An assembly including an aircraft turbine engine and a mounting pylon therefor

Abstract

The assembly comprises an aircraft turbine engine (10) and a pylon (32) for mounting the turbine engine on elements of the aircraft, the pylon comprising members (34, 36) for suspending the turbine engine (10), these members being connected to the turbine engine in at least one plane (PI) perpendicular to the axis (A) and located upstream of the combustion chamber (18) of the turbine engine, so that the turbine engine (10) is attached to the pylon (32) in a cantilevered manner, characterized in that the assembly further comprises at least one damper (40) connecting the turbine engine (10) to the pylon (32) and located in a plane (P2) perpendicular to the axis (A), this plane being located downstream of the combustion chamber (18), said damper (40) being configured to limit the relative displacement between the turbine engine (10) and the pylon (32) without transmitting forces.

Description

Technical Field

[0001] This invention relates to an assembly comprising an aircraft turbine engine and its mounting bracket. Background Technology

[0002] The prior art specifically includes documents FR-A1-2 969 700, FR-A1-2 987 401 and FR-A1-3053661.

[0003] An aircraft turbine engine includes a gas generator that, from upstream to downstream, typically comprises at least one compressor, an annular combustion chamber, and at least one turbine, referring to the flow of gas within the turbine engine. In the case of a twin-body low-pressure and high-pressure turbofan 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 drives a propulsion propeller, typically located upstream of the gas generator, to rotate.

[0005] When the propeller is streamlined and thus enclosed in an annular casing, it is called a fan and generates an airflow that surrounds the gas generator. Even when the propeller is not streamlined, it still generates an airflow that surrounds the 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 and includes beams extending parallel to the longitudinal axis of the turbine engine. If the turbine engine is attached under the wing of the aircraft, the pylon is positioned at the 12 o'clock position, similar to the face of a clock.

[0007] In current technology, the hanger comprises an upstream suspension member and a downstream suspension member for the turbine engine. However, this configuration leads to its disadvantages. During operation, the gas generator ensures stress is transferred between the upstream and downstream attachment points on the hanger, resulting in deformation of the generator and gas, and altering the clearance between the rotor and stator of the gas generator. Consequently, the gas generator is subjected to torque generated by axial stresses (off-axis thrust and thrust recovery). The turbine engine is also subjected to torques generated by the asymmetry of axial stresses on the fan propeller blades, as well as stresses generated by air trapping by the turbine engine (sleeve stress).

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

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

[0010] However, mounting a turbine engine in a cantilever configuration has its disadvantages:

[0011] - The rear section of the turbine engine no longer has any support, so it will not stop if the turbine engine is moved too far.

[0012] - 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 loss of fan blades);

[0013] - 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 major displacement (especially major displacement at the downstream end of the turbine engine).

[0014] - Issues concerning the equipment housing and related items; etc.

[0015] The present invention proposes improvements to the prior art that enable the resolution of at least some of the aforementioned problems and disadvantages. Summary of the Invention

[0016] The present 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 comprising, from upstream to downstream in the gas flow orientation, at least one compressor, an annular combustion chamber, and at least one turbine. The pylon has a generally elongated shape along the axis and includes components for suspending the turbine engine, these components being connected to the turbine engine in at least one plane perpendicular to the axis and located upstream of the combustion chamber, such that the turbine engine is cantilevered to the pylon. The assembly is characterized in that it further includes at least one damper connecting the turbine engine to the pylon and located in a plane perpendicular to the axis, downstream of the combustion chamber, the damper being configured to limit relative displacement between the turbine engine and the pylon without transmitting stress.

[0017] Therefore, this invention proposes to dampen the rear or downstream portion of a turbine engine, which is cantilevered and mounted to a hanger, via the front or upstream portion of the turbine engine. Although the damper is located downstream of the turbine engine, particularly downstream of the combustion chamber, its purpose is not to transfer stress from the turbine engine to the hanger, but rather to dampen and limit the relative displacement and vibration of the turbine engine only during operation. Thus, unlike conventional suspension members that form rigid connections, the damper forms a flexible connection.

[0018] Components according to the invention may include one or more of the following features, which may be adopted independently of each other or in combination with each other:

[0019] The damper is configured to limit displacement (between the turbine and the hanger) within the plane during a first predetermined stroke along a first direction contained in the plane and a second predetermined stroke along a second direction contained in the plane and perpendicular to the first direction, without transmitting stress. The damper is also configured to prevent any displacement of the turbine relative to the hanger beyond these strokes by transmitting stress. To avoid excessive displacement of the turbine beyond certain displacement strokes of the turbine, the turbine is blocked by the damper, which then ensures stress transmission between the turbine and the hanger. Therefore, the damper has different functions depending on the operation of the turbine; during normal operation, the damper functions as a conventional damper (or a flexible connection), while during final operation, the damper functions as a stress-transmitting (or rigid connection) connection.

[0020] - The damper includes at least two links, a first link including a first end hinged to the hanger, a second link including a first end hinged to the turbine engine, the first link and the second link including a second end connected together by a damping device, the first end and the second end of the link having a hinge or connecting axis that is generally parallel to each other and parallel to the longitudinal axis of the turbine engine.

[0021] - The first end of the connecting rod is connected to the hanger and the turbine engine respectively via a slewing connection;

[0022] - The damping device is a tray-type damping device and includes a cage fixed to one of the first and second links and a tray fixed to the other of the first and second links, the tray being rotatable in a cavity filled with damping fluid in the cage;

[0023] - Damping fluids are non-Newtonian fluids;

[0024] - Damping fluids are Newtonian fluids;

[0025] - The cage cavity includes a stop that is configured to limit the angular deflection of the tray within the cavity;

[0026] - The tray includes two radially opposed ears configured to support the stop;

[0027] - The damping device is a piston-type or spring-type damping device;

[0028] - The damping device is equipped with a motion sensor. Attached Figure Description

[0029] Other features and advantages of the invention will become apparent from the following detailed description, and with reference to the accompanying drawings for understanding the description, in which:

[0030] [ Figure 1 ] Figure 1 This is a very schematic view of an assembly including an aircraft turbine engine and its mounting bracket, based on the technology prior to this invention.

[0031] [ Figure 2 ] Figure 2 This is a very schematic view of an aircraft turbine engine based on the prior art of the present invention, and shows the attachment and suspension points on the mounting bracket;

[0032] [ Figure 3 ] Figure 3 This is a very schematic view of another aircraft turbine engine based on the prior art of the present invention, and shows the attachment and suspension points on the mounting bracket;

[0033] [ Figure 4 ] Figure 4 This is a schematic perspective view of a component according to the invention, which includes an aircraft turbine engine and its mounting bracket;

[0034] [ Figure 5 ] Figure 5 This is a schematic front view of a damper according to a first embodiment of the present invention;

[0035] [ Figure 6 ] Figure 6 yes Figure 5 A schematic view of the tray-type damping device shown;

[0036] [ Figure 7 ] Figure 7 yes Figure 4 A schematic side view of the damper shown;

[0037] [ Figure 8 ] Figure 8 Is with Figure 7 Similar view, and with Figure 7This demonstrates the damper's ability to shift and articulate;

[0038] [ Figures 9a-9b ] Figure 9a and Figure 9b Is with Figure 5 A similar view, showing the two extreme positions of the damper;

[0039] [ Figures 9c-9d ] Figure 9c and Figure 9d Is with Figure 5 A similar view, showing the other two extreme positions of the damper;

[0040] [ Figures 10a-10b ] Figure 10a and Figure 10b Is with Figure 9a and Figure 9d A similar view is shown, and a variant embodiment is illustrated;

[0041] [ Figures 11a-11b ] Figure 11a and Figure 11b Is with Figure 9a and Figure 9d A similar view is shown, and another variation of the embodiment is illustrated;

[0042] [ Figure 12 ] Figure 12 It is similar to Figure 5 The view shows another variation of the invention; and

[0043] [ Figure 13 ] Figure 13 It is similar to Figure 5 The view shows another variation of the invention. Detailed Implementation

[0044] Figure 1 A turbine engine 10 for use in an aircraft is shown. The turbine engine 10 is a turbofan engine with a dual-flow and dual-body configuration.

[0045] Axis A is the longitudinal axis of the turbine engine. The orthogonal reference coordinate system XYZ is used in some diagrams (including...). Figure 1 As shown in the diagram. Direction X is parallel to axis X and oriented towards the upstream or front of turbine engine 10, axis Z is oriented upward, and axis Y is oriented to one side.

[0046] The turbine engine 10 includes a gas generator 12, which, with reference to the flow of gas along axis A from upstream to downstream, includes an LP 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 an LP or low-pressure turbine 22.

[0047] Despite Figure 1 As not visible in the image, the rotor of HP compressor 16 is connected to the rotor of HP turbine 20 via a high-pressure shaft, and the rotor of LP compressor 14 is connected to the rotor of LP turbine 22 via a low-pressure shaft that passes through the high-pressure shaft and drives the propulsion propeller located upstream of gas generator 12 to rotate. This propulsion propeller is surrounded by an annular housing of fan housing 24.

[0048] 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.

[0049] Gas generator 12 defines a main annular flow channel for the first airflow (referred to as the main flow). The gas generator is surrounded by a secondary annular flow channel for the secondary airflow (referred to as the secondary flow).

[0050] The airflow entering the fan is divided into a portion to form the mainstream. The air in this mainstream is compressed in LP 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 LP turbine 22 and finally flow into exhaust nozzle 30.

[0051] Another portion of the airflow entering the fan forms a secondary flow and is designed to mix with the mainstream flow downstream of nozzle 30.

[0052] The turbine engine 10 is attached to the aircraft by means of a gantry 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.

[0053] Figures 1 to 3 Prior art prior to this invention is shown.

[0054] exist Figure 1 and Figure 2 In the first scenario shown, the hanger 32 is attached to three points or regions of the turbine engine 10. Two points are located upstream or in the front plane P1 perpendicular to axis A, and the last point is located downstream or in the rear plane P2 perpendicular to axis A.

[0055] At plane P1, the first attachment member 34 ensures that the pylon 32 is connected to the fan housing 24. At plane P2, the attachment member 38 ensures that the pylon 32 is attached to the turbine or exhaust housing 40. 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.

[0056] exist Figure 3In the second case shown, only the two attachment points in plane P1 exist, so the turbine engine is cantilevered to the hanger 32. In this case, at plane P1, attachment member 34 ensures that the hanger 32 is connected to the fan housing 24, and thrust recovery rod 36 ensures that the hub 28 of the intermediate housing 26 is connected to the hanger 32 by means of attachment member (not shown), which is attached to the hanger but not to the turbine engine.

[0057] Figure 4 The general principle of the invention is illustrated schematically. In addition to providing suspension members 34, 36 of the turbine engine 10 located upstream of the combustion chamber 18, the general principle of the invention also provides at least one damper 40, which connects the turbine engine 10 to the hanger 32 and is located downstream of the combustion chamber 18.

[0058] 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 components 34 or 36.

[0059] The damper 40 extends in a plane P2 perpendicular to axis A, which, for example, passes through the turbine or exhaust casing of the turbine engine 10. The damper 40 is configured to limit the relative displacement between the turbine engine 10 and the hanger 32 without transmitting stress, as will be explained below.

[0060] Figures 5 to 9d A first embodiment of the damper 40 (in this case, a tray-type damper) is shown.

[0061] The damper 40 includes at least two links 42 and 44. The first link 42 includes a first end 42a hinged to a U-shaped clamp 45a of the hanger 32, and the second link 44 includes a first end 44a hinged to a U-shaped clamp 45b of the turbine engine 10.

[0062] Links 42 and 44 include second ends 42b and 44b connected together by a damping device 46. For example... Figure 7 and Figure 8 As shown, especially Figure 7 In the visible static position, the ends 42a, 42b, 44a, 44b of the connecting rods 42 and 44 have hinged or connecting axes B, C, and D that are generally parallel to each other and parallel to the longitudinal axis A of the turbine engine 10.

[0063] The ends 42a and 44a of the connecting rods 42 and 44 are preferably connected to the hanger 32 and the turbine engine 10 respectively via a slewing connection, so that the damper 40 allows the turbine engine 10 to be relatively displaced relative to the hanger 32 along axis A (see...). Figure 8This prevents constraints caused by operating temperature during engine extension.

[0064] In this first embodiment, the damping device 46 is a tray-type damping device and includes a cage 48 fixed to the connecting rod 42 and a tray 50 fixed to the connecting rod 44. The tray 50 can rotate within a cavity 52 in the cage 48, which is filled with damping fluid.

[0065] In the example shown, the cavity 52 of the cage 48 includes stops 54a, 54b, 56a, 56b, which are configured to limit the angular deflection of the tray 50 within the cavity (see [reference]). Figure 6 This restricts the relative displacement of the turbine engine 10 relative to the hanger 32. The tray 50 includes two radially opposing lugs 50a and 50b configured to rest on stops 54a, 54b, 56a, and 56b. Lug 50a is displaceable between stops 54a and 54b, and lug 50b is displaceable between stops 56a and 56b. A first restricted position of the tray 50 in the cavity 48 corresponds to the lugs 50a and 50b abutting against stops 54a and 56b, respectively. A second restricted position of the tray 50 in the cavity 48 corresponds to the lugs 50a and 50b abutting against stops 54b and 56a, respectively.

[0066] Figures 9a to 9d The extreme positions of the turbine engine 10 relative to the hanger 32 are shown in the plane P2 extended by the damper 40.

[0067] exist Figure 9a In the middle, the turbine engine 10 is shifted upward (shift along axis Z: Z+) and also shifted to one side (shift along axis Y = Y-), opposite the hanger 32. The end 44a of the connecting rod 44 (in particular, the connection axis between this end and the turbine engine) is located at the first corner of the square of the shift of this end in the aforementioned plane.

[0068] The length of one side of this square corresponds to the predetermined stroke displacement of the rear portion of the turbine engine relative to the pylon. The square has two parallel sides oriented in the Z direction and two parallel sides oriented in the Y direction.

[0069] exist Figure 9b In the middle, the turbine engine 10 is shifted downwards (shifted along axis Z: Z-), and in relation to... Figure 9a The same side is shifted (shifted along axis Y: Y-), opposite to hanger 32. End 44a of link 44 is located at the second corner of the shifted square.

[0070] exist Figure 9cIn the middle, the turbine engine 10 is shifted upward (shifted along axis Z: Z+) and also shifted upward on the opposite side (shifted along axis Y: Y+), opposite the hanger 32. The end 44a of the connecting rod 44 is located on the triangular portion of the shifted square.

[0071] Finally, Figure 9d In the middle, the turbine engine 10 is shifted downwards (shifted along axis Z: Z-), and in relation to... Figure 9c The same side is shifted (shifted along axis Y: Y+), opposite to hanger 32. The end 44a of link 44 is located at the fourth and final corner of the shifted square.

[0072] Outside of these strokes, the pallet rests against the cage stop to prevent any further displacement of the turbine engine and, if necessary, to ensure that stress is transferred from the turbine engine to the hanger.

[0073] The liquid contained in cavity 48 for displacing tray 50 can be Newtonian or non-Newtonian.

[0074] Non-Newtonian fluids are fluids whose viscosity varies with the applied constraint or force and with time. The most common example of a non-Newtonian fluid is cornstarch dissolved in water.

[0075] The properties of Newtonian fluids (such as water) can only be described by temperature and pressure. However, the physical properties of non-Newtonian fluids depend on the forces acting on them over time. For example, the viscosity of some fluids increases with increasing applied force. These fluids are called rheologically thickening fluids. Conversely, the viscosity of some fluids decreases when subjected to force; these fluids are also rheologically thickening fluids.

[0076] If a Newtonian fluid is used, the fluid can be oil, silicone fluid, water, etc.

[0077] In preferred cases where a non-Newtonian fluid is used, the fluid is preferably rheologically thickened and, for example, a fluid composed of coarse particles in the liquid. When there are far more particles than liquid, the particles slide against each other when at rest, and the mixture has a very low viscosity. However, when constraints are applied, the particles block one another, and the liquid can no longer flow between the particles. The mixture solidifies, and the viscosity of the mixture increases.

[0078] like Figures 9a to 9d As shown, the turbine engine shifts in directions Z and Y, depending on the constraints applied to the stator (flight influence range, vibration, takeoff, landing, etc.).

[0079] The four maximum positions create an angular difference between the two links 42 and 44 of the damper 40. This angular difference with the non-Newtonian fluid in the cavity 48 produces various reactions:

[0080] - When tray 50 moves slowly and / or the force applied by the tray is low, the viscosity of the liquid remains low.

[0081] The rapid movement of tray 50 increases the viscosity of the liquid, which inhibits and restricts this movement.

[0082] Therefore, the use of non-Newtonian fluids causes the damping rate to vary with velocity and stress transmitted by tray 50.

[0083] Therefore, particularly under extreme dynamic loads (loss of the ring gear portion of the propeller blade or turbine blade) or under extreme static maneuvers (G / gyroscope coefficient), the present invention enables protection of the structure from significant loads while retaining the benefits of the suspension system selected as a cantilever to protect the performance of the turbine engine under nominal load conditions.

[0084] Figure 10a and Figure 10b A second embodiment of the damper 40 (in this case, a piston 60 type damper) is shown.

[0085] The damper 40 includes four links 42 and 44. Each of the two first links 42 includes a first end 42a hinged to the bracket 32, and each of the two second links 44 includes a first end 44a hinged to the turbine engine 10.

[0086] The second end 42b of one of the connecting rods 42 and the second end 44b of one of the connecting rods 44 are hinged together. The second end 42b of the other connecting rod 42 and the second end 44b of the other connecting rod 44 are connected together by a damping device 46 with a piston 60.

[0087] The ends 42a and 44a of the connecting rods 42 and 44 are preferably connected to each other by a slewing connection and are respectively connected to the hanger 32 and the turbine engine 10.

[0088] The damping device 46 includes a piston 60 fixed to the connecting rod 42 and a cylinder 62 fixed to the connecting rod 44, wherein the piston 60 can slide. The piston 60 is capable of sliding within the cylinder in the aforementioned plane P2.

[0089] As described above, cylinder 62 is filled with damping fluid.

[0090] The stroke of damper 40 is calculated as a margin with a functional range and a two-stroke termination stop with a piston.

[0091] Figure 11a and Figure 11b A second embodiment of the damper 40 (in this case, a spring-type damper, i.e., having an elastically deformable element) is shown.

[0092] The damper 40 includes four links 42 and 44. Each of the two first links 42 includes a first end 42a hinged to the bracket 32, and each of the two second links 44 includes a first end 44a hinged to the turbine engine 10.

[0093] The second end 42b of one of the links 42 and the second end 44b of one of the links 44 are hinged together. The second end 42b of the other link 42 and the second end 44b of the other link 44 are connected together by a damping device 46 with a spring 70.

[0094] The ends 42a and 44a of the connecting rods 42 and 44 are preferably connected to each other by a slewing connection and are respectively connected to the hanger 32 and the turbine engine 10.

[0095] The damping device 46 includes a spring 70 fixed to the link 42 and a guide 72 fixed to the link 44, wherein the spring 70 can deform in or on the link 44. The spring 70 is deformable in the plane P2 described above.

[0096] The stroke of damper 40 is calculated as a margin with a functional range and a spring-loaded two-stroke termination stop.

[0097] Figure 12 and Figure 13 Other variations are shown, including equipping the damping device 46 with a motion sensor 80.

[0098] exist Figure 12 In the process, the damping device 46 with tray 50 is equipped with a rotational motion sensor 80, which measures the angular deflection of tray 50 about its axis of rotation.

[0099] exist Figure 13 In the case shown, the damping device 46 with piston 70 or spring 80 is equipped with translation sensor 80, which measures the deflection of the piston or the deformation of the spring.

[0100] During nominal operation, the relative displacement and vibration of the turbine engine are damped by dampers. This is the case, for example, during normal turbine engine operation and aircraft maneuvering. There is no stress transfer between the turbine engine and the pylon.

[0101] In final operation (such as blade loss and final maneuvering of the aircraft), the damper prevents and ensures the transmission of stress.

[0102] Therefore, the present invention proposes to position the damper 40 between the turbine engine 10 and the hanger 32 so that:

[0103] - Create a stop that can only be actuated under large displacement conditions (system size is a certain value).

[0104] - Limiting resonance in cantilevered turbine engines through dampers, and

[0105] - No stress is generated on the cantilever turbine engine during normal operation (limited displacement and low vibration frequency).

Claims

1. An assembly comprising an aircraft turbine engine (10) and a pylon (32), the pylon being used to mount the aircraft turbine engine to components of an aircraft. The aircraft turbine engine has a longitudinal axis (A) and includes a gas generator (12) which, from upstream to downstream, includes at least one compressor (14, 16), an annular combustion chamber (18), and at least one turbine (20, 22). The suspender has a generally elongated shape along the longitudinal axis (A) and includes components (34, 36) for suspending the aircraft turbine engine (10), which are connected to the aircraft turbine engine in at least one first plane (P1) perpendicular to the longitudinal axis (A) and located upstream of the annular combustion chamber (18), such that the aircraft turbine engine (10) is cantilevered to the suspender (32). Its features are, The assembly also includes at least one damper (40) that connects the aircraft turbine engine (10) to the pylon (32) and is located in a second plane (P2) perpendicular to the longitudinal axis (A) downstream of the annular combustion chamber (18). The damper (40) is configured to limit relative displacement between the aircraft turbine engine (10) and the pylon (32) in the second plane (P2) without transmitting stress.

2. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 1, wherein, The damper (40) is configured to limit displacement within the second plane (P2) on a first predetermined stroke in a first direction (Z) and a second predetermined stroke in a second direction (Y) without transmitting stress, the first direction being contained in the second plane (P2) and the second direction being contained in the second plane (P2) and perpendicular to the first direction (Z), and the damper is configured to prevent any displacement of the aircraft turbine engine relative to the pylon beyond the first predetermined stroke and the second predetermined stroke by transmitting stress.

3. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 1 or 2, wherein, The damper (40) includes at least two links, a first link (42) including a first end (42a) hinged to the hanger (32), and a second link (44) including a first end (44a) hinged to the aircraft turbine engine (10). The second ends (42b) of the first link (42) and the second ends (44b) of the second link (44) are connected together by a damping device (46). The first ends (42a) and the second ends (42b) of the first link (42) and the first ends (44a) and the second ends (44b) of the second link (44) have connecting axes (B, C, D) that are generally parallel to each other and parallel to the longitudinal axis (A) of the aircraft turbine engine.

4. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 3, wherein, The first end (42a) of the first link (42) and the first end (44a) of the second link (44) are respectively connected to the gantry (32) and the aircraft turbine engine (10) via a swivel connection.

5. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 3, wherein, The damping device (46) is a tray-type damping device and includes a cage (48) fixed to one of the first and second links and a tray (50) fixed to the other of the first and second links, the tray (50) being rotatable in a cavity (52) filled with damping fluid in the cage (48).

6. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 5, wherein, The damping fluid is a non-Newtonian fluid.

7. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 5, wherein, The damping fluid is a Newtonian fluid.

8. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to any one of claims 5 to 7, wherein, The cavity (52) of the cage (48) includes stops (54a, 54b, 56a, 56b) configured to limit the angular deflection of the tray (50) in the cavity (52).

9. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 8, wherein, The tray (50) includes two radially opposing ears (50a, 50b) configured to be supported on the stops (54a, 54b, 56a, 56b).

10. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 3, wherein, The damping device (46) is a piston-type or spring-type damping device.

11. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 3, wherein, The damping device (46) is equipped with a motion sensor (80).

12. The assembly comprising an aircraft turbine engine (10) and a pylon (32) according to claim 3, wherein, The connecting axes (B, C, D) are hinge axes.

Images