Active radial or axial play control system for an abradable seal of an aeronautical turbomachinery element

Abstract

This application relates to a turbomachine assembly (1) centered on an X-axis, comprising: a rotor, a stator, a seal (22) comprising a stator seal portion (25) mechanically fixed to the stator and a rotating seal portion (26) mechanically fixed to the rotor, the assembly comprising at least one heating element (2), preferably resistive, disposed opposite or fixed to a radially external surface (252) of the stator seal portion (25), and in that each heating element (2) is configured to heat the radially external surface (252) of the stator seal portion (25). Figure for the abstract: Fig. 3

Description

Title of the invention: Active control system for radial or axial clearances for an abradable seal of an aeronautical turbomachinery component technical field

[0001] This application relates to aircraft turbomachinery. More particularly, it relates to radial or axial clearance control systems at a sealing joint in a turbomachine element such as a compressor or turbine. STATE OF THE ART

[0002] An aeronautical propulsion system generally comprises, from upstream to downstream in the direction of gas flow, a blower section, a compressor section which may include a low pressure compressor and a high pressure compressor, a combustion chamber and a turbine section which may include a high pressure turbine and a low pressure turbine.

[0003] When the propulsion system is in operation, the high-pressure compressor is driven in rotation by the high-pressure turbine via a high-pressure shaft. The blower and, where applicable, the low-pressure compressor are driven in rotation by the low-pressure turbine via a low-pressure shaft.

[0004] A compressor or turbine stage of an aeronautical turbomachine comprises a plurality of fixed blades and a plurality of moving blades. In the case of a turbine, a fixed casing surrounds the blades so as to define a flow channel through which hot air from the combustion chamber of the turbomachine flows.

[0005] With reference to [Fig. 1], certain turbines 1' further comprise, within a cavity 17' arranged axially between a drive shaft 10' and discs 16' on which movable blades 4' are mounted, a sealing gasket 22' for sealing the cavity 17'. Such a gasket comprises a stator gasket portion 25', fixed relative to the drive shaft X and integral with a turbine stator, and a rotating gasket portion 26' configured to rotate integrally with the movable blades. The stator gasket portion further comprises an abradable coating 27' opposite the rotating gasket portion, designed to be eroded by scuffs on the rotating gasket portion.

[0006] The cavity 17' is designed to allow the intake of air 12' that is cooler than the air located in the flow channel 18' of the turbine, as well as to allow an outlet air 14', sometimes called "purge", from cavity 17' to flow vein 18'.

[0007] During the operation of the turbomachine, the temperature variation in the duct according to the engine speed affects the thermal expansion or contraction behavior of the rotor and stator of the turbine.

[0008] When the turbomachine is shut down following a flight, the difference in thermal expansion behavior between the turbine stator and the rotor further reduces the radial clearances between them, as the stator cools down faster than the rotor. Thus, when the turbomachine is shut down, initially, the diameter of a radially internal surface of the stator seal portion 25' of the seal—and therefore its distance from the engine axis X—decreases due to this cooling, while the distance between the rotating seal portion 26' of the seal and the engine axis X decreases less rapidly. Therefore, for a few hours, contact may exist between the contact edges of the rotating seal portion 26' of the seal and the abradable coating 27' of the stator seal portion 25', which helps prevent the turbomachine from restarting, a phenomenon known as "rotor lock."It is therefore necessary to wait for the 26' rotary joint portion of the seal to retract sufficiently to reintroduce radial clearance between it and the 25' stator joint portion. This prevents the immediate restart of the turbomachine, and thus the initiation of a new flight phase, for several hours. This is detrimental to airlines, which must then schedule relatively long downtimes between successive flights and can therefore operate a limited number of flights in a given day. As for aircraft engine manufacturers, they can partially mitigate this rotor lock-up phenomenon by incorporating this reduction in radial clearance into turbine design, but this requires larger radial clearances between the stator joint portion and the rotary joint portion, which degrades turbine efficiency during all flight phases.

[0009] Note that axial contact between the stator seal portion and the rotating seal portion can also contribute to rotor lockup, as this clearance is also due to the difference in thermal expansion behavior of these components. Such axial contact can damage the seal. During a turbomachine restart, friction occurs due to this axial contact, which can exacerbate the rotor lockup phenomenon.

[0010] A solution to this problem of rotor stalling at startup implemented in the prior art is to provide, prior to starting the turbomachine, for a lapping of the abradable coating 27' of the static seal portion 25'. This lapping creates a radial clearance between the rotating seal portion 26' and this coating 27', thus allowing the free rotation of the rotating joint portion 26' and the start-up of the turbomachine. However, although implementing such a break-in procedure reduces the waiting time required to start the turbomachine without rotor lock-up, the break-in itself requires a certain amount of time, potentially several hours. Furthermore, breaking in a portion of the abradable coating 27' is undesirable: this coating is designed to be eroded during the various phases of flight to obtain the smallest possible radial clearance between the rotating joint portion and the stator joint portion without damaging the stator joint portion, and breaking in would cause premature wear of this coating.

[0011] Due to differences in thermal behavior between the stator seal portion 25' and the rotating seal portion 26', the seal portions move relative to each other along the axial direction during the various phases of flight or engine shutdown. Consequently, a groove formed by the scuffs of the rotating seal portion in the abradable coating of the stator seal portion, as flights progress, develops a significant axial extent, which can compromise the sealing of the seal and that of the cavity 17'. This can promote the return of hot air from the flow stream into the cavity 17, in which case the rotor blade discs 16' are not ventilated but heated. This results in a reduced service life for these discs 16'. EXPOSED

[0012] One purpose of the present disclosure is therefore to enable control of radial and axial clearances between a stator seal portion and a rotating seal portion of a sealing joint in a turbomachine element such as a compressor or turbine, so as to limit a loss of sealing of the joint and of a cavity whose sealing is ensured by the joint.

[0013] Another objective of the present disclosure is to provide an active control system for radial and axial clearances in a turbomachine element that makes it possible to reduce the downtime required between two successive flights to allow a restart of the turbomachine.

[0014] Another purpose of the present disclosure is to prevent a rotor blockage phenomenon at the start-up of the turbomachine without requiring the running-in of an abradable coating provided on an internal surface of the fixed casing, between it and the moving blades.

[0015] To this end, a turbomachine assembly centered on an X-axis is proposed according to a first aspect, comprising: - a rotor, - a stator, - a sealing gasket comprising a stator gasket portion mechanically fixed to the stator and a rotating gasket portion mechanically fixed to the rotor,

[0016] the turbomachine assembly comprising at least one heating element, preferably resistive, arranged opposite or attached to a radially external surface of the stator joint portion and in that each heating element is configured to heat the radially external surface of the stator joint portion.

[0017] The use of heating elements to heat the radially external surface of the stator seal portion allows for the selective increase of the internal diameter of this portion, thereby increasing the radial and / or axial clearance between a radially internal surface of the stator seal portion and a radially external surface of the rotating seal portion, thus preventing rotor lock-up during turbomachine startup. This makes it possible to start the turbomachine quickly following a first flight, eliminating a waiting period required, for state-of-the-art turbomachines, between this first flight and a second flight to allow for rotor cooling, or reducing the duration of such a waiting period.Furthermore, the improved sealing of the cavity, made possible by the active control system, prevents the rotor blade discs from heating up due to hot air returning from the flow stream to the cavity, and consequently increases the lifespan of the discs.

[0018] According to one embodiment, the assembly is a turbine or a compressor.

[0019] According to one embodiment, the assembly includes a battery configured for powering each heating element, and preferably including an alternator suitable for being mechanically driven by a drive shaft and configured to supply the battery with electrical energy.

[0020] According to one embodiment, the alternator includes first splines complementary to second splines of the drive shaft, allowing a mechanical coupling between the alternator and the drive shaft.

[0021] According to one embodiment, the assembly comprises at least two heating elements, preferably four resistive heating elements arranged at distinct axial positions.

[0022] According to one embodiment, the assembly includes a controller configured to control a variation of electrical power transmitted to each heating element.

[0023] According to one embodiment, a radially internal surface of the stator seal portion comprises an abradable coating positioned radially opposite a lip of the rotating seal portion

[0024] According to one embodiment, the rotating joint portion comprises between one and ten licks, preferably between four and six licks.

[0025] According to one embodiment, each heating element is arranged radially opposite a lick.

[0026] The present disclosure relates, according to a second aspect, to a method of actively controlling a clearance between, on the one hand, a radially internal surface of the stator seal portion of the sealing joint of a turbomachine assembly as defined above and, on the other hand, a radially external surface of the rotating seal portion of the sealing joint, the active control method comprising a step of heating the radially external surface of the stator seal portion so as to increase the clearance between the radially external surface of the rotating seal portion and the radially internal surface of the stator seal portion.

[0027] According to one embodiment, the method includes a step of modulating an electrical power transmitted to the heating element so that the radial clearance between the radially external surface of the rotating joint portion and the radially internal surface of the stator joint portion reaches a predefined clearance value.

[0028] According to one embodiment, the method includes a step of collecting information by a controller to determine a gap between, on the one hand, the radial clearance between the radially external surface of the rotating joint portion and the radially internal surface of the stator joint portion and, on the other hand, the predefined clearance value.

[0029] According to one embodiment of the method, the step of modulating the electrical power transmitted to the heating element is implemented so as to maintain the radial clearance between the radially external surface of the rotating joint portion and the radially internal surface of the stator joint portion constant during a turbomachine start-up phase prior to take-off.

[0030] The present disclosure relates, according to a third aspect, to a turbine or turbomachine compressor comprising a turbomachine assembly as defined above. DESCRIPTION OF THE FIGURES

[0031] Fig. 1 schematically represents a conventional low-pressure turbomachine turbine.

[0032] Figure 2 schematically represents a turbomachine comprising a low pressure turbine equipped with an active radial and / or axial play control system as proposed.

[0033] Fig. 3 schematically represents the low-pressure turbine of the turbomachine shown in Fig. 2, according to a first embodiment.

[0034] Fig. 4 schematically represents a first embodiment of a sealing joint for the low-pressure turbine of Fig. 3.

[0035] Fig. 4 schematically represents a second embodiment of a sealing joint for the low-pressure turbine of Fig. 3.

[0036] Throughout the figures, similar elements are designated by identical reference numerals. DETAILED DESCRIPTION OF IMPLEMENTATION METHODS

[0037] Figure 2 represents a turbomachine 17 with axis X. The turbomachine 17 comprises a turbomachine assembly, selected from a turbine or a compressor. In what follows, and for the sake of clarity, we will take as an example a turbomachine assembly that is a low-pressure turbine 1. However, the description that follows applies equally to any other turbine or to a turbomachine compressor.

[0038] The turbomachine includes the low-pressure turbine 1 and a low-pressure compressor 102, mechanically driven by the low-pressure turbine 1 via a low-pressure shaft 10. The turbomachine 17 may further include a blower 101, a high-pressure compressor 103, a combustion chamber 104, and a high-pressure turbine 105, the high-pressure turbine 105 mechanically driving the high-pressure compressor 103 via a high-pressure shaft 108. Where applicable, the blower 101 is driven by the low-pressure shaft 10.

[0039] Figure 3 represents the low-pressure turbine 1. This turbine comprises a rotor equipped with one or more disks 16 on which one or more rotor blades 4 are mounted. The disks 16 and their blades 4 are distributed over one or more turbine stages along an axial direction defined by an axis X of the low-pressure turbine 1. Static or straightening blades 3 may be arranged between the rotor blades 4 so as to straighten the flow in the turbine 1 between two successive rotor stages. A radial direction is further defined as any direction perpendicular to and passing through the axis X of the turbomachine assembly.

[0040] During operation of the turbomachine, the rotor blades 4 are rotating around the turbine X axis.

[0041] In order to ventilate the rotor blade discs 16 of the rotor blades 4, a cavity surrounding the drive shaft X and radially internal to the flow channel 18 is configured to allow air to enter through at least one air inlet 12. The air supplied to the cavity 17 is taken from a section of the turbomachine that is cooler than the low-pressure turbine 1, so that the admitted air is cooler than the air in the turbine 1 and thus serves to cool the rotor blade discs 16 of the rotor blades 4. For example, the air supplied to the cavity 17 is taken from the high-pressure compressor 103, and may therefore have a temperature approximately 400 to 500 °C lower. at the air temperature in the flow channel 18 of the turbine 1. The air passing through the air intake 12 allows, by ventilating the discs 16 of the rotor blades 4, to extend the life of these discs 16.

[0042] At least one air outlet 14 allows air to be purged from the cavity 17 to the flow vein 18, so as to prevent a return of hot air from the flow vein 18 into the cavity 17, which could undesirably heat the discs 16.

[0043] According to one embodiment, the air outlet 14 is located directly upstream of a first stage of the turbine 1 in the normal direction of gas flow. Alternatively, the air outlet 14 may be located in an intermediate position of the turbine, with some turbine stages located upstream of the air outlet 14 in the normal direction of gas flow.

[0044] The sealing of cavity 17 is ensured by a sealing gasket 22, shown in detail in Figures 4 and 5. The gasket 22 comprises a stator gasket portion 25, integral with the stator of the turbine 1 and fixed relative to the drive axis X during operation of the turbomachine. The gasket 22 further comprises a rotating gasket portion 26, integral with the rotor of the turbine 1, and which therefore rotates around the drive axis X with the rotor during operation of the turbomachine.

[0045] The turbine 1 includes an active control system for the radial clearance and / or axial displacement between the stator seal portion 25 and the rotating seal portion 26 of the seal 22. The active control system includes at least one heating element 2 disposed either opposite and near a radially external surface 252 of the stator seal portion 25, or in direct contact with this external surface 252. For example, the heating element 2 may be disposed at a distance from the radially external surface 252 greater than or equal to 1 mm and less than or equal to 5 mm. Alternatively, the heating element 2 is received on a portion of the radial dimension of the radially external surface 252, in a recess in the radially external surface 252 provided for this purpose. The heating element 2 may, in particular, be a resistive element suitable for heating by Joule heating.The heating element 2 is powered by an electrical voltage source, and heats the stator joint portion 25 of the joint 22 by heat transfer with its radially external surface 252. This causes an expansion of the stator joint portion 25 without significantly affecting the behavior of the rotating joint portion 26. Following a first flight of an aircraft comprising a turbomachine equipped with the low-pressure turbine 1, the heating element can be used to cause an expansion of the stator joint portion 25, i.e. an increase in the diameter of a radially internal surface 251 of the stator joint portion 25 by thermal expansion, and thus an increase in the radial clearance between the radially internal surface 251 and the rotating joint portion. 26. Thus, the active control system prevents excessive thermal contraction of the stator seal portion 25, so that the duration of an engine stall phase, during which the stator seal portion 25 is in contact with the radially internal surface 251, can be reduced or even eliminated, and a second flight can be performed quickly after the first flight. For an airline, the proposed active control system therefore allows for more flights to be performed in a single day.

[0046] According to certain embodiments, the seal 22 comprises, on the radially internal surface 251 of its stator seal portion 25, an abradable coating 27. The rotating seal portion 26 comprises one or more blades 28 projecting from a base of the rotating seal portion 26. During operation of the turbomachine, the blades 28 rotate around the drive shaft X in contact with the abradable coating 27 so as to ensure the sealing of the cavity 17. The blades thus erode the abradable coating 27.

[0047] The seal 22 may, for example, comprise six slats 28, as shown in [Fig. 4]. The seal 22 may, in particular, in a manner known per se, comprise a plurality of slats arranged in groups of slats 30, each group of slats 30 comprising one or more axially adjacent slats 28 having substantially the same radial dimension. Correspondingly, the abradable coating then comprises steps 29 of different radial dimensions, so that the abradable coating has a stepped shape. The abradable coating comprises a number of steps 29 equal to the number of groups of slats 30. Each step 29 is provided, in at least one phase of flight of the turbomachine (for example, a takeoff, cruise, or landing phase), to be arranged radially opposite a group of slats 30.In the event of a significant radial displacement of the stator joint portion 25 relative to the rotating joint portion 26, the groups of blades 30 and the steps 29 allow this clearance to be reduced, the axial displacement of the stator joint portion 25 relative to the rotating joint portion 26 changing which group of blades 30 is opposite which step 29, so as to place in front of a given step 29 a group of blades 30 of radial dimension greater than that of the group of blades 30 which was previously opposite this step 29.

[0048] According to other embodiments, shown in [Fig. 5], the sealing joint 22 comprises fewer than six flaps 28. The joint 22 may, for example, comprise two flaps 28 of substantially identical radial dimensions, the abradable coating 27 then having a substantially constant radial dimension, and does not exhibit any steps. Indeed, the reduction of the axial displacement of the stator seal portion 25 relative to the rotating seal portion 26 during flight, made possible by the active control operated by means of the heating elements 2, allows sufficient control of the radial clearance between the blades 28 and the abradable coating 27, so that it is not necessary to provide steps 29 which ensure a change in the radial clearance between the blades 28 and the abradable coating 27. Compared with conventional seals, for which it is necessary to provide steps 29 in the abradable material 27, as well as a large number of blades 28, the proposed seal is more compact and lighter, and therefore allows for better performance of the turbomachine.

[0049] The joint 22 can alternatively comprise between four and six licks 28. The joint 22 can in particular comprise six licks 28.

[0050] According to some embodiments, the heating element 2 extends substantially radially in relation to a lip 28, so as to allow effective expansion of the portion of stator seal 25 at this radial position.

[0051] The active control system preferably includes a controller 11. This controller 11 can in particular modulate an electrical power transmitted to the heating element 2. It is thus possible to adjust a dilation of the stator seal portion 25, which makes it possible to control the radial clearance between the rotating seal portion 26 and the radially internal surface 251 of the stator seal portion 25 for any phase of flight, and thus optimize the performance of the low-pressure turbine 1.

[0052] In one embodiment, the heating element 2 is electrically powered by a battery 6. Preferably, the battery 6 is connected via an electrical link 8 to an alternator 7, which is itself mechanically connected to the shaft 10 of the low-pressure turbine 1. Thus, it is not necessary to provide a separate power supply for the battery 6 specific to the active control system. However, in some embodiments, the battery is not connected to the shaft 10 of the turbomachine assembly.

[0053] The controller 11 may have an interface with the battery 6 and / or with the alternator 7.

[0054] According to one embodiment, the controller 11 is located near the battery 6 so as to allow signals emitted by the controller 11 to the battery 6 to pass quickly between these two elements.

[0055] It is possible to provide that only a portion of the electrical energy from battery 6 is allocated to the operation of the active control system, and that other functions of the turbomachine are supplied with electrical power from battery 6.

[0056] In one embodiment, the alternator 7 is driven by the low-pressure turbine 1. To this end, a rotor of the alternator 7 is mechanically connected to the shaft 10 of the low-pressure turbine 1, for example, by a splined system. The alternator 7 comprises first splines 9, which are complementary to second splines (not shown) arranged on the shaft 10 of the turbomachine assembly 1. Any other equivalent mechanical connection method between the alternator 7 However, the shaft 10 of the low-pressure turbine 1 falls within the scope of the proposed system. This could, for example, involve bolted connections or a shrink-fit connection.

[0057] When the active control system includes several heating elements 2, the controller 11 can be designed to allow independent modulation of the electrical power supplying the different heating elements 2. Indeed, it may be desirable to induce a different expansion of the stator joint portion 25 of the seal 22 at different axial positions. The heating requirement of each rotor stage can be determined by numerical simulation, in particular by finite element analysis. For example, it is possible to model a heating system that allows the retraction of the stator joint portion 25 to be identical to the retraction of the rotating joint portion 26, so as to maintain a constant radial clearance between them.

[0058] According to other aspects, the present disclosure relates to a method for actively controlling clearances in a compressor or turbomachine turbine 1, which may be axial or radial clearance as defined above. The method includes at least one step of heating a radially external surface 252 of a stator seal portion 25 of a seal 22 of the compressor or turbine, thereby increasing the diameter of the radially internal surface 251 of the stator seal portion 25 by thermal expansion, and consequently increasing the clearance between the rotating seal portion 26 of the seal 22 and the radially internal surface 251 of the stator seal portion 25.The heating step of the stator joint portion 25 is implemented by means of a heating element 2 arranged opposite the radially external surface 252, which may in particular be a heating element having an electrical resistance and capable of heating by Joule effect, a heat exchange taking place between the heating element 2 and the radially external surface 252.

[0059] For example, the heating step is implemented prior to starting the turbomachine, so as to prevent a blockage of the seal 22 by increasing the radial clearance between the stator seal portion 25 and the rotating seal portion 26.

[0060] According to one embodiment of the method, a desired clearance value is predefined, and the electrical power transmitted to the heating element 2 is modulated so that the axial or radial clearance reaches this predefined value. Optionally, prior to the commissioning of the turbomachine or a flight, a controller 11 may be calibrated to transmit electrical power to the heating element 2 in such a way as to determine what electrical power is required to achieve a given axial or radial clearance value.

[0061] The controller 11 may, according to one embodiment of the method, be capable of receiving information relating to radial or axial play, in which case the modulation of the electrical power transmitted by the controller 11 to the heating element 2 may take into account This information is used to increase or decrease the electrical power. Specifically, this information may include a difference between the desired preset clearance value and the actual clearance value in the compressor or turbine 1. The controller 11 can then modulate the electrical power to reduce this difference. These steps of gathering information and reducing the difference between the desired preset clearance value and the actual clearance value can be repeated, for example, until the difference is less than an acceptable error.

[0062] The active control method can be implemented during a turbomachine startup phase prior to takeoff. For example, the active control method can be implemented to maintain the radial or axial clearance between the stator seal portion 25 and the rotating seal portion 26 of the seal constant, in particular to maintain it at the actual value of this clearance when the turbomachine is shut down, at the end of a previous flight. This prevents contact between the stator seal portion 25 and the rotating seal portion 26, which could lead to a blockage of the seal 22 and therefore of the rotor, and prevent or delay a turbomachine restart.

Claims

Demands

1. Assembly (1) of a turbomachine centered on an axis X, comprising: a rotor, a stator, a seal (22) comprising a portion of stator seal (25) mechanically attached to the stator and a portion of rotating seal (26) mechanically attached to the rotor, characterized in that it comprises at least one heating element (2), preferably resistive, disposed opposite or attached to a radially external surface (252) of the portion of stator seal (25) and in that each heating element (2) is configured to heat the radially external surface (252) of the portion of stator seal (25).

2. Assembly (1) of turbomachine according to claim 1, the assembly being a turbine or a compressor.

3. Assembly (1) of turbomachine according to claim 1 or 2, comprising a battery (6) configured to power each heating element (2), and preferably comprising an alternator (7) suitable for being mechanically driven by a drive shaft (10) and configured to supply the battery (6) with electrical power.

4. Assembly (1) of turbomachine according to claim 3, wherein the alternator (7) comprises first splines (9) complementary to second splines of the drive shaft, allowing mechanical coupling between the alternator (7) and the drive shaft (10).

5. Assembly (1) of turbomachine according to any one of claims 1 to 4, comprising at least two heating elements (2), preferably four resistive heating elements arranged at distinct axial positions.

6. Assembly (1) of turbomachine according to any one of claims 1 to 5, comprising a controller (11) configured to control a variation of electrical power transmitted to each heating element (2).

7. Turbomachine assembly (1) according to any one of claims 1 to 6, wherein a radially internal surface (251) of the stator seal portion (25) comprises an abradable coating (27) positioned radially opposite a lip (28) of the rotating seal portion (26)

8. Assembly (1) of turbomachine according to claim 7, wherein the rotating joint portion (26) comprises between one and ten blades (28), preferably between four and six blades (28).

9. Assembly (1) of turbomachine according to any one of claims 7 and 8, wherein each heating element (2) is arranged radially opposite a nozzle (28).

10. A method for actively controlling a clearance between, on the one hand, a radially internal surface (251) of the stator seal portion (25) of the seal (22) of a turbomachine assembly according to any one of claims 1 to 9 and, on the other hand, a radially external surface of the rotating seal portion (26) of the seal, the active control method comprising a step of heating the radially external surface (252) of the stator seal portion (25) so as to increase the clearance between the radially external surface of the rotating seal portion (26) and the radially internal surface (251) of the stator seal portion (25).

11. A method according to claim 10, comprising a step of modulating electrical power transmitted to the heating element (2) so that the radial clearance between the radially external surface of the rotating joint portion (26) and the radially internal surface of the stator joint portion (25) reaches a predefined clearance value.

12. Method according to claim 11, comprising a step of collecting information by a controller to determine a gap between, on the one hand, the radial clearance between the radially external surface of the rotating joint portion (26) and the radially internal surface of the stator joint portion (25) and, on the other hand, the predefined clearance value.

13. A method according to any one of claims 11 or 12, wherein the step of modulating the electrical power transmitted to the heating element (2) is implemented so as to maintain the radial clearance between the radially external surface of the rotating joint portion (26) and the radially internal surface (251) of the stator joint portion (25) constant during a turbomachine start-up phase prior to take-off.

14. Turbomachine turbine or compressor comprising a turbomachine assembly (1) according to any one of claims 1 to 9.

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