Active control system for radial or axial clearances in an aeronautical turbomachine compressor or turbine

Abstract

This disclosure relates to an X-axis turbomachine assembly (1), comprising: - a rotor including rotor blades (4), and - a housing (13) including a wall surrounding the rotor, the wall having a radially internal surface (131) facing the rotor blades (4) and a radially external surface (132), the assembly including at least one heating element (2), preferably resistive, disposed opposite or attached to the radially external surface (132) of the housing wall (13), and in that each heating element (2) is configured to heat a thermally driven zone of the housing wall (13). Abbreviated figure: Figure 2

Description

Title of the invention: Active control system for radial or axial clearances in an aeronautical turbomachine compressor or turbine. FIELD OF THE INVENTION

[0001] The invention relates to aircraft turbomachinery. More particularly, the invention relates to radial or axial clearance control systems between rotating blades and a fixed casing in a turbomachine assembly 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] Within a turbine, during turbomachine operation, the temperature variation in the flow stream according to engine speed affects the thermal expansion or contraction behavior of the moving blades as well as the fixed turbine casing. In order to minimize radial clearances between the tips of the moving blades and the fixed casing, active control systems are used. Indeed, it is desirable to reduce such radial clearances to force the air to flow in the flow stream rather than in a space located between the moving blade and the casing, thereby improving the efficiency of the turbomachine.

[0006] Such active control systems are also designated as LPTACC or HPTACC depending on whether they are provided around the low-pressure turbine or the high-pressure turbine of a turbomachine (for Low Pressure Turbine Active Clearance). Control and High Pressure Turbine Active Clearance Central, i.e. active control of low pressure / high pressure turbine clearance respectively).

[0007] When the turbomachine shuts down following a flight, the difference in thermal expansion behavior between the fixed turbine casing and the moving blades further reduces the radial clearances between them, as the fixed casing cools down faster than the moving blades. Thus, when the turbomachine shuts down, initially, the diameter of an internal surface of the fixed casing—and therefore its distance from the engine shaft—decreases due to this cooling, while the distance between the tips of the rotor blades and the engine shaft decreases less rapidly. This is primarily due to the fact that the external surface of the fixed casing is in contact with air that is cooler than that found in the main airflow channel defined between the blades and the fixed casing.Thus, for several hours, contact can exist between the blade tips and the fixed casing or an abradable material bonded to its inner surface, preventing the turbomachine from restarting due to a phenomenon known as "rotor lock." It is therefore necessary to wait for the moving blades to retract sufficiently to reintroduce radial clearance between them and the fixed casing, which prevents the turbomachine from immediately restarting, and thus from initiating 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 reduce this rotor lock-up phenomenon by taking into account this reduction of radial clearances in the design of the turbines, but this requires providing larger radial clearances between the moving blades and the fixed casing, which degrades the efficiency of the turbine during all phases of flight.

[0008] It should be noted that rotor lock-up can also occur due to axial contact between the rotor blades and the fixed housing, also resulting from the difference in thermal expansion behavior of these components. Such axial contact can damage or even cause rotor blades to break. During a turbomachine restart, friction occurs due to this axial contact, which can exacerbate the rotor lock-up phenomenon.

[0009] A solution to this problem of rotor blockage at start-up implemented in the prior art is to provide, prior to the start-up of the turbomachine, a break-in of an abradable coating fixed on the internal surface of the fixed housing. This break-in process creates radial clearance between the moving blades and the coating, thus allowing the blades to rotate freely and the turbomachine to start. However, although implementing such a break-in process reduces the waiting time required to start the turbomachine without rotor lock-up, the The break-in process itself requires a certain amount of time, potentially lasting several hours. Furthermore, breaking in part of the abradable coating is not desirable: this coating is designed to be eroded during the various phases of flight to achieve the smallest possible radial clearance between the moving blades and the fixed housing without damaging the blades, and breaking in would cause premature wear of this coating.

[0010] To a lesser extent, contacts between moving blades and turbine stator can also occur due to the movements of these elements along the axial direction, also due to differences in thermal expansion behavior of these elements when the turbomachine is stopped.

[0011] The aforementioned LPTACC / HPTACC type active control systems cannot overcome this rotor lock-up problem at startup. Indeed, these systems allow cooling of the turbine housing during certain flight phases to reduce the radial clearance between the rotor and the housing as much as possible. At startup, these systems cannot prevent rotor lock-up, which is caused by zero radial clearance, because these systems cannot increase this radial clearance. Furthermore, even if sufficient radial clearance were provided in the turbomachine design, such LPTACC / HPTACC systems would not allow this radial clearance to be reduced to the minimum necessary for efficient turbine operation, as the pressure difference between the intake and intake points would then be insufficient to allow the air to effectively cool the housing. EXPOSED

[0012] One objective of the present disclosure is therefore to provide an active control system for radial and axial clearances in a turbomachine turbine that makes it possible to avoid a blockage of the turbomachine rotor at start-up, without negatively affecting the efficiency of the turbomachine during its subsequent phases of use.

[0013] Another objective of the present disclosure is to provide an active control system for radial and axial clearances in a turbomachine turbine 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 first aspect of a turbomachine assembly is proposed, comprising: - a rotor comprising rotor blades, and - a housing comprising a wall surrounding the rotor, the wall having a radially internal surface facing the rotor blades and a radially external surface, the assembly comprising at least one heating element, preferably resistive, arranged facing or attached to the radially external surface of the housing wall and in that each heating element is configured to heat a thermal control zone of the housing wall.

[0016] The use of heating elements to heat the external surface of the fixed casing allows for the selective increase of the casing's internal diameter, thereby increasing the radial and / or axial clearance between an internal casing surface and a blade tip, thus preventing rotor lock-up during turbomachine startup. This makes it possible to start the turbomachine quickly after 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 turbomachine design can incorporate smaller clearances between the fixed casing and the rotor blades at rest, improving engine performance during all flight phases and thus significantly reducing fuel consumption.

[0017] According to one embodiment, the assembly is a turbine or a compressor.

[0018] According to one embodiment, the assembly includes a power supply device in cooling air, intended to cool the crankcase wall by impact jet on the radially external surface of the crankcase wall, the cooling air supply device being mounted radially opposite the radially external surface and being configured so as to cool each thermal pilot zone via the cooling air.

[0019] According to one embodiment, each heating element and the cooling air supply device are each designed to control a clearance between the rotor and the housing, the clearance being an axial and / or radial clearance.

[0020] According to one embodiment, the assembly includes a battery configured to power each heating element, and preferably comprising an alternator suitable for being mechanically driven by a drive shaft and configured to supply the battery with electrical energy.

[0021] 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.

[0022] According to one embodiment, each heating element is arranged radially opposite a stage of rotor blades or a stage of stator blades.

[0023] According to one embodiment, the assembly comprises at least two heating elements, each arranged radially opposite a rotor blade stage or a stator blade stage.

[0024] According to one embodiment, the assembly comprises a number of heating elements less than or equal to the sum of the number of rotor blade stages and the number of stator blade stages in the assembly, each heating element extending radially in relation to a rotor blade stage or radially in relation to a different stator blade stage.

[0025] According to one embodiment, the housing includes a hook extending radially inwards from the radially internal surface of the housing, the hook being configured to allow mounting of a stator blade of the assembly on the radially internal surface of the housing, and in which a heating element is arranged radially opposite the hook.

[0026] According to one embodiment, the assembly includes a controller configured to control a variation of electrical power transmitted to each heating element.

[0027] According to one embodiment, each heating element is fixed on a support, the support being mounted on the housing radially outside the radially external surface of the housing.

[0028] According to one embodiment, the assembly includes an abradable element mounted on the hook, the abradable element being positioned radially opposite the rotor blade over at least part of a rotation of the rotor blade around the axis, each heating element being configured to allow control of a clearance between the rotor blade and the abradable element.

[0029] The present disclosure relates, according to a second aspect, to a method of actively controlling a clearance between, on the one hand, the radially internal surface of the casing wall of a turbomachine assembly as defined above and, on the other hand, a rotor blade of the turbomachine assembly, the active control method comprising a step of heating the casing wall of the turbomachine assembly so as to increase the clearance between the rotor blade of the assembly and the radially internal surface of the casing.

[0030] According to one embodiment of the method, the turbomachine assembly includes a cooling air supply device for cooling the casing wall by impact jet on the radially external surface of the casing wall, the cooling air supply device being mounted radially opposite the radially external surface, the method comprising a step of cooling the casing wall of the turbomachine assembly by means of the air supply device cooling so as to reduce the clearance between the rotor blade of the turbomachine assembly and the radially internal surface of the casing.

[0031] 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 rotor blade and the radially internal surface of the housing reaches a predefined clearance value.

[0032] 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 rotor blade and the radially internal surface of the housing and, on the other hand, the predefined clearance value.

[0033] 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 rotor blade and the radially internal surface constant during a turbomachine start-up phase prior to take-off. DESCRIPTION OF THE FIGURES

[0034] Other features, purposes and advantages of the invention will become apparent from the following description, which is purely illustrative and not limiting, and which should be read in conjunction with the accompanying drawings on which:

[0035] Fig. 1 schematically represents a turbomachine comprising a low-pressure turbine equipped with an active radial and / or axial clearance control system as proposed.

[0036] Fig.2 schematically represents the low-pressure turbine of the turbomachine shown in Fig.1, according to a first embodiment.

[0037] Fig. 3 schematically represents the low-pressure turbine of the turbomachine shown in Fig. 1, according to a second embodiment.

[0038] [Fig.4] schematically represents an aircraft equipped with a turbomachine as shown in [Fig.1].

[0039] Throughout the figures, similar elements bear identical references. DETAILED DESCRIPTION OF THE INVENTION

[0040] Figure 1 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, a turbomachine assembly consisting of a low-pressure turbine 1 will be used as an example. However, the following description applies equally to any other turbine or turbomachine compressor.

[0041] 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.

[0042] Figure 2 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.

[0043] During operation of the turbomachine, the rotor blades 4 rotate about the turbine axis X. The low-pressure turbine 1 also includes a fixed housing 13 with a wall, the wall comprising a radially internal surface 131 opposite the ends of the rotor blades 4, and a radially external surface 132 on the opposite side. The wall is substantially cylindrical with revolution about the turbine axis X.

[0044] The proposed active control system comprises a heating element 2 disposed near the radially external surface 132 of the fixed housing 13. For example, the heating element 2 may be disposed at a distance from the radially external surface 132 greater than or equal to 1 mm and less than or equal to 5 mm. The heating element may also be in contact with the radially external surface 132, and in particular, be received over a portion of its radial dimension in a recess in the radially external surface 132, as shown in [Fig. 2]. 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 wall of the housing 13 by heat transfer with the radially external surface 132. This causes the fixed housing 13 to expand, without significantly affecting the behavior of the rotor blades 4.Following the first flight of an aircraft comprising a turbomachine equipped with the low-pressure turbine 1, the heating element 2 can be used to cause expansion of the fixed casing 13, i.e., an increase in the diameter of its radially internal surface 131 by thermal expansion, and thus an increase in the radial and / or axial clearance between the radially internal surface 131 and the tips of the rotor blades 4. In this way, the duration of an engine stall phase, during which the tips of the rotor blades 4 are in contact with the radially internal surface 131, can be reduced or even eliminated, and a second flight can be carried out quickly after the first flight. For . For an airline, the proposed active control system therefore allows for more flights to be carried out on a given day.

[0045] The active control system may also include a power supply device 21. This device can, during certain flight phases, project a cooling fluid, preferably air, by means of an impact jet onto the radially external surface 132 of the fixed housing 13 in order to cool it. This causes thermal contraction of the fixed housing 13, and therefore a reduction in the diameter of its radially internal surface 131, without significantly affecting the behavior of the rotor blades 3. This makes it possible, during the relevant flight phases, to reduce the radial or axial clearance between the fixed housing 13 and the rotor blades 4 if it is excessive. Thus, as much air as possible flows in the flow stream and the efficiency of the compressor or turbine 1 is improved.The cooling air supply device 21 can be configured to supply air drawn from a position of the turbomachine located further upstream than the compressor or turbine 1 and cooler than the air flowing in the flow channel, which is in contact with the radially internal surface 131. For example, if the active control system is located around a turbine 1, the air supplied by the cooling air supply device 21 is drawn from a compressor. The cooling air supply device 21 may include an air distribution box 27, to which several distribution tubes 28 are connected. The cooling fluid is supplied to the distribution box 27 and then distributed by impact jet onto the radially external surface 132 of the fixed housing 13 via the distribution tubes 28.

[0046] According to other embodiments, the active control system includes the heating element 2 but does not include a cooling air supply system 21.

[0047] According to one embodiment, the heating element 2 extends substantially radially with respect to a rotor blade 4, so as to allow optimal expansion of the fixed housing 13 at this radial position. Alternatively, the heating element 2 can extend substantially radially with respect to a static blade 3, or even at another axial position intermediate between the respective axial positions of a static blade 3 and a rotor blade 4.

[0048] The cooling air supply device 21 and the heating elements 2 together allow the clearance between the rotor blades 4 and the radially internal surface 131 to be increased or decreased, and thus precise control of this clearance.

[0049] The active control system preferably includes a controller 11. This controller 11 can, in particular, modulate the electrical power transmitted to the heating element 2. It is thus possible to adjust the expansion of the fixed housing 13, This allows control of the radial clearance between the ends of the rotor blades 4 and the radially internal surface 131 for any phase of flight, and thus improves the performance of the low-pressure turbine 1.

[0050] 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.

[0051] The controller 11 may have an interface with the battery 6 and / or with the alternator 7. 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.

[0052] 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.

[0053] 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. However, any other equivalent mechanical connection method between the alternator 7 and 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 shrink fittings.

[0054] According to one embodiment, the active control system comprises at least two heating elements 2. Each heating element 2 may, in particular, be arranged in an axial position located in a radial plane passing through a given rotor blade stage 4. In this case, it is possible that the heating elements 2 are arranged opposite only some of the rotor stages, the other rotor stages having no heating element 2 arranged opposite them in their axial position.Alternatively, the active control system may include a number of heating elements 2 less than or equal to the number of rotor stages present in the low-pressure turbine 1, in particular equal to the number of rotor blades 4, so as to allow independent control of clearances for several rotor stages of the low-pressure turbine 1 (for a number of heating elements 2 strictly less than the number of rotor stages) or even for each rotor stage of the low-pressure turbine 1 (for a number of heating elements 2 equal to the number of rotor stages).

[0055] According to another embodiment, the active control system comprises a number of heating elements 2 less than or equal to, in particular equal to, the sum of the number of rotor blades 4 and the number of static blades 3 present in the low pressure turbine 1.

[0056] According to one embodiment, and as shown in [Fig.2], heating elements 2 are arranged at least radially opposite hooks forming part of the radially internal surface 131 of the wall of the fixed housing 13. The hooks provide a mechanical connection of the radially internal surface 131 with the distributors 3. The hooks include downstream hooks 24, arranged directly downstream of the distributors 3, and upstream hooks 22, arranged directly upstream of the distributors 3. This makes it possible to control the radial play at the level of the hooks 22, which can be subject to particularly large thermal gradients in operation of the turbomachine.

[0057] According to an embodiment illustrated in [Fig. 3], the heating elements 2 can be arranged on supports 23 attached to the fixed housing 13. The supports 23 are attached to the fixed housing 13 on either side of the housing along its axial dimension. In order not to obstruct the passage of impact jets from the cooling air supply device 21, the supports 23 can comprise a plurality of branches 29 extending circumferentially around the fixed housing 13, the branches being offset along the axial direction relative to the distribution tubes 28. The branches can, in particular, be connected to each other at specific points, at a given circumferential position of the support 23. The heating elements 2 are then also offset axially from the cooling tubes 28.For example, the cooling tubes can be arranged opposite the hooks 22, 24 so as to allow optimal cooling of these hooks, and the heating elements 2 are then arranged opposite the rotor blades 4.

[0058] The active control system may have, in addition to the heating elements arranged radially opposite the rotor blades 4, one or more heating elements 2 opposite one or more static blades 3, so as to also allow selective expansion of the fixed housing 13 at the axial positions where the static blades 3 are located.

[0059] 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 different expansions of the fixed housing 13 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 fixed housing 13 to contract in the same way as the rotor stages.

[0060] According to one embodiment, an abradable material 15 is disposed in contact with the radially internal surface 131 of the fixed housing 13, at least radially opposite one end of a rotor blade 4 during a portion of its rotation around the motor shaft. The abradable material 15 may, in particular, be disposed in contact with an upstream hook 22. The abradable material 15 and the adjacent upstream hook 22 are then received in a receiving portion 25 of the distributor 3, the receiving portion 25 engaging with the abradable material 15 and the upstream hook 22 downstream of the distributor 3. Such an abradable material 15 is intended to be lapped by the rotor blades 4 during certain phases of flight in order to obtain the smallest possible—but not zero—radial clearance between the ends of the rotor blades 4 and the radially internal surface 131.In this embodiment, the active control system prevents excessive running-in of the abradable material 15 at turbomachine startup, which would result from the thermal contraction of the fixed housing 13 before that of the rotor blades 4. Such running-in at startup presents several disadvantages: firstly, it prematurely damages the abradable material 15 and necessitates its replacement earlier in the turbomachine's service life. Secondly, it prevents the achievement of optimal radial clearance in subsequent flight phases, since excessive radial clearance may then exist between the tips of the rotor blades 4 and the already eroded abradable material. Finally, the procedure for running-in the abradable material at turbomachine startup, sometimes used in the prior art, has a considerable duration that delays the point at which the turbomachine can be used again.

[0061] The embodiments of the proposed system comprising an abradable material 15 are therefore particularly interesting during a take-off phase, which immediately follows the start-up of the turbomachine and for which the premature erosion of the abradable material in the prior art turbomachines has a particularly deleterious effect.

[0062] In other respects, and as shown in [Fig. 4], this disclosure also relates to a turbomachine 17 equipped with a turbine or compressor 1 comprising the active control system described above. This disclosure further relates to an aircraft 18 comprising a fuselage 19 and optionally wings 20, the aircraft 18 being equipped with the turbomachine 17 thus defined.

[0063] According to other aspects, the present disclosure relates to a method for actively controlling clearances in a compressor or turbine 1 of a turbomachine, which may be axial or radial clearance as defined above. The method includes at least one step of heating a wall of a fixed housing 13 of the compressor or turbine 1, thereby increasing the diameter of a radially internal surface 131 of the wall of the fixed housing 13 by thermal expansion, and consequently increasing the clearance between a blade 4 of a rotor of the compressor or turbine 1 and the radially internal surface 131. internal 131. The heating step of the wall of the housing 13 is implemented by means of a heating element 2 arranged opposite the radially external surface 132, which can in particular be an 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 132.

[0064] The method further includes a step of cooling the wall of the fixed housing 13 of the compressor or turbine 1, which reduces the diameter of the radially internal surface of the fixed housing 13 by thermal shrinkage, and which consequently reduces the clearance between the rotor blade 4 and the internal surface of the fixed housing 13. The step of cooling the wall of the fixed housing 13 is carried out by means of an impact projection of a cooling fluid, preferably air, onto the radially external surface 132 of the wall of the fixed housing 13.

[0065] The cooling and heating stages of the wall of the fixed housing 13 can be implemented during different phases of aircraft operation. For example, the heating stage is implemented prior to the start of the turbomachine, so as to prevent a blockage of the rotor, while the cooling stage is implemented during flight phases subsequent to start, so as to reduce the clearance between the rotor blade 4 and the radially internal surface 131 of the fixed housing 13 and thus force a greater proportion of the flow into the flow channel.

[0066] 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.

[0067] The controller 11 may, according to one embodiment of the method, be capable of receiving information relating to radial or axial clearance, in which case the modulation of the electrical power transmitted by the controller 11 to the heating element 2 may take this information into account in order to increase or decrease the electrical power. In particular, this information may include a difference between the predefined clearance value that one wishes to achieve and the actual clearance value in the compressor or turbine 1. The controller 11 may then modulate the electrical power so as to reduce this difference. These steps of collecting information and reducing the difference between the predefined clearance value that one wishes to achieve and the actual clearance value may be repeated, for example, until the difference is less than an acceptable error.

[0068] 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 a constant radial or axial clearance between the blade 4 and the radially internal surface 131 of the fixed housing 13, 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 blade 4 and the radially internal surface 131 of the wall of the fixed housing 13, which could lead to rotor blockage and prevent or delay a turbomachine restart.

Claims

Demands

1. Assembly (1) of an axis (X), comprising: - a rotor including rotor blades (4), and - a casing (13) including a wall surrounding the rotor, the wall having a radially internal surface (131) opposite the rotor blades (4) and a radially external surface (132), characterized in that it includes at least one heating element (2), preferably resistive, disposed opposite or attached to the radially external surface (132) of the casing wall (13) and in that each heating element (2) is configured to heat a thermal control zone of the casing wall (13).

2. Turbomachine assembly (17) according to claim 1, the assembly being a turbine or a compressor.

3. Assembly (1) of turbomachine according to any one of claims 1 and 2, comprising a cooling air supply device (21), intended to cool the casing wall (13) by impact jet on the radially external surface (132) of the casing wall (13), the cooling air supply device (21) being mounted radially opposite the radially external surface (132) and being configured so as to cool each thermal pilot zone via the cooling air.

4. Assembly (1) of turbomachine according to claim 3, wherein each heating element (2) and the cooling air supply device (21) are adapted to drive each a clearance between the rotor and the casing, the clearance being axial and / or radial.

5. Assembly (1) of turbomachine according to any one of claims 1 to 4, 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.

6. Assembly (1) of turbomachine according to claim 5, in which the alternator (7) comprises first splines (9) complementary to second splines of the drive shaft (10), allowing a mechanical coupling between the alternator (7) and the drive shaft (10).

7. Turbomachine assembly (1) according to any one of claims 1 to 6, wherein each heating element (2) is arranged radially opposite a rotor blade stage (4) or a stator blade stage (3).

8. Assembly (1) of turbomachine according to claim 7, comprising at least two heating elements (2), each arranged radially opposite a rotor blade stage (4) or a stator blade stage (3).

9. Turbomachine assembly (1) according to claim 8, comprising a number of heating elements (2) less than or equal to the sum of the number of rotor blade stages (4) and the number of stator blade stages (3) in the assembly (1), each heating element (2) extending radially opposite a different rotor blade stage (4) or radially opposite a different stator blade stage (3).

10. Turbomachine assembly (1) according to any one of claims 1 to 6, wherein the housing (13) includes a hook (22, 24) extending radially inwards from the radially internal surface (131) of the housing (13), the hook (22, 24) being configured to permit mounting of a stator blade (3) of the assembly on the radially internal surface (131) of the housing (13), and wherein a heating element (2) is arranged radially opposite the hook (22, 24).

11. Assembly (1) of turbomachine according to any one of claims 1 to 10, comprising a controller (11) configured to control a variation of electrical power transmitted to each heating element (2).

12. Assembly (1) of turbomachine according to any one of claims 1 to 11, wherein each heating element (2) is fixed to a support (23), the support (23) being mounted on the housing (13) radially outside the radially external surface (132) of the housing (13).

13. Turbomachine assembly according to any one of claims 10 to 12, comprising an abradable element (15) mounted on the hook (22, 24), the abradable element (15) being positioned radially opposite the rotor blade (4) over at least a portion of a rotation of the rotor blade (4) around the axis (X), each heating element (2) being configured to allow control of a clearance between the rotor blade (4) and the abradable element (15).

14. Method of actively controlling a clearance between, on the one hand, the radially internal surface (131) of the wall of the casing (13) of a turbomachine assembly (1) according to any one of claims 1 to 13 and, on the other hand, a rotor blade (4) of the turbomachine assembly (1), the active control method comprising a step of heating the wall of the casing (13) of the turbomachine assembly (1) so as to increase the clearance between the rotor blade (4) of the assembly (1) and the radially internal surface (131) of the casing (13).

15. An active control method according to claim 14, wherein the turbomachine assembly includes a cooling air supply device (21), intended to cool the casing wall (13) by impact jet on the radially external surface (132) of the casing wall (13), the cooling air supply device (21) being mounted radially opposite the radially external surface (132), the method comprising a step of cooling the casing wall (13) of the turbomachine assembly (1) by means of the cooling air supply device (21) so as to reduce the clearance between the rotor blade (4) of the turbomachine assembly (1) and the radially internal surface (131) of the casing (13).

16. A method according to any one of claims 14 and 15, comprising a step of modulating electrical power transmitted to the heating element (2) so that the radial clearance between the rotor blade (4) and the radially internal surface (131) of the housing (13) reaches a predefined clearance value.

17. Method according to claim 16, comprising a step of collecting information by a controller (11) to determine a deviation between, on the one hand, the radial clearance between the rotor blade (4) and the radially internal surface (131) of the housing (13) and, on the other hand, the predefined clearance value.

18. A method according to any one of claims 16 to 17, 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 rotor blade (4) and the radially internal surface (131) constant during a turbomachine start-up phase prior to take-off.

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