Turbomachine assembly incorporating a variable propeller blade pitch control device

Abstract

The invention relates to an aircraft turbomachine assembly comprising a propeller constituting a non-fan, the turbomachine comprising a cyclic pitch control device (50) for the propeller blades, the cyclic pitch control device (50) comprising at least two actuators (52), each actuator being fixed to a stator (70), the stator (70) being rotationally fixed relative to a bearing support supporting the rotating shaft, the stator (70) comprising a rim (70C) connected to a hub (70D) by a set of spokes (70E), the rim (70C) being fixed to the bearing support, the hub being rotationally mounted about the rotating shaft, at least two sets of two adjacent spokes (70E) being configured for the attachment of an actuator (52) of the variable pitch control device, each actuator (52) being at least partially inserted between the two corresponding adjacent spokes (70E). Figure 3.

Description

Title of the invention: Turbomachine assembly comprising a variable propeller blade pitching device. Technical field

[0001] The present invention relates to the field of aeronautics, and more particularly to a propulsion assembly comprising a double-flow turbomachine and an unfaired fan. State of the art

[0002] Aircraft are known that are propelled by at least one propulsion unit comprising a turbomachine, such as a turbofan engine. Each propulsion unit is attached to the aircraft by a pylon located generally under or on a wing, or at the level of the aircraft fuselage. A turbofan engine primarily comprises a gas generator and a fan.

[0003] The fan may be enclosed, in which case the turbojet is housed in a nacelle. The fan may also be unenclosed, as is the case with turbojets known as "open fan" or "unducted single fan".

[0004] The gas generator includes, in particular, from upstream to downstream with respect to the direction of gas flow, a rectifier, a low pressure compressor and a high pressure compressor.

[0005] During operation, an airflow is accelerated by the fan, then splits into a primary flow and a secondary flow. The primary flow flows in a primary gas circulation channel through the turbojet's gas generator.

[0006] In the case of a shrouded fan, the secondary flow runs in a secondary channel surrounding the gas generator. The secondary channel is delimited, radially inward, partly by an internal structure of the nacelle that encloses the gas generator, and radially outward, partly by an external structure of the nacelle that surrounds the turbojet. A portion of the secondary channel is further delimited radially outward by a fan casing surrounding the fan, and by an intermediate casing located downstream of the fan casing. In the case of an unshrouded fan, the secondary flow is also generated by the fan, but is open and flows around the gas generator.

[0007]

[0008] In the case of an unshod fan, it may be advantageous to provide a variable fan blade pitching system. In such a case, the fan blades are movable around a radial axis, and a mechanism allows The adjustment of each blade's position allows for varying the blade's angle relative to a radial plane. Furthermore, a cyclic blade adjustment system can be implemented.

[0009] The known fan blade pitch control systems have complex architectures, not allowing for more than one actuator.

[0010] There is therefore a need for an architecture that allows for a control system comprising several actuators.

[0011] The objective of the present invention is to propose a turbomachine that meets this need. Description of the invention

[0012] To this end, the invention relates to an aircraft turbomachine assembly comprising a propeller constituting an unfaired fan generating a secondary flow, the turbomachine comprising a gas generator through which a gas flow, called primary flow, the propeller being driven in rotation about a central axis of the turbomachine by a rotating shaft, the turbomachine comprising a cyclic pitch control device for the propeller blades, the cyclic pitch control device comprising at least two actuators, each actuator being fixed to a stator, the stator being fixed in rotation relative to a bearing support supporting the rotating shaft, the stator comprising a rim connected to a hub by a set of spokes, the rim being fixed to the bearing support, the hub being mounted for rotation about the rotating shaft, at least two sets of two adjacent spokes being configured for the attachment of an actuator of the variable pitch control device,each actuator being at least partially inserted between the two respective adjacent beams.

[0013] Thus, by providing a fixed wheel, or stator having adjacent spokes configured by the attachment of an actuator of the variable pitching device, it is possible to provide for the attachment of several actuators. Each actuator can thus be controlled individually so as to apply a non-uniform pitching to all the blades of the propeller.

[0014] The turbomachine assembly according to the invention may include one or more of the following optional features, considered alone or in all possible combinations.

[0015] According to one characteristic, each actuator is mobile in rotation relative to the spokes to which it is fixed, at least along an axis of tangential direction relative to the central axis.

[0016] According to one characteristic, each actuator is mobile in rotation relative to the spokes to which it is fixed, along a first axis of tangential direction and along a second axis of radial direction relative to the central axis.

[0017] According to one characteristic, each actuator is connected to the spokes to which it is attached by means of one or more axes.

[0018] According to one characteristic, the axis or axes allowing the fixing of an actuator are ball-jointed or ball-rolling axes.

[0019] According to one feature, each actuator of the cyclic timing device is a linear actuator connected to a respective actuating lever, each actuating lever being connected to a motion transfer bearing.

[0020] According to a characteristic, each actuator can be controlled individually.

[0021] According to one feature, the motion transfer bearing is mounted on a ball joint.

[0022] According to one characteristic, at least two adjacent radii define a space allowing the passage of easements.

[0023] The invention also relates to an aircraft comprising at least one propulsion unit including a turbomachine assembly conforming to that defined above. Brief description of the drawings

[0024] [Fig-1] Fig. 1 represents an aircraft equipped with a propulsion system conforming to the invention.

[0025] [Fig.2] Fig.2 represents a schematic cross-sectional view of a propulsion assembly according to the invention.

[0026] [Fig.3] The [Fig.3] is a schematic cross-sectional view of the cyclic timing device of the propulsion assembly of the [Fig.1].

[0027] [Fig.4] The [Fig.4] is a partial perspective view of the cyclic timing device. Detailed description

[0028] Figure 1 represents an aircraft 100, in this example an airplane, equipped with two propulsion units 10, namely one propulsion unit 10 per wing 101, with only one propulsion unit 10 and one wing 101 shown in Figure 1. According to one embodiment, the aircraft 100 can be equipped with more than one propulsion unit 10 per wing 101, each wing 101 having the same number of propulsion units 10. The reference symbol "A" designates the axis of the fuselage 102 of the aircraft 100. The propulsion unit 10 can be configured to propel the aircraft 10 at a cruising speed between Mach 0.7 and Mach 0.9.

[0029] Figure 2 shows a schematic cross-sectional view of the propulsion assembly 10, according to plane II of Figure 1. The propulsion assembly 10 extends along a central axis X, and comprises a propulsion module 20, a gas generator 30, and, in the example, a speed reduction device 40 and a cyclic pitch control device 50. When the propulsion unit 10 is mounted on the aircraft 100, the central axis X is not necessarily parallel to the axis A.

[0030] The propulsion module 20 has a propeller 22, constituting an unfaired fan. The propeller is provided with a plurality of blades 22A. The propulsion module also includes a stator 24 provided with a plurality of blades 24A, and a propeller shaft 26 configured to drive the propeller 22 in rotation. The propeller shaft 26 can extend along the central axis X. The blades 22A of the propeller 22 can be made entirely or partially of composite material. The blades 24A of the stator 24 can be made entirely or partially of composite material. The propeller 22 can comprise between 8 and 14 blades 22A and the stator 24 can comprise the same or fewer number of blades 24A, for example, between 8 and 14 blades 24A. The setting of the 24A blades of the 24 rectifier can be fixed or variable.

[0031] The gas generator 30 has a drive shaft 33A. The drive shaft can extend along the central axis X. The propeller shaft 26 can be coaxial with the drive shaft 33A, and their respective axes of rotation can coincide with the central axis X of the propulsion assembly 10. This allows for an annular air inlet within the gas generator 30 coaxial with the central axis X, thanks to which the outer casing of the gas generator has a relatively simple shape and exhibits a certain rotational symmetry, which tends to reduce possible airflow disturbances. In this example, the gas generator 30 comprises, from upstream to downstream, the gases flowing within the propulsion assembly 100, from upstream to downstream, a compressor 32 (or compressor section 32), a combustion chamber 34, and a turbine 36 (or turbine section 36).

[0032] The gas generator 30 may be of the twin-spool type and comprise a low-pressure spool 30A and a high-pressure spool 30B. The low-pressure spool 30A may comprise a low-pressure compressor 32A rotationally coupled to a low-pressure turbine 36A via a low-pressure shaft 33A, which may form the drive shaft of the gas generator 30. The high-pressure spool 30B may comprise a high-pressure compressor 32B located downstream of the low-pressure compressor 32A and upstream of the combustion chamber 34, and a high-pressure turbine 36B located downstream of the combustion chamber 34 and upstream of the low-pressure turbine 36A, and rotationally coupled to the high-pressure compressor 32B via a high-pressure shaft 33B. The compressor 32 of the gas generator 30 may comprise the low-pressure and high-pressure compressors 32A and 32B.The turbine 36 of the gas generator 30 can comprise the low-pressure and high-pressure turbines 36A and 36B. The low-pressure and high-pressure shafts 33A and 33B can be coaxial. The high-pressure shaft 33B can receive a portion of the low-pressure shaft 33A. According to one variant, the low-pressure shaft 33A and high-pressure shaft 33B can be... Co-rotating, i.e., configured to rotate relative to each other in the same direction around the central axis X. According to another variant, the low-pressure shafts 33A and high-pressure shafts 33B can be contra-rotating, i.e., configured to rotate relative to each other in opposite directions around the central axis X. The rotational speed of the low-pressure shaft 33A can be lower than the rotational speed of the high-pressure shaft 33B.

[0033] According to an alternative (not shown), the propulsion unit may be of the three-shaft type. The turbine 36 may include an intermediate turbine arranged axially between the high-pressure turbine 36B and the low-pressure turbine 36A and configured to drive an intermediate compressor arranged axially between the low-pressure compressor 32A and the high-pressure compressor 32B via an intermediate shaft. The intermediate shaft may be located between the low-pressure shaft 33A and the high-pressure shaft 33B. The intermediate shaft and the low-pressure shaft 33B may rotate co- or counter-rotating with respect to each other.

[0034] Each compressor 32A, 32B and turbine 36A, 36B may comprise a plurality of stages, each stage comprising an impeller, respectively 32AA, 32BA, 36AA, 36BA, movable in rotation about the central axis X (or rotor) and an impeller, respectively 32AB, 32BB, 36AB, 36BB, fixed about the central axis X (or stator). In this example, the low-pressure compressor 32A may have at least 2 stages and at most 5 stages, for example 2 stages, the high-pressure compressor 32B may have between 8 and 11 stages (only two stages being shown for clarity of the figure), the high-pressure turbine 36B may have 2 stages, and the low-pressure turbine 36A may have between 3 and 8 stages (only two stages being shown for clarity of the figure). A rectifier 37, or fixed paddle wheel rotating around the central axis X, can be arranged downstream of the combustion chamber 34 and upstream of the high-pressure turbine 36B.

[0035] A speed reduction device 40 can indirectly couple the drive shaft 33A in rotation with the propeller shaft 26. The speed reduction device 40 can be configured to drive the propeller shaft 26 at a rotational speed lower than the rotational speed of the drive shaft 33A. The drive shaft 33A connects the low-pressure turbine 36A (or the low-pressure housing 30A) to an inlet of the speed reduction device 40, while the propeller shaft 26 connects an outlet of the speed reduction device 40 to the propeller 22. The propeller 22 is therefore driven by the low-pressure turbine 36A (or the low-pressure housing 30A) via the drive shaft 33A (or low-pressure shaft), the speed reduction device 40, and the propeller shaft 26. In this example, the speed reduction device 40 can be arranged, considered along the central axis X, between an upstream end of the drive shaft 33A and a downstream end of the propeller shaft 36.

[0036] For example, the speed reduction device 40 may be an epicyclic gear train reduction device, for example of the "epicyclic" or "planetary" type, according to the terminology sometimes used by those skilled in the art. Such a mechanism may comprise one stage, two stages, or more than two stages.

[0037] In operation, an airflow F (see [Fig.2]) entering the propulsion assembly 10 passes through the propeller 22 and is then divided into a primary flow Fl and a secondary flow F2, which circulate from upstream to downstream within the propulsion assembly 10.

[0038] The primary airflow Fl flows in a channel called the "primary channel" inside the gas generator 30, sometimes also called the primary body, passing successively through the low-pressure compressor 32A, the high-pressure compressor 32B, the combustion chamber 34, the high-pressure turbine 36B, the low-pressure turbine 36A, and then through the outlet nozzle. The expansion of the combustion gases downstream of the combustion chamber 34 within the turbine 36 provides the energy to drive the rotation of the high-pressure and low-pressure turbines 36B, 36A, and therefore the shafts 33A and 33B.

[0039] The secondary airflow F2 flows through the rectifier 24, then along the gas generator 30, outside the gas generator 30. This secondary airflow F2 provides, by reaction, the majority of the thrust generated by the propulsion assembly 10. The secondary airflow F2 can also be used to cool the gas generator 30 from the outside.

[0040] The cyclic timing device for the propeller blades 50 is described in more detail with reference to Figures 3 and 4. The cyclic timing device 50 comprises at least two actuators 52 distributed around the central axis X, and for example, four actuators 52 (only one actuator 52 being visible in Figures 3 and 4). The actuators 52 are, in the example, cylinders 52. The cylinders 52 may be of the "single chamber" type. According to an alternative embodiment not shown, the cyclic timing device 50 may comprise exactly three cylinders 52, at least one of the cylinders 52, or even all of the cylinders 52, being of the "double chamber" type. The cylinders 52 can be regularly distributed circumferentially around the central axis X, for example every 90° in the case where the cyclic stabilization device 50 comprises exactly four cylinders 52.The cylinders 52 can be configured to adopt a position P within a total collective shimming stroke CT, and to allow a cyclic shimming stroke CC of between ±40 mm, for example between ±20 mm, for example between ±16 mm, for example between ±9.6 mm, around said position P. The cylinders 52 are, for example, hydraulic cylinders. Each cylinder 52 can be supplied with pressurized oil via a supply line 51, specific to each cylinder. 52 and independent for each cylinder 52. These independent supply lines 51 can extend via arms 55A of the intermediate housing 55. In other words, the supply lines 51 can bypass the speed reduction device 40 and not extend through it. The pressurized oil supply circuit for the cylinders 52 can be independent and separate from the oil supply circuit for the speed reduction device 40. Such a configuration allows for a degree of modularity and facilitates maintenance as well as the assembly / disassembly of the cyclic timing device 50. Furthermore, the independent supplies for the cylinders 52 reduce the risk of failure of the entire cyclic timing device 50 and prevent a risk of collective failure of the operability of the cylinders 52.In the event of a failure of one of the cylinders 52 or its power supply circuit, three cylinders 52 remain operational, which is sufficient to control the chocking ring 54 described below.

[0041] The cylinders 52 may each have a cylinder 52A mounted on a stator 70 of the propulsion assembly 10, for example on a bearing support housing 70A of the propeller shaft 26, as described in detail below. The cylinders 52 may each have a rod 52B sliding relative to the cylinder 52A, each rod having, for example, a distal end sliding over a total stroke C equal to the total collective trim stroke CT plus the cyclic trim stroke CC.

[0042] The cyclic shimming device 50 may include a shimming ring 54 which is stator-shaped like the cylinder 52 and which is ball-jointed about the central axis X and slides parallel to the central axis X. According to a variant 50' shown in [Fig.3], the shimming ring 54 may be mounted to slide on the propeller shaft 26.

[0043] The shim ring 54 may include a ball-joint ring 54A mounted to slide about the central axis X on the propeller shaft 26 (see [Fig. 3]), for example, by means of rollers 53. These rollers 53 may be configured to allow the ball-joint ring 54A to slide about the central axis X and to rotate about the central axis X relative to the propeller shaft 26 (see [Fig. 3]), and thus form a sliding bearing. The ring 54A has a convex, ball-joint-shaped external surface 54A1.

[0044] The shim ring 54 may include an inner ring 54B, for example forming an internal stator plate, arranged radially outside the ball-joint ring 54A. The inner ring 54B has a concave inner face 54B1 configured to cooperate by complementary shape with the convex outer surface 54A1 of the ball-joint ring 54A. This configuration is an example of providing the ball-jointed character about the central axis X of the shim ring 54. The inner ring 54B is connected to a plurality of actuating levers 54B2, extending axially and / or radially. Each actuating lever 54B2 is connected to an actuator 52 respective. The inner ring 54B may comprise an outer face 54B3 assembled with a radially internal portion of a rolling bearing 56.

[0045] The shim ring 54 may include an outer ring 54C, for example forming an outer rotor plate, arranged radially outside the inner ring 54B. The outer ring 54C may include an inner face 54C1 assembled with a radially external portion of the bearing 56. In other words, the bearing 56 may be arranged radially between the inner ring 54B and the outer ring 54C. The inner ring 54B and the outer ring 54C can thus be rotationally decoupled about the central axis X. In other words, the inner ring 54B and the outer ring 54C are rotationally mobile relative to each other about the central axis X. The outer ring 54C may include an outer face 54C2 provided with a plurality of attachments 54C3, for example, eye brackets, to ensure a mechanical connection between the outer ring 54C and each of the blades 22A. For example, the outer ring 54C may include as many attachments 54C3 as there are blades 22A.

[0046] An oil circuit not shown supplies oil to the sliding bearing formed by the rollers 53, the ball joint formed by the rings 54A and 54B, and the bearing 56 of the shim ring 54.

[0047] The shim ring 54 may have an internal radius RR, which corresponds to the internal radius of the ball joint ring 54A, between 150 mm and 450 mm, for example 225 mm. The shim ring 54 may be ball jointed over an angular range C2 between ± 30°, for example between ± 15°, for example between ± 10.5°, for example approximately ± 4.0°.

[0048] Each blade 22A is pivotally mounted about its radially extending Z axis on the hub 22B of the propeller 22, so as to allow adjustment of the pitch angle of each of the blades 22A (see arrow AC in [Fig. 3]), i.e. of the aerodynamic profile 22A2 of each of the blades 22A. More particularly, in this example, the foot 22A3 of each of the blades 22A is mounted on the hub 22B of the propeller 22 via a bearing 22C.

[0049] The foot 22A3 of each of the blades 22 can be provided with a pitch control lever 22D, configured to rotate the blade 22 about its Z-axis. The pitch control lever 22D can be connected to the pitch control ring 54 via a connecting rod 25. The connecting rod 25 can be connected to the pitch control ring 54 via a pivot joint or via a ball joint. The connecting rod 25 can be connected to the pitch control lever 22D via a pivot joint or via a ball joint.

[0050] When the propeller 22 rotates, the blades 22A drive, via their respective lever 22D and connecting rod 25, the outer ring 54C in rotation around the central axis X. The inner ring 54B is rotationally decoupled from the outer ring 54C via the bearing 56, and rotationally coupled around the central axis X to the stator 70 via the connecting rods. The actuation device 54B2 and the actuators 52 remain stationary while rotating about the central axis X. Depending on the inclination (provided by an angular stroke less than or equal to the angular range C2) of the inner ring 54B around the ball-joint ring 54A, imposed by the actuators 52, the outer ring 54C follows the inclination imposed on the inner ring 54B during its rotation about the central axis X. Thus, the pitch of each of the blades 22B varies during the rotation of the propeller 22. At each predefined angular position about the central axis X, the blades 22A, during their successive passage through these predefined angular positions during the rotation of the propeller 22, acquire a pitch associated with each of these predefined angular positions. This constitutes an example of a cyclic pitching device.

[0051] The cyclic timing device 50 may include a pressure accumulator 58 (see [Fig. 2]) configured to provide safety control energy to bring all the blades to the feathered position. For example, the pressure accumulator 58 may be a vacuum reservoir configured to purge oil from the cylinders 52 so as to bring them to a neutral configuration corresponding to the feathered position of the blades 22A. For example, the pressure accumulator 58 may be disposed in an enclosure E housing the speed reduction device 40.

[0052] According to the invention, each actuator 52 is fixed to the stator 70. The stator 70 is rotationally fixed relative to a bearing support 70B supporting the propeller shaft 26. The stator 70 has a rim 70C connected to a hub 70D by a set of spokes 70E. The rim 70C is fixed to the bearing support 70B, the hub 70D being rotationally mounted around the propeller shaft 26. At least two sets of two adjacent spokes 70E are configured for attaching an actuator 52 of the variable timing device, each actuator 52 being at least partially inserted between the two respective adjacent spokes 70E. The gap between the adjacent spokes 72E configured for attaching an actuator 52 is compatible with the insertion of the cylinder 52A between these two adjacent spokes 70E. Each actuator 52 is thus arranged between two adjacent spokes 72E.In the example, the gap between two adjacent spokes 72E intended to accommodate an actuator 52 is greater than the gap between two adjacent spokes not intended for the attachment of an actuator 52. Thus, in the example, the angular distribution of the spokes 70E is not uniform.

[0053] The configuration of the stator 70 allows two functions to coexist effectively, namely a function of transmitting the forces from the propeller shaft 26 to the bearing support 70B, and a function of housing the actuators and transmitting the forces related to the operation of these actuators.

[0054] Advantageously, each actuator 52 is mounted to rotate freely relative to the spokes 70E to which it is fixed, along at least one axis, in particular a steering axis tangential, for example via axes 52C. Advantageously, each actuator 52 is mounted to rotate freely about a first tangential axis and about a second radial axis. Each actuator 52 can be fixed to the stator 70, for example, via a joint or a ball joint.

[0055] In operation, all the cylinders 52 can be controlled simultaneously and according to the same command, so that the shim ring 54 is moved along the central axis X (see double arrow Cl in [Fig.3]), thereby changing the shim position of all the blades 22A to achieve an identical shim angle for all the blades 22A. This allows the collective shim positioning of the blades 22A to be controlled and regulated. Each cylinder 52 can also be controlled independently of the other according to a setpoint which is specific to it and different from that of the other cylinders, so that the ball-jointed shim ring 54 pivots around a radial direction (see for example the double arrow associated with the angular range C2 on [Fig.3]), by which means the shim of each of the blades 22A is modified in a specific way according to its angular position during the rotation of the propeller 22. This makes it possible to control and regulate the cyclic shim of the blades 22A.The cyclic pitch angle can be within ±30°, e.g., ±6° around the collective pitch position. For example, the angular range of the ball joint C2 of the pitching ring 54 can be configured so that the cyclic pitch angle is within ±30°, e.g., ±6° around the collective pitch position. For example, the cyclic pitch stroke CC of each of the cylinders 52 can be configured so that the cyclic pitch angle is within ±30°, e.g., ±6° around the collective pitch position. In this example, the pitching device 50 can allow both cyclic and collective pitching.

Claims

Demands

1. An aircraft turbomachine assembly comprising a propeller (22) constituting an unfaired fan generating a secondary flow (F2), the turbomachine comprising a gas generator (30) through which a gas flow, referred to as the primary flow (F1), passes, the propeller (22) being driven in rotation about a central axis (X) of the turbomachine by a rotating shaft (26), the turbomachine comprising a cyclic timing device (50) for the blades (22A) of the propeller (22), the cyclic timing device (50) comprising at least two actuators (52), each actuator (52) being fixed to a stator (70), the stator (70) being rotationally fixed relative to a bearing support (70B) supporting the rotating shaft (26), the stator (70) comprising a rim (70C) connected to a hub (70D) by a set of spokes (70E), the rim (70C) being fixed to the bearing support (70B), the hub (70D) being mounted for rotation around the rotating shaft (26),at least two sets of two adjacent spokes (70E) being configured for the attachment of an actuator (52) of the variable positioning device, each actuator (52) being at least partially inserted between the two respective adjacent spokes (70E).

2. Turbomachine assembly according to the preceding claim, wherein each actuator (52) is rotationally movable relative to the radii (70E) to which it is fixed, at least along one axis of tangential direction relative to the central axis (X).

3. Turbomachine assembly (10) according to the preceding claim, wherein each actuator (52) is rotationally movable relative to the radii (70E) to which it is fixed, along a first tangential direction axis and along a second radial direction axis relative to the central axis (X).

4. Turbomachine assembly (10) according to any one of claims 2 and 3, wherein each actuator (52) is connected to the spokes (70E) to which it is fixed by means of one or more shafts (52c).

5. Turbomachine assembly (10) according to the preceding claim, wherein the shaft(s) (52C) ​​for attaching an actuator (52) are ball-jointed or spherical shafts.

6. Turbomachine assembly according to any one of the preceding claims, wherein each actuator (52) of the cyclic timing device (50) is a linear actuator connected to a lever

7.

8.

9.

10. respective actuation levers (54B2), each actuation lever (54B2) being connected to a motion transfer bearing (54). Turbomachine assembly according to any one of the preceding claims, wherein each actuator (52) can be individually controlled. Turbomachine assembly according to claim 6 or according to claims 6 and 7, wherein the motion transfer bearing (54) is mounted on a ball joint. Turbomachine assembly according to any one of the preceding claims, wherein at least two adjacent radii define a space allowing passage of servitudes (51). Aircraft (100) comprising at least one propulsion unit (10) comprising a turbomachine assembly conforming to one of the preceding claims.

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