HYDRAULIC ACTUATION SYSTEM FOR AN AIRCRAFT

Abstract

An aircraft hydraulic actuation system for retracting an aircraft landing gear. The actuation system includes a supply line arranged to carry hydraulic fluid pressurized by a pump, a return line arranged to return hydraulic fluid to a reservoir, and a hydraulic actuator 128. In a first mode of operation, a first chamber 130 of the actuator 128 is supplied with pressurized hydraulic fluid from the supply line such that a piston 134 is moved in a first direction so as to move a load such as a landing gear. In a second mode of operation, the first chamber 130 is taken out of fluid communication with the supply line and a second chamber 132 is in fluid communication with the return line, such that the piston 134 is able to be moved under the influence of the load, for example when the landing gear extends under gravity.

Claims

1. An aircraft hydraulic actuation system comprising: a supply line arranged to carry hydraulic fluid pressurized by a pump, a return line arranged to return the hydraulic fluid to a reservoir, and a hydraulic actuator comprising a first chamber and a second chamber separated by a piston, wherein the first chamber is fluidly connectable to the supply line, and the second chamber is fluidly connectable to the return line; wherein the aircraft hydraulic actuation system has a first mode of operation in which the first chamber is supplied with the hydraulic fluid from the supply line such that the piston is moved in a first direction to move a load; and wherein the aircraft hydraulic actuation system has a second mode of operation in which the first chamber is taken out of fluid communication with the supply line and the second chamber is in fluid communication with the return line, such that the piston is configured to be moved in a second direction, opposite to the first direction, under the influence of the load.

2. The aircraft hydraulic actuation system according to claim 1, further comprising a hydraulic fluid flow path between the first chamber and the second chamber such that, in the second mode of operation, the second chamber is filled with the hydraulic fluid from the first chamber.

3. The aircraft hydraulic system according to claim 2, wherein the hydraulic fluid flow path between the first chamber and the second chamber comprises a flow restrictor valve configured to restrict a rate of fluid flow to a predetermined level.

4. The aircraft hydraulic actuation system according to claim 2, wherein the aircraft hydraulic actuation system comprises a first hydraulic line feeding into the first chamber, a second hydraulic line feeding into the second chamber, and a third hydraulic line connecting the first hydraulic line and the second hydraulic line, wherein the third hydraulic line is configured to provide the hydraulic fluid flow path between the first chamber and the second chamber.

5. The aircraft hydraulic actuation system according to claim 4, wherein the third hydraulic line is integral with the hydraulic actuator.

6. The aircraft hydraulic actuation system according to claim 2, wherein the hydraulic fluid flow path between the first chamber and the second chamber is within the piston.

7. The aircraft hydraulic actuation system according to claim 1, wherein the hydraulic actuation system is arranged such that the supply line is arranged to supply the hydraulic fluid at a pressure of 100 Bar or greater.

8. The aircraft hydraulic actuation system according to claim 1, wherein the hydraulic actuation system is arranged such that the hydraulic fluid in the return line has a pressure of 10 Bar or less.

9. The aircraft hydraulic actuation system according to claim 1, wherein the pump is driven by an aircraft engine.

10. The aircraft hydraulic actuation system according to claim 1, wherein the pump is arranged to draw the hydraulic fluid from the reservoir.

11. The aircraft comprising an aircraft hydraulic actuation system according to claim 1.

12. The aircraft according to claim 11, wherein the load is an aircraft component which, during use, is arranged to move under an external force in a direction that causes the piston to move in the second direction.

13. The aircraft according to claim 12, wherein the aircraft component is an aircraft landing gear, the hydraulic actuator is a landing gear retraction actuator, wherein the aircraft hydraulic actuation system is configured to be in the first mode for retraction of the landing gear and configured to be in the second mode for extension of the landing gear.

14. The aircraft according to claim 12, wherein the aircraft component is an aircraft door.

15. A method of moving an aircraft component in an aircraft according to claim 13, the method comprising: configuring the hydraulic actuation system into the first mode of operation, supplying the hydraulic fluid from the supply line into the first chamber to move the component from a first position to a second position under the action of the hydraulic actuator; and configuring the hydraulic actuation system into the second mode of operation and allowing the component to move under an under the external force from the second position to a first position.

16. An aircraft landing gear extension and retraction system comprising: a supply line arranged to carry hydraulic fluid pressurized by a pump, a return line arranged to return hydraulic fluid to a reservoir, and a hydraulic landing gear retraction actuator for connection to a landing gear, the actuator comprising a first chamber and a second chamber separated by a piston, wherein the piston is moved during extension and retraction of the landing gear; the landing gear extension and retraction system being operable in a first mode to retract the landing gear, wherein in the first mode the first chamber is supplied with the hydraulic fluid from the supply line such that the piston is urged in a direction that causes the aircraft landing gear to be urged towards the retracted position; the landing gear extension and retraction system being operable in a second mode to allow the landing gear to extend, wherein in the second mode the first chamber is taken out of fluid communication with the supply line and the second chamber is in fluid communication with the return line such that the piston is able to move in a direction that allows the landing gear to extend under gravity.

17. A method of extending a landing gear of an aircraft, the aircraft comprising an aircraft landing gear extension and retraction system according to claim 16, the method comprising: configuring the landing gear extension and retraction system into the second mode of operation; unlocking the landing gear from the retracted position; allowing the landing gear to extend under gravity; and locking the landing gear in its extended position.

18. The method according to claim 17, further comprising a step of refilling the first chamber with hydraulic fluid from the second chamber.

19. An aircraft hydraulic actuation system comprising: a hydraulic actuator for retraction of a landing gear, the hydraulic actuator comprising a first chamber and a second chamber separated by a piston; and a fluid flow path between the first chamber and the second chamber that is open for a flow of hydraulic fluid from the first chamber to the second chamber when the piston is moved in a first direction.

Description

DESCRIPTION OF THE DRAWINGS

[0032] Embodiments of the present invention will now be described by way of example only with reference to the accompanying schematic drawings of which:

[0033] FIG. 1 shows a side view of an aircraft comprising a landing gear;

[0034] FIG. 2 shows a side view of a landing gear and example arrangement for retracting and extending the landing gear;

[0035] FIGS. 3a and 3b show a first prior art hydraulic actuator and the connections thereto, FIG. 3a shows the actuator retracting and FIG. 3b shows the actuator extending;

[0036] FIGS. 4a and 4b show a second prior art hydraulic actuator and the connections thereto, FIG. 4a shows the actuator retracting and FIG. 4b shows the actuator extending;

[0037] FIG. 5 shows a plan view of an aircraft according to a first embodiment of the invention;

[0038] FIG. 6 shows a side view of an aircraft according to a first embodiment of the invention;

[0039] FIG. 7 shows a landing gear and a landing gear extension and retraction system according to a first embodiment of the invention;

[0040] FIGS. 8a and 8b show a hydraulic actuator and connections thereto according to a first embodiment of the invention, FIG. 8a shows the actuator retracting and FIG. 8b shows the actuator extending;

[0041] FIG. 9 shows a hydraulic actuator according to a second embodiment of the invention.

DETAILED DESCRIPTION

[0042] FIGS. 5 and 6 show an aircraft 102 according to a first embodiment of the invention. The aircraft 102 comprises a hydraulic actuation system comprising an engine driven pump 104 arranged to pressurize hydraulic fluid to a supply pressure of 200 Bar. The engine driven pump 104 is supplied with hydraulic fluid from a reservoir 106. A back-up electric pump 108 is provided in parallel to the engine driven pump 104 to pressurize the hydraulic fluid in the hydraulic system in the event that the engine driven pump 104 cannot operate. In alternative embodiments, only an electric pump or pumps are provided.

[0043] A hydraulic fluid supply line 110 extends from the pumps 104, 108 and carries the pressurized hydraulic fluid to various actuators in the aircraft 102, for example, actuators associated with flight control surfaces, brakes and landing gear retraction. The hydraulic actuation system further comprises an accumulator 112 that feeds into the supply line 110 downstream of the hydraulic pumps 104, 108. A return line 111 returns hydraulic fluid to the reservoir. The pressure of hydraulic fluid in the return line 111 is approximately 5 to 10 Bar.

[0044] The aircraft 102 further comprises a landing gear 114. As shown in FIG. 7, the landing gear 114 comprises a landing gear leg 116 comprising a shock absorber 118 and a torque link 120, a set of wheels 122 are mounted to the bottom of the leg 116. In alternative embodiments, the landing gear may comprise a bogie pivotally mounted to the landing gear leg, the bogie comprising a plurality of sets of wheels.

[0045] A landing gear extension and retraction system 124 is coupled to the landing gear 114. The landing gear extension and retraction system 124 comprises various components that also form part of the aircraft's hydraulic actuation system, including a hydraulic actuator 128 associated with extension and retraction of the landing gear 114. The landing gear extension and retraction system 124 also comprises a landing gear up-lock mechanism and a landing gear down-lock mechanism (not shown), which each comprise their own hydraulic actuators.

[0046] FIGS. 8a and 8b show the hydraulic actuator 128 and its associated hydraulic connections in further detail. The hydraulic actuator 128 comprises a housing 142 containing a first chamber 130 and a second chamber 132 separated by a piston 134. The piston 134 is attached to a piston rod 136, the free end of which is pivotally mounted to the landing gear leg 116. The end of the housing 142 that is distal from the free end of the piston rod 136 is pivotally mounted to a structural part of the aircraft 102.

[0047] A first hydraulic line 138 feeds into the first chamber 130 and a second hydraulic line 140 feeds into the second chamber 132. The first hydraulic line 138 is in fluid communication with the second hydraulic line 140 via a third hydraulic line 144 which creates a flow path between the first chamber 130 and the second chamber 132. A flow restrictor 146 in the third hydraulic line 144 restricts the rate at which hydraulic fluid can flow through the third hydraulic line 144.

[0048] The hydraulic actuation system has a first mode of operation (FIG. 8a) for retraction of the landing gear 114. In the first mode of operation, a selector valve places the first hydraulic line 138, and therefore the first chamber 130, into fluid communication with the supply line 110. Pressurized hydraulic fluid can thereby be supplied into the first chamber 130. The second hydraulic line 140, and therefore the second chamber 132, is in fluid communication with the return line 111. Hydraulic fluid can thereby exit the second chamber 132 and may return to the reservoir 106.

[0049] In use, in the first mode, the difference between the hydraulic pressure in the supply line 110 and the hydraulic pressure in the return line 111 causes the piston 134 to be urged in a direction which retracts the piston rod 136 into the housing 142. The overall length of the actuator 128 is thereby reduced and the landing gear 114 is retracted. The presence of the hydraulic line 144 causes some hydraulic fluid to pass directly from the supply to the return. The flow restrictor 146 is selected such that the rate of fluid flow has no impact on the effective operation of the actuator 128 when retracting the landing gear 114.

[0050] The hydraulic actuation system has a second mode of operation (FIG. 8b) for extension of the landing gear 114. In the second mode of operation, the selector valve places the first hydraulic line 138, and therefore the first chamber 130, into fluid communication with the return line 111. Hydraulic fluid can thereby exit the first chamber 130 and may return to the reservoir 106. The second hydraulic line 140, and therefore the second chamber 132, is also in fluid communication with the return line 111.

[0051] When the hydraulic actuation system is in the second mode of operation, and the system in a substantially static state, the hydraulic pressure in the first chamber 130 and the second chamber 132 is substantially the same. The piston 134 is thereby not urged in any particular direction by hydraulic pressure.

[0052] When the landing gear is allowed to extend by disengagement of the up-lock and opening of the landing gear doors, the weight of the landing gear is such that it drops under gravity. The landing gear is also arranged such that, in flight, drag caused by the flow of air over the landing gear also urges the landing gear towards an extended position. As the landing gear extends, the piston rod 136 is pulled out from the housing 142 and the piston 134 is moved in a direction that reduces the volume of the first chamber 130 and increases the volume of the second chamber 132.

[0053] Downstream of the actuator 128, the return line 111 comprises a check valve (not shown) to ensure one way flow towards the reservoir 106, therefore very little, if any, hydraulic fluid is able flow back down the return line 111 so as to fill the second chamber 132. Instead, the second chamber 132 is filled, via the hydraulic line 144, by hydraulic fluid from the first chamber 130 and by hydraulic fluid from other hydraulic elements of the landing gear extension and retraction system 124.

[0054] The flow restrictor 146 is selected to allow a relatively low rate of fluid flow so that the fluid flow through the third hydraulic line 144 does not to have a negative impact on the ability of the hydraulic system to pressurize the first chamber 130 on retraction of the landing gear 114. As a consequence, during extension of the landing gear 114, the volume of fluid flowing through the flow restrictor 146 is insufficient to fill the increasing size of the second chamber 132. Therefore, as the landing gear 114 extends, the pressure in the second chamber 132 drops towards a vacuum.

[0055] When the landing gear 114 has reached full extension, the down-lock is engaged to lock the landing gear 114 in its fully extended position and the pressure in the second chamber slowly increases as it is filled via the hydraulic line 144. As the second chamber 132 is not in fluid communication with the supply line 110, the actuator 128 does not transfer high loads into the aircraft structure, landing gear 114 and associated attachment points at full extension.

[0056] FIG. 9 shows a hydraulic actuator of a hydraulic actuation system according to a second embodiment of the invention. In the second embodiment of the invention, the first hydraulic line 238 and the second hydraulic line 240, that feed straight into the first and second chambers 230, 232 of the actuator 228, are not directly linked via a third hydraulic line. Instead, the piston 234 comprises a through-hole 244 that provides a flow path between the first chamber 230 and the second chamber 232. A flow restrictor 246 is located in the though-hole 244 so as to restrict the rate at which hydraulic fluid can flow between the first chamber 230 and the second chamber 232. In alternative embodiments, the through-hole is sized (e.g. during manufacture) so as to restrict the rate of flow through the through-hole 244 to the required level.

[0057] 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.

[0058] In an alternative embodiment of the invention, the flow restrictor in the hydraulic line between the first hydraulic line and the second hydraulic line is replaced by a valve arranged to selectively close the flow path between the first chamber and the second chamber. The valve is closed when the hydraulic actuation system is in the first mode, and opened when the hydraulic actuation system is in the second mode. In embodiments, operation of the valve is computer controlled.

[0059] In an alternative embodiment of the invention, the aircraft hydraulic actuation system is arranged such that in the first mode of operation the piston is urged in a direction that extends the length of the actuator. This may, for example, be achieved by swapping around the various supply line and return line connections that feed into the actuator. Corresponding changes may need to be made to the linkages in the landing gear extension and retraction system such that extension of the actuator results in retraction of the landing gear.

[0060] In an alternative embodiment of the invention, the actuator comprises a housing comprising a first and a second internal hydraulic line formed in the housing. Each internal hydraulic line extends from an input/output port to a chamber of the actuator. The input/output ports are arranged for connection to an external hydraulic line. A third internal hydraulic line provides a hydraulic fluid flow path between the first internal hydraulic line and the second internal hydraulic line.

[0061] It is envisaged that the present invention may have non-aircraft applications. References in the description and claims to aircraft hydraulic actuation systems could be replaced by references to hydraulic actuation systems in the general sense.

[0062] Where in the foregoing description, integers or elements are mentioned which have known, obvious or foreseeable equivalents, then such equivalents are herein incorporated as if individually set forth. Reference should be made to the claims for determining the true scope of the present invention, which should be construed so as to encompass any such equivalents. 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 are 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 some embodiments of the invention, may not be desirable, and may therefore be absent, in other embodiments.