System for cooling a rotating part of a dual-bleed turbomachine

Abstract

The present disclosure relates to a turbomachine compressor (10), comprising: at least one rotor stage provided with a rotor disc (5), and a cooling system comprising a pipe having an air bleed opening (3) located on an inner shroud (20) of the compressor (10), wherein the cooling system comprises a flap (9) and a return element (11), wherein the flap (9) is configured to adopt: a first position, in which it allows, through the opening (3) at a first rotational speed of the turbomachine, a first air flow to pass, and a second position, in which it allows, through the opening (3) at a second rotational speed of the turbomachine higher than the first speed, a second air flow or no flow to pass, and wherein the compressor is configured to enable the flap (9) to move from the first position to the second position through a change in a centrifugal force acting on the flap (9).

Description

[0001] Cooling system for a rotating part of a double-bleed turbomachine

[0002] TECHNICAL FIELD

[0003] The present disclosure relates to the field of temperature control of rotating parts of turbomachines. More particularly, it relates to cooling systems for rotating parts of a compressor or a turbine, in particular within an aeronautical turbomachine.

[0004] STATE OF THE ART

[0005] An aeronautical propulsion system is known which comprises, from upstream to downstream in a normal direction of gas flow, a fan section, a compressor section which may comprise a low pressure compressor and a high pressure compressor, a combustion chamber and a turbine section which may comprise a high pressure turbine and a low pressure turbine.

[0006] When the propulsion system is in operation, the high-pressure compressor is rotated by the high-pressure turbine via a high-pressure shaft. The fan and, if applicable, the low-pressure compressor are rotated by the low-pressure turbine via a low-pressure shaft.

[0007] Part of the gases, called the "main flow", flows downstream of the fan through a primary flow which successively passes through the high-pressure then low-pressure compressors, the combustion chamber and the high-pressure and low-pressure turbines. Another part of the gases, called the "secondary flow", passes through a secondary flow whose inlet is located downstream of the fan, this secondary flow not passing through the compressors, the combustion chamber and the turbines, and joining the main flow directly at a gas ejection nozzle.

[0008] An aeronautical turbomachine compressor or turbine stage comprises a plurality of fixed blades and a plurality of moving blades. The moving blades form, with a rotor disk, a rotor stage of the compressor or turbine, a compressor or turbine being able to comprise a plurality of such stages. For a given stage, the moving blades are fixed to a radially external surface of the rotor disk. In use of the turbomachine, the temperature of certain parts of the disks, which are in contact with the hot gases flowing in the duct, increases significantly, and it is necessary to limit this increase to prevent the appearance of significant temperature gradients in the disks, which can weaken the material forming them. For this purpose, bores are provided in the disks in order to allow the passage of a cooling flow, colder than the gases flowing in the duct, so as to cool the rotor disks.

[0009] In order to allow the rotor discs of a high-pressure compressor to be cooled, one solution is to take air from a position in the turbomachine located upstream of the high-pressure compressor and therefore not having undergone the same temperature increase as the air flowing into it. This temperature increase is linked to the compression undergone by the flow at each stage of the high-pressure compressor.

[0010] For example, it is possible to provide a sampling orifice at a first position in an internal shell located in the radially internal part of the vein, at an intermediate stage of the low-pressure compressor, the colder air flowing, due to the pressure differential between the sampling orifice and an outlet orifice located downstream of the rotor discs which need to be cooled, through the bores of these discs. However, the use of a single path for conveying cold air to the discs does not allow the flow rate or temperature of the cooling air for the discs to be controlled.

[0011] Patent FR 2 690 482 B1 discloses a turbomachine in which ventilation means are provided comprising two separate sources for sampling cooling gas, located respectively upstream of the high-pressure compressor and at an intermediate stage of the high-pressure compressor, as well as conduit means for independently bringing the gases sampled at the first source and the gases sampled at the second source into the bores of the rotor disks. The conduit means comprise inlets in the primary stream, and pass through an intermediate casing located radially between the main and secondary streams before joining the secondary stream. A portion of each conduit means thus extends into the secondary stream and comprises a valve for regulating the flow rate in the conduit.These valves make it possible to control the temperature of the cooling flow, since this is obtained by combining the flows coming from the two conduit means, these being crossed by gases whose temperatures are different. Downstream of the valves according to the direction of flow of the cooling gases, the conduit means are reinjected into the primary vein. Patent FR 3 097 907 B1 also discloses a turbomachine comprising a means for conducting a cooling flow in which the flow rate of the flow is controlled by a valve located in the secondary vein.

[0012] Such cooling systems have significant drawbacks. Indeed, it is necessary for the conduit means to include a portion in the secondary flow path, the primary flow path not being wide enough to accommodate the valves. Thus, the conduit means have a significant length, because they must cross, from upstream to downstream, a portion of the primary flow path, the intermediate casing a first time, a portion of the secondary flow path, the intermediate casing a second time and the primary flow path before emerging on the flow path which crosses the bores of the rotor disks. This length of the conduit means creates aerodynamic losses which compromise the efficiency of the cooling flow at the moment when it crosses the bores.Furthermore, the fact that part of the conduit crosses the primary and secondary streams also causes aerodynamic losses in the primary and secondary flows themselves. The length of the conduits as well as the presence of valves also weigh down the cooling system. It should also be noted that the presence of valves that deteriorate over the flight cycles creates a weak point in the cooling system, which can ultimately compromise its robustness.

[0013] Recently, some aircraft engine manufacturers have developed turbomachines comprising hybridization, i.e. for which part of the power generated by a low-pressure turbine is used to power, via a low-pressure shaft mechanically driven by the low-pressure turbine, one or more electrical generators which can then power certain functions of the turbomachine. For this type of turbomachine, the cooling systems described above have another disadvantage: for lower engine speeds, the power draw from the low-pressure shaft can result in the flow pressure difference between the point at which the cold air is drawn and the point at which it is injected into the flow path being small, or even negative.Indeed, for an engine whose high-pressure spool speed is regulated, the power draw from the low-pressure shaft reduces the low-pressure spool speed at an equivalent high-pressure spool speed. This therefore results in a reduction in the compression ratio of the low-pressure compressors. As a result, the draw is carried out at a lower pressure, while the outlet point of the air thus drawn does not change, being located downstream of the low-pressure turbine. This therefore results in a reduction in the pressure differential between the draw point and the outlet point of the drawn air, which may cause a reversal of the direction of the cooling flow, in which case the air passing through the bores of the rotor discs comes from a section of the turbomachine located downstream of the discs, where the air is hotter than the discs, thus compromising the efficiency of the ventilation of the discs by the cooling flow.

[0014] STATEMENT OF THE INVENTION

[0015] There is therefore a need for a rotor disk cooling system for a rotating part of a turbomachine which allows efficient cooling of the disks even when the direction of flow of a cooling flow is reversed, particularly for lower engine speeds.

[0016] There is also a need for a rotor disk cooling system for a rotating part of a turbomachine that provides adequate cooling while minimizing air bleeds from the primary stream.

[0017] There is also a need for a rotor disk cooling system for a rotating part of a turbomachine that allows control of the temperature of a cooling flow with reduced aerodynamic losses within the cooling system itself.

[0018] There is also a need for a cooling system for rotor discs of a rotating part of a turbomachine with reduced aerodynamic losses in the main and / or secondary veins of the turbomachine.

[0019] To this end, according to a first aspect of the present disclosure, there is proposed a turbomachine compressor, comprising: at least one rotor stage provided with a rotor disk, and a cooling system comprising a pipe having an air bleed orifice located on an internal shell of the compressor, the cooling system comprising a flap and a return element, the flap being configured to adopt: a first position in which it allows, through the orifice at a first speed of the turbomachine, a flow of a first air flow, and a second position in which it allows, through the orifice at a second speed of the turbomachine higher than the first speed, a flow of a second air flow or a zero flow, the compressor being configured to allow a passage of the flap from the first position to the second position by a variation of a centrifugal force undergone by the flap.

[0020] The proposed cooling system thus allows precise regulation of the air flow taken to cool the rotor discs, including for low engine speeds for which the flow through the main pipe is reversed. The passive use of centrifugal forces to control the flow regulator eliminates the need for an active control valve, allowing a lighter cooling system, with lower aerodynamic losses because it is entirely contained in a portion of the compressor located radially between the machine axis and the internal shell.

[0021] According to one embodiment, the compressor has a main axis and comprises a plurality of orifices at different circumferential positions of the shell, preferably between four and twenty orifices.

[0022] According to one embodiment, the orifice is a secondary orifice and the pipe is a secondary duct, the cooling system comprises a main pipe having a main air sampling orifice located on the internal shell, the main orifice being located upstream of the secondary orifice in a normal direction of flow of the gases in the turbomachine.

[0023] According to one embodiment, the compressor has a main axis, and in the second position, at least a portion of a radially external surface of the flap is flush with a radially external surface of the shell, and / or at least a portion of a radially internal surface of the flap is flush with a radially internal surface of the shell.

[0024] According to one embodiment, the compressor has a main axis, and the flap comprises an overlap portion configured to abut a radially inner surface of the shell in the second position, the overlap portion preferably comprising a seal which is interposed between the radially inner surface of the shell and the overlap portion in the second position.

[0025] According to one embodiment, the compressor has a main axis, and the cooling system comprises between four and twenty flaps located at the same axial position.

[0026] According to one embodiment, the return element comprises a torsion spring.

[0027] According to one embodiment, the compressor comprises, from upstream to downstream in a normal flow direction of the gases in the turbomachine, at least one axial compressor stage and at least one centrifugal compressor stage. The present disclosure further relates to a turbomachine comprising a compressor as described above, in which a turbine is connected to the compressor via a shaft.

[0028] The present disclosure further relates to a method for cooling rotor disks of a turbomachine compressor, the method comprising: a step of cooling the rotor disks at a first speed of the turbomachine, comprising a flow, through a pipe having an air bleed orifice located on an internal shell of the compressor, of a first air flow rate, and a step of cooling the rotor disks at a second speed of the turbomachine, comprising a flow through the pipe of a second air flow rate or a zero flow rate, in which a flap passes, under the effect of a variation of a centrifugal force, from a first position in which the flap allows the flow of the first air flow rate to a second position in which the flap allows the flow of the second air flow rate or the zero flow rate, the flap being articulated against a return element.

[0029] According to one implementation of the method, the pipe is a secondary pipe and the orifice is a secondary orifice, and the step of cooling the rotor disks at the second speed comprises a flow of air through a main cooling pipe having a main air sampling orifice located on the shell, the main orifice being located upstream of the secondary orifice in a normal direction of flow of the gases in the turbomachine.

[0030] DESCRIPTION OF FIGURES

[0031] Figure 1a schematically represents an aircraft comprising turbomachines according to the present disclosure.

[0032] Figure 1b shows an axial sectional view of a turbomachine of the aircraft.

[0033] Figure 2 schematically represents a turbomachine compressor comprising a cooling system, according to a first position of a flap in a low speed configuration of the turbomachine.

[0034] Figure 3 schematically represents a flow regulator comprising the compressor flap in its first position. Figure 4 schematically represents the turbomachine compressor comprising the cooling system, according to a second position of the flap in a high speed configuration of the turbomachine.

[0035] Figure 5 schematically represents the flow regulator including the compressor flap in its second position.

[0036] Figure 6 schematically represents a specific embodiment of the compressor illustrated in Figures 2 and 4.

[0037] Figure 7 schematically represents a specific embodiment of the shutter illustrated in Figures 3 and 5.

[0038] Figure 8 schematically represents a specific embodiment of the shutter illustrated in Figures 3 and 5.

[0039] In the following, similar elements are designated by the same reference signs.

[0040] DETAILED DESCRIPTION OF EMBODIMENTS

[0041] Figure 1a represents an aircraft 100, equipped with a fuselage 101 and wings 102. One or more turbomachines 1 are fixedly mounted on the wings 102 of the aircraft 100.

[0042] The turbomachine 1, shown schematically in Figure 1b, comprises, from upstream to downstream with reference to a normal direction of gas flow, a compressor, a combustion chamber and a turbine mechanically connected to the compressor by a shaft, so that the turbine can mechanically drive the compressor, as well as a gas exhaust nozzle. The turbomachine may be, as in this case, a twin-spool turbomachine, comprising, from upstream to downstream, a fan 106 and possibly a low-pressure compressor 110, a high-pressure compressor 103, a combustion chamber 104, a high-pressure turbine 105 and a low-pressure turbine 107, and the exhaust nozzle. A portion of the gases, called the primary flow, then flows through the compressors and the turbines in a primary flow path 18, while the remainder of the gases joins a secondary flow path 19 downstream of the fan through which it directly joins the exhaust nozzle.In this case, the compressor 10 and the shaft 8 considered in the following description may be, respectively, either the low pressure compressor 110 and the low pressure shaft 109 connecting the low pressure turbine 107 to the low pressure compressor 110, or the high pressure compressor 103 and the high pressure shaft 108 connecting the high pressure turbine 105 to the high pressure compressor 103. In all that follows, and in the absence of a contrary mention, the terms “upstream” and “downstream” will be used with reference to the normal direction of flow of the gases in the turbomachine.

[0043] The compressor 10 is shown schematically in Figure 2. It extends around a compressor axis X, concentric with the compressor shaft 8. The compressor 10 comprises at least one rotor stage, which comprises a rotor blade 6 mounted on a rotor disk 5. The rotor blade 6 extends in the primary vein 18, which is defined between an inner shroud 20 and an outer shroud 21. The rotor disk 5 therefore extends radially internally of the primary vein 18. Advantageously, the compressor 10 may comprise static blades 7, also called rectifiers or distributors, arranged axially between the rotor blade stages 6.

[0044] The compressor 10 may comprise a main air sampling orifice 2, which passes through the internal shell 20. This main orifice 2 constitutes an inlet of a main cooling pipe, which is configured to bring a first cooling flow S1 taken from the primary vein 18 through the main orifice 2 through bores 22 provided in one or more rotor discs 5, thus making it possible to ventilate them.

[0045] According to one embodiment, the compressor 10 is a high-pressure compressor and the main air bleed port 2 is arranged axially between the low-pressure compressor and the high-pressure compressor 10, so that the first cooling flow passes through the bores 22 of the discs 5 of all the rotor blades 6 of the high-pressure compressor 10. Alternatively, the main air bleed port 2 is located further downstream of the low-pressure compressor, between a row of rotor blades 6 and a row of stator blades 7 of the high-pressure compressor 10 or between a row of stator blades 7 and a row of rotor blades 6, this order of enumeration referring to the normal direction of flow of the gases in the high-pressure compressor 10.For example, the main port 2 may be located between the first stage and the second stage of the high pressure compressor 10, each stage consisting of a row of rotor blades 6 and an adjacent row of stator blades 7.

[0046] The high-pressure compressor 10 comprises a secondary air sampling orifice 3, which passes through the inner shell 20. The secondary orifice 3 is located downstream of the main orifice 2. In particular, the secondary orifice 3 may be located downstream of the high-pressure axial compressor 10, or between the last and penultimate stages of the high-pressure axial compressor as illustrated schematically in FIG. 2. It constitutes an inlet of a secondary cooling pipe, which is configured to bring a second cooling flow S2 into contact with the bores 22 of the discs 5 of the rotor blades 6.

[0047] The secondary orifice 3 is provided with a flow regulator 4, shown in more detail in Figure 3. The flow regulator 4 comprises a flap 9 pivotally mounted on an axis 27 arranged on or near one of the two ends of the secondary orifice 3. In addition, the regulator 4 comprises a return element 11 which exerts a return force so as to return the flap 9 to a nominal position in which it is located away from the secondary orifice 3, and therefore allows the passage of a non-zero flow rate of the second cooling flow S2.

[0048] The compressor 10 may comprise several flow regulators 4 distributed circumferentially around the motor axis X. For example, the compressor may comprise at least four flow regulators 4, preferably between four and twenty. Providing at least four flow regulators 4 ensures uniformity of the flow of the second cooling flow S2 through the compressor 10. Furthermore, a regular circumferential distribution of the flow regulators 4 around the motor axis X makes it possible not to create an imbalance and therefore does not significantly impact the balancing of the compressor 10.

[0049] The secondary orifice 3 may, according to certain embodiments, extend over the entire circumference of the internal ferrule 20.

[0050] According to one embodiment, the return element 11 is a torsion spring. The return element 11 then exerts a return torque on the rotation axis 27 of the flap 9.

[0051] Referring to Figure 2, at a first speed of the turbomachine, the secondary orifice 3 and where appropriate the main orifice 2 both allow the passage of a cooling flow. The first speed is a low speed, in which the compression ratio across the different stages of the compressor 10 is relatively low. As a result, the flow velocity of the first cooling flow S1 is also relatively low. For certain types of turbomachines that fall within the scope of the present disclosure, it is possible that the flow velocity in the primary stream is not sufficient to compensate for the static pressure difference between the main orifice 2 and an outlet point of the main cooling pipe located downstream of the main orifice 2, so that a reversal of the flow direction takes place.The first cooling flow S1 then flows towards the main orifice 2 from upstream, and therefore comes from a position located downstream in the primary vein 18. Thanks to the presence of the secondary orifice 3, the first cooling flow S1 is then merged with the second cooling flow S2, and corresponds to a part of the second flow S2 reinjected into the primary vein 18 through the main orifice 2. It allows, despite the inversion of flow direction, adequate ventilation of the discs 5.

[0052] At this first speed, the flap 9 of the flow regulator 4 is positioned away from the secondary orifice 3 under the action of the return element 11. A second non-zero cooling flow S2 can therefore flow into the second pipe. In the event that a reversal of the flow direction of the first cooling flow S1 takes place, the second flow S2 - in the extension of which the first flow S1 is then located - can flow upstream from the secondary orifice 3 thanks to the static pressure at the secondary orifice 3, which is higher than the static pressure further upstream. The second flow S2 can ventilate the rotor discs 5 located downstream of the secondary orifice 3. Furthermore, the second flow S2 flows through all the bores 22 of rotor discs 15 located axially between the main orifice 2 and the secondary orifice 3.Ventilation of the discs 5 is therefore ensured, despite the inversion of the flow direction of the first flow S1 which is then in the extension of the second flow S2. This makes it possible to reduce the thermal gradient across the bores 22 of the rotor discs 5, and thus to extend their service life.

[0053] The specific choice of the position, in the axial direction, of the secondary orifice 3, is a compromise between, on the one hand, the temperature of the fluid taken, which increases from upstream to downstream in the compressor, and which it is desired to maintain at a sufficiently low value to allow effective cooling of the rotor discs 5, and on the other hand, the static pressure of the fluid taken, which also increases from upstream to downstream as the air is compressed by the moving blades 6. However, it is desirable for this pressure to be sufficiently high to allow a flow speed of the second flow S2 which also allows effective ventilation of the rotor discs 5.

[0054] The exact flow rate allowed by the flow regulator 4 at a given speed of the turbomachine can be controlled during the design of the compressor 10 by modifying in particular the dimension of the flap 9, which affects the centrifugal force seen by it at said speed, as well as the stiffness of the return element 11. In particular, it is necessary to provide a return element 11 sufficiently stiff to allow the flap 9 to be kept in the open position at the first speed of the turbomachine 1. It is also necessary to take into account the inclination of the internal shell 20 at the level of the flow regulator 4, which also affects the centrifugal force seen by the flap 9 - the more the flap 9 is perpendicular to the engine axis X in its closed position, the less the centrifugal forces seen by the flap 9 can contribute to bringing it back to this position.

[0055] According to one embodiment, the secondary air sampling orifice 3 is located downstream of the row of rotor 6 or stator 7 blades furthest downstream of the high-pressure compressor 10. Thus, in the event of reversal of the flow direction of the first flow S1, the second flow S2 can ventilate all the rotor discs 5 located between the main 2 and secondary 3 orifices.

[0056] With reference to figures 4 and 5, at a second speed of the turbomachine, higher than the first speed, the centrifugal forces exerted on the flap 9 and tending to return it to the closed position are sufficiently high for the flap 9 to close the secondary orifice 3, despite the restoring force exerted by the restoring element 11 on the flap 9. At this second speed, the second flow rate S2 is not necessary: ​​due to the higher engine speed, the flow of the first flow rate S1 is not reversed, and the air flows from upstream to downstream passing through the bores 22 of the rotor discs 5, thus allowing their ventilation.

[0057] Alternatively, it is possible to provide that the second speed of the turbomachine corresponds to a speed for which the centrifugal forces acting on the flap 9 are higher than for the first speed, while remaining below a value for which the flap 9 comes into a fully closed position. In other words, at the second speed, the flap 9 adopts a position in which it allows a flow of the second flow rate S2 whose flow rate is lower than the flow rate of the second flow rate S2 at the first speed of the turbomachine, while remaining non-zero. It is particularly advantageous for the flow rate of the second flow rate S2 to be non-zero at the second speed when the secondary orifice 3 is located further downstream of the compressor 10 so as to increase the dynamic pressure of the flow in the primary stream 18 at the secondary orifice, as described previously.Indeed, for such a position of the secondary orifice, a complete closure of the flap 9 could cause undesirable aerodynamic turbulence in the primary vein 18.

[0058] More generally, it is possible for the turbomachine 1 to be configured to operate at a first speed, for which the flap 9 is in an open position, a second speed for which the flap 9 is in a fully closed position, and intermediate speeds between the first and second speeds for which the flap 9 adopts intermediate positions for which the flow rate of the second flow rate S2 gradually decreases with increasing speed.

[0059] The cooling system thus formed, with its two cooling ducts, has several advantages. It allows adequate cooling of the rotor disks 5 at all speeds of the turbomachine between the first and second speeds, without it being necessary to pass one or more ducts through the primary vein 18 and / or the secondary vein 19, the main and secondary ducts being entirely included in the space radially internal relative to the internal shell 20. Furthermore, in comparison with the solutions of the prior art for which a flow control valve is necessary for some of the ducts, the proposed system is lighter and has better robustness due to the absence of such a valve.In addition, the portions of the cooling pipes of the proposed system located upstream of the rotor disks 5 according to the flow of gases in these pipes can have a shorter length. This reduces the aerodynamic losses in these pipes.

[0060] Taking air for the secondary cooling line from the primary vein 18 also makes it possible to limit pollution of the flow by particles (for example, sand) which could compromise the quality of the flow in the case of taking air from the external vein 19.

[0061] Furthermore, taking samples from the main stream penalizes the thermodynamic cycle of the turbomachine, and induces excess fuel consumption, which increases when the sample is taken further downstream of the compressor. The compressor

[0062] 10 proposed makes it possible to limit these overconsumptions when the turbomachine speed is high, that is to say when the second flow S2 does not contribute or contributes little to the ventilation of the discs 5.

[0063] The regulation of the flow rate of the second flow S2 is entirely ensured by a passive system, using the variation of the centrifugal forces seen by the flap 9, which allows a very low or even zero impact of this regulation on the fuel consumption of the turbomachine 1.

[0064] It may be desirable, for certain configurations of the turbomachine 1, to prevent the second cooling flow S2 from flowing downstream from the secondary orifice 3. Indeed, when the temperature of the air at the secondary orifice 3 is particularly high, and when the rotor disks 5 located between the main orifice 2 and the secondary orifice 3 are cold - for example when starting the turbomachine - the second flow S2 could heat these disks 5. According to certain embodiments, and as shown in FIG. 6, a cover 28 is therefore provided so as to allow the flow of the second flow S2 from the secondary orifice 3 only downstream, and to prevent it from returning upstream.A spacing 29 is provided between the cover 28 and the internal vein 20 so as not to obstruct the passage of the first flow S1 towards the compressor stages located downstream of the secondary orifice 3 for the speeds of the turbomachine 1 at which the flow of the first flow S1 is not reversed.

[0065] According to certain embodiments of the flow regulator 4, illustrated in Figures 3 and 5, in the fully closed position of the flap 9, the flap 9 is flush with the inner shell 20. A radially outer surface 13 of the flap 9 may be flush with a radially outer surface 15 of the inner shell 20, or a radially inner surface 12 of the flap 9 may be flush with a radially inner surface 14 of the inner shell 20. Preferably, each of the radially outer 13 and radially inner 12 surfaces of the flap is flush with the corresponding surface of the inner shell 20. Thus, there is no significant discontinuity between the flap 9 and the inner shell 20 that could cause undesirable aerodynamic losses.

[0066] Alternatively, and as shown in Figure 7, the flap 9 may have an overlap portion 16 which abuts the radially inner surface 15 of the inner shell 20 in the fully closed position of the flap 9. The overlap portion 16 makes it possible to ensure better sealing of the flap 9 when the latter is in its fully closed position. The overlap portion 16 may optionally comprise, on the radially outer surface 13 of the flap 9, a seal 17, which also contributes to improved sealing of the flap 9.

[0067] The geometry of the flap may be configured to combine the two embodiments described above, i.e. to have an overlapping portion 16 and so that the radially external surface 13 of the flap 9 is flush with the radially external surface 15 of the internal shell 20, and / or so that the radially internal surface 12 of the flap 9 is flush with the radially internal surface 14 of the internal shell 20 in the fully closed position of the flap 9.

[0068] According to one embodiment, and with reference to Figure 8, a downstream edge 31 of the secondary orifice 3 is offset in the radial direction, in a direction opposite to the engine axis X, so as to form an air intake scoop. This means that the internal shell 20 has, near the secondary orifice 3, a downstream edge 31 which moves closer, radially, to the engine axis X as one moves away, axially, from the secondary orifice 3, thus forming the intake scoop. In its closed position, the flap 9 is then located away from this downstream edge 31, allowing a non-zero air flow to flow. The intake scoop makes it possible to capture a greater dynamic pressure of the second flow S2, advantageously increasing the flow rate of this second flow.

[0069] According to other aspects, the present disclosure relates to a compressor 10 and a turbomachine 1 comprising this compressor 10, the compressor being provided with a cooling system comprising the main cooling line and the secondary cooling line as described previously.

[0070] According to one embodiment, the high-pressure compressor 10 is an axial compressor provided with a plurality of rotor stages with rotor blades 6 mounted on rotor discs 5, each rotor stage being located axially upstream of a stator blade stage 7. In this case, the secondary orifice 3 may be located at an axial position of the compressor 10 intermediate between two successive blades, or else be located downstream of the axial compressor 10, that is to say downstream of the row of stator blades 7 furthest downstream of the compressor.

[0071] According to another embodiment, the compressor 10 is of the axial-centrifugal type. Such a compressor comprises axial compressor stages, each provided with a rotor stage with at least one rotor blade 6 mounted on a rotor disk 5, and a static blade stage 7. Downstream of the axial compressor stages, the axial-centrifugal compressor 10 comprises one or more centrifugal compressor stages (not shown). In this case, the secondary orifice 3 is advantageously positioned, axially, between the most downstream blade of the axial compressor stages and the most upstream centrifugal compressor stage. Such a positioning of the secondary orifice 3 is advantageous: the axial-centrifugal compressor 10 comprises, between the axial compressor and centrifugal compressor stages, an elbow without blades which constitutes a relatively long space in the axial direction, and which makes it possible to house the secondary orifice 3 and the flow regulator 4.

[0072] According to other aspects, the present disclosure relates to a method for cooling rotor disks 5 of a compressor 10. The method is implemented by means of a cooling system as described previously. The method comprises, at a first speed of the turbomachine, a first step of cooling the rotor disks 5 of the compressor 10 by flowing a first air flow from the internal vein 18 through the secondary orifice 3 and through the secondary cooling duct. The method also comprises, at a second speed higher than the first speed, a second step of cooling the rotor disks 5 by flowing a second air flow from the internal vein 18 through the secondary orifice 3 and through the secondary cooling duct, the second flow rate being lower than the first flow rate.The flow rate through the secondary orifice 3 is reduced, between the first and second regimes, by the at least partial closing of the flap 9 due to the increase in the centrifugal forces to which it is subjected. The air flows through the secondary duct make it possible to ventilate the rotor discs 5 by passing through bores 22 provided for this purpose in the discs 5, along the secondary duct. The cooling method also comprises, at least for the second cooling step at the second regime of the turbomachine 1, a flow of air coming from the internal vein 18 through the main orifice 2 and through the main cooling duct, thus contributing to the ventilation of the rotor discs 5 by passing through the bores 22 provided for this purpose. According to certain embodiments, the second air flow rate is zero.In this case, at the second speed of the turbomachine 1, the flow regulator 4 does not allow any air flow through the secondary orifice 3, and only the main flow through the main duct contributes to the cooling of the rotor discs 5.

Claims

CLAIMS 1. Turbomachine compressor (10), comprising: at least one rotor stage provided with a rotor disk (5), and a cooling system comprising a secondary duct having a secondary air sampling orifice (3) located on an internal shell (20) of the compressor (10), the cooling system comprising a flap (9) and a return element (11), the flap (9) being configured to adopt: a first position in which it allows, through the secondary orifice (3) at a first speed of the turbomachine, a flow of a first air flow, and a second position in which it allows, through the secondary orifice (3) at a second speed of the turbomachine higher than the first speed, a flow of a second air flow or a zero flow, the compressor being configured to allow a passage of the flap (9) from the first position to the second position by a variation of a centrifugal force undergone by the flap (9),in which the cooling system further comprises a main pipe having a main air sampling orifice (2) located on the internal shell (20), the main orifice (2) being located upstream of the secondary orifice (3) in a normal direction of flow of the gases in the turbomachine., 2. Turbomachine compressor according to claim 1, having a main axis, the compressor comprising a plurality of secondary orifices (3) at different circumferential positions of the shell (20), preferably between four and twenty secondary orifices (3).

3. Turbomachine compressor according to either of claims 1 and 2, having a main axis, compressor in which, in the second position, at least a portion of a radially external surface (13) of the flap is flush with a radially external surface (15) of the shell (20), and / or at least a portion of a radially internal surface (12) of the flap is flush with a radially internal surface (15) of the shell (20).

4. Turbomachine compressor according to any one of claims 1 to 3, having a main axis, compressor in which the flap (9) comprises an overlap portion (16) configured to come into abutment with a radially internal surface (15) of the shell (20) in the second position, the overlap portion (16) preferably comprising a seal (17) which is interposed between the radially inner surface (14) of the ferrule (20) and the overlap portion (16) in the second position.

5. Turbomachine compressor according to any one of claims 1 to 4, having a main axis, compressor in which the cooling system comprises between four and twenty flaps (9) located at the same axial position.

6. Turbomachine compressor according to any one of claims 1 to 5, the return element (11) comprising a torsion spring.

7. Turbomachine compressor according to any one of claims 1 to 6, comprising, from upstream to downstream in a normal flow direction of the gases in the turbomachine (1), at least one axial compressor stage and at least one centrifugal compressor stage.

8. Turbomachine (1) comprising a compressor (10) according to any one of claims 1 to 7, in which a turbine (105, 107) is connected to the compressor (10) via a shaft (108, 109).

9. Method for cooling rotor disks (5) of a compressor (10) of a turbomachine (1), the method comprising: a step of cooling the rotor disks (5) at a first speed of the turbomachine (1), comprising a flow, through a secondary pipe having a secondary air bleed orifice (3) located on an internal shell (20) of the compressor, of a first air flow rate, and a step of cooling the rotor disks (5) at a second speed of the turbomachine (1), comprising: a flow through the pipe of a second air flow rate or a zero flow rate, and an air flow through a main cooling pipe having a main air bleed orifice (2) located on the shell (20), the main orifice (2) being located upstream of the secondary orifice (3) in a normal direction of flow of the gases in the turbomachine, in which a flap (9) passes, under the effect of a variation of a centrifugal force,from a first position in which the flap (9) allows the flow of the first air flow to a second position in which the flap (9) allows the flow of the second air flow or the zero flow, the flap being articulated against a return element (11).,

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