that the plunger (or slide) is operated via a cam and linkage mechanism directly attached to the landing gear leg. This ensures that when the gear starts to move the door starts to, or is in the process of opening.

2.7 SAFETY BARS

On some aircraft with hydraulically sequenced doors if the hydraulics system was to fail, to allow the landing gear to lower, the wheels will forcibly open the doors. This is done by the landing gear legs pushing against safety bars which are fitted to the doors. The doors will open without being damaged and once operated the doors will remain open.

PLUNGER PRESSURE IN TO DOOR ACTUATOR

TO MAIN ACTUATOR VALVE SEAT

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LANDING GEAR

3 INDICATIONS AND WARNING

It is essential for the pilot to know the position of the landing gear, so all aircraft fitted with retractable gears will have some method of indicating whether the landing gear is locked up, in transit or locked down or whether there is a fault in the system.

The main method used is by visual indications on the instrument panel. There is usually one indication for each of the landing gear legs. Red triangles illuminate when the landing gear is in transit (either moving up or down). Green triangles illuminate when the landing gear is locked down.

All illuminations are extinguished when the landing gear is locked up (black cockpit concept). If there is a mix of green or red illuminations this will indicate a fault in the system and associated audio warnings and warning lights will flash to alert the pilot. Each pilot may have his own indicator panel on the instrument panel.

Other methods can be mechanical indicators outside the aircraft, visible from the cockpit. There may be painted indicator lines on the landing gear legs toggle levers which align when the gear is down and locked. (Figure 18)

Landing Gear Down Locked Visual Indicator Figure 18

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Some aircraft have pop up indicators which stand proud on the upper wing surface when the gear is down and locked (Figure 19). These are plunger operated through a cable linkage attached to the toggle levers. When the landing gear extends and is locked down a plate attached to the toggle lever operates a spring loaded plunger which by cable connection moves the indicator from its housing, proud of the airframe skin. The indicator returns under spring pressure into its housing when the landing gear is retracted

Landing Gear pop Up Indicator Figure 19 POP UP INDICATOR PLUNGER TOGGLE LEVERS SIDE STRUT UNLOCK ACTUATOR AIRFRAME SKIN POP UP INDICATOR PLUNGER TOGGLE

LEVERS SIDE STRUT UNLOCK ACTUATOR AIRFRAME SKIN

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LANDING GEAR

To prevent the pilot from landing with his under carriage retracted there may be a warning system connected to the centralised warning panel with associated warning lights and audio warnings. The warning system may be activated when the aircraft descends to a certain height above the ground detected by the radio altimeter, or when the landing configuration is incorrect ie, when the engine power levers or flaps are set incorrectly.

Landing Gear Selector Lever Safety Interlock Figure 20

Landing Gear Selector Lever Safety Interlock Figure 21 SAFETY SOLENOID DE-ENERGISED LANDING GEAR LEG EXTENSION LIMIT SWITCHES DOWN UP LANDING GEAR SELECTOR LEVER SAFETY LATCH PIN

CONTROL UNIT LANDING GEAR SELECTOR LEVER SAFETY SOLENOID ENERGISED CONTROL UNIT DOWN UP SAFETY LATCH PIN

LANDING GEAR LEG EXTENSION LIMIT SWITCHES

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The landing gear may have an electro-mechanical safety device, which prevents operation of the selector lever on the ground. When all the landing gear legs are compressed a safety solenoid is de-energised which moves a latch pin under the landing gear selector lever. So long as the solenoid remains de-energised the latch pin prevents the selector lever from operating.

As soon as each landing gear leg is fully extended the limit switch is made which sends a signal to the control unit. When the control unit receives signals from all the landing gear legs an earth is made and the safety solenoid is energised. The latch pin is withdrawn from beneath the selector lever allowing gear up when selected. (Figures 20 and 21)

3.1 SAFETY SWITCHES

Proximity switches on each landing gear leg will indicate that the landing gear leg is either downlocked or is in transit. The switch will be made when the target on the landing gear leg comes into alignment with the switch probe indicating that the landing gear is downlocked. The gap between the probe and target is set in accordance with the maintenance manual for the aircraft. When the proximity switche probes are out of alignment with their targets, the switches are broken and it is sensed that the landing gear leg is in transit.

The signals will be sent to an electronic control unit or computer where they are processed and will illuminate an associated green triangle on the landing gear panel when locked down and a red triangle when the landing gear is in transit.

Limit micro-switches on the uplocks will sense when the landing gear is locked up and limit switches on the oleos will sense when the oleo leg is fully extended. The signals will be sent to an electronic control unit or a computer where they are processed. When the landing gear is locked up the limit switch will change the red triangles to black. When the oleos are fully extended the limit switches will allow the landing gear to be retracted.

The proximity switches and limit switches form part of the weight on wheels, weight off wheels squat switch system and will prevent inadvertent retraction of landing gear on the ground and will only allow retraction when certain circumstances are met. This mainly being that all 3 landing gear legs are weight off wheels and are fully extended, and the downlocks have been unlocked.

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LANDING GEAR

4 WHEELS, BRAKES, ANTISKID AND AUTOBRAKING

4.1 WHEELS

The wheels on the landing gear leg provides some form of suspension and adhesion between the aircraft and the ground. Early wheels and tyres were of the bicycle type with spoke rims and with the tyres fitted using tyre levers. Most light aircraft have fixed flange one piece forged or cast wheels (Figure 22).

Fixed Flange Wheel Figure 22

Modern tyres are much more rigid, due to the load-bearing requirements, which results in the wheels having to be of two piece construction (Figure 23). The two piece wheel construction, are of 2 types, removable rim or split wheel. The removable rim wheel has an inner tube where as the split wheel is tubeless and requires a perfect seal between the halves. An O ring is located between the mating surfaces. To be as light and strong as possible they are usually constructed from alluminium or magnesium alloys and may be cast or forged. The inboard wheel section is fitted with key ways that allows the brake discs to slot into. These key ways drive the brake discs with the wheels. Larger aircraft wheels have one or more fusible plugs fitted. These plugs have a centre hole which is filled with a low melting point alloy. In the event of the tyre overheating, when a temperature limit is reached the low melting point alloy melts and allows the tyre to safely deflate.

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Some tyres are fitted with over-inflation safety valves so that when a tyre is over inflated to a dangerous pressure the safety valve ruptures and bleeds off any excess pressure and deflates the tyre before there is a chance of the tyre exploding.

Split Wheel Assembly Figure 23

The outboard wheel section houses the anti skid braking mechanism if fitted and also the charging valve. If an inner tube is fitted there is a hole in the wheel through which the charging valve stem protrudes.

4.2 TYRES

Tyres with patterned tread became important when aircraft got effective brakes that could be used for slowing the aircraft during landing. At first the treads were a diamond pattern that provided good braking on wet grass but the ribbed tread proved to be more suitable for operation on hard surface runways. Today almost all aircraft tyres have a ribbed tread that consists of straight grooves, which run around the tyres circumference.

4.2.1. Tyre Construction

Figure 24 show a typical aircraft tyre and its major components.

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O RING

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JAR 66 CATEGORY B1 MODULE NO 11.13 LANDING GEAR • The Bead •

• The bead gives the tyre its strength and stiffness to assure a firm mounting on the wheel. The bead is made up of bundles of high strength carbon steel wire with two or three bead bundles on each side of the tyre. Rubber strips

streamline the round bead bundles to allow the fabric to fit smoothly around them without any gaps. The bead bundles are enclosed in layers of

rubberised fabric, to insulate the carcass plies from the heat absorbed in the bead wires.

•

• The Carcass

The carcass (or chord body) is the body of the tyre that is made up of layers of rubberised fabric cut in strips with the threads running at an angle of about 45 degrees to the length of the strip. These strips extend completely across the tyre around the bead and partially up the side. Each ply is put on in such a way that the threads cross each other at about 90 degrees to that of the adjacent ply. This type of construction is known as bias ply.

Aircraft Tyre Construction Figure 24

BEAD BUNDLE

CARCASS

TREAD

SIDEWALL

BEAD WIRES

CHAFING STRIPS

PLIES

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The cords of the ply fabric were originally cotton, then nylon and now aramid fibres (kevlar) are used. This is stronger than nylon, polyester or fibreglass and even strong pound for pound than steel.

Chafing strips are rubberised strips of fabric that wrap around the edges of the carcass plies and enclose the bead area. The chafing strips provide a smooth chafe resistant surface between the tyre and the bead seat of the wheel.

The undertread is a layer of compound rubber between the plies and the tread rubber that provides good adhesion between the tread and the carcass. On top of the undertread are more plies of strong fabric that strengthen the tread and oppose centrifugal forces that try to pull the tread from the carcass during high speed rotation.

The inner liner is a thin coating of rubber over the inside plies. For tubeless tyres it is made from a compound which is less permeable than other rubbers used. It seals the tyre and reduces the amount of leakage. On tyres with inner tubes the liner is very smooth to help prevent chafing.

• The Tread

The tread is the thick layered rubber around the outer circumference of the tyre that serves as a wearing surface. The tread has a series of moulded grooves moulded into its surface to give optimum traction with the runway surface.

4.3 BRAKES

Aircraft brake systems convert kinetic energy from the motion of the aircraft into heat energy, which is generated by the fiction between the brake linings and the brake drum or disc.

There are two types of brakes in use energising (servo) and none energising. Energising brakes use the friction developed between the rotating and stationary parts to produce a wedging action that uses the momentum of the aircraft to increase the braking force which reduces the pilots effort needed in producing the required braking action. None energising brakes do not use this wedging action. 4.3.1 Energising Brakes

Energising brakes used on some smaller aircraft have a single servo action and only operate with forward motion. Energising brakes have their shoes and linings mounted on a torque plate in such a way that they are free to move out against the rotating drum. When the brakes are applied two pistons move out and push the linings against the drum that rotates with the wheel. Rotation of the brake drum wedges the linings against it. When the hydraulic pressure is released, a retracting spring pulls the linings form the drum and releases the brakes.

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4.3.2 None Energising Brakes

This is the most common type of brake used on aircraft. These brakes are actuated by hydraulic pressure and the amount of braking action depends on the pressure applied. Expander tube, single disc and multiple disc brakes are the main types of none energising brakes used.

4.3.3 Expander Tube Brakes

Expander brakes use a heavy neoprene tube. Hydraulic fluid from the master cylinder is directed into the expander tube which is located on the circumference of a torque flange. When this tube is expanded it pushes the brake block linings out against the brake drum and the friction between the linings and the drum slows the aircraft.

The heat generated in the linings is kept from damaging the expander tube by stainless steel heat shields placed between each of the lining blocks. As soon as the brake pedal is released, the return springs between the brake lining blocks collapse the expander tube and force the fluid back into the cylinder reservoir.

4.3.4 Single Disc Brakes

This is most common on light aircraft. The brakes are actuated by hydraulic pressure from a master cylinder and friction is produced when the rotating disc is squeezed between the brake linings in the brake caliper.

There are two types of single disc brakes, one has the disc keyed into the wheel and it is free to move in and out as the brake is applied. This type is called floating disc fixed caliper. The second type of brake disc is rigidly attached to the wheel and the caliper moves in and out on anchor bolts. This type is called fixed disc floating caliper.

Some single disc brakes have automatic adjusters and wear indicators. The automatic adjusting pin is pulled through the grip when brakes are applied. When the brakes are released the piston and the linings move back only under pressure of the return spring. The protrusion on the adjuster pin indicates lining wear. In general, when the pin is flush with the housing the linings are replaced.

4.3.5 Multi Disc Brakes

The gross weight of the aircraft and the speed at the time of brake application determines what size brakes are required. As the aircrafts size, weight and landing speed increases there is a need for greater braking surfaces and heat dissipation.

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Segmented rotor, multiple disc brakes are standard on most modern high performance aircraft. The segmented disc brake has three rotating discs keyed on to the wheel. The rotors are segmented to allow for cooling and for expansion caused by the high temperatures generated during braking.

Between each disc is a stator plate or brake lining disc, keyed on to the axle shaft. Rivetted on to each side of the stator plates are the brake linings. A pressure plate is located on the inboard side of the axle shaft and a backing plate is located on the outboard side.

Automatic adjusting pins are pulled through the grip when brakes are applied. When the brakes are released the pressure plate moves back under pressure of their return springs. The protrusion on the adjuster pins indicates lining wear. In general, when the pin is flush with the housing the linings are replaced.

Multi Disc Brake Unit Figure 25

The brakes used on most large jet aircraft use a number of brake cylinders instead of a single annular cylinder. (Figure 25) Each cylinder has a piston which presses against the pressure plate when hydraulic pressure is applied. Each cylinder will be supplied from separate hydraulic systems so if one fails full braking can be applied from the other system.

Some aircraft may have their brake discs made from carbon fibre. These are lighter in weight and they can function at higher temperatures. They are expensive to use and generally only used on transport aircraft where the weight saving makes them more cost effective.

CYLINDERS

WEAR PINS

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4.4 ANTI SKID SYSTEMS

Anti skid systems work on the principle of releasing the brake on a wheel that has or is just about to lock. There are two types of systems in use, mechanical or electronic. The mechanical system uses the principle of inertial weights the electronic system uses wheel speed signals sent to a computer which in turn transmits feedback signals to the brake unit.

The advantages of using anti-skid systems are: • Maximum braking efficiency

• Reduction of the landing run.

• No skidding or locking of the wheels • Increased tyre life

• The plilot can apply maximum braking effort without locking the wheels.

4.4.1 Mechanical Anti-Skid Systems

This type of system uses components that are situated around the wheel area. They are self contained systems which when required will change the supply of metered brake pressure going to the brake unit. They cab be externally or axle mounted. They are usually referred to as MAXARETS (maximum retardation units) and the principle of operation is identical.

Externally mounted maxarets are mounted on the brake torque plate, the leg or the bogie. The maxaret is driven by a small rubber tyred wheel which contacts the aircraft wheel and rotates with the aircraft wheel. Axle mounted maxarets are mounted inside the wheel axles and are less susceptible to damage. They are driven by the aircraft wheel via a hub cap and a flexible drive.

• Operation

The maxaret (Figure 26) is connected to the hydraulic line between the brake metering valve and the brake unit. The unit is sensitive to the angular deceleration which occurs when braking and just before the wheel locks the energy in the flywheel is released.

As the wheel speed decreases the flywheel keeps on rotating due to its inertia. The flywheel drives thrust balls, up a cam which operates a push rod inside the brake unit. The push rod acts on a lever which pivots and shuts off the pressure supply and opens the brake unit return, which in turn releases the brake pressure.

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When the wheel regains speed it realigns itself with the flywheel. The thrust balls move back down the cam. The push rod realigns under spring pressure and the pivot lever reopens the supply line. The brake reapplies until the angular deceleration is such that the system senses a lock and the system operates again.

Mechanical Anti Skid System Figure 26

4.4.2 Electronic Anti-Skid Systems

Most modern aircraft have electronic anti-skid systems which prevent the wheels from locking and skidding by releasing the brake pressure. They are more

sensitive than mechanical systems and they can modulate the brake pressures for optimum braking efficiency. They are more reliable than mechanical systems and require less maintenance. (Figure 27)

Only the wheel speed transducer is located around the wheel area, the remainder is usually located remotely within the aircrafts fuselage. The system is made up of:

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• a wheel speed transducer. • an anti-skid control unit. • an anti-skid valve.

Electronic Anti Skid System Figure 27

Speed transducers are speed sensing devices whose rotation creates an ac signal, the voltage produced being proportional to the wheel speed. The speed of the wheel is sensed by the transducer and a signal is sent to control unit. In the control unit the ac voltage is converted to dc voltage. This voltage is compared to a reference voltage which has been set at the maximum deceleration rate for the aircraft.

A rapid deceleration of wheel speed caused under braking will reduce the transducer output voltage down to the reference voltage. The control unit senses this drop and when the reference voltage is reached it sends a signal to the anti-skid control valve to release the brake. The wheel will then regain speed, the transducers output voltage increases, the control unit de-energises the valve and the brakes are re-applied.

The anti-skid valve receives the signals from the control unit. The metered brake pressure goes to the pressure side of the flapper valve. If no signal is present from the anti-skid control unit the flapper covers the return port. The pressure pushes the spool valve across to allow the metered pressure to the brake unit. Also the feedback chamber is fed with brake pressure ready for immediate brake release when required.

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When the transducer senses a wheel skid, signals are sent from the control unit which moves the anti-skid control valve flapper to cover the pressure port. The return port is opened and the spool valve is pushed by the feedback chamber to open the return port and release the brakes.

Some electronic systems incorporate a safety device to prevent landing with the brakes applied, the brake pressure being held off until a pre-determined wheel speed is reached. Some systems allow the aircraft to land with brakes applied.

4.5 AUTOBRAKING

Some modern aircraft have auto-braking systems. A selector switch on the instrument panel allows the pilot to select a deceleration rate that will be controlled automatically after landing. On landing the auto-braking system will smoothly apply the brakes to achieve the selected deceleration rate down to a complete stop without any further action from the aircrew. This allows the aircrew to concentrate on other activities during landing.

The auto brake system utilises the normal anti-skid and brake units but instead of using pressure from the brake metering valve, hydraulic pressure is sent via solenoid valves which allow a pre-determined amount of pressure through the anti-skid valves to the brake units.

4.5.1 Selector Panel

The selector panel consists of a solenoid latched switch which will hold a selected position only if all the arming conditions for that setting are met. If the system cannot be armed the switch will automatically return to the DISARM position and a warning will illuminate on the local panel and centralised warning panel. The panel will have a number of settings that the pilot can select depending on the rate of deceleration that is required.

4.5.2 Auto-Brake Control Unit

Selection on the auto brake selector panel will send an electrical signal to the auto-brake control unit. The signal is processed by the control unit, which commands the solenoid valve to direct pressure to the brake units.

The brake pressure must be gradually built up and released to prevent brake snatch and jerking. To prevent this a time delay and an electrical ramp are used. The time delay ensures that the aircraft is firmly on the ground before the system activates. The terminology used to indicate the auto-brake operation is:

• On Ramp – A gradual build up of brake pressure to the amount required for the selected rate of deceleration.

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• Off Ramp – A gradual decrease in pressure down to zero at the end of the landing run or cancellation of auto-brake.

• Drop Out – Instantaneous pressure release to zero (go around mode).

4.5.3 Auto Brake Solenoid Valve

These valves are electrically controlled, hydraulic valves that allow pressure to the brake units at a specific setting. The greater the deceleration rate the higher the setting. These valves are fitted just upstream of the anti-skid valves.

The solenoid will open when all the arming conditions are met and the aircraft is weight on wheels. It is also the solenoid valves that immediately shuts on Drop Out.

A solenoid servo valve modulates the brake pressure to regulate the deceleration rate. A pressure switch is connected to the DISARM warning light to monitor zero pressure when auto-brakes are armed.

4.5.4 System Operation

Once the aircraft lands and is weight on wheels the anti skid transducers send signals to the control unit. When the wheels have achieved a certain speed or after a pre-determined time delay the brakes will be applied “Up The Ramp”. Once the selected rate of deceleration is reached the auto-brake pressure is modulated to hold that rate.

As the wheel speed slows down to more than the deceleration rate, the servo valve will close slightly reducing the brake pressure causing the wheel to speed up. Once the aircraft has come to a stop or the aircraft is below a certain speed the auto-brakes will switch off to enable the aircraft to taxi.

4.5.5 Auto Brake Termination

Auto-Brake can be cancelled at any time. Depending on the aircraft, the system can be over-ridden by:

• The pilot moving the selector lever to disarm or off. • The pilot using manual braking.

Auto-brake needs to be immediately cancelled if the pilot has to initiate a go round procedure. The following actions will cause immediate DROP OUT:

• The thrust levers are advanced from the idle gate.

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4.6 STEERING

To improve the ground operation of aircraft nose wheel systems are used. These improve tyre life through less scrub, reduce brake wear, save fuel and engine life as brakes and engine thrust are no longer required to turn the aircraft.

Most nose wheel steering systems use servo jack operated scissor links attached to a collar on the landing gear leg, the collar being driven by the servo jacks which rotates the nose wheel leg via the scissor links. Steering inputs to the servo jacks come from a tiller on the pilots side of the cockpit. Inputs can also come from the rudder pedals.

Apart from mechanical steering systems there are three basic methods of operation:

• Single Servo Jack. T

This system is used on smaller light aircraft (Figure 28). Both ends of the jack ram are attached to the landing gear leg. Fluid is directed to move the jack body along its ram. A cam and link assembly is attached to the jack body. Movement of the jack body operates the link which rotates the cam and turns the wheel. Action of the shock absorber is unaffected as the shock absorber is splined on to the steering shaft to allow the compression and extension of the absorber.

Single Servo Steering Figure 28 SPLINED SHAFT CAM LINK JACK BODY PISTON STRUT AXLE

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• Double Servo Jack.

Larger aircraft use a two servo jack system (Figure 29). The two jacks are fixed to a steering collar, which is free to rotate around the landing gear leg. The steering collar is attached to the upper scissor link. When the servo jacks are actuated they rotate the wheels and axle through the scissor link. assembly

Double Servo Jack Figure 29

• Rack and Pinion

Some aircraft use a rack and pinion steering system. Hydraulically operated racks rotate a pinion which rotates the wheel and axle. A mechanical linkage from the cockpit tiller operates a servo valve in a hydraulic metering valve. The servo valve when operated directs fluid to one side or the other of the rack piston. The rack then moves and rotates the pinion and turns the aircraft nose wheel in the required direction.

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4.6.1 Steering Mechanisms

On some small aircraft the nose wheel is steered by direct linkages from the rudder pedals, or on small retractable landing gear aircraft, from the rudder pedals to a steering bar which locates against a steering arm on the landing gear leg. (Figure 30) Once the wheel is stowed the mechanism is ineffective.

Nose Wheel Steering Mechanism Figure 30

The nose wheels or tail wheels on light aircraft mat be steerable or castoring. A castoring nose wheel aircraft is steered by the independent use of the brakes and rudder inputs. Some light aircraft have limited tail wheel steering via a mechanism interlinked with the rudder pedals. The tail wheel will brake out if the turning circle is too small to allow the tail wheel to castor. Once centralised the tail wheel becomes steerable again. Some aircraft have a tail skid mechanism (Figure 31)

STEERING ARMS

STEERING BAR

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Tail Skid Mechanism Figure 31

Inputs to the hydraulic control valves which direct pressure to the steering jacks are carried out by a mechanical system of cables, bellcranks, levers and gearboxes from the hand operated tiller and the rudder pedals. The input has a follow up action through interconnected links or cables which neutralise the nose wheel movement when the desired rate of turn has been achieved.

Rudder pedals movement can also be inputted to the control valve, but this is usually restricted to a small degree of movement either side of the aircraft centre line. Rudder pedal steering is normally used on take off or landing and is isolated when the aircraft is airborne.

4.6.2 Nose Wheel Self Centreing

It is important that when a steerable nose wheel is being retracted that the wheel is centred so that it fits into the wheel well to prevent any damage to the aircraft structure as well as the landing gear. This is done by a centreing cam inside the oleo strut. When the strut is compressed the piston cam disengages from the cylinder cam receptacle to allow the wheel to be steered. On take off when the strut extends the piston cam is forced into the cylinder receptacle to hold the wheel in the desired position for stowing.

Easa Part 66 - Module 11.13 - Landing Gear

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CONTENTS

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PAGE

INTENTIONALLY

BLANK

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1

LANDING GEAR

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2

EXTENSION AND RETRACTION SYSTEMS

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3

INDICATIONS AND WARNING

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4

WHEELS, BRAKES, ANTISKID AND AUTOBRAKING

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BEAD BUNDLE

CARCASS

TREAD

SIDEWALL

BEAD WIRES

CHAFING STRIPS

PLIES

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