Domein Techniek

Aviation Studies
Preface

In the first period of the second year of Aviation Studies, project group 2A2R was assigned to
investigate the landing gear of a modern commercial aircraft. The landing gear of the A320 has been
chosen. During this project the members investigated the operation, construction and materials of
the landing gear and landing gear systems, the forces and stresses on the landing gear construction
and the possible failures which can occur in the landing gear systems.

During this project the project group was supported by Pieter van Langen. The project group would
like to thank Pieter van Langen for the support and guidance of the group during this project.

Project group 2A2R; Stefanie van Doren, Paul Heslinga, Julien van Ingen, Niels van Klink, Michel van
Leeuwen, Eric de Veen en Rob Vermeij.

15 October 2009




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Index
Summary ................................................................................................................................................. 1
Introduction ............................................................................................................................................. 2
1. Landing gear A320 ............................................................................................................................... 3
1.1 General description ....................................................................................................................... 3
1.1.1 Tricycle gear layout ................................................................................................................. 3
1.1.2 Nose and main landing gear ................................................................................................... 4
1.2 Air/ground logic ............................................................................................................................. 4
1.3 Extension and Retraction .............................................................................................................. 5
1.3.1 Main gear ................................................................................................................................ 5
1.3.2 Nose gear ................................................................................................................................ 7
1.3.3 Free fall extension .................................................................................................................. 8
1.4 Steering .......................................................................................................................................... 8
1.4.1 Steering system ...................................................................................................................... 8
1.4.2 Limitations .............................................................................................................................. 9
1.4.3 Display .................................................................................................................................. 10
1.5 Wheel systems ............................................................................................................................. 10
1.5.1 Wheels .................................................................................................................................. 10
1.5.2 Tires ...................................................................................................................................... 10
1.6 A320 Brakes ................................................................................................................................. 11
1.6.1 Brake components ................................................................................................................ 11
1.6.2 Normal brake system ........................................................................................................... 12
1.6.3 Anti skid ................................................................................................................................ 13
1.6.4. Other brake systems ............................................................................................................ 13
1.6.5 Alternate brake systems ....................................................................................................... 14
1.6.6 Brake system display ............................................................................................................ 15
1.7 Damping system .......................................................................................................................... 16
1.7.1 Shock absorber ..................................................................................................................... 16
1.7.2 Torque link damper .............................................................................................................. 17
1.8 Regulations .................................................................................................................................. 18
1.8.1 Landing gear regulations ...................................................................................................... 18
1.8.2 Airworthiness regulations .................................................................................................... 19
2. Landing gear construction ................................................................................................................. 21
2.1 Rest .............................................................................................................................................. 21
2.1.1 Free body diagrams and assumptions .................................................................................. 21
2.1.2 Shear stress .......................................................................................................................... 25
2.2 Rejected take off ......................................................................................................................... 27
2.2.1 Free Body Diagram and assumptions ................................................................................... 27



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2.2.2 Motion equations ................................................................................................................. 29
2.2.3 Conclusion ............................................................................................................................ 29
2.3 Landing ........................................................................................................................................ 30
2.3.1 Free body diagrams and assumptions .................................................................................. 30
2.3.2 Landing force calculation ...................................................................................................... 32
2.3.3 Landing shear force .............................................................................................................. 32
2.4 Materials ...................................................................................................................................... 33
3. Landing gear operations .................................................................................................................... 35
3.1 MMEL & MEL ............................................................................................................................... 35
3.2 A320 failures ............................................................................................................................ 36
3.2.1 Nose landing gear failure ...................................................................................................... 36
3.2.2 Shock absorber failure .......................................................................................................... 36
3.3 Maintenance and costs ............................................................................................................... 37
3.3.1 Maintenance ......................................................................................................................... 37
3.3.2 Costs ..................................................................................................................................... 39
3.4 Conclusion ................................................................................................................................... 40
Bibliography ........................................................................................................................................... 41



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Summary
The project consists of three different analyses. In the first chapter the landing gear and the
subsystems of an A320 are described. The second one is an analysis of the landing gear forces in
different flight phases and the relation to material choice and in the third analysis an advice is given
to the board of Amsterdam Leeuwenburg Airlines if it is profitable to increase the maintenance, so
failures occur less frequently.

The analysis of the first chapter is build up like a landing, with a short general description as first
paragraph. The different systems are described in the sequence of a landing, what happens with the
landing gear between the approach and the stop at the gate. Only the shock absorber is an
exception, because this system is different due that it is not possible to control this system out of the
cockpit.
The landing gear used on the A320 it has a tricycle gear layout. The main landing gears are mounted
under the wing and retract sideways into the fuselage. The retraction of the landing gear legs is done
by hydraulic actuators. The main difference between the main gears and the nose landing gear is the
forward retraction motion into the fuselage. The information of the landing gear and systems like the
brakes are displayed on the electronic centralized aircraft monitor. In case of a failure the extension
can be done with a free fall extension. When the aircraft is on ground again, the nose wheel steering
system controls the direction of the aircraft. This control is needed during taxiing, landing and taking
off. To decelerate the A320 brakes on the landing gear can be used. During manual operation of the
brakes, the normal brake system is used. Each main landing wheel is equipped with one brake and to
get the most efficient use out of these brakes and to stop the aircraft from skidding an anti skid
system is installed. The shock absorber must soften the impact of the touchdown. The main landing
gear is in normal conditions the first part of the aircraft which touch the ground. The regulations for
this project are found in Certification Specification for Large Aeroplanes (CS-25), which is written by
the European Aviation Safety Agency (EASA).

The different flight phases that are compared are rest, rejected take off and landing. This comparison
is necessary to find the biggest load on the landing gear. The forces on the aircraft in this situation
are calculated using free body diagrams. When the aircraft is in rest, the forces are in equilibrium.
The maximum deceleration during a rejected take off does not only come from the brakes because
several other forces are involved. With the assumptions made for these forces, the friction force is
calculated. When an aircraft is approaching the forces of the aircraft are equilibrium, an extra
impulse is created during the touchdown. In these different flight phases the sheer stress is
calculated, which is needed for the choice of the landing gear materials. These materials have to be
strong enough to carry the loads, but the weight is another property that has to be taken into
account. Also some possible reasons are given for the difference of the calculations and the expected
values.

To advice the board of Amsterdam Leeuwenburg Airlines different aspects of containing
airworthiness and maintenance are described. The Master minimum equipment list and the
minimum equipment list are used to inform about the airworthiness of the aircraft when a failure
occurred. It also is used to increase profitability. The documents can be used as handbook to decide
if an aircraft can dispatch or need to stay on the ground and go into maintenance. Every failure that
can occur does not only have a direct impact on a certain system since many systems have
interaction with other systems, multiple failures need to be avoided at all cost since these result in an
unsatisfied level of safety. When the nose wheel steering is 90 rotated, one of the causes is an
incorrect installation, which can be prevented by better maintenance. Also poor maintenance can be
a cause of a shock absorber hydraulic leak. By a better maintenance programme the costs for the
maintenance are high, but the failure costs are decreased. In order to the project objective the
conclusion is that it is profitable to perform more maintenance, which leads to lower failure costs.

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Introduction

The first project for Aviation studies 2
nd
year is about landing gear. Project group 2A2R exists of seven
years Aviation students at the Hogeschool van Amsterdam, which analyses a current landing
system of a modern commercial aircraft. An analyses is made of the landing gear of the Airbus A320,
which is used by Amsterdam Leeuwenburg Airlines (ALA). ALA has a new board, this board wants a
analyze what kind of failures can occur in the landing gear and if it is profitable to increase the
frequency of maintenance, to prevent possible failures.
Beside this the project group need to analyse how these failures has a affect at the aircraft
airworthiness by means of the Minimum Equipment List (MEL).
The project is written according to the standard project layout. Because of this it is arranged in three
chapters in the sequence of the research, namely Landing gear A320, Landing gear construction and
Landing gear operations.

An exactly analysis of the landing gear, as designed on modern aircrafts, is necessary to understand
the construction and systems of the landing gear of the Airbus A320. Operation of the landing gear is
made possible through air/ground logic, extension/retraction, steering, braking and damping
systems. In order to maintain the safety of the aircraft, the design of the landing is satisfied to the
legislation of the European Aviation Safety Agency (1).

With the knowledge of the landing gear construction of the Airbus A320, the forces on the
construction calculated during different flight phases. In these flight phases the aircraft endures
several forces. The materials that are used depends on the forces at the aircraft (2).

Then, with a good insight of the A320s landing gear the project group is able to provide an in detail
overview of the common faults and problems of the A320s landing gear. These faults and problems
have consequences for the aircrafts airworthiness. Change in the aircrafts airworthiness requires
maintenance with inevitable costs for the airline (3).

The used main sources, serve as information to learn how the landing gear of the Airbus A320
operates. To give the report a good appearance the note of Tilly Wenzel (2008) is used. Furthermore
the following main sources are used: Technical training manual landing gear (2000), Static voor
dynamici (2006) and Aircraft maintenance manual (2004). A extensive bibliography can be found at
page 44
The pyramid model can be found in the appendices which are collected in a appendices book. The
division of the report can be found there. The appendices begin with the project assignment, the
pyramid model and the process report (appendix I-III).


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1. Landing gear A320
The landing gear of the A320 is controlled by several landing gear systems, but the basis lay-out is the
same as any other aircraft of the same class (1.1). There are several systems that control the landing
gear during different flight phases, these systems must somehow be activated at the right time. This
is done by the air/ground ground logic system (1.2).
The basis lay-out slowly evolved to a retractable lay-out which is standard for large civil aircraft these
days (1.3). The landing gear is also the only way to control the aircraft direction on the ground (1.4).
Due to the excessive use of the landing gear on the ground, the tires and wheels must be able to
endure this stress for longer periods of time. But if they need to be replaced, it must be a easy
routine job which does not require the disassembly of the landing gear (1.5).
One of the main causes of wear to the tires is braking, because the tires are kept on the skid limit
(1.6). Before the brakes are used during landing, the aircraft has made a touchdown. This touchdown
requires another landing gear system, namely the shock absorbers which reduce the impact on the
rest of the aircraft (1.7). Some of these systems are required by law, especially for civilian aircraft
(1.8).
The source that is most used for this chapter is: Technical training manual (1999-2000)

1.1 General description
The landing gear or undercarriage of the A320 has four main functions. It separates the aircraft from
the ground, allows the aircraft to maneuver on the ground with a steerable nose wheel, it softens the
shock during landing and slows the aircraft down by using the brakes. The landing gear used on the
A320 it has a tricycle gear layout (1.1.1). It has one nose landing gear leg and two main landing gear
legs that are placed under the wing and retract into the fuselage (1.1.2).

1.1.1 Tricycle gear layout
The tricycle gear (figure 1.1) consists of one nose landing gear leg (1) and two main landing gear legs
(2). When using a tricycle gear, around 13% of the total weight of the airplane acts down on the nose
landing gear and 87% acts down on the main landing gear. The weight distribution depends on the
position of the center of gravity. The center of gravity on the A320 must be between 17% and 38,5%
on the Mean Aerodynamic Chord (MAC). The location of the MAC can be calculated using the
aircrafts dimensions (appendix V).


Figure 1.1 Airbus 320 Landing gear layout

The advantage of the tricycle gear is the stabilizing momentum during touchdown when the aircrafts
longitudinal axis is not parallel to the runway. The drag created by the two main gear wheels is
parallel to the motion direction (figure 1.2). This means there is no momentum created however the

1 Nose landing gear

2 Main landing gear

1
2

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drag that works on the two main gear wheels to the side creates a negative momentum. This
momentum turns the aircraft parallel to the runway again.















Fig 1.2 Tricycle gear

Another advantage of the tricycle gear is that the aircraft is in a horizontal position when parked on
the ground. Unlike the conventional landing gear configuration which has a tail wheel instead of a
nose wheel. With the tail wheel causes the airplane to stand in a nose up position therefore making
loading the airplane more difficult.

The disadvantage of the tricycle gear is that it is heavier than a conventional configuration. Also
when loading an aircraft equipped with a tricycle gear you cannot load the aircraft in such a way that
the center of gravity is shifted to behind the two main gear wheels. This causes the aircraft to tip
over and hit the ground with its tail.

1.1.2 Nose and main landing gear
The A320 uses a four bar linkage on both nose and mean landing gears (figure 1.3). The four bar
linkage is used. This four bar linkage has one degree of freedom it can extend (1) and retract (2).
When using the four bar linkage it does not matter if it is retracted straight or retracted sideways.



1 Extended four bar linkage
2 Retracted four bar linkage
Figure 1.3 extended and retracted four bar linkage

1.2 Air/ground logic
The air/ground logic system determines whether the aircraft is on the ground or in the air. This
system is implemented to activate or disable systems during different phases of the flight. The brakes
for instance cannot be activated before landing because the aircraft would skid immediately during
landing.

The system consists of proximity sensors (appendix VI). When the aircraft is on the ground both main
landing gear struts are compressed by the weight of the aircraft, so the proximity sensors transmits a

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signal that the aircraft is on the ground. During a take-off, when the weight of the wheel and bogie
assembly extends the oleo, the proximity sensors provided a signal that the aircraft is in the air. The
signals from the proximity sensors are going to the Landing Gear Control and Interface Unit (LGCIU).
The LGCIU gives an output if the aircraft is on the ground or in the air.

1.3 Extension and Retraction
The landing gear is used to soften the impact during landing. The main gear is the first part that hits
the ground (1.3.1). Beside the main gears, the nose gear is used to stabilize the aircraft on the ground
(1.3.2). These two systems work with electric and hydraulic components, but in case of a failure an
alternate way to control the landing gear is used (1.3.3).

1.3.1 Main gear
The main landing gears are mounted under the wing and retract sideways into the fuselage. The
retraction of the landing gear legs is done by hydraulic actuators. The gears are kept retracted in the
landing gear bay by an up lock mechanism. A lock stay is used for the down locking of the main
landing gear (A). The landing gear is electrically controlled and hydraulically operated by the green
hydraulics system of the A320. Two LGCIUs control and monitor the landing gear and landing gear
doors. The hydraulic actuators move the landing gear to its up/down position (B). The landing gear
indication panel and the Electronic Centralised Aircraft Monitor (ECAM) display the position of the
landing gear during extension and retraction (C).

A. Mechanical construction
The main landing gear of the A320 consists of two inboard retracting main gear legs. The gear is
fitted on both wings, retraction goes side-ways into the landing gear bay that is situated in the
fuselage. A two-piece side stay assembly holds the gear in extended position. The side stay assembly
connects the leg fitting of the mean landing gear to the wing structure, this will prevent that the
main landing gear is making sideways movements. A lock spring is attached between the side stay
and the leg.

When the mean landing gear retracts (appendix VII) the lock actuator (1) pulls the lock stay (2) so the
lock spring (3) starts to fold. The hydraulic actuator (4) extends and retracts the main landing gear
sideways into the fuselage.

The gears are locked in the fuselage by an up lock mechanism (appendix VIII). The up lock is done by
a hook (1), which locks the pin (2) mounted on the landing gear. The hook is controlled by the up lock
actuator (3) and a tension spring (4) holds the hook in an over-centred and locked position. When the
up lock actuator retracts, it moves the hook out of the up lock which unlocks the up lock mechanism.

B. Operation
The landing gear in normal operation is controlled by the landing gear control lever, situated at the
first officers side of the instrument panel. The landing gear lever in UP-position states that the
landing gear is retracted and when the lever is in DOWN-position the gear is or will be extended.

The first officer puts the landing gear lever in the DOWN-position when the aircraft is going to land
and the aircraft is flying under the 260 knots. A safety valve that is connected to the Air Data Inertial
Reference Unit (ADIRU) blocks the pressure of the system when the aircraft is flying over 264 knots.
The landing gear lever moves in a slot and internal locks in the unit keep the arm in the fully UP- or
DOWN-position. The lever sends electrical signals, these signals are created by the switches inside
the unit. A solenoid-operated baulk mechanism, which is a locking device that is controlled by
electronic signals from the LGCIU, prevents the retraction. When the shock absorbers are not fully

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extended before the LGCIU sends an electrical signal to release the baulk mechanism. The landing
gear control system consists of two independently working electrical circuits.

The electrical signal is sent to the LGCIU. The LGCIU is an interface unit based on ARINC that controls
and review the sequence of the landing gear. The landing gear control system has two independently
working LGCIUs, the LGCIU will be switched at every sequence, when extension starts LGCIU1 is
working and LGCIU2 when retraction starts. Only one LGCIU controls the system at the time. When a
LGCIU does not work anymore the other LGCIU takes the control of the landing gear over.

The LGCIU sends a signal to the landing gear and landing gear doors electrical selector valves
(appendix IX) to open the doors. The two selector valves are mounted on a manifold block. Both
selector valves have two operating solenoids A and B. If the doors solenoid B (1) energizes the ball
valve (2) stays at his position to block the valve (3). This stops the hydraulic fluid to stream back to its
system it creates a higher pressure on the left side of the end cap (4). So the spool (5) will moves to
right to open port 2 (6), which let the hydraulic fluid stream to the door actuators to open the doors.

The up lock from the door releases and the hydraulically operated actuating cylinder opens the main
door. The solenoid B of the doors is energized until the gears are locked in down position. The
proximity sensors on the door detect the open door and will send this back to the LGCIU.
The LGCIU receives the signal from the proximity sensor that the doors are open. The LGCIU sends a
signal to the landing selector valves. The landing gear selector valves works in the same way as the
selector valve of the doors. The fluid goes through to the main landing gear actuating cylinder.
The main landing gear actuating cylinder (figure 1.4) has three primary contents: The cylinder (1), the
piston (2) and two restrictor valves (3). When the piston rod is in damping mode the piston blocks
the control orifice (4). So when port A (5) is open for the hydraulic fluid and port B (6) for return of
the hydraulic fluid. The hydraulic fluid enters port A, the pressure will directly create the pressure to
the restrictor valve, the control orifice is blocked, so the restrictor valve is moved to let fluid goes
through. This creates the pressure on the piston, which will move the piston to the right. The
hydraulic fluid, with a lower pressure, at the other side of the piston will return in the hydraulic
system. The two-restrictor valves are closed when the landing gear is operated. This decreases the
flow and controls the speed.

1. Cylinder
2. Piston
3. Restrictor
valve
4. Control
orifice


Figure 1.4 Landing Gear Actuator

The doors close, and lock, and the doors solenoid A stays energized whenever electrical power is
applied. The LGCIU will receive an electrical signal from the proximity sensors that the doors are
closed. The LGCIU will send this information to the ECAM.

C. Indication
The indication of the landing gear in the cockpit is done in two ways (appendix X):
A
B
2
1
3
4

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1. Landing gear indicator panel
2. ECAM

Ad 1. Landing gear indicator panel
The landing gear indicator panel indicates the landing gear positions monitored by LGCIU1. The
landing gear indicator panel can only be used as back up.

Ad 2. ECAM
The ECAM is the primary device to display landing gear position and indication (appendix XI). The
landing gear indicator panel only gets information of LGCIU1, ECAM receives information from both
LGCIUs. Both indications of the LGCIUs are displayed on the ECAM. The ECAM gets information about
the position of the landing gear, locked or unlocked. ECAM also receives information from the
position of the doors.

1.3.2 Nose gear
The main difference between the main gears is the forward retraction motion into the fuselage (A).
The nose landing gear is also electrically controlled and hydraulically operated by the green hydraulic
system. The difference with the main gear is the neutral position check by the LGCIU (B). The landing
gear indication panel and the ECAM display the position of the landing gear during extension and
retraction (C).

A. Mechanical construction
The nose gear is equipped with two wheels and retracts forward into the front fuselage of the
airplane (appendix XII). A two-piece drag strut assembly with a lock stay keeps the gear locked in
extension position. A hydraulic actuator in normal operation controls the lock stay. If the hydraulic
actuator has a failure the two springs will pull the gear down for down lock.

When the nose gear retracts the hydraulic cylinder of the lock stay retracts which will fold the lock
stay. The nose gear actuator retracts the gear into the front fuselage.

The nose gear up lock holds the gear in retracted position. The main components of the up lock are
the hook and the lever. The hook locks the pin, which is mounted on the nose landing gear leg, for
locking the nose gear. The hydraulic actuator pushes the lever forward, the lever pivots by the
springs.

B. Operation
The nose landing gear operates almost identically as the main landing gear. The landing gear lever is
used for every gear of the aircraft. The LGCIUs also control and monitor the nose landing gear. When
the nose landing gear is extracted normally the gear is in the neutral position, if that fails the LGCIU
sends no electrical signal to the solenoid baulk device, so the lever cannot be put in UP- position.

The gears and doors are both operated by the LGCIU, the hydraulic operation is done by a selector
valve. Hydraulic actuators control the gears and doors and the proximity sensors that are placed on
the lock stay and the up lock do the position information.

C. Indication
The position of the nose landing gear is displayed on the landing gear indication panel and on the
ECAM. The position indication works the same as the main landing gear.


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1.3.3 Free fall extension
The free fall extension is controlled by a gravity extension crank handle in the cockpit. When this
handle is operated, it activates a mechanical cable system that releases the gear door up lock
(appendix XIII). In normal conditions the latch assembly (1) cannot rotate, because of the hydraulic
power from the actuator (2). The target (3) at the end of the latch assembly makes contact with the
proximity sensor (4). Now the LGCIU indicates that the landing gear doors are closed. The spring
mechanism (5) and the manual release lever (6) are also locked by the latch assembly.
When the gravity handle is turned, the hydraulic power of the actuator is released from the landing
gear system by the cut off valve and the manual release lever rotates by a mechanic cable
movement. By this movement the latch assembly is also rotating, which has the result that the target
is not contacting the proximity sensor and the forces of the spring let the hook (7) rotate. Now the
door is unlocked and will open by gravity forces.

The gear release (appendix XIV) almost works the same as the door release. In normal position the
hydraulic pressurized actuator (1) holds the latch assembly (2) in the same position. The target (3) is
contacted to the proximity sensor (4), so an electrical signal locked is given to the LCGIU that the
gear is retracted. By the hydraulic pressure the latch assembly and the hook (5) are not able to
rotate. Also the tension spring (6) is in a constant position.
In case of using the gravity extension crank handle the release lever (7) is rotated clockwise by
mechanic cables. Through this rotation the bearing (8) of the latch assembly is lifted and the spring
force rotates the lock hook. Also the connection between the target and the proximity sensor is
broken, resulting in an unlocked signal.

The nose landing gear is lowered in extended position by aerodynamic forces. This is possible due the
fact that the nose gear extends in the direction of the air flow. The main gears are lowered by gravity.
The main gear is locked by spring lock forces, the nose gear is locked by aerodynamic forces. These
spring locks are not the same as the down lock, which is hydraulically powered. These spring locks
are mechanical locks.
Normal landing gear operation is restored after rotating the crank handle three times counter
clockwise.

1.4 Steering
During taxiing, landing and takeoff, the pilots need directional control of the aircraft. The nose
landing gear steering system controls the direction of the aircraft on ground (1.4.1). The steering
system has limitations to prevent dangerous situations or damage to the aircraft (1.4.2). The ECAM
display shows information about the steering system (1.4.3).

1.4.1 Steering system
The steering system is hydraulically powered and controlled by an electrical servomechanism from
the cockpit via the Braking and Steering Control Unit (BSCU). The steering system (appendix XV) is
controlled by two hand wheels during taxi, or with the rudder pedals during landing and takeoff. The
pilots give steering commands from the cockpit. The signals are processed by the BSCU (A) which
controls the hydraulic block (B). The hydraulic block powers the steering actuator (C) that turns the
nose landing gear.

A. Braking and Steering Control Unit
The BSCU gets commands from the hand wheels or the rudder pedals. The orders are converted to
an electrical signal and sent to the BSCU. The orders are processed before they reach the hydraulic
block.



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1. Aircraft speed
2. Anti skid and nose wheel steering switch
3. LGCIU
4. Towing control lever

Ad 1. Aircraft speed
The aircraft reference speed for the nose wheel steering is received by the ADIRUs. The information
is used for a correct steering angle at the relative aircraft speed (appendix XVI).

Ad 2. Anti skid and nose wheel steering switch
The anti skid and nose wheel steering switch, which is placed in the cockpit, must be in the ON
position to send the signals to the hydraulic block.

Ad 3. LGCIU
When the nose landing gear is extended and the nose gear doors are closed, the hydraulic block is
pressurized. Also the main landing gear has to be compressed. When both demands are met, both
chambers of the steering actuator are supplied and the nose wheels return to an angle of 0 for
landing.

Ad 4. Towing control lever
The towing control lever (appendix XVII), which is located on the strut of the nose landing gear, must
be in the NORMAL position. Only then it is possible for the pilots to manoeuvre with the aircraft.
The towing control lever has to be deactivated for towing purposes. When it is deactivated, the BSCU
will not send orders to the hydraulic block. It is one of the safety measures in the steering system.
Because of this it is impossible to damage the steering system of the nose wheel landing gear.

B. Hydraulic block
The BSCU sends signals to the hydraulic block (appendix XVIII, 1). The hydraulic block is placed at the
rear of the nose landing gear, and is powered by the hydraulic green system (2). With a pressure of
4000 Pressure per square inch (Psi) from the hydraulic green system, it is not possible to accurately
control the steering actuator (3), because the bypass valve is opened and tries to decrease the
pressure in the hydraulic block. The hydraulic block calculates the correct pressure by using the signal
from the BSCU. The adjustable diaphragm (4), located on each output line of the servo valve (5),
adjusts the flow to the actuating cylinder chamber and thus the wheel steering speed.

C. Steering actuator
With a hydraulic flow from the hydraulic block, the steering actuator pushes the rotating tube, which
is part of the nose landing gear structure, via a rack-and-pinion assembly to the correct position.
There is one anti-shimmy valve (6) per steering actuator camber, which prevents that the steering
wheel is influenced.

1.4.2 Limitations
Limitations are set to prevent damage at the nose wheel steering system (appendix XIX). There are a
few maximum steering angles depending on a few factors. When the aircraft is towed, the maximum
steering angle is 95. When the aircraft is taxiing and uses the hand wheels for steering, the
maximum steering angle is 74. When the aircraft is steering with the rudder pedals the maximum
steering angel is set on 6. These limitations are set in the BSCU. By using the alternate gear
extension, the hydraulic pressure is lost. As result of this the steering is lost.


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1.4.3 Display
The ECAM WHEEL page displays the steering indications. The steering indication appears amber in
case of green hydraulic failure, nose wheel steering failure or when the anti-skid and nose wheel
steering switch is set to OFF.

1.5 Wheel systems
The six tires on the A320 are mounted on divided wheels (1.5.1). The tires have different zones and
have a complicated design (1.5.2).

1.5.1 Wheels
The A320 has divided wheels because the tires are mounted tight on the wheel for an air tight seal.
Replacing the tires would be impossible without the divided wheel because pulling the tire over the
wheel is impossible. The divided wheel consists of two forged light alloy halves and is connected with
high tensile steel bolts and nuts (figure 1.5, 1). The tires are mounted directly on the wheel, so an
airtight seal is needed between the two halves. This seal is made by using an O ring (2). The tire is
filled with nitrogen, because by using this, the air pressure is also maintained for a longer period of
time. The inflating of the tire is done using an inflation valve (3). When braking the aircraft, heat is
building up in the brake packs. This heats up the wheel and tire causing high pressure and result in
the possibility of tire blowout. To prevent tire blowout due to heat two sets of fuse plugs have been
installed. The first set placed inside the wheel keys melts at 300 C (4). The second set placed on the
wheel web blows out at 183 C (5).
Figure 1.5 Wheel construction

1.5.2 Tires
Tires are made of natural rubber because of its better elastic properties during temperature changes.
The downside of natural rubber is its poor resistance against fuel such as kerosene.

Tires can be divided into different sections. Each of these sections has a different function.
Underneath all of these sections is the tire carcass. The carcass is used to hold the form of the tire
and it gives the tire the strength to withstand the high pressure inside. The carcass is made out of
rubber and cotton, rayon or nylon cords. These cords are not woven together but arranged parallel
to each other. They are held together by a thin layer of rubber. This also prevents other plies from
cutting each other. The plies are placed at a 90 angle of each other for added strength. When the
carcass can be seen the tire has to be replaced.

1 Steel nuts and bolts
2 O ring
3 Inflation Valve
4 Firs set of fuse plugs
5 Second set of fuse plugs
1
2
3
4
5

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The parts that make contact with the runway are the crown and shoulder (figure 1.6, 1-2) that wears
out the fastest. Therefore this part has a thick layer of rubber over it. In this layer the tread of the tire
is added to avoid aquaplaning. To protect the casing of the tire, breaker strips are also imbedded in
this layer. When due to tire wear these breaker strips are visible it is time to replace the tire. When a
tire has to be replaced, it can be recovered for reuse if the carcass is not damaged.

The shoulder area does not only make contact with the runway but a shine can also be added in this
area for a nose wheel tire. This is done to direct water away from the engine intakes to prevent
engine shutdown due to water ingestion. Above the shoulder area is the sidewall area (3). In this
section there is the most friction, because when braking the tire is on a skidding limit, so it
sometimes skids. The last section of the tire is the tire bead (4). In the bead of the tire steel wire
cords are added to keep the form of the tire. Without these cords the tire could not be placed on the
rim.


1 Crown
2 Shoulder
3 Sidewall area
4 Bead

Figure 1.6 Tire sections

1.6 A320 Brakes
The A320 has brakes on the main landing gear only (1.6.1). During manual operation of the brakes,
the normal brake system is used (1.6.2). To get the most efficient use out of these brakes and to stop
the aircraft from skidding an anti skid system is installed (1.6.3). However some phases of the flight
or on the ground require a different operation of the brakes (1.6.4). If a failure occurs in one of these
brake systems, the A320 has several alternate brake systems, so it is still able to decelerate or stop
safely to prevent other damage to the aircraft and its load (1.6.5). The pilots must know when the
normal systems or one of the alternate systems is active and where the failure is located, this is
projected on several displays in the cockpit (1.6.6).

1.6.1 Brake components
The brakes of an A320 are hydraulic multiple disc brakes. Each main landing wheel is equipped with
one brake, so this means that there are a total of four brakes on the aircraft. The braking principle is
based on friction between two different parts of the brake. The plates that are used for the brakes
are made of carbon because carbon wears less than normal steel. The multi disc brake (figure 1.7)
consists of a piston house (1), which is connected to the axe of the tire. Each brake consists of
fourteen pistons (2), seven of these are used for the normal braking and the others are used for the
alternate braking. In the hydraulic system there is also a bleed valve (3). The purpose of this valve is
to get rid of air in the hydraulic fluid. The carbon heat packs (rotor) (4) is attached to the wheel and
rotate together. The carbon plates (stator) (5) are attached to the piston house. The measurement of
the wear pin indicator (6) is based on the amount of extension of the pin. The further the extension is
the less wear there is.
The pistons driven by hydraulic fluid push the carbon heat pack on the carbon plates, the green
system is used as normal and yellow as alternate. When the heat pack and the plates are compressed
1
2
3
4

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by the pistons, friction is created between these two parts and the aircraft slows down. By creating
friction between the two parts, the temperature increases. Therefore a temperature sensor (7) is
added to the brakes. The brakes are not allowed to get a higher temperature then 300C when the
brake fans are off. In order to lower the temperature of the brake fans can be enabled. The
maximum temperature of the brakes is 150C in this case. The brake temperature indicator measures
the potential difference between the two different metals.



1. Piston house
2. Pistons
3. Bleed valve
4. Carbon heat pack (rotor)
5. Carbon plates(stator)
6. Wear pin indicator
7. Temperature sensor


Figure 1.7 Multi brake disc

1.6.2 Normal brake system
To control the normal brakes, the system has an electronic system (A) and a hydraulic system (B).

A. Electronic brake system
The electronic system (appendix XX) is used to control the hydraulic system and to collect and
process the feedback it gains from the hydraulic system. When the brake pedals are used, the
mechanical signal of the pedals is transformed into an electronic signal by the brake pedal
transmitter unit. The electronic signal is send to the Brake and Steering Control Unit (BSCU).
The BSCU is connected to the LGCIU, the ADIRU and the Spoiler Elevator Computers (SECs) that
supply data to the BSCU like wheel rotation speed and deceleration/acceleration data. The BSCU
consists of two exactly the same systems which work simultaneously (figure 1.8). If one system fails,
the BSCU can still operate with the other one.


1. System one
2. System two
Figure 1.8 Brake and Steering Control Unit
3
2
1
5
4
6
7
1
2

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B. Hydraulic brake system
The BSCU processes the electronic signal and activates the selector valve which allows green
hydraulic pressure into the system (appendix XXI). The pressure goes through the throttle valve to
the automatic selector valve which puts the yellow system on standby and allows green pressure into
the brakes manifold. The throttle valve prevents quick returning of the automatic selector valve
when it is de-energized. This is to avoid instantaneous release of the brakes.
The normal brake manifold consists of two normal servo valves and a hydraulic fuse, there is one
normal manifold for each wheel. Depending on the brake force needed by the pilots, the BSCU sends
a certain current to the normal servo valves. These valves convert this current to a certain amount of
braking pressure by which braking is realized. The hydraulic fuse plugs the line in case of a leakage.
Two master cylinders are also present to give an artificial feel at the braking pedals.

1.6.3 Anti skid
The purpose of the anti skid is to keep the maximum braking efficiency on each wheel in order to
avoid a skid situation. The BSCU provides anti-skid control during normal and alternate braking, the
yellow or green system is used for this. The BSCU provides this information to two systems, system
one or system two, which work simultaneously.
The speed indication of the wheels is giving by the tachometer. The tachometer measures the
rotation speed of the wheels. The tachometer consists of two rings. One ring is attached to the axle
and the other is attached to the wheel itself. The rotation speed is measured by variations in the
induction frequency, this variations appear because a piece of metal moves along the ring that is
attached on the axle.
This speed is compared to the speed that is supplied by the air date inertial reference unit. As soon as
the indication of the tachometer goes below 0.87 of the speed that is provided by the ADIRU, the
brakes are released. When the speed of a single wheel comes below the speed of 0.87 this single
wheel, orders are given to release the brakes on this wheel to avoid skidding.
A situation where a form of skidding happens pretty often is hydroplaning, what happens in this case
is that a layer of water comes between the tires and the ground. The aircraft will lose traction and
because of this it will not slow down when the brakes are used. Anti skid can be disabled by putting
the A/SKID & N/W STR switch to off.

1.6.4. Other brake systems
In order to get the best braking efficiency during landing the auto brakes can be used instead of
manual braking (A). The auto-brakes can only be used during movement of the aircraft, but the
aircraft needs a parking brake too so it cannot roll of the platform or roll away during loading at the
gate (B). Another problem occurs after take-off when the wheels are still spinning but the pilots want
to retract the landing gear. If the pilots do so with spinning wheels it could cause structural damage
to the aircraft and excessive wear to the tires, to prevent this, the A320 is equipped with in flight
brakes (C).

A. Auto-brake
The auto-brakes (appendix XXII) are mainly used during landing and Rejected Take Off (RTO). This
system automatically brakes when two of the three ground spoilers are deployed. It has three
modes: LO, MED and MAX (table 1.1) which pilots can select in the cockpit on the AUTO BRK
panel. By selecting an auto-brake mode, the mode is also activated in the BSCU. The braking is
realized the same way as during normal braking after the BSCU. The Air Data Inertial Reference
System (ADIRS) is monitoring the amount of actual deceleration and sends the data to the BSCU
which compares this data with the selected deceleration mode and regulates the amount of braking
pressure.

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Auto-braking can be stopped by putting the mode switch on the OFF-position, retracting the
ground spoilers or applying sufficient force on the pedals.
The auto-brakes can be used only when green pressure is available, anti skid is ON and no failures
in the normal braking system.


* After deployment of two of the three ground spoilers
** During take-off MAX mode is always selected to be prepared for a possible RTO
Auto-brake mode LO MED MAX
**
Deceleration 2 m/s
2
3 m/s
2
Maximum breaking
pressure
Deployment
*
4 sec 2 sec immediately
Table 1.1 Different auto-brake modes

B. Parking brake
The parking brake system (appendix XXIII) is used when an aircraft is parked, like a parking brake in a
car. Pilots can turn the parking brake on by setting the parking brake switch to the ON-position.
The BSCU shuts down all other brake systems and anti skid system and the parking brake control
valve is energized. The automatic selector is energized and cuts of the green system and allows
yellow pressure into the parking brake system. Pressure could also be supplied by the brake
accumulator which can give a maximum of 145 bar pressure and holds pressure for up to 12 hours.
After the pressure passes the parking brake control valve, it reaches the dual shuttle valve which
allows the alternate brake system to be supplied by the parking brake system. When passed through
the dual shuttle valve, the parking brake is realized by the alternate braking manifold which contains
two alternate braking servo valves and two hydraulic fuses. There is one alternate manifold for each
wheel.

C. In flight braking
Before the landing gear is fully retracted the wheel must stop rotating, this is done by in-flight
braking. The wheels of the main landing gear are stopped by releasing a preprogrammed braking
pressure. This pressure is released on the normal brakes when the gear lever is in the UP-position
for three seconds or when the nose gear is no longer down locked.
The wheels of the nose gear are stopped by using a brake band just before they are fully retracted.

1.6.5 Alternate brake systems
When there is a problem with the normal brake system, the alternate system is automatically
switched on. In order to still brake as efficient as possible anti skid is still active (A). When the anti
skid function is lost, the aircraft needs a longer distance to decelerate because the pilots need to
brake more careful to prevent skidding (B). If this alternate system fails, the parking brake could be
used as an emergency brake (C).

A. Alternate braking with anti skid
When green pressure is lost the automatic selector pressurizes the alternate system (appendix XXIV)
with yellow pressure. The mechanical signal from the pedals is no longer converted into an electronic
signal but directly into low hydraulic pressures by the auxiliary low pressure control system. The
pressure goes to the dual valve which converts this in the proper amount of braking pressure for the
left and right-hand brakes using yellow pressure. This pressure passes the dual shuttle valve which is
now in alternate selection instead of parking brake selection. Braking is realized by the alternate
braking manifold. The anti skid is regulated by the BSCU using the alternate brakes.


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B. Alternate braking without anti skid
Anti skid is no longer available when the anti skid is turned off, when there is a power supply failure
or when only accumulator pressure is available. If the alternate system is supplied by the
accumulator only, then only seven full brake actions can be performed. The anti skid is not available
anymore, pilots need to reduce the amount of braking to avoid skidding. This means that the aircraft
decelerates slower and needs a longer distance to reduce speed.

C. Emergency parking brake
As a last resource, the parking brake could be used in an emergency. The dual shuttle valve selects
the parking brake system which then supplies the alternate brakes and brake pressure is released.
This is only for emergency situations because it can cause skidding, overheating and damage to the
wheels and brake systems.

1.6.6 Brake system display
The information regarding the braking system is displayed on the electronic centralized aircraft
monitor (figure 1.9). The alternate braking (1) will turn green if this system is operative. The A/SKID
(2) will turn amber in case of a failure of the BSCU or if the BSCU detects another anti skid failure. The
auto brake (3) will turn green when the auto brake is on, below this is displayed what mode is
selected LO, MED or MAX. The wheel numbers (4) give an overview of every specific wheel that
is a part of the landing gear. The release indicators (5) give an indication when the landing gear is
extended, which appear just after touchdown to indicate that the antiskid system is working
correctly. The temperature (6) on the ECAM system gives an indication of the each brake this number
is displayed in green, it will turn to amber when a wheel exceeds 300C.


1. Alternate brake
2. Anti Skid
3. Auto brake
4. Wheel number
5. Release indicators
6. Brake temperature



Figure 1.9 ECAM system

There is also a yellow pressure triple indicator (figure 1.10) located in the cockpit. This instrument
shows the pressure that is still left in the accumulator (1) and the pressure on the left (2) and right-
hand (3) alternate brakes.
6
4
2
1
5
3

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1. Accumulator pressure
2. Left-hand brake pressure
3. Right-hand brake pressure
Figure 1.10 yellow pressure triple indicator

1.7 Damping system
During the landing, the landing gear endures huge forces. The shock absorbers soften the forces that
are involved (1.7.1). Torque links are placed at the rear of the nose landing gear. Torque link dampers
decrease the vibrations through the torque links during a landing (1.7.2).

1.7.1 Shock absorber
The shock absorber must soften the impact of the touchdown. The main landing gear is in normal
conditions the first part of the aircraft which touch the ground (A). The nose landing gear is the last
part of the landing gear which touches the ground (B).

A. Main landing gear
The shock absorber makes the touchdown more smoothly and gradually. The shock absorber is a
telescopic oleo-pneumatic unit. Oleo-pneumatic means that the shock absorber works with hydraulic
fluid and pneumatics combined. The reason that hydraulics and pneumatics are combined is that if
there are only pneumatics, the aircraft gets a recoil. With only hydraulics, there will be no
compression at all. When the shock absorber compresses, the forces of the touchdown are
transmitted to the hydraulic fluid and nitrogen gas. The recoil stroke is slow to make sure that the
aircraft does not become airborne again.
The shock absorber is a two stage unit and contains four chambers (appendix XXV).

First stage gas chamber contains a gas at low pressure and some hydraulic fluid (1).
Recoil chamber that contains hydraulic fluid (2).
Compression chamber that contains hydraulic fluid (3).
Second stage gas chamber contains a gas at high pressure (4).

The damping tube (5), which contains the first stage orifice (6), attaches to the head of the second
stage cylinder and has a fluid connection. The movement of the damping tube through the orifice
block decreases the fluid flow in the first stage damping. This increases the damping effect. A floating
piston (7) in the second stage cylinder separates the hydraulic fluid of the compression chamber and
the gas of the second stage chamber. During compression, the floating piston does not move down
until the gas pressures of the first stage and the second stage chambers are equal.

As noted earlier, the shock absorbers of the main landing gear are running a few stages when the
aircraft makes a touchdown.

1. Compression
2. Recoil
3. Compression and recoil second stage gas chamber


1
2
3

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Ad 1. Compression
During the compression, the sliding tube (8) slides into the main fitting (9). The volume of the shock
absorber reduces, which compress the gas.
The fluid flows from the compression chamber to the first stage gas chamber, through the first stage
orifices (10) of the damping tube. While the shock absorber compresses, the damping tube and the
first stage orifices go through the damping head (appendix XXVI). The flow through the first-stage
orifices stops, and the flow limits to that through the compression-orifice plate. This produces a two-
stage damping effect. The gas compression in the second stage chamber, pushes on the floating
piston to help the damping. It also helps to make the damping effect and the compression effect of
the oleo smooth.
At the same time, the fluid goes from the first stage gas chamber into the recoil chamber, through
the openings in the upper bearing. This flow of fluid moves the recoil-orifice plate against the flange
of a retaining ring, to let the fluid flow fully. The gas compression and the fluid transfer absorb the
shock-loads from the main landing gear.

Ad 2. Recoil
There is still energy in the gas of the first and second stage gas chambers. This gas wants to expand.
The fluid goes through the recoil-orifice and the compression-orifice plates. The flow of fluid moves
these plates to their almost closed position, so that the fluid movement is slow. This decreases the
speed of the recoil travel.
A flow of fluid through the first-stage orifices in the damping tube only occurs if the recoil-orifice and
the compression-orifice plates go into the compression chamber again. The gas in the second stage
chamber helps to make the extension effect of the shock absorber smooth.

Ad 3. Compression and recoil second stage gas chamber
The compression and the expansion of the gas in the second stage gas chamber help to make the
effect of the shock absorber smooth. This is transmitted through the floating piston to the oil and
then to the remaining parts of the shock absorber assembly. This procedure helps to make a smooth
landing.

B. Nose landing gear
The shock absorber in the nose landing gear (appendix XXVII) is a single chamber type and handles
the shock at the touchdown in two stages. The reason that the nose landing gear is not a four
chamber shock absorber is because it does not have to catch the huge forces of the main landing
gear.
The shock absorber is filled with hydraulic fluid and nitrogen through a single standard servicing valve
in the upper part of the leg.

1.7.2 Torque link damper
To prevent the inner cylinder from rotating in the outer cylinder of the shock strut, a torque
link is present. The torque link prevents the wheel from shimmies. The best example of a
shimmy is a shopping trolley where the wheels do not move straight but are changing
direction continuously. The torque link does not block the vertical movement of the shock
strut. The torque link consists of two identical smaller constructions which form together a v-
shape which are connected by a hinge in the middle of the v-shape (figure 1.11). These
smaller constructions are called the upper torque link (1) which is connected to the outer
cylinder and the lower torque link (2) which is connected to the inner cylinder. To the hinge a
torque link damper (3) is connected, which decreases the landing vibrations in both torque
links.



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1. Upper torque link
2. Lower torque link
3. Torque link damper



Figure 1.11 Torque link

The torque link damper (appendix XXVIII) is a spring centered, two way hydraulic unit, with its own
hydraulic reservoir. The hydraulic fluid of the damper is shown by the extension of the reservoir
when it is pressurized. When the contents of the reservoir is correct, the level indications FULL (1) or
REFILL (2) are in view. In the top of the reservoir of the torque link damper there is a bleed screw (3)
and a bleed plug (4) mounted. The torque link damper has a hydraulic reservoir (5) at the top and a
damper body (6) with a check valve (7) at the bottom. Inside the torque link damper a piston moves
from side to side. The hydraulic fluid inside the damper will corrected this to give enough pressure to
both sides to compensate the effect.

1.8 Regulations
The regulations for this project are found in Certification Specification for Large Aeroplanes (CS-25),
which is written by the European Aviation Safety Agency (EASA). In CS-25 regulations are set for
landing gear and landing conditions (1.8.1). Next to this, the safety of the aircraft has to be
guaranteed by the airworthiness of an aircraft (1.8.2).

1.8.1 Landing gear regulations
The retracting system is able to carry loads in flight and on the ground (A). Nose wheel steering must
be safe when it is used in extreme situations (B). The wheels (C) and tires (D) cannot cause hazardous
situations because of some prevention systems. Brakes are used to bring the aircraft in rest, what
also can be done in case of a failure (E). Landing and braking limits must be set by tests (F). These can
be found in the original regulations (appendix XXIX). The requirements are split up in variable and
fixed regulations (appendix XXX).

A. Retracting mechanism
The landing gear retracting mechanism must be able to carry the loads occurring when the gear is in
retracted position. In flight the landing gear is also designed for extra loads during special
manoeuvres. During landing and flight it has to be possible to keep the landing gear in extended
position. During flight it also has to be possible to keep the landing gear and doors in retracted
position, unless it can be shown that extended gear and doors are not hazardous at any speed. In
case of a failure in the hydraulic, electric or other energy supply, it must be possible to extend the
landing gear in a mechanic way.
3
2
1

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B. Nose-wheel steering
In case of a failure in the nose-wheel steering, it is necessary that the nose-wheel position cannot
cause a hazardous situation. It must be possible to control the steering system during landing and
take-off without exceptional skills, even in case of cross-wind.

C. Wheels
In the wheels a system must be presented that prevents wheel failure and tire burst by excessive
pressurization during landing. This system is called overpressure burst prevention. Also the wheels
have to be able to carry the loads during landing.

D. Tires
Each wheel must be fitted with a suitable tire. The tires must be able to carry the load during landing.
Each tire installed on a retractable landing gear system must have a clearance to surrounding
structure and systems that is enough to prevent contact between the tire and other systems. Also it
must be shown that tires do not produce a flammable gas (because of the gas combination in the
tires) and the tires are prevented from reaching unsafe heating levels.

E. Brakes
When a failure occurs in any electrical, pneumatic hydraulic or mechanical part, it has to be possible
to bring the aircraft in rest in a certain distance. A hydraulic fluid loss from the hydraulic braking
system may not cause a hazardous fire in flight or on the ground. When parking brake is used, the
pilot must be able to overrule this brake by using the manual braking system. An over temperature
burst prevention ensures that no tire burst or wheel failures can occur as a result of high elevated
temperatures.

F. Landing and braking limits
The maximum loads that occur during the landing must be tested with the maximum velocity
combined with the maximum weight of the aircraft, to ensure that it is safe in extreme conditions.
This also is necessary for braking situations as rejected take-off.

1.8.2 Airworthiness regulations
To maintain the airworthiness of an aircraft, it is necessary to ensure the safety of it. In the
handbooks of EASA different laws are set for maintaining airworthiness. A major aspect of the
airworthiness is the maintenance of the aircraft (A). In case of damage the safety of an aircraft is not
always doubt full according to the damage-tolerance of an aircraft (B).

A. Maintenance
Maintenance is a major aspect to keep the airworthiness of an aircraft. Therefore it is necessary that
an aircraft is maintained in accordance with a maintenance programme approved by a competent
authority, which shall be periodically reviewed and amended accordingly. The maintenance
programme shall contain details, including frequency, of all maintenance to be carried out, including
any specific tasks linked to specific operations. The maintenance of large aircraft shall be carried out
by an approved maintenance organisation. The maintenance shall be carried out on regular basis,
scheduled basis (e.g. after exceeding flight hours) or non-scheduled basis (e.g. in case of incidents).

B. Damage-tolerance
An evaluation of the strength, detail design and fabrication must show that failures due to damage
will be avoided throughout operational life of the aircraft. This means that every single part of the
structure which contribute to the landing gear or other parts have to be checked on failures and an
analysis must ensure that it is safe to operate the aircraft. There are four different analyses:

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1. Damage-tolerance (fail-safe) evaluation
2. Fatigue (safe-life) evaluation
3. Sonic fatigue strength
4. Damage-tolerance (discrete source) evaluation

Ad 4. Damage-tolerance (fail-safe) evaluation
This evaluation shows probable locations and modes of damage due to fatigue, corrosion or
accidental damage.

Ad 5. Fatigue (safe-life) evaluation
It has to be tested that all the structures in the aircraft are able to carry the repeated loads during its
service life without detectable damages. So the different materials in structures are tested by
analysis to ensure a life time quality.

Ad 6. Sonic fatigue strength
An analysis must show, supported by test evidence or by the service history that sonic fatigue cracks
are not probable in any part of the flight structure.

Ad 7. Damage-tolerance (discrete source) evaluation
When an aircraft is damaged occurred as a result of bird impact, it must be able to complete the
flight. The pilot can control the aircraft safely to its destination by limiting maneuvers, avoiding
turbulence and reducing the speed of the aircraft.

These evaluations must ensure the safety in case of damage.



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2. Landing gear construction
The landing gear construction is the mechanical construction of the landing gear. There are several
situations where the stress on the construction increases, the three situations when the stress
increase the most are analyzed in this chapter. For the stress in parking position the centre of gravity
needs to be known this point can be calculated (2.1). As soon as the aircraft is not in static position
anymore more forces will have its effect on the aircraft. During a rejected take off, the forces on the
nose landing gear will increase compared to the static position (2.2). When an aircraft touches the
ground an extra impulse due to the landing is created this impulse have its effect on the sheer stress
in the wheel axle (2.3). The landing gear construction consist of different materials each material has
an maximum value of stress that it can handle so the right choice of material is essential (2.4).
The source that is most used for this chapter is: Statica voor mechanici (2005).

2.1 Rest
In rest the aircraft is in parked position. The free body diagram is used to calculate the forces on the
aircraft in this situation (2.1.1). By using the calculated forces on the landing gear the shear stress in
the landing gear wheel axles is calculated (2.1.2).

2.1.1 Free body diagrams and assumptions
The forces on the nose landing gear and the main landing gear is calculated. (A). With this
information the forces in every strut of the main landing gear (B) and in the nose landing gear (C) are
calculated.

A. Free body diagrams
When the aircraft is in rest only the weight of the aircraft acts down on the landing gear (Figure 2.1).
Two normal forces counteract this weight. To determine the size of these normal forces, the
following calculations must be done:

1. Total mass of the aircraft
2. Centre of gravity
3. Normal forces on the main landing gear
4. Normal forces on the nose landing gear






Figure 2.1 Forces on the landing gear.

Ad 8. Total mass of the aircraft
F
MLG
F
NLG
12.64 m
10.88 m
Fz

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The total weight of the aircraft during rest is equal to its Maximum takeoff weight is 77000 kg. The
weight can be calculated by using the gravity acceleration (formula 2.1).

=

F
Z
: Gravitational force [N]
m : total mass of the airplane [kg]
g : gravity acceleration [m/s
2
] (9,81 on Earth)
Formula 2.1 Total mass formula


By filling in the mass, F
Z
is calculated:

= 77000 9,81 =

Ad 9. Centre of gravity
The position of the centre of gravity and the distance from the nose landing gear leg is calculated
using the Mean Aerodynamic Cord (MAC). The MAC is calculated using the following method
on a schematic top view of the A320:

1 Measure the root chord and tip chord.
2 Parallel to the tip cord the length of the wing root should be added in front and behind. The
same should be done for the tip but now adding the wing root cord.
3 Now draw two diagonal lines from one tip to the other. A cross is created.
4 At the lines intersect the wing cord should be measured. Using the scale (1:225) of the
drawing the MAC is calculated.

In the finished drawing the MAC can be measured by using the scale noted on the drawing (appendix
XXXI). The length of the MAC according to the drawing is 7,1 * 225/100 = 3,6 M. The distance of the
nose landing gear to the mean aerodynamic cord is 1,6 * 225/100 = 15,975 M.

Now because, stated in H1.1, the center of gravity must be placed between 17% and 37,5% on the
mean aerodynamic cord. Assuming the center of gravity is in its most forward position the center of
gravity position can be calculated (formula 2.2).



=

L
MAC
= Length of the MAC [m]
Cg = Center of gravity position on the MAC
D
M
= Distance to the beginning of the MAC [m]
Formula 2.2 Position of the center of gravity

Filling the previous mentioned elements, the distance is calculated:

= 3,6 0,17 15,975 = 16,587

However according to the airbus manual the center of gravity is not 16,587 meters from the nose of
the airplane but 15,95 M. This deviation in the answer is due the unclear definition of the MAC that
Airbus uses. The position found using the MAC calculation is an estimate. In all following calculations
the true position of the centre of gravity is used which is 10,88 M.
The distance from the nose landing gear to the main landing gear leg is 12,64 M (appendix V).


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Ad 10. Normal forces on the main landing gear
Now that the position of the centre of gravity is determined, the normal force on the aircrafts main
landing gear can be calculated. This is done by using a momentum equation of the momentum on
the main landing gear leg (formula 2.3).



=

= 0

D
Cg
= Distance to centre of gravity from the
nose landing gear leg
F
Z
= Total weight of the airplane
D
MLG
= Distance to main landing gear legs
F
MLG
= Normal force on main landing gear legs

Formula 2.3 Normal force on main landing gear

10,88 755370 +12,64

= 0

= ,

Ad 11. Normal force on nose landing gear
With the normal force on the main landing gear the total Y components of the free body diagram
should be 0. Resulting in the formula for the normal force on the nose landing gear (formula 2.4).

= 0

Fz = Total weight of the aircraft
Fmlg = Normal force in main landing gear legs
Fnlg = Normal force in the nose landing gear legs
Formula 2.4 Normal force nose landing gear

755370 +650191.8987 +

= 0

= ,

B. Free body diagram main landing gear
Using a free body diagram of the main landing gear (figure 2.2) the forces in every strut can be
calculated.
















30
Point D
F
c

F
MLG1
F
MLG2

F
Dy

F
Dx


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Figure 2.2 Free body diagram
main landing gear.
All forces must be determined in a certain order to be able to know all forces in the free body
diagram:

1 F
c

2 F
Dx

3 F
MLG1
/ F
MLG2

4 F
Dy


Ad 12. F
c

By placing the momentum in point D components F
Dx
and F
Dy
are eliminated. The Fn1 and Fn2
components cancel out each other because they create the same momentum in opposite direction.
Only the Fc force is now left
= 0
This means that F
c
= 0

Ad 13. F
Dx

= 0
sin30 = 0
Fx = 0

Ad 14. F
MLG1
and F
MLG2

The total normal force on the main landing gear was already calculated on the main landing gear in
2.1.1A this was 650191.8987 N. Because there are 2 main landing gears and 2 wheels on each main
landing gear this value is divided by 4.

650191.8987
4
= 162547.9747

Ad 15. F
Dy

= 0
162547.9747 +162547.9747 = 0
F
Dy
= 325095.9494 N

C. FBD nose landing gear
Because the nose landing gear forces are calculated the exact same way as the main landing gear.
Only the input values are different from the main landing gear. The F
n1
and F
n2
values of the nose
landing gear is
104845,356
2
= 52422.678 N
Now following the calculation of 2.1.1A the F
Dy
= 104845.356 N


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2.1.2 Shear stress
The forces on the landing gear cause a shear stress. The shear stress is calculated for the nose
landing gear (A) and the main landing gear (B).

A. Nose landing gear
The shear stress in the landing gear wheel axle can be calculated using a free body diagram (figure
2.3).


Figure 2.3 Free body diagram landing gear wheel axle

The shear stress is calculated with the following steps:

1. F
NLG

2. V
3. N
4. Momentum
5. Shear stress



Ad 16. F
NLG

The normal force on the wheel is calculated in 2.1.1.B:

=
104845,356
2
= 52422,078

Ad 17. V
The V is calculated with the sum of the forces in y direction (formula 2.5):

= 0

= 0

Fn = Normal force on the wheel
V = #dwarskracht#
Formula 2.5 Sum of the forces in y direction

= 0
= 0
52422,078 = 0
= 52422,078

Ad 18. N

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= 0
N = 0 N

Ad 19. Momentum
= 0
52422,078 0,231775 = 0
= 12150,1271

Ad 20. Shear stress
By using the V force and the surface of the pen (formula 2.6), the shear stress in the landing gear
wheel axle is calculated (formula 2.7):

=
2


A= Surface
r= Radius
Formula 2.6 Surface

The diameter of the pin in the wheel axle is 11,6 cm.
A= 0,058 = 0,0105683177 m
2


Shear stress =

V= #dwarskracht#
A= Surface
Formula 2.7 Calculation shear stress

=
52422,078
( 0,058
2
)
4960304.899 Pa = 4.960304899 Mpa
Shear stress = 4.960304899 Mpa

B. Main landing gear
The shear stress is calculated with the following steps:
1. Fn
2. V
3. N
4. Momentum
5. Shear stress

Ad 21. Fn
=
3255224,5535
2
= 162612,2768

Ad 22. V
= 0
= 0
162612,2768 = 0
= 162612,2768

Ad 23. N
= 0
N = 0 N


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Ad 24. Momentum
= 0
162612,2768 0,231775 = 0
= 37689,46046

Ad 25. Shear stress

=
162612,2768
( 0,058
2
)
= 15386770,31 = 15,38677031
Shear stress = 15,38677031 Mpa

2.2 Rejected take off
During RTO the A320 has a deceleration of 5 m/s
2
because the auto brakes are in MAX mode. This
deceleration does not only come from the brakes because several other forces are involved (2.2.1).
With the assumptions made for these forces, the friction force is calculated (2.2.2) which has
influence on the shear and pressure stress thickest part of the main landing gear leg (2.2.3).

2.2.1 Free Body Diagram and assumptions
During RTO several forces have influence on the A320 (figure 2.4).


Figure 2.4 Free Body Diagram of a A320 during RTO

The following assumptions had to be made, because specific data about these forces aren't available:

1. Body drag
2. Ground spoiler drag
3. Flap drag
4. Reverse thrust
5. Friction coefficient of the nose landing gear



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Ad 26. Body drag
The body drag (F
B
) and all the other drags can be calculated by using a drag coefficient (CD)(formula
2.8).

0,5
2



=

(0,5
2
)

C
D
: Drag coefficient
F
B
: body drag
: air density
V : Speed
A : Frontal surface

Formula 2.8 Body drag and drag coefficient

The best approach for the body of an A320 is the C
D
of a sphere which is 0,47. The frontal surface is
2,07
2
= 13,46
2
, the speed is the speed of V
1
which is 131 knots = 67,39 m/s. The density is
1,225 which is the ISA density at Mean Sea Level (MSL).
So the F
B
is:

= 0,47 (0,5 1,225 67,39

2
13,46) = ,

Ad 27. Ground spoiler drag
The ground spoiler drag (F
GS
) is calculated by using the same formula as for the body drag. The best
approach for the ground spoilers is the C
D
of a flat plate which is 1,28. The frontal surface of both
sets of ground spoilers is 15 0,5 2 = 15
2
. So the F
GS
is:

= 1,28 (0,5 1,225 67,39

2
15) = ,

Ad 28. Flap drag
The flap drag (F
FL
) is calculated by using a variant of the previous formula, the frontal surface is now
replaced with the wing surface (S) which is 244.84 m
2
for both wings. The best approach for the C
D
is
that of a wing with extended plain flaps which is 0,01 so a wing with fowler flaps has a C
D
of
approximately 0,03. So the F
FL
is:

= 0,03 (0,5 1,225 67,39

2
244,84) = ,

Ad 29. Reverse thrust
The reverse thrust force (F
REV
) is approximately 30% of take-off thrust. For the A320 with CFM56-5B4,
the maximum take-off thrust is 27000 lb which is 120143,01N per engine. So F
REV
is:

= 0,3 (120143,01 2) = ,

Ad 30. Friction coefficient of the nose landing gear
The friction of the nose landing gear (F
NLG
) and main landing gear (F
MLG
) is calculated by using a
friction coefficient (formula 2.9).

F
FRICTION
= Friction force
= Friction coefficient
F
NLG/MLG
= Normal force

Formula 2.9 Friction force

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The nose landing gear can be approached as a free rolling wheel, which has a of 0,03. The main
landing gear friction coefficient cannot be estimated and therefore needs to be calculated.

2.2.2 Motion equations
The friction of the nose and main landing gear are calculated by using the motion equations. To
calculate F
MLG
a moment equation on point N is done:

3,11 =

12,64 +

10,88 +

0,58 +

3,87 +

2,14 +

3,11

After filling in the assumptions made in the previous paragraph F
MLG
is:

= ,

Now F
MLG
is known, F
NLG
can be calculated:

= 0

= ,

So F
FRICTION-NLG
is:

= 0,03 172456,96 = ,

With F
FRICTION-NLG
, F
FRICTION-MLG
can be calculated:

= =

= ,

The F
FRICTION
per main landing gear leg is:
,

= ,

2.2.3 Conclusion

MLG
can now be calculated using the friction formula:

=
,
,
= ,

By using the method described in 2.1, the shear stress and the pressure stress in the thickest part of
the main landing gear leg during RTO can be calculated. The stresses in the axle are not calculated as
the friction force has no influence on the stress in the axle when calculating from the same direction
as in 2.1 (figure 2.2):

The shear stress during RTO is (figure 2.5):

=

=
291459,52
( 0,085
2
)
12840753,86 Pa = ,


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And the pressure stress during RTO is:

=

=
108252,45
( 0,085
2
)
4769249,14 Pa = ,




















2.3 Landing
When the forces during landing gear are calculated, some values are unknown and cannot be found.
Therefore assumptions have to be made to define the forces on the landing gear (2.3.1). With these
assumptions the force that appears during touchdown is calculated with help of Newtons second
law (2.3.2). When the aircraft is landing the new impulse brings a shear force that is needed for
constructors (2.3.3).

2.3.1 Free body diagrams and assumptions
When an aircraft is approaching the forces of the aircraft are equilibrium (A). When the touchdown is
made, the approach forces are still the same, but a new impulse occurs (B).

A. Approach assumptions
The A320 endures several forces during the flight and in approach (figure 2.6). The basic forces of the
aircraft are F
Lift
, F
Drag
, Weight and F
Thrust
. The maximum landing weight (MLW) of the A320 is 66000 kg.
The assumption for the touchdown is that the aircraft is at a constant speed, so there is no
acceleration. The forces on the aircraft in descent with no acceleration are all in equilibrium. The
situation of the touchdown is that the aircraft touches the ground with the maximum allowable
descent velocity that is stated by CS-25 of 3.05 m/s. The situation is also stated in the most desirable
conditions and the standard ISA conditions, there is no crosswind or shears through the touchdown.


Figure 2.5 Stresses in the main landing gear leg

M
V
N

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Figure 2.6 Forces in approach A320

B. Touchdown assumptions
Movement of the aircraft between first touch with the ground and the complete compress of the
shock absorber is the touchdown. At the moment the aircraft touches the ground (figure 2.7)
vertical velocity is gone and this impact creates energy, the impulse. The vertical velocity needs to be
transformed into an opposite force, the normal force. The other forces are still all in equilibrium.

For the calculation it is assumed that the shock absorbers will be compressed with a constant speed
till 0.5 m during the touchdown. The shock strut velocity is 3.05 m/s in the beginning and at the end
it is 0 m/s. Because it is not possible to find the real deceleration of the shock strut, this is calculated
with a constant average velocity: 1.525 m/s. The landing is performed at a paved runway with a
friction coefficient of 0.03.



Figure 2.7 Free motion body diaphragm Touchdown

F
Lift


Weight
F
Thrust

F
Drag

F
Lift


Weight
F
drag
F
Thrust

Lift
Thrust
Weight
Drag
F
normal

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2.3.2 Landing force calculation
The second law of Newton can be written so that it is helpful to calculate the landing forces (formula
2.10).



=
1
0
1
0


1 = time (second measuring) [s]
0 = time (first measuring) [s]
= sum of forces [N]
= delta time [s]
1 = velocity (second measuring) [m/s]
0 = velocity (first measuring) [m/s]
= mass [kg]
= delta velocity [m/s]
Formula 2.10 Impulse formula

The touchdown time can be calculated (formula 2.11).


= /
= time [s]
= distance [m]
= average velocity [m/s]
Formula 2.11 Time formula

With the time of the touchdown (0.328 s) the sum of the forces (F
normal
) can be calculated:

(0.328 0) = 66000 0 66000 3.05

= = 613719.5


The F
friction
is calculated with a friction force formula (formula 2.12).


=
= friction force [N]
= friction coefficient [no dimension]
= normal force [N]
Formula 2.12

= 0.03 613719.5 = 18411

2.3.3 Landing shear force

The force used for the shear stress is the F
normal
plus the weight of the aircraft. The force per main
gear wheel is:

=
613719.5 + 647460
4
= 315294.88

Now the shear stress of the axis during landing is:

=

=
315294.88
( 0,058
2
)
= 29833970.68 Pa = .


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2.4 Materials
The landing gear is made of different materials. Which type of material that is used, depends on
several properties of the material (A) which must be able to withstand the shear and pressure
stresses(B).

A. Types of materials
Which material Airbus uses depends on the application of the material. In the designing phase Airbus
determine the forces on the landing gear in unique situations. With these forces Airbus can consider
which material has the best properties.

The A320 landing gear consists of:

1. Steel
2. Aluminium
3. Titanium
4. Magnesium
5. Composites

Ad 31. Steel
Steel is an alloy of iron and carbon. The more carbon, the stronger the steel will be, however it will
be more brittle too. Because steel is relative cheap and strong, it is the most used material in the
landing gear for the main fitting.
The landing gear leg and axle are made out of a variant of this material, namely stainless steel. These
parts have to endure huge forces and stresses during landing and RTO, as seen in previous
paragraphs. Stainless steel combines the advantageous properties of steel like high strength en low
ductility with stainless properties.

Ad 32. Aluminium
Pure aluminium is not very strong but it has the advantage of low weight and natural protection
against corrosion. To give aluminium strength, it can be alloyed with copper or zinc. This material is
not widely used at the landing gear.

Ad 33. Titanium
Titanium is only used as an alloy which consists of aluminium, cobalt, copper, iron and nickel. It is
used to make materials more stiffness.

Ad 34. Magnesium
Magnesium is used for the wheels. Nowadays magnesium is generally regarded as unacceptable, due
to the high risk of fire and the susceptibility of corrosion.

Ad 35. Composites
Composites are materials of two or more constituent materials with different physical or mechanical
properties. It is a relative light material, and it is used for pistons, side braces and torque arms.

B. Deviation
The size of the shear and pressure stresses in the previous paragraphs where different then the real
shear stress in the parts. Airbus has, during the designing phase, also used a safety factor. This safety
factor is ,in aviation, 1.5 times the shear stress. Another reason could be that the Airbus aircraft are
designed with hollow axles. The surface of the axles will decrease in that case, which resulted in a
higher shear stress. The exact shape en construction of the axle is impossible to find in the standard

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A320 manuals. The last conceivable reason is that the dimension of the axle in reality is less than
expected. Because of this the surface also decreases.


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3. Landing gear operations
During the operational time of the A320s landing gear malfunctions will occur in the landing gear. To
prevent that all malfunctions ground the aircraft and create high costs, the Master minimum
equipment list and Minimum equipment lists have been created (3.1). Two of these malfunctions,
the 90 turned nose wheel and leaking shock absorber, have been analyzed for cause and damage to
the landing gear (3.2). The 90 turned nose wheel and the leaking shock absorber cause the airplane
to go into maintenance and create large costs. However by doing proper maintenance by following
procedures the safety of the airplane increases and costs will decrease because of less malfunctions
(3.3). Following from this information a conclusion has been made on how the maintenance effects
costs and its positive effect on costs in the long term (3.4).
The most used sources for this chapter are Aircraft maintenance manual (2004) and the Master
minimum equipment list (2009).

3.1 MMEL & MEL
The Master minimum equipment list (MMEL) and the minimum equipment list (MEL) inform the
operator or mechanic about the airworthiness of the aircraft when a failure occurred. The
documents are used as handbook to decide if an aircraft can dispatch or need to stay on the ground
and undergo maintenance. Each failure has a different impact on the system. For some failures it is
easy to recognize if the aircraft is still able to dispatch, for example a leak in the green hydraulic
system it is more difficult to determine.

During the development process of the MMEL different technicians work together, in order to create
the MMEL in such a way that the acceptable level of safety is guaranteed while there is no loss in
profitability of the aircraft. Every failure that can occur has impact on several systems instead on only
the failed system, multiple failures need to be avoided at all cost since these result in an
unacceptable level of safety.

The difference between a MMEL and the MEL, is that the MMEL is made by the manufacturer the
moment the aircraft is designed. The MEL is made by the operator which includes his own demands,
that are as strict or stricter then the MMEL. The MEL is also specified to the layout of operators
aircraft.

The MMEL has a standard lay-out (figure 3.1). In the column item (1), the items are listed for which a
failure can occur. The rectification interval (2) can be classified as A,B,C or D. When an item is
classified as A there is no maximum timeframe when the rectification should take place, for B there is
a maximum of 3 days, C 10 days and 120 days for D. The column for number installed (3) refers to the
amount of installed items there are in total on the aircraft. Number required for dispatch (4) refers to
the amount of items that need to be operative in order to have a safe dispatch. The column remarks
or exceptions (5) gives an overview of all exceptions there can be for a certain item that must be met
before a dispatch can take place.


1. The item
2. Rectification
interval
3. Number installed
4. Number required
for dispatch
5. Remarks or
exceptions

1
5
4
3
2

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Figure 3.1 example of a MMEL

3.2 A320 failures
There can occur several failures on the A320 which have an effect on the airworthiness of the A320.
After the take-off a JetBlue A320 occurred of a nose wheel retracting failure, this problem was
created by a wrong installed shock absorber (3.2.1). Another failure that can occur on the A320 is a
hydraulic leak on the shock absorber, which can arise through poor maintenance (3.2.2).

3.2.1 Nose landing gear failure
On the 21 of September 2005 a JetBlue A320, on a flight from California to New York, had a problem
with its nose wheel. After take-off, when the pilot tried to retract the landing gear, two warning
lights light up. These lights indicated that there are problems with the shock absorber and steering
on the nose gear. The problem was that the nose landing gear turned 90. During the emergency
landing when the nose gear touched the ground, the rubber tires shredded away until the metal
wheel scraped the runway pavement. This friction caused a trail of white smoke and finally sparks
and flames.

The failure could be a result of a malfunction in one of the control systems. Malfunctions in other
mechanical components can also result in this failure. In this case, the real cause was a wrongly
installed hydraulic shock absorber. During the installation of the shock absorber the upper part of the
inner cylinder, which contains anti-rotation lugs, was wrong installed. Anti-rotations lugs are installed
on the top of the shock absorber inner cylinder to protect the inner cylinder against rotating and
twisting. The anti-rotation lugs must correctly seat in the slots. Because of the wrong installation the
anti-rotation lugs were rotated 20-30, so they are not properly fitted in the back plate slots.

The A320 cannot be airworthy with a wrong installed hydraulic shock absorber. The nose wheel
cannot be retracted and because of the 90 nose wheel turn the landing cannot be performed
smoothly. There will be severe damage to the nose landing gear and the aircraft.

3.2.2 Shock absorber failure
The shock absorber could have a hydraulic leakage which results in less effective or no absorption
during landing. This failure could occur because of poor maintenance.

There could be a leakage of hydraulic oil in several points (figure 3.2). Leaks between the main fitting
(1) and the gland housing (2) or between the gland housing and the sliding tube (3) are possible. One
of the leakage causes is poor maintenance of the chrome sliding tube. In this case the chrome is
damaged by dust that is not removed during maintenance. When the chrome is damaged, the sliding
tube is not sealed any more. It is also possible that the leak is caused by damaged grand seals (4)
which separates the hydraulic fluid from the sliding tube through the main fitting.

The aircraft is not airworthy with a hydraulic leak in the shock absorber. The shock absorber cannot
bear the forces during touchdown, so the safety is not guaranteed.


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4



3

1

2
1. Main Fitting
2. Gland housing
3. Sliding tube
4. Gland seals






Figure 3.2 Leakage points of a shock absorber

3.3 Maintenance and costs
To solve previous failures, maintenance is needed. But the failures could have been prevented if the
previous maintenance checks where done correctly (3.3.1). This non-scheduled maintenance
increases the already existing maintenance costs (3.3.2).

3.3.1 Maintenance
To ensure and maintain the safety of the passengers and personnel, maintenance of the aircraft is
required. Maintenance is divided in three different categories (A). When a failure occurs, the last of
the three categories is needed which is non-scheduled maintenance. During this maintenance, failure
specific procedures are used (B).

A. Maintenance categories
There are three maintenance categories:

1. Inspections
2. Scheduled maintenance
3. Non-scheduled maintenance

Ad 36. Inspections
Inspections are visual and carried out at daily or weekly basis. The purpose of these inspections is to
discover visual damage to the aircraft. These inspections are carried out by a qualified mechanic,
which has to sign for approval. The walk around of the pilot before each flight is also a form of
inspection.


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Ad 37. Scheduled maintenance
Scheduled maintenance is carried out at a fixed time or on-condition basis. A component is replaced
or overhauled after a number of flight hours or a cycle at a fixed time. There are three types of
scheduled maintenance, the A/B-check, the C-check and the D-check.
The A check is done after 550 flight hours (approximately three months) or 330 flight cycles. During
the A/B check the oil filters are replaced which takes at least 16 hours. The C-check is more complex
and is done after 4000 flight hours (approximately eighteen months) or after 3000 cycles. During this
check all crucial elements of the landing gear and the aircraft are inspected and if necessary
replaced.
The D-check is very time consuming, it takes about four weeks. During this period the entire aircraft
is separated and each part is inspected. This check is done after 24.000 flight hours (approximately
eight years) or after 18.000 cycles.

Ad 38. Non-scheduled maintenance
When a failure occurred during flight, or when a failure is found during inspection non-scheduled
maintenance is required. Sometimes replacement of components can be deferred, if allowed by the
minimum equipment list (MEL).

B. Failure specific procedures
Each failure requires a different maintenance procedure, the two previous described failures are no
exception:

1. JetBlue accident
2. Hydraulic leakage in a shock absorber

Ad 39. JetBlue accident
When the nose wheel is 90 rotated during a landing, a non-scheduled maintenance is needed. The
maintenance will be done in a hangar. The nose landing gear must be replaced, because the damage
to the nose landing gear is irreparable and the airworthiness could not be guaranteed with just a
repair. Before the landing gear can be removed the front side of the aircraft has to be propping up.
Then all hydraulic liquid and gasses has to be removed. The landing gear may finally be dismantled.
After the dismantling, the aircraft can be checked on damages at the connection between the
landing gear and the aircraft. If there is no significant damage, the new nose gear can be placed.
Then the hydraulic liquid can be refilled in the systems. Before the aircraft returns in to operation,
the nose landing must be certified.

The incident with the JetBlue A320 was not the only incident with the nose wheel at 90 has caused
that the manufacturer changed the AMM. The changes that are made comes with the removal and
disassembly of the shock absorber. For the removal of the shock absorber the cylinder of the shock
absorber will be removed and there must be made a mark to show the position of the different parts.

For the assembly of the shock absorber the cylinder must be installed in a way that the axis XX
parallel is placed to the axis YY and the axis ZZ perpendicular to the axis YY (appendix XXXII).
This will ensure that the shock absorber is placed right. When the shock absorber will finally be
installed, the piston lugs must be placed in line with slots from the shock strut back plate. The shock
absorber will now be in the correct position, so the wheel axle will be in line with the centreline of
the A320. The final check of the installation is that the shock absorber cannot be turned around
because of the right installation of the anti-rotation lugs.

Ad 40. Hydraulic leakage in a shock absorber
The maintenance of the shock absorber is done by a scheduled maintenance and inspection. Next to
these inspections the scheduled maintenance can be done once a month by dismantle the shock

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absorber and clean up the parts of it. To contain airworthiness the other seal can be activated during
a after flight inspection as a short term solution and the whole shock absorber is replaced during a
non-scheduled maintenance task as a long term solution.

Good flight inspection is enough to check if the shock absorber is free of dust. To avoid a shock
absorber hydraulic leak, it is necessary that the chrome of the sliding tube is well maintained.

3.3.2 Costs
The costs plays a important role for the ALA, these costs can divided into direct costs and indirect
costs (A). The repair and replacement of the new nose wheel brings extra costs (B). The hydraulic
leak in the shock absorber mainly effects the costs in the long term (C)

A. General costs
Finances play a crucial role in the aviation industry. Airliners are resolute to aim for the most efficient
way to perform operations in order to obtain maximum profit from the flights. Costs of the airliners
in relation to the landing gear can be split in to two different types of costs:

1. Direct costs
2. Indirect costs

Ad 1. Direct costs
Direct costs are defined as costs in relation with the purchase of the landing gear and the following
maintenance procedures. In addition, aircraft on ground costs, labour costs (salaries of certified
personnel to execute maintenance procedures), costs for equipment, shipping of equipment and
certification are a part of direct costs.

Ad 2. Indirect costs
Indirect costs are defined as costs, which are not directly related to the operation of the landing gear
itself. Training costs for maintenance personnel, rent for a hangar are part of indirect costs for the
landing gear.

B. Costs 90 rotated nose wheel
The damage of the nose wheel landing gear has to be repaired directly. Because the maintenance is
non-scheduled the aircraft creates aircraft on ground costs. For the maintenance of the nose wheel
landing gear a hangar has to be rented. In the hangar the equipment for the maintenance has to
present. This equipment is at the location or may be flown with the maintenance personnel of the
airliner. Before the aircraft gets in operation the aircraft has to certificate. This certification causes
extra costs (table 3.1).

Price Number Total
Aircraft on ground 7000,- 12 84.000,-
Labour costs 375,- 12 4500,-
Equipment costs 50.000,- 1 50.000,-
Test flight per hour 2500,- 3 7500,-
Hangar rent per hour 150,- 15 2250,- +
Total 148.250,-
Table 3.1

Aircraft on ground costs are based on a turnover of 5000,00 per hour. Added to this the costs for
personnel and aircraft parking will bring it to a total of 7000,00 per hour.

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Domein Techniek
Aviation Studies
The labour costs based on a wage of 75,00 per hour for each employee. Assumed is that five
employees carry out the maintenance.

C. Costs hydraulic leakage
For the short term solution no other costs than a normal inspection are needed. After this solution a
new certification is not needed. For the long term solution extra costs have to be taken into account.
Direct costs are the purchase of a new shock absorber, the rent of a hangar and the maintenance
personnel which complete the replacing of the shock absorber. Indirect costs are fewer incomes,
because the aircraft cannot be used for flights (aircraft on ground). After the replacing of a shock
absorber, the certification causes extra costs (appendix XXXIII).

3.4 Conclusion
After the three analyses that have been done regarding the construction, the different forces on the
construction and the failures that can occur on the landing gear. The team can conclude the
following points:

Landing gear is one of the most essential parts of the aircraft, it has influence on different
systems as well as the design of the aircraft.
The materials are able to sustain the stress that is on the different components in the
different phases of the flight.
Maintenance is required to prevent malfunctions, when this maintenance is not done in a
sophisticated way as for example in the case of the JetBlue where the engineers did not
rotate the shock absorber around its vertical axis before tightening the shock absorbers
upper bolt, this malfunction would have been avoided.
The leakage of the shock absorber is another failure that can be avoided by correct
maintenance, when the chrome sliding tube is kept clean so that dirt damage is avoided
there I s a slighter chance this failure will occur.
A correct maintenance program does not only have a positive influence, the downside is that
the aircraft can be less in the air. Beside this, a good maintenance program does not
guarantee that no failure will occur anymore.

Based on these points the advice of the team is to increase the frequency of the maintenance by
300% compared to the old maintenance program.


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Domein Techniek
Aviation Studies
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Domein Techniek
Aviation Studies
Flight physics
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