GB2614877A - Aircraft landing gear shock absorber strut - Google Patents

Abstract

An aircraft landing gear shock absorber strut 24 comprising: first and second bearings 54, 56 and a mechanical support linkage SL extending between the bearings. The mechanical support linkage comprises: a shock absorber 60 comprising outer and inner cylinders 26, 28, the inner cylinder moves along a longitudinal axis LA of the bore between a positions which compress and extend the shock absorber and is biased by a spring force to assume the extended condition; a movable member 62 configured to move in a first direction which reduces the distance between the bearings when the shock absorbing strut is supporting the weight of the aircraft; and in a second direction which increases the distance between the bearings when the shock absorbing strut is not supporting the weight of the aircraft; and a locking device 64 operable between a locking condition in which the locking device inhibits movement of the movable member; and a passive condition in which the locking device permits movement of the movable member.

Description

AIRCRAFT LANDING GEAR SHOCK ABSORBER STRUT

Background to the Invention

It is common for an aircraft landing gear assembly to include a main hydraulic shock absorber strut having an upper end arranged to be pivotally coupled to the underside of the aircraft and a lower end coupled to a wheel and brake assembly.

Such shock absorber struts can comprise an outer cylinder and an inner cylinder arranged to telescope relative to the outer cylinder. The shock absorber strut can be compressed and extended as the inner cylinder moves relative to the outer cylinder. The two portions are coupled together to define a chamber containing oil and in some cases a gas. As the shock absorber is compressed, oil within the chamber is forced through damping orifices and, where gas is also provided, the gas is compressed, in order to dampen landing loads. The compressed gas serves as a spring to lengthen the shock absorber as applied external load decreases. Recoil damping orifices can be provided to restrict the flow of oil to the annulus as the shock absorber extends.

An aircraft landing gear bay is a space within an aircraft that is configured to accommodate a stowed landing gear. The landing gear may have been designed specifically for the aircraft.

It can be desirable for the main strut of a landing gear to be longer than a standard landing gear main strut for a particular aircraft in order to raise the aircraft higher to improve ground clearance. However, unless the landing gear bay is redesigned to accommodate the increase in length, the landing gear must still be capable of retracting into the existing space within the bay. Hence the landing gear is required to shorten on retraction.

Various means are known by which to shorten an aircraft landing gear shock absorber strut. This can involve a dedicated shortening mechanism which must be attached to a structural part of the landing gear bay. In another example, as illustrated by EP 1 921 341, fluid is pumped into the shock absorber strut to shorten it.

The present inventor has devised a new type of aircraft landing gear shock absorber strut which can be shortened for stowage and can be simpler and/or lighter than existing solutions.

Summary of Invention

According to a first aspect of the invention, there is provided an aircraft landing gear shock absorber strut comprising: a first bearing arranged to movably couple the shock absorber strut to an aircraft; a second bearing arranged to couple the shock absorber strut to a ground contacting assembly; and a mechanical support linkage extending between the first and second bearings, wherein the mechanical support linkage comprises: a shock absorber, the shock absorber comprising: an outer cylinder having a bore defining an opening; an inner cylinder having a first end region movably coupled within the bore and a second end region which projects out of the opening, the inner cylinder being arranged to move along a longitudinal axis of the bore between a first condition in which the shock absorber strut is compressed and a second condition in which the shock absorber strut is extended, the inner cylinder being biased by a spring force to assume the second condition; a movable member configured to move: in a first direction which reduces the distance between the first and second bearings when the shock absorbing strut is supporting the weight of the aircraft; and in a second direction which increases the distance between the first and second bearings when the shock absorbing strut is not supporting the weight of the aircraft; and a locking device operable between: a locking condition in which the locking device inhibit movement of the movable member; and a passive condition in which the locking device permits movement of the movable member.

Thus, the shock absorber strut according to the first aspect includes a means to shorten the overall length of the shock absorber strut without relying on hydraulic or electrical power from the aircraft. Instead, the energy to shorten the shock absorber strut can come from the weight of the aircraft. When deployed for taxiing to the runway and take-off, the landing gear can compress under the weight of the aircraft via the locking device being in the passive condition. This allows the shock absorber to be in a shortened condition before the landing gear is retracted and stowed. After being compressed, the locking device can be operated to the locking condition to maintain the shortened length. When the landing gear is to be deployed, the locking device can be operated to the passive condition which allows extension of the shock absorber strut due to the weight of the unsprung mass. In contrast, known shortening assemblies add complexity (weight, cost, reduced reliability) to the gear and also require additional energy from the aircraft to shorten the gear during or prior to retraction.

The movable member can comprise a rod arranged to move coaxially with the longitudinal axis of the bore.

The rod can comprise a piston having a greater diameter than the rod and wherein the mechanical support linkage can further comprise a hydraulic chamber sized such that the piston moves in sealing engagement with an inner sidewall of the hydraulic chamber as the rod moves in the first direction and the second direction such that the piston separates the hydraulic chamber into a first sub-chamber and a second sub-chamber, the inner wall of the hydraulic chamber can be provided with first and second end stops to limit movement of the piston in the first and second directions respectively, the shock absorber strut can further comprise a fluid return passage arranged to enable fluid communication between the first and second sub-chambers such that as the piston moves in the first direction, fluid from first sub-chamber can move to the second sub-chamber via the fluid return passage and as the piston moves in the second direction, fluid from second sub-chamber can move to the first sub-chamber via the fluid return passage, wherein the locking device can comprise a valve located within the fluid return passage and can be configured in the locking condition to inhibit fluid flow through the fluid return passage and can be configured in the passive condition to permit fluid flow through the fluid return passage.

The rod can have a first end and a second end, and the piston is located between the first and second ends, the second end extending through a port in an axial face of the hydraulic chamber and can be arranged to move in sealing engagement with an inner sidewall surface of the port as the piston moves between the extremities defined by the first and second end stops.

The aircraft landing gear shock absorber strut according to the first aspect can further comprise a reservoir external to the hydraulic chamber and can be arranged in fluid communication with the fluid return passage between the valve and the first sub-chamber.

The hydraulic chamber can be located at an axial end of the outer cylinder of the shock absorber opposite to the opening of the bore and the inner cylinder can define a blind bore in fluid communication with the bore of the outer cylinder and can have an a third end stop projecting radially into the blind bore to define a passage which is narrower than the blind bore, wherein the second end of the rod can be provided with an out stop piston including one or more damping orifices providing fluid communication between the blind bore and the bore, the out stop piston can have a diameter which is greater than the diameter of the passage such that the third end stop inhibits the out stop piston moving out of the blind bore.

The mechanical support linkage can further comprise a structural cylinder defining a second bore within which the outer cylinder is slidably housed with a free end of outer cylinder projecting from the structural cylinder and the inner cylinder can extend into the second bore towards the hydraulic chamber, wherein the rod can be coupled to the free end of the inner cylinder.

The aircraft landing gear shock absorber strut according to the first aspect can further comprise a controller operable to: change the locking device from the passive condition to the locking condition when the shock absorber strut is supporting the weight of the aircraft on the ground; and change the locking device from the locking condition to the passive condition following deployment of the shock absorber strut from the aircraft and before landing.

According to a second aspect of the invention, there is provided an aircraft landing gear assembly comprising: the aircraft landing gear shock absorber strut according to the first aspect; and a wheel or other ground contacting assembly coupled to the shock absorber strut.

The landing gear assembly can be a main landing gear assembly, in contrast to a nose gear, and the wheel assembly can be mounted on a multi axle bogie beam.

According to a third aspect of the invention, there is provided a method of controlling a length of aircraft landing gear shock absorber strut according to the first aspect or an aircraft landing gear assembly according to the second aspect comprising: in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground, changing the locking device from the passive condition to the locking condition; and in response to determining the shock absorber strut has been deployed from the aircraft for landing, and before landing, changing the locking device from the locking condition to the passive condition.

The method can include: prior to the step of in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground, changing the locking device from the passive condition to the locking condition, a step of changing the locking device from the locking condition to the passive condition to allow the weight of the aircraft to shorten the shock absorbing strut.

The method can include: following the step of in response to determining the shock absorber strut has been deployed from the aircraft for landing, and before landing, changing the locking device from the locking condition to the passive condition, and before landing, a step of changing the locking device from the passive condition to the locking condition to lock the shock absorbing strut in the extended condition for landing.

The step of in response to determining the shock absorber strut has been deployed from the aircraft for landing, and before landing, can comprise waiting a period of between one and five seconds following a signal indicating completion of deployment.

The step of in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground can comprise waiting for a signal that aircraft speed is less than 10 kilometres per hour.

Brief Description of the Drawings

Embodiments of the invention will now be described, strictly by way of example only, with reference to the accompanying drawings, of which: Figure 1 is a diagram of an aircraft; Figures 2a to 2e are diagrams of an aircraft landing gear assembly; Figure 3 is a schematic diagram of an aircraft landing gear shock absorber strut according to a first embodiment of the invention in a normal condition; Figure 4 is a schematic diagram of the aircraft landing gear shock absorber strut of Figure 3 in a shortened condition; Figure 5 is a schematic diagram of an aircraft landing gear shock absorber strut according to a second embodiment of the invention in a normal condition; Figure 6 is a schematic diagram of the aircraft landing gear shock absorber strut of Figure 5 in a shortened condition; Figure 7 is a schematic diagram of an aircraft landing gear shock absorber strut according to a third embodiment of the invention in a normal condition; Figure 8 is a schematic diagram of the aircraft landing gear shock absorber strut of Figure 7 in a shortened condition; Figure 9 is a schematic diagram of an aircraft landing gear shock absorber strut according to a fourth embodiment of the invention in a normal condition; Figure 10 is a schematic diagram of the aircraft landing gear shock absorber strut of Figure 9 in a shortened condition; Figure 11 is a flow chart illustrating a method of controlling the length of an aircraft landing gear shock absorber strut according to an embodiment of the invention; and Figure 12 is an example spring curve illustrating load versus stroke for shock absorber struts according to embodiments of the invention.

Description of Embodiments

Figure 1 is a diagram of an aircraft 10. The aircraft 10 includes assemblies such as a nose landing gear 12, main landing gear 14 and engines 16. The landing gear 12, 14 each includes a shock absorber strut for damping landing loads and supporting the weight of the aircraft 10 when it is on the ground. The term aircraft as used herein can include aeroplanes, helicopters and the like having a mass more than 450Kg.

Referring now to Figures 2a to 2e, an aircraft assembly, namely an aircraft landing gear assembly, is shown generally at 14. Figures 2a to 2e are an example of an aircraft landing gear assembly which can include a shock absorber strut according to an embodiment of the invention. It will however be appreciated that shock absorber struts according to embodiments of the invention can be used in a range of types of aircraft landing gear.

The landing gear assembly 14 includes a foldable stay 18, a lock link 20 and a down lock spring assembly 22 mounted to the stay 18 and arranged to urge the lock link 20 to assume a locked state. The landing gear assembly also includes a main shock absorber strut 24, comprising an outer cylinder 26 and an inner cylinder 28, as well as a wheel and brake assembly 30.

The aircraft landing gear assembly is movable between a deployed condition, for take-off and landing, and a stowed condition for flight. An actuator (not shown) is provided for moving the landing gear between the deployed condition and the stowed condition. This actuator is known in the art as a retraction actuator, and more than one can be provided. A retraction actuator can have one end coupled to the airframe and another end coupled to the outer cylinder such that extension and retraction of the actuator results in movement of the outer cylinder between deployed and stowed conditions.

The stay 18 serves to support the orientation of the outer cylinder 26 when the landing gear is in the deployed condition. The stay 18 generally includes a two bar linkage that can be unfolded to assume a generally straight or aligned, over centre condition in which the stay 18 is locked to inhibit movement of the outer cylinder, as shown in Figures 2c and 2e. When the stay is broken, it no longer prevents pivotal movement of the outer cylinder 26 and the outer cylinder 26 can be moved by the retraction actuator towards the stowed condition, as shown in Figure 2a. During flight the stay 18 is arranged in the folded condition, while during take-off and landing the stay 18 is arranged in the generally straight or aligned condition. Some main landing gear assemblies include a pair of stays coupled to a common shock absorber strut.

The stay 18 has an elongate upper stay arm 18a having a lower end defining a pair of lugs pivotally coupled via a pivot pin 32 to a pair of lugs defined at an upper end of an elongate lower stay arm 18b. The stay arms 18a and 18b can therefore pivotally move relative to one another about the pivot pin 32. The upper end of the upper stay arm 18a defines a pair of lugs that are pivotally coupled to a lug of a connector 34 which in turn is pivotally coupled to the airframe 11. The lower end of the lower stay arm 18b defines a pair of lugs pivotally coupled to a lug of a connector 36 which in turn is pivotally coupled to the outer cylinder 26.

The lock link 20 has an elongate upper link arm 20a having a lower end pivotally coupled to an upper end of an elongate lower link arm 20b via a pivot pin 38. The link arms 20a, 20b can therefore pivotally move relative to one another about the pivot pin 38. An upper end of the upper link arm 20a defines a pair of lugs that are pivotally coupled to a lug of a connector 40 which in turn is pivotally coupled to the outer cylinder 26. A lower end of the lower link arm 20b defines a lug that is pivotally coupled to lugs of the stay arms 18a, 18b via the pivot pin 32. Lugs of the upper stay arm 18a are in this example disposed between the lugs of the lower stay arm 18b and the lugs of the lower link arm 20b.

When the lock link 20 is in the locked condition, as illustrated in Figures 2d and 2e, the upper and lower link arms 20a, 20b are generally longitudinally aligned or coaxial, and can be 'over-centre', such that the lock link 20 is arranged to oppose a force attempting to fold the stay 18, so as to move the landing gear assembly from the deployed condition towards the stowed condition. The lock link 20 must be broken to enable the stay 18 to be folded, thereby permitting the outer cylinder 26 to be moved by the retraction actuator towards the stowed condition.

One or more down lock springs 22 are generally provided to assist in moving the landing gear assembly to the deployed condition and locking it in that state by making the lock link. Down lock springs 22 also inhibit the lock link accidentally being unlocked. Down lock springs 22 are generally metal coil springs, which can be coupled between the lock link and another part of the landing gear assembly, such as an arm of the stay assembly, as shown in Figures 2b and 2e.

The spring assembly 22 is arranged to bias the lock link 20 towards the locked condition by way of spring tension. A distal end of the spring 22a is coupled to the lower stay arm 18b via a lower engagement formation 22b which in turn is coupled to an anchor point defined by the lower connector 22c.

The coil spring of the spring assembly 22 is at its shortest when the landing gear assembly is in the deployed condition, as shown in Figure 2e, and at its longest when the landing gear assembly approaches the stowed condition, as shown in Figure 2b. As the landing gear assembly is retracted towards the stowed condition, the spring of each spring assembly extends, resulting in increased spring load and torsional stress.

Referring to Figure 2e, a lock stay actuator 42 is coupled between the upper stay arm 18a and lower link arm 20b and arranged to pivotally move the link arms 20a, b so as to 'lock' and 'unlock' the lock link 20, as illustrated in Figure 2c. The actuator 42 can break the lock link 20 against the down lock spring bias, allowing the landing gear assembly to be folded and stowed as described previously.

Referring to Figures 3 and 4, an aircraft landing gear shock absorber strut according to a first embodiment of the invention is shown generally at 24.

The shock absorber strut 24 has a first bearing 54 arranged to movably couple the shock absorber strut 24 to an aircraft 10, a second bearing 56 arranged to couple the shock absorber strut to a ground contacting assembly 58 and a mechanical support linkage SL extending between the first and second bearings 54, 58.

The mechanical support linkage SL is arranged to absorb landing loads and support the weight of the aircraft 10 when on the ground and includes a shock absorber 60, a movable member 62 and a locking device 64. In embodiments of the invention the shock absorber 60 is of a functionally conventional design. The present inventor has realised that by including the movable member 62 and the locking device 64 as part of the mechanical support linkage SL, the weight of the aircraft 10 can be utilised to shorten the shock absorber strut 24 and the weight of the un-sprung mass of the shock absorber strut 24 (and in some embodiments the internal spring force of the shock absorber 28) can be utilised to lengthen the shock absorber strut 24 to its normal length.

In the first embodiment, the shock absorber 60 has an outer cylinder 26 having a bore B defining an opening 0. The outer cylinder 26 is elongate. An upper end of the outer cylinder 26 is structurally coupled to the first or primary mounting bearing 54, via which the shock absorber strut 24 is arranged to be movably mounted to an aircraft 10 to move between a deployed condition for take-off and landing and a stowed condition for flight. When in the stowed condition, the shock absorber strut 24 can be received within a landing gear bay BA within the aircraft 10.

The shock absorber 60 also has an inner cylinder 28 having a first end region, which defines a radially enlarged piston head 66, movably coupled within the bore B and a second end region which projects out of the opening 0. The inner cylinder 28 is arranged to move along a longitudinal axis LA of the bore B between a first condition in which the shock absorber 60 is compressed (shown in Fig. 4) and a second condition in which the shock absorber strut 60 is extended, as shown in Fig. 3.

The inner cylinder 28 is biased by a spring force to assume the second condition. The bore B and blind bore BB together define a chamber which contains hydraulic fluid such as oil and a gas such as nitrogen which is compressed as the shock absorber strut 24 is compressed and provides the spring force biasing the shock absorber strut 24 to extend. In other embodiments a conventional separator piston can be provided within the blind bore BB to separate the gas from the hydraulic fluid.

An annular gland member 68 is provided at the opening 0 to close the shock absorber 60 and can include bearings and dynamic seals which act on the other surface of the inner cylinder 28 to support it as it moves and confine oil within the shock absorber 60.

The free end of the inner cylinder 28 is provided with the second bearing 56 for coupling to a wheel and brake assembly 58, bogie beam or the like.

In this embodiment the movable member 62 is in the form of an extended orifice support tube or rod having a radially enlarged out stop piston 70 at its lower end (when the gear is deployed) configured to slide within the blind bore BB and having one or more damping orifices 72 extending axially through the out stop piston 70. An end stop formation 74 projects radially into the blind bore BB to define a passage which is narrower than the blind bore BB. The out stop piston 70 has a diameter which is greater than the diameter of the passage such that the end stop formation 74 inhibits the out stop piston 70 moving out of the blind bore BB. As such, the out stop piston 70 can engage the end stop formation 74 to inhibit the shock absorber 60 from extending.

The locking device is in the form of a hydraulic chamber 64 which forms a structurally integral part of the mechanical support linkage SL. In this embodiment the hydraulic chamber is formed in an upper region of the outer cylinder 26 but could instead be a separate chamber that is mechanically coupled to the top of the outer cylinder 26.

The rod has a first end and a second end and the piston 76 is located between the first and second ends. The piston 76 has a greater diameter than the rod 62 and the hydraulic chamber 64 is sized such that the piston 76 moves in sealing engagement with an inner sidewall 78 of the hydraulic chamber 64 as the rod 62 moves in first and second directions, towards and away from the aircraft 10. The piston 76 separates the hydraulic chamber 64 into a first sub-chamber 64a and a second sub-chamber 64b, each of which can be filled with hydraulic fluid. The inner wall 78 is provided with first and second end stops 80, 82 to limit movement of the piston 76 in the first and second directions respectively. The middle region of the rod 62 extends through a port P in an axial face of the hydraulic chamber 64 and is arranged to move in sealing engagement with an inner sidewall surface of the port P as the piston moves between the extremities defined by the first and second end stops. As such, hydraulic fluid within second sub-chamber 64b is inhibited from entering the bore B via the port P. The shock absorber strut 24 also includes a fluid return passage 84 arranged to provide fluid communication between the first and second sub-chambers 64a, 64b such that, as the piston 76 moves in the first direction, fluid from first sub-chamber 64a moves to the second sub-chamber 64b via the fluid return passage 84 and, as the piston 76 moves in the second direction, fluid from second sub-chamber 64b moves to the first sub-chamber 64a via the fluid return passage 84. The locking device also includes a valve 86 located within the fluid return passage 84 and configured in a locking condition to inhibit fluid flow through the fluid return passage 84 and configured in a passive condition to permit fluid flow through the fluid return passage 84.

Thus, the rod 62 is configured to move in the first direction which reduces the distance between the first and second bearings 54, 58 (shown in Figure 4) when the shock absorbing strut 24 is supporting the weight of the aircraft and move in the second direction which increases the distance between the first and second bearings 54, 58 (shown in Figure 3) when the shock absorbing strut 24 is not supporting the weight of the aircraft. The locking device 64 is operable between the locking condition, when the valve 86 is closed, in which the locking device 64 inhibit movement of the rod 62 due to hydraulic fluid either side of the piston 76, and a passive condition, when the valve 86 is open, in which the locking device 64 permits movement of the rod 62 due to the piston 76 being able to force hydraulic fluid past the valve 86. As such, the valve 86 can be operated during an aircraft weight on wheels condition to lock the shock absorbing strut 24 in a shortened condition and can be opened when the gear is deployed to allow the shock absorbing strut 24 to lengthen for landing.

In the illustrated embodiment the second end of the rod 62 extends through a second port P2 in an upper axial face of the hydraulic chamber 64 and is arranged to move in sealing engagement with an inner sidewall surface of the second port P2 as the piston moves between the extremities defined by the first and second end stops 80, 82. As such, hydraulic fluid within first sub-chamber 64a is inhibited from leaving the first sub-chamber 64a via the second port P2. This arrangement has an advantage that the volume of hydraulic chamber 64 occupied by the rod 62 remains the same as the rod moves, meaning that the volume of hydraulic fluid within the chamber remains constant.

A controller C can be provided which is operable to change the locking device from the passive condition to the locking condition when the shock absorber strut is supporting the weight of the aircraft on the ground and change the locking device from the locking condition to the passive condition following deployment of the shock absorber strut from the aircraft and before landing. Alternatively, the valve can be controlled manually from the aircraft.

An example of the sequence for the landing gear system is as follows: Sequence Shock Absorber Condition Aircraft in flight, gear retracted Shortened Aircraft in flight, gear extends Shortened Aircraft in flight, shortening valve opens Extends due to the weight of the unsprung mass Aircraft in flight, shortening valve closes Extended Landing Extended Aircraft on ground, taxi-in Extended Aircraft on ground, shortening valve opens Shortens due to the weight of the aircraft Aircraft on ground, shortening valve closes Shortened Aircraft on ground, taxi-out Shortened Take-off Shortened Aircraft in flight, gear retracts Shortened In a second embodiment, shown in Figures 5 and 6, the shock absorber strut 124 is the same as the shock absorber strut 24 except that the rod 162 terminates at the piston 176. This means that the upper axial face of the hydraulic chamber does not require the second port P and, in use, the rod 162 does not extend from the top of the shock absorber strut 124. However, as the volume of hydraulic chamber 164 occupied by the rod 162 increases as the rod 162 moves into the chamber 164, that the volume of hydraulic fluid within the chamber 164 reduces. As such, a reservoir 190 is provided external to the hydraulic chamber 164 and arranged in fluid communication with the fluid return passage between the valve 186 and the first sub-chamber 164a.

The shock absorber strut 124 is shown shortened in Figure 6 and at a normal length in Figure 5.

Figures 7 and 8 show a shock absorber strut 224 according to a third embodiment which is similar to the shock absorber strut 24 of the first embodiment, but in which the shock absorber is inverted as a "capsule type" shock absorber. In this embodiment, the mechanical support linkage further comprises a structural cylinder 228 or "main fitting" defining a second bore B2 within which the outer cylinder 226 is slidably housed with a free end of outer cylinder projects from the structural cylinder. The inner cylinder 292 extends into the second bore towards the hydraulic chamber. The rod 262 is coupled to the free end of the inner cylinder.

Figures 9 and 10 show a shock absorber strut 324 according to a fourth embodiment which is similar to the shock absorber strut 224 of the third embodiment except that, as with the second embodiment, the rod 362 terminates at the piston 376 and therefore a reservoir 390 is provided.

Figure 11 illustrates a method 400 of controlling a length of an aircraft landing gear shock absorber strut or an aircraft landing gear assembly according to embodiment of the invention, the method comprising: in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground 410, changing the locking device from the passive condition to the locking condition 420. In preferred embodiments where the locking device has locked the strut in the extended state prior to landing, the method can include a preceding step of changing the locking device from the locking condition to the passive condition to allow the weight of the aircraft to shorten the strut. The method 400 further includes a step of, in response to determining the shock absorber strut has been deployed from the aircraft for landing 430, and before landing, changing the locking device from the locking condition to the passive condition 440. The method can include a subsequent step of, prior to landing, changing the clocking device from the passive condition to the locking condition to lock the strut in the extended condition for landing; this advantageously provides the full shock absorber stroke for landing.

The step of in response to determining the shock absorber strut has been deployed from the aircraft for landing, and before landing, can comprise waiting a period of between one and five seconds following a signal indicating completion of deployment.

The step of in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground 410 can comprise waiting for a signal that aircraft speed is less than 10 kilometres per hour.

Figure 12 illustrates the shock absorber spring curve for the first and second embodiments, (outstop style shock absorber) and for the third and fourth embodiments (capsule type shock absorber). The spring curve demonstrates the potential stroke available for a shock absorber against the load exerted on the shock absorber. In both the outstop type and capsule type, the full range of stroke is available on landing 520, as the landing gear is extended under its own weight. During take-off, the shock absorber is locked in the shortened state, meaning that less stroke is available compared to a conventional non-shortened shock absorber strut. However, a reduced stroke can be acceptable during on-ground operation such as during taxi-in, taxi-out and take off, particularly for main landing gear. For the capsule type, the lower end of the spring curve is available on ground 540.

However, for the outstop type the top end of the spring curve is available 560, which is the portion of the spring curve that would be available on-ground in a conventional landing gear, meaning that both the landing performance and the on-ground performance match that of with a conventional landing gear.

Thus, embodiments of the invention utilise the weight of the aircraft for shortening the gear prior to take-off, rather than using an external power source, such as the aircraft hydraulic or electrical systems. Similarly the weight of the unsprung mass of the gear (and the shock absorber pressure for the outstop shortening concept) is used to lengthen the gear following extension, rather than using hydraulic or electrical power from the aircraft.

Components of the aircraft landing gear and/or shock absorber strut struts according to embodiments of the invention can be implemented from conventional aerospace materials, such as titanium, aluminium and/or steel for structural members, polymer or metal bearings etc. It should be noted that the above-mentioned embodiments illustrate rather than limit the invention, and that those skilled in the art will be capable of designing many alternative embodiments without departing from the scope of the invention as defined by the appended claims. In the claims, any reference signs placed in parenthesis shall not be construed as limiting the claims. The word "comprising" does not exclude the presence of elements or steps other than those listed in any claim or the specification as a whole. The singular reference of an element does not exclude the plural reference of such elements and vice-versa. Parts of the invention can be implemented by means of hardware comprising several distinct elements. In a device claim enumerating several parts, several of these parts can be embodied by one and the same item of hardware. The mere fact that certain measures are recited in mutually different dependent claims does not indicate that a combination of these measures cannot be used to advantage.

Claims (14)

1. CLAIMS1. An aircraft landing gear shock absorber strut comprising: a first bearing arranged to movably couple the shock absorber strut to an aircraft; a second bearing arranged to couple the shock absorber strut to a ground contacting assembly; and a mechanical support linkage extending between the first and second bearings, wherein the mechanical support linkage comprises: a shock absorber, the shock absorber comprising: an outer cylinder having a bore defining an opening; an inner cylinder having a first end region movably coupled within the bore and a second end region which projects out of the opening, the inner cylinder being arranged to move along a longitudinal axis of the bore between a first condition in which the shock absorber strut is compressed and a second condition in which the shock absorber strut is extended, the inner cylinder being biased by a spring force to assume the second condition; a movable member configured to move: in a first direction which reduces the distance between the first and second bearings when the shock absorbing strut is supporting the weight of the aircraft; and in a second direction which increases the distance between the first and second bearings when the shock absorbing strut is not supporting the weight of the aircraft; and a locking device operable between: a locking condition in which the locking device inhibit movement of the movable member; and a passive condition in which the locking device permits movement of the movable member.
2. 2. The aircraft landing gear shock absorber strut according to claim 1, wherein the movable member comprises a rod arranged to move coaxially with the longitudinal axis of the bore.
3. 3. The aircraft landing gear shock absorber strut according to claim 2, wherein the rod comprises a piston having a greater diameter than the rod and wherein the mechanical support linkage further comprises a hydraulic chamber sized such that the piston moves in sealing engagement with an inner sidewall of the hydraulic chamber as the rod moves in the first direction and the second direction such that the piston separates the hydraulic chamber into a first sub-chamber and a second sub-chamber, the inner wall of the hydraulic chamber being provided with first and second end stops to limit movement of the piston in the first and second directions respectively, the shock absorber strut further comprising a fluid return passage arranged to enable fluid communication between the first and second sub-chambers such that as the piston moves in the first direction, fluid from first sub-chamber moves to the second sub-chamber via the fluid return passage and as the piston moves in the second direction, fluid from second sub-chamber moves to the first sub-chamber via the fluid return passage, wherein the locking device comprises a valve located within the fluid return passage and configured in the locking condition to inhibit fluid flow through the fluid return passage and configured in the passive condition to permit fluid flow through the fluid return passage.
4. 4. The aircraft landing gear shock absorber strut according to claim 3, wherein the rod has a first end and a second end and the piston is located between the first and second ends, the second end extending through a port in an axial face of the hydraulic chamber and being arranged to move in sealing engagement with an inner sidewall surface of the port as the piston moves between the extremities defined by the first and second end stops.
5. 5. The aircraft landing gear shock absorber strut according to claim 3, further comprising a reservoir external to the hydraulic chamber and arranged in fluid communication with the fluid return passage between the valve and the first sub-chamber.
6. 6. The aircraft landing gear shock absorber strut according to any of claims 3 to 5, wherein the hydraulic chamber is located at an axial end of the outer cylinder of the shock absorber opposite to the opening of the bore and the inner cylinder defines a blind bore in fluid communication with the bore of the outer cylinder and having an a third end stop projecting radially into the blind bore to define a passage which is narrower than the blind bore, wherein the second end of the rod is provided with an out stop piston including one or more damping orifices providing fluid communication between the blind bore and the bore, the out stop piston having a diameter which is greater than the diameter of the passage such that the third end stop inhibits the out stop piston moving out of the blind bore.
7. 7. The aircraft landing gear shock absorber strut according to any of claims 3 to 5, wherein the mechanical support linkage further comprises a structural cylinder defining a second bore within which the outer cylinder is slidably housed with a free end of outer cylinder projecting from the structural cylinder and the inner cylinder extending into the second bore towards the hydraulic chamber, wherein the rod is coupled to the free end of the inner cylinder.
8. 8. The aircraft landing gear shock absorber strut according to any preceding claim, further comprising a controller operable to: change the locking device from the passive condition to the locking condition when the shock absorber strut is supporting the weight of the aircraft on the ground; and change the locking device from the locking condition to the passive condition following deployment of the shock absorber strut from the aircraft and before landing.
9. 9. An aircraft landing gear assembly comprising: the aircraft landing gear shock absorber strut according to any preceding claim; and a ground contacting assembly coupled to the shock absorber strut.
10. 10. A method of controlling a length of an aircraft landing gear shock absorber strut according to any of claim 1 to 8 or an aircraft landing gear assembly according to claim 9, the method comprising: in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground, changing the locking device from the passive condition to the locking condition; and in response to determining the shock absorber strut has been deployed from the aircraft for landing, and before landing, changing the locking device from the locking condition to the passive condition.
11. 11. The method of claim 10, further comprising: prior to the step of in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground, changing the locking device from the passive condition to the locking condition, changing the locking device from the locking condition to the passive condition to allow the weight of the aircraft to shorten the shock absorbing strut.
12. 12. The method of claim 10 or claim 11, further comprising: following the step of in response to determining the shock absorber strut has been deployed from the aircraft for landing, and before landing, changing the locking device from the locking condition to the passive condition, and before landing, changing the locking device from the passive condition to the locking condition to lock the shock absorbing strut in the extended condition for landing.
13. 13. The method according to any of claims 10 to 12, wherein the step of in response to determining the shock absorber strut has been deployed from the aircraft for landing, and before landing, comprises waiting a period of between one and five seconds following a signal indicating completion of deployment.
14. 14. The method according to any of claims 10 to 13, wherein the step of in response to determining that the shock absorber strut is supporting the weight of the aircraft on the ground comprises waiting for a signal that aircraft speed is less than 10 kilometres per hour.