GB2623951A

GB2623951A - An aircraft power system

Info

Publication number: GB2623951A
Application number: GB2216061.8A
Authority: GB
Inventors: David Sadler Andrew
Original assignee: Airbus Operations Ltd
Current assignee: Airbus Operations Ltd
Priority date: 2022-10-31
Filing date: 2022-10-31
Publication date: 2024-05-08
Also published as: WO2024094352A1; GB202216061D0

Abstract

An aircraft power system 800 for driving wheels 120, 130 of an aircraft for taxiing comprises a pressurised hydraulic reservoir 801, 803 and a bidirectional hydraulic pump 212 for pumping hydraulic fluid to and from the reservoir, and connectable to drive the wheels. The system further comprises an electric motor 230 able to drive the wheels, such that the taxiing can be powered by the electric motor and/or the hydraulic pump. The system is operable in an electrical generation mode in which the hydraulic pump 212 is connected to the electric motor 230 e.g. via a transfer box 220 such that the hydraulic pump drives the electric motor as a generator, to supply electrical energy.

Description

AN AIRCRAFT POWER SYSTEM

TECHNICAL FIELD

100011 The present invention relates to an aircraft power system, an aircraft landing gear drive system, an aircraft and a method of operating all aircraft.

BACKGROUND

[0002] Typically, an aircraft taxis to and from a runway using thrust from its engines.

In an effort to reduce fuel usage, there has been an effort to develop alternative means to enable aircraft taxiing. It is well known in the prior art to use an aircraft power system to drive a driveable component of an aircraft (for example, landing gear extension/retraction mechanisms, high lift device mechanisms, or landing gear brakes). Such driveable components are typically hydraulically actuated, with the aircraft power system providing hydraulic power to actuate the driveable component.

I00031 It has been proposed to use a hydraulic aircraft power system to drive the wheels of an aircraft to enable the aircraft to taxi. For example, Chinese patent no. CN106741877 describes a hydraulic aircraft power system for driving the wheels of the aircraft to enable taxiing. The hydraulic aircraft power system includes a hydraulic fluid reservoir, which provides a store of hydraulic power. The aircraft wheels are each connected to a hydraulic motor, which is connected to the hydraulic reservoir such that the hydraulic aircraft power system can be used to drive the hydraulic motor (and thereby the wheels) to enable taxiing. As hydraulic fluid from the reservoir is used to diive the hydraulic motors, the reservoir is depleted. The hydraulic aircraft power system therefore also includes an electric motor connected to a hydraulic pump. Using electrical power from an aircraft electrical power system, the electric motor drives the hydraulic pump to pump hydraulic fluid back into the reservoir. Thus, the electric motor and hydraulic pump operate to convert electrical power into hydraulic power. In addition, the hydraulic motors are configured to also operate as hydraulic pumps. Thus, when the aircraft is braking, the hydraulic motors use the kinetic energy of rotation of the wheels to pump hydraulic fluid into the reservoir. Thus, the hydraulic motor also operates to recover kinetic energy to recharge the hydraulic reservoir.

[0004] EP 4005920A1 discloses a hybrid aircraft power system comprising a hydraulic reservoir, a bidirectional hydraulic pump for pumping hydraulic fluid to and from the reservoir, and an electric motor. The electric motor is connectable to a first driveable component (one or more landing gear wheels) of the aircraft such that the electric motor is arranged to drive the first driveable component of the aircraft. The hydraulic pump is connectable to the first driveable component of the aircraft such that the hydraulic pump is arranged to pump hydraulic fluid from the reservoir to drive the first driveable component of an aircraft. Thus, in a first driveable mode of operation, the first driveable component is driven by both the electric motor and the hydraulic pump.

[0005] Such a system allows for both hydraulic and electric power to be utilised to drive a driveable component of the aircraft. Thus, the driveable component can be driven by both the hydraulic pump and the electric motor at the same time. This enables the electric motor to be designed for a "steady state" or normal power requirement of the driveable component, because the hydraulic pump can provide an additional power requirement to the driveable component on a temporary/short-duration basis, when needed (e.g. to provide a peak power demand of 90 kW). This means that it is possible, in embodiments, for the electrical motor to be optimised for its most usual power output. The electric motor may also be designed for much lower required levels and so means that the system is much smaller than would otherwise be required. In embodiments, this enables maintenance, weight and cost savings, as well as a reduction in emissions and fuel burn.

[0006] However, it is desired to optimise such a system as far as possible to ensure the aircraft is as efficient as possible, in terms of overall weight.

[0007] The present invention seeks to mitigate the above-mentioned problems.

Alternatively or additionally, the present invention seeks to provide an improved aircraft power system.

SUMMARY

[0008] A first aspect of the present invention provides an aircraft power system comprising a hydraulic reservoir, a bi-directional hydraulic pump for pumping hydraulic fluid to and from the reservoir, and an electric motor, wherein the hydraulic pump is connectable to the electric motor such that the hydraulic pump is arranged to drive the electric motor as a generator, such that, in an electrical generation mode of operation, the electric motor is driven by the hydraulic pump.

[0009] The aircraft power system may further comprise a first driveable component of the aircraft, wherein the hi-directional hydraulic pump is arranged such that, in a regenerating mode of operation, the hi-directional hydraulic pump is externally driven by the first driveable component to pump hydraulic fluid into the hydraulic fluid reservoir. This enables energy (e.g. kinetic energy) from the first driveable component to be stored as hydraulic energy. This energy can then he later used to drive the electric motor.

[0010] The first driveable component may be a moveable component, and thus have a kinetic energy that may be recovered (e.g. to the hydraulic fluid reservoir). For example, the first driveable component may comprise one or more landing gear wheels of a landing gear drive system.

[0011] Such a system allows the hydraulic energy to be effectively converted to electrical energy. This provides for a more flexible, efficient power system that can be lighter weight hut still suitably sized for the different power demands of the aircraft. For example, charging the electrical system is a means to deplete the hydraulic accumulator without just converting the energy to heat. This may be especially useful prior to braking. For example, an empty hydraulic accumulator (or at least one with some unused capacity) would then allow the hydraulic motor to be used during braking to increase the deceleration capability and further reduce brake use and brake wear.

[0012] The electrical generation mode may be known as an extended regenerative mode.

[0013] The electric motor arrangement may have a torque density of greater than 40NM per kg at 1000RPM and greater than 18NM per kg at 7500RPM. The electric motor arrangement may have a power density of greater than 7kW per kg.

[0014] The hydraulic reservoir may comprise a hydraulic fluid accumulator. Thus, the hydraulic fluid reservoir may be arranged to hold hydraulic fluid under pressure. Thus, the system allows hydraulic energy to be stored in the reservoir, for example in a hydraulic replenishing mode of operation. In such a mode of operation, the hydraulic pump may operate to store hydraulic power in the reservoir. This provides a means of storing surplus power as hydraulic power for later use to, for example drive the electric motor and/or drive a driveable component (for example, in circumstances where excess power is temporarily required by the driveable component of the aircraft). Hence, such a system is able to store surplus energy for later use.

[OW 5] The hydraulic arrangement may have a working pressure of up to around 400bar.

The hydraulic arrangement may have an energy density of greater than 3.2Wh per kg.

[0016] The electric motor and hydraulic pump may be connected to a transfer unit.

Thus, it may be that the transfer unit connects the electric motor to the hydraulic pump.

[0017] The transfer unit may be connectable to a first driveable component of the aircraft. It will he appreciated that such a connection enables the transfer of mechanical power. The transfer unit may be configured to connect and disconnect the hydraulic pump from at least one (for example, both) of the electric motor and the first driveable component. The transfer unit may be configured to connect and disconnect the electric motor from at least one (for example, both) of the hydraulic pump and the first driveable component. Thus, the transfer unit may be configured to enable each of the hydraulic pump, the electric motor, and the first driveable component to be independently connected and disconnected from the other two of the hydraulic pump, the electric motor, and the first driveable component.

[0018] Preferably, the aircraft power system further comprises an electric battery and wherein electrical energy generated in the electrical generation mode of operation is used to replenish the battery.

[0019] This allows the electrical energy generated to be stored for later use. Hence, it allows a hydraulic energy store to be converted to an electrical energy store.

[0020] Preferably, the aircraft power system further comprises an electric distribution network and wherein electrical energy generated in the electrical generation mode of operation is distributed by the electric distribution network.

[0021] For example, the electrical energy could be distibuted around the aircraft to a variety of different electric devices.

[0022] Preferably, the hydraulic reservoir is connectable to the electric motor by a switch arrangement, such that the switch arrangement can select the electrical generation mode.

[0023] The switch arrangement may select the electrical generation mode based on the status of the hydraulic reservoir (for example, if it is full). The switch arrangement may select the electrical generation mode based on the status of the aircraft (for example, if it has decelerated after landing and is being taxi driven by its engines). The switch arrangement may be able to select any number of other modes of operation as well.

[0024] Preferably, the aircraft power system further comprises a first driveable component of the aircraft and wherein the hydraulic pump and/or the electric motor is connectable to the first driveable component of the aircraft.

[0025] The first driveable component may be disconnectable from both the hydraulic reservoir and the electric motor. This disconnection could occur in the electrical generation mode.

[0026] Such a connection allows the hydraulic pump and/or electric motor to be used to drive the first driveable component, for example in a driveable mode of operation. Such a connection allows the hydraulic pump and/or electric motor to be driven by the first driveable component, for example in a regenerating mode of operation.

[0027] More preferably, the electric motor is connectable to the first driveable component of the aircraft such that the electric motor is arranged to drive the first driveable component of the aircraft, and wherein the hydraulic pump is connectable to the first driveable component of the aircraft such that the hydraulic pump is arranged to drive the first driveable component, such that, in a first driveable mode of operation, the first driveable component is driven by both the electric motor and the hydraulic pump.

[0028] Such a system allows for both hydraulic and electric power to be utilised to drive a driveable component of the aircraft. Thus, the driveable component can be driven by both the hydraulic pump and the electric motor at the same time. This enables the electric motor to be designed for a "steady state" or normal power requirement of the driveable component, because the hydraulic pump can provide an additional power requirement to the driveable component on a temporary/short-duration basis, when needed (e.g. to provide a peak power demand of 90 kW). This means that it is possible, in embodiments, for the electrical motor to be optimised for its most usual power output. The electric motor may also be designed for much lower required levels and so means that the system is much smaller than would otherwise be required. In embodiments, this enables maintenance weight and cost savings, as well as a reduction in emissions and fuel burn.

[0029] Even more preferably, the electric motor and hydraulic pump are connected to each other in parallel such that, in the first driveable mode of operation, both arc arranged to supply power to the driveable component independently of the other.

[0030] Thus, in the first driveable mode of operation, the electric motor and hydraulic pump may each be capable of supplying, and arranged to supply, power to the driveable component independently of the other. This ensures that, if one of the electric motor and hydraulic pump should fail, power is still provided to the driveable component by the other. It also means that each can he efficiently designed for their respective power outputs, rather than for the combined power output. Thus, the aircraft power system may be configured to operate in a first driveable mode of operation. Operating the aircraft power system in the first driveable mode of operation may comprise connecting the electric pump and the hydraulic pump to the first driveable component (for example, by use of the transfer unit). It may be that the hydraulic pump is configured to operate (for example, in the first driveable mode of operation) as a hydraulic motor under the action of hydraulic fluid flowing from the hydraulic fluid reservoir.

[0031] Preferably, the electric motor is connectable to the hydraulic pump such that, during a hydraulic replenishing mode of operation, the electric motor is arranged to drive the hydraulic pump to pump hydraulic fluid into the reservoir.

[0032] Thus, in the hydraulic replenishing mode of operation, electrical energy (for example, from a battery or an electric distribution network) is used to store hydraulic power by driving the hydraulic pump to pump hydraulic fluid to the hydraulic reservoir. This may be through the transfer unit, if present. In certain embodiments, the electric motor and the hydraulic pump may be connected to one another directly. Thus, the aircraft power system may be configured to operate in a hydraulic replenishing mode of operation. Operating the aircraft power system in the hydraulic replenishing mode of operation may comprise connecting the electric motor to the hydraulic pump (for example, by use of the transfer unit). Operating the aircraft power system in the hydraulic replenishing mode of operation may comprise disconnecting the electric motor and the hydraulic pump from the first driveable component (for example, by use of the transfer unit). Alternatively, or additionally, the hydraulic pump is electrically connectable to a second electric motor such that the second electric motor is arranged to drive the hydraulic pump to pump hydraulic fluid into the reservoir. The different electric motor may be provided with electrical power from the aircraft APU, a hydrogen fuel cell, a battery, or the aircraft engines.

[0033] Preferably, in a regenerating mode of operation, the electric motor is externally driven by a moveable component of the aircraft to generate electricity.

[0034] The electric motor may also function as a generator. In such a case, it may be that, in a regenerating mode of operation, one or both of the electric motor and the hydraulic pump is externally driven by a moveable component of the aircraft (for example, the first driveable component) to store the kinetic energy from the moveable component.

[0035] The moveable component may he the first driveable component. Thus, the aircraft power system may be configured to operate in a regenerating mode of operation. Operating the aircraft power system in the regenerating mode of operation may comprise connecting at least one (for example, both) of the electric motor and the hydraulic pump to the first driveable component (for example, by use of the transfer unit). Operating the aircraft power system in the regenerating mode of operation may comprise disconnecting the electric motor and the hydraulic pump from the first driveable component (for example, by use of the transfer unit).

[0036] It may be that operating in the regenerating mode of operation comprises prioritising energy recovery by the hydraulic pump over energy recovery by the electric motor. Thus, it may be that energy recovered by operating in the regenerating mode is preferentially stored as hydraulic energy (for example, stored in the hydraulic fluid reservoir), rather than electrical energy (for example, stored in the battery). It may be that more than 50%, preferably more than 60%, more preferably more than 70%, yet more preferably more than 80% of the recovered energy is stored as hydraulic energy.

[0037] Prioritising energy recovery by the hydraulic pump over energy recovery by the electric motor can enable recovery of a greater percentage of the kinetic energy of the first driveable component. The effectiveness of electrical energy recovering using the electric motor is limited by the characteristics of the electric motor. For example, energy recovery using the electric motor requires a minimum commutation speed of the electric motor shaft. Furthermore, energy recovery using the electric motor is also constrained by the size and power rating of the electric motor. By contrast, it is possible for energy recovery using the hydraulic pump to be effective at all rotor speeds. Therefore, in embodiments. prioritising hydraulic energy recovery over electrical energy recovery enables a more efficient and complete recovery of kinetic energy from the first driveable component.

[0038] The aircraft power system may further comprise a hydraulic sensor (for example, a pressure sensor) configured to determine a state of charge of the hydraulic fluid reservoir. The hydraulic sensor may be configured to generate hydraulic sensor data indicating the determined state of charge of the hydraulic fluid reservoir. The aircraft power system may further comprise a battery sensor configured to determine a state of charge of the battery. The battery sensor may be configured to generate battery sensor data indicating the determined state of charge of the battery. The aircraft power system may be configured to operate on the basis of one or both of the hydraulic sensor data and the battery sensor data Such data may, at least in part, be provided as digital data. Such data may, at least in part, be provided as analogue signals.

[0039] The aircraft power system may be configured to prioritise energy recovery by one of the hydraulic pump and the electric motor (for example, on the basis of one or both of the hydraulic sensor data and the battery sensor data). The aircraft power system may be configured to preferentially recover energy via the hydraulic pump (for example, unless or until the hydraulic sensor data indicates that the determined state of charge of the hydraulic fluid reservoir exceeds a predetermined threshold). The aircraft power system may be configured to preferentially recover energy via the hydraulic pump by recovering energy using only the hydraulic pump. The aircraft power system may be configured to recovery energy only via the hydraulic pump until the state of charge of the hydraulic fluid reservoir exceeds a predetermined threshold. It may be that, once the state of charge of the hydraulic fluid reservoir exceeds the predetermined threshold, the aircraft power system ceases to recover energy using the hydraulic pump. The aircraft power system may subsequently recover energy using the electric motor. Thus, the prioritisation of hydraulic energy recovery may be performed by exclusively recovering energy via the hydraulic pump for a first period of time before exclusively recovering energy via the electric motor for a second period of time. It will be appreciated that the first and second periods of time may be defined in terms of a quantity of energy recovered, rather than as fixed time periods.

[0040] It may be that the aircraft power system is configured in certain embodiments to simultaneously recover energy electrically and hydraulically whilst prioritising hydraulic energy recovery. In such cases, it may be that the transfer box is configured to perform torque splitting to direct the kinetic energy from the first driveable component to the hydraulic pump and the electric motor according to a predetermined torque splitting ratio. For example, the transfer box may be configured to perform torque splitting such that 70% the kinetic energy of the first driveable component is directed to the hydraulic pump (i.e. a torque splitting ratio between the hydraulic pump and the electric motor of 7:3). It will be appreciated by the skilled person that other torque split ratios are also equally possible. The torque splitting ratio may be in the range of 95:5 to 55:45, for example, in favour of prioritising hydraulic energy recovery -for at least some of the time. It may be that the aircraft power system is configured to adjust the torque splitting ratio (for example, on the basis of the hydraulic sensor data and the battery sensor data). The aircraft power system may be configured to alter the torque splitting ratio on the basis of one or both of the hydraulic sensor data and the battery sensor data. For example, the aircraft power system may be configured to alter the torque splitting ratio in response to the hydraulic sensor data indicating that the hydraulic fluid reservoir has reached a predetermined state of charge. In such a case, it may be that the predetermined state of charge is fully charged, and the alteration to the torque splitting ratio is such that the aircraft power system ceases to recover energy via the hydraulic pump.

[0041] The aircraft power system may comprise a processor and associated memory.

The processor may be configured to (for example, by execution of instructions stored in the associated memory) control operation of the aircraft power system (including, for example, operation of the hydraulic sensor and the battery sensor).] [0042] Preferably, the hydraulic pump is also connectable to other driveable components of the aircraft.

[0043] Hence, in a second driveable mode of operation, the hydraulic pump can provide hydraulic power to one or more other driveable components. These other driveable components may include: landing gear extension/retraction mechanisms, landing gear bay door mechanisms, cargo door mechanisms, landing gear braking, high lift device mechanisms (for example, flaps and/or slats), flight control surfaces or a hybrid propulsion system.

[0044] Preferably, the electric motor is connectable to the hydraulic pump such that the electric motor is arranged to drive the hydraulic pump to pump hydraulic fluid out of the reservoir.

[0045] Hence, in the second driveable mode of operation, the electric motor may drive the hydraulic pump to provide hydraulic power to the other driveable components. It may be that this occurs and for example is configured to so occur, when the first driveable component does not need to be driven by the electric motor.

[0046] Preferably, the first driveable component comprises one or more landing gear wheels.

[0047] Hence, in the first driveable mode of operation the electric motor (and hydraulic pump) drives the landing gear wheels to move the aircraft over the ground (this is known as e-taxiing). Here, the electric motor can be optimised for driving the landing gear wheels at a constant velocity over level ground, for example. In such embodiments, if there is an uphill slope, or when the landing gear wheels arc to accelerate (for example, when starting to e-taxi the aircraft) it may then be the case that the hydraulic pump is used to provide additional power.

[0048] In a regenerating mode of operation (in the case where there are one or more driven landing gear wheels) it may be that the electric motor is externally driven by the landing gear wheel(s), for example after aircraft landing (e-taxi retardation) when the wheels are turning with excess kinetic energy. In embodiments, this may help to reduce brake wear by slowing the aircraft by regenerative braking, as well as traditional carbon brakes. This may, as a result, reduce the cost of brake maintenance, as well as reduce the environmental impact of the brake manufacture required.

[0049] In certain embodiments, it may be that the first driveable component comprises an open rotor propeller or a drive shaft of a high or low pressure turbine.

[0050] Preferably, the electric motor also functions as a generator such that, in a regenerating mode of operation, the electric motor is externally driven by the first driveable component to generate electricity and the hi-directional hydraulic pump is arranged such that, in the regenerating mode of operation, the hi-directional hydraulic pump is externally driven by the first driveable component to pump hydraulic fluid into the hydraulic fluid reservoir.

[0051] According to a second aspect of the invention there is also provided an aircraft landing gear drive system comprising the aircraft power system of the first aspect, wherein the first driveable component comprises one or more landing gear wheels of the landing gear drive system and wherein the electric motor and the hydraulic pump are both connected to the one or more landing gear wheels, such that the electric motor and the hydraulic pump are both able to drive the one or more landing gear wheels.

[0052] Preferably, the first driveable component comprises one or more landing gear wheels of the landing gear drive system, and wherein the electric motor and the hydraulic pump are both connected to the one or more landing gear wheels, such that the electric motor and the hydraulic pump arc both able to drive the one or more landing gear wheels.

[0053] According to a third aspect of the invention there is also provided an aircraft comprising the aircraft power system of the first aspect or the landing gear drive system of the second aspect.

[0054] It may be that the aircraft is a passenger aircraft.

[0055] According to a fourth aspect of the invention there is also provided a method of operating an aircraft, the aircraft being as per the third aspect (or comprising the aircraft power system of the first aspect or the landing gear drive system of the second aspect).

[0056] According to a fifth aspect of the invention there is also provided a method of operating an aircraft, the aircraft being as per the third aspect wherein, in an electrical generation mode of operation, the hydraulic pump is connected to the electric motor such that the hydraulic pump drives the electric motor as a generator, such that the electric motor is driven by the hydraulic pump.

[0057] Preferably, the aircraft power system further comprises an electric battery and wherein electrical energy generated in the electrical generation mode of operation is used to replenish the battery.

[0058] This allows the electrical energy generated to he stored for later use.

[0059] Preferably, the aircraft power system further comprises an electric distribution network and wherein electrical energy generated in the electrical generation mode of operation is distributed by the electric distribution network.

[0060] For example, the electrical energy could be distibuted around the aircraft to a variety of different electric devices.

[0061] Preferably, the hydraulic reservoir is connectable to the electric motor by a switch arrangement, and wherein the switch arrangement has selected the electrical generation mode.

[0062] The switch arrangement may have selected the electrical generation mode based on the status of the hydraulic reservoir (for example, if it is full). The switch arrangement may have selected the electrical generation mode based on the status of the aircraft (for example, if it has decelerated after landing and is being taxi driven by its engines). The switch arrangement may be able to select any number of other modes of operation as well.

[0063] It will of course be appreciated that features described in relation to one aspect of the present invention may be incorporated into other aspects of the present invention. For example, the method of the invention may incorporate any of the features described with reference to the apparatus of the invention and vice versa.

BRIEF DESCRIPTION OF THE DRAWINGS

[0064] Embodiments of the invention will now be described, by way of example only, with reference to the accompanying drawings, in which: [0065] Figure 1 shows a schematic view of an aircraft landing gear according to a first example embodiment of the invention; [0066] Figure 2 shows a schematic view of the aircraft landing gear of Figure 1 in an extended regenerative mode of operation; [0067] Figure 3 shows a schematic view of an aircraft having the landing gear of Figure 1; and [0068] Figure 4 is a diagram showing a method utilising the first embodiment of the invention.

DETAILED DESCRIPTION

[0069] Figure 1 shows a schematic view of an aircraft landing gear 700 having a hybrid power system 800 according to a first example embodiment of the invention.

[0070] The landing gear 700 comprises a landing gear leg 110. An axle 140 is mounted on and extends outwards from the landing gear leg 110. A first landing gear wheel 120 is mounted for rotation on a first end of the axle 140. A second landing gear wheel 130 is mounted for rotation on an opposite second end of the axle 140. When the landing gear 700 is supporting the weight of an aircraft on the ground, the rotation of the first and second landing gear wheels 120, 130 facilitates movement of the aircraft across the ground.

[0071] The landing gear 700 further comprises a power system 800.

[0072] The power system 800 comprises an electric motor 230. The electric motor 230 is connected to the landing gear wheels 120, 130 via a transfer box 220. A mechanical connection 231 connects the electric motor 230 to the transfer box 220. The transfer box 220 is connected to the axle 140 (and thereby to the wheels 120, 130) by a drive shaft 221.

Mechanical connection 231, transfer box 220, drive shaft 221, and axle 140 together enable the transfer of kinetic energy from the electric motor 230 to the wheels 120, 130.

[0073] Thus, the electric motor 230 can drive the wheels 120, 130. The electric motor 230 is therefore capable of converting electrical energy (for example, from an aircraft auxiliary power unit (APU), a hydrogen fuel cell, a battery, or the aircraft engines) into kinetic energy in the form of rotation of the wheels 120, 130. The electric motor 230, mechanical connection 231, transfer box 220, drive shaft 221, and axle 140 can all therefore be considered to form part of a landing gear drive system.

[0074] The power system 800 further comprises a hydraulic fluid reservoir 210, configured to contain a store of hydraulic fluid. The hydraulic fluid reservoir 210 is connected to a hi-directional hydraulic pump 212 by a hydraulic line 211, such that the hydraulic line 211 provides fluid communication between the hydraulic fluid reservoir 210 and the hydraulic pump 212. Thus, the hydraulic pump 212 is operable to pump hydraulic fluid to and from the hydraulic fluid reservoir 210.

[0075] A mechanical connection 213 connects the hydraulic pump 212 to the transfer box 220. Thus, the hydraulic pump 212 is also connectable to the wheels 120, 130. Mechanical connection 213, transfer box 220, drive shaft 221, and axle 140 together enable the transfer of kinetic energy from the hydraulic pump 212 to the wheels 120, 130.

[0076] The hydraulic fluid reservoir 210 comprises a high-pressure accumulator, such that hydraulic fluid stored in the hydraulic fluid reservoir 210 is held under pressure. Thus, the hydraulic fluid reservoir 210 acts as a store of hydraulic power (for example, for use in driving the first driveable component). The hydraulic fluid reservoir 210 also comprises a valve (not shown) which is operable to control the release of hydraulic fluid from the hydraulic fluid reservoir 210. When stored hydraulic fluid is released from the hydraulic fluid reservoir 210 (for example, by opening the valve), it is driven from the hydraulic fluid reservoir 210 by action of the pressure at which it is held. Thus, the hydraulic fluid reservoir 210 is operable to provide a pressurised stream of hydraulic fluid.

[0077] Pressurised hydraulic fluid released from the hydraulic fluid reservoir 210, can be used to drive one or more driveable components of the aircraft. In particular, the flow of pressurised hydraulic fluid from the hydraulic fluid reservoir 210 can be used to turn the hydraulic pump 212. The hi-directional hydraulic pump 212 thereby acts as a hydraulic motor, converting the hydraulic power from the flow of hydraulic fluid through the hydraulic pump 212 into kinetic energy, which is transferred by the connection 213, transfer box 220, drive shaft 221, and axle 140 to the wheels 120, 130. Thus, hi-directional hydraulic pump 212 and mechanical connection 213 can also he considered to form part of a landing gear drive system.

[0078] The transfer box 220 is operable to selectively connect and disconnect the mechanical connections 213, 231 and the drive shaft 221. In this case, the transfer box comprises a clutching mechanism (not shown) connected to each of the mechanical connections 213, 231 and the drive shaft 221. Thus, by operating the clutching mechanisms the mechanical connections 213, 231 and the drive shaft 221 can each be independently connected and disconnected. Thus, the transfer box 220 controls the transfer of energy between the electric motor 230, the hydraulic pump 212, and the wheels 120, 130. Thus, the electric motor and hydraulic pump can be said to be connected to each other in parallel such that both the electric motor and the hydraulic pump can each supply power to the wheels 120, 130 independently of the other. The power system 800 can therefore be said to be a hybrid power system (in that the system utilises both hydraulic and electrical power). The transfer box 220 is also operable to connect the electric motor 230 to the hi-directional pump, such that the electric motor 230 can be used to drive the hi-directional pump (for example, to pump hydraulic fluid to or from the hydraulic fluid reservoir 210).

[0079] The hydraulic pump 212 is also connectable to one or more further driveable components (not shown) of the aircraft. By pumping hydraulic fluid to and from the wheels 120, 130, it is possible to operate the wheels 120, 130. Similarly, the hydraulic pump 212 is operable to pump hydraulic fluid from the reservoir to drive the one or more further driveable components. Examples of such driveable components include landing gear extension/retraction mechanisms, landing gear bay door mechanisms, cargo door mechanisms, landing gear braking high lift device mechanisms (for example, flaps and/or slats), and flight control surfaces.

[0080] The hydraulic fluid reservoir 210 is shown as comprising a hydraulic fluid reservoir 801 separate from a hydraulic accumulator 803. Hydraulic sensor 240 is configured to monitor a state of charge of the hydraulic accumulator 803. The hydraulic sensor is configured to deterinine a state of charge of the hydraulic fluid reservoir 210 and to generate hydraulic sensor data indicating the determined state of charge of the hydraulic fluid reservoir 210. The power system 800 also comprises a battery sensor (not shown). The battery sensor is configured to determine a state of charge of the battery and to generate battery sensor data indicating the determined state of charge of the battery. The power system 800 further comprises a processor (not shown) and associated memory (not shown). The processor is configured to execute instructions stored in the associated memory to control operation of the power system 800 (including, for example, operation of the hydraulic sensor 240 and the battery sensor).

[0081] This landing gear 700 is similar to the system described in relation to Figures 6a to Od in EP 4005920A1. However, there arc some important differences.

[0082] In EP 4005920A1, the power system 800 is capable of operating in three distinct modes of operation: a driving mode, a regenerative mode, and a replenishing mode.

[0083] Operation of the hybrid power system 800, and in particular the mode of the hydraulic system, in the first driveable, regenerating, and replenishing modes of operation is controlled by use of a three-position switch 805. A first (middle) position of the switch 805 causes the hybrid power system 800 to operate in the first driveable mode of operation (i.e. where at least the hydraulic system is providing power e.g. to drive the wheels). A second (right hand side) position of the switch 805 (shown in Figure 1) causes the hybrid power system 800 to operate in the regenerating mode of operation (i.e. where the hydraulic system (reservoir) is replenished). A third (left hand side) position of the switch 805 causes the hybrid power system 800 to operate in the replenishing mode of operation (i.e. where the hydraulic system (reservoir) is replenished and where hydraulic energy is supplied to an additional hydraulic system of the aircraft).

[0084] In the driving mode of operation, the wheels 120, 130 are driven by one or both of the electric motor 230 and the hydraulic pump 212. Thus, in the driving mode, one or both the electric motor 230 and the hydraulic pump 212 are connected to the wheels 120, 130 by the transfer box 220. When the electric motor 230 is connected to the wheels 120, 130 and the hydraulic pump 212 is disconnected from the wheels 120, 130, the electric motor 230 delivers electric drive power 320 to the transfer box 220, which delivers the electric drive power to the drive shaft 221 to turn the wheels. When the electric motor 230 is disconnected from the wheels 120, 130 and the hydraulic pump 212 is connected to the wheels 120, 130, the hydraulic pump 212 delivers hydraulic drive power 310 to the transfer box 220, which delivers the hydraulic drive power to the drive shaft 221 to turn the wheels. When both the electric motor 230 and the hydraulic pump 212 arc connected to the wheels 120, 130, the electric motor delivers electric drive power 320 to the transfer box 220 and the bi-directional hydraulic pump 212 delivers hydraulic drive powcr 310 to the transfer box 220. The transfer box 220 delivers the combined drive power 330 to the drive shaft 221 to turn the wheels 120, 130, enabling the aircraft to taxi (represented by item 300).

[0085] In the regenerative mode of operation, the power system 800 operates to recover kinetic energy from the wheels 120, 130. The energy recovery can be performed when the electric motor is externally driven by the wheels 120, 130 (for example, when the aircraft is braking). In the regenerative mode, the wheels 120, 130 arc connected by the transfer box 220 to one or both of the electric motor 230 and the hydraulic pump 212. Which of the electric motor 230 and the hydraulic pump 212 the wheels 120, 130 are connected to is determined by the desired means for storing the recovered energy. The transfer box 220 can operate to connect the electric motor 230 to the wheels 120, 130 and disconnect the hydraulic pump 212 from the electric motor 230 and the wheels 120, 130. Alternatively, the transfer box 220 can operate to connect the hydraulic pump 212 to the wheels 120, 130 and disconnect the electric motor 230 from the hydraulic pump 212 and the wheels 120, 130. The transfer box 220 can also operate to connect the electric motor 230 and the hydraulic pump 212 to the wheels 120, 130.

[00861 When in the regenerative mode of operation, the kinetic energy 400 of the rotation of the wheels 120, 130 is transferred to the transfer box 220. The transfer box 220 then transfers the kinetic energy to one or both of the electric motor 230 and the hydraulic pump 212. When the wheels 120, 130 are connected to the electric motor 230, the kinetic energy of the wheels is transferred (represented by arrow 420) to the rotor of the electric motor 230. The electric motor 230 operates as a generator to convert the kinetic energy into electrical energy. This electrical energy can be stored (for example, in a battery) for later use, either by the electric motor 230 or by another aircraft subsystem. When the wheels 120, 130 are connected to the hydraulic pump 212, the kinetic energy of the wheels is transferred (represented by arrow 410) to the hydraulic pump 212. The hydraulic pump 212 is driven, using the kinetic energy, to pump hydraulic fluid into the hydraulic fluid reservoir 210. The kinetic energy is thereby stored as hydraulic energy in the hydraulic fluid reservoir 210 for later use, either to drive the wheels 120, 130 or another driveable component of the aircraft. It will be appreciated that the wheels 120, 130 can be connected to both the electric motor 230 and the hydraulic pump 212 at the same time, allowing recovered kinetic energy to be stored both electrically and hydraulically simultaneously. At other times, only one of the electric motor 230 and the hydraulic pump 212 is connected to the wheels 120, 130.

[0087] Thus, the power system 800 enables the recovery and storage (electrically and/or hydraulically) of kinetic energy from the wheels 120, 130. Further, performing regenerative braking in such a way reduces the braking forces that must be applied by the actual brakes, and therefore reduces the wear on those brakes.

[0088] In the replenishing mode of operation, the wheels 120, 130 are mechanically disconnected from the electric motor 230 and the hydraulic pump 212 by the transfer box 220, but the electric motor 230 and the hydraulic pump 212 are connected. The electric motor 230 converts electrical energy into kinetic energy (in the foim of rotation of the rotor) which is transferred (represented by arrow 510) to the transfer box 220. The transfer box 220 transfers (represented by arrow 520) that kinetic energy on to the hydraulic pump 212, where it is used to operate the hydraulic pump 212 to pump hydraulic fluid.

[00891 The electric motor 230 can drive the hydraulic pump 212 to pump hydraulic fluid into the hydraulic Iluid reservoir 210, thereby "replenishing" the store ofhydraulic fluid. The electric motor is also operable to drive the hydraulic pump 212 to pump hydraulic fluid out of the hydraulic fluid reservoir 210 (for example, in order to operate one or more further driveable components connected to the hydraulic pump 212). Thus, in the replenishing mode, the electric motor 230 can be said to be operable to control one or more further driveable components hydraulically connected to the hydraulic pump 212.

[0090] In the present embodiment, the power system 800 is capable of operating in four distinct modes of operation: the driving mode, regenerative mode, and replenishing mode of EP 4005920A1, as described above and in that publication, and also a fourth extended regenerative mode.

[0091] Figure 2 shows the power system 800 operating in the fourth extended regenerative mode of operation. In the fourth extended regenerative mode the wheels 120, 130 are mechanically disconnected (represented by cross 610) from the electric motor 230 and the hydraulic pump 212 by the transfer box 220, but the electric motor 230 and the hydraulic pump 212 are connected.

[00921 The hydraulic pump 212 is driving the electric motor 230 as a generator using high pressure hydraulic fluid from the accumulator 803. The hydraulic energy is transferred (represented by arrow 620) to the transfer box 220. The transfer box 220 transfers (represented by arrow 630) that energy on to the electric motor 230. Thus, in the extended regenerative mode, the electric energy stored in a battery, for example, can be replenished by the hydraulic fluid.

[0093] The second position of the switch 805 is used in the extended regenerative mode of operation.

[0094] The extended regenerative mode is typically used when the aircraft is being propelled on ground with the aircraft engines. This is because, after an aircraft deceleration phase, the hydraulic energy within the accumulator is high and so can be used to then drive the electrical generator/motor, via the hydraulic motor/pump. The electrical energy can be stored in a battery and/or can be used to provide additional power to an aircraft electrical distribution network.

[0095] Figure 3 shows an aircraft 1000 comprising the landing gear 700. The operation of the landing gear 700 and the hybrid power system 800 is now described.

[0096] Whilst the aircraft is on the ground before take-off, the power system operates in a first driveable mode of operation, driving the wheels 120, 130 using both the electric motor 230 and the hydraulic pump 212 to enable the aircraft to taxi. The aircraft may also drive the wheels 120, 130 using only one of the electric motor 230 and the hydraulic pump 212 (for example, once the aircraft is in motion across the ground and only drive power sufficient to maintain the motion is required).

[0097] Once the aircraft is in the air, the power system 800 may operate in the replenishing mode of operation, wherein the electric motor 230 drives the hydraulic pump 212 to pump hydraulic fluid into the hydraulic fluid reservoir 210. Optionally, the power system 800 may also operate the hydraulic pump 212 to provide hydraulic power to one or more other driveable components of the aircraft 1000 (for example, by controlling the electric motor 230 to drive the hydraulic pump 212 to provide hydraulic power to the other driveable components).

[0098] After the aircraft lands (when braking during taxiing), the power system may operate in the regenerating mode of operation, wherein the wheels 120, 130 are used to drive at least one (for example, both) of the electric motor 230 and the hydraulic pump 212 to recover and store kinetic energy (as electrical and hydraulic energy, respectively). This stored energy can then be used in subsequent operation (for example, in the driving mode or to operate the one or more further driveable components). By using the stored energy to drive the wheels 120, 130, the hybrid power system 800 can recover energy during braking and subsequently use the recovered energy to accelerate again. Such operation can enable more efficient taxiing (for example, where the aircraft is taxiing in a queue, so is repeatedly braking and accelerating).

[0099] When the aircraft is being propelled on the ground by its engines and/or when the hydraulic energy store is full or nearly full, hydraulic energy within the accumulator is high and so can be used to then drive the electrical generator/motor, via the hydraulic motor/pump. The electrical energy can be stored in a battery and/or can be used to provide additional power to an aircraft electrical distribution network. By using the stored hydraulic energy to drive the electric motor 230, the hybrid power system 800 can recover electrical energy and subsequently use the recovered energy. Such operation can enable more efficient power and energy use.

[0100] Figure 4 is a diagram 2000 illustrating schematically a method of operating the aircraft of Figure 3.

[0101] Thus, Figure 4 shows a first driveable mode of operation 2001 in which the landing gear wheels 2120 are driven by both the electric motor 2230 and the hydraulic pump 2212.

[0102] Figure 4 also shows a replenishing mode of operation 2002 in which the electric motor 2230 drives the hydraulic pump 2212 to pump hydraulic fluid into the hydraulic reservoir 2210.

[0103] Figure 4 also shows a second driveable mode of operation 2003, in which the electric motor 2230 drives the hydraulic pump 2212 so that it provides hydraulic power to one or more other driveable components 2004.

[0104] Figure 4 also shows a regenerating mode of operation 2004 in which the landing gear wheels 2120 are externally driven to drive the electric motor 2230 and/or the hydraulic pump 2212.

[0105] Figure 4 also shows an extended regenerative mode 2005 in which the hydraulic pump 2212 drives the electric motor 2230 as a generator using hydraulic fluid from the reservoir 2210.

[0106] Whilst the present invention has been described and illustrated with reference to particular embodiments, it will be appreciated by those of ordinary skill in the art that the invention lends itself to many different variations not specifically illustrated herein. By way of example only, certain possible variations will now be described.

[0107] The driving modes may alternatively, or additionally be used For reverse movement of the aircraft (known as pushback), as well as forward taxiing. One or more of the wheels may he driven in a reverse direction. There may he provided additional sensors and/or systems to ensure the aircraft does not tip back during pushback.

[0108] The driving modes may additionally be used for steering the aircraft by driving the wheels at different speeds (differential speed control), which may include driving wheels in opposite directions).

[0109] During, and/or before take-off of the aircraft, for example when the aircraft is being driven by its engines and/or not being driven by the wheels, wheel driving may be inhibited or blocked by die arrangement.

[0110] The accumulator may be discharged in situations where peak acceleration or deceleration performance is required. This provides optimised performance of the hydraulic pump and/or motor.

[0111] Off-board (as well as on-board) energy storage/supply mat he provided. This off-board energy storage/supply may comprise any suitable electrical storage or supply, including a supply connected to a mains supply, an electrical battery 262, and/or an electrical capacitance device 263.

[0112] The electrical motor 230 is connected to the off-board energy storage/supply by a connection. This connection may be a wireless (inductive) connection, a wired connection or both.

[0113] The above embodiments are to be understood as illustrative examples of the invention. Equivalents and modifications not described above may also be employed without departing from the scope of the invention, which is defined in the accompanying claims.

[0114] It is to be understood that any feature described in relation to any one embodiment may be used alone, or in combination with other features described, and may also be used in combination with one or more features of any other of the embodiments, or any combination of any other of the embodiments.

[0115] It will also be appreciated by the reader that integers or features of the invention that are described as preferable, advantageous, convenient or the like arc optional and do not limit the scope of the independent claims. Moreover, it is to be understood that such optional integers or features, whilst of possible benefit in seine embodiments of the invention, may not be desirable, and may therefore be absent, in other embodiments.

[(11 16] It should be noted that throughout this specification, "or" should be interpreted as "and/or".

[0117] Although the invention has been described above mainly in the context of a fixed-wing aircraft application, it may also be advantageously applied to various other applications, including but not limited to applications on vehicles such as helicopters, drones, trains, automobiles and spacecraft.

Claims

CLAIMS: 1. An aircraft power system comprising: i) a hydraulic reservoir, ii) a hi-directional hydraulic pump for pumping hydraulic fluid to and from the reservoir, iii) an electric motor, and iv) a first driveable component of the aircraft, wherein the hi-directional hydraulic pump is arranged such that, in a regenerating mode of operation, the hi-directional hydraulic pump is externally driven by the first driveable component to pump hydraulic fluid into the hydraulic fluid reservoir, and wherein the hydraulic pump is connectable to the electric motor such that the hydraulic pump is arranged to drive the electric motor as a generator, such that, in an electrical generation mode of operation, the electric motor is driven by the hydraulic pump.
2. An aircraft power system as claimed in claim 1, wherein the aircraft power system further comprises an electric battery and wherein electrical energy generated in the electrical generation mode of operation is used to replenish the battery.
3. An aircraft power system as claimed in claim 1 or claim 2, wherein the aircraft power system further comprises an electric distribution network and wherein electrical energy generated in the electrical generation mode of operation is distributed by the electric distribution network.
4. An aircraft power system as claimed in any preceding claim, wherein the hydraulic reservoir is connectable to the electric motor by a switch arrangement, such that the switch arrangement can select the electrical generation mode.
5. An aircraft power system as claimed in any preceding claim, wherein the electric motor is connectable to the first driveable component of the aircraft.
6. An aircraft power system as claimed in claim 5, wherein the electric motor is connectable to the first driveable component of the aircraft such that the electric motor is arranged to drive the first driveable component oldie aircraft, and wherein the hydraulic pump is connectable to the first driveable component of the aircraft such that the hydraulic pump is arranged to drive the first driveable component, such that, in a first driveable mode of operation, the first driveable component is driven by both the electric motor and the hydraulic pump.
7. The aircraft power system of claim 6, wherein the electric motor and hydraulic pump are connected to each other in parallel such that, in the first driveable mode of operation, both are arranged to supply power to the driveable component independently of the other.
8. The aircraft power system of any preceding claim, wherein the electric motor is connectable to the hydraulic pump such that, during a hydraulic replenishing mode of operation, the electric motor is arranged to drive the hydraulic pump to pump hydraulic fluid into the reservoir.
9. The aircraft power system of any preceding claim, wherein, in a regenerating mode of operation, the electric motor is externally driven by a moveable component of the aircraft to generate electricity.
10. The aircraft power system of any preceding claim, wherein the hydraulic pump is also connectable to other driveable components of the aircraft.
11. The aircraft power system of any preceding claim, wherein the electric motor is connectable to the hydraulic pump such that the electric motor is arranged to drive the hydraulic pump to pump hydraulic fluid out of the reservoir.
12. The aircraft power system of any preceding claim, wherein the first driveable component comprises one or more landing gear wheels.
13. The aircraft power system of any preceding claim, wherein: the electric motor also functions as a generator such that, in a regenerating mode of operation, the electric motor is externally driven by the first driveable component to generate electricity.
14. An aircraft landing gear drive system comprising the aircraft power system of any preceding claim, wherein the first driveable component comprises one or more landing gear wheels of the landing gear drive system and wherein the electric motor and the hydraulic pump are both connected to the one or more landing gear wheels, such that the electric motor and the hydraulic pump are both able to drive the one or more landing gear wheels.
15. An aircraft landing gear system as claimed in claim 14, when comprising the aircraft power system of claim 5, wherein the first driveable component comprises one or more landing gear wheels of the landing gear drive system, and wherein the electric motor and the hydraulic pump are both connected to the one or more landing gear wheels, such that the electric motor and the hydraulic pump are both able to drive the one or more landing gear wheels.
16. An aircraft comprising the aircraft power system of any of claims 1 to 13 or the landing gear drive system of claim 14 or 15.
17. A method of operating an aircraft, the aircraft being as per claim 16, wherein, in an electrical generation mode of operation, the hydraulic pump is connected to the electric motor such that the hydraulic pump drives the electric motor as a generator, such that the electric motor is driven by the hydraulic pump.
18. The method of claim 17, wherein the aircraft power system further comprises an electric battery and wherein electrical energy generated in the electrical generation mode of operation is used to replenish the battery.
19. The method of claim 17 or claim 18, wherein the aircraft power system further comprises an electric distribution network and wherein electrical energy generated in the electrical generation mode of operation is distributed by the electric distribution network.
20. The method of any of claims 17 to 19, wherein the hydraulic reservoir is connectable to the electric motor by a switch arran2ement, and wherein the switch arrangement has selected the electrical generation mode.