Module 11.13_b1_rev 00.pdf

Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

MODULE 11 SUB MODULE 11.13 LANDING GEAR (ATA 32)

Rev. 00 Oct 2006

11.13 For Training Purposes Only

Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL Table of Contents

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Table of Contents

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Function........................................................................................ 2

Electrical Retraction System ....................................................... 30

Landing Gear Classification .......................................................... 2

Landing Gear Control.................................................................. 32

Landing Gear Configuration (Wheel Arrangement)....................... 4

Landing Gear Position Indication ................................................ 34

Conventional Or Tail Wheel-Type Landing Gear .......................... 4

Safety Features........................................................................... 36

Tricycle-Type Landing Gear ......................................................... 4

Additional Features ..................................................................... 40

Tandem Landing Gear.................................................................. 4

Maintenance ............................................................................... 42

Retractable Vs. Fixed ................................................................... 6

Powered Steering........................................................................ 46

Fixed Landing Gear ...................................................................... 8

Hydraulic Steering System.......................................................... 48

Fixed Landing Gear Shock Absorbers .......................................... 8

Steering Control .......................................................................... 50

Spring Steel Legs ....................................................................... 10

Maintenance ............................................................................... 52

Rubber Cord ............................................................................... 10

Aircraft Wheels............................................................................ 56

Oleo-Pneumatic Shock Absorber................................................ 12

Introduction ................................................................................. 56

Shimmy Dampers ....................................................................... 16

Construction ................................................................................ 58

Retractable Landing Gear........................................................... 18

Split Hub Wheel Construction ..................................................... 60

Landing-Gear Components ........................................................ 20

Wheel Inspection – On The Aircraft ............................................ 62

Bogie Undercarriages ................................................................. 24

Wheel Inspection – Off The Aircraft ............................................ 63

Hydraulic Retraction System ...................................................... 26

Aircraft Tires And Tubes ............................................................. 70

Simple Hydraulic Retraction System........................................... 28

Tire Classification........................................................................ 70

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL Table of Contents

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Type............................................................................................ 70

Inflation ....................................................................................... 82

Size............................................................................................. 70

Tread Condition........................................................................... 84

Ply Rating ................................................................................... 72

Sidewall Condition....................................................................... 86

Tube Or Tubeless ....................................................................... 72

Tire Inspection Off Of The Aircraft............................................... 88

Tire Construction ........................................................................ 74

Tire Retreading ........................................................................... 90

Bead ........................................................................................... 74

Tire Storage ................................................................................ 90

Carcass (Bias Tire) ..................................................................... 74

Aircraft Tubes.............................................................................. 92

Carcass (Radial Tire).................................................................. 76

Tube Construction And Selection................................................ 92

The Tread ................................................................................... 76

Tube Inspection........................................................................... 92

Deflector ..................................................................................... 78

Tube Storage .............................................................................. 94

The Sidewall ............................................................................... 78

Tire Mounting – Tube-Type Tires ................................................ 96

Tire Markings .............................................................................. 80

Tire Balancing ............................................................................. 97

Part Number ............................................................................... 80

Aircraft Brakes............................................................................. 98

Serial Number............................................................................. 80

Internal Expanding-Shoe Brakes............................................... 100

Speed Rating .............................................................................. 80

Expander-Tube Brakes ............................................................. 102

Other Markings ........................................................................... 80

Single-Disk Brakes.................................................................... 106

Retreads ..................................................................................... 80

Multiple-Disk Brakes ................................................................. 108

Tire Inspection On The Aircraft ................................................... 82

Segmented Rotor-Disk Brakes.................................................. 110

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL Table of Contents

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Carbon Composite Brakes........................................................ 112

Auto-Brake System ................................................................... 164

Aircraft Brake Systems ............................................................. 114

Brake Controls .......................................................................... 170

Independent Brake Systems..................................................... 116

Brake Indications....................................................................... 171

Power Assisted Or Boosted Brake Systems............................. 120

Brake Temperature ................................................................... 172

Power Brake System ................................................................ 122

Brake Pressure ......................................................................... 174

Multiple Power Brake-Actuating Systems ................................. 126

Brake Accumulator Pressure .................................................... 174

Emergency Brake System ........................................................ 128 Braking Heat Energy................................................................. 130 Brake Inspection And Service................................................... 132 Typical Brake Malfunctions And Damage ................................. 136 Bleeding Of Brakes................................................................... 138 Skid Protection.......................................................................... 140 Anti-Skid Requirements ............................................................ 142 Mechanical Anti-Skid Systems.................................................. 144 Electronic Anti-Skid Systems .................................................... 148 System Tests ............................................................................ 160 System Maintenance ................................................................ 162 Auto Brakes .............................................................................. 164

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

“The training notes and diagrams are compiled by SriLankan Technical Training and although comprehensive in detail, they are intended for use only with a Course of instruction. When compiled, they are as up to date as possible, and amendments to the training notes and diagrams will NOT be issued”.

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

FUNCTION

LANDING GEAR CLASSIFICATION

The functions of a landing gear are to support an aircraft during ground manoeuvres, dampen vibration, and absorb landing shocks; when required, it also performs the functions of steering and braking.

A landing gear usually takes the form of two or more main undercarriage units in the wings or fuselage, and an auxiliary undercarriage unit at the nose or tail, which carries only a small, proportion of the total load and is used for steering purposes.

These objectives are achieved by many different designs, depending on the type of aircraft to which the landing gear is fitted and the degree of sophistication required.

The Main Landing Gear (or undercarriage) provides the main support of the aircraft on land or water. It may include a combination of wheels, floats, skis, shock absorbing equipment, brakes, retracting mechanism, cowling, fairing, and structural members needed for attachment to the primary structure of the aircraft. The Auxiliary Landing Gear consists of tail or nose landing-wheel installations, skids, outboard pontoons, outrigger wheels, etc., with necessary cowling and reinforcements. Landing gear is classified by the configuration (or wheel arrangement), provisions for cushioning the landing shock, and if the landing gear is retractable or not.

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CATEGORY B1 - MECHANICAL

TRICYCLE-TYPE LANDING GEAR

LANDING GEAR CONFIGURATION (WHEEL ARRANGEMENT)

Tricycle landing gear is characterized by having two main gear assemblies, one on each side of the aircraft located behind the airplane's center of gravity and the nose of the airplane supported by the nose wheel assembly. Control on the ground for small airplanes is provided by steering the nose wheel through connections to the rudder pedals, but large airplanes have hydraulic steering cylinders to control the direction of the nose wheel.

CONVENTIONAL OR TAIL WHEEL-TYPE LANDING GEAR The so-called “conventional", Tail wheel or Tail dragger type landing gear is normally associated with older aircraft and those designed for rough field operations. In this configuration two main wheel assemblies are located on each side of the aircraft, ahead of the aircraft center of gravity, and the tail is supported by a much smaller wheel.

This arrangement places the aircraft fuselage in a level attitude when the aircraft is on the ground. In this attitude the pilot has good forward visibility and the cabin area is level, making it easier for passengers to move inside the cabin when the aircraft is on ground. This configuration also makes the aircraft stable during ground operations and easy to control, which is especially important during takeoff and landing. (Figure B)

Steering on the ground is done by moving the tail wheel through a connection with the rudder pedals. Some aircraft have no provision for steering the tail wheel, but the wheel is locked in line with the fuselage for takeoff and landing and is unlocked, making it free to swivel, for taxiing. Control on the ground is then achieved by differential use of the brakes.

TANDEM LANDING GEAR

This arrangement has the advantage of reduced drag in the air and reduced landing-gear weight. There is some loss of forward visibility for the pilot when maneuvering on the ground due to the aircraft nose-high attitude. This configuration is less stable on the ground and requires more skill when taxiing and during takeoff and landing when compared to a tricycle-geared aircraft as there is a tendency to cause the airplane to ground loop. The pilot must be careful to keep the airplane rolling straight, or the center of gravity will swing around ahead of the wheels, causing the airplane to spin around on the ground. (Figure A)

Rev. 00 Oct 2006

The tandem wheel arrangement is seldom used on civilian aircraft, but some of the military aircraft use it. The main wheels are located in line under the fuselage and the wings are supported by outrigger wheels. (Figure C )

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Fig. A – Conventional Landing gear Fig. B – Tricycle Landing Gear

Fig. C – Tandem Landing Gear

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

RETRACTABLE VS. FIXED With slow, light aircraft, and some larger aircraft on which simplicity is of prime importance, a fixed (non-retractable) landing gear is often fitted; the reduced performance caused by the drag of the landing gear during flight is offset by the simplicity, reduced maintenance and low initial cost. With higher performance aircraft, drag becomes progressively more important, and the landing gear is retracted into the wings or fuselage during flight; there are, however, penalties of increased weight, greater complication and additional maintenance.

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CATEGORY B1 - MECHANICAL

FIXED LANDING GEAR SHOCK ABSORBERS

FIXED LANDING GEAR

There are three main types of fixed landing gear; those which have a spring steel leg, those which employ rubber cord to absorb shocks, and those which have an oleo-pneumatic strut to absorb shocks. Exceptions include aircraft with rubber in compression, spring coil, and liquid spring struts.

Fixed (Non-retractable) landing gear is generally attached to structural members of the airplane with bolts, but it is not actually ‘fixed’, because it must absorb stresses; therefore, the wheels must move up and down while landing or taxiing in order to absorb shocks. The landing gear is often equipped with a fairing where it joins the fuselage or wing to reduce the drag (air resistance). Chafing strips are used to prevent excessive wear between the sections of the fairing, because there is usually some motion between these sections. Wheel pants are often used to cover the wheel and tire to reduce their drag. Fixed landing gear may have bracing, or it may be of the cantilever type without any additional bracing. Fixed landing gear is usually found on small aircraft and aircraft where aerodynamic cleanliness for an efficient cruise configuration is not a major factor.

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Fig. A – Fixed Landing gear

Fig. B – Wheel Pants

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RUBBER CORD

SPRING STEEL LEGS

When rubber cord is used as a shock absorber, the undercarriage is usually in the form of tubular struts, designed and installed so that the landing force is directed against a number of turns of rubber in the form of a grommet or loop.

Spring steel legs are usually employed at the main undercarriage positions. The leg consists of a tube, or strip of tapered spring steel, the upper end being attached by bolts to the fuselage and the lower end terminating in an axle on which the wheel and brake are assembled.

Rubber cord is colour coded to indicate the date of manufacture and the specification to which it conforms, by replacing some of the fibres in the outer cotton covering with coloured threads wound in a spiral.

MAINTENANCE Spring steel undercarriages should be inspected regularly for damage and corrosion. The aircraft should be jacked up periodically, so that all loads are taken off the wheels, and the security of each undercarriage checked by attempting to move it against the restraint of its attachments to the airframe structure. If there are signs of looseness, the bolts should be removed for detailed inspection and the bolt holes should be checked for cracks or fretting. Axle fittings should be similarly inspected, and all nuts and bolts should be tightened to the specified torque.

Rev. 00 Oct 2006

MAINTENANCE The undercarriage should be examined for damage, corrosion, wear or cracks at the pivot points, and bent pivot bolts, and should be lubricated as specified in the approved Maintenance Schedule. The rubber cord should be inspected for chafing, necking, or other deterioration, and it is advisable to replace it if it is more than five years old, regardless of its external condition.

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Fig. A – Spring Steel Main Gear

Fig. B – Rubber Cord or Bungee Type Shock Absorbers

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OPERATION

OLEO-PNEUMATIC SHOCK ABSORBER Some fixed main undercarriages, and most fixed nose undercarriages, are fitted with an oleo-pneumatic shockabsorber strut. The design of individual struts varies considerably, and reference should be made to the appropriate Maintenance Manual for a particular type, but operation and maintenance procedures for a typical design are covered in the following paragraphs. CONSTRUCTION Figure B shows the construction of a simple oleo-pneumatic strut, in this instance a nose undercarriage, which also includes a steering mechanism. The outer cylinder is fixed rigidly to the airframe structure by two mounting brackets, and houses an inner cylinder and a piston assembly, the interior space being partially filled with hydraulic fluid and inflated with compressed gas (air: or nitrogen). The inner cylinder is free to rotate and move up and down within the outer cylinder, but these movements are limited by the torque links, which connect the inner cylinder to the steering collar. The steering collar arms are connected through spring struts to the rudder pedals, and a shimmy damper is attached to the steering collar.

Rev. 00 Oct 2006

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Under static conditions the weight of the aircraft is balanced by the strut gas pressure and the inner cylinder takes up a position approximately midway up its stroke.

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Under compression (e.g. when landing), the strut shortens and fluid is forced through the gap between the piston orifice and the metering rod, this restriction limiting the speed of upward movement of the inner cylinder.

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As the internal volume of the cylinders decreases, the gas pressure rises until it balances the upward force.

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As the upward force decreases, the gas pressure acts as a spring and extends the inner cylinder. The speed of extension is limited by the restricted flow of fluid through the orifice.

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NOTE: On some struts an additional valve is fitted to the piston or inner cylinder, to further restrict the flow of fluid during extension, and prevent violent extension of the strut if upward force is suddenly released, such as when a bounce occurs.

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Normal taxiing bumps are cushioned by the gas pressure and dampened by the limited flow of fluid through the orifice.

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Movement of the rudder pedals turns the nose wheel to facilitate ground manoeuvres, the spring struts being provided to allow for vertical movement of the nose wheel, and prevent shocks from being transmitted through the rudder control system.

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Fig A – Oleo-pneumatic Shock Absorber

Rev. 00 Oct 2006

Fig B

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CATEGORY B1 - MECHANICAL MAINTENANCE

SERVICING STRUTS

Oleo-pneumatic undercarriages should be subjected to inspections similar to those recommended for spring leg and rubber cord types, such as examinations for cracks or damage to mounting structure, corrosion, and wear at pivot points. In addition, the following maintenance is necessary:

When it becomes necessary to check the fluid level in a strut, the following procedure should be carried out:

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1. Jack up aircraft to take the weight off the strut. 2. Remove inflation valve cap and release air pressure completely.

Machined surfaces of the strut inner cylinder should be wiped free of dust or dirt at frequent intervals, to prevent damage to the lower cylinder seals. A lint-free cloth, soaked in the fluid used in the strut, should be used for this purpose.

3. Remove valve housing. 4. Compress strut and check fluid level is at bottom of filler hole; if not, top-up with the approved fluid.

The extension of the inner cylinder, i.e. the length of the visible portion of the inner cylinder, should be checked frequently against the centre of gravity/loading graphs provided in the approved Maintenance Manual.

5. Extend and compress strut several times to expel any trapped air, then repeat step 4. 6. With strut compressed, replace valve housing and inflate strut to specified gas pressure, checking that the leg extends completely.

NOTE: Because of the tightness of the sealing glands in the strut, it may be necessary to rock the aircraft to free the inner cylinder and obtain the true extension. -

The strut should be inspected frequently for fluid leaks. If leaks are due to faulty glands the glands may be replaced, but if they are due to a scored inner cylinder, the strut should be changed.

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Torque links, steering arms, and damper attachments should be checked for security, and for cracks, wear or any other damage.

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All moving parts of the undercarriage should be lubricated on assembly, and at the intervals specified in the approved Maintenance Schedule.

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7. NOTE: It is usually recommended that a new seal be fitted when replacing the valve. 8. Lower aircraft and check that extension of the inner cylinder is in accordance with the tables or graphs supplied by the manufacturer, for the particular aircraft weight and centre of gravity position. (Figure A)

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Fig A – Oleo-pneumatic Shock Absorber Servicing

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MAINTENANCE

SHIMMY DAMPERS

Friction disc dampers should be inspected for security, damage, and wear of the friction material. Piston type dampers will not operate satisfactorily if air is present in the cylinder, and should be inspected frequently for oil leaks; they should be removed at the periods specified in the approved Maintenance Schedule, and the oil level should be checked.

Most nose and tail wheels are fitted with shimmy dampers to prevent rapid oscillation during ground manoeuvres. A simple damper consists of two friction discs, one connected to a fixed part of the undercarriage and the other connected to the oscillating part. The discs are held in contact by spring pressure and resist relative movement between the parts to which they are connected. Piston Type The piston rod is connected to the steering collar and the cylinder attached to a fixed part of the strut. The cylinder is completely filled with fluid, and small holes in the piston allow a restricted flow of fluid when force is applied to the piston rod. Movement of the nose undercarriage is therefore slowed down, and oscillations damped. (Figure A) Vane Type Vane-type dampers are designed with a set of moving vanes and a set of stationary vanes. The moving vanes are mounted on a shaft, which extends outside the housing. When the shaft is turned, the chambers between the vanes change in size, thus forcing hydraulic fluid from one to the other. The fluid must flow through restricting orifices, providing a dampening effect to any rapid movement of the vanes in the housing. The body or housing of the vane-type damper is usually mounted on a stationary part of the nose landing gear, and the shaft level is connected to the turning part. Thus any movement of the wheel right or left causes a movement of the vanes in the shimmy damper. (Figure B)

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Fig. A – Piston type Damper

Fig. C – Piston type Damper Installation

Fig. B – Vane type Damper

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

RETRACTABLE LANDING GEAR The majority of modern transport aircraft, and an increasing number of light aircraft, are fitted with a retractable landing gear, for the purpose of improving aircraft performance. Retraction is normally effected by a hydraulic system, but pneumatic or electrical systems are also used. In some instances power is used for retraction only, extension being effected by gravity and slipstream. Retractable landing gear is also provided with mechanical locks to ensure that each undercarriage is locked securely in the retracted and extended positions; devices to indicate to the crew the position of each undercarriage; and means by which the landing gear can be extended in the event of failure of the power source. In addition, means are provided to prevent retraction with the aircraft on the ground, and to guard against landing with the landing gear retracted. Undercarriage wells are normally sealed by doors for aerodynamic reasons, but one particular aircraft type employs inflatable rubber bags to seal the main undercarriage wells. Retractable undercarriages normally consist of an oleopneumatic shock-absorber strut supported in a trunnion bearing which is fixed to a spar or strengthened box section in the wings or fuselage; the strut is braced longitudinally by drag struts, and laterally by sidestays. In some designs the drag strut or sidestay is in two parts, and hinges about the centre point to provide a means of retraction, while in others, the retraction jack operates on an extension of the shock absorber strut housing. Figure A shows a typical retractable undercarriage unit, which is hydraulically operated in both directions and locked by means of a geometric (overcentre) lock.

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Fig. A – Typical Retractable Undercarriage Unit

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Fig. B –Retractable Gear 19 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

DRAG LINK

LANDING-GEAR COMPONENTS

A drag link (also called a drag strut) is designed to stabilize the landing-gear assembly longitudinally. If the gear retracts forward or aft, the drag link will be hinged in the middle to allow the gear to retract.

Landing-gear assemblies are made up of various components designed to support and stabilize the assembly. The following terms identify many of these components. The terms are presented here as they relate to retractable landing-gear systems. When the terms are used with fixed landing gear, the exact use of the components may vary. Be aware that different manufacturers will occasionally use different terms for the same basic components.

SIDE BRACE LINK A side brace link is designed to stabilize the landing-gear assembly laterally. If the gear retracts sideways, the side brace link is hinged in the middle to allow the gear to retract. This is also called a side strut.

TRUNNION The trunnion is the portion of the landing-gear assembly attached to the airframe. The trunnion is supported at its ends by bearing assemblies, which allow the gear to pivot during retraction and extension. The landing-gear strut extends down from the approximate center of the trunnion.

OVERCENTER LINK An overcenter link is used to apply pressure to the center pivot joint in a drag or side brace link. This prevents the link from pivoting at this joint except when the gear is retracted, thus preventing collapse of the gear during ground operation. The overcenter link is hydraulically retracted to allow gear retraction. This component is also called the jury strut or downlock.

STRUTS The strut is the vertical member of the landing-gear assembly that contains the shock-absorbing mechanism. The top of the strut is attached to, or is an integral part of, the trunnion. The strut forms the cylinder for the air-oleo shock absorber. The strut may also be called the outer cylinder.

UP-LOCKS Landing Gear and L/G Doors are usually maintained locked in the retracted and closed position by means of up-locks. The Uplock is an assembly that is mounted to the aircraft structure and consists of a hook that engages a roller attached to the landing gear or door. The hook itself is maintained in the lock position by a pin or a roller that engages a suitable provision on the hook.

The piston is the moving portion of the air-oleo shock absorber. This unit fits inside the strut, and the bottom of the piston is attached to the axle or other component on which the axle is mounted. Other terms used for the piston are piston rod, piston tube, and inner cylinder.

Rev. 00 Oct 2006

Generally the up-lock engagement is mechanical but the release is either hydraulic or mechanical. (Figure B)

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Fig. A - Retractable Main L/G

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Fig. B - MLG Up-lock

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL TORQUE LINKS

SWIVEL GLAND

Torque links or torque arms, often referred to as a scissors assembly, are two A-frame-type members used to connect the strut cylinder to the piston and axle. The torque links restrict the extension of the piston during gear retraction and hold the wheels and axle in a correctly aligned position in relation to the strut.

A swivel gland is a flexible joint with internal passages that route hydraulic fluid to the wheel brakes and the bungee cylinder of a landing gear. Swivel glands are used where the bend radius is too small or space limitations prevent the use of coiled hydraulic lines. A swivel gland, which is illustrated in Figure B, may be mounted on a bracket secured to the main-gear trunnion fitting. The gland remains stationary and is the terminus of the stationary hydraulic lines. The movable portion of a swivel gland is connected to the hydraulic lines that are routed down the strut of the bungee cylinder and the wheel brakes.

The upper torque link is connected to a clevis fitting on the lower forward side of the shock strut. The lower torque link is connected to a clevis fitting on the axle. The upper and lower torque links are joined together, as shown in the drawing, by a bolt and nut spaced with washers- Each link is fitted with flanged bushings. The gap between the flanged ends of the bushings is taken up by a spacer washer. This spacer establishes the alignment of the wheel.

In Figure C, it can be seen that the gland consists of annular grooves separated from one another by means of slipper rings and packing to isolate the pressure fluid from the return fluid. Thus the gear can be raised and lowered without disturbing the fluid passage to and from the brakes and bungee cylinder.

TRUCK The truck is located on the bottom of the strut piston and has the axles attached to it. A truck is used when wheels are to be placed in tandem (one behind the other) or in a dual-tandem arrangement. The truck (often referred to as a bogie) can tilt fore and aft at the piston connection to allow for changes in aircraft attitude during takeoff and landing and during taxiing.

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Fig B – Swivel Gland Installation

Fig C – Swivel Gland

Fig A

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MAINTENANCE

BOGIE UNDERCARRIAGES

In addition to the lubrication, testing and maintenance of landing gear described in previous paragraphs, particular care and higher standards of workmanship are necessary with bogie undercarriages. Since this type of undercarriage is fitted to heavy aircraft, the materials used are of very high strength, and great care is taken in the manufacture, heat treatment and finish of the components. However, these materials are usually more susceptible to failure from scratches, indentations or corrosion, than materials of lower strength. All servicing functions should, therefore, be carried out with special care, particularly with regard to lubrication, the lack of which could result in corrosion or hydrogen embrittlement. If any surface damage is found during inspection, it should be repaired strictly in accordance with the instructions and limitations specified in the manufacturer's manuals, or, if no adequate guidance is given, in accordance with an approved repair scheme.

On heavy aircraft, the need to spread the weight over a large area has resulted in the use of multiple wheel undercarriages. A typical four-wheeled bogie is illustrated in Figure A, but a larger number of wheels are used on some undercarriages. The undercarriage unit normally consists of a shock-absorber strut, at the lower end of which a bogie beam is pivoted, and the axles are attached to each end of the beam. On some aircraft the rear pair of wheels swivels on the bogie beam, and castors when the nose wheel is turned through a large angle; on others, the upper torque link member is replaced by a pair of hydraulic jacks, which, when nose wheel steering is applied, rotates the whole bogie. Castoring or steering prevents excessive torque on the undercarriage leg and minimises tyre scrubbing during turns. For normal operation, the swivelling pair of wheels is locked in line with the fixed pair. Brake torque at each wheel is transmitted through compensating rods to the shock-absorber strut, thus preventing excessive loads on the bogie beam.

When changing wheel or brake assemblies, the axle should be fitted with a protective sleeve to prevent damage, and the surface and threads should be inspected for damage and corrosion before re-assembling the wheel or brake.

On retractable landing gear a levelling strut, truck positioner or 'hop damper' provides a means of positioning the bogie beam at suitable angles for retraction and landing; this strut is usually connected into the hydraulic system to prevent retraction if the bogie is not at a suitable angle, and combines the functions of hydraulic ram and damper unit.

Rev. 00 Oct 2006

When carrying out retraction or steering tests, operation of the levelling strut and locking/unlocking of the swivelling wheels should be checked in accordance with the appropriate Maintenance Manual.

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CATEGORY B1 - MECHANICAL

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CATEGORY B1 - MECHANICAL HYDRAULIC RETRACTION SYSTEM A hydraulic system for retracting and extending a landing gear normally takes its power from engine driven pumps, alternative systems being available in case of pump failure. On some light aircraft a self-contained 'power-pack' is used, which houses a reservoir and selector valves for the landing gear and flap systems; an electrically driven pump may also be included, or the system may be powered by engine driven pumps. This type of system normally provides for powered retraction of the landing gear, extension being by 'free-fall', with the assistance of spring struts. Figure A on page 27 is a schematic diagram of a simple hydraulic retraction system. The various components shown illustrate operation of this system, but are not intended to represent a typical design; actual components often operate in a different manner, but their purpose is the same.

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PAGE INTENTIONALLY LEFT BLANK

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CATEGORY B1 - MECHANICAL SIMPLE HYDRAULIC RETRACTION SYSTEM RETRACTION

EXTENSION

-

When the landing gear selector is moved to the 'up' position, fluid under pressure is directed to the 'up' line and fluid from the 'down' line is directed back to the hydraulic reservoir.

-

When the landing gear selector is moved to the 'down' position, fluid under pressure is directed to the 'down' line, and fluid from the 'up' line is directed back to the reservoir.

-

Fluid flows to the sequence valves (SV3, SV4), retraction jacks (MJI, MJ2, NJI), main undercarriage down-lock jacks (LJ1, LJ2), and nose undercarriage down-lock (NL2); it cannot pass the sequence valves, which are closed, but operates the retraction jacks and down locks.

-

-

The locks operate first, releasing the landing gear and allowing the retraction jacks to raise each undercarriage, the nose undercarriage engaging its spring-loaded up-lock (NL1) first, because of the jack's smaller size.

Fluid flows to the sequence valves (SV1, SV2), door jacks (DJI, DJ2), door locks (DL1, DL2), nose undercarriage retraction jack (NJ1) and the nose undercarriage up-lock (NL1). The sequence valves are closed, so fluid pressure releases all the door locks and the nose undercarriage uplock, and the doors and nose undercarriage extend, the nose undercarriage engaging its down-lock (NL2) at the end of its travel.

-

When the doors are fully open, the door jacks strike the plungers of their associated sequence valves (SV1, SV2) and open the valves, allowing fluid to flow through the restrictor valves (RI, R2) to the main undercarriage up-locks (ML1, ML2) and retraction jacks (MJI, MJ2).

-

These locks are released, and the retraction jacks lower the main undercarriage fully, the spring-loaded lock-jacks (LJ1, LJ2) imposing a geometric lock on the sidestays. Main undercarriage doors are held open by fluid pressure.

-

At the end of upward travel of the main undercarriage units, a striker on each leg contacts the plunger of its associated sequence valve (SV3, SV4), and opens the valve, allowing fluid to flow to the door jacks (DJI, DJ2).

-

The main undercarriage engages the up-locks (ML1, ML2) and the doors close, engaging locks DL1, DL2. Fluid in the 'down' lines returns to the reservoir, flowing unrestricted through the restrictor valves (RI, R2) and overcoming the small restriction of the spring loading of the sequence valves (SV1, SV2).

NOTE: Restrictor valves are normally fitted to limit the speed of lowering of the main undercarriage units, which are influenced in this direction by gravity. The nose undercarriage often lowers against the slipstream and does not need the protection of a restrictor valve.

NOTE: The nose undercarriage doors are operated mechanically by linkage to the nose shock absorber housing.

Rev. 00 Oct 2006

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Fig A – A Simple Hydraulic Retraction System Rev. 00 Oct 2006

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CATEGORY B1 - MECHANICAL ELECTRICAL RETRACTION SYSTEM An electrical retraction system is often fitted to light aircraft which do not otherwise require the use of a high-pressure fluid system. The main and nose undercarriage units are similar to those used in fluid retraction systems, but push and pull forces on the retraction mechanism are obtained by an electric motor and suitable gearing. Figure A illustrates a typical system, in which a single reversible electric motor provides the power to retract and extend the landing gear. -

The motor operates a screw jack, which provides angular movement to a torque tube; a push-pull rod from the torque tube acts on the drag strut of the nose undercarriage, and cables and rods from the torque tube act on the main undercarriage sidestays, rubber cord being used to assist extension of the main undercarriage units.

-

Down-locks are imposed by over-centering of the drag strut and sidestays during final movement of the operating mechanism, with the assistance of springs.

-

Limit switches on the drag strut and sidestays cut off electrical power and brake the motor when the down-locks have engaged, while a limit switch on the torque tube stops and brakes the motor when the landing gear is fully retracted.

-

Undercarriage doors are operated by linkage to the shock absorber housings.

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CATEGORY B1 - MECHANICAL

ALTERNATE EXTENSION

LANDING GEAR CONTROL Landing gear control includes cockpit controls and control systems that manage following functions of the landing gear. -

Normal retraction/extension

-

Alternate extension

-

Gravity / Emergency extension

Alternate extension is a function available on some aircraft in which the landing gear can be extended with the aid of an alternate hydraulic system / accumulator or the pneumatic system in case of a normal hydraulic system failure. For activation of the alternate extension system the landing gear selector lever usually has to be lowered and an alternate system activation switch or lever should be selected. This will lead to the opening of a valve that will supply hydraulic or pneumatic pressure from an alternate system to extend the landing gear. After an alternate extension, all hydraulically operated gear doors will remain open.

NORMAL RETRACTION / EXTENSION Normal Retraction and Extension is commanded via a Landing Gear selector lever that is located on the right hand side of the main instrument panel on most aircraft. The L/G selector lever can have either two or three selectable positions. -

-

GRAVITY / EMERGENCY EXTENSION Gravity or emergency extension is available to extend the landing gear in case of a normal system failure.

In a system with only two selectable positions, the positions will be identified as L/G ‘Retract’ or ‘Up’ when the lever is raised and L/G ’Extend’ or ‘Down’ when the lever is lowered.

Gravity extension system can be either a mechanical or an electromechanical system. When activated by selecting a lever or switch the hydraulic supply is cut-off landing gear circuit is rearranged to form a run-around-path or connected to return and the landing gear and door up-locks are released.

In a system with three selectable positions an additional centre position that is marked ’Neutral’ will be available for positioning the selector lever after gear retraction to depressurise the landing gear hydraulic circuit.

Landing gear retraction/extension may be mechanically commanded and controlled, electrically commanded and mechanically controlled or electrically commanded and controlled by a system incorporating a control computer where the sequencing is also achieved by the same control computer.

-

In a mechanical system the movement of cables or linkage attached to the emergency selector lever achieves this.

-

In an electromechanical system switches will supply power to electromechanical actuators that will drive a linkage perform the extension.

A gravity extension system is discussed in page 38.

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L/G Selector Lever

Electrical

Electrical or Hydraulic Supply

Position Feedback

L/G Control Module Hydraulic Supply

Fig. A – Mechanically or Electrically commanded Mechanically controlled system

Rev. 00 Oct 2006

Hydraulic Pressure to Landing Gear and L/G Doors

Hydraulic Pressure to Landing Gear and L/G Doors

Fig. B – Electrically commanded Electrically controlled system

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL LANDING GEAR POSITION INDICATION

On some transport aircraft, provision is also made for the crew to examine the locks during flight in the event of failure or incorrect operation of the indicating system. Whichever indicating system is used, it is important that the micro switches are adjusted so that operation of the lights coincides with the corresponding position of the landing gear.

Although the landing gear, when selected down, may be visible from the crew compartment, it is not usually possible to be certain that each undercarriage is securely locked. An electrical indicating system is used to provide a positive indication to the crew of the operation of the locks and of the position of the landing gear. The system usually consists of micro switches, on the up-locks and down-locks, which make or break when the locks operate, and which are connected to a landing gear position indicator on the instrument panel. A mechanical indicator may also be provided, to show that the landing gear is down and locked when the electrical system is inoperative. (Figure A) On British manufactured aircraft, the electrical undercarriage indicating system operates in such a manner that a green light is displayed when the undercarriage is locked down, a red light is displayed when the undercarriage is in transit, and no lights are visible when the undercarriage is locked up: bulbs are usually duplicated to avoid the possibility of false indications as a result of bulb failures. On other aircraft, similar indications may be obtained by the use of magnetic indicators or lights, but on some light aircraft a single green light indicates that all undercarriages are locked down, and an amber light indicates that all undercarriages are locked up.

Rev. 00 Oct 2006

Fig. A – Landing Gear Mechanical Indication

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Fig. B - L/G Position Sensing

Fig. C - L/G Cockpit Indication

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CATEGORY B1 - MECHANICAL

SELECTOR LOCK

SAFETY FEATURES Since the correct operation of the landing gear is of the utmost importance, a number of safety features are included in the retraction system to ensure its correct operation under all conditions.

To prevent inadvertent retraction of the landing gear when the aircraft is resting on its wheels, a safety device is incorporated which prevents movement of the selector lever; mechanical ground locks are also provided for servicing purposes. The safety lock consists of a spring-loaded plunger, which retains the selector in the down position and is released by the operation of a solenoid. Electrical power to the solenoid is controlled by a switch mounted on the shock-absorber strut; when the strut is compressed the switch is open, but as the strut extends after take-off, the switch contacts close and the electrical supply to the solenoid is completed, thus releasing the selector lever lock and allowing the landing gear to be selected up. A means of overriding the lock, such as a separate gated switch to complete the circuit, or a mechanical means of avoiding the locking plunger, is provided for emergency use and for maintenance purposes. (Figure B)

NOSE-WHEEL CENTERING To avoid damage to the airframe structure, the nose wheel must always be aligned in a fore and aft direction during retraction, and a number of methods are used to ensure that this happens automatically. One method utilizes a cam and cam track between the inner and outer cylinders on the shock absorber. The cam is fixed to the top of the inner cylinder, and the track to the bottom of the outer cylinder. As the strut extends under internal gas pressure after take-off, the cam engages the track and centres the nose undercarriage before it retracts. A second method is the use of a peg located at the top of the shockabsorber strut, which engages a track fixed to the strut housing or in the wheel bay, and this device centres the undercarriage as it retracts. (Figure A)

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CATEGORY B1 - MECHANICAL

Selector Lever Lock Manual Release

Fig. A - Nose Landing Gear Centering Mechanism

Rev. 00 Oct 2006

37 For Training Purposes Only

Fig. B - L/G Selector Incorporating Safety Lock

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL WARNING DEVICES

EMERGENCY EXTENSION

To guard against landing with the landing gear retracted or unlocked, a warning horn is incorporated in the system and connected to a throttle-operated switch. If one or more throttle levers are less than approximately one third open, as would be the case during approach to land, the horn sounds and the red warning lamp illuminates if the landing gear is in any position other than down and locked. A horn isolation switch is often provided to allow certain flight exercises and ground servicing operations to be carried out without hindrance.

A means of extending the landing gear and locking it in the down position is provided to cater for the eventuality of main system failure. On some aircraft the up-locks are released manually or by means of an emergency pneumatic system; the landing gear 'free-falls' under its own weight and the down locks are engaged by spring jacks. On other aircraft the landing gear is extended by an emergency pressure system which often uses alternative pipelines to the jacks. Pressure for the emergency system may be supplied by a hydraulic accumulator, a hand pump, a pneumatic storage cylinder, or an electrically powered pump.

Rev. 00 Oct 2006

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Fig. A - Landing Gear Gravity Extension System

Rev. 00 Oct 2006

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL ADDITIONAL FEATURES AIR / GROUND SENSING

LANDING GEAR SHORTENING

Air/Ground or Flight/Ground sensing system is used for declaring aircraft ‘on ground’ or ‘in air’ (airborne) status.

Landing gear shortening is one option that is available to designers when the landing gear cannot be accommodated in the available wheel well, when retracted, because of excessive height of the shock absorber when it is extended after lift-off.

Aircraft ‘on ground’ or ‘in air’ status is essential for activation/deactivation of certain functions in the landing gear and other aircraft systems (such as air conditioning, pressurization, interior compartment lights, etc.).

Such a landing gear will incorporate a landing gear shortening mechanism to shorten the landing gear during retraction. This mechanism usually does not require any additional power but is designed to be actuated by the movement of the gear during retraction.

Aircraft ‘on ground’ or ‘weight on wheels’ condition is monitored by using mechanical means in older generations aircraft and in later model aircraft using ‘squat’ switches (proximity sensors).

Safety systems are in-built to prevent retraction in case of a malfunctioning shortening mechanism to prevent gear or structural damage. Also warnings may get activated to inform the pilot of a malfunctioning shortening mechanism during gear extension (lengthening fault).

The aircraft is considered to be on ground when the landing gear is compressed under the aircraft weight, which is usually detected by the angle of the torque links. On aircraft equipped with a mechanical sensing system one end of the mechanical linkage is connected to the torque links with the other end connected to the system requiring air/ground status On later model aircraft equipped with electrical sensing system the proximity sensors and their targets are installed on the torque links and the signal is either send to a switch control unit or the landing gear control unit. The other systems receive the aircraft air/ground status from the above control units electrically. Suitably equipped aircraft rely on this air/ground sensing to prevent landing gear lever selection to retract until the aircraft is in air.

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Fig A – Air/Ground Sensors

Fig B – L/G Shortening

Rev. 00 Oct 2006

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL MAINTENANCE The landing gear performs an important function and every care should be taken to ensure that the instructions for its inspection and maintenance contained in the relevant Maintenance Manual and approved Maintenance Schedule are correctly carried out.

-

Care should be taken to ensure that the fluids used for topping up the hydraulic system or shock-absorber strut are perfectly clean. Funnels and containers must be kept clean and should be rinsed in clean fluid before use.

GENERAL PRECAUTIONS

-

Fluid bled or drained from the system, or used for flushing, must be discarded.

The following precautions are relevant to most types of landing gear, and will help ensure the safety of personnel and correct operation of the system.

-

Care should be taken to prevent spillage of fluid, which may have a detrimental effect on paint, rubber, cable insulation, etc. Some fluids are also irritant to the skin and eyes.

-

Air pressure should be released slowly, particularly in confined spaces.

-

Ground equipment used for replenishing fluids, or for providing hydraulic power or air pressure, should be kept scrupulously clean and should be serviced at stipulated intervals.

-

Ground locks should be fitted whenever the aircraft is out of service, and the appropriate circuit breakers tripped, or fuses removed, when work is carried out on the system.

-

Replacement or adjustment of components in the retraction system should be followed by a retraction test.

-

Components should never be removed while the system is under pressure, i.e. by hydraulic accumulator or pneumatic supply bottle.

-

When components are removed, the open pipelines should be properly blanked; rags or masking tape must not be used for this purpose. .

Unless otherwise specified, components should usually be installed using the appropriate lubricant or anti-seize compound on mating surfaces.

-

Only the recommended lubricants and fluids should be used, and any tests necessary should be carried out strictly in accordance with the relevant Maintenance Manual.

-

-

Rev. 00 Oct 2006

New components should be inspected for cleanliness before installation, and it is usually recommended that components containing fluid should be completely filled before installation, or primed and bled after installation.

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Fig A - Landing Gear Ground Lock Installation

Rev. 00 Oct 2006

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL ROUTINE SERVICING

COMPONENT INSTALLATION

-

At the periods specified in the approved Maintenance Schedule, the landing gear should be lubricated and the relevant inspections carried out.

-

The retraction mechanism should be inspected for security, damage, wear of moving parts, fluid leaks and chafing of pipelines and electrical cables.

-

Doors and wheel bays should be inspected for damage resulting from debris thrown up by the wheels, or witness marks from the tyres indicating faulty adjustment or damaged linkage.

-

Minor damage may usually be blended out and the part reprotected as appropriate, but cracks, kinks in pipelines, or wear beyond the limits specified in the Maintenance Manual are not acceptable.

Whenever a new component is installed in the retraction system, it should be carefully adjusted to prevent physical damage and ensure correct operation. A common method of adjusting components and linkage after installation is to jack-up the aircraft, install ground locks on the undercarriages not being worked on, make the system electrically safe, and operate the individual retraction jack using a hand pump rig. This ensures slow, controlled operation, and allows individual adjustments to be made to the mechanism in accordance with measurements quoted in the relevant Maintenance Manual. Mter adjustment, the system should be reconnected and bled, and retraction tests carried out.

-

Some leakage from the components of a pneumatic system is usually permissible, since the operating medium is replaceable, but serious leaks could affect operation of the system. Leakage from a hydraulic system may sometimes be corrected by cleaning and remaking a connection, but a component with a persistent leak should be replaced.

Rev. 00 Oct 2006

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL RETRACTION TESTS Retraction tests should be carried out following replacement of a faulty component, whenever incorrect operation is reported or suspected, and after a hard or overweight landing. The sequence of operations will depend on the particular installation and type of retraction system concerned, and full details should be obtained from the relevant Maintenance Manual. The following procedure is applicable to most retractable landing gears.

-

Locks, switches, warning devices and mechanical indicators for correct operation.

-

Freedom from fouling during retraction or extension, especially of flexible pipes.

-

General smooth operation of the mechanism.

9. Raise the aircraft so that the wheels are clear of the ground, and lock the lifting jacks. Ensure that no ground equipment or personnel are in the vicinity of the undercarriages and doors.

13. Finally tighten and lock any equipment installed immediately prior to the test.

12. Remove servicing equipment, lower aircraft and fit ground locks.

NOTE: Retraction tests following initial assembly, replacements or significant adjustments, should be carried out with the wheel doors disconnected from their operating struts, and, if necessary, the sequence valves operated by hand; loose operating rods should be guided clear of structure. This procedure will permit direct inspection for clearance and alignment, and will also permit adjustment of mechanical stops, sequence contact points, up and down locks, and over-centre linkage.

NOTE: In some aircraft the arc described by the wheels during retraction brings them nearer to the ground, and additional ground clearance must be allowed in these instances. 10. Connect electrical power and external hydraulic or pneumatic servicing equipment as appropriate. 11. Carry out several retractions and extensions, initially at low power to ensure slow operation, and using both the normal and emergency systems, and check the following: -

Undercarriages for proper operation.

-

Doors for correct operation and fit.

-

Clearance in the wheel bays with the landing gear retracted, making due allowance for the effects of centrifugal force on tyre diameter.

-

Linkage for correct operation and adjustment.

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CATEGORY B1 - MECHANICAL

POWERED STEERING Light aircraft generally employ a simple steering system, in which the nose wheel is mechanically linked to the rudder pedals. Larger aircraft require powered steering arrangements, in which the nose wheel is turned by hydraulic, pneumatic, or electrical power. A powered steering system generally includes a cockpit steering wheel or tiller, a control valve, steering cylinders to actuate the nose undercarriage, a follow-up device to hold the nose wheel at the correct angle, and a power source.

Rev. 00 Oct 2006

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Fig A - Powered Steering System

Rev. 00 Oct 2006

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CATEGORY B1 - MECHANICAL

CASTORING

HYDRAULIC STEERING SYSTEM

Whenever the control valve is in its neutral position, fluid is free to flow between the steering jacks, thus allowing the aircraft to be towed, or the nose wheel to return to the central position after a turn has been initiated with the steering wheel. Angular movement of the nose wheel during towing will be transmitted through the follow-up linkage to the steering wheel. Some form of quick-release pin is often provided to enable the steering jacks to be disconnected so that the nose wheel may be turned through large angles during ground servicing.

Main operating pressure is derived from the under carriage 'down' line, and a limited emergency supply is provided by a hydraulic accumulator. In the system shown in Page Figure A, hydraulic pressure passes through a changeover valve, which ensures that the steering system is only in operation when the nose undercarriage is down. STEERING OPERATION Pressure is directed through the control valve to the steering jacks, which retract or extend to rotate the nose shock-absorber strut within its housing. Movement of the steering wheel is transmitted through mechanical linkage to the control valve, in accordance with the amount and direction of turn required. Follow-up linkage from the nose undercarriage gradually resets the control valve as the nose wheel turns, and when the selected angle is reached a hydraulic lock is formed between the control valve and the steering jacks, preventing further movement. When the steering wheel is released, the control valve returns to neutral under the action of its centering springs, and the nose wheel is free to castor.

DAMPING Restrictors in the pipelines between the control valve and the steering jacks, provide damping for the nose undercarriage.

An inner cylinder in each steering jack is connected to the landing gear 'up' line and is supplied with fluid under pressure when the landing gear is selected up. The steering jacks extend equally to centralise the nose wheel before pressure is applied to the nose retraction jack, and the by-pass valve allows fluid from the steering jacks to flow to the return line.

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CATEGORY B1 - MECHANICAL

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CATEGORY B1 - MECHANICAL

RUDDER PEDAL STEERING

STEERING CONTROL Steering control involve systems used to transmit steering commands from the cockpit to the steering control valve that will meter the hydraulic fluid to the steering actuators.

During takeoff and landing the aircraft can be steered using commands from the rudder pedals. Rudder pedal authority for steering is limited compared to hand wheel input.

Steering commands can be generated by

-

On older aircraft with mechanical controls, rudder pedal steering commands are parsed with hand wheel commands by a mechanical mixer and sent to the servo using the same mechanical control path.

-

On later generation aircraft with electronic steering, rudder pedal input is sent to the steering computer in an electronic format. In this system, rudder pedal authority is progressively reduced with increasing speed.

-

Steering hand wheels or tillers

-

Rudder pedals

-

Autopilot (on latest model aircraft)

STEERING HAND WHEELS Steering hand wheels are provided for the Captain or both members of the flight crew for steering the aircraft on ground during taxiing. Maximum authority of steering is available through the steering hand wheels but in latest model aircraft the authority is steadily reduced with forward speed. -

-

Rev. 00 Oct 2006

AUTOPILOT STEERING On later model aircraft, steering commands are sent from autopilot computers to the steering control computer to control the aircraft on ground during an autoland with automatic rollout.

On older generation aircraft the steering commands from the hand wheels are transmitted to the steering control servo via a mechanical linkage that usually comprises of cables. A feedback is provided from the gear to the servo to hold the gear in the selected position until a new command is received. (Figure A on page 47)

STEERING DEACTIVATION FOR TOWING To facilitate aircraft towing, the steering system can usually be deactivated by installing a steering by-pass pin that selects the actuators to by-pass mode. On aircraft equipped with mechanical controls, the steering bypass pin mechanically locks the control path to the servo control when installed.

On later model aircraft, steering commands from the hand wheels are transmitted to the steering control servo as an electrical signal via a steering control computer that computes the necessary commands. The position feedback from the gear is sent to the computer in such a system for holding the gear in the selected position. (Figure A)

On aircraft with electronic steering, the pin is used to lock a lever on a control box in a steering deactivated position that sends a signal to the computer to select a bypass mode on the servo.

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Figure B - Mechanical Steering De-activation

Autopilot Computers

Figure C - Electronic Steering De-activation

Figure A- A Modern steering control system

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CATEGORY B1 - MECHANICAL MAINTENANCE The lubrication and inspection requirements of the steering system are broadly similar to those detailed in for retractable landing gear. Installation and adjustment of the mechanical linkage, and functional testing of the system are described in the following paragraphs. MECHANICAL LINKAGE Proper adjustment of the mechanical linkage is most important, since slackness or faulty installation could lead to incorrect operation of the steering system. To facilitate installation of components, rigging pins are usually inserted through jig-drilled holes in the steering wheel, drum assembly and follow-up linkage in order to fix their positions. The nose wheel can then be centralised, and the cables and rods fitted and adjusted, accordingly. Cables should be tensioned using a tension-meter, and rods adjusted so that the connecting pins and bolts can be easily fitted. When new pulleys or cables are fitted, it is usually recommended that they are 'bedded-in' by operating the steering wheel a number of times over its full range of travel; cables should then be re-tensioned.

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CATEGORY B1 - MECHANICAL FUNCTIONAL TEST The following test is applicable to the system described earlier and contains the basic essentials for tests on similar steering systems. The hydraulic installation on a particular aircraft may necessitate additional operations, and these will be fully described in the appropriate Maintenance Manual.

8. Carry out further retractions to check that the steering is only operative when the nose undercarriage is down.

1. Ensure that the shock strut is correctly serviced.

9. NOTE: Operations (6), (7), and (8) could lead to extensive damage if malfunction occurs, and should be performed with the test rig adjusted to give a slow rate of operation of the retraction system.

2. Jack the aircraft so that the wheels are clear of the ground and ensure that no ground equipment or personnel are in the vicinity of the landing gear.

10. Check that the stand-by accumulator is correctly charged with air pressure and operate the test rig to pressurise the accumulator.

3. Depressurise the main hydraulic system and check that the nose wheel has freedom of movement over the full castoring range.

11. Select stand-by steering and check that the nose wheel can be steered satisfactorily. This check may involve a specified number of turns before the accumulator is exhausted or the stand-by system low pressure warning lights illuminate.

4. Connect a hydraulic test rig and ground electrical power, and set controls and switches for normal hydraulic operation.

12. Set the stand-by selector to off, and disconnect the test rig and external electrical power.

5. Operate the steering wheel over its full range, and check that the nose wheel follows smoothly and stops at selected positions.

13. Lower the aircraft and finally lock any components installed prior to the test.

6. Set the nose wheel a few degrees to one side and select the landing gear up, checking that the nose wheel centres before the down-lock breaks. 7. Lower landing gear and repeat operation (6) with the nose wheel displaced inthe opposite direction.

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CATEGORY B1 - MECHANICAL WHEEL ALIGNMENT Alignment of the main gear wheels is very important in that misalignment adversely affects landing and take-off, roll characteristics, tire wear, and steering during ground operations. Severe misalignment can cause malfunction and failure of some of the major components of the landing gear system.

Alignment check on an aircraft equipped with a spring steel landing gear is difficult carry out since the landing alignment changes with the changes in aircraft weight. Therefore prior to an alignment check, roll each wheel onto a pair of aluminum plates with grease between them. If the aircraft is rocked back and forth a bit before the measurement is taken, the greased plates will allow the wheels to assume their true position of alignment. (Figure D)

Alignment consists of checking and adjusting the toe and the camber of the gear. The aircraft maintenance manual normally specifies the amount of toe and camber the landing gear should have.

Wheel alignment of oleo-equipped landing gear is adjusted by means of shim washers installed between the torque links at the joint between the upper and lower torque links. (Figure F)

Toe is the amount wheels are angled from the horizontal axis. An aircraft's wheels are configured in a toe-in position if lines drawn through the center of the two wheels, perpendicular the axles, cross ahead of the wheels or toe-out if the lines cross behind the wheels. (Figure B)

For spring steel landing gear alignment is adjusted by the use of tapered shim plated installed between the landing gear leg and the axle. (Figure E)

Camber is the amount wheels are tilted from the vertical. If the top of the wheel leans outward, the camber is positive. If it leans inward, the camber is negative. In order to measure toe-in, a carpenter's square is held against a straightedge placed across the front of the main wheels. The straightedge should be perpendicular to the longitudinal axis of the aircraft. If this is correct, then the distance between the blade of the carpenter's square and the front and rear flanges of the wheel will indicate toe-in or toe-out. (Figure C) Camber can be measured by using a spirit level protractor held vertically against the outboard flanges of the wheel. (Figure A)

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Fig. C

Fig. D

Fig. F

Fig. E

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AIRCRAFT WHEELS INTRODUCTION Wheels and Tires provide the means by which the aircraft can move and maneuver while maintaining contact with ground. In addition wheels and tires cushion some of the shock loads encountered during landing and taxiing. Aircraft wheels operate in some of the harshest environments known and the extremely high loading and landing speeds of the transport category aircraft exacerbate the effects further. Therefore a high standard of maintenance and inspection is essential at all times to ensure the continued serviceability of the tires. In all cases where doubt exists regarding the condition of aircraft tires or wheels, the wheel should be changed and the manufacturer's representative should be consulted.

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CONSTRUCTION Heat generated by braking action is dissipated by radiation and conduction through the wheel and tire, and every effort is made to keep heat transference to a minimum. Wheels are designed to permit optimum ventilation, and cylindrical stainless steel heat shields may be installed around the brake unit. On some aircraft, an electric motor mounted within the axle or the brake housing, drive fans, which provides forced cooling of the brake. (Figure C)

There are three main types of aircraft wheels, known as well-base, detachable flange and split hub. (Figure A) Tires used on many of the small aircraft are flexible enough that they could be forced over the wheel rim with tire tools in much the same way we force tires on automobile wheels today. Therefore these aircraft are usually equipped with well-base wheels, which are of single piece construction.

To prevent the danger of tire explosion, the main wheels of many modern aircraft are fitted with fusible plugs which melt at a predetermined temperature (approximately 150°C), allowing a piston to be blown out of the plug bore and thus deflating the tire.

But modem aircraft tires are normally so stiff they cannot be forced over the rims, and as a result, almost all modem wheels are of two-piece construction.

Some aircraft wheels are also fitted with a pressure relief valve, the purpose of which is to prevent over-inflation of the tire.

The development of tubeless tires promoted the development of two-piece wheels that are split in the center and made airtight with an a-ring seal placed between the two halves. Today, this form of wheel is the most popular for all sizes of aircraft, from the small trainer up through the large jet transports. (Figure B)

In general wheels are mounted on ball or roller bearings, which fit directly on to the axle, or on to a bearing sleeve, which is keyed to the axle. Some nose wheels are mounted rigidly on to a ‘live’ axle, which itself rotates within bearings in the nose wheel leg.

Since aircraft wheels must be both lightweight and strong, most of them are made of either aluminum alloy or magnesium alloy and, depending upon their strength requirements, may be either cast or forged.

Nose wheels for aircraft are often smaller in diameter and width than the main wheels, and only a few aircraft use brakes on the nose wheel but in some instances all wheels on an aircraft are interchangeable for ease of provisioning.

The bead seat area is the most critical part of a wheel, and to increase its strength against the surface tensile loads applied by the tire, these areas are usually rolled to pre-stress their surface with a compressive stress.

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If a nose wheel is not designed to accept a brake assembly, a fusible plug is not used, but a pressure relief valve may still be used.

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Fig. A

Fig. C

Fig. B

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CATEGORY B1 - MECHANICAL SPLIT HUB WHEEL CONSTRUCTION INBOARD WHEEL HALF

OUTBOARD WHEEL HALF

This is the half of a two-piece wheel into which the brake fits. Rotating brake disks are driven either by tangs on the disk which ride in steel-reinforced keyways, or by steel keys bolted inside the wheel which mate with slots in the periphery of the disk.

This half of the wheel bolts to the inboard half, and it also holds a shrunk-in bearing cup in which a tapered bearing cone rides. A seal protects the roller and bearing surfaces from water and dirt, as well as retaining the lubricant in the bearing.

Stainless steel heat shields are installed between the drive keys to form a barrier around the brake unit to reduce the amount of heat transferred to the wheel. (Figure A)

A cap is held in place with a retaining ring to cover the end of the axle shaft and the bearing. If the aircraft is equipped with an antiskid system, this cap is fitted with a bracket to drive the wheel-speed sensor, which is mounted in the axle.

A polished-steel bearing cup is shrunk into the bearing cavity of the wheel, and a tapered bearing cone slides over the landing gear axle to support the wheel. A grease seal covers the bearing to hold grease in the bearing and prevent any dirt or water getting to its surfaces.

An inflation valve in the outboard wheel half allows air to be put into a tubeless tire, or if a tube-type tire is installed, a hole in this half of the wheel is provided for the valve stem of the tube to stick through. Wheels used on aircraft equipped with tire pressure remote indication systems are equipped with an adaptor similar to the valve stem for the installation of the tire pressure sensing transducer.

One or more fusible plugs are installed in the inboard half of the main wheels of jet aircraft to release the air from the tire in the event of an extreme overheat condition, such as might exist from the heavy braking required during an aborted takeoff.

Provision may also be available for mounting a fairing cover on the wheel for aircraft equipped with retractable landing gear without a corresponding landing gear door. A cutout is available on the fairing to provide access to the inflation valve. Wheels installed on aircraft equipped with brake cooling fans require the installation of a cover that acts as an air guide to properly direct the air to the brake unit.

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Fig. B

Fig. A

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WHEEL INSPECTION – ON THE AIRCRAFT CHECK FOR AXLE NUT TORQUE

CHECK FOR DAMAGES

If too little torque is used on the axle nut, it is possible for the bearing cup to become loose and spin, enlarging its hole and requiring a rather expensive repair to the wheel. On the other hand, if the torque is too high, the bearing can be damaged because the lubricant will be forced out from between the mating surfaces. The amount of torque required varies with the installation, and the procedure used for installing and securing the axle nut must be that recommended by the airframe manufacturer.

Check for any damages to the wheel such as cracks, nicks, visible corrosion, and for possible heat damage due to the high temperatures developed during severe brake applications. During inspection pay special attention to broken or loose wheel tie bolts, because wheel tie bolt failure is common on heavily loaded aircraft such as freighters. Also check for loose or missing items such as hubcap and balance weight fasteners. CHECK FOR PROPER INSTALLATION Visually check the wheel for proper installation and clearances between wheel and the brake unit or adjacent structure. It is possible with some types of wheel and brake assemblies that the wheel can be installed with the disk drive tangs between the drive slots, rather than mating with the slot itself. Be sure when inspecting the wheel that the brake is correctly installed and everything is in its proper place.

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WHEEL INSPECTION – OFF THE AIRCRAFT LOOSEN THE TIRE FROM THE RIM Before a wheel can be inspected, the tire must be removed, and before loosening the wheel half retaining bolts, be sure the tire is completely deflated. For safety, let the air out of the tire by using a deflator cap screwed onto the valve and, after most of the air is out, removeand discard the valve core. After all of the air is out, break the bead of the tire from the wheel by applying an even pressure to the tire as close to the wheel as you can get the tool. Never use a screwdriver or any type of tire tool to pry the bead away from the rim, as it is easy to nick or damage the soft wheel in the critical bead area, and any damage here will cause a stress concentration, which can lead to wheel failure. DISASSEMBLE THE WHEEL Place the wheel on a clean, flat surface and remove the bearing seals and cones from both wheel halves. Remove the nuts from the wheel bolts so the halves can be separated.

Fig. A

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CATEGORY B1 - MECHANICAL CLEAN THE WHEEL ASSEMBLEY

INSPECT THE WHEEL HALVES

Use Stoddard solvent or some similar dry cleaning fluid to remove any grease or dirt from the wheel. A soft bristle brush will aid in removing stubborn deposits, but don't use any kind of scraper that will remove any of the protective finish from the wheel. After all of the parts have been cleaned, dry them with compressed air.

The most difficult area of an aircraft wheel to inspect is the bead seat region. This area, which is highly stressed by the inflated tire, can be distorted or cracked by a hard landing. But, when all of the forces are removed and the tire dismounted, these cracks may close up so tightly, especially on forged wheels that no penetrant can get in. Therefore, eddy current inspection should be used for the bead seat area, as it is better able to detect failures such as tight cracks. (Figure A)

CLEAN THE BEARINGS Use clean solvent to wash the wheel bearings. Soak them to soften the grease and any hardened deposits in the bearings, and then brush them with a soft bristle brush to remove all of the residue.

Dye penetrant inspection procedures can be used to check for cracks in the key slot areas, as cracks there have no tendency to close up. Examine the entire wheel for indication of corrosion, which is most likely to form at any place where moisture is held trapped against the surface of the wheel. Such as the rim area where moisture could be trapped between the tire and the wheel and the surface of the wheel that is inside the tire when it is mounted. (Figure B)

Dry the bearings by blowing them out with low-pressure dry compressed air. Don't Spin The Bearings As You Dry Them, since rotating dry metal against dry metal will surely damage both the rollers and the races. Bearings should never be cleaned with steam, as the heat and excess oxygen will cause a premature breakdown of the bearing surface.

Any corrosion pits that are found must be completely dressed out, without removing more metal than the manufacturer's service manual allows. After all traces of the damage have been removed, treat the surface of the metal to prevent further corrosion. A good visual inspection is one of the best ways to locate a defect in a part. Examine all of the suspect areas in the wheel magnifying glass. Check all of the dimensions specified in the overhaul manual to be sure that there are no parts worn to the extent that they will require repair or replacement.

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Fig. A

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Fig. B

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CATEGORY B1 - MECHANICAL INSPECT THE WHEEL BOLTS

CHECK THE FUSIBLE PLUGS

Inspect the bolts by magnetic particle inspection. Pay particular attention to the junction of the head and the shank and to the end of the threaded area. These two locations are where the cross-sectional area of the shank changes and are most likely locations for cracks to form.

Carefully examine the condition of the fusible plugs in the wheel and that none of them show any sign of core melting. All the plugs must be replaced even if only one indicates any deformation. (Figure C) CHECK THE BALANCE WEIGHTS

If self locking nuts are used to secure the bolts their self-locking feature should be checked prior to them being reused. Selflocking feature is considered serviceable if not more than one and a half threads are visible on the bolt section protruding from the nut after engaging the nut on the bolt by hand.

Almost all wheels having a diameter of more than 10” are statically balanced when they are built, and they have balance weights attached. If the weights have been removed for any reason, be sure that they are put back in their original position. The final balancing of the wheel is done after the tire is mounted, and the weights are usually installed around the outside of the rim of the wheel or at the wheel bolt circle. (Figure D)

INSPECT THE KEY AND KEY SCREWS The disk drive keys are subject to some of the most severe forces acting on a wheel, as they try to rotate the disks against the friction in the brake. No play can be tolerated between the keys and the wheel half, so they must be checked for looseness, for cracks, and for excessive wear. (Figure A & D) Each of the key attachment screws is staked to prevent it becoming loose in service, and these stakes must be checked. INSPECT THE HEAT SHIELD Segments of the heat shield should be checked for condition and security of installation. There have been many occurrences of displaced heat shield segments coming into contact with the brake unit rotors and causing damages to the brake unit and the wheel. These segments either become loose or warp due to high temperatures and the failure will become evident only during the removal of the wheel. (Figure B)

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Fig. B

Fig. C

Fig. E

Fig. A

Fig. D

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CATEGORY B1 - MECHANICAL INSPECT THE BEARINGS

The bearing cup should be carefully inspected for any signs of water marks, discoloration, rust, or brinell marks, which are shallow, smooth depressions made by the rollers forced against the cup by excessive pressure. Any of these is a cause for rejection. (Figure C)

If a bearing is difficult to remove from the axle shaft, it should be removed with a special puller, but it should never be driven from the shaft with any form of drift. Bearings that have been difficult to remove from the shaft often have indications of galling on their inner bore, and this galling is generally cause for rejection of the bearing cone.

If it is necessary to remove the bearing cup, heat the wheel in boiling water or in an oven. When the wheel is at the proper temperature, tap the cup out of its cavity with a fiber drift. Install a new cup by again heating the wheel and chilling the cup with dry ice. Coat the outside of the cup with zinc chromate primer and tap it in place with a fiber drift. When heating the wheel, be extremely careful that it is not overheated, as this will impair its heat treatment.

Water stains in a bearing may not look bad, but they are an indication of intergranular corrosion in the surface of the rollers or the races, and any bearing showing these marks should be rejected. (Figure A) Any damage to the large end of the rollers is a reason for rejection, but minor flattening at the small end will usually cause no problem, since this end does not provide rolling contact.

GREASE THE BEARINGS

Any spalling, which is a slight chipping of the rolling surface of the bearing, is a reason to reject the bearing. Also, any grooves on the roller surface that are deep enough to feel with your fingernail will require replacement of the bearing.

Use the grease specified by the aircraft manufacturer. If equipment is available for pressure-packing the bearings, it should be used, as it uniformly covers all of the rollers. But if the bearings must be packed by hand, be sure to work the grease up around each of the individual rollers until they are all uniformly covered, and check to be sure that the inner cone is completely covered. After the bearing is thoroughly greased, wrap the inner cone and roller assembly in clean waxed paper and store it until it is ready for installation. Spread a light film of grease on the bearing cups in the wheel halves, and protect them from dirt or damage until the wheel is reassembled.

Any indication of bearing overheating, which shows up as discoloration of the rollers, or of dry operation, which is indicated by rust on the rolling surface, is cause for rejection of the bearing. (Figure B) The bearing cages hold the rollers spaced away from each other and aligned on the races, and if there are any bends or distortion of the cage, it is likely that there is some hidden damage in the bearing. Since the cage is not a wearing surface, if it shows any signs of wear, it is an indication of bearing damage, therefore the bearing should be replaced.

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Fig. A

Fig. B

Fig. C

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CATEGORY B1 - MECHANICAL SIZE

AIRCRAFT TIRES AND TUBES

A tire size could be shown on a tire as follows: (ref. Figure C)

TIRE CLASSIFICATION

-

A only (for example, 44”), or

Aircraft tires are classified according to their type, size, and ply rating and whether they use tubes or are tubeless.

-

B-C only (for example, 8.50-10), or

-

AxB only (for example, 49x17), or

The United States Tire and Rim Association has established nine types of aircraft tires, but only three of these types are of primary concern today.

-

AxB-C (for example, 49x19.0-20) or

-

AxBRC (for example, 46x17R20) for radial-ply tires.

The size specifications shown above are the standards used for tires of new design.

TYPE Type III is the most popular low-pressure tire found today on piston-powered aircraft. The section width is relatively wide in relation to the bead diameter. This allows lower inflation pressure for improved cushioning and floatation. (Figure A)

A tire size could have a prefix (for example, H49x19.0-22). The prefix tells you that the wheel rim on which the tire is to be installed must have a specified width between its flanges (expressed as a percentage of the nominal width of the tire; also known as width ratio). Available prefixes are shown on the table in figure D.

Type VII extra-high pressure tires are the standard for jet aircraft. They have exceptionally high load-carrying ability. Type VIII tires are used for high performance jet aircraft with their extremely high takeoff speeds. They use extra high inflation pressure and have a low profile. Because of their strength and rigidity, such tires, whether tubed or tubeless, are normally fitted only to divided or detachable-flange wheels. (Figure C)

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Dimensions A, B and C could be inches or millimeters. For example, for a tire size specified as 750x230-15, dimension A = 750 mm, dimension B = 230 mm, and dimension C = 15 in. An inner tube size is specified the same as the size of its related tire.

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Fig. C Fig. A

Fig. B

Prefix Width Ratio Bead Ledge C 50%-60% 150 B 50%-60% 150 H 60%-70% 50 No Prefix 70% - Higher 50 Width Ratio – Section Width / Rim Width, Expressed as a Percentage Bead Ledge – Angle at the Base of the Bead

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CATEGORY B1 - MECHANICAL PLY RATING

TUBE OR TUBELESS

In the past, tires were rated for strength by the number of fabric plies used in the construction of their carcass. But newer materials have comparatively so much greater strength that fewer actual layers of material are needed to get the same strength. Therefore, now tires are given a ply rating, rather than specifying the actual number of layers of fabric material used in the carcass.

Aircraft tires are manufactured as both tube type and tubeless, with the basic difference between the two being the inner liner. Tubeless tires have an inner liner that is about one-tenth of an inch thick that serves as a container for the air, while tube-type tires have no such liner, but are somewhat smoother on the inside so the tube will not be damaged by chafing against the inside of the tire.

The ply rating of a tire relates to its maximum static load and its inflation pressure.

Some of the advantages derived from the use of tubeless tires include about 7½% saving in weight compared with using a tire and tube, a reduction in permeability losses, cooler running by about 100C, less danger of deflation due to puncture, and the elimination of tube troubles. Tubeless tires are identified by the word TUBELESS on their sidewall, and the lack of identification signifies that a tube should be used in the tire. Because it is necessary to keep the bead areas in good condition, tubeless aircraft tires are not fitted to well-base wheels.

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CARCASS (BIAS TIRE)

TIRE CONSTRUCTION

The carcass of the tire is made up of layers or plies of rubbercoated nylon-cord fabric. This fabric is cut into strips on the bias, meaning that the cords of the fabric run at an angle of about 450 to the length of the strip. These strips extend completely across the tire and lap back over the beads to form the ply turnups. Each successive ply of the fabric is put on in such a way that the cords cross each other at about 90 degrees, so the strength of the carcass will be balanced.

Automobile and truck tires are required to operate for long periods of time, carrying a relatively large but steady load at reasonably high rotational speeds. Therefore these tires are allowed to have only a relatively small amount of deflection. Airplane tires, on the other hand, must be strong enough to absorb the tremendous loads encountered on touchdown, and while they must operate at very high speeds, far higher than that of a car or truck, their ground rolls are of a limited duration. Because of these severe requirements, aircraft tires are allowed a deflection of more than twice that allowed for automobile tires.

Chafers of fabric and rubber wrap around the edges of the carcass plies and enclose the entire bead area. They also provide good chafe resistance between the bead and the wheel.

BEAD

A breaker is an added layer of fabric, which gives more strength. The breaker layers are laid directly on the top casing ply across the width of the tread.

The most important part of a tire is the bead, for it is the bead that anchors the carcass and provides a firm mounting for the tire on the wheel. Ultimately, all of the ground forces on the tire terminate in the bead.

An undertread, which is a layer of specially compounded rubber, is placed over the carcass to provide good adhesion between the tread and the carcass.

The beads are made of bundles of high strength carbon steel wire, with one, two, or three of these bundles used in each side of the tire.

The tread reinforcement, made of one or more plies of nylon fabric, strengthens the tread and opposes the centrifugal forces that try to pull the tread from the carcass during high-speed operation. It also stabilizes the tread on the cord body and prevents its squirming or moving.

Rubber apex strips streamline the round bead bundles, so the fabric will fit smoothly around them with no voids. And flippers, which are layers of rubber and fabric, enclose the bead bundles to insulate the carcass plies from the bead wires. Since the greatest amount of heat in the tire is in the bead area, this insulation increases the durability of the tire.

Tubeless tires are lined with a special compound of rubber that is less permeable than the rubber used in the rest of the tire. In tube-type tires, a thin coating of rubber over the inner ply cords protects the tube from chafing.

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Fig. B

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There have been a number of basic tread patterns used on aircraft tires, the most familiarof them are:

CARCASS (RADIAL TIRE) The casing of a radial-ply tire is not the same as that in a bias tire. Casing fabric plies of a radial-ply tire is laid so that the cord direction is from bead to bead (at approximately 900). (Figure A)

PLAIN TREAD The plain, or smooth tread was popular for tires used on airplanes with no brakes, or for aircraft whose brakes were used primarily as a taxi aid, rather than for slowing the aircraft in its landing roll. Today, this type of tread is found only on some helicopters and on very light airplanes. (Figure B)

The breaker package (also known as the belt plies) is a number of plies attached below the tread of a radial-ply tire. The breaker package gives stiffness in the tread area, and helps to keep the circumference constant as the tire expands when it is inflated.

ALL-WEATHER TREAD

THE TREAD The tread is the wearing surface on the outer circumference of the tire. It is made of specially compounded rubber and has a pattern of grooves molded into its surface to give the tire the required traction characteristics with the type of runway surface the aircraft will encounter.

A diamond-shaped tread pattern is effective for aircraft operating on grass or hard-packed dirt, but good braking action can be obtained on both hard surfaced and dirt runways with a tire having a rib tread in the center and diamonds molded into the shoulders. (Figure C)

One of the problems encountered with modem jet aircraft is that of hydroplaning. This condition occurs when the tire rides on the surface of water or slush, rather than contacting the runway surface. The brakes stop the wheel, and the tire rides on the water, in much the same way as a water ski. When this happens, braking action is nil, and extreme heat is built up in a highly localized area of the tire. Hydroplaning is minimized by the proper design and placement of the grooves in the tread, so the water can pass under the footprint of the tire, and the tread can contact the runway surface.

RIB TREAD This is the most popular tread pattern found on aircraft today. It is designed especially for use on hard surface runways and gives long tread wear, good traction, and exceptionally good directional stability. (Figure D) The width and depth of the grooves and their placement on the tread are factors used to adapt the tire to the operating conditions of the aircraft for which they are designed. Some nose wheels are fitted with tires having twin-contact tread, which consists of a large circumferential rib at each side of the crown, which is designed to assist in preventing shimmy.

Water dispersing treads, which have many small holes incorporated in the crown and shoulder rubber, are also fairly common as a means of helping to prevent aquaplaning.

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Fig. B

Fig. C

Fig. A

Fig D

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Fig. E

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THE SIDEWALL

DEFLECTOR

The sidewall is a rubber covering that extends from the tread down to the bead heel to protect the carcass from such minor injuries as cuts or bruises and from exposure to moisture and ozone.

Jet aircraft with aft-mounted engines have a problem with water or slush being thrown up by the nose wheel and entering the engines, causing damage or flameout. To prevent this, tires used on the nose wheels of these aircraft have chines, or deflectors, molded in the upper sidewall to deflect the water or slush away from the engine intakes. Tires for dual nose wheel installations have chines on one side only, while single nose wheel installations have dual deflectors, one on either side of the tire. (Figure A)

Tubeless tires have an inner liner that is designed to hold air, but some air does seep through. If the sidewall trapped this escaping air in the body plies, it could expand when the tire was heated and cause ply separation and possibly allow the tread to be thrown from the tire. To prevent this, tubeless tires have vent holes in their lower with a green or gray spot and must be kept open when the tire is retreaded. Tube-type tires are also vented to release air that is trapped between the tire and the tube when the tire is mounted.

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Fig. A

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OTHER MARKINGS

TIRE MARKINGS

Other markings that may be found on new tires include the following:

Tires have certain markings imprinted on their sidewalls for identification purposes. These markings vary according to the manufacturer but usually include size, part number, serial number, date of manufacture, tubed/tubeless, speed rating, ply rating, and the type and number of retreads carried out PART NUMBER

-

A Green or gray spots indicating the positions of the awl vents.

-

A red spot or triangle indicating the light part of the tire.

RETREADS

The part number usually includes the manufacturer's identification, the drawing to which the tire is manufactured, and letters to indicate the tread type, and whether it is tubed or tubeless. The part number is the only positive means of identifying a tire, and size markings alone should not be used for this purpose. Example: DR 7153 T.

Retreaded tires are usually marked in accordance with a system peculiar to each manufacturer. The markings usually include the tire part number, the name of the retreader, the number and date of the last retread, and in the case of retreads in which the sidewalls are covered with new rubber, the tire serial number, manufacturer, speed, size and ply rating.

SERIAL NUMBER The serial number is usually marked in conjunction with the date of manufacture, which may be in the form of a code indicating the day, week, or month, and the year. Example: 2283 Nov 72 or 23202283. SPEED RATING Most high-speed tires (those which may be used at speeds over 160mph) have the speed rating imprinted on the tire to indicate the maximum speed for which they are designed, e.g. 200mph.

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TIRE INSPECTION ON THE AIRCRAFT The proper inflation pressure is that specified by the airframe manufacturer in his service manuals and, for the same tire, varies from one aircraft type to another. The pressure specified in the airframe manufacturer's manual is for a loaded tire. When the tire is subjected to this load, it will be deflected the designed amount, and the volume of its air chamber will be decreased enough to raise the pressure by about four percent.

INFLATION The greatest enemy of an aircraft tire is heat, either the heat that is generated within the tire as it flexes when rolling over the ground, or that from such external sources as the brakes or hot runway surfaces. It is the internally generated heat that causes damage that is not likely to be discovered until it results in a tire failure. Aircraft tires are designed to withstand the heat generated by normal flexing for a reasonable amount of time.

Inaccurate pressure gages are one of the major causes for chronic inflation problems. To be sure that your gage is accurate, have it periodically calibrated.

Deflection due to under-inflation of the tire will cause excess heat to be generated within the tire that it is not designed to withstand and therefore can cause internal carcass damage. Tires that have been operated with low inflation pressure will have their tread worn away on the shoulders more than in the center, and any tire showing this pattern of wear should be carefully examined for evidence of hidden damage.

Inflation pressure should always be measured when the tire is cold, and for this reason you should allow two to three hours to elapse after a flight before you measure the pressure. (Figure B)

Over-inflation causes accelerated centerline wear on the tread while leaving rubber on the shoulder. When a tire is worn in this way, it has much less resistance to skidding than it has when its tread wears uniformly.

Inflation pressure of a tire varies with the ambient temperature by about one percent for every five degrees Fahrenheit. If an airplane is to fly into an area where the temperature is much lower than that of the departing point, the pressure should be adjusted before the airplane leaves. The airframe manufacturer's manual should be consulted before adjusting the pressure.

On some aircraft, each wheel is equipped with a tire pressure gauge that is integral with the inflation valve. This built-in pressure gauge allows monitoring and quick check of the tire pressure at anytime. (Figure C)

The importance of maintaining the proper inflation pressure in a tire makes pressure checks one of the most important parts of routine preventive maintenance. Tire pressure should ideally be checked daily and before each flight. (Figure A)

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Nylon tires will stretch when they are first inflated and will increase their volume enough to cause a pressure drop of about 5-10% of the initial pressure in the first 24 hours. Their pressure should therefore be adjusted 12 to 24 hours after installation.

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Fig. B

Fig. A

Fig. C

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CATEGORY B1 - MECHANICAL TREAD CONDITION TREAD DEPTH AND WEAR PATTERN

If the center ribs are worn away while the shoulder ribs an appreciable depth, the tire has been operated in inflated condition, and as such is highly susceptible to bruises. It should be carefully checked for this type of (Figure C)

Since the basic strength of the tire is in its carcass, a tire loses none of its strength as long as the tread does not wear down into the body plies of the carcass. But when the tread is worn away, tire traction characteristics are seriously affected. A tire that has been properly maintained and operated with the correct inflation pressure will wear the tread uniformly, and it should be removed for retreading while there is still at least 1/32” of tread left at its most shallow point. When the tire is removed at this point, there is still enough tread left to provide traction and handling during wet runway operation. (Figure B)

still have an overcuts and damage.

Under-inflation will cause the shoulder ribs to wear more than those in the center. Any tire showing this wear pattern should be carefully inspected for signs of bulges, which could indicate, ply separation. (Figure D) Tread that has been worn until the body plies are visible indicates poor maintenance. If it is worn only to the point that the tread reinforcement is showing, it is possible that the tire can be salvaged by retreading. But if it is worn into the body plies, it has to be scrapped. (Figure E) Uneven tread wear can indicate that the landing gear is out of alignment, and this should be checked and adjusted so it conforms to the specifications of the aircraft service manual. If the uneven wear is slight and the landing gear is not adjusted, the tires may be dismounted, carefully inspected, and reversed so the wear will be evened out. Tread wear in spots can be caused by malfunctioning brakes, improperly torqued bearings, or worn shock struts.

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Fig. B

Fig. C

Fig. D

Fig. A

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Fig. E

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CATEGORY B1 - MECHANICAL TREAD DAMAGE

SIDEWALL CONDITION

Any time the tread is cut more than halfway across a rib, or any of the carcass plies are exposed, the tire should be removed. (Figure A)

The main purpose of the sidewall of a tire is to protect the carcass plies from damage, either from mechanical abrasion or from deterioration by chemicals or by the sun. Small snags or cuts or weather checking in the sidewall rubber that do not expose the cords are not normally considered a cause for removal of the tire, but if any of the ply material is exposed, the tire must be removed.

If there are bits of glass, rock, or metal embedded in the tread, they should be carefully pried out with a blunt awl or a small screwdriver. Be very careful when you do this that you don't puncture the tire. It is also a good policy to hold one hand over the foreign matter to prevent its flying up into your eye when it pops out.

The liner of a tubeless tire is there to hold the air, but some of the air diffuses through into the body plies. The sidewalls of these tires are vented to allow this air to escape, but if the vents do not adequately relieve the pressure, ply separation may occur.

When a wheel locks up on a water-covered runway and rides on the surface of the water, a tremendous amount of heat builds up at the point of contact and actually bums the rubber. Tires showing this type of damage should be removed from service. (Figure B)

NOTE Mark any damage or suspect area of the tire with a light colored crayon before deflating the tire, because when the air is out, these areas will be almost impossible to locate.

Operating on grooved runways can cause chevron-shaped cuts across the ribs of a tire, and, as with any type of cut, if they extend across more than one-half of the rib, the tire should be removed from service. (Figure D)

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Any tread damage should be carefully evaluated by a licensed retreading agency to determine whether or not the tire is repairable.

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Fig. A – Cut Damage

Fig. D – Chevron Cutting

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Fig. B – Wet Braking Flats

Fig. E – Tread Chunking

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Fig. C – Dry Braking Flats

Fig. F – Lateral Scoring

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Punctures that do not exceed ¼” on the outside of the tire and 1/8” on the inside and injuries that do not penetrate more than 40% of the actual body plies can be repaired when the tire is retreaded.

TIRE INSPECTION OFF OF THE AIRCRAFT Any tire that has been involved in an aborted takeoff or severe braking, or has been exposed to enough heat that the fusible plug in the wheel has blown and deflated the tire, should be replaced. This excessive heat has caused damage to the tire that, even though it is not obvious, has weakened the tire enough that it will likely fail in service.

If any bulges were marked when the tire was inflated, carefully check them to determine whether they are ply separations or separations between the tread and the carcass. If it is a ply separation, the tire must be scrapped, but tread separation may possibly be repaired by retreading.

If one tire in a dual installation fails, there have been enough extra stresses put into the other tire that it should be discarded too, even though no visible damage has been done to it.

Carefully examine the sidewall for condition, and if any of the cords have been damaged or exposed, the tire cannot be repaired, as it is reasonable to suspect that the exposed cords have been weakened by exposure to the elements.

Carefully spread the beads apart so you can inspect the inner liner. Don't concentrate the force used to spread the beads, and don't spread them more than the section width of the tire. The use of improper procedure when breaking the bead or when spreading it can kink the wire bundles so the bead cannot seat against the wheel when it is reinstalled, and a tire with a kinked bead should be scrapped.

Carefully examine all of the bead and the adjacent area for indication of damage from tire tools or from chafing against the rim. Any severe damage here would require the tire to be scrapped, but if the damage is only through the chafer, it can be repaired when the tire is retreaded.

Carefully examine the inner liner of tubeless tires for any bulges or blisters, and have any suspect areas evaluated by a retreading agency. (Figure A)

Damage from excessive heat will usually show up on the bead area because the heat can build up here faster than it can be dissipated. If any of the bead area is damaged or has an unusual appearance or texture, the tire cannot be repaired. (Figure E)

Probe all of the suspected areas that were marked when the tire was inflated. When checking any cuts, open them up enough that you can see into their depth, but be sure you don't puncture the tire. (Figure B)

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The bead surface from the wheel flange to the toe of the bead is the sealing surface for a tubeless tire, and if it has been damaged by tire tools or by slipping on the wheel it will not seal. Bare chafer cords, however, if they are not broken, will not normally cause a tire to leak, and they are not necessarily a cause for removing the tire from service.

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Fig. A – Bulge

Fig. B – Cut Damage

Fig. C – Burst Tire

Fig. E – Heat Damage

Fig. D – Sidewall Cracking

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TIRE STORAGE

TIRE RETREADING The great amount of abrasion a tire experiences on its tread each time the airplane lands or taxis on a hard surfaced runway wears the tread away long before the carcass is worn out, and so it is standard practice for commercial aircraft tires to be retreaded.

All new and retreaded tires should be stored in a cool, dry area, out of direct sunlight and away from any electrical machinery. Fluorescent lights, electric motors, generators, and battery chargers all convert oxygen into ozone, which is very harmful to rubber.

When a tire is received by the retreading agency, it is thoroughly inspected. The tread, sidewalls, and beads are checked for cuts, bruises, other damage, or wear, and air is injected into the sidewall to check for any ply separation. The tire is checked for fabric fatigue and for any indication of contamination by oil, grease, or hydraulic fluid. The tires that pass this inspection then have their old tread rubber removed by contour buffing, which produces a smooth shoulder-to-shoulder surface. New tread rubber and reinforcement are then applied to the buffed carcass, and the tire is placed in a heated mold and cured. After it is taken from the mold, balance patches are bonded to the inside of the tire to achieve the proper static balance, and the tire is then given a final inspection.

The storage room should not have extremes of temperature, but should be maintained between 320 and 800 F (00 and 270 C). Special care should be taken to assure that no grease, oil, hydraulic fluid, or any other hydrocarbon compound comes in contact with the stored tire, as all of these compounds will attack the rubber to some degree. Whenever possible the tires should be stored vertically in racks, with the tire supported on a flat surface which is at least three or four inches wide. If tubeless tires are stacked horizontally, the bottom tires in the stack may be distorted so much that the beads will not seat on the wheel unless a special bead-seating tool is used. If it is necessary, however, to stack them horizontally, don't stack them more than five tires high, for tires with a diameter of up to 40”, four tires high for those between 40”-49”, and three high for tires larger than 49”.

The tire is identified as a retreaded tire and a record made of the number of times it has been retreaded. Some tire manufacturers may specify the number of time a tire can be retreaded but for some tires there is no specific limit to the number of times, but is determined by the condition of the carcass.

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Fig. A

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AIRCRAFT TUBES TUBE INSPECTION

TUBE CONSTRUCTION AND SELECTION

If a tube is suspected of leaking, first check the valve by spreading a drop of water over the end of the valve, and watch to see if a bubble forms. If a bubble does form, the valve core should be replaced. If the leak is not in the valve, the tire must be deflated and demounted and the tube removed. If the tube is not too large, inflate it and submerge it in water to find the source of bubbles. If the tube is too large for the available water container, flow water over the surface of the tube as you look for the leak.

A great number of aircraft tires, ranging, from some of the small type Ill up to large type VII are of the tube type. Tubes for these tires are available in either non reinforced rubber for normal applications or as a special heavy-duty reinforced tube that has a layer of nylon fabric molded to its inside circumference to protect it from chafing against the rim and from heat caused by brake application. All aircraft tubes are made of a specially compounded natural rubber that holds air with a minimum of leakage. There are two primary causes for an aircraft tube leaking: a hole in the tube or a defective valve.

It is extremely important, when inflating a tube that is not in a tire, that you do not put more air into it than is required to just round it out. Check the tube carefully around the valve stem and the valve pad for any indication of the pad pulling away from the tube.

It is extremely important that only the tube recommended for a particular tire be used with it. If the tube is too small for the tire, its splices will be overstressed and the tube will be weakened.

Examine the inside circumference of the tube for evidence of chafing against the toe of the bead or by corrosion on the wheel. Any tube that is chafed enough to lose some of its thickness in spots should be replaced. Examine the inner circumference of the tube for any indication that it has been heated enough for the rubber to have lost its smooth contour and taken a set or developed square corners. Any tube that is deformed in this way should be replaced. Reinforced tubes should be used on installations where there is enough heat to damage a regular tube.

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Fig. B Fig. A

Fig. C

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CATEGORY B1 - MECHANICAL TUBE STORAGE Tubes should be stored in their original cartons whenever possible, but if their cartons are not available, they should be dusted with tire talcum and wrapped in heavy paper. Tubes may also be stored inflated by putting them in the proper size tire and inflating them just enough to round them out. The inside of the tire and the outside of the tube should be dusted with tire talc to prevent the tube sticking to the tire. Tubes should never be stored by hanging them over nails or pegs, or supporting them in any way that would cause a sharp fold or crease, as these creases will eventually cause the rubber to crack. Tubes with creases should not be put into service. Tubes, like any other rubber product, should be stored in a cool dry, dark area, away from any electrical equipment that would produce rubber-damaging ozone.

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TIRE MOUNTING – TUBELESS TIRES Now, very carefully place the tire over the inboard wheel half with the red dot indicating the tire's light point adjacent to the wheel valve. Now place the outboard wheel half inside the tire and line up the boltholes.

Most modem aircraft use the split-type wheel which makes tire mounting far easier than it was with either the single-piece dropcenter wheel or with wheels having a removable flange held on with a locking ring.

Apply an anti-seize compound such as Lubtork to the threads of the bolts, to both sides of the washers, and to the bearing surface of the nuts. Install the bolts and nuts, and draw all of the nuts up in a criss-cross fashion to one-half of the required torque. Then go back and bring them all up to the full torque. It is important when torquing the wheel bolts to use an accurate hand torque wrench. Never use an impact wrench on any bolt where the torque is critical, because its torque is applied in a series of blows or jerks, and the actual stresses to which the bolt is subjected are considerably greater than the bolt is designed to take.

The fact that the wheel is so highly stressed makes it extremely important that the manufacturer's service information be followed in detail when mounting and demounting the tires. Before the tire is mounted on a wheel, the wheel must be carefully inspected to be sure that there are no damages and that thermal fuse plugs, inflation air valve and any balance weights are properly installed. Clean the bead seat area and the O-ring seal area with a cloth dampened with isopropyl alcohol, and place the inboard wheel half on a clean, flat surface.

Place the wheel and tire assembly in a safety cage, adjust the air pressure regulator to the recommended tire pressure, and, using a clip-on chuck, inflate the tire gradually. Watch while the tire is inflating to be sure the beads seat against the wheel flange.

Lubricate the O-ring and carefully place the seal in the groove without stretching or twisting it. Check to be sure that the tire is approved for the aircraft on which it is being mounted and that the word TUBELESS is marked on the sidewall. Be sure there is no foreign material inside the tire, and wipe the bead area with rag damp with isopropyl alcohol.

All nylon tires stretch when they are initially inflated and should be allowed to remain for 12 to 24 hours with no load applied. This stretch may cause a five- to ten-percent decrease in pressure, and the pressure should be adjusted after this period. Continue to monitor the inflation pressure daily. There will be some pressure loss, but it should not exceed 5% in any 24-hour period.

Because of the tight fit between the bead of the tire and the wheel, apply a little tire talc to the toe, or inner edge, of the bead. Be sure, however, that no powder gets between the bead and the wheel flange.

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TIRE MOUNTING – TUBE-TYPE TIRES Lubricate the bolts with anti-seize compound and tighten the nuts in a criss-cross fashion to one-half of the required torque. Then, go back and bring all of the nuts up to the recommended value with a good smooth pull on the handle of the torque wrench.

Before mounting a tube-type tire on a wheel check that the tire and tube are both correct for the installation. Check that the wheel is clean and free from oil or grease and any damage. Before mounting the tire, clean the bead seat area with a rag damp with isopropyl alcohol.

Put the tire in a safety cage and, using a clip on chuck, gradually bring the air pressure up to the recommended value to seat the beads and then deflate the tire. Now, re-inflate it to the correct pressure. This inflation, deflation, and re-inflation procedure allows the tube to straighten itself out inside the tire and will remove any wrinkles from the tube.

Check the inside of the tire to be sure that it is clean and free of all foreign matter, and then dust it with an approved tire talcum powder. Fold the inner tube and dust it with talc and slip it inside the tire with the valve sticking out on the side of the tire having the serial number. Inflate the tube just enough to round it out and adjust it inside the tire so the yellow mark that indicates the heavy point of the tube aligns with the red dot on the tire indicating its light point. If there is no balance mark on the tube, you can assume that the valve is the heavy point.

The air pressure in a tube-type tire will drop after its initial inflation because the nylon plies stretch in the same way they do in a tubeless tire, and there may also be air trapped between the tube and the tire. When all this trapped air leaks out around the valve, under the beads, and through the sidewall vents, the inflation pressure will drop. All of this air should be out within the initial 12- to 24-hour period, and the pressure may then be adjusted and the tire put in service.

Install the tire and tube on the outboard wheel half so the valve stem sticks out through the hole in the wheel. If you wish, a bit of tire talc may be rubbed on the toe of the bead to help it slide over the wheel and seat itself. Now you can place the inboard half of the wheel in the tire, two opposite bolts can be inserted to guide it into position: making sure that tube is not pinched between the wheels. On pressing the two halves of the wheel together, a metallic noise should be heard when they meet; this is a good indication of whether or not the tube has been nipped.

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CATEGORY B1 - MECHANICAL TIRE BALANCING As aircraft takeoff speeds increase, the vibration caused by unbalanced wheels becomes annoying. And this vibration is especially noticeable on nose wheels, since they extend quite a distance below the airplane on a slender strut, and they usually do not have a brake to help dampen the vibrations. After the tire is mounted on the wheel, inflated, and allowed to take its initial stretch, the assembly is mounted on a balancing stand with the cones of the balancing shaft seating firmly against the bearing cups in the wheel. Place the shaft on the balancing stand and allow the wheel to rotate until its heavy point comes to a rest at the bottom. (Figure A) Counterbalance the wheel with test weights until the assembly is balanced, and then install the correct amount of weight on the wheel at the location identified by the test weights.

Fig. B

Some balance weights are installed on special brackets that mount under the head of the wheel bolts, and others fasten to the wheel rim by a cotter pin through holes that have been drilled in the rim for that purpose. (Figure B) Fig. A

Many of the smaller wheels do not have provisions for mechanically attaching balance weights, and for these wheel lead strips having an adhesive backing may be used. (Figure C) Be sure to use only the type of weight that is approved for the particular wheel you are balancing, and follow the instructions in the aircraft service manual for the installation of these weights.

Fig. C

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AIRCRAFT BRAKES The various brake designs discussed in this section reflect the variety of braking capabilities required for different size aircraft. Light aircraft can rely on a single disk brake or a simple shoe brake because the landing speeds are slow and the aircraft is light in weight. Large aircraft, such as transports, land at high speeds and weigh several tons. These aircraft require very powerful multi-surfaced brakes in order for the brakes to be effective at slowing the aircraft.

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CATEGORY B1 - MECHANICAL INTERNAL EXPANDING-SHOE BRAKES Shoe brakes will still be encountered on older aircraft and some current models of home built aircraft. The types of internal expanding-shoe brakes are; 1. One-way, or single servo, type 2. Two-way, or dual-servo, type Servo action in a brake of this type means that the rotation of the brake drum adds braking energy to the brake shoes and makes them operate more effectively and with less effort by the pilot. In single-servo type brakes, the servo action is effective when the wheel rotates in one direction only, in contrast with a dual servo, or reversible, type, which is able to give servo action in either direction. Brake-shoe assemblies are attached to the landing-gear strut flange by means of bolts through the torque plate on the axle. The brake drum is attached to the wheel and rotates with it. A dual-servo brake assembly is shown in the drawing of Figure A. As explained previously, dual-servo brakes are effective for either direction of wheel rotation; therefore they can be interchanged between the left and right wheels of the airplane and are effective for both forward and backward motion of the airplane. These brakes may be operated hydraulically, mechanically, or pneumatically.

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Fig. A

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CATEGORY B1 - MECHANICAL EXPANDER-TUBE BRAKES Expander-tube is type of brake that was used on older aircraft.

OPERATION

CONSTRUCTION

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Each expander-tube brake consists of four main parts: brake frame, expander tube, return springs and brake blocks

The expander-tube brake is hydraulically operated and can be used with any conventional hydraulic brake system.

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The single-type brake has one row of blocks around the circumference and is used on small aircraft. The duplex-type, expander-tube brake has two rows of brake blocks and is designed for larger aircraft.

When the brake pedal is pressed, the fluid is forced into the expander tube. The frame prevents any expansion either inward or to the sides. The pressure of the fluid in the tube forces the blocks radially outward against the brake drum.

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When the pressure is released, the springs in the ends of the blocks tend to force the fluid out of the expander tube and to pull the blocks away from contact with the brake drum. This action is increased by the tube itself, since it is molded slightly smaller in diameter than the brake frame and tends to contract without the help of the springs.

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Each block is independent in its action; therefore, there is no buildup of servo action and no tendency to grab.

An inner fairing, or shield, fits between the torque flange on the axle and the brake frame to protect the frame against water. The brake expander tube is a flat tube made of synthetic rubber compound and fabric. It is stretched over the circular brake frame between the side flanges, and it has a nozzle that is connected with the hydraulic-fluid line by means of suitable fittings.

Figure A shows the side and end views of an expander-tubetype brake and Figure B illustrates the principle of operation.

The brake blocks are made of a material similar to that used for molded brake linings. The blocks have notches at each corner to engage with lugs on the brake frame and to prevent movement with the brake drum as it rotates. There are grooves across the ends of each block, and flat return springs are inserted in these grooves. The ends of the springs fit into slots in the side flanges of the brake frame, holding the blocks firmly against the expander tube and keeping them from dragging when the brake is released.

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Fig. A

Fig. B

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CATEGORY B1 - MECHANICAL MAINTENANCE During the inspection and servicing of expander-tube brakes, the Technician must make sure that no hydraulic pressure is applied to the brakes when a brake is not enclosed in its drum. If the brake blocks are not restrained by the drum when hydraulic power is applied, the retaining grooves at the ends of the blocks will brake and the blocks will pop out. When the brake blocks are worn to their allowable limits, they are easily replaced. The return springs that retain the blocks are removed by pressing down on one end clip with a screwdriver or other tool and sliding the springs out of the rectangular holes in which they are held. When all the blocks have been removed, the entire assembly is cleaned and inspected before installing new blocks. The new blocks are installed one at a time with the return springs to hold them in place. If the expander tube is found to be damaged upon removal of the brake blocks, the tube must be replaced.

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CATEGORY B1 - MECHANICAL SINGLE-DISK BRAKES

Most hydraulic brake systems require a method for the removal of air from the system. In Figs. A and C, items 16 through 20 and 15 through 17, respectively, comprise the brake bleeder valve assembly.

One of the most popular types of brakes, especially for smaller aircraft, is the single-disk brake. An exploded view of such a brake is shown in Figure A, and a cutaway is shown in Figure B. CONSTRUCTION

In order to bleed the air from the brakes, the valve is opened slightly and hydraulic fluid under pressure is applied to the piston. A complete discussion of brake bleeding is found at the end of this chapter.

The main disk (1) of the brake shown in Figure A is held in the wheel by means of teeth or keys around the outer rim of the disk, causing it to turn with the wheel but allowing it limited in-and-out movement on the keys. On each side of the disk linings (2) are located, which bear against the disk when the brakes are applied, causing the wheel to slow down or stop.

Although modem aircraft operate the single-disk brake with hydraulic power, some older aircraft are designed with mechanically operated single-disk brakes.

One lining of the brake is mounted in a recess in the plate attached to the main axle structure. The other lining (2) is mounted against the piston (11) and moves according to the amount of hydraulic pressure applied to the piston. In Figure A, three pistons are incorporated in the brake housing (25); therefore, three linings must be mounted on the opposite side of the disk to back up the movable linings. Single-disk brakes may be constructed with as many separate pistons and linings as deemed advisable for the airplane for which they are designed. Each piston is equipped with separate sets of linings, which bear against the brake disk (1) when the brakes are applied.

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with each outer surface faced with a bronze sheet. The bronze sheet takes the wear, whereas the stainless steel wear pads on the stator plates remain comparatively stable.

MULTIPLE-DISK BRAKES Multiple-disk brakes are used on large aircraft where a substantial amount of braking force is required. They are operated by hydraulic pressure. The construction of a typical multiple-disk brake assembly is shown in Figure A.

The torque tube is the structural drum-type member that links the carrier with the back plate. It provides the keys for the stator plates and ties the brake assembly to the axle flange. The torque tube is the structural backbone of the brake assembly and transfers braking forces to the axle and gear strut.

The brake assembly in Figure A is described as a dual-system, five-rotor disk-type power brake. Each brake contains two independent cylinder and passageway systems; and each system contains seven brake pistons, two bleed valves, and one hydraulic pressure port and its associated passageways. Each brake also contains one carrier, four stator plates, one pressure plate, one back plate, five rotor plates, and one torque tube.

Braking action is produced by hydraulic pressure forcing the pistons against the pressure plate, which, in turn, forces the disk stack together and creates friction between the rotating and stationary disks. An organic insulator fitted to each piston prevents brake-heat transfer to the pistons and carrier.

The carrier houses the two independent internally drilled, hydraulic passageway systems; 14 hydraulic pistons, 7 for each system; the brake-return assemblies; 4 bleed valves; and 2 pressure ports, 1 for each system.

When hydraulic-fluid pressure is released from the brake pistons, the return springs and pins pull the pressure plate and the pistons to the full OFF position, thus releasing the brakes. A self-adjusting mechanism on the return pins maintains a constant running clearance throughout the life of the brake. No adjustment is necessary. (Figure B)

The 4 stator plates and the pressure plate are keyed to the torque tube of the brake. The stator plates consist of a steel heat-sink-type core with 14 stainless steel pads riveted onto both outer surfaces. The pads are comparable with brake linings on conventional disk-type brakes but resist wear more effectively. The heat-sink feature helps to absorb and carry the heat away from the stainless steel brake pads.

Although not shown in this example of a multiple-disk brake, wear indicators are often included in the brake design. These rods are extensions of the return pins. As the brake linings wear and the adjuster operates, the wear indicator moves into the assembly. When the exposed length of the rod has decreased to a predetermined minimum length, the brakes must be replaced.

The five rotor plates are keyed to the wheel and rotate with the wheel. Each rotor plate consists of a steel heat sink-type core

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Fig. A

Fig. B

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CATEGORY B1 - MECHANICAL SEGMENTED ROTOR-DISK BRAKES

Alternating stator plates, with brake-lining material on both sides, and rotor assemblies are installed until the proper number of each is in place.

Segmented rotor-disk brakes are heavy-duty brakes designed for use with high-pressure hydraulic systems using power brake control valves or power boost master cylinders. Braking action results from several sets of stationary, high-friction-type brake linings making contact with rotating (rotor) segments. This action is the same as occurs with multiple-disk brakes. An exploded view of this type brake is shown in Figure A, and a cross section of the brake is shown in Figure B.

After the last rotor segment plate is in position, a compensating shim is installed to space the back plate out from the carrier, and then the back plate is installed. The back plate contains brake lining on the side toward the rotors. The compensating shims allow the brake linings to wear down until the piston is out of travel. The shims are then removed, causing the pistons to be moved back into the cylinder, and more of the available brake lining can now be used.

A carrier assembly is the brake component that is attached to the landing-gear shock strut flange and on which all of the other components are mounted.

Because of the gap between the rotor segments and the space between the lining sections, more brake cooling can be achieved than is possible with the multiple-disk brake, allowing more braking action to be achieved before a limiting temperature is reached.

The piston cups and pistons are placed in two grooves, which act as cylinders, in the carrier assembly. The automatic adjusters, which compensate for lining wear, are threaded into holes equally spaced around the face of the carrier. Each adjuster is composed of an adjuster pin, adjuster clamp, return spring, sleeve, nut, and clamp hold-down assembly. The pressure plate is notched to fit over the stator drive sleeve. This component is stationary on the sleeve. An auxiliary stator plate fits next to the pressure plate and has brake-lining material attached to the side away from the pressure plate. The rotor segment plate is installed next. This part is notched on the outside to mate with the wheel and rotate with it. The plate is made up of several segments, as shown.

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Fig. B

Fig. A

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CATEGORY B1 - MECHANICAL CARBON COMPOSITE BRAKES

At the present time carbon brakes are more expensive to produce and are generally found only on long-range aircraft, where weight savings can have a significant impact on cruise performance.

Carbon composite brakes have become a long-life, lightweight alternative to steel brakes since their introduction to commercial aviation as standard equipment on the Concorde. A carbon fiber rotor is illustrated in Figure A.

Carbon also offers reduced maintenance costs. Steel brakes normally last about 1000 landings before they must be replaced. Carbon brakes will last longer (1200-1500 landings) and, due to fewer brake changes, significantly reduce maintenance requirements.

Carbon composites possess several unique properties that permit the combining of the overall brake disc functions of friction surface, heat sink, and structural member into a single unit. The material has the unique property that its strength does not decrease at elevated temperatures. This property, coupled with low thermal expansion, yields a brake heat sink whose operating temperature is limited effectively only by the temperature limits of the surrounding structure. When rubbed against itself, a carbon composite can perform excellently as a friction material. It provides a high heat-storage capability for each pound. This is important because in a rejected takeoff, steel brakes are designed to reach a temperature of about 2000°F. Carbon brakes have the capability of exceeding 3000°F. The maximum temperature for carbon brakes ranges from 3200 to 3400°F. In addition, good thermal conductivity characteristics serve to dissipate the heat rapidly. Another major factor is weight. Carbon brakes weigh 40 percent less than aircraft brakes with conventional steel rotors and linings, meaning greater fuel efficiency and the ability to carry a heavier payload in commercial aircraft.

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AIRCRAFT BRAKE SYSTEMS Brake-actuating systems for aircraft can be classified as mechanically operated, hydraulically operated, or pneumatically operated. All brake-actuating systems provide for applying brakes either on one side of the aircraft or all the aircraft brakes by operating foot pedals or hand levers. Mechanical brakes are found on only a few of the older, small airplanes. A mechanical brake-actuating system includes pulleys, cables, and bell cranks for connecting the foot pedals to the brake-shoe operating mechanism. In some airplanes, the hydraulic brake system is a subsystem of the main hydraulic system. In other airplanes, there is an entirely independent brake system. Pneumatic brake systems utilize air pressure instead of fluid pressure to operate the brakes. Some hydraulic brake systems are arranged with a pneumatic backup system for operation in case of hydraulic-fluid loss or failure of hydraulic pressure.

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It is a simple but effective unit, normally connected by a linkage to the brake pedal mounted on the rudder pedal. The hydraulic fluid enters the master cylinder through the inlet port and compensating port from the external reservoir, which supplies the master cylinders for both the right and left brake systems. The application of the brake forces the piston into the cylinder and causes hydraulic fluid to flow toward the brake-actuating cylinder in the wheel.

INDEPENDENT BRAKE SYSTEMS An independent brake system, such as is shown in Figure A, is usually found on small aircraft. This system is self-contained and independent of the aircraft's main hydraulic system. The basic components of this type of system are a reservoir, a master cylinder operated by the brake-control pedal or handle, a brake assembly on the wheel, and necessary lines, hoses, and fittings. Expander-tube, shoe, or disk-brake assemblies may be used with this type of system.

The illustration shows the cylinder in the horizontal position, but when it is installed in the aircraft, it is in a vertical position with the eye of the piston rod downward. When the piston moves against the return spring, the compensating port is closed and the fluid in the cylinder is trapped under pressure. Continued pressure applied through the brake pedal forces the fluid pressure to the brake-actuating cylinder and applies the brake. When the force is removed from the brake pedal, the piston is returned to the OFF position by means of the return spring, and the compensating port is again open to the reservoir. The compensating port permits the fluid to flow toward or away from the reservoir as temperature changes, thus preventing a build-up of pressure when the brake is off.

The reservoir is a storage tank that supplies the fluid to compensate for small leaks in the connecting lines or cylinders. The reservoir may be a part of the master cylinder or it may be a separate unit, as shown in the drawing. It is vented to the atmosphere to provide for feeding the fluid to the master cylinders under the force of gravity; therefore, the fluid must be kept at the correct level, or air will enter the system and reduce its effectiveness. The master cylinder is the energizing unit. There is one for each main landing-gear wheel. The master cylinder is actually a footoperated, single-action reciprocating pump, the purpose of which is to build up hydraulic fluid pressure in the brake system. One type of master cylinder for light aircraft is illustrated in Figure B.

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With this type of master cylinder, the brakes are locked in the ON position for parking by means of a ratchet-type lock that is constructed as part of the mechanical linkage between the foot pedal and master cylinder. If an increase of temperature occurs, expansion increases the volume of fluid. This is compensated for by means of a spring built into the linkage. To unlock the brakes, the pilot applies enough force to the brake pedals to unload the ratchet-type lock.

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Fig. B

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CATEGORY B1 - MECHANICAL Another type of master cylinder is used on a number of light aircraft. This cylinder is illustrated in Figure A. These cylinders are mounted on the rudder pedals, as shown in the drawing of Figure B.

OPERATION As pressure is applied to advance the piston rod into the cylinder, the piston remains stationary until the lock-o-seal is seated on the piston. When the lock-o-seal is seated, fluid cannot pass the piston, and with continued movement of the piston rod forcing the piston into the cylinder, pressure in the cylinder is increased. At any time during the stroke if the force on the piston is eased, the piston return spring will tend to keep the piston seated against the lock-o-seal, maintaining pressure in the cylinder.

CONSTRUCTION In the illustration of Figure A, this type of master cylinder incorporates a fluid reservoir (8) on the top of the cylinder (11) within the same body (7). A plastic filler plug (18) is used to close the opening in the cover (4), which is threaded into the body. The filler plug is not vented because sufficient ventilation is provided by clearance between the piston rod (3) and the pistonrod opening through the cover boss (6).

As the force is further eased, allowing the piston-return spring to force the piston to retreat, the upper end of the compensating sleeve will contact the cover boss; thus the piston is forced to unseat itself from the lock-o-seal. This allows additional fluid from the reservoir to enter the cylinder. This positive unseating also allows unrestricted passage of fluid from the cylinder to the reservoir while the piston is in the static position. This action is to compensate for any excess fluid that may be present in the system due to pumping or from thermal expansion.

With the exception of the piston return spring (12), all internal operating parts are assembled onto the piston rod. These parts are the piston (15), piston spring (14), "lock-o-seal" (16), and compensating sleeve (17). A seal between the piston and the cylinder walls is provided by the a-ring (9) installed in a groove around the piston.

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CATEGORY B1 - MECHANICAL POWER ASSISTED OR BOOSTED BRAKE SYSTEMS There is a size of airplane which requires more braking force than can be applied with an independent master cylinder, yet it does not require the complex system of a power brake; the boosted brake fills this need. The boosted brake master cylinder of Figure B is typical for this type of operation. This cylinder is mounted on the rudder pedal and attached to the toe brake pedal in such a way that depressing the pedal pulls on the rod and forces fluid out to the brake cylinder. If the pilot needs more pressure on the brakes than he can apply with the pedal, he continues to push, and as the toggle mechanism straightens out, the spool valve is moved over so it will direct hydraulic system pressure behind the piston where it assists the pilot in forcing fluid out to the brake. When the pedal is released, the spool valve moves back to its original position and vents the area on top of the piston back to the system reservoir. At the same time the compensator poppet unseats and vents the brakes to the reservoir.

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acting on the end of the rod creates a return force tending to close the valve. This return force varies with the intensity of braking force and provides "feel" at the pedals. The desired braking effort is obtained by depressing the pedals a greater or lesser distance. Cable stretch and adjustment of pedal position permits the valve rod to move back until both pressure and return ports are closed. At this point the braking effort remains constant. This condition is shown schematically in Figure A. Releasing the brake pedals allows the pressure in the compensating chamber to move the valve rod out and open the brake line to the return line. When the pressure in the brake line falls, the brakes are released and return force on the valve rod is also reduced.

POWER BRAKE SYSTEM A power brake system is used to operate the brakes of large aircraft where the independent and power boost systems are not adequate. The pilot initiates braking by depressing the brake pedal. This causes a power brake-control valve to direct hydraulic system pressure to the brakes and operate the brake assembly. The brake pedal is connected to the power brake control valve through an arrangement of cables, pulleys, bell cranks, and linkages. The power brake control valve for a transport aircraft is illustrated in Figure B. These are also called brake metering valves. One metering-valve assembly is used for each main landing-gear brake.

Automatic braking to stop the rotation of the wheels during retraction is provided by a small-diameter piston actuating cylinder attached to the metering valve. The cylinder is connected to the landing-gear-retract hydraulic line. When the landing-gear control is placed to ‘UP’, hydraulic pressure is directed to the automatic cylinder and the piston extends. One end of the piston rod rests on the valve rod; therefore, extension of the piston opens the metering valve and applies the brakes.

In a typical system, four hydraulic lines are attached to each valve. These lines are for pressure, return, brakes, and automatic braking. Valve ports are opened or closed by operating a circular grooved, sliding valve rod (spool). The linkage end of the valve rod projects beyond the valve body, whereas the opposite end is supported in a sealed compensating chamber. When the brake pedals are depressed, an inward movement is imparted to the metering valve rod through the mechanical linkage and cables. As the rod moves in, the return port is closed, and the pressure port is opened to direct hydraulic fluid pressure to the brakes. A passage through the valve rod permits the hydraulic fluid under pressure to enter a compensating chamber enclosing the inner end of the valve rod. Pressure

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Fig. B

Fig. A

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CATEGORY B1 - MECHANICAL DEBOOSTER VALVE

The debooster has a pin-operated ball valve that allows fluid in the line to the brakes to be replenished if there should be a leak in the line. If the debooster piston should move down enough for the pin to push the ball off its seat, fluid under system pressure will flow into the lower chamber and replenish the lost fluid. As soon as enough fluid enters the chamber, the piston will rise and the ball will reseat.

A brake debooster valve is installed in systems where the high pressure of the hydraulic system is used to operate brakes that are designed to work with lower pressure. This valve is positioned in the hydraulic line between the power brake-control valve and the brake assembly. These are primarily pressure-reducing valves that operate on the basis of a pressure differential being produced by an area differential. The principle of the debooster is illustrated in Figure A, where we have a system pressure of 1,500psi applied to a piston having an area of two square inches. This generates 3,000 pounds of force. If the other end of the piston has an area of ten square inches, the 3,000 pounds of force will produce a pressure of 300psi.

Lockout debooster, such as the one shown in Figure B allow the piston to go all the way to the bottom. The pin pushes the ball off its seat, but the spring-loaded valve prevents fluid entering the lower chamber until the reset handle is lifted.

The other function of the debooster is to increase the volume of fluid going to the brakes. When the 1,500psi system pressure moves the small piston down one inch, two cubic inches of fluid is used, but this same travel of the larger piston moves ten cubic inches of fluid to the brakes.

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CATEGORY B1 - MECHANICAL MULTIPLE POWER BRAKE-ACTUATING SYSTEMS The brake-actuating system for a transport turbine aircraft involves many components and a number of subsystems and is described briefly here to provide the technician with a general understanding of how such a system operates. This information also emphasizes the need for careful work on the part of the technician while servicing, maintaining, and repairing such a system.

The schematic diagram of Figure A shows the operating components and subsystems of the DC-10 brake system. As can be seen in the drawing, the brake pedals of the airplane are mechanically connected to the brake valves, which control hydraulic pressure to the brakes. Each brake pedal controls two cable systems. These cable systems operate a pair of corresponding dual-brake-control valves that are located in the right and left main-gear wheel wells. The brakes are operated by two completely independent hydraulic power systems. The no.1 hydraulic system supplies pressure to the no.1 brake system, and the no.3 hydraulic system supplies pressure for the no.2 brake system. Each wheel brake is actuated by power from both systems through independent pressure-metering valves. Each brake system consists of a dual-brake-control valve, pressure accumulator, brake-pressure transmitter and indicator, brake system manifold, eight skid-control valves, eight fluid quantity-limiter valves, a skid-control manifold for each gear, and a parking-brake valve, all of which contribute to the actuation of the independent cylinders in the eight main wheel brakes. Although both brakepressure systems are normally used at all times, either system is capable of stopping the airplane on a maximum-gross-weight landing.

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Fig. A

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CATEGORY B1 - MECHANICAL EMERGENCY BRAKE SYSTEM In the case of a total failure of the hydraulic system, the pilot of most large aircraft can operate a pneumatic valve on the instrument panel and direct compressed air or nitrogen into the brake system. When the pilot turns the handle, he is actually adjusting a regulator that controls air pressure to the brake. When sufficient pressure reaches the brake line, the piston moves up against the force of the control spring and shuts off the inlet valve. The compression of the spring determines the amount of pressure supplied to the brake. When the brake handle is rotated in the direction to release the brakes, the air is exhausted overboard. Rather than allow compressed air to enter the wheel cylinders, which would require the entire brake system to be bled of this air, the emergency air may be directed into a transfer tube. The air forces hydraulic fluid from this tube into the brake system.

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CATEGORY B1 - MECHANICAL BRAKING HEAT ENERGY Stopping a high-speed aircraft, either upon landing or as required for an aborted takeoff, involves the conversion of a great amount of kinetic energy to heat at the brakes and main wheels. This energy may be specified in foot-pounds or joules.

These are as follows:

Brake limitation charts have been prepared for some airplanes to give crew members and maintenance personnel a means of determining how to deal with hot brake situations safely and effectively. One such chart for a comparatively small jet airplane is shown in Figure A. The specified purpose of the chart is to avoid in-flight fires and to ensure adequate brake capacity at all times for a rejected takeoff.

Zone II, normal zone: 1.0 - 2.05 million ft-lb [3.8 x105 - 2.8x105J]

Zone I, normal zone: Below 1.0 million ft-lb [3.8x105J] 1. No special requirement under normal operations. 2. Delay subsequent takeoff as indicated by chart. Zone Ill, caution zone: 2.05 - 4.0 million ft-lb [2.8x105 - 5.5x105J] 3. Move the airplane to clear the active runway, because uneven braking could cause one or more tires to deflate if energy is in the upper range.

Note in the chart of Figure A that the factors used in determining the amount of energy absorbed by the brakes in a given situation are (1) the indicated airspeed in knots at the time the brakes are applied, (2) the gross weight of the airplane, and (3) the density altitude at the airport where the braking occurs. The proper use of the chart will establish a condition zone for any particular braking event. For each zone, a particular set of requirements is set forth.

4. Use brakes sparingly to maneuver. 5. Do not set parking brake. 6. Allow brakes to cool for the time indicated by chart. 7. After brakes have cooled, make a visual check of brakes. Zone IV, danger zone: Over 4.0 million ft-lb [5.5 x105J] 8. Clear the runway immediately as fusible plugs will blow 2 to 30 min after stop. 9. Do not apply dry chemical or quench until fusible plugs have released tire pressure. 10. Do not approach for ½ hour or until fusible plugs have blown. 11. When artificial cooling is not used, 2 to 3 hours are required for brakes to cool enough for safe removal. 12. Tires, wheels, and brakes must be replaced.

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CATEGORY B1 - MECHANICAL The dotted lines in the chart of Figure A are to show how a particular braking event can be evaluated and the appropriate zone located. The steps are as follows: 1. Locate aircraft gross weight on the chart. 2. Project vertically to the indicated airspeed line. 3. Project from the airspeed line horizontally to intersect the zero-altitude line. 4. Project parallel to the nearest density line to intersect the correct altitude vertical line. 5. Project horizontally to give correct zone, kinetic energy, and ground cooling time. The example demonstrated on the chart is for an airplane with a gross weight of 15,000Ib [6804kg]. The indicated airspeed at the time the brakes were applied was 110kg. The density altitude was 5000ft [1524m]. When these values are applied to the chart, it is revealed that the kinetic energy is 2,350,000ft-lb [3186172J] in the caution zone, and the cooling time is 28 min. If an airplane is equipped with thrust reversers, the energy absorbed by the brakes is reduced from the values shown here. The amount of reduction depends upon the time that the thrust reversers are operated.

Fig. A

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BRAKE INSPECTION AND SERVICE Flexible hoses used in the brake system must be examined for swelling, sponginess, leakage, and wear of the outer covering. A brake hose that has become soft or swollen is likely to cause "spongy" brakes, because the hose will expand and take up some of the fluid volume. The effect is similar to the presence of air in the system.

The technician should always refer to the aircraft manufacturer's instructions when inspecting and maintaining brake systems. The information presented here is meant as an introduction to brake maintenance tasks. WARNING: Brake systems may include compounds using asbestos in the construction of the brake-lining material. Use appropriate respiratory and other health-protective measures when working with these systems.

During the inspection of brakes and brake systems, the technician should follow up on any discrepancies, defects, and malfunctions reported by the pilot or other crew member.

Brake inspection is accomplished at the same time that wheels are removed for other inspections.

CHECK FOR LINING WEAR The most important inspection for an installed brake is that regarding the wear of the brake lining material.

The brake disks should be inspected for pitting and grooving. If these surface defects exceed the manufacturer's limits, the disk must be resurfaced or replaced.

This check should be done with the brakes applied unless instructed otherwise. The manufacturer's service manual specifies the minimum extension allowed before the brake should be disassembled and the linings replaced.

Check the retaining clips that hold the brake disk in the wheel on some types of disks. If the disk is distorted or the clips are broken, the brake disk will not stay in the correct position and brake failure will result.

The amount of wear on single-disk brakes is usually indicated by the amount the return pin of the automatic-adjusters sticks through the bushing.

During major inspections brakes are disassembled and disks examined for wear and for damage resulting from heat. The lugs or keys holding the rotor disks in the wheels are examined for wear and security. The actuating parts, such as the pistons and pressure plates, must also be examined for condition. Leaking pistons must be repaired as directed.

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If the brake does not have automatic adjusters, the lining wear can be measured by measuring the distance between the disk and the housing. The total stack wear of the disks in a multiple-disk brake may be checked by applying the brake and measuring the amount of clearance between the back of the pressure plate and the edge of the housing.

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Fig. A

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CATEGORY B1 - MECHANICAL CHECK FOR AIR IN THE SYSTEM

Almost all brake housings are made of either cast aluminum or cast magnesium alloy, and it is possible that porosity of the casting could cause a seep-type leak. If it is determined that this is the cause of fluid on the outside of the brake, the casting should be returned to the manufacturer for repairs.

Spongy brake action is nearly always in indication of air in the system, and if any air is found, it must be removed by bleeding before proper braking action can be restored. CHECK THE FLUID QUANTITY AND TYPE

CHECK FOR THE PROPER BOLT TORQUE

The hydraulic system of modem aircraft may use either a synthetic fluid or mineral-base. Before servicing the system, be extremely sure that you sure the proper fluid, as the use of the wrong type will damage any of the seals it contacts and will contaminate the entire system.

All of the components of an aircraft landing gear are subject to hammering action, and it is vital that all of the bolts have the proper torque, as any looseness will accelerate wear and damage. After a bolt has been properly torqued, a small touch of paint should be placed across the end of the bolt and the nut or across the bolt head and the adjacent structure. Any break in the paint will show if there has ever been any motion between the fastener and the structure. If this paint line is unbroken, you can be reasonable sure that the torque is still adequate. If there is no paint mark, the brake attachment bolts should be checked for the proper torque. Use a torque wrench of known accuracy and recent calibration, to be sure that the bolt has the torque recommended in the aircraft service manual.

On aircraft equipped with independent or boosted brake systems the fluid quantity of the brake reservoir should be checked and serviced regularly. On some aircraft the reservoir may be integral with the brake master pump. If frequent fluid servicing is required a leak check of the system should be performed. Separate brake system servicing is not required on aircraft equipped with powered brakes using hydraulic system pressure CHECK THE ENTIRE SYSTEM FOR LEAKS Brake system leakage can be an indication of hidden or impending damage, and the cause of the leakage must be found and corrected. Leaking fluid line connections deserve special care, as tightening the fitting is not necessarily a sure way to stop the leak. If a fitting is leaking, remove the pressure from the system and check the fitting for proper torque. If it is obviously loose, tighten it and re-check it with pressure.

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GRABBING

TYPICAL BRAKE MALFUNCTIONS AND DAMAGE OVERHEATING

As the brakes are applied, build up of the braking action should normally be gradual and linear. Upon application if the braking action is sudden and uneven then the brakes are said to be grabbing. Grabbing brakes are usually caused by oil or some other foreign matter on the disks and linings. In addition, worn disks and drums can also cause grabbing.

One of the most demanding requirements for the modern aircraft brake is that of its ability to absorb tremendous amounts of heat energy. Any time a brake shows signs of overheating, or if it has been involved in an aborted takeoff, it must be removed from the aircraft and given a complete inspection and rebuilt.

CHATTERING OR SQUEALING An extremely annoying condition can exist when the brakes are applied if, instead of exerting a smooth and even friction between the surfaces of the disks, the friction is applied and released many times a minute. This produces a loud chattering noise, and if the frequency of this chattering is quite high, the noise will be a loud squeal. Chattering and squealing brakes are not only annoying to hear, but the vibration generated is harmful to the landing gear and brake structure.

DRAGGING When a brake fails to completely release after the pressure is removed, it is said to be dragging, and if this condition is not immediately corrected, it will lead to excessive heating and wear of the disks. Dragging may be caused by disks that have become dished from overheating or by weak or missing return springs or by a slipped pin in the return system. Air trapped in the system will also cause brakes to drag as air expands when heated keeping the brakes partially applied.

Warped or glazed disks will cause chattering, as will any unparallel condition of the surface of the disk stack. If the disks have been overheated, there is a possibility that some of the mix has been transferred from a rotating disk to a stationary disk, and the uneven friction caused by this transfer will produce chattering.

When brakes are released, all hydraulic pressure should be released from the brake cylinders; however, if a valve sticks closed or are plugged, the pressure may not be released and the brakes will drag.

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CATEGORY B1 - MECHANICAL SPONGY Spongy brakes are caused by air in the brake hydraulic system. This problem occurs because the air contracts and expands as pressure is applied or released, thus creating the spongy feeling and reducing the positive pressure that should be available for brake operation. FADING Fading brakes are usually caused by the condition of the lining. If the brakes have been overheated and the linings burned, glazing is likely to take place on the surface of the linings, and this condition can result in brake fade. Any drum-and-shoe-type brake in which a great deal of heat is generated has a problem of fading. As the drum is heated, it expands in a bell-mouth fashion and loses a great deal of its effectiveness. EXCESSIVE PEDAL TRAVEL Excessive brake pedal travel can be caused by worn brakes, lack of fluid in the brake system, air in the system, or improperly adjusted mechanical linkages. In some very large aircraft, the brake-control cylinders are located in the wheel wells in the wings. The control cylinders are connected to the brake pedals through a system of levers, cables, and pulleys. If this system is worn or not adjusted properly, excessive pedal travel will result. A leaking piston seal in a brake master cylinder will cause the pedal to slowly creep down while pedal pressure is applied. In such a case, the master cylinder should be overhauled.

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Gravity Method

BLEEDING OF BRAKES

This type of brake may also be bled by forcing trapped air out of the system at the wheel cylinder. Slip one end of a section of flexible tubing over the bleeder valve on the brake, and immerse the other end of the tubing in a container of clean hydraulic fluid. Check to be sure that the reservoir is full of fluid, and depress the master cylinder. Hold the piston down and crack the bleeder valve on the wheel. Allow the piston in the master cylinder to go all of the way down, and then close the bleeder valve before you release the master cylinder. Continue this procedure, being sure to keep the reservoir full until fluid flows from the brake with no bubbles.

MASTER CYLINDER BRAKES There are two general methods of bleeding Master cylinder equipped brakes. They are the pressure method and the gravity method. Pressure Method For this type of bleeding, brake fluid is forced through the system by means of a pressurized reservoir connected to the system brake-bleed fitting. Fluid is forced through the system and back to the reservoir until no bubbles appear in the fluid outflow. This is illustrated in Figure B.

POWER BRAKES

To bleed this type of brake, remove the screw from the bleeder valve and connect the brake bleeding pressure pot to the brake with a flexible hose. Before tightening the line on the bleeder valve, be sure to purge all of the air from it.

Remove the bleeder screw from the brake's bleeder valve and install a length of flexible hose. Place the end of the hose in a clean container of hydraulic fluid. Open the bleeder valve and, with the hydraulic system pressure in the proper operating range, very carefully apply the brakes. Allow the fluid to flow from the brake until there are no more air bubbles. Close the bleeder valve, remove the hose, and replace the bleeder screw.

CAUTION: Be sure that the pressure pot is filled with the proper type of hydraulic fluid, as improper fluid can cause serious damage to the system. Attach a flexible hose to the brake reservoir vent and place the end of the hose in a clean container. Open the bleeder valve with pressure applied from the pot to force fluid up through the brake, the master cylinder, and reservoir into the container. Allow the flow to continue until there are no more air bubbles. Close the bleeder valve and the valve on the pressure pot and remove the line from the brake. Replace the bleeder screw, and remove the line from the reservoir. Check to be sure that there is the proper amount of fluid in the reservoir.

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It is especially important that the airframe manufacturer's service manual be followed in detail when power brakes are bled, as there may be antiskid components, deboosters, or hydraulic fuses in the lines that must be given special consideration. The bleeding procedure is not complete until both the main and the emergency, or backup, systems are free of air.

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Fig. A

Fig. B

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Fig. C

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SKID PROTECTION BRAKING ACTION AND SKID PREVENTION

INTRODUCTION

For maximum brake effectiveness, the friction between the tire and the runway surfaces should closely relate to the friction in the brake so that the peripheral speed of the tire will be just slightly less than the speed of the aircraft. When this is true, the tire will grip the runway surface and slip just a little. This produces the maximum tire drag.

Our modern high-speed jet aircraft usually have more than one wheel on each side, and all of the brakes on one side are controlled with one pedal. With this arrangement, the pilot has no way to know when one of these wheels begins to skid so that he can take corrective action. But, if corrective action is not taken within a few seconds to release a locked-up wheel, the tire is likely to blowout and control of the aircraft can be lost.

Maintaining this optimum friction is no easy matter, because if the brake pressure is held constant after the slip starts and the wheel begins to decelerate, the brake friction will rapidly increase to the point that the wheel will lock up. The tire will skid over the runway and produce very little effective braking.

When the aircraft touches down on a water covered runway and the pilot applies the brakes, the friction on the runway surface is so much less than that generated in the brake that the wheel locks up and the tire hydroplanes down the runway, supported on the surface of the water in much the same way a water skier is supported.

We use a very simple form of manual anti skid control in an automobile when we drive on ice. For the most effective stopping, we pump the brakes, applying them only enough to slow the wheel, but releasing them before the wheel decelerates enough to lock up.

All braking action and directional control is lost for a hydroplaning wheel.

This same on-and-off type of operation was employed in some of the early aircraft antiskid systems, but it has a major drawback if the control valves do not operate fast enough.

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CATEGORY B1 - MECHANICAL In Figure A, we see the way this problem comes about. -

When the brakes are applied, the pressure rises until the wheel starts to slip, but not skid (Point A). This is the ideal condition, but the pilot has no indication that it has been reached, so he continues to increase the force on the brake pedal.

-

A pressure is soon reached which produces enough friction in the brake to cause the tire to start to skid on the runway (Point B).

-

The wheel now decelerates fast enough that the pilot can feel it, so he releases the pedal. But since the braking force needed becomes less as the wheel slows down, the wheel continues to decelerate even though the brake pressure is decreasing.

-

At point C, the wheel has completely locked up, and the pressure continues to drop.

-

At point D, the pressure is low enough for the friction between the tire and the runway surface to start the wheel rotating again, and soon after this the brake pressure drops to zero.

-

The wheel then comes back up to speed.

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Fig. A

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CATEGORY B1 - MECHANICAL ANTI-SKID REQUIREMENTS All anti-skid systems work on the principle of releasing the brake on a wheel that has, or is about to lock up, this is termed skid anticipation. Information of a wheel about to skid (rapid deceleration) can be received by either mechanical or electrical devices. The mechanical system uses the principle of inertial weights; the electrical system uses wheel speed signals sent to a computer which in turn transmits signals to the anti-skid valves to release the appropriate brake unit. The advantages of anti-skid systems are: -

maximum braking efficiency,

-

reduces the landing run,

-

no skidding or locking of the wheels, increased tire life,

-

the pilot can apply any degree of braking effort without fear of loss of directional control of the aircraft or tire scrub damage.

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Axle mounted maxarets are mounted inside the wheel axles. They are driven by the aircraft wheel via the hub cap and a flexible drive to prevent damage in case of seizure of any of the components. The hydraulic connections are made via a self sealing hydraulic coupling on installation. An extractor tool is normally used when removing the maxaret. The advantages of axle mounted maxarets over external types are:

MECHANICAL ANTI-SKID SYSTEMS This type of anti-skid system uses components that are situated around the wheel area, they are self contained systems which can when required, interrupt the supply of metered brake pressure going to the brake unit. There are two types of mechanical systems in common use: -

externally mounted

-

axle mounted

These mechanical units are referred to as MAXARETS (Maximum Retardation Units). The principle of operation is the same for both types so we will only discuss one type.

-

that they are less susceptible to damage as the entire unit is enclosed by the axle,

-

that they are easier to replace as no pipeline connections are involved,

-

that there are no drive problems, such as oil on the aircraft wheel to cause slip

INSTALLATION Figure A shows the installation of externally mounted and axle mounted maxaret units.

OPERATION The maxaret unit is connected to the hydraulic line between the pilot metered pressure from the brake metering valve and the associated brake unit. The unit is sensitive to the angular deceleration which occurs when entering a skid, at such a moment sufficient energy is released by the flywheel to operate the valve system and release the brake. When the wheel regains speed the flywheel setting is restored and the brake reapplies until there is a further tendency to skid.

Externally mounted maxarets are mounted either on the brake torque plate, the leg or the bogie beam. The maxaret is driven by a rubber tired wheel which contacts the aircraft wheel. The maxaret is adjusted so the tire will firmly contact the aircraft wheel, about 1" of tire contact is normal. This adjustment can be carried out by two methods, depending on the type. The methods of adjustment are: -

shimmed installation,

-

self adjusting (by spring loading)

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Fig. A

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CATEGORY B1 - MECHANICAL SEQUENCE OF OPERATION Figure A shows the non skid condition with brakes applied. The aircraft wheel and the maxaret are spinning at the same speed and pressure is allowed to pass to the brake unit. Figure B shows the assemblies in a skid condition, the following events have taken place: -

the wheel has slowed down (impending skid) and the outer tired drum has decelerated too as it is being driven by the aircraft wheel,

-

the flywheel, due to inertia, continues to rotate against the main spring, and travels about 60 relative to the drum,

-

the thrust balls are driven up the cam profile and operate the push rod,

-

the inlet valve is closed and the exhaust valve is opened thus releasing the brake,

-

the flywheel is gradually slowed down by the drag of the main spring and with the brake off the landing wheel regains speed,

-

the drive drum begins to drive the flywheel again and the thrust balls return to base of the cam and normal braking is returned.

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Fig. A

Fig. B

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CATEGORY B1 - MECHANICAL ELECTRONIC ANTI-SKID SYSTEMS Many modern aircraft have electronic anti-skid systems, which also prevent skids by releasing the brake pressure, they are much more sensitive and can modulate the brake pressure for optimum braking efficiency. The electronic systems are more reliable and require less maintenance than the mechanical types. Only one item of this system is around the wheel area, the transducer, all the rest of the components are out of harm's way from stones and other debris. An anti skid system consists basically of three components; the wheel speed sensors, the control box, and the control valves. There are two types of systems in use, an AC system and a DC system. They are essentially alike except for the wheel speed sensors and one circuit in the control box.

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Fig. A

Fig. B

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CATEGORY B1 - MECHANICAL WHEEL SPEED SENSORS

DC Sensor

The purpose of the wheel speed sensor is to relay to the control unit the speed of the wheel to which it is fitted, there is one sensor per wheel/brake assembly. Sensors are fitted in the axle and driven by the hub cap as shown in Figure A.

The DC sensor is essentially a small permanent-magnet directcurrent generator whose voltage output is directly proportional to the rotational speed of its armature. When this type of sensor is used, there is no need for the converter in the control box, and there is less danger of stray voltages induced into the system causing brake interference.

AC Sensor The AC sensor is a variable-reluctance AC generator, which uses a permanent magnet surrounded by a pickup coil in the axle of the landing gear (Figure B). The outside of this sensor has four equally spaced poles with teeth cut into their periphery.

The shaft of the armature is fitted with a blade driven by a bracket in the wheel hubcap and rotates with the wheel.

A soft iron exciter ring with internal teeth is mounted in the hubcap of the wheel so that it rotates around the sensor. The two sets of teeth pass near each other and, as the exciter ring rotates, the teeth approach each other and then separate. As the distance between the teeth changes, the reluctance of the magnetic circuit is alternately increased and decreased, and each time the distance changes, the amount of magnetic flux cutting across the coil changes and induces an alternating current in the pickup coil. The faster the wheel turns, the higher the frequency of the induced current. The control box used with the AC sensor converts the varying frequency AC into a DC signal whose voltage is proportional to the frequency of the AC.

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Fig. B

Fig. A

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CATEGORY B1 - MECHANICAL CONTROL VALVES A three-port antiskid control valve is located in the pressure line between the brake valve and the brake cylinder, with a third line connecting the control valve to the system return manifold.

The DC signal from the control box flows through a coil around the armature of the flapper valve. This armature is free to pivot and is centered between two permanent magnets.

For normal operation of the brakes, when no skid is being indicated, the valve allows the brake fluid to flow into and out of the brake, with the valve serving only as a passage. But, if the wheel speed sensor determines that one of the wheels is beginning to decelerate fast enough to cause a skid, its changing output voltage is measured in the control box, and a DC signal is sent to the control valve to close off the pressure port and open the passage between the brake and the system return.

When the signal from the control box indicates that no skid is impending, and the braking action should be normal, the magnetic field of the coil reacts with the fields of the permanent magnets and holds the flapper centered between the nozzles. Fluid from the brake valve flows through the filter and discharges equally from each nozzle. Since the amount of flow is the same through each orifice, the pressure drop across the orifices will be the same, and the second-stage spool valve will assume a position that allows free passage between the brake valve and the brake.

This valve operates fast enough to maintain an output pressure that is directly proportional to the amount of signal current from the control box.

Rev. 00 Oct 2006

But when the control box receives a signal from the wheel speed sensor telling it that a skid is impending, it sends current through the coil of the armature to create a polarity in the armature that causes it to pivot and unbalance the flow from the nozzles.

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Fig. A Fig. B - direct-current signal from the control box energizes the coil on the armature of the flapper valve, and the movement of the flapper changes the pressure drop across the fixed orifices. This varies the pressure on the lands of the second-stage spool to control the flow of fluid through the valve.

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CATEGORY B1 - MECHANICAL In our illustration of Figure B, the flapper has moved over, restricting the flow from the left nozzle and opening the flow from the one on the right. There is now more flow through orifice O2 and therefore a greater pressure drop across it, leaving P1 greater then P2. This imbalance of pressures moves the second stage spool over, shutting off the flow of fluid from the brake valve to the brake, and opening a passage from the brake to the return manifold. The extremely fast reaction time for this type of valve allows it to maintain a pressure at the brake that is directly proportional to the amount of current flowing in the armature coil.

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Fig. A - When the flapper is centered between the nozzles, the pressure drop across the two orifices is the same, and P1 is equal to P2.

Rev. 00 Oct 2006

Fig. B - When the armature of the flapper valve is energized, the flapper moves over and restricts the flow through orifice O1 while increasing it through O2. The increased pressure drop across O2 causes P1 to be greater than P2

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A deceleration threshold of around 20 feet per second per second, with a wheel speed that is at least six miles per hour below the speed of the airplane, is designed into the skid detector circuit as a reference. Any time a wheel decelerates at a rate greater than this threshold value, a signal is sent to the amplifier and then to the control valve to dump the brake pressure. At this time, the skid detector sends a signal to the modulator which, by measuring the width of the skid detector signal, automatically establishes the amount of current that will continue to flow through the valve after the wheel has recovered from the skid. When the amplifier receives its signal from the modulator, it maintains this current, which is just enough to hold the flapper over to prevent the pressure being completely dumped, but will maintain a pressure slightly less than that which caused the skid. A timer circuit in the modulator then allows this pressure to increase slowly until another skid starts to occur and the cycle repeats itself.

CONTROL BOX Touchdown Protection Before the airplane touches down, the locked-wheel detector sends a signal into the amplifier, which causes the control valve to open the passage between the brakes and the system return manifold. This makes it impossible to land with the brakes applied. As soon as the airplane touches down, the squat switch opens and the wheel starts to spin up. By the time it reaches a speed of about 20 miles per hour, the voltage generated in the wheel speed sensor is high enough to cause the locked wheel detector to remove the touchdown control signal from the amplifier, and the control valve will allow full pressure to be applied to the brakes. Anti-Skid Protection When the airplane is on the ground and the wheels are turning more than 20 miles per hour, almost all of the antiskid control is by the skid detector and the modulator. The AC voltage is converted for comparator purposes to DC which is then a voltage reference for wheel speed. This voltage is compared to a reference voltage which has been set at the maximum deceleration rate for the aircraft. If any rapid reduction occurs during braking this will be reflected as a sudden fall in transducer output voltage. This information from the control box is sent to an anti-skid valve to release the brake, the wheel will then regain speed, the transducer voltage rises and the control box de-energizes the valve and the brake will then be re-applied.

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Fig. B - On touchdown, the squat switch removes the ground from the locked wheel arming circuit, and the wheel speed sensor generates a signal which allows the control valve to send full pressure to the brakes.

Fig. A – Before landing, the squat switch completes the locked

wheel arming circuit. The locked wheel detector sends a signal into the amplifier which supplies current to hold the control valve in the dump position.

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CATEGORY B1 - MECHANICAL Locked Wheel Protection

AC system

When the aircraft is operated on a wet or icy runway, the antiskid system will hold the wheels in the slip area. But if one wheel begins to hydroplane or hits a patch of ice and slows down to less than ten miles per hour, while its mated reference wheel is still rolling more than 20 miles per hour, the locked wheel detector gets into the act. The timer measures the width of the skid detector signal, and if it is more than about one-tenth of a second. It will send a "full dump" signal to the valve, and the valve will remain in the full dump position until the wheel spins back to more than ten miles per hour.

The control box for antiskid systems using AC sensors operates in the same way as those using DC generators, the only difference being the addition of a converter circuit. This circuit receives the varying frequency alternating current and converts it into a varying voltage direct current. The changes in the DC voltage follow exactly the frequency changes in the AC.

Low Speed Disarming When all of the wheels are turning at less than 20 miles per hour, the locked wheel arming circuit becomes inoperative, giving the pilot full braking action for low-speed taxiing and parking.

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Fig. A – When the airplane is on the ground and all wheels are rotating more than 20 miles per hour, the skid detector and the modulator provide signals for the amplifier.

Fig. C – The difference between the control box of an antiskid system using an AC wheel speed sensor and one using a DC sensor is in the converter between the sensor and the control circuit.

Fig. B – When the airplane is on the ground and all three wheels are rotating at less than 20 miles per hour, the locked wheel arming circuit is inoperative and the pilot has full control for low speed taxiing and parking

Rev. 00 Oct 2006

159 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

IN-FLIGHT TEST

SYSTEM TESTS

A part of the pre-landing cockpit check gives an indication of the condition of the entire antiskid system. With the airplane configured for landing, depress the brake pedals and the brake lights should remain off. This indicates that the control valves are holding the brakes fully released.

Because it is so important that the pilot know the exact condition of his brake system before he uses it, antiskid systems are equipped with test circuits and control switches. These allow the pilot to test the integrity and operation of the entire system and to disable the antiskid system without affecting the normal braking action if any faults are found.

Depress the test switch and the brake lights should come on. Depressing this switch sends a signal through the wheel speed sensors, simulating a wheel speed of greater than 20 miles per hour. If the system is operating properly, this will cause the control valve to direct normal pressure to the brake. The light should remain on as long as the test switch is held depressed.

GROUND TEST The integrity of the antiskid system can be tested on the ground before flight. With the antiskid control switch on, depress the brake pedals, and both the left and right brake lights should come on, indicating that all of the pressure from the brake valves is going to the brakes.

When the switch is released, the two brake lights should go out, indicating that the anti skid system is holding all of the pressure off of the brakes.

Now, with the brakes still applied, press the test switch and hold it for a few seconds. This sends a signal through the wheel speed sensors into the control box to simulate a wheel speed of more than 20 miles per hour. The lights should remain on.

Any time a system fails either the ground test or the in-flight portion, the antiskid system may be disabled without affecting the normal braking in any way.

Release the test switch, and the two brake lights should go out and stay out for a couple of seconds and come back on. This simulates a wheel lockup that causes a release of pressure, and then the pressure is restored. This test checks the continuity of all of the wiring and the operation of the locked wheel circuits, the amplifiers, and the control valves.

Rev. 00 Oct 2006

160 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Fig. A

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161 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

A wheel speed sensor that passes both the resistance and the polarity test and shows no physical signs of damage is probably good and any trouble must be elsewhere. Be sure the drive tube in the hubcap is tight, that there are no dents in the cap, and that the alignment rivet is properly seated in the recess in the wheel.

SYSTEM MAINTENANCE Before condemning the antiskid system for a brake malfunction, be sure that no relevant faults exist in the rest of the system. If the fault is identified to be of the anti-skid system then the most logical place to start troubleshooting is with the wheel speed sensor.

CONTROL BOX The control box is checked by the substitution method. Remove both of the connector plugs from the box and swap them left to right. If, for example, the trouble indication was originally on the left side of the airplane, when the leads from the box are switch and the indication remains on the left side, in all probability the trouble is not with the control box. But if the indication switches over to the right side, you have a good reason to suspect the control box, and it should be replaced with one that is known to be good.

WHEEL SPEED SENSOR The DC wheel speed sensor can be checked in-situ by first removing the wheel hubcap to expose the blade of the sensor. Secondly with the brakes applied and the antiskid switch on give the blade a sharp spin in its normal direction of rotation with a finger. It will not turn more than about 180 degrees; it is not the amount of rotation that is important, however, but the rate at which it is turned. If the system is operating properly, the brakes should momentarily release and then reapply. You can tell if this is happening by watching the disk stack; it should relax and then tighten back up.

Any time the leads have been switched, be sure that they are reinstalled on the proper receptacles and are properly secured before returning the aircraft to service.

If this "tweak" test does not cause the brakes to release, check out the sensor before going to any other part of the system. Remove the connector plug from the back of the sensor, and measure the resistance of the armature while rotating the blade through a full 360 degrees. Rotate the blade slowly and note the amount of resistance, as well as the evenness of the reading. It must be within the tolerance specified in the manufacturer's service manual. If the resistance is correct, then check for the proper polarity by placing the meter on its lowest DC voltage scale, with the positive lead to pin B, and the negative lead to pin A. Tweak the blade in a clockwise direction, as viewed from its drive end, and the meter should read upscale.

Rev. 00 Oct 2006

162 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL CONTROL VALVE If the trouble remains after checking the two devices that were the easiest to get to, all that remains in the antiskid system is the control valve. These valves are electro-hydraulic, and the trouble could be either electrical or hydraulic. The easiest check is the electrical resistance of the coil. Remove the connector plug and measure the resistance of the coil with an accurate ohmmeter. It should measure within the tolerance specified in the service manual. If the trouble has been traced to the control valve and it is not electrical, the valve must be removed, as the difficulty is most probably in the hydraulic portion of the valve. The extremely close tolerances used in the manufacturer of this valve make the use of absolutely clean fluid imperative. To be sure that no contaminants reach the inside the valve, a fifteen micron steel-mesh screen is installed in the line before the orifices. If this screen clogs, the valve may malfunction. Check with the manufacturer's service manuals to see if it is possible to replace this filter in the field. If it is allowed, follow the service instruction in detail. If any field servicing is allowed on the valve, it must be done in an area that is free from any contamination, and be sure to follow, in detail, the manufacturer's latest service information.

Fig. A

Rev. 00 Oct 2006

163 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

AUTO BRAKES Let's look at some of the components in the auto-brake system. INTRODUCTION

They are the:

An automatic braking system can be of tremendous value to pilots under normal and adverse landing conditions, for at any time prior to landing the pilot can select an aircraft deceleration rate to be achieved by the brakes.

-

selector panel,

-

auto-brake control box

-

solenoid valves

Once a setting has been made the auto-brake system on landing will smoothly apply the brakes to achieve the selected deceleration rate down to a complete stop, without any further action from the crew. AUTO-BRAKE SYSTEM The auto-brake system represents a major advancement in landing and take-off safety. It allows the pilot to concentrate on other activities such as reverse thrust. The auto-brake system utilizes the normal anti-skid and brake units, but instead of using pressure from the brake metering valves, hydraulic pressure is sent via solenoid valves which will allow a predetermined amount of pressure through the anti-skid valves to the brake units. Auto-brake is only available on aircraft equipped with electrical anti-skid systems. Figure A illustrates the inter-relationship between the brake pedals and the anti-skid in auto-brake mode.

Rev. 00 Oct 2006

164 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Fig. A

Rev. 00 Oct 2006

165 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL SELECTOR PANEL

AUTO-BRAKE CONTROL BOX

A typical selector panel is shown in Figure A. It consists of a solenoid latched switch which will hold in a selected position only if all the arming requirements are met. If the auto-brake system cannot be armed the switch will return to the disarm position and the warning light will illuminate to warn the pilot that the system will not function or is not armed.

Selection on the panel will send an electrical signal to the autobrake control box, which is normally in the aircraft's electrical bay. The control box is a microprocessor which commands the solenoid valve to port pressure to the brake units. Because the pressure must be gradually built up and released to prevent brake snatch and jerking, a time delay and an electrical ramp (Figure B) are used to achieve these ends. The terminology used to indicate the operation of the auto-brake is:

Let us consider the various switch positions on the auto-brake selector panel. 1, 2, 3, 4, Max Auto - These are preset deceleration rates which the auto-brake control box will signal to the solenoid valve to port pressure to the brakes.

-

ON RAMP

-

OFF RAMP

-

DROP OUT

OFF - The system is off. DISARM - The system is unable to be armed.

ON RAMP - A gradual build up of brake pressure to the amount required for the selected deceleration rate.

RTO - (Rejected Take-Off) This setting is armed on the ground prior to take off. If after a certain speed has been achieved during the take off run the throttles are retarded to idle, the brakes will automatically receive full pressure and bring the aircraft to a rapid stop. If the aircraft makes a normal take off, as soon as the air/ground sensor detects an air mode the RTO mode will be cancelled and the switch will move to OFF.

Rev. 00 Oct 2006

OFF RAMP - A gradual decrease in pressure down to zero at the termination of the landing run or cancellation of auto-brake. DROP OUT - Instantaneous pressure release to zero e.g. 'go around' mode.

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Fig. A

Fig. B

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL 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 must be the pressure allowed to the brake units. These valves are usually fitted in the undercarriage bays just upstream of the anti-skid valves. The solenoid valve will open when all the arming requirements have been met and the aircraft is on the ground. It is also this solenoid valve that 'snaps shut' on DROP OUT. The servo valve modulates the brake pressure to regulate the deceleration rate. The pressure switch is connected to a disarm warning light to monitor zero pressure when the auto-brakes are armed. Figure A shows a typical auto-brake solenoid valve.

Rev. 00 Oct 2006

168 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Fig. A

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169 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

BRAKE CONTROLS Most airplanes have the master cylinders mounted on the rudder pedals, although a few aircraft have the master cylinders mounted remotely from the pedals. Such a configuration utilizes a system of rods, levers, bell cranks, and cables to carry the mechanical energy to the master cylinders. Other control arrangements make use of heel brakes, operated by the pilot pressing the heel on the brake pedal, and a central hand-brake lever, which operates all brakes at the same time. PARKING BRAKE CONTROLS The parking-brake mechanism is a subassembly of the usual hydraulic brake system. The parking brake is usually applied by a pull handle or lever. When the brake pedals are depressed and the parking-brake lever is pulled back by hand, the brakes are locked in the ON position. Depressing the brake pedals or releasing a parking-brake handle, depending on the system, again releases the brakes. Depending upon the type of master cylinder used, depressing the pedals will either build up enough pressure to unseat the parking valve, or it will unload a ratchet-type parking lock. Note: With respect to parking brakes, the setting of these brakes when the main brakes are hot may cause serious damage. Hot brakes should be allowed to cool before the parking brakes are applied.

Rev. 00 Oct 2006

170 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

BRAKE INDICATIONS The following parameters are indicated in the cockpit with respect to brake systems. -

Brake Temperature

-

Brake Pressure

-

Brake Accumulator Pressure

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL BRAKE TEMPERATURE MEASUREMENT -

Brake unit temperature thermocouples.

is

typically

measured

-

The hot junction of the thermocouples is affixed to the brake unit; usually accommodated in a recess in the torque tube. This is to measure the heat stack temperature as accurately as possible.

Since the temperature measured are that of the torque tube and not that of the heat pack it takes some time before the correct brake temperature is displayed.

using

-

The cold junction is located in the indicator in simple systems.

-

In more complex systems the cold junction is located within a control box that is located on or near the landing gear. The control box converts the measured voltage into an analogue or digital signal and transmits it to be indicated in the cockpit.

Manufacturers’ manuals usually specify certain actions based on the indicated brake temperature. If the temperature is above a certain value take-off may need to be delayed unless the brake unit is force cooled. If the temperature is at the danger levels the aircraft should be taxied only to clear the runway and personnel should be kept clear of the landing gear as a possibility of a tire explosion exists. Unusually high brake temperatures, especially when compared with the indication of other brake units are a good indicator of a faulty brake unit.

INDICATION On aircraft equipped with conventional instruments brake temperature is indicated individually for each brake unit on a mechanical indicator with cautionary and danger areas marked. On aircraft equipped with electronic instruments brake temperature indication is in a numerical format on a display unit and cautionary and danger levels are indicated by the change in color of the numerical indication or by a symbol such as an arc appearing close to the indication of the affected wheel.

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172 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Fig. B

Fig. A

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173 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL BRAKE PRESSURE

BRAKE ACCUMULATOR PRESSURE

Brake pressure is not a standard indication of the brake system. Brake pressure, which essentially means the fluid pressure supplied to the brakes are indicated on aircraft where the crew is supposed to limit applied brake pressure to prevent an unsafe condition; Such as nose over of a tail-wheel type aircraft or to prevent skidding.

Aircraft equipped with power brakes supplied from hydraulic systems usually utilize accumulators to provide pressure for back-up or emergency brakes if the hydraulic system is not powered. Accumulators will also be utilized to maintain pressure in the brakes when the parking brakes are applied. MEASUREMENT

MEASUREMENT

A pressure transducer will measure the pressure of the brake accumulators and transmit the signal to the cockpit for display.

Brake supply pressure is usually measured by a pressure transducer downstream of the power brake control valve and the electrical signal is sent to the indicator directly or via a control box.

INDICATION Brake pressure is indicated utilizing mechanical indicators which show pressure of fluid in the accumulator.

INDICATION

Since the accumulator supplies the back-up or emergency brakes manufacturers’ specify a minimum accumulator pressure for departure, for setting the parking brakes and for brakes prior to an engine ground run.

Brake pressure is indicated utilizing mechanical indicators which shows the brake pressure supplied to the right hand and left hand brakes individually. A green arc will indicate the safe margin for brake application and amber or red for pressures to be avoided.

Rev. 00 Oct 2006

174 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

Backup Accumulator Pressure

Left Brake Pressure

Right Brake Pressure

Fig. A

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175 For Training Purposes Only

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Module 11 – AIRCRAFT AERODYNAMICS AND STRUCTURES Sub Module 11.13 – LANDING GEAR

CATEGORY B1 - MECHANICAL

PAGE INTENTIONALLY LEFT BLANK

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