Cooling-air injection casing for cooling a rotor disc of a turbine and comprising bypass passages

Abstract

The invention relates to a cooling-air injection casing (1) for cooling a rotor disk of a turbine, which casing extends about a longitudinal axis and comprises a radially outer wall (4), a radially inner wall (5) and a plurality of channels (6a, 6b) distributed about the longitudinal axis, each channel forming an air injector extending axially from an inlet port (7) opening through the radially outer wall, to an outlet port (8) positioned at a downstream end (42) of the radially inner wall, each channel being configured to guide a first air flow from an annular space for bypassing a combustion chamber to a ventilation cavity formed between the disc and a sealing flange, the casing (1) comprising at least one bypass passage (11) extending radially between two channels and configured to divert a portion of a second air flow coming from a high-pressure compressor and guide it to a purge cavity.

Description

Cooling air injection housing for a turbine rotor disc, including bypass passages Technical Field

[0001] The present invention relates to the field of ventilation of a high-pressure turbomachine turbine and more particularly to an air injection housing for cooling a rotor disc of such a turbine. Previous technique

[0002] A turbomachine includes a high-pressure turbine which is positioned at the outlet of a combustion chamber to recover energy from the combustion gas flow and thus drive in rotation, a high-pressure compressor, located upstream of the combustion chamber and supplying the latter with pressurized air.

[0003] In the following description and claims, the terms "upstream" and "downstream" are to be taken into consideration with respect to the direction of air flow inside the high-pressure turbine, as well as inside the cooling air injection housing.

[0004] Typically, a high-pressure turbine comprises a rotor disc, positioned at the outlet of a combustion chamber and on which turbine blades are mounted, driven into rotation by a flow of gas ejected from this combustion chamber.

[0005] Due to the high temperatures reached by the combustion gases, the rotor disc and the turbine blades it supports are subjected to significant thermal stresses that can induce expansion. To limit the negative impact of these thermal stresses on the lifespan of the turbine blades, the latter are equipped with internal cooling circuits that include ducts through which ventilation air is drawn from the bottom of the combustion chamber.

[0006] This ventilation air is typically supplied to an annular cavity by ventilation air injectors distributed circumferentially around the longitudinal axis of the turbomachine. The injectors are arranged around a cooling air injection housing, extend radially beneath the combustion chamber, and are fluidically connected to an annular cavity that carries ventilation air from the bottom of the compressor to the turbine.

[0007] The ventilation air, exiting the injectors, enters the annular cavity located upstream of the rotor disc, passing through orifices formed in a sealing flange positioned upstream of the rotor disc. The cavity communicates with internal cooling circuits located inside the turbine blades.

[0008] Documents FR2943092 and W02023047055 describe examples of high-pressure turbines and blades.

[0009] Part of the air from the injector also circulates to an upstream purge, first passing through a downstream seal which ensures the seal between the blade and the flange at the location of an external part of the turbine.

[0010] Simultaneously, an airflow taken downstream of the last stage of the high-pressure compressor circulates through a first upstream seal and then through a second upstream seal, positioned downstream of the first upstream seal, ensuring the seal between the blade and the flange at the location of an internal part of the turbine.

[0011] Typically, the first and second upstream seals are labyrinth seals comprising flaps attached to the rotor and a ring made of abradable material connected to the housing, featuring a honeycomb structure. This airflow, drawn downstream of the last stage of the high-pressure compressor, mixes with the airflow from the injectors before venting the upstream purge and the blades.

[0012] The airflows are injected tangentially with respect to the longitudinal axis. The tangential velocity of the air circulating in the internal circuits of the turbine is a A key parameter for its cooling is the speed of the tangentially injected airflow, which is closer to the rotor's rotational speed. The closer the speed of the tangentially injected airflow is to the rotor's rotational speed, the better the cooling. The airflow exiting the injectors has a significant tangential speed that is equal to, or even greater than, the rotor's rotational speed for the same radius.

[0013] However, the airflow from the first and second upstream seals has, on the contrary, a low tangential velocity and a higher temperature. This results in a decrease in the effectiveness of this airflow in cooling the upstream purge. Description of the invention

[0014] The invention therefore aims to resolve at least in part these drawbacks by proposing a cooling air injection housing for a turbine rotor disc that allows the upstream purge to be cooled more efficiently.

[0015] The invention relates to a cooling air injection housing for a bladed rotor disc of a turbine, in particular a high-pressure turbine, of a turbomachine.

[0016] The casing extends around a longitudinal axis and comprises a radially external wall, a radially internal wall and several channels distributed circumferentially around the longitudinal axis, each channel forming an air injector extending axially and at least partially along the radially internal wall from an inlet opening through the radially external wall to an outlet opening through a downstream end wall, each channel being configured to guide a first airflow from an annular bypass space of a turbine combustion chamber to a ventilation cavity formed between the rotor disc and a sealing flange positioned upstream of the rotor disc.

[0017] The housing includes at least one bypass passage. Each bypass passage is positioned between two circumferentially adjacent channels and extends in a generally radial direction, from a radially internal inlet to a radially external outlet. The passage of The bypass is configured to divert a portion of air from a second airflow originating from a high-pressure compressor of the turbine and guide said portion of the second airflow to a purge cavity of the turbine.

[0018] The invention thus provides a cooling air injection housing for a turbine rotor disc, enabling the upstream purge and blades to be cooled more efficiently.

[0019] The bypass passage allows the air portion of the second airflow to bypass the channel outlet and arrive at a high tangential velocity in the turbine purge cavity. The tangential velocity of this air portion corresponds to approximately 20% of the rotor speed.

[0020] The temperature of the air intended to cool the upstream purge is reduced by approximately 15°C compared to prior art solutions as described in document W02023047055A.

[0021] Furthermore, this housing configuration eliminates the first upstream seal, specifically a disc, flanges, an abradable cartridge, and an abradable support, resulting in a weight reduction. Removing the disc also reduces the tensile forces applied to the rotor shaft.

[0022] This solution also minimizes the air flow taken from the high-pressure compressor in order to improve the overall performance of the turbomachine.

[0023] In some embodiments, the housing includes several bypass passages arranged circumferentially around the longitudinal axis.

[0024] In some embodiments, circumferentially adjacent channels are separated by inter-channel connecting sections, each inter-channel connecting section being crossed by a bypass passage.

[0025] In some embodiments, each bypass passage has a circumferentially measured length that is greater than its axially measured width.

[0026] In some embodiments, each bypass passage is delimited axially by an upstream wall connected to an upstream portion of the radially internal wall and a downstream wall, opposite the upstream wall, and connected to a downstream portion of the radially internal wall, each bypass passage being delimited circumferentially by a first circumferential end wall of a first channel and a second circumferential end wall of a second channel.

[0027] In some embodiments, the housing comprises a first abradable element which extends radially inward from a first radially internal face of the upstream portion of the radially internal wall and intended to cooperate with at least one first lip of the sealing flange to form a first sealing device, and a second abradable element which extends radially inward from a second radially internal face of the downstream portion of the radially internal wall and intended to cooperate with at least one second lip connected to the first lip to form a second sealing device, the radially internal inlet of each bypass passage opening between the first and second abradable elements.

[0028] In some embodiments, the housing includes a downstream end ring which extends radially outward from the radially internal wall and which is delimited axially by the downstream end wall and the downstream wall, and radially, on the inside, by the downstream portion of the radially internal wall and, on the outside, by a radially external ring wall, opposite the downstream portion of the radially internal wall.

[0029] In some embodiments, the radially external wall of the crown includes a third abradable element which extends radially outwards and is intended to cooperate with a third lip connected to a radially external part of the sealing flange to form a third sealing device.

[0030] In some embodiments, the radially internal wall is connected to the downstream portion of the radially internal wall by connecting channel portions, each portion of connecting channel circumferentially separating two bypass passages.

[0031] In some embodiments, the radially internal wall comprises an upstream portion along which a primary section of the channel extends from the inlet opening and a downstream portion along which a secondary section of the channel extends to the outlet opening, the upstream portion being connected to the downstream portion by a curved portion.

[0032] In some embodiments, the bypass passages are positioned axially upstream of the curved portion.

[0033] The invention also relates to a turbine for a turbomachine comprising a casing as defined above.

[0034] The aforementioned features and advantages, as well as others, will become apparent upon reading the detailed description that follows, along with examples of housing construction. This detailed description refers to the attached drawings. Brief description of the drawings

[0035] The attached drawings are schematic and are primarily intended to illustrate the principles of the presentation.

[0036] In these drawings, from one figure to another, identical elements (or parts of elements) are identified by the same reference symbols.

[0037] [Fig. 1] Figure 1 schematically represents an axial cross-sectional view of a turbomachine according to the invention;

[0038] [Fig. 2] Figure 2 schematically represents an axial cross-sectional view of part of a high-pressure turbine of the turbomachine including a cooling air injection housing, according to an embodiment of the invention;

[0039] [Fig. 3] Figure 3 schematically represents a perspective view of an angular portion of the crankcase;

[0040] [Fig. 4] Figure 4 schematically represents another perspective view of an angular portion of the crankcase;

[0041] [Fig. 5] Figure 5 schematically represents a view of the inside of the casing of figure 4;

[0042] [Fig. 6] Figure 6 schematically represents the inside of a channel in the crankcase along an axial section. Description of the implementation methods

[0043] To make the explanation more concrete, an example of a housing 1 is described in detail below, with reference to the attached drawings. It should be noted that the invention is not limited to this example.

[0044] The invention applies to turbines 3, in particular high-pressure turbines, of a twin-spool turbomachine, such as the high-pressure turbine of an aircraft turbojet 40 shown in Figure 1. The turbojet 40 extends around a longitudinal axis X, corresponding to the axis of revolution of the turbojet 40 and the axis of rotation of the rotor.

[0045] The turbojet 40 comprises, from left to right, i.e. from upstream to downstream with reference to the gas flow during operation in the turbomachine: a fan 37, a high-pressure compressor 38, a combustion chamber 36, the high-pressure turbine 3 and a low-pressure turbine 3'. The high-pressure turbine 3 is equipped with blades 39.

[0046] As is known, the high-pressure turbine 3 is positioned at the outlet of the combustion chamber 36 to recover energy from a flow of combustion gases from this chamber and drive in rotation the high-pressure compressor 38 located upstream of the chamber and supplying the latter with pressurized air.

[0047] The high-pressure turbine 3 comprises a rotor disc 2 centered on the longitudinal axis X and which is disposed at the outlet of the combustion chamber 36. The turbine blades 39 are mounted on the rotor disc 2 and driven in rotation by the gas flow ejected from this combustion chamber 36.

[0048] The turbine 3 also includes a sealing flange 10 which is centered on the longitudinal axis X and arranged upstream of the rotor disc 2. This sealing flange 10 is movable in rotation and rotates with the rotor disc 2.

[0049] Furthermore, the sealing flange 10, together with the rotor disc 2, defines an annular ventilation cavity 9 designed to receive ventilation air and direct it to internal cooling circuits for the blades 39. For this purpose, channels 6a, 6b, forming ventilation air injectors, are regularly distributed around the longitudinal axis X, following the circumferential direction. The channels 6a, 6b are connected upstream to an annular bypass space 12 around the combustion chamber 36, also called the combustion chamber bottom, to convey ventilation air from the high-pressure compressor 38 to the ventilation cavity 9. The annular space 12 extends around the combustion chamber 36.

[0050] As shown in Figures 2 to 6, the turbine 3 includes an annular cooling air injection housing 1 extending around the longitudinal axis X and comprising a radially external wall 4 and a radially internal wall 5 which acts as a stiffener. The housing 1 is fixed to the stator of the turbine 3. The sealing flange 10 is rotatably mounted relative to the stator. The housing 1 is positioned between the annular space 12 and the sealing flange 10.

[0051] The radially external wall 4 has the function of separating the annular bypass space 12 located radially below the combustion chamber 36 from cavities, such as the ventilation cavity 9, intended to cool the blades 39.

[0052] The radially internal wall 5 extends downstream from the radially external wall 4. The radially internal wall 5 is connected to the radially external wall 4 at an upstream end 60 of the housing 1. More precisely, the radially internal wall 5 comprises an upstream portion 25, which extends along the longitudinal axis X from an internal face 52 of the radially external wall 4 to a curved portion 29, inclined relative to the upstream portion 25. The curved portion 29 extends downstream from the upstream portion 25. The radially internal wall 5 includes a downstream portion 26, which extends from the curved portion 29 along the longitudinal axis X and downstream.

[0053] Each channel 6a, 6b extends axially, that is to say globally along the direction of the longitudinal axis X from an inlet opening 7 opening through the radially external wall 4 to an outlet opening 8 positioned at a downstream end 42 of the housing 1, opposite the upstream end 60. Each channel 6a, 6b extends along a radially external face 23 of the radially internal wall 5, opposite a first radially internal face 15a of the radially internal wall 5.

[0054] The housing 1 includes a downstream end ring 30 positioned at the downstream end 42 of the housing 1. The downstream end ring 30 is delimited axially by a downstream end wall 46 and an upstream wall 59, opposite the downstream end wall 46, and radially by the downstream portion 26 of the radially internal wall 5 and a radially external ring wall 33, opposite the downstream portion 26. The outlet openings 8 of the channels 6a, 6b open through the downstream end wall 46. The downstream end wall 46 and the upstream wall 59 extend radially, that is, in a radial direction Z, orthogonal to the longitudinal axis X.

[0055] The housing 1 includes a flange 32 extending radially from an upstream end 60 of the radially internal wall 5 and towards the longitudinal axis X.

[0056] The downstream end ring 30 is positioned opposite the flange 32.

[0057] The radially external wall of the crown 33 is annular and positioned radially above a portion of canals 6a, 6b. It extends axially for a length equivalent to approximately 1 / 3 of the length of canals 6a, 6b, for example. The radially external wall of the crown 33 extends axially from the downstream end wall 46 to the upstream wall 59.

[0058] Each channel 6a, 6b is traversed by a first ventilation airflow DI from the annular space 12 bypassing the combustion chamber 36 to supply air to the ventilation cavity 9.

[0059] More specifically, the ventilation air exiting the channels 6a, 6b enters the ventilation cavity 9 through orifices 41 formed in the sealing flange 10. The ventilation cavity 9 communicates with the internal cooling circuits arranged inside the blades 39.

[0060] As shown in Figure 5, each channel 6a, 6b comprises a primary segment I extending along the upstream portion 25 of the radially internal wall 5 along the longitudinal axis X from the inlet opening 7 to a bend 61 positioned in the curved portion 29 of the radially internal wall 5 and a secondary segment 28 extending along the downstream portion 26 of the bend 61 to the outlet opening 8.

[0061] Channels 6a, 6b are inclined as described in application W02023047055.

[0062] The channels 6a, 6b include a bottom 34 delimited by the radially internal wall 5.

[0063] The inlet opening 7 has an oblong section, preferably rectangular, and a central axis parallel to the longitudinal axis X. The outlet opening 8 also has an oblong section, preferably rectangular.

[0064] As shown in Figure 6, the secondary section 28 exhibits a progressive variation in its orientation along a tangential direction Y between the section of the elbow 61 and the section of the outlet opening 8. The tangential direction Y is perpendicular to the longitudinal axis X and to the radial direction Z. By "progressive variation in orientation", we mean a variation in the orientation of a vector normal to the center of a section of the channel 6a, 6b and originating from the center of said section.

[0065] Preferably, the outlet section of the outlet opening 8 of each channel 6a, 6b extends tangentially in a plane perpendicular to the longitudinal axis X. In addition, the elbow 61 is advantageously oriented so that the airflow exiting the outlet opening 8 flows tangentially in the same direction as the direction of rotation of the rotor disc 2 facing it.

[0066] Each channel 6a, 6b includes a neck 31 forming a reduction in cross-section. The neck 31 corresponds to the point in channel 6a, 6b with the smallest cross-section. The neck 31 extends from the elbow 61 to the outlet opening 8.

[0067] The throat 31 forms a calibrating section in the sense that it is this section that calibrates the flow rate through channels 6a and 6b. The throat 31 also has an orientation with a tangential component relative to the longitudinal axis X. This allows the air passing through channels 6a and 6b to obtain a tangential velocity close to the rotor's rotational speed. The more inclined the throat 31, the higher the tangential velocity. The cross-sectional ratio between the inlet opening 7 and the throat 31 is at least 2.

[0068] The downstream end crown 30 is axially traversed by a portion of the channels 6a, 6b and includes at least part of the neck 31.

[0069] The radially internal wall 5 includes a first annular abradable element 16 intended to cooperate with at least one first lip 19 connected to an internal part 22 of the sealing flange 10 to form a first sealing device. More precisely, the first abradable element 16 is supported by a first radially internal face 15a of the upstream portion 25 of the radially internal wall 5.

[0070] The radially internal wall 5 also includes a second annular abradable element 17 intended to cooperate with at least one second lip 20 connected to the first lip 19 to form a second sealing device. More specifically, the second abradable element 17 is supported by a second radially internal face 15b of the downstream portion 26 of the radially internal wall 5 and thus by the downstream end ring 30. The second abradable element 17 is positioned radially below the neck 31.

[0071] The first and second sealing devices are contiguous and ensure sealing between the radially internal wall 5 of the housing 1 and the internal part 22 of the sealing flange 10.

[0072] The radially external wall of the crown 33 of the downstream end crown 30 comprises an external surface 63 supporting a third abradable element 18 annular intended to cooperate with a third lick 21 connected to a radially external part 24 of the sealing flange 10 to form a third sealing device ensuring the seal between the radially internal wall 5 of the housing 1 and the radially external part 24 of the sealing flange 10.

[0073] The collar 31 is positioned radially between the second and third abradable elements 17, 18.

[0074] The third abradable element 18 is positioned radially above the neck 31, at the downstream end 42 of the downstream end ring 30. The second and third abradable elements 17, 18 run along the downstream end wall 46 of the downstream end ring 30.

[0075] The first, second and third sealing devices are labyrinth seals.

[0076] The first, second, and third abradable elements 16, 17, 18 are annular in shape, project radially from the radially internal wall 5, and are made of an abradable material, particularly a honeycomb structure, and come into contact with the slats 19, 20, 21. The first, second, and third abradable elements 16, 17, 18 may each have one or more tiers. The slats 19, 20, 21 may be straight or inclined and / or tiered.

[0077] In the example of Figures 2 to 5, the first sealing device comprises a first abradable element 16 having two first abradable surfaces 43 in a stepped arrangement, i.e., offset from each other along the radial direction Z, each cooperating with a first lip 19 inclined with respect to the first abradable surfaces 43. The second sealing device comprises a second abradable element 17 having two second abradable surfaces 44 in a stepped arrangement, each cooperating with a second lip 20 inclined with respect to the second abradable surfaces 44. The first and second lips 19, 20 are supported by a first arm 45 connected to the inner part 22 of the sealing flange 10.

[0078] The third sealing device comprises a single third abradable surface 47 in contact with the third slit 21. The third slit 21 is straight, i.e. orthogonal to the third abradable surface 47.

[0079] The third lick 21 is supported by a second arm 48 connected to the radially external part 24 of the sealing flange 10.

[0080] The turbine 3 includes a fourth sealing device comprising three inclined fourth blades 49 supported by the second arm 48 and cooperating with a fourth abradable element 51 supported by an internal face 52 of the radially external wall 4. The fourth abradable element 51 is stepped and comprises three fourth abradable surfaces 50 offset from each other along the radial direction Z. Each fourth abradable surface 50 is in contact with one of the fourth blades 49 which is inclined with respect to the fourth abradable surface 50.

[0081] The fourth sealing device ensures the seal between the radially external part 24 of the sealing flange 10 and the radially external wall 4 of the housing 1.

[0082] In another embodiment, all the sealing devices can be brush seals.

[0083] The housing 1 includes at least one bypass passage or conduit 11 positioned between two channels 6a, 6b and extending in a general radial direction Z, from a radially internal inlet 55 to a radially external outlet 56.

[0084] Bypass passage 11 also extends transversely in relation to channels 6a, 6b.

[0085] The bypass passage 11 is configured to guide a portion of a second airflow D2 from the high-pressure compressor 38 of the turbine 3 to a purge cavity in the turbine 3. The bypass passage 11 forms an air passage through which a diverted ventilation airflow D2' from the second airflow D2 flows. The diverted ventilation airflow D2' flows in the radial direction Z.

[0086] Inter-channel connecting sections 13 are formed between circumferentially adjacent channels 6a and 6b. Each inter-channel connecting section 13 includes a bypass passage 11.

[0087] Several bypass passages 11 are arranged circumferentially around the longitudinal axis X so as to form a ventilation ring 14 extending circumferentially around the longitudinal axis X.

[0088] Each bypass passage 11 is in the form of a slot with a generally parallelepiped cross-section in tangential section. Each bypass passage 11 has an elongated shape in the circumferential direction. In other words, each bypass passage 11 has a circumferential length that is greater than its axial width. The width is optionally at most 50%, or optionally 30%, of the length.

[0089] Each bypass passage 11 is axially delimited by the upstream wall 59 connected to the upstream portion 25 of the radially internal wall 5 and the downstream wall 54, opposite the upstream wall 59, and connected to the downstream portion 26 of the radially internal wall 5. Each bypass passage 11 is circumferentially delimited by a first circumferential end wall 53a of a first channel 6a and a second circumferential end wall 53b of a second channel 6b.

[0090] The circumferential length of the bypass passages 11 is as large as possible while maintaining a sufficient circumferential end wall thickness 53a, 53b. The thickness of the circumferential end walls 53a, 53b separating each bypass passage 11 from a channel 6a, 6b is at least 0.5 mm.

[0091] Preferably, there are as many bypass passages 11 as there are channels 6a, 6b. The number of channels 6a, 6b must be sufficient to deliver the necessary airflow to the blades for cooling and to avoid generating excessive pressure heterogeneities downstream of the downstream end ring 30.

[0092] The bypass passages 11 cross radially through the ventilation ring 14 and the channels 6a, 6b cross axially through the ventilation ring 14.

[0093] The ventilation ring 14 and therefore the bypass passages 11 are positioned upstream of the curved portion 29 connecting the upstream portion 25 to the downstream portion 26 of the radially internal wall 5, to an upstream part of the neck 31.

[0094] The radial projection of the ventilation ring 14 towards the longitudinal axis is positioned between the first and second abradable elements 16, 17.

[0095] As shown in Figure 6, the bypass passages 11, and more specifically the radially internal inlets 55, open radially between the first and second abradable elements 16, 17. In other words, the bypass passages 11 extend around the central axis A which is positioned between the first and second abradable elements 16, 17. The abradable elements 16, 17 are axially spaced and the radially internal inlets 55 open radially between the abradable elements 16, 17.

[0096] Each bypass passage 11 extends from the radially internal inlet 55, which is in fluidic communication with the last stage of the high-pressure compressor 38 of the turbine 3, to the radially external outlet 56, which is in fluidic communication with the purge cavity of the turbine 3.

[0097] The first and second abradable elements 17 are separated by an annular groove 57, as shown in Figure 6. The radially internal inlets 55 open radially into the annular groove 57.

[0098] The axial width of the bypass passages 11 corresponds approximately to the spacing between the first and second abradable elements 16, 17. The spacing between the first and second abradable elements 16, 17 is on the order of 5 mm, for example.

[0099] Channels 6a, 6b extend through the ventilation ring 14 from the upstream portion 25 of the radially internal wall 5, forming connecting channel portions 58, each positioned between two bypass passages 11. The ventilation ring 14 is formed by an alternation of bypass passages 11 and portions of connecting channel 58.

[0100] The connecting channel portions 58 are curved along the tangential direction Y and extend from the upstream wall 59 to the downstream wall 54 of the downstream end ring 30. The connecting channel portions 58 are each delimited by the two opposing circumferential end walls 53a, 53b, a radially external wall 62 and a radially internal wall 64, opposite the radially external wall 62. The circumferential end walls 53a, 53b and the radially external and internal walls 62, 64 extend axially from the upstream wall 59 to the downstream wall 54 of the downstream end ring 30. The first and second circumferential end walls 53a, 53b are curved.

[0101] The following describes the circulation of airflows in the ventilation circuits of turbine 3, regardless of the example presented previously.

[0102] As shown in Figure 2, each channel 6a, 6b is traversed by the first ventilation airflow DI from the annular bypass space 12 of the combustion chamber 36.

[0103] The second airflow D2 from the last stage of the high-pressure compressor 38 of the turbine 3 passes through the first sealing device formed by the first abradable element 16 and the first two blades 19.

[0104] The second airflow D2 then splits into two airflows, between the first and second sealing devices, including a diverted airflow D2' passing through the bypass passages 11 and a third airflow D3 passing through the second sealing device formed by the second abradable element 17 and the two second flaps 20. The diverted airflow D2' represents between 95% and 80% of the second airflow D2, and preferably 90%.

[0105] The third airflow D3 then mixes with the first airflow DI to form an airflow that splits into a fourth airflow D4 and a fifth airflow D5. The fifth airflow D5 represents between 5% and 15% of the airflow generated by the mixing of the third airflow D3 and the first airflow D1.

[0106] The fourth airflow D4 passes through the orifices 41 formed in the sealing flange 10 to enter the ventilation cavity 9. The ventilation cavity 9 communicates with the internal cooling circuits arranged inside the blades 39. The fourth airflow D4 therefore supplies air to the internal cooling circuits of the blades 39.

[0107] The fifth airflow D5 then passes through the third sealing device formed by the third abradable element 18 and the third lick 21 to mix with the diverted airflow D2', generating a sixth airflow D6 supplying the turbine 3 purge cavity to cool it.

[0108] The sixth airflow D6 mixed with the diverted airflow D2' according to the invention has a non-zero tangential velocity which limits the temperature rise of the air cooling the purge cavity of the turbine 3. The cooling efficiency of the ventilation circuit of the turbine 3 is therefore improved.

[0109] Although the present invention has been described with reference to specific embodiments, it is evident that modifications and changes can be made to these examples without departing from the general scope of the invention as defined by the claims. In particular, individual features of the various embodiments illustrated / mentioned can be combined in additional embodiments. Therefore, the description and drawings should be considered in an illustrative rather than restrictive sense.

Claims

Demands

1. Cooling air injection housing (1) for a bladed rotor disc (2) of a high-pressure turbomachine turbine (3), the housing (1) extending about a longitudinal axis (X) and comprising a radially external wall (4), a radially internal wall (5), and several channels (6a, 6b) distributed circumferentially about the longitudinal axis (X), each channel (6a, 6b) forming an air injector extending axially and along the radially internal wall (5) from an inlet opening (7) opening through the radially external wall (4) to an outlet opening (8) opening through a downstream end wall (46), each channel (6a, 6b) being configured to guide a first airflow (Dl) from an annular bypass space (12) of a combustion chamber (36) of the turbine (3) to a ventilation cavity (9) formed between the rotor disc (2) and a sealing flange (10) positioned upstream of the rotor disc (2),characterized in that: the casing (1) comprises at least one bypass passage (11), each bypass passage (11) being delimited by two circumferentially adjacent channels (6a, 6b) extending in a general radial direction from a radially internal inlet (55) to a radially external outlet (56), the bypass passage (11) being configured to divert a portion of a second airflow (D2) from a high-pressure compressor (38) of the turbine (3) and guide said portion of the second airflow (D2) to a purge cavity of the turbine (3).

2. Carter (1) according to claim 1, comprising several bypass passages (11) arranged circumferentially in a regular manner around the longitudinal axis (X).

3. Carter (1) according to claims 1 or 2, wherein the circumferentially adjacent channels (6a, 6b) are separated by inter-channel connecting sections (13), each inter-channel connecting section (13) being traversed by a bypass passage (11).

4. Carter (1) according to any one of claims 1 to 3, wherein each bypass passage (11) has a circumferentially measured length that is greater than its axially measured width.

5. Carter (1) according to any one of claims 1 to 4, wherein each bypass passage (11) is axially delimited by an upstream wall (59) connected to an upstream portion (25) of the radially internal wall (5) and a downstream wall (54), opposite the upstream wall (59), and connected to a downstream portion (26) of the radially internal wall (5), each bypass passage (11) being circumferentially delimited, on the one hand, by a first circumferential end wall (53a) of a first channel (6a) and, on the other hand, by a second circumferential end wall (53b) of a second channel (6b).

6. Housing (1) according to claim 5, comprising a first abradable element (16) extending radially inward from a first radially internal face (15a) of the upstream portion (25) of the radially internal wall (5) and intended to cooperate with at least one first lip (19) of the sealing flange (10) to form a first sealing device, and a second abradable element (17) extending radially inward from a second radially internal face (15b) of the downstream portion (26) of the radially internal wall (5) and intended to cooperate with at least one second lip (20) connected to the first lip (19) to form a second sealing device, the radially internal inlet (55) of each bypass passage (11) opening between the first and second abradable elements (16, 17).

7. A housing (1) according to any one of claims 5 or 6, comprising a downstream end ring (30) extending radially outward from the radially internal wall (5) and bounded axially by the downstream end wall (46) and the downstream wall (54), and radially on the inside, by the downstream portion (26) of the radially internal wall (5), and on the outside on the outside, by a radially external crown wall (33), opposite the downstream portion (26) of the radially internal wall (5).

8. Housing (1) according to claim 7, wherein the radially external wall of the crown (33) has a third abradable element (18) which extends radially outwards and is intended to cooperate with a third slat (21) connected to a radially external part (24) of the sealing flange (10) to form a third sealing device.

9. Carter (1) according to any one of claims 5 to 8, wherein the upstream portion (25) of the radially internal wall (5) is connected to the downstream portion (26) of the radially internal wall (5) by connecting channel portions (58), each connecting channel portion (58) separating circumferentially between two bypass passages (11).

10. A housing (1) according to any one of claims 1 to 9, wherein the radially internal wall (5) comprises an upstream portion (25) along which extends a primary section (27) of the channel (6a, 6b) from the inlet opening (7) and a downstream portion (26) along which extends a secondary section (28) of the channel (6a, 6b) to the outlet opening (8), the upstream portion (25) being connected to the downstream portion (26) by a curved portion (29).

11. Housing (1) according to claim 10, wherein the bypass passages (11) are positioned axially upstream of the curved portion (29).

12. Turbine (3) for a turbomachine comprising a casing (1) as defined according to any one of claims 1 to 11.

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