Including aircraft turbine engines and their mounting brackets.

Abstract

The present invention relates to an assembly comprising an aircraft turbine engine (10) and a pylon (32) for mounting the turbine engine to an element of an aircraft, the assembly further comprising at least one system (40) for limiting relative movement between the turbine engine (10) and the pylon (32), the system comprising a first fixing member (42) and a second fixing member (46), the first fixing member being connected to the pylon (32) and including at least one first stop element (44), the second fixing member being connected to the turbine engine (10) and including at least one second stop element (48), the first stop element and the second stop element (44, 48) being parallel to an axis (A) and separated from each other by at least one elastically deformable device (54) to engage with each other by abutment and limit relative movement between the turbine engine (10) and the pylon (32).

Description

Technical Field

[0001] The present invention relates to an assembly comprising an aircraft turbine engine and a mounting bracket for the aircraft turbine engine. Background Technology

[0002] Prior art includes, in particular, documents FR-A1-2969700, FR-A1-2987401, FR-A1-3118992, US-B2-11,560,840, US-A11-2021 / 284348, FR-A1-3114129, and FR-A1-3118992.

[0003] An aircraft turbine engine includes a gas generator, which typically comprises at least one compressor, an annular combustion chamber, and at least one turbine, relative to the gas flow in the turbine engine from upstream to downstream. In the case of a low-pressure and high-pressure twin-shaft turbojet engine, the gas generator includes a low-pressure compressor, a high-pressure compressor, a combustion chamber, a high-pressure turbine, and a low-pressure turbine. The gas generator defines an annular flow channel for the airflow through the compressor, combustion chamber, and turbine.

[0004] The rotor of the high-pressure compressor is connected to the rotor of the high-pressure turbine via a high-pressure shaft. The rotor of the low-pressure compressor is connected to the rotor of the low-pressure turbine via a low-pressure shaft that passes through the high-pressure shaft and rotates the propeller / fan blades, which are typically located upstream of the gas generator.

[0005] When the propeller is ducted and therefore surrounded by an annular casing, the propeller is called a fan and generates airflow around a gas generator. When the propeller is unducted, it also generates airflow around a gas generator.

[0006] The turbine engine is attached to components of the aircraft, such as the wing or fuselage, via a mounting pylon also known as a mast. This pylon is generally elongated in shape and includes beams extending parallel to the longitudinal axis of the turbine engine. If the turbine engine is mounted under the wing of the aircraft, the pylon is positioned at what resembles the 12 o'clock position (12 hours) on a clock face.

[0007] In current technology, the mounting bracket includes an upstream suspension member for suspending the turbine engine and a downstream suspension member for suspending the turbine engine. However, this configuration has certain drawbacks. During operation, the gas generator transmits forces between the upstream and downstream attachment points on the mounting bracket, causing the gas generator to deform and altering the clearance between the rotor and stator. Therefore, the gas generator is subjected to torque generated by axial forces (off-axis thrust and thrust reaction). The turbine engine is also subjected to torque generated by the asymmetry of axial forces on the propeller's fan blades, as well as forces generated by the turbine engine trapping air (duct forces). In the case of a ductless propeller, the turbine engine is subjected to a yaw force known as Mode 1P.

[0008] Therefore, it is understandable that the performance and operability of a turbine engine may be affected by these forces.

[0009] One solution to this problem is to attach the turbine engine to a mounting bracket in a cantilever manner. This means suspending the front or upstream portion of the turbine engine to the mounting bracket, while leaving the rear or downstream portion of the turbine engine (such as its turbine housing) free.

[0010] However, mounting a turbine engine in a cantilever configuration has its disadvantages: - There is no longer any support for the rear part of the turbine engine, so there is no stop if the turbine engine is displaced too far; - The cantilever section will have a bending mode that rotates at a relatively low frequency, which can be excited by unbalanced loads, especially unbalanced loads caused by losses of fan blades. - The loads associated with the loss of low-pressure turbine blades will be transmitted along the entire line of the cantilever housing, resulting in significant loads and vibrations at the bottom of the cantilever, as well as large displacements, particularly at the downstream end of the turbine engine. - Issues related to shell strength and shell-related equipment; etc.

[0011] The applicant has proposed a solution to this problem in document FR-A1-3118992. This solution involves installing a damper between the turbine engine and the mounting bracket, located downstream of the combustion chamber. The damper is configured to limit relative movement between the turbine engine and the mounting bracket without transmitting forces.

[0012] This invention proposes improvements to the current technology that allow for the resolution of at least some of the aforementioned problems and disadvantages. Summary of the Invention

[0013] This invention relates to an assembly comprising an aircraft turbine engine and a pylon for mounting the turbine engine to components of an aircraft. The turbine engine has a longitudinal axis and includes a gas generator, which, from upstream to downstream in the direction of airflow, includes at least one compressor, an annular combustion chamber, and at least one turbine. The mounting bracket has a generally elongated shape along the axis and includes suspension members for suspending the turbine engine, all of which are connected to the turbine engine in at least a first plane such that the turbine engine is cantilevered to the mounting bracket, the first plane being perpendicular to the axis and located upstream of the at least one turbine. The assembly also includes at least one system for limiting relative movement between the turbine engine and the mounting bracket, the system being connected to the turbine engine in at least one second plane perpendicular to the axis and located downstream of the combustion chamber. The limiting system is characterized in that it includes a first fixing member and a second fixing member, the first fixing member being connected to the mount and including at least one first stop element, the second fixing member being connected to the turbine engine and including at least one second stop element, the first stop element and the second stop element being parallel to the axis and capable of limiting relative movement between the turbine engine and the mount, and the first stop element and the second stop element being separated from each other by at least one elastically deformable device capable of damping the relative movement.

[0014] Therefore, the present invention proposes a system for limiting relative movement between a turbine engine and a mounting bracket, the system being simplified and equipped with an elastic member acting as a damper. The orientation of the element parallel to the longitudinal axis of the turbine engine is advantageous because it allows relative movement between the turbine engine and the mounting bracket in a direction parallel to that axis. When the rear of the turbine engine moves in one or more other directions, the element is configured to engage with it by abutment to prevent relative movement between the turbine engine and the mounting bracket.

[0015] The components and parts of the limiting system are not designed to transmit forces from the turbine engine toward the mounting bracket, but only to prevent relative movement of the turbine engine in certain directions during operation.

[0016] In the context of this invention, the "device" or each elastically deformable "device" may be a stud, block, pad, layer, coating, strap, etc.

[0017] Components according to the invention may include one or more of the following features, individually or in combination with each other: - The restraining system is configured such that the first stop element and the second stop element have at least one degree of freedom relative to each other in a direction parallel to the axis; --The first plane is located upstream of the combustion chamber; - One of the first stop element and the second stop element includes a hollow tube, and the other of the first stop element and the second stop element includes a finger engaged within the hollow tube, the finger being separated from the hollow tube by the at least one means extending around the finger; - The finger is capable of moving axially within the hollow tube, for example, a distance of 50 mm or greater; - The at least one device is formed of an annular layer of elastically deformable material, the annular layer covering the interior of the hollow tube and attached to the hollow tube; - The at least one device is formed of an annular layer of elastically deformable material, the annular layer covering the finger and attached to the finger; - The first stop element includes a first plate, and the second stop element includes a second plate, the first plate and the second plate being separated by the at least one device; - One of the first stop element and the second stop element includes a third plate, the plate of the first stop element being inserted between the two plates of the other stop element and being separated from those plates by the at least one means; - The plate is tangent to a circumference centered on the axis; - Two plates of one of the stop elements are connected together by a bottom wall, which is separated from the plate of the other stop element by the at least one device; - The first stop element includes a finger engaged in a receiving portion in the second stop element; -The interior of the finger or the receiving portion is coated with an elastically deformable material layer forming the at least one device; - The finger is connected to the interior of the receiving portion via the at least one device; --The first component is rigidly connected to the bracket; --The first component is connected to the bracket via a V-shaped or X-shaped link; --The second element is rigidly connected to the turbine engine; --The second element is connected to the turbine engine via a V-shaped or X-shaped connecting rod; --The limiting system is located in a plane passing through the axis of the mounting bracket and the turbine engine; --When the bracket is at 12 o'clock (similar to a clock face), the limiting system is at 12 o'clock; --The limiting system is located between the mounting bracket and the turbine engine, for example, at the 3 o'clock or 9 o'clock position; --The component includes two limiting systems located at 2-3 o'clock and 9-10 o'clock, respectively. Attached Figure Description

[0018] Other features and advantages of the invention will become apparent from the following detailed description, in which reference is made to the accompanying drawings, in which: [ Figure 1 ] Figure 1 This is a very schematic view of an assembly including an aircraft turbine engine and its mounting bracket, according to prior art of the present invention; [ Figure 2 ] Figure 2 This is a very schematic view of an aircraft turbine engine according to the prior art of the present invention, and shows the attachment points and suspension points on the mounting bracket; [ Figure 3 ] Figure 3 This is a very schematic view of another aircraft turbine engine according to the prior art of the present invention, and shows the attachment points and suspension points on the mounting bracket; [ Figure 4 ] Figure 4 It is a schematic perspective view of the components including the aircraft's turbine engine and its mounting bracket; [ Figure 5 ] Figure 5 This is a very schematic side view of an assembly including an aircraft turbine engine and its mounting bracket according to an embodiment of the present invention; [ Figure 6 ] Figure 6 yes Figure 5 A schematic front view of the components shown in the figure. [ Figure 7 ] Figure 7 This is a very schematic side view of an assembly including an aircraft turbine engine and its mounting bracket, according to a variant embodiment of the present invention. [ Figure 8 ] Figure 8 This is a very schematic side view of an assembly including an aircraft turbine engine and its mounting bracket, according to another embodiment of the present invention; [ Figure 9 ] Figure 9 This is a very schematic view of a system for limiting relative movement between a turbine engine and a mounting bracket, according to another variant embodiment of the invention; and [ Figure 10 ] Figure 10 This is a very schematic view of a system for limiting relative movement between a turbine engine and a mounting bracket, according to yet another variant of the invention. Detailed Implementation

[0019] Figure 1 A turbine engine 10 for use in an aircraft is shown. The turbine engine 10 is a turbojet engine with a dual-flow and dual-shaft configuration.

[0020] Axis A is the longitudinal axis of the turbine engine. (Including...) Figure 1 Some of the figures show an orthogonal reference frame XYZ. The X direction is parallel to axis A and oriented toward the front or rear of the turbine engine 10, axis Z is oriented upwards, and axis Y is oriented to one side.

[0021] The turbine engine 10 includes a gas generator 12, which, relative to the flow of gas along axis A from upstream to downstream, includes a BP or low-pressure compressor 14, an HP or high-pressure compressor 16, an annular combustion chamber 18, an HP or high-pressure turbine 20, and a BP or low-pressure turbine 22.

[0022] Although this is in Figure 1 Not shown, but the rotor of HP compressor 16 is connected to the rotor of HP turbine 20 via a high-pressure shaft, and the rotor of BP compressor 14 is connected to the rotor of BP turbine 22 via a low-pressure shaft. The low-pressure shaft passes through the high-pressure shaft and drives the fan blades of the propulsion propeller located upstream of gas generator 12. The propulsion propeller is surrounded by an annular housing called fan housing 24.

[0023] The fan housing 24 is connected to the gas generator 12 via an intermediate housing 26, which includes a central hub 28 and a series of radial arms connecting the hub 28 to the fan housing 24.

[0024] Gas generator 12 defines a main annular flow duct for the first airflow (referred to as the main flow). The gas generator is used to surround a second annular flow duct for the second airflow (referred to as the secondary flow).

[0025] The airflow entering the fan is divided into a portion that forms the mainstream. The air in this mainstream is compressed in BP compressor 14 and HP compressor 16, then mixed with fuel and burned in combustion chamber 18. The combustion gases from the mainstream then expand in HP turbine 20 and BP turbine 22 and finally flow into exhaust nozzle 30.

[0026] Another portion of the airflow entering the fan forms a secondary flow and is configured to mix with the mainstream downstream of nozzle 30.

[0027] The turbine engine 10 is attached to the aircraft via a pylon 32, which has a generally elongated shape along axis A and includes components 34, 36, and 38 for attaching and suspending the turbine engine 10.

[0028] Figures 1 to 4 Prior art prior to this invention is shown.

[0029] exist Figure 1 and Figure 2 In the first scenario shown, there are three points or regions where the mount 32 is attached to the turbine engine 10. Two points are located in the upstream or front plane P1 perpendicular to axis A, and the last point is located in the downstream or rear plane P2 perpendicular to axis A.

[0030] The first plane P1 is located upstream of the turbines 20 and 22, and preferably upstream of the combustion chamber 18. The second plane P2 is located downstream of the combustion chamber 18.

[0031] At plane P1, the first attachment member 34 ensures the connection between the pylon 32 and the fan housing 24. At plane P2, the attachment member 38 ensures the attachment between the pylon 32 and the turbine or exhaust housing 39. This attachment member 38 is also connected to the hub 28 of the intermediate housing 26 via thrust recovery rods 36. These rods 36 ensure that thrust is transmitted from the turbine engine 10 to the pylon 32 and thus to the aircraft.

[0032] exist Figure 3 In the second scenario shown, there are only two attachment points in the aforementioned plane P1, so the turbine engine is cantilevered to the mount 32. In this case, at plane P1, the attachment member 34 ensures the connection between the mount 32 and the fan housing 24, and the thrust reaction link 36 ensures the connection between the hub 28 of the intermediate housing 26 and the mount 32 via an attachment member (not shown), which is attached to the mount but not to the turbine engine.

[0033] Figure 4 The principle described in document FR-A1-3118992 is schematically shown, which, in addition to the suspension members 34, 36 located upstream of the combustion chamber 18 for suspending the turbine engine 10, also provides a system 40 located downstream of the combustion chamber 18 for limiting the relative movement between the turbine engine 10 and the mounting bracket 32.

[0034] Components 34 and 36 absorb loads in the Y and Z directions, as well as moments Mx, My, and Mz in all directions. The thrust of the turbine engine in the X direction is recovered by a system integrated into or independent of component 34 or 36.

[0035] System 40 is located at plane P2, which is perpendicular to axis A and passes through, for example, the turbine or exhaust casing of turbine engine 10.

[0036] This invention is an improvement on this principle, and its embodiments are as follows: Figure 5 And as shown below.

[0037] Generally, in the context of this invention, the restraint system 40 includes a first fixing member 42 connected to the hanger 32 and including at least a first stop element 44, and a second fixing member 46 connected to the turbine engine 10 and including at least a second stop element 48.

[0038] The first stop element and the second stop elements 44 and 48 can be engaged together by abutting to limit the relative movement between the turbine engine 10 and the mount 32.

[0039] The first and second stop elements 44 and 48 are parallel to axis A.

[0040] The first stop element and the second stop elements 44 and 48 are separated from each other by at least one elastically deformable device 54 capable of damping the aforementioned relative movement.

[0041] Figure 5 and Figure 6 A first embodiment of the invention is shown, wherein one of the stop elements 44, 48 includes a hollow tube 50, and the other of the stop elements 44, 48 includes a finger 52 engaged within the hollow tube 50, the finger 52 being separated from the hollow tube 50 by the aforementioned means 54 extending around the finger 52 and taking the form of an annular layer.

[0042] In the example shown, the finger 52 is connected to the turbine engine 10, and the hollow tube 50 is connected to the mount 32. However, the reverse is also possible.

[0043] The connection between the finger 52 and the turbine engine 10 can be rigid. Alternatively, this connection can be achieved through a suspension system, such as crossed X- or V-shaped bars or linkages.

[0044] The hollow tube 50 can be rigidly connected to the turbine engine 10. Alternatively, this connection can be made via a suspension system, such as crossed X- or V-shaped bars or linkages.

[0045] Preferably, the finger 52 is axially movable within the hollow tube 50 due to the respective orientations of the finger and the hollow tube.

[0046] The interior of the hollow tube 50 may be coated with an annular layer of a device 54 for forming an elastically deformable material. This layer is then attached to the hollow tube 50.

[0047] Alternatively, the finger 52 may be coated with an annular layer of the device 54 forming an elastically deformable material. The layer 54 is then attached to the finger 52.

[0048] During operation, the finger 52 is capable of sliding axially within the hollow tube 50 to allow relative movement of the turbine engine 10 relative to the mount 32 along axis A. When the turbine engine 10 moves relative to the mount 32 in a direction transverse to axis A, these movements are damped and limited by the contact between the finger 52 and the hollow tube 50.

[0049] Figure 7 A variant embodiment of the invention is shown, wherein the first stop element 44 includes a first plate 56 and the second stop element 48 includes a second plate 58, and the first plate and the second plates 56, 58 are separated by the aforementioned device 54.

[0050] In the example shown, each of components 42, 46 includes a single plate 56, 58. These plates 56, 58 are parallel to axis A, and more specifically, tangent to a circumference centered on axis A. Plates 56, 58 are separated by a single device 54, which may take the form of a flat layer.

[0051] During operation, plates 56 and 58 can move freely axially relative to each other to allow relative movement of the turbine engine 10 relative to the mount 32 along axis A. When the turbine engine 10 moves radially toward the mount 32, this movement is damped and limited by the contact of plates 56 and 58 with each other.

[0052] Figure 8 Another embodiment of the invention is shown, wherein the first stop element 44 includes two plates 56, 60, and the second stop element 48 includes a plate 58 inserted between the two plates 56, 60 and separated from these plates by a damping device 54.

[0053] These plates 56 and 58 are parallel to axis A, and more specifically, tangent to a circumference centered on axis A. Plates 56 and 58 are separated, for example, by two devices 54 or layers.

[0054] Alternatively, the turbine engine 10 may be connected to two plates, with the plate connected to the mount 32 located between the two plates.

[0055] The two plates 56 and 60 can be connected together by the bottom wall 62, which is separated from the plate 58 by the device 54 or a separate device.

[0056] In fact, the plates can be connected to each other and to the bottom wall 62 by using several independent damping devices 54 or, conversely, a single damping device 54 that can occupy all or part of the space between the plate and the bottom wall 62.

[0057] During operation, the movement of plate 58 is damped and restricted by the cooperation of plates 56 and 60 and bottom wall 62.

[0058] Figure 8 and Figure 9 Another embodiment of the invention is shown, wherein the first stop element 44 includes a finger 64 that engages in a receiving portion 66 in the second stop element 48.

[0059] like Figure 10 As shown, the interior of the finger 64 or the receiving portion 66 may be coated with an elastically deformable material layer that forms the aforementioned damping device 54.

[0060] These variant operations are similar to Figure 8 The operation of the variant shown.

[0061] As shown in the example in the accompanying drawings, the limiting system can be located at 12 o'clock (12 hours), similar to the dial of a clock centered on axis A. Alternatively, the components according to the invention may include two limiting systems located at 2-3 o'clock and 9-10 o'clock respectively.

[0062] Generally speaking, the elastically deformable material used in this invention can be an elastomer.

Claims

1. An assembly comprising an aircraft turbine engine (10) and a pylon (32) for mounting said turbine engine to components of an aircraft, The turbine engine (10) has a longitudinal axis (A) and includes a gas generator (12) which includes at least one compressor (14, 16), an annular combustion chamber (18) and at least one turbine (20, 22) in the direction of airflow from upstream to downstream. The mounting bracket (32) has a generally elongated shape along the axis (A) and includes suspension members (34, 36) for suspending the turbine engine (10), all of which are connected to the turbine engine in at least a first plane (P1) such that the turbine engine (10) is cantilevered to the mounting bracket (32), the first plane being perpendicular to the axis (A) and located upstream of the at least one turbine (20, 22). The assembly also includes at least one system (40) for limiting relative movement between the turbine engine (10) and the mounting bracket (32), the system (40) being connected to the turbine engine (10) in at least one second plane (P2) perpendicular to the axis (A) and located downstream of the combustion chamber (18). characterized in that The limiting system (40) includes a first fixing member (42) and a second fixing member (46), the first fixing member being connected to the hanger (32) and including at least one first stop element (44), the second fixing member being connected to the turbine engine (10) and including at least one second stop element (48), the first stop element and the second stop element (44, 48) being parallel to the axis (A) and capable of limiting relative movement between the turbine engine (10) and the hanger (32), and the first stop element and the second stop element (44, 48) being separated from each other by at least one elastically deformable device (54) capable of damping the relative movement.

2. The assembly of claim 1, wherein, The restraint system (40) is configured such that the first stop element and the second stop element (44, 48) have at least one degree of freedom relative to each other in a direction parallel to the axis (A).

3. The assembly of claim 1 or 2, wherein, One of the first stop element and the second stop element (44, 48) includes a hollow tube (50), and the other of the first stop element and the second stop element (48, 44) includes a finger (52) engaged within the hollow tube (50), the finger (52) being separated from the hollow tube (50) by at least one means (54) extending around the finger (52).

4. The assembly of claim 3, wherein, The finger (52) is axially movable within the hollow tube (50).

5. The assembly of claim 3 or 4, wherein, The at least one device (54) is formed of an annular layer (54) of elastically deformable material, the annular layer covering the interior of the hollow tube (50) and attached to the hollow tube (50).

6. The assembly of claim 3 or 4, wherein, The at least one device (54) is formed of an annular layer (54) of elastically deformable material, the annular layer covering the finger (52) and attached to the finger (52).

7. The assembly of claim 1 or 2, wherein, The first stop element (44) includes a first plate (56), and the second stop element (48) includes a second plate (58), the first plate and the second plate (56, 58) being separated by the at least one device (54).

8. The assembly of claim 7, wherein, One of the first stop element and the second stop element (44, 48) includes a third plate (60), and a plate (58) of one of the stop elements (44, 48) is inserted between the two plates (56, 60) of the other stop element (48, 44) and separated from those plates (56, 60) by the at least one device (54).

9. The component according to claim 8, wherein, Two plates of one of the stop elements are connected together by a bottom wall, which is separated from the plate of the other stop element by the at least one device.

10. The component according to any one of claims 7 to 9, wherein, The plate is tangent to a circumference centered on the axis.

11. The component according to any one of claims 1 to 3, wherein, The first stop element (44) includes a finger (64) engaged in a receiving portion (66) in the second stop element (48).

12. The component of claim 11, wherein, The interior of the finger (64) or the receiving portion (66) is coated with an elastically deformable material layer (54) forming the at least one device (54).

13. The component of claim 11, wherein, The finger (64) is connected to the interior of the receiving portion (66) via the at least one device (54).

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