CN110271662B

CN110271662B - Method for driving an aircraft landing gear between a deployed position and a retracted position

Info

Publication number: CN110271662B
Application number: CN201910203266.9A
Authority: CN
Inventors: B·尤扎特; M·奎纳切杜; B·杜巴赫; P·亨利昂; S·杜波依斯
Original assignee: Safran Landing Systems SAS
Current assignee: Safran Landing Systems SAS
Priority date: 2018-03-16
Filing date: 2019-03-18
Publication date: 2023-06-30
Anticipated expiration: 2039-03-18
Also published as: CA3036788C; FR3078942A1; CA3036788A1; FR3078943A1; FR3078943B1; CN110271662A; FR3078942B1

Abstract

The invention relates to a method of driving an aircraft landing gear between a deployed position and a retracted position, the landing gear comprising a leg (2) hinged to the structure of the aircraft so as to be movable between the two positions, which is stabilized in the deployed position by means of a strut member (10) comprising two branches (11, 12) hinged to each other, one branch being coupled to the leg and the other branch being coupled to the structure of the aircraft, the branches being brought into an aligned position when the leg is in the deployed position. According to the invention, a rotary actuator (20) is arranged on the aircraft, the actuator comprising a first crank (22) and a second crank (24), the first crank and the second crank being mounted to rotate freely about a common axis of rotation, but having relative angular positions which can be controlled, and the first crank and the second crank being arranged to stabilize the branches of the strut member in a substantially aligned position.

Description

Method for driving an aircraft landing gear between a deployed position and a retracted position

Technical Field

The present invention relates to a method of driving an aircraft landing gear between a deployed position and a retracted position.

Background

Aircraft landing gears are known that include legs that are hinged to the structure of the aircraft so as to be movable between a deployed position and a retracted position. The leg is stabilized in the deployed position by means of a strut member, which generally comprises two branches hinged together, one branch being coupled to the leg and the other branch being coupled to the structure of the aircraft, the two branches being held in a substantially aligned position by a stabilizer member forming a lock which can be unlocked so as to enable the leg to be lifted from the deployed position to the retracted position. For this purpose, such landing gears typically include an unlocking actuator for unlocking the stabilizer member and for disrupting the alignment of the strut member, and a driving actuator for lifting the leg toward the retracted position.

However, it is also possible to use a single actuator performing both functions. For example, document FR 2 946 319 proposes the use of a driving actuator of the rotary electromechanical type, coupled to one arm of the member for the purpose of driving the landing gear and unlocking the stabilizer member, for stabilizing the stay in the aligned position.

Maintaining the leg in the retracted position typically requires the use of an upturned lock box (upsilon box) secured to the structure of the aircraft and including a hook that engages a roller secured to the leg when the leg reaches the retracted position. However, realignment schemes are also known in which, when the landing gear is in the retracted position, the two branches of the brace member or the two branches of the stabilizer member are in an aligned position, making it possible to omit the stow position lock box.

Disclosure of Invention

Object of the Invention

The present invention seeks to propose a method of driving an aircraft landing gear between a deployed position and a retracted position using only a single actuator.

Summary of The Invention

In order to achieve this object, a method of driving an aircraft landing gear between a deployed position and a retracted position is provided, the landing gear having a leg that is hinged to a structure of the aircraft so as to be movable between the deployed position and the retracted position, which is stabilized in the deployed position by means of a strut member comprising two branches hinged to each other, one branch being coupled to the leg and the other branch being coupled to the structure of the aircraft. According to the invention, a rotary actuator is arranged on an aircraft, the actuator having a first crank and a second crank, the first crank and the second crank being mounted to rotate freely about a common axis of rotation but assuming controllable relative angular positions, the first crank being connected to a strut member by a first link and the second crank being connected to a leg by a second link, such that each crank assumes:

-a first relative angular position, wherein the first crank and the first link are in a first alignment when the leg is in the deployed position, thereby stabilizing the branches of the brace member in a substantially aligned position; and

-a second relative angular position, wherein the second crank and the second link are in a second alignment while the leg is in the retracted position, thereby stabilizing the leg in the retracted position.

This arrangement relates the angular position of the leg relative to the structure of the aircraft to the relative angular position of the two cranks in a one-to-one correspondence. In the present invention, it is ensured that the deployed position of the leg corresponds to a first alignment of the first crank and the first link, and the retracted position corresponds to a second alignment of the second crank and the second link. The first alignment serves to stabilize the brace member in the aligned position and thus the legs in the deployed position, while the second alignment stabilizes the legs in the retracted position, thereby eliminating any need for an upper position lockbox.

In this context, the crank and associated connecting rod are referred to as being "aligned" in the sense that when the two elements are positioned such that their axes of articulation with each other, the axis of articulation between the connecting rod and the landing gear, and the axis of rotation of the crank are contained substantially in a single plane. However, and as is well known, alignment may be stabilized by slightly exceeding the fully aligned position, thereby bringing the two elements into a slightly off-aligned position (i.e., an "over-centered" position) defined by the abutments between the elements. These rest positions are maintained by the residual torque of the actuator when the actuator is not energized (electromagnetic torque due to permanent magnets for electromagnetic actuators, or fluid trapped in the actuator chamber for hydraulic actuators). This arrangement thus makes it possible to avoid any need for an auxiliary locking member, since its function is replaced by each of the aligned crank and connecting rod pairs.

Drawings

The invention may be better understood in light of the following description of specific embodiments thereof with reference to the accompanying drawings in which:

figures 1 and 2 are landing gear according to the invention equipped with an actuator having two cranks, the landing gear being shown in its deployed and its retracted positions, respectively;

figures 3 to 6 are side views of the landing gear of figure 1 in the deployed position, when initially lifted, when being lifted and in the retracted position, respectively;

FIG. 7 is a perspective view showing details of the alignment between the crank and the associated connecting rod;

figure 8 is a cross-sectional view of the resilient abutment mounted on the roof to co-operate with the strut when the landing gear is in the retracted position; and

figure 9 is a side view of the landing gear in the retracted position, in which the brace bears against the resilient abutment.

Detailed Description

With reference to fig. 1 and 2, the invention is applied in this example to a landing gear 1, the landing gear 1 comprising a leg 2 carrying a wheel 3 at its bottom end and hinged to the structure of the aircraft about a hinge axis X1, the hinge axis X1 being substantially horizontal in operation. The legs are movable between a deployed position shown in fig. 1 and a retracted position visible in fig. 2. The leg 2 is stabilized in the deployed position by means of a brace member 10, the brace member 10 comprising two branches hinged together, in particular a panel 11 hinged to the structure of the aircraft about a hinge axis X2 and an arm 12 hinged to the leg 2 and the panel 11 about respective hinge axes X3 and X4. In the deployed position, the arm 12 and the panel 11 are in a substantially aligned position.

According to the invention, a rotary actuator 20 is provided, the rotary actuator 20 being free to rotate on the structure of the aircraft about a rotation axis X5 parallel to the hinge axes X1 to X4. The rotary actuator comprises a housing 21, the housing 21 having an appendage forming a first crank 22 and comprising a shaft 23, the shaft 23 being mounted to rotate about a rotation axis X5 and carrying a second crank 24. The relative angular position between the two

cranks

22 and 24 can be modified by powering the actuator, causing the shaft 23 to rotate relative to the housing 21, and when it is unpowered, it can be fixed and maintained by the residual torque of the actuator 20. In this example, the first crank 22 is coupled to the panel 11 of the brace member 10 by means of a first link 25 (in this example two links extend on either side of the end of the panel 11), the first link 25 being hinged to the first crank 22 about a hinge axis X6 and to the panel 11 about a hinge axis X7, while the second crank 24 is coupled to the leg 2, and more particularly to a horn 26 on the leg, by means of a second link 27, the second link 27 being hinged to the second crank 24 about a hinge axis X8 and to the horn 26 about a hinge axis X9. In this example, all axes X1 to X9 are mutually parallel.

This configuration correlates the relative angular positions of the

cranks

22 and 24 with the angular position of the leg 2 relative to the structure of the aircraft one to one. In the position shown in fig. 3, in which the leg 2 is in the deployed position, the relative angular positions of the cranks are such that the first crank 22 and the first link 25 are in a substantially aligned position, which is referred to as a first alignment. More precisely, the first alignment is a position obtained by: causing the first crank 22 and the first link 25 to slightly exceed their geometrically aligned positions (as defined by the complete alignment of the axes X5, X6 and X7 in a single plane), causing them to reach the respective abutments. As can be seen in fig. 7, the abutment is in the form of a finger 28, the finger 28 extending from the end of the first crank, bearing against obstacles 29 fixed to the first link 25, which extend facing the finger 28 and form a stop defining the first alignment. Similarly, the second link 27 has a finger 36, the finger 36 bearing against an obstacle 37 (in particular a pin 37) fixed to the second crank 24. In fig. 7, the second crank 25 and the second connecting rod 27 are in a second aligned position, wherein the finger 36 is supported against the obstacle 37.

Returning to fig. 3, it can be seen that since the actuator 20 is not powered, the first crank 22 cannot rotate, so its residual torque prevents the housing 21 and the shaft 23 from rotating relative to each other, thereby constraining them to rotate together. However, the shaft 23 is blocked by the second crank 24, the second crank 24 itself being blocked from rotation by its coupling with the leg 2 via the second connecting rod 27, the second connecting rod 27 being out of alignment with the second crank 24.

To raise the leg 2 towards the retracted position, the actuator 20 is energized, causing the shaft 23 to rotate, thereby modifying the relative angular position between the

cranks

22 and 24. As shown in fig. 4, a first effect of this rotation is to disrupt the alignment between the first crank 22 and the first link 24, thereby disrupting the alignment between the panel 11 and the arm 12 of the brace member 10. The landing gear 2 is then no longer stable in its deployed position, and the legs may be lifted towards the retracted position. Continuing to power the actuator 20, the leg 2 continues to lift and the leg 2 reaches an intermediate position as shown in fig. 5 in which the first link 25 pulls the panel 11 while the second link 27 pushes against the leg 2, thereby having the effect of lifting the leg towards a retracted position as shown in fig. 6 in which the cranks are in a relative angular position such that the second crank 24 and the second link 27 reach their positions of substantial alignment, referred to as a second alignment. In the same way as the first alignment, the second alignment is in particular a position obtained by: causing the second crank 24 and the second connecting rod 27 to slightly exceed the geometric alignment (as defined by the complete alignment of the axes X5, X8 and X9 in the same plane), causing them to engage the respective abutments. This alignment blocks the leg 2 in the retracted position, making this position stable without the use of an upper position lock box.

According to a particular aspect of the invention, the landing gear has doors 30 coupled thereto, these doors 30 being hinged to the structure of the aircraft about axis X10 and serving to close the compartment housing the landing gear in the retracted position, this operation being achieved by means of a link 31, the link 31 being directly coupled to a lug 32 projecting from the panel 11 of the strut member 10. To facilitate an understanding of the invention, the figures do not show fairings that are coupled to the legs and cooperate with the two doors 30 to close the cabin when the landing gear is in the retracted position and to remain open when the landing gear is in the extended position. In the two positions of the landing gear shown in figures 1, 2, 3 and 6, the doors 30 can be seen in a closed position, but as can be seen in figures 4 and 5, these doors open when the landing gear is driven. Maintaining the doors in the closed position requires that a pre-stress be placed on the doors 30 to ensure that they do not open under aerodynamic forces in flight. For this purpose, the length of the coupling 31 is determined so that, in both the retracted and the deployed position of the landing gear, the door 30 bears against an abutment 35 (visible in fig. 1) fixed to the structure of the aircraft, slightly before the alignment of the connecting rod and crank assembly, and in particular slightly before the aforementioned passage of the latter through the geometrically aligned position. The abutment of the door 30 with the abutment 35 then entails pulling the coupling in order to pre-stress the door 30, which then acts as a return spring, ensuring that the crank and connecting rod assembly is in its internal abutment position and thus its alignment.

To increase the pre-stress induced by closing the door, or to replace it in the case of a door not coupled to the landing gear, it is possible to use another external pre-stress source, such as an elastic abutment 50, as shown in fig. 8. The resilient abutment has a base 51 for fastening to the roof of the cabin. The base 51 carries a hollow cylinder or body 52, the hollow cylinder or body 52 being provided with an end opening carrying a guide 53. The piston 54 is slidably mounted in the guide 53 to protrude from the main body 52. As shown in fig. 8, the piston 54 is urged to the protruding position by a belleville washer 55, the belleville washer 55 being located inside the hollow body 52 and exerting a pre-stress on the piston 54. In order to cause the piston 54 to retract, a force is required to be exerted thereon that is not less than the prestress force exerted by the belleville washers 55.

As can be seen in fig. 9, the resilient abutment 50 is positioned on the roof of the landing gear bay such that when the landing gear reaches the retracted position, a portion of the strut member, and in particular the strut panel 11, abuts the piston 54, after which the second crank 24 and the second link 27 are geometrically aligned before the "second aligned" position is achieved. Thus, to go through and beyond the geometric alignment to reach the second alignment, the strut panel 11 needs to push the piston 54 back against the pre-stress applied by the belleville washers 55. In this way, the prestress from the belleville washer is transferred to the stay panel 11 in the same way as the prestress is transferred from the door to the stay panel via the nose angle 32.

Of course, the prestressing from the door and/or from the resilient abutments may be transferred to a location other than the strut panel on the landing gear, for example directly to the strut leg of the landing gear. The prestressing force may also be applied by one or more internal springs, which ensure that the second crank 24 and the second connecting rod 27 are in their second alignment.

The invention is not limited to the above description but instead covers any modifications brought within the scope defined by the claims.

In particular, although in this example the hinge axes are all parallel to each other, the invention naturally applies to linkages having axes that are not parallel, so long as each crank and linkage assembly of the actuator is aligned when the leg is in one or the other of its deployed and retracted positions.

Claims

1. A method of driving an aircraft landing gear between a deployed position and a retracted position, the landing gear having a leg (2) hinged to a structure of the aircraft so as to be movable between the deployed position and the retracted position, the leg being stabilised in the deployed position by means of a brace member (10) comprising a first limb (11) and a second limb (12) hinged to each other, one of the first and second limbs being coupled to the leg and the other of the first and second limbs being coupled to the structure of the aircraft, and the first and second limbs being brought to an aligned position when the leg is in the deployed position, the method being characterised in that a rotary actuator (20) is arranged on the aircraft, the actuator having a first crank (22) and a second crank (24), the first and second cranks being mounted to rotate freely about a common rotational axis but assuming a controlled relative angular position, the first crank (22) being connected to the first crank (24) by a first link (25) to the second crank (27):

-a first relative angular position, wherein when the leg is in the deployed position, the first crank (22) and the first link (25) are in a first alignment, thereby stabilizing each of the branches of the brace member in a substantially aligned position; and

-a second relative angular position, wherein the second crank (24) and the second link (27) are in a second alignment while the leg is in the retracted position, thereby stabilizing the leg in the retracted position.

2. The method of claim 1, wherein the first alignment and the second alignment are defined by an abutment between the crank and an associated connecting rod.

3. The method of claim 1, wherein a door (30) is coupled to the strut member, the door closing a bay that houses the landing gear in the retracted position, the door being in a closed position when the landing gear is in the deployed position and when the landing gear is in the retracted position.

4. A method according to claim 3, characterized in that when the door is in the closed position, a prestressing force is established on the door (30) such that the prestressing force ensures the alignment of the crank with the connecting rod assembly concerned.

5. A method according to claim 1, characterized in that a resilient abutment (50) is placed on the structure of the aircraft so that the panel (11) of the landing gear is in contact with the abutment and exerts a pre-stress on the landing gear, ensuring the alignment of the second link and the second crank when the landing gear is in the retracted position.

6. A method according to claim 5, wherein the abutment comprises a piston (54) which is held protruding from the body (52) by the action of a belleville washer (55) which exerts a pre-stress on the piston (54) against which the panel (11) of the strut member bears when the landing gear is in the retracted position.