Hydraulic Systems

Hydraulic Systems


A hydraulic system uses a fluid under pressure to drive machinery or move mechanical components.


Virtually all aircraft make use of some hydraulically powered components. In light, general aviation aircraft, this use might be limited to providing pressure to activate the wheel brakes. In larger and more complex aeroplanes, the use of hydraulically powered components is much more common. Depending upon the aircraft concerned, a single hydraulic system, or two or more hydraulic systems working together, might be used to power any or all of the following components:

A hydraulic system consists of the hydraulic fluid plus three major mechanical components. Those components are the “pressure generator” or hydraulic pump, the hydraulically powered “motor” which powers the component concerned and the system “plumbing” which contains and channels the fluid throughout the aircraft as required.

Hydraulic Fluid

Fluid is the medium via which a hydraulic system transmits its energy and, theoretically, practically any fluid could be utilized. However, given the operating pressure (3000 to 5000 psi) that most aircraft hydraulic systems generate in combination with the environmental conditions and strict safety criteria under which the system must operate, the hydraulic fluid that is used should have the following properties:

  • High Flash Point. In the event of a hydraulic leak, fluid ignition should not occur at the normal operating temperatures of the surrounding components. Special hydraulic fluids with fire resistant properties have been developed for aviation use. These fluids are phosphate esters and, unlike mineral oil based hydraulic fluids, they are very difficult to ignite at room temperature. However, if the fluid is heated to temperatures in excess of 180 degrees C, it will sustain combustion. The auto-ignition temperature of most aviation hydraulic fluids is in the range of 475 degrees C.
  • Adequate Viscosity. Aircraft hydraulic systems must work efficiently over a broad temperature spectrum. The fluid used must flow easily at very low temperatures but must also maintain adequate viscosity at high temperatures. The ideal hydraulic fluid will have a very low freezing point and a very high boiling point.
  • Lubricant Properties. The hydraulic fluid acts as a lubricant for the pumps, actuators and motors within the system. The fluid should have anti-corrosion properties and be thermally stable.
  • Thermal Capacity/Conductivity. Hydraulic fluid acts as a system coolant. The fluid must be able to readily absorb and release heat.

Hydraulic Pumps

Several types of hydraulic pumps driven by a variety of power sources can be found in aviation applications. Pumps include:

  • Gear Pumps. Gear pumps use meshing gears to pump fluid. Gear pumps are fixed displacement type pumps as they move a specific amount of fluid per rotation. Gear pumps may be used on low pressure systems (under 1500 psi) but are generally not suitable for high pressure applications
  • Fixed Displacement Piston Pumps. Piston pumps utilize a piston moving in a cylinder to pressurize a fluid. A fixed displacement pump moves a specific amount of fluid with each stroke.
  • Variable Displacement Piston Pumps. This is the most common type of pump on large aircraft. The variable displacement design allows the pump to compensate for changes in the system demand by increasing or decreasing the fluid output. This allows near constant system pressure to be maintained.

The motive power for these pumps can be generated by a wide variety of options inclusive of:

  • Manual. In many light aircraft, a manual hydraulic pump provides pressure for wheel brakes or flap extension and retraction.
  • Engine Driven. Pumps are frequently mounted on the engine accessory gear box.
  • Electric. Both AC and DC motors are utilized to power hydraulic pumps with three phase AC motors being most common.
  • Pneumatic. Bleed air powered motors are utilized on some aircraft to drive hydraulic pumps.
  • Hydraulic. A Power Transfer Unit (PTU) allows the hydraulic pressure of one hydraulic system to drive a pump to pressurize a second hydraulic system without any transfer of hydraulic fluid. Depending upon the installation, a PTU can be single or bi-directional.
  • Ram Air Turbine. In the event of an emergency, some aircraft have a Ram Air Turbine (RAT) that can be extended into airstream to generate hydraulic pressure.

Hydraulic Motors and Cylinders

Hydraulic motors and cylinders utilize pressurized fluid to do mechanical work.

  • Hydraulic Motors. A hydraulic motor is a mechanical device that converts hydraulic pressure and flow into torque and angular displacement or rotation. Various types of hydraulic motors such as gear, vane and radial piston motors are available. On aircraft, hydraulic motors are most often used to drive jackscrews which can in turn be utilized to power flaps, stabilizer trim and some vertically extending landing gear applications such as found on the LOCKHEED C-130 Hercules aircraft.
  • Hydraulic Cylinders. A hydraulic cylinder, sometimes referred to as a linear hydraulic motor or a hydraulic actuator, is a mechanical actuator that is used to provide a reversible force in a single direction. The hydraulic cylinder consists of a cylinder barrel within which a piston connected to a piston rod ues hydraulic pressure to move back and forth. Aircraft applications include landing gear extension and retraction, cargo door operation and movement of flight control surfaces.

System "Plumbing" Components

Aviation hydraulic systems, in general, are of the "open loop" variety drawing fluid from a reservoir, pressurizing it and making it available to the various user components before returning the fluid to the reservoir. The primary components of the "plumbing" portion of the hydraulic system include the following:

  • Reservoir. Hydraulic fluid reservoirs are required by most aircraft systems to provide a ready source of fluid for the hydraulic pump(s) and to contain a varying volume of fluid. This variance results from differential actuator volume (dependent upon whether the actuator is extended or retracted) and for fluid thermal contraction or expansion. The reservoir size is optimized so that only the amount of fluid needed for proper function is carried. In many installations, bleed air is used to pressurize or "bootstrap" the reservoir to help prevent hydraulic pump cavitation.
  • Filters. Hydraulic fluid cleanliness is essential to proper system function. In-line filters are incorporated into the hydraulic system to remove any contaminants from the fluid.
  • Shut Off Valves. Hydraulic shut off valves are usually installed at the engine firewall. In the event of an engine fire, the shutoff valve is closed to prevent possible ignition of the hydraulic fluid.
  • Control Valves. Hydraulic motors and actuators have an associated control valve which is positioned in response to a manual or automated system selection such as moving the flap lever. The control valve responds to that selection by positioning to allow pressurised hydraulic fluid to flow into the motor or actuator in the appropriate direction.
  • Pressure Relief Valve. In some systems, especially those utilising a fixed displacement pump, pressure relief valves are incorporated to ensure that nominal system pressure is not exceeded. If system pressure becomes too high, the relief valve opens and fluid is returned to the reservoir.
  • Hydraulic Fuses. Hydraulic fuses are in-line safety devices designed to automatically seal off a hydraulic line if pressure becomes too low.
  • Accumulators. A hydraulic accumulator is a pressure storage reservoir in which hydraulic fluid is held under pressure by an external source of energy. The external source can be a spring or a compressed gas. An accumulator enables a hydraulic system to cope with extremes of demand using a less powerful pump and to respond more quickly to a temporary demand. It also acts as a system shock absorber by smoothing out pulsations. In the event of a hydraulic pump failure, the energy stored in an accumulator can provide a limited number of brake applications after landing.

Hydraulic System Redundancy

Hydraulic system redundancy is achieved by two primary means - multiple systems and multiple pressure sources within the same system.

  • Multiple Pressure Sources. Hydraulic systems often have more than one pump available to pressurise the system. It is quite common for a system to have one or more engine driven pumps plus one or more electric pumps. In some cases, a manual pump is also incorporated. Some systems only use the electric or manual pumps while on the ground when the engines are not operating. Others use the electric pump(s) to provide an additional pressure source during high demand situations such as gear retraction or as the primary pressure source in the event of the loss of the engine driven pump(s). When an electric pump is used as the primary pressure source, a second electric pump or a Ram Air Turbine might be incorporated into the system as a backup source of hydraulic pressure. Provision of multiple pressure sources helps to ensure that the entire hydraulic system is not lost in the event of a single component failure.
  • Multiple Hydraulic Systems. In many aircraft, flight control surfaces are hydraulically actuated. In these cases, multiple actuators on each surface, powered from multiple hydraulic systems, are essential to ensure that the failure a hydraulic system will not result in loss of control. In modern commercial aircraft, it is common to power the flight control surfaces from three independant hydraulic systems. The control surface architecture allows for failure of two of those systems without compromising control.


Hydraulic systems are subject to several significant threats. These include:

  • System Overheat. The system exceeds its maximum allowable operating temperature and must be de-energized.
  • Loss of System Pressure. Loss of system pressure can occur in two different ways; loss of fluid or failure of a hydraulic pump.
  • Hydraulic Fluid Contamination. Contamination can be chemical or particulate in nature and can be caused during fluid production, by improper servicing of the hydraulic system or by a component failure.


Hydraulic system overheat, loss of pressure or fluid contamination can all result in the loss of the hydraulic system and the loss of function of those components that it powers. Fluid contamination can also result in loss of hydraulic system efficiency, fluid leaks, excessive component wear and premature component failure.


The primary defense against hydraulic fluid contamination lies in robust maintenance practices. Any fluids used to service the system must be as specified in the AOM and fluid types should not be mixed. Care should be taken to ensure that the fluid is not contaminated prior to use and that no contaminants are introduced to the system while topping up the fluid. System filters should be cleaned or replaced as per manufacturer's guidelines.

In the event of a System Overheat or Loss of Pressure, following the Quick Reference Handbook (QRH) or ECAM checklists may result in recovery of the system. If the loss of pressure was as a result a total of loss of hydraulic fluid, the system is not recoverable.

Typical Scenarios

  • A transatlantic flight gradually loses hydraulic fluid from one of the hydraulic systems. When the Electronic Centralized Aircraft Monitor (ECAM) annunciates the system Loss of Pressure, appropriate actions are completed to secure the system.The Captain contacts Company maintenance to discuss options. Due to the fact that there is no significant loss of aircraft capability due to system redundancy, that a few hours of flight time to destination remain and that there were ample diversion airfields enroute, it is decided to continue to the planned destination. The flight lands without further incident.
  • Just after top of climb, a turboprop aircraft suffers a complete loss of the Utility hydraulic system due to a rudder actuator failure. Controlability of the aircraft is not an issue as the Booster hydraulic system also provides pressure to the flight control surfaces. The Captain electes to make a turnback. An alternate (gravity) gear extension is carried out and a flapless approach and landing is made. The Auxillary hydraulic system is used to provide pressure for the brakes and full reverse propeller pitch is used to help bring the aircraft to a stop. The aircraft is shut down on the runway and then towed to the ramp.

Accidents & Incidents

On 7 April 2022, a Boeing 757-200F returning to San Jose after a left side hydraulics failure and MAYDAY declaration suddenly veered off the right hand side of the landing runway there during deceleration and passage over uneven ground led to landing gear collapse and significant fuselage structural damage. This runway excursion immediately followed simultaneous advancement of both thrust levers after their prior asymmetric movement earlier in the landing roll and resulted in high left thrust concurrent with idle thrust on the right. With no airworthiness aspect identified, the excursion was attributed to unintended thrust lever selection by the crew.

On 29 September 2019, an Airbus A330-200 received simultaneous indications of low pressure in two hydraulic systems soon after takeoff. An emergency was declared and a return to land was followed by a stop on the runway due to a burst main wheel tyre. A manual valve for one of the hydraulic systems located in the left main gear wheel well had completely detached and impact-damaged a pipe in a nearby but separate hydraulic system. Both systems lost their fluid with valve detachment attributed to fatigue failure of the attachment screws, a risk addressed by an un-adopted non-mandatory Service Bulletin. 

On 15 December 2019, an Airbus A330-200 turned back to Sydney shortly after departure when a major hydraulic system leak was annunciated. The return was uneventful until engine shutdown after clearing the runway following which APU use for air conditioning was followed by a gradual build up of hydraulic haze and fumes which eventually prompted an emergency evacuation. The Investigation found that fluid leaking from ruptured rudder servo hose had entered the APU air intake. The resulting evacuation was found to have been somewhat disorganised with this being attributed mainly to a combination of inadequate cabin crew procedures and training.

On 23 July 2011, a Boeing 737-300 being operated by Jet2.com on a passenger flight from Leeds/Bradford to Paris CDG experienced violent vibration from the main landing gear at touch down in normal day visibility on runway 27R at a normal speed off a stabilised approach. This vibration was accompanied by lateral acceleration that made directional control difficult but the aircraft was kept on the runway and at a speed of 75 knots, the vibrations abruptly stopped. Once clear of the runway, the aircraft was stopped and the engines shutdown prior to a tow to the gate. None of the 133 occupants were injured.

On 15 October 2015 a Boeing 747-300 experienced significant vibration from one of the engines almost immediately after take-off from Tehran Mehrabad. After the climb out was continued without reducing the affected engine thrust an uncontained failure followed 3 minutes later. The ejected debris caused the almost simultaneous failure of the No 4 engine, loss of multiple hydraulic systems and all the fuel from one wing tank. The Investigation attributed the vibration to the Operator's continued use of the engine without relevant Airworthiness Directive action and the subsequent failure to continued operation of the engine after its onset.

On 1 November 2011, a Boeing 767-300 landed at Warsaw with its landing gear retracted after declaring an emergency in anticipation of the possible consequences which in this event included an engine fire and a full but successful emergency evacuation. The Investigation attributed inability to achieve successful gear extension using either alternate system or free fall to crew failure to notice that the Battery Busbar CB which controlled power to the uplock release mechanism was tripped. Gear extension using the normal system had been precluded in advance by a partial hydraulic system failure soon after takeoff from New York.

An announcement by the Captain of a fully-boarded Boeing 757-200 about to depart which was intended to initiate a Precautionary Rapid Disembarkation due to smoke from a hydraulic leak was confusing and a partial emergency evacuation followed. The Investigation found that Cabin Crew only knew of this via the announcement and noted subsequent replacement of the applicable procedures by an improved version, although this was still considered to lack resilience in one respect. The event was considered to have illustrated the importance of having cabin crew close to doors when passengers are on board aircraft on the ground.

On 4 October 2014, the fracture of a hydraulic hose during an A330-200 pushback at night at Karachi was followed by dense fumes in the form of hydraulic fluid mist filling the aircraft cabin and flight deck. After some delay, during which a delay in isolating the APU air bleed exacerbated the ingress of fumes, the aircraft was towed back onto stand and an emergency evacuation completed. During the return to stand, a PBE unit malfunctioned and caught fire when one of the cabin crew attempted to use it which prevented use of the exit adjacent to it for evacuation.

On 26 February 2013, the crew of a Boeing 752 temporarily lost full control of their aircraft on a night auto-ILS approach at Keflavik when an un-commanded roll occurred during flap deployment after an earlier partial loss of normal hydraulic system pressure. The origin of the upset was found to have been a latent fatigue failure of a roll spoiler component, the effect of which had only become significant in the absence of normal hydraulic pressure and had been initially masked by autopilot authority until this was exceeded during flap deployment.

On 17 January 2007, a Bombardier CRJ 100 being operated by French airline Brit Air on a scheduled night passenger flight from Paris CDG to Southampton could not be directionally controlled after touchdown on a dry surface in normal visibility and almost calm winds and departed the side of the runway during the landing roll. There were no injuries to any of the 36 occupants and there was no damage to the aircraft.

On 22 June 2009, an Airbus A340-300 being operated by Finnair suffered a single tyre failure during take off on a scheduled passenger flight to Helsinki and malfunction assessed as consequential by the flight crew occurred to the hydraulic system. The flight proceeded to destination and carried out a daylight landing there in normal visibility without any further aircraft damage. Because of a further deterioration in the status of the aircraft hydraulic systems during the landing roll, the aircraft was stopped on the runway and then towed into the gate. No persons were injured in this incident.

On August 12, 1985 a Boeing 747 SR-100 operated by Japan Air Lines experienced a loss of control attributed to loss of the vertical stabiliser. After the declaration of the emergency, the aircraft continued its flight for 30 minutes and subsequently impacted terrain in a mountainous area in Gunma Prefecture, Japan.

On 18 June 1998, the crew of a Swearingen SA226 did not associate directional control difficulty and an extended take off ground run at Montreal with a malfunctioning brake unit. Subsequent evidence of hydraulic problems prompted a decision to return but when evidence of control difficulties and fire in the left engine followed, a single engine diversion to Mirabel was flown where, just before touchdown, the left wing failed upwards. All occupants were killed when the aircraft crashed inverted. The Investigation found that overheated brakes had caused an engine nacelle fire which spread and eventually caused the wing failure.

On 19 July 1989, a GE CF6-6D-powered Douglas DC-10-10 at FL370 suffered a sudden explosive failure of the tail-mounted number 2 engine and a complete loss of hydraulics so that the aircraft could only be controlled by varying thrust on the remaining two engines. With only limited flightpath control, the subsequent Sioux City emergency landing led to the destruction of the aircraft by impact and fire. The Investigation attributed the engine failure to non-identification of a fan disc fatigue crack arising from a manufacturing defect and the loss of hydraulics to debris dispersal which had exceeded the system s certification protection.

On 16 May 1995, an RAF BAe Nimrod on an airworthiness function flight caught fire after an electrical short circuit led indirectly to the No 4 engine starter turbine disc being liberated and breaching the No 2 fuel tank. It was concluded by the Investigation that the leaking fuel had then been ignited by either the electrical arcing or the heat of the adjacent engine. After the fire spread rapidly, the risk of structural break up led the commander to ditch the aircraft whilst it was still controllable. This was successful and all seven occupants were rescued.

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