Table of Contents
Hydraulics system................................................................................................................................................ 3
Advantages: ..................................................................................................................................................... 3
Primary flight controls:.................................................................................................................................... 3
Secondary flight controls: ............................................................................................................................... 3
Utility systems: ................................................................................................................................................ 3
User requirement parameters: ....................................................................................................................... 4
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Pressure ............................................................................................................................................... 4
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Integrity ............................................................................................................................................... 4
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Flow rate.............................................................................................................................................. 4
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Duty cycle ............................................................................................................................................ 4
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Emergency or reversionary use ........................................................................................................... 4
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Heat load and dissipation .................................................................................................................... 4
Components of Hydraulics system ...................................................................................................................... 4
Major functions ........................................................................................................................................... 5
Redundancy ................................................................................................................................................. 5
Hydraulic Fluids ............................................................................................................................................... 6
Fluid Pressure .................................................................................................................................................. 6
Fluid Temperature........................................................................................................................................... 6
Fluid Flow rate ................................................................................................................................................. 7
Hydraulic Piping............................................................................................................................................... 7
Hydraulic Pumps.............................................................................................................................................. 7
Working principle of a piston pump................................................................................................................ 8
Fluid conditioning ............................................................................................................................................ 9
Cooling......................................................................................................................................................... 9
Cleaning ....................................................................................................................................................... 9
Hydraulic reservoir ........................................................................................................................................ 10
Warnings and status...................................................................................................................................... 10
Emergency power sources ............................................................................................................................ 10
Proof of design .............................................................................................................................................. 11
Aircraft system applications .......................................................................................................................... 12
The Avro RJ Hydraulic System ................................................................................................................... 12
The BAE SYSTEMS Hawk 200 Hydraulic System ........................................................................................ 16
Civil Transport Comparison ....................................................................................................................... 18
Landing Gear Systems ................................................................................................................................... 21
Nose Gear .................................................................................................................................................. 21
Main Gear .................................................................................................................................................. 22
Braking Anti-Skid and Steering .................................................................................................................. 23
Electronic Control ...................................................................................................................................... 25
Automatic Braking ..................................................................................................................................... 26
Multi-Wheel Systems ................................................................................................................................ 26
Brake Parachute ........................................................................................................................................ 29
Steering ..................................................................................................................................................... 29
Figure 1 Hydraulic system loads .......................................................................................................................... 3
Figure 2 A simple hydraulic system ..................................................................................................................... 4
Figure 3 A typical dual channel hydraulic system .............................................................................................. 5
Figure 4 Characteristic curve for a 'constant pressure' pump ............................................................................ 8
Figure 5 working principle of a piston pump ...................................................................................................... 8
Figure 6 external view of a piston pump............................................................................................................. 9
Figure 7 Cross section view ................................................................................................................................. 9
Figure 8 Filter units.............................................................................................................................................. 9
Figure 9 Hydraulic test rig ................................................................................................................................. 11
Figure 10 Hydraulic system testing process ...................................................................................................... 12
Figure 11 RJ146 regional jet hydraulic system .................................................................................................. 12
Figure 12 Yellow system components............................................................................................................... 13
Figure 13 The BAE Systems Hawk 200 hydraulic system .................................................................................. 16
Figure 14 The Panavia Tornado hydraulic system............................................................................................. 17
Figure 15 Tornado hydraulic system ................................................................................................................. 18
Figure 16 Simplified A320 Hydraulic system ..................................................................................................... 19
Figure 17 a Boeing 767 Hydraulic system ......................................................................................................... 20
Figure 18 The Raytheon 1000 nose landing-gear.............................................................................................. 21
Figure 19 Landing gear sequence ...................................................................................................................... 22
Figure 20 The Raytheon Main landing gear ...................................................................................................... 22
Figure 21 Brake control systems - Functional elements ................................................................................... 23
Figure 22 The Dunlop Maxaret anti-skid system .............................................................................................. 24
Figure 23 Electronic anti-skid system with adaptive pressure control ............................................................. 25
Figure 24 An Automatic Brake control system.................................................................................................. 26
Figure 25 Simplified B 777 braking configuration ............................................................................................. 27
Figure 26 Airbus A-380 Landing gear configuration .......................................................................................... 27
Figure 28 A380 brake control system ............................................................................................................... 28
Figure 29 F117 deploying brake parachute....................................................................................................... 29
Figure 27 A380 steering control system ........................................................................................................... 29
Hydraulics system
The principal requirements are
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3.
4.
low weight,
low volume,
low initial cost,
High reliability and low maintenance.
Advantages:
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Self-lubricating system
low weight per unit power
Advancement in micro-processors: ‘smart’ pumps and valves
Primary flight controls:
-Elevators – (1)
-All-moving tail surfaces (military)
-Rudders – (2)
-Ailerons – (3)
-Flaperons – (4)
-Canards
Secondary flight
controls:
-Flaps – (5)
-Slats – (7)
-Spoilers – (8)
-Airbrakes – (9)
-Stabilizer trim – (10)
Utility systems:
-Landing gear
-Brakes
-Gear steering
-Aerial refueling probes (military)
-Cargo doors
-Loading ramp (military)
-Passenger stairs
Figure 1 Hydraulic system loads
User requirement parameters:
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Pressure – primary pressure of the system? Determined by the appropriate standards and the
technology of the system
Integrity – Is the system flight safety critical or can its loss or degradation be tolerated? This
determines the number of independent sources of hydraulic power that must be provided, and
determines the need for a reversionary source of power
Flow rate – What is the rate of the demand, in angular or linear motion per second, or in liters per
second in order to achieve the desired action?
Duty cycle – What is the ratio of demand for energy compared to quiescent conditions. This will be
high for continuously variable demands such as primary flight control actuation on an unstable
aircraft (throughout the flight), whereas it will be low for use as a source of energy for undercarriage
lowering and retraction (twice per flight)
Emergency or reversionary use – Are there any elements of the system that are intended to
provide a source of power under emergency conditions for other power generation systems? An
example of this is a hydraulic powered electrical generator. Is there a need for a source of power in
the event of main engine loss to provide hydraulic power which will demand the use of reversionary
devices?
Heat load and dissipation – The amount of energy or heat load that the components of the system
contribute to hydraulic fluid temperature
Components of Hydraulics system
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A source of energy – engine, auxiliary power unit or ram air turbine
A reservoir
A filter to maintain clean hydraulic fluid
A multiple redundant distribution system – pipes, valves, shut-off cocks
Pressure and temperature sensors
A mechanism for hydraulic oil cooling
A means of exercising demand – actuators, motors, pumps
A means of storing energy such as an accumulator
Figure 2 A simple hydraulic system
Major functions
The primary source of power on an aircraft is the engine, and the hydraulic pump is connected to the engine
gearbox. The pump causes a flow of fluid at a certain pressure, through stainless steel pipes to various
actuating devices. A reservoir ensures that sufficient fluid is available under all conditions of demand.
Redundancy
This simple system is unlikely to satisfy the condition stated above, and in practice most aircraft contain
multiple pumps and connections of pipes to ensure that single failures and leaks do not deplete the whole
system of power.
To achieve the levels of safety described above requires at least two hydraulic circuits as shown in Figure 3.
The degree of redundancy necessary is very largely controlled by specifications and mandatory regulations
issued by the national and international bodies charged with air safety. The requirements differ considerably
between military and civil aircraft. Military aircraft frequently have two independent circuits, large civil
transports and passenger aircraft invariably have three or more.
In both types additional auxiliary power units and means of transferring power from one system to another
are usually provided.
Figure 3 A typical dual channel hydraulic system
Hydraulic Fluids
Mineral based fluid known variously as:
• DTD 585 in the UK • MIL-H-5606 in the USA • AIR 320 in France • H 515 NATO
Advantages.
Freely available throughout the world, reasonably priced
Low rate of change of viscosity with respect to temperature compared to other fluids.
Unfortunately, being a petroleum based fluid, it is flammable and is limited to a working temperature of
about 130 C. One of the rare departures from DTD 585 was made to overcome this upper temperature limit.
This led to the use of DP 47, known also as Silcodyne, in the ill-fated TSR2.
Reduced flammability
MIL-H-83282, an entirely synthetic fluid, now adopted for all US Navy aircraft. It is miscible with DTD 5858
and, although slightly more viscous below 20 C, it compares well enough.
In real terms the designer of military aircraft hydraulic systems has little or no choice of fluid since defense
ministries of the purchasing nations will specify the fluid to be used for their particular project. Most
specifications now ask for systems to be compatible with both DTD 585 and MIL-H-5606.
Commercial aircraft make use of phosphate ester fluids which are fire resistant, e.g.
• Solutia Skydrol LD-4, Skydrol 500B-4 or Skydrol 5
• Exxon Type IV HJ4AP or Type V HJ5MP
These fluids are not fireproof – there are certain combinations of fluid spray and hot surfaces which will
allow them to ignite and burn.
Fluid Pressure
Similarly little choice is available with respect to working pressure. Systems have become standardized at
3000 psi or 4000 psi.
Interestingly DTD 585 cannot be used above 5000 psi because of shear breakdown within the fluid.
A detailed study would show that the optimum pressure will differ for every aircraft design.
Fluid Temperature
With fast jet aircraft capable of sustained operation above Mach 1, there are advantages in operating the
system at high temperatures, but this is limited by the fluid used.
For many years the use of DTD 585 has limited temperatures to about 130 C, and components and seals
have been qualified accordingly.
The use of MIL-H-83282 has raised this limit to 200 C and many other fluids have been used from time to
time, for example on Concorde and TSR2, to allow high temperature systems to be used.
A disadvantage to operating at high temperatures is that phosphate ester based fluids can degrade as a
result of hydrolysis and oxidation. As temperature increases, so the viscosity of the fluid falls. At some
point lubricity will be reduced to the extent that connected actuators and motors may be damaged.
Fluid Flow rate
Nominal system pressure = Stall pressure
I.e. no flow will be present in the circuit apart from some very low quiescent leakage.
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Incorporating realistic pressure drop. 20-25 % of the nominal pressure is considered.
Pressure drops across each actuator will be known.
Requirement for simultaneous action and movement speed will be known.
Sum of these will give the maximum flow rate demand of system.
Flow demands at various phases of the flight – take-off, cruise etc. – Graphical representation.
For sizing of pumps, flow required on approach provides the design case. I.e. at low engine rpm, pump rpm.
Absolute max. Flow demand is of very short duration, involving very high velocities of very small volumes of
oil. So, an accumulator can be used to augment the flow available.
Accumulator contains a compressed gas cylinder, which provide energy to augment system pressure.
When the flow demanded will exceed the pump capabilities the system pressure is controlled by the
accumulator, not the pump? However frequency of max demand must be known, and time must be
available for the pump to recharge the accumulator.
Hydraulic Piping
After system architecture is defined for all aircraft systems using hydraulic power, then it’s possible to
design the pipe layout in the aircraft.
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Separate piping to avoid common mode failures as result of accidental/battle damage.
Once layout is obtained, pipes length and flow rate calculation in each section/branch is possible.
25 % Allowable pressure drop will be further divided between pressure pipes, return pipes and
components and pressure drop values are chosen across each component.
Alternatively it can be achieved without excessive penalties, being incurred by over-large porting or
body sizes.
Once pipe lengths, flow rates and permissible pressure drops are known, pipe diameters can be
calculated using the normal expression governing friction flow in pipes.
Hydraulic Pumps
A system will contain one or more pumps depending on: type of aircraft, need for safety and redundancy.
Mounted on an engine-driven gearbox.
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For civil, Mounted on an accessory gearbox attached on engine casing.
For military, mounted on an Aircraft Mounted Accessory Drive (AMAD) attached on the airframe.
Pump speed is directly related to engine speed and must be capable of working over a wide speed range.
Typical max continuous speed for modern military aircraft is 6000 rpm, but this is largely influenced by
pump size, the smallest pumps running fastest.
Universally used pump type is variable delivery, constant pressure.
Demand on the pumps continuous throughout a flight, but frequently varying.
Within pump flow range, these pumps can maintain pressure within 5 % of nominal except during short
transitional stages from low flow to high flow.
Figure 4 Characteristic curve for a 'constant pressure' pump
Working principle of a piston pump
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Pumps are designed to sense outlet pressure and feedback this signal to a plate carrying the
reciprocating pistons.
Pistons are free to move at an angle to the longitudinal axis of the rotating drive shaft.
When angle is 90 ° to the linear axis (drive shaft), no linear displacement of the pistons.
Normally 9 – pistons arranged diametrically around the plate.
Position of the plate varies the amount of reciprocating movement of each piston.
Up to its maximum limit the plate will move to displace the volume needed to main nominal system
pressure.
When flow demands beyond maximum displacement are made the system pressure drops rapidly to
zero.
For short periods pressure can be maintained by means of an accumulator.
Figure 5 working principle of a piston pump
Figure 7 Cross section view
Figure 6 external view of a piston pump
Fluid conditioning
Under normal working conditions hydraulic fluids needs
Cooling
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Bleed off any air accumulating in the reservoir (de-aerate by connection of ground equipment)
For cooling purposes the fuel/hydraulic heat exchanger is used. (on ground)
Air/fluid cooling may be provided once the aircraft is in flight.
Heat exchangers are low pressure devices; they are situated in the return line of actuator before
entry to reservoir.
Adequate strength must be present to prevent external burst (avoiding any failure) by determining
max pressure experienced across fuel/hydraulic matrix.
Heat due to friction carried away by pump case drain line, when running offload.
Cleaning
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Servo-valve introduction (very fine clearances) requires the need for very clean fluids.
Leading to filter elements made of resin bonded paper supported by arrangements of metal tubes
and wire mesh.
This produces filter elements of high strength capable of withstanding differential pressures of one
and half times the system pressure.
Capable of stopping all particles of contaminant above 5 – microns size and high % of particles
below this size.
‘Beta’ rating; ratio of particles upstream and downstream of the filter.
Typically size ranges: 5-15 microns, 15-25 microns, 25-50 microns & 50-100 microns.
Mostly adequate level of cleanliness can be achieved by using a 5 – micron absolute return line filter
in combination with a 15 – micron pressure line filter, gives acceptable element life.
Filters are not used in pumps inlet line.
Electronic automatic counters can be used to measure the cleanliness level achieved within 5 %
repeatability.
Figure 8 Filter units
Hydraulic reservoir
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For military aircraft the reservoir must be fully aerobatic.
o Fluid must be fully contained, with no air/fluid interfaces
o Fluid supply must be maintained in all aircraft attitudes and g conditions.
Reservoir must be sufficient to accelerate a full charge of fluid into each cylinder, in order to
achieve a good volumetric efficiency from the pump.
Volume of reservoir includes all differential volumes, allowance for thermal expansion a generous
emergency margin.
When the reservoir level falls below a predetermined point, isolation of certain parts of the system
required. I.e. to isolate leaks within the system and to provide further protection for flight safety
critical subsystems.
The cut-off point must ensure sufficient volume for the remaining system under all conditions.
Warnings and status
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3.
4.
5.
Pressure Transducers: Monitor system pressure and display it in the cockpit.
Pressure Switches: Warn of low pressure on the central panel.
Filter Blockage Indicators: Show filter condition to ground personnel.
Fluid Temperature Warning: Alerts aircrew about high fluid temperature.
Microprocessor-Based Units: Provide detailed health data to ground crews.
Emergency power sources
Hydraulic systems in aircraft have emergency power sources, often in the form of accumulators. These
accumulators serve various critical functions:
1. Wheel-Brake Systems: Standby accumulators provide power for brake applications when other
power sources fail. This ensures braking capability even in emergencies.
2. Cockpit Canopies: Accumulators allow emergency hydraulic opening and closing of cockpit canopies.
3. Flight Control Actuation: In total system failure, accumulators can move flight control surfaces
enough to stabilize the aircraft, allowing safe crew ejection.
4. Electric Motor-Driven Pumps: For longer emergency power, electric motor-driven pumps may be
used. These are limited by battery size and weight, typically providing power for only a few minutes.
5. One-Shot Batteries: These batteries exploit the latest technology without recharge capabilities.
They automatically activate based on pressure switches.
6. Ram Air Turbine (RAT): RATs provide continuous emergency power. However, they have drawbacks,
including space requirements, complexity, and the need for a small accumulator to deploy the
turbine.
Additional Facts:
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RATs are small wind-driven turbines that generate power during emergencies. They are typically
deployed when other power sources fail, providing essential functions like flight control and
communication systems.
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Hydraulic accumulators store energy in the form of pressurized fluid. They are crucial for
maintaining hydraulic pressure during emergencies, ensuring critical systems remain operational.
Proof of design
The process of designing a hydraulic system for an aircraft involves rigorous testing to ensure its
functionality and safety. Here are the key steps:
1. Component Qualification: Each individual component undergoes tests to verify that it meets
specifications. These tests include proof and burst pressure tests, fatigue, vibration, acceleration,
and functional assessments.
2. Declaration of Design and Performance: Upon successful component testing, a formal Declaration
of Design and Performance Certificate is issued. This certificate is signed by both the specialist
company responsible for component design and manufacture and the company designing the
aircraft.
3. Hydraulic Test Rig: The entire hydraulic system is assembled into a test rig. This rig mimics the
aircraft’s structure, with correct piping dimensions, shape, and length. Flight-standard pumps
provide the necessary flows and pressures. Loading devices simulate aerodynamic and other loads
on actuators like the undercarriage.
4. Strain Gauging and Load Techniques: Strain gauges and other load measurement techniques are
used to assess forces and stresses during testing. The rig is “flown” for several hundred hours before
actual flight testing on the prototype aircraft.
5. Qualification and Certification: Before an aircraft is accepted into service, the hydraulic system
must be fully qualified. Evidence from both the rig and flight testing supports this qualification.
Despite the considerable cost and effort, a well-designed hydraulic test rig is essential for the formal
qualification and certification of an aircraft.
The iron bird is a ground-based test rig that simulates realistic flight conditions for aircraft systems. It
combines actuators and landing gear, providing a faithful representation of the system. Iron birds allow
safer and more economical testing, leading to approval for actual aircraft testing. This approach aligns
with other major subsystems.
Figure 9 Hydraulic test rig
Figure 10 Hydraulic system testing process
Aircraft system applications
Aircraft hydraulic systems vary based on aircraft type. Examples span single and multiple engine civil and
military aircraft.
The Avro RJ Hydraulic System
The Avro RJ aircraft family includes the RJ70, RJ85, and RJ100 models. These regional jet airliners have a
hydraulic system designed for worldwide operations. It combines the simplicity of a two-engine design with
the backup capabilities of a four-engine system. The system operates at a nominal pressure of 3000 psi and
features controls and annunciations on the pilot’s overhead panel. Fault warnings are indicated through an
amber caption and an audio chime.
Figure 11 RJ146 regional jet hydraulic system
The aircraft’s hydraulic system consists of two independent systems: Yellow and Green. Here’s how they
operate:
1. Each system is normally pressurized by a self-regulating engine-driven pump on the inboard
engines.
2. Both systems have their own independent hydraulic reservoirs, pressurized by regulated air bleed
from their respective engines.
3. Flare-less pipe couplings with swaged fittings ensure reliability and ease of repair.
4. The Yellow and Green systems are geographically segregated within the aircraft. Yellow is on the
left, and Green is on the right. (pilot’s viewpoint)
5. Backup power:
o
For the Yellow system: An AC electric pump serves as backup.
o
For the Green system: A power transfer unit (PTU), driven by the Yellow system, provides
backup power.
6. In case of failures in both systems, an electrically operated DC pump (fed from a segregated
hydraulic supply) enables emergency lowering of the landing gear and operation of the brakes.
7. All these components are housed in a pressurized and vented hydraulic equipment bay, fully
protected from foreign object damage.
The primary power generation components of the Yellow system are:
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Engine Driven Pump (EDP) on No. 2 engine
Standby AC powered hydraulic pump
Emergency DC powered hydraulic pump
Accumulator
Reservoir
All these components, except for the EDP, are located in the hydraulics equipment bay. The components are
shown in Figure 4.16.
Figure 12 Yellow system components
Yellow Hydraulic System
The Yellow system powers the following services:
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1 flap motor
Flap asymmetry brakes
Roll spoilers
2 lift spoilers (inner spoilers on the left and right wing)
1 rudder servo control
Standby fuel pumps (left and right)
Landing gear emergency lock down
Wheel brakes including park brake
Air stairs through the AC pump
Power transfer unit (PTU)
Yellow System Standby AC Pump
Certainly! In the event of an EDP (Engine Driven Pump) failure, the Yellow hydraulic system relies on
a standby AC pump. Let’s break down the details:
1. The standby AC pump is designed to maintain the system pressure at 3000 psi.
2. It serves as a backup when the primary EDP is unavailable.
3. The pump is controlled by a three-position switch located on the hydraulics overhead panel on the
flight deck.
4. This panel also includes amber pump high temperature and failure annunciators for monitoring
pump status.
5. The standby pump can be manually selected on or off by the flight crew.
6. However, it typically operates in automatic mode.
7. In automatic mode: A pressure switch in both the Yellow and Green hydraulic systems monitors the
EDP delivery pressure. If the EDP delivery pressure falls below 1500 psi, the standby
pump automatically activates.
8. The standby pump directly supports the Yellow system and indirectly assists the Green system via
the PTU (Power Transfer Unit).
Yellow System Backup DC Pump
The Yellow System Backup DC Pump serves as an emergency solution when both the Yellow and Green
hydraulic systems fail.
Emergency Functions:
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Locks down the main landing gear.
Operates the Yellow system wheel brakes.
On the ground, it provides brake pressure for parking, starting, or towing.
Components:
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DC powered hydraulic pump
Fluid supply
Accumulator
Segregated reservoir
The Yellow system accumulator is connected to the wheel brakes.
It’s protected by non-return valves, ensuring isolation from other services.
The accumulator is pressurized by the Yellow EDP (Engine Driven Pump), AC pump, or DC pump.
Yellow System Reservoir
A 15.5 liter reservoir is provided for the Yellow system. It is pressurized by bleed air regulated to 50 psi from
the engine HP compressor. The reservoir incorporates the following:
• A pressure gauge
• A sight glass
• An air low pressure switch
• Inward and outward relief valves
• A bursting disc to protect against manual failure of the outward relief valve
• A ground charge connection and manual pressure release lever
• A contents transmitter
Indications of tank contents are provided on the flight deck overhead panel that also includes amber low
quantity and high temperature annunciators.
Engine Driven Pump
The Yellow system Engine Driven Pump (EDP) is mounted on the left inner engine auxiliary gearbox at the
bottom of the engine to ensure easy maintenance access. The EDP has an associated motorized isolation
valve. When the valve is closed it isolates the pump from the tank and provides an idling circuit to offload
the pump. If the engine fire handle is pulled to its fullest extent the valve closes automatically, preventing
more fluid reaching the pump.
A two-position switch on the overhead hydraulic panel controls the position of the EDP isolation
valve. An amber annunciator on the overhead panel illuminates when the valve is travelling and remains on
until it reaches the selected position. The EDP also has an associated relief valve which opens to allow excess
pressure back to the tank at 3500 psi.
Green Hydraulic System
The primary power generation components of the Green system are:
• Engine Driven Pump (EDP) on No. 3 engine
• Power Transfer Unit (PTU)
• Hydraulic reservoir
• Accumulator
All components, except for the EDP, are located in the hydraulic equipment bay. The Green system power
the following:
• 1 flap motor
• 4 lift spoilers (center and outer spoilers on the left and right wing)
• Airbrakes
• Landing gear – normal
• Nose gear steering
• Wheel brakes excluding park brake
Green System Standby PTU
The Power transfer unit (PTU) is an alternative power source for the Green system. The PTU is a back-toback hydraulic motor and pump. It can support all Green system services except for the standby AC/DC
generator. The motor is powered by the Yellow system pressure and is connected by a drive shaft to a pump
in the Green system.
The PTU is controlled from the hydraulics overhead panel by a two-position switch. When the switch is in
the on position it is automatically activated if Green system pressure falls below 2600 psi. With the switch in
the off position, the motor is isolated from the Yellow system by a motorized valve. Movement of the valve
is indicated by an amber PTU VALVE annunciator on the flight deck hydraulics panel. The PTU may also be
used during ground servicing to pressurize the Green hydraulic system, provided the hydraulic reservoir is
fully charged with air.
Green System Standby AC/DC Generator
• The Green hydraulic system can support the electrical system in the event of low electrical power.
• A standby AC/DC generator, driven by a hydraulic motor is powered by the Green system and is
controlled by a three-position switch on the flight deck overhead electrical panel.
• The generator can be selected on or off manually but is usually in automatic standby (ARM) mode.
• The generator is normally isolated from the system pressure by a solenoid operated selector valve.
• When the standby AC/DC pump is operating its selector valve is opened, and at the same time
Green system services are isolated by their shut-off valve.
• Green system services are therefore not available while the generator is operating and the Green
system LO PRESS annunciator is indicated by a white light on the overhead electrical panel
Green System Reservoir
The Green system reservoir has the same capacity as the Yellow system and is charged with bleed air from
No. 3 engine. Its features are exactly the same as the Yellow system reservoir.
Accumulator
The Green system accumulator is identical to the Yellow system accumulator. It maintains stability in the
Green systems during operation of the PTU and also assists the EDP for initial run-up of the standby AC/DC
generator.
The BAE SYSTEMS Hawk 200 Hydraulic System
The BAE SYSTEMS Hawk 200 is a single-engine, single-seat multi-role attack aircraft in which the hydraulic
power is provided by two independent systems. Both power the flying controls by means of tandem
actuators at the ailerons and tail-plane.
The number 1 system provides power to the rudder, which can also be manually operated. The number 1
system also provides power for utility services such as flaps, airbrakes, and landing gear and wheel brakes.
The number 2 system is dedicated to the operation of the flying control surfaces. In the event of engine or
hydraulic pump failure, a ram air turbine driven pump automatically extends from the top rear fuselage into
the airstream. This powers the flying control system down to landing speed.
A pressurized nitrogen accumulator is provided to operate the flaps and landing gear in an emergency, and
wheel brake pressure is maintained by a separate accumulator.
Figure 13 The BAE Systems Hawk 200
hydraulic system
Tornado Hydraulic System
The Tornado is a twin-engine, two-seat, high-performance aircraft designed for ground attack as the IDS
version, or for air defense as the ADV version.
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The Tornado aircraft features a 4000 psi fully duplicated hydraulic system.
The high operating pressure allows the use of small diameter piping.
Low system weight despite the duplicated pipe routings needed for battle damage tolerance.
The two pumps are mounted on the engine gearboxes.
Depressurization: During engine start, the hydraulic system is depressurized to reduce engine
power offtake and allow rapid engine starting.
Cross-Drive Mechanism: A cross-drive is provided between the two RB 199 engines, enabling either
engine to power both hydraulic pumps in case of engine failure.
Accessory Drive Gearboxes (AMADs): The hydraulic pumps in the Tornado aircraft are driven by two
independent accessory drive gearboxes or AMADs. One AMAD is connected by a power offtake shaft
to the right-hand engine, and the other is similarly connected to the left-hand engine.
Hydraulic System Location: This design allows the hydraulic pumps, along with the fuel pumps and
independent drive generators, to be mounted on the airframe and separated from the engine by a
firewall. As a result, the Tornado’s hydraulic system is completely contained within the airframe.
This not only enhances safety but also improves engine change time, since the engine can be
removed without the need to disconnect hydraulic pipe couplings.
Dual System Feeds: The engine intake ramp, taileron, wing-sweep, flap, and slat actuators are all fed
from both systems. In case any part of the utility system becomes damaged, isolating valves operate
to give priority to the primary control actuators.
Undercarriage and Emergency Measures: The undercarriage is powered by the number 2 system. In
the event of a failure, the gear can be lowered by means of an emergency nitrogen bottle.
Additionally, a hand pump is provided to charge the brake and canopy actuators.
Pressure Gauges and Filters: Skin-mounted pressure and contents gauges are conveniently provided
adjacent to the charging points, and all filters are hand-tightened.
Figure 14 The Panavia Tornado hydraulic system
Figure 15 Tornado hydraulic system
Civil Transport Comparison
Examining different philosophies a comparison is made between an Airbus narrow body – the A320 family
and a Boeing wide body – the B767. It is usual for three independent hydraulic systems to be employed,
since the hydraulic power is needed for flight control system actuation.
Hydraulic pumps are driven by:
• Engine driven
• Electrically driven
• Air turbine/bleed air driven
• Ram air turbine driven
Airbus A320
The aircraft is equipped with three continuously operating hydraulic systems called Blue, Green and Yellow.
Each system has its own hydraulic reservoir as a source of hydraulic fluid.
1. The Green system (System 1) is pressurized by an Engine Driven Pump (EDP) located on No. 1 engine
which may deliver 37 gallon per minute or 140 L/min
2. The Blue system (System 2) is pressurized by an electric motor-driven pump capable of delivering
6.1 gpm or 23 L/min. A Ram Air Turbine (RAT) can provide up to 20.6 gpm or 78 L/min at 2175 psi in
emergency conditions
3. The Yellow system (System 3) is pressurized by an EDP driven by No. 2 Engine. An electric motor
driven pump is provided which is capable of delivering 6.1 gpm or 23 L/min for ground servicing
operations. This system also has a hand-pump to pressurize the system for cargo door operation
when the aircraft is on the ground with electrical power unavailable.
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Each channel has the provision for the supply of ground-based hydraulic pressure during
maintenance operations.
Each main system has a hydraulic accumulator to maintain system pressure in the event of
transients.
Each system includes a leak measurement valve (shown as L in a square on the diagram), and a
priority valve (shown as P in a square).
The leak measurement valve is positioned upstream of the primary flight controls and is used for
the measurement of leakage in each flight control system circuit. They are operated from the
ground maintenance panel.
In the event of a low hydraulic pressure, the priority valve maintains pressure supply to essential
systems by cutting of the supply to heavy load users.
The bi-directional Power Transfer Unit (PTU) enables the Green or the Yellow systems to power
each other without the transfer of fluid.
In flight in the event that only one engine is running, the PTU will automatically operate when the
differential pressure between the systems is greater than 500 psi.
On the ground, while operating the yellow system using the electric motor driven pump, the PTU
will also allow the Green system to be pressurized.
The RAT extends automatically in flight in the event of failure of both engines and the APU.
In the event of an engine fire, a fire valve in the suction line between the EDP and the appropriate
hydraulic reservoir made be closed, isolating the supply of hydraulic fluid to the engine.
Pressure and status readings are taken at various points around the systems to be shown on the
Electronic Crew Alerting and Monitoring (ECAM).
Figure 16 Simplified A320 Hydraulic system
Boeing 767
The B767 also has three full-time independent hydraulic systems to assure the supply of hydraulic pressure
to the flight controls and other users. These are the left, right and center systems serviced by a total of eight
hydraulic pumps.
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The left system (Red system) is pressurized by an EDP capable of delivering 37.5 gpm or 142 L/min. A
secondary or demand electric motor driven pump capable of delivering 7 gpm or 26.5 L/min is
turned on automatically in the event that the primary pump cannot maintain pressure.
The right system (Green system) has a similar configuration to the left system.
The center system (Blue system) uses two electric driven motor pumps, each with the capability of
delivering 7 gpm or 26.5 L/min as the primary supply.
An Air-Driven Pump (ADP) with a capacity of 37 gpm or 140.2 L/min is used as a secondary or
demand pump for the center system. The center system also has an emergency RAT rated at 11.3
gpm or 42.8 L/min at 2140 psi.
Primary flight control actuators, autopilot servo-valves and spoilers receive hydraulic power from each of
the three independent hydraulic systems. The stabilizer, yaw dampers, elevator feel units and the brakes are
operated from two systems. A Power Transfer Unit (PTU) between the left and right systems provides a third
source of power to the horizontal stabilizer.
A motorized valve (M) located between the delivery of ACMP #1 and ACMP #2 may be closed to act an
isolation valve between the ACMP #1 and ACMP #2/ADP delivery outputs.
Hydraulic systems status and a synoptic display may be portrayed on the Engine Indication & Crew Alerting
System (EICAS) displays situated between the Captain and First Officer on the instrument console. A number
of maintenance pages may also be displayed.
Figure 17 a Boeing 767 Hydraulic system
The RAT supplies emergency power in flight once the engine speed (N2) has fallen below 50 % on both
engines and the airspeed is in excess of 80 kts. The RAT may only be re-stowed on the ground.
Landing Gear Systems
The Raytheon/BAE 1000 is representative of many modern aircraft; its landing gear is shown in Figure. It
consists of the undercarriage legs and doors, steering and wheels, brakes and anti-skid system. All of these
functions can be operated hydraulically in response to pilot demands at cockpit mounted controls.
Nose Gear
The tricycle landing gear has dual wheels on each leg. The hydraulically operated nose gear retracts forward
into a well beneath the forward equipment bay.
Figure 18 The Raytheon 1000 nose landing-gear
The nose-wheel doors are designed to be normally closed and only open when the nose gear is being
lowered or retracted. This design has two main advantages:
Protection of the undercarriage bay
By keeping the doors normally closed, the undercarriage bay is shielded from spray during take-off and
landing. This helps to protect the internal components of the undercarriage from potential damage or wear
caused by debris or water.
Reduction in drag
Closed doors present a smoother surface to the airflow, reducing aerodynamic drag. This can improve the
aircraft’s performance and fuel efficiency.
A small panel on the leg completes the enclosure when the gear is retracted, providing further protection
and aerodynamic efficiency. A mechanical indicator on the flight deck provides a visual confirmation of the
gear’s locked position, ensuring the safety and proper functioning of the landing gear system.
Main Gear
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The main gear is hydraulically operated and retracts inwards into wheel bay.
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Once retracted, the main units are fully enclosed by fairings attached to the legs and by hydraulically
operated doors.
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Each unit is operated by a single jack.
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A mechanical linkage maintains the gear in the locked position without hydraulic assistance.
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The main wheel doors jacks are controlled by a sequencing mechanism that closes the doors when
the gear is fully extended or retracted.
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The clean lines of the nose wheel bay with the doors shut highlight the aerodynamic efficiency of
this design.
Figure 19 Landing gear sequence
Figure 20 The Raytheon Main landing gear
Braking Anti-Skid and Steering
Stopping an aircraft safely at high landing speeds on a variety of runway surfaces and temperatures, and
under all weather conditions demands an effective braking system. Its design must take into account tire to
ground and brake friction, the brake pressure/volume characteristics, and the response of the aircraft
hydraulic system and the aircraft structural and dynamic characteristics.
Simple systems are available which provide reasonable performance at appropriate initial and maintenance
costs.
More complex systems are available to provide minimum stopping distance performance with features such
as auto-braking during landing and rejected take-off, additional redundancy and self-test.
The normal functions of landing, deceleration and taxying to dispersal or the airport gate require large
amounts of energy to be applied to the brakes. Wherever possible, lift dump and reverse thrust will used to
assist braking.
However it is usual for a large amount of heat to be dissipated in the brake pack. This results from the
application of brakes during the initial landing deceleration, the use of brakes during taxying, and the need
to hold the aircraft on brakes for periods of time at runway or taxiway intersections.
When the aircraft arrives at the gate the brakes, and the wheel assembly will be very hot. This poses a
health and safety risk to ground crew working in the vicinity of the wheels during the turnaround. This is
usually dealt with by training.
Figure 21 Brake control systems - Functional elements
A more serious operational issue is that the aircraft cannot depart the gate until the brake and wheel
assembly temperature cools to a value that will not support ignition of hydraulic fluid.
This is to ensure that, during the taxi back to the take-off runway, further brake applications will not raise
the temperature of the brake pack to a level that will support ignition if a leak of fluid occurs during
retraction.
Departure from the gate, therefore, may be determined by brake temperature as indicated by a sensor in
the brake pack rather than by time taken to disembark and embark passengers.
Some aircraft address this issue by installing brake cooling fans in the wheel assembly to ventilate the
brakes. An alternative method is to install fire detection and suppression systems in the wheel bay.
There are events that can raise the temperature of the brakes to the extent that a fire may occur and the
tires can burst.
 Brake Overheating: Certain events, such as an aborted take-off or a heavy landing, can cause the
brakes of an aircraft to overheat to the point of causing a fire or bursting the tires. In such situations,
thermal plugs are used to deflate the tires and fire crews attend to the aircraft to extinguish the fire
while passengers disembark.
 Anti-skid System: The Dunlop Maxaret unit, a widely known anti-skid system. This system consists
of a hydraulic valve assembly regulated by a spring-loaded flywheel. The system detects skid
conditions and allows hydraulic pressure to dissipate, providing optimal braking force.
Dunlop Maxaret anti-skid system working
 The flywheel in the system rotates with the wheel hub, allowing the entire unit to be housed within
the axle and protecting the unit from weather and stones.
 Skid conditions are detected when the flywheel overruns, which opens the Maxaret valve to reduce
hydraulic pressure and release the brakes.
 The system uses flow-sensitive hydraulic units and switches in the oleo leg to modulate pressure for
optimal braking force.
 The system prevents inadvertent brake application before touchdown by only becoming active after
the oleo switches have sensed that the oleo is compressed, a condition known as ‘weight-onwheels’.
 Without this protection, landing with full braking applied could lead to loss of control of the aircraft
or burst tires.
Figure 22 The Dunlop Maxaret anti-skid system
Electronic Control
Electronic control of braking and anti-skid systems has been introduced in various forms to provide different
features.
Figure 23 Electronic anti-skid system with adaptive pressure control
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The electronic control box contains individual wheel deceleration rate skid detection circuits with
cross reference between wheels.
It has changeover circuits to couple the control valve across the aircraft should the loss of a wheel
speed signal occur.
If a skid develops, the system disconnects braking momentarily.
The adaptive pressure coordination valve ensures that brake pressure is re-applied at a lower
pressure after the skid than the level which allowed the skid to occur.
A progressive increase in brake pressure between skids attempts to maintain a high level of pressure
and braking efficiency.
The adaptive pressure control valve dumps hydraulic pressure from the brake when its first stage
solenoid valve is energized by the commencement of a skid signal.
On wheel speed recovery, the solenoid is de-energized and the brake pressure re-applied at a
reduced pressure level, depending on the time interval of the skid.
Brake pressure then rises at a controlled rate in search of the maximum braking level, until the next
incipient skid signal occurs.
Automatic Braking
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The system allows an aircraft to land and stop without pilot braking intervention.
During automatic braking, a two-position three-way solenoid valve is energized after wheel spinup.
System pressure is fed directly to the anti-skid valves, which modulate pressure at the brakes.
Auto-braking is overridden by pilot intervention in the anti-skid control circuit.
Prerequisites for auto-braking include:
o Auto-brake switch must be on with the required deceleration selected.
o Anti-skid switch must be on and operative.
o Throttle must be correctly positioned.
o Hydraulic pressure must be available.
o Brake pedals must not be depressed.
o Wheels must be spun up.
Auto-braking will retard the aircraft at a predetermined rate unless overridden by anti-skid activity.
Pilot can override auto-braking by advancing throttle levers for go-around or applying normal
brakes.
Figure 24 An Automatic Brake control system
Multi-Wheel Systems
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Large aircraft (e.g., B747-400, B777) feature multi-wheel bogies with more than two main gears.
B747-400 has 4 – main oleos with four wheels each; B777 has two main bogies with six wheels each.
Complex systems use multi-lane dual redundant control for enhanced reliability.
Boeing B777 Landing gear configuration
B777 main gear exemplifies complexity, with wheels grouped in 4 lines of 3-wheels for independent
control channels.
 Brake System Control Unit (BSCU) houses dual redundant controllers for each line of three wheels.
 BSCU interfaces with aircraft through left and right A629 data buses.
 System supplied by Hydro-Aire (Crane Aerospace) showcases sophistication in modern brake
systems.
Figure 25 Simplified B 777 braking configuration
Airbus A380 Landing gear configuration:
 Airbus A380 landing gear includes two 6-wheel under-fuselage and two 4-wheeled wing-mounted
landing gear. The wing-mounted landing gear is slightly forward of the fuselage-mounted gear.
 Goodrich provides carbon brakes on main landing gear wheels for Airbus A380, enhancing braking
efficiency.
Figure 26 Airbus A-380 Landing gear configuration
Nose Landing Gear:
 Supplied by Messier-Dowty.
 Steering control through the nose gear and rear axle of the fuselage landing gear.
 Enables U-turn maneuvers on a 60 m-wide runway.
 Hydraulically steerable aft axle improves maneuverability for tight turns without imposing excessive
torsion loads on the main oleo.
Ground Maneuverability:
 Aircraft can maneuver on 23 m-wide taxiways and 45 m-wide runways.
 Latecoere, based in Toulouse, developed the External and Taxi Aid Camera System (ETACS).
 ETACS includes five video cameras on the tailfin and under the fuselage, relaying data to cockpit
displays for ground maneuvers.
 Integrated with Honeywell terrain guidance and on-ground navigation systems linked to the
aircraft's flight management system.
A380 Braking System:
 Braking system provided by Messier-Bugatti.
 Based on self-adapting braking algorithms from the A340-500/600.
 Optimizes braking by managing wheel-by-wheel and landing-by-landing functions based on runway,
tire, and brake conditions.
 Each wheel is continuously and independently controlled in real-time, considering individual
parameters and environment.
 Utilizes three dedicated computers (RDCs - Remote Data Concentrator) connected to the IMA via a
digital bus.
Figure 27 A380 brake control system
Brake Parachute
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Essential for high-speed landings on short runways in military aircraft.
System can be armed in-flight and activated by a weight-on-wheels switch upon main wheel
touchdown.
Figure displays an F-117 with the brake parachute deployed.
The parachute is jettisoned onto the runway and must be retrieved before the next landing attempt.
Enhances deceleration, allowing for effective use of limited runway space during military operations.
Figure 28 F117 deploying brake parachute
Steering
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Normally disengaged during landing; rudder used until forward speed diminishes its effectiveness.
Manual or automatic engagement occurs when rudder becomes ineffective.
Steering motors respond to rudder pedal demands when nose wheel steering is selected.
Angular range and rate of change of steering angle are calibrated to prevent over-steering or tire
scrubbing during runway and taxiway maneuvers.
Airbus A380 Steering System:
 A380 utilizes nose-wheel steering and after wheels of the main gear.
 Enables the aircraft to complete a 180-degree turn within 56.5 m, well within the standard 60 m
runway width.
 Enhances maneuverability and allows precise control during ground operations.
Figure 29 A380 steering control system