Assembly for an aircraft turbine engine comprising means for axially and radially retaining a fan

Abstract

The invention relates to an assembly (50) for an aircraft turbine engine, comprising a fan (15), a reduction gear (20) and an intermediate housing (25), the intermediate housing comprising an inner annular structure (52) housing an epicyclic gear train equipped with a planet carrier (42). According to the invention, the assembly comprises a secondary retaining device (70) for the ring gear (60) of the planet carrier, the device comprising: first projections (72a) projecting radially inwards from the inner annular structure (52), each first projection (72a) having a recess (74) opening radially inwards, axially downwards and circumferentially in a first direction; and second projections (72b) projecting radially outwards from the ring gear (60) of the planet carrier and cooperating with the first projections (72a) such that each second projection is partially housed in a recess (74) of a first projection associated with and downstream of the second projection.

Description

Technical Field

[0001] The present invention relates to the field of aircraft turbine engines, in particular to twin-flow and / or twin-body turbine engines, such as turbojets.

[0002] The invention relates in particular to a turbine engine comprising a fan driven by a speed reducer. Background Art

[0003] In aircraft turbine engines, it is known to provide a fan driven by a speed reducer to limit the fan's rotational speed relative to the low-pressure body. In a so-called planetary configuration, the speed reducer comprises an inner ring driven by the low-pressure body, a planet carrier attached to the turbine engine stator, and an outer gear ring engaged with the fan shaft. The planetary gears mesh with the inner and outer rings of the epicyclic gear train. This configuration is known, for example, from WO 2019 / 158883.

[0004] The fan shaft is typically held axially by a bearing, which itself is supported by a bearing support connected to the stator portion of the turbine engine. The bearing includes a stop that prevents the fan shaft from moving axially in the downstream to upstream direction.

[0005] In the event of a bearing or bearing support failure, there is a risk that the fan will move axially in an upstream direction. Therefore, it is necessary to provide a configuration that limits / avoids this risk. Summary of the Invention

[0006] In order to solve this problem, the subject of the invention is first of all an assembly for an aircraft turbine engine, which assembly includes a fan; a reducer that drives the fan and is located downstream of the fan; an intermediate casing, which includes an internal annular structure that internally delimits a cavity that at least partially accommodates the reducer, the reducer including an epicyclic gear train equipped with a planet carrier, one ring of which is fixed to the internal annular structure of the intermediate casing by main retaining means.

[0007] According to the invention, the assembly comprises a secondary retaining device of the annular element of the planet carrier relative to the inner annular structure of the intermediate casing, the secondary retaining device comprising:

[0008] - first projections projecting radially inwards from the inner annular structure of the intermediate casing, the first projections being circumferentially spaced apart from one another around the longitudinal axis of the turbine engine, each first projection comprising a recess opening radially inwards, opening axially downstream and opening circumferentially in a first circumferential direction and being delimited by three faces, namely, an axial retaining face, a radial retaining face and a circumferential retaining face;

[0009] - second projections projecting radially outwards from the annular element of the planetary carrier, the second projections being circumferentially spaced apart from one another around the longitudinal axis of the turbine engine and cooperating in pairs with the first projections so that each second projection is partially housed in a recess of the first projection associated with it, located downstream of the second projection.

[0010] The present invention solves this problem in a simple and reliable manner by establishing a cogging-type connection between radial projections. Specifically, by axially and radially securing the planet carrier of the reducer, the fan is also held axially and radially relative to the inner annular structure of the intermediate casing in the event of a failure of the fan shaft support bearing or of the bearing supporting the bearing. Thus, the present invention cleverly enables the planet carrier of the reducer to act on the fan to prevent / limit axial and radial movement of the fan in the event of a failure.

[0011] Furthermore, the fit between the first and second radial protrusions advantageously ensures the transfer of reducer / fan torque to the inner annular structure of the intermediate housing. This circumferential torque transfer via the radial protrusions can be observed continuously or only in the event of a malfunction. The circumferential gap between the first and second protrusions is preferably zero or very small, but alternatively, it may be larger without departing from the scope of the present invention.

[0012] However, the axial gap and the radial gap between the first radial protrusion and the second radial protrusion are preferably larger in order to limit the transmission of vibrations to the speed reducer during normal operation of the turbine engine.

[0013] The present invention preferably provides at least one of the following optional features considered individually or in combination.

[0014] As described above, the assembly is configured so that, in the normal operating configuration of the turbine engine, a radial gap is defined between each second protrusion and a radial retaining surface that defines a recess of the first protrusion with which the second protrusion cooperates, and / or an axial gap is defined between each second protrusion and an axial retaining surface that defines the recess, and / or a circumferential gap is defined between each second protrusion and a circumferential retaining surface that defines the recess.

[0015] Preferably, the inner annular structure of the intermediate housing and the first protrusion are made in one piece, preferably by casting.

[0016] Preferably, the main retaining means comprises a flexible half-segment annular member in the general shape of a U opening radially outwards, the upstream leg of the U being attached to the annular member of the planet carrier and the downstream leg of the U being attached to a flange of the inner annular structure of the intermediate casing, the flange protruding radially inwards.

[0017] Preferably, the number of first protrusions is between three and fifteen, preferably between six and ten.

[0018] Preferably, the epicyclic gear train comprises an outer ring gear meshing with the planet gears of the planet carrier, the outer ring gear being engaged with the fan shaft.

[0019] Preferably, the fan shaft is supported by bearings which ensure axial retention of the fan shaft in the direction from downstream to upstream.

[0020] Preferably, the bearing is supported by a bearing support fixed to the stator part of the assembly, preferably to the inner annular structure of the intermediate housing.

[0021] Preferably, the intermediate casing further comprises an outer annular structure connected to the inner annular structure by radial arms, the annular space between the two inner annular structures and the outer annular structure forming part of the main flow path of the turbine engine.

[0022] Another subject of the invention is an aircraft turbine engine, preferably a twin-flow, twin-body turbojet engine, comprising such an assembly.

[0023] Other characteristics and advantages of the invention appear in the non-limiting detailed description given below. BRIEF DESCRIPTION OF THE DRAWINGS

[0024] The description will be given with reference to the accompanying drawings, in which:

[0025] [ Figure 1 ] shows a schematic side view of a turbojet engine according to the invention;

[0026] [ Figure 2 ] shows in more detail an enlarged view of the speed reducer, which is assembled to the turbojet engine shown in the figure above;

[0027] [ Figure 3 ] is a front view of an intermediate housing forming an integral part of an assembly according to a preferred embodiment of the present invention, the assembly being used in Figure 1 In turbojet engines;

[0028] [ Figure 4 ] is a perspective view of a portion of an assembly;

[0029] [ Figure 5 ] are perspective views of a portion of an assembly viewed from different angles; and

[0030] [ Figure 6 ] is a perspective view of a portion of an assembly viewed from another angle. DETAILED DESCRIPTION

[0031] Reference Figure 1 , shows a twin-flow, twin-body turbojet engine 1, preferably with a high bypass ratio. The turbojet engine 1 typically has a gas generator 2, on either side of which are arranged a low-pressure compressor 4 and a low-pressure turbine 12. The gas generator 2 includes a high-pressure compressor 6, a combustion chamber 8, and a high-pressure turbine 10. In the following text, the terms "front" and "rear" are defined along a direction 14 opposite to the main flow direction of the turbojet's internal gases, which direction 14 is parallel to the turbojet's longitudinal axis 3. Conversely, the terms "upstream" and "downstream" are defined according to the main flow direction of the turbojet's internal gases (referenced by 15).

[0032] The low-pressure compressor 4 and the low-pressure turbine 12 form a low-pressure body and are connected to each other by a low-pressure shaft 11 centered on the axis 3. Similarly, the high-pressure compressor 6 and the high-pressure turbine 10 form a high-pressure body and are connected to each other by a high-pressure shaft 13 centered on the axis 3 and arranged around the low-pressure shaft 11. The shafts are supported by roller bearings (not shown), which are lubricated by being arranged in an oil chamber. The same applies to the fan shaft 17, also called the fan wheel shaft, which is supported by a plurality of roller bearings 19, one of which is in Figure 1 17 . This roller bearing 19 thus supports the fan shaft 17 and includes an axial stop 21 that prevents the shaft 17 and the fan 15 from moving axially upstream relative to the stator section of the turbojet. This roller bearing is therefore the bearing 19 that holds the fan 15 axially in the axial direction and is also known as a thrust bearing. This roller bearing itself is supported by a bearing support 44 that is fixed to the stator section of the turbojet.

[0033] The turbojet engine 1 also includes a single fan 15, located in front of the gas generator 2 and the low-pressure compressor 4, and arranged directly behind the engine's air inlet cone. The fan 15 is rotatable about the axis 3 and is enclosed by a fan housing 9. The fan 15 is driven by a speed reducer 20 via a fan shaft 17, which enables it to rotate at a lower speed than the low-pressure body.

[0034] Furthermore, the turbojet engine 1 defines a main flow path 16 intended for a main flow 16 a to flow through, and a secondary flow path 18 intended for a secondary flow 18 a situated radially outside the main flow, in which the flow coming from the fan is divided.

[0035] Downstream of the fan 15, in the secondary flow path 18, a ring of guide vanes is provided, in this case outlet guide vanes (OGV) 24. These stator vanes 24 connect the outer casing element (here, the outer casing 23) to the inner casing element (here corresponding to the intermediate casing 25), which is arranged axially between the two compressors 4 and 6. Preferably, the bearing support 44 is fixed to the intermediate casing 25, preferably at or near the upstream end of the inner annular structure of the intermediate casing 25.

[0036] Reference Figure 1 and Figure 2 The speed reducer 20 will now be described. It includes an epicyclic gear train. More specifically, the epicyclic gear train includes planetary gears 34 meshing with inner planetary gears 36 (also known as sun gears or inner annular members). Planetary gears 34 also mesh with outer planetary gears 38 (also known as outer annular members). Inner and outer planets 36, 38 are coaxial with the turbojet engine's axis 3. Each planetary gear 34 is mounted so as to rotate freely about a pivot 40, which is integral with a planet carrier 42.

[0037] In this planetary reducer 20, the inner ring 36 is connected in rotation with the low-pressure shaft 11, which forms the input of the gear train. The planet carrier 42 is fixed to the stator part of the turbojet engine, while the outer ring gear 38 is engaged to the fan shaft 17, thus forming the reduction output of the epicyclic gear train.

[0038] The present invention relates to a turbojet engine assembly 50 comprising a fan 15, a speed reducer 20 arranged downstream of the fan, and an intermediate casing 25. Figure 1 and Figure 3 The intermediate casing 25 has an inner annular structure 52 that internally delimits a cavity 54 centered on the axis 3 and that houses at least a portion of the reducer 20. The casing 20 also includes an outer annular structure 56 connected to the inner annular structure by radial arms 58 that are circumferentially spaced apart from one another. The annular space between the inner annular structure 52 and the outer annular structure 56 forms part of the main flow path 16 of the turbojet engine.

[0039] Reference Figures 3 to 6First, it should be noted that the ring 60 of the planet carrier 42 is fixed to the inner annular structure 52 by means of a primary retaining device 62. This comprises a flexible, generally U-shaped half-ring that opens radially outward. The upstream leg 64 of the U-shaped member is fixed (preferably by a circumferential row of bolts) to the ring 60 of the planet carrier. Similarly, the downstream leg 66 of the U-shaped member is fixed (also preferably by a circumferential row of bolts) to a flange 68 of the inner annular structure 52 of the intermediate casing 25. The flange 68 is located at the downstream end of the inner annular structure 52 and projects radially inward, thereby defining the cavity 54 downstream.

[0040] Due to the flexibility of the generally U-shaped main retaining device 62, vibrations from the stator portion of the turbojet engine are partially filtered out and not transmitted to the reducer 20, which advantageously increases the life of the reducer, particularly with regard to tooth wear. During normal operation of the turbojet engine, axial and radial vibrations are essentially filtered out by the generally U-shaped main retaining device 62 and not transmitted to the ring 60 of the planet carrier 42.

[0041] In the event of a failure in the axial retention of the fan 15 (for example due to a failure of the roller bearing 19 or its bearing support 44), there is a risk that the fan will move forward. In order to overcome this disadvantage, one of the features of the invention is the use of a secondary retaining device of the ring 60 of the planetary carrier, which is able to retain the fan axially upstream in the event of a failure such as those described above. In addition, this secondary retaining device 70 is not only configured to limit the axial movement of the ring 60 / fan 15 relative to the internal annular structure 52 of the intermediate casing, but also to limit radial displacement in the event of a significant imbalance of the fan. In addition, this secondary retaining device 70 is also designed to ensure the transmission of the torque of the reducer / fan towards the internal annular structure 52 of the intermediate casing 25. This torque transmission in the circumferential direction can be observed continuously or only in the event of a failure of the type described above.

[0042] To implement the secondary retaining device 70, first protrusions 72a are provided so as to project radially inward from the inner annular structure 52 of the intermediate housing. The first protrusions 72a are circumferentially spaced apart from one another about the axis 3. The first protrusions have the form of cogs or teeth evenly spaced apart in the circumferential direction, for example, six to ten cogs or teeth. The first protrusions 72a are preferably manufactured as a single piece with the entire inner annular structure 52 of the intermediate housing 25 by casting.

[0043] Each first protrusion 72a has a recess 74 formed at its distal end. The recess 74 opens radially inward, axially downstream, and circumferentially in a first circumferential direction (e.g., clockwise in a front view). Facing these three openings, the recess 74 is bounded by three faces: an axial retaining face 76, a radial retaining face 78, and a circumferential retaining face 80. These three adjacent retaining faces 76, 78, and 80 are substantially perpendicular to one another, forming a kind of core angle designed to mate with the second protrusion 72b, thereby forming an integral part of the secondary retaining device 70.

[0044] In practice, the second projections 72 b project radially outwards from the planet carrier annular element 60 and are preferably made in one piece with this same annular element. The second projections 72 b are spaced circumferentially from one another around the axis 3. The second projections also have the form of cogs or teeth, which are provided in the same number as the first projections 72 b and are uniformly spaced in the circumferential direction.

[0045] The second protrusions mate with the first protrusions 72a in pairs such that the distal end of each second protrusion 72b is partially received in the recess 74 of the first protrusion 72a associated with the second protrusion located downstream of the second protrusion.

[0046] The angular extensions of the first and second protrusions 72a, 72b are preferably the same or similar.

[0047] In the normal operating configuration of the turbojet engine, a radial gap R1 is defined between each second protrusion 72b and the radial retaining surface 78 of the associated first protrusion. This gap R1 is very small, for example, between 0.5 mm and 5 mm. Similarly, an axial gap R2 of the same size is defined between each second protrusion 72b and the axial retaining surface 76 of the associated first protrusion. However, preferably, no circumferential gap is provided between each second protrusion 72b and the circumferential retaining surface 80 defining the recess 74, or only a small gap that is smaller than gaps R1 and R2 is provided.

[0048] In the event of a failure in the roller bearing 19 or the bearing support 44 supporting the fan shaft 17, the fan 15 tends to move forward relative to the stator portion of the turbojet engine. This movement initially causes the stator and rotor elements of the reducer 20 to come into contact with each other, dissipating some of the fan's energy. This fan movement is primarily transmitted to the ring 60 of the planetary carrier 42, where its forward movement is quickly stopped by utilizing the axial clearance R2 between the first and second protrusions 72a, 72b, which form an effective cog-like connection after the observed failure. The same applies to the use of radial clearance R1 if such a failure results in a greater fan imbalance, which is still limited by the function of the secondary retaining device 70. Furthermore, the secondary retaining device 70 continues to ensure torque transmission in the event of a failure through the cooperation between the second protrusion 72b and the circumferential retaining surface 80 that delimits the recess 74.

[0049] The invention thus makes it possible to provide a compact and simple solution to the problem of failure of the axial retention of the fan.

[0050] Naturally, various modifications may be made by a person skilled in the art to the invention described merely as a non-limiting example, the scope of the invention being defined by the appended claims.

Claims

1. An assembly (50) for an aircraft turbine engine, comprising a fan (15); a speed reducer (20) driving the fan and located downstream of the fan; an intermediate casing (25), the intermediate casing comprising an inner annular structure (52) delimiting a cavity (54) at least partially housing the speed reducer, the speed reducer comprising an epicyclic gear train equipped with a planet carrier (42), one of the rings (60) of the planet carrier being fixed to the inner annular structure (52) of the intermediate casing by means of primary retaining means (62), It is characterized by: The assembly comprises a secondary retaining device (70) of the planet carrier annular member (60) relative to the inner annular structure (52) of the intermediate housing, the secondary retaining device comprising: - first projections (72a) projecting radially inwards from the inner annular structure (52) of the intermediate casing, the first projections being circumferentially spaced apart from one another around the longitudinal axis (3) of the turbine engine, each first projection (72a) comprising a recess (74) opening radially inwards, opening axially downstream and opening circumferentially in a first circumferential direction and being delimited by three faces, namely an axial retaining face (76), a radial retaining face (78) and a circumferential retaining face (80); - second projections (72b) projecting radially outwards from the annular element (60) of the planetary carrier, said second projections being circumferentially spaced apart from one another around the longitudinal axis (3) of the turbine engine and cooperating in pairs with said first projections (72a) so that each second projection (72b) is partially housed in a recess (74) of a first projection associated with said second projection, located downstream of said second projection.

2. The assembly according to claim 1, characterized in that The assembly is configured such that, in a normal operating configuration of the turbine engine, a radial gap (R1) is defined between each second protrusion (72b) and a radial retaining surface (78) defining a recess (74) of the first protrusion that cooperates with the second protrusion.

3. The assembly according to claim 1, characterized in that The inner annular structure (52) of the intermediate housing (25) and the first protrusion (72a) are made into an integral piece.

4. The assembly according to claim 1, wherein The primary retaining device (62) comprises a U-shaped flexible half-segment annular member opening radially outward, the upstream leg (64) of the U-shaped member being fixed to the annular member (60) of the planet carrier, and the downstream leg (66) of the U-shaped member being fixed to a radially inwardly projecting flange (68) of the inner annular structure (52) of the intermediate housing.

5. The assembly according to claim 1, characterized in that The number of the first protrusions (72a) is between three and fifteen.

6. The assembly according to claim 1, characterized in that The epicyclic gear train comprises an outer ring gear (38) meshing with the planet gears (34) of the planet carrier, the outer ring gear (38) being connected to the fan shaft (17).

7. The assembly according to claim 6, characterized in that The fan shaft (17) is supported by a bearing (19) which ensures axial retention of the fan shaft in the direction from downstream to upstream.

8. The assembly according to claim 7, characterized in that The bearing (19) is supported by a bearing support (44) which is attached to the stator portion of the assembly.

9. The assembly according to claim 1, wherein The intermediate casing (25) also includes an outer annular structure (56) connected to the inner annular structure (52) by radial arms (58), the annular space between the inner and outer annular structures forming part of the main flow path (16) of the turbine engine.

10. The assembly according to claim 1 or 2, characterized in that The assembly is configured such that, in a normal operating configuration of the turbine engine, an axial clearance (R2) is defined between each second protrusion (72b) and an axial retaining surface (76) delimiting the recess (74).

11. The assembly according to claim 1 or 2, characterized in that The assembly is configured such that, in a normal operating configuration of the turbine engine, a circumferential gap is defined between each second projection (72b) and a circumferential retaining surface (80) delimiting the recess.

12. The assembly according to claim 5, characterized in that The number of the first protrusions (72a) is between six and ten.

13. The assembly according to claim 8, wherein The bearing support is attached to an inner annular structure (52) of the intermediate housing (25).

14. An aircraft turbine engine (1) comprising an assembly (50) according to claim 1.

15. Aircraft turbine engine according to claim 14, characterized in that The turbine engine is a twin-flow, twin-body turbojet engine.

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