TURBOMACHINE STAGE INCLUDING AT LEAST ONE SEALING RING

Abstract

A turbomachine stage, particularly for aircraft, having a main axis (AA) and comprising a bladed wheel (2A) having a disc carrying at its periphery blades (4), these blades (4) having platforms (6) forming upstream an annular spoiler (6A) extending around the axis (AA) and oriented axially upstream and a distributor (2B) having an annular ferrule (12) extending around the axis (AA) and blades (11) extending radially outwards from this ferrule (12). The stage also includes at least one first sealing ring (17) extending around the axis (AA) and arranged radially inside said ferrule (12), this first sealing ring (17) being located axially upstream and opposite said spoiler (6A) and being made of an abradable material so as to be able to wear away by friction with the spoiler (6A). Figure for the abbreviation: Figure 4

Description

Title of the invention: TURBOMACHINE STAGE COMPRISING AT LEAST ONE SEAL RING Technical field of the invention

[0001] The present invention relates to a turbomachine stage, particularly for aircraft, comprising at least one sealing ring. The present invention also relates to a turbomachine comprising such a stage. Technical background

[0002] Generally, an aircraft turbomachine, whether for a civil or military transport aircraft, a helicopter, etc., comprises, from upstream to downstream with respect to the general direction of gas flow, a plurality of compressors that compress the air entering a nacelle. The air, which may have been previously accelerated by a fan located upstream of the compressors, is then introduced into a combustion chamber where it is mixed with fuel and burned. The combustion gases pass through one or more turbines consisting of several stages which, in turn, drive the compressor(s). The stages of the turbine(s) have a principal axis of rotation coaxial with the longitudinal axis of the turbomachine. Each stage includes a rotating wheel and a distributor axially adjacent to it.

[0003] Typically, a rotating wheel comprises a plurality of blades, each with a blade extending radially between an outer casing and an inner platform. A distributor comprises a plurality of fixed blades extending radially between the outer casing and an inner ring. In the turbomachine, the space between the outer casing on one side and the inner ring and inner platform on the other forms a combustion gas flow channel called the turbomachine's primary stream. Furthermore, the space radially internal to the ring and platforms defines cavities called purge zones outside the stream. These cavities belong to the secondary air circuit, from which a portion of the gases from the compressors is drawn to cool the moving parts.This portion of gas, also called the purge flow, passes radially through the purge zone from the inside out before being discharged into the primary channel at the level of the inner shell and platforms.

[0004] To limit the passage of combustion gas from the primary stream into the purge zone, and in other words to maximize the amount of gas passing through the distributor blades, it is possible to modify the geometry of the wheel and / or the distributor by arranging the labyrinths along the path of the purge flow while limiting interference between the inner shell and the platforms. However, depending on the compromises Due to dimensional adjustments, axial and radial clearances between the impeller and the distributor in the purge zone can increase. These clearances can lead to an increase in the flow of combustion gases into the purge zone. To counteract this hot air flow in the purge zone, it is necessary to increase the purge flow rate of the secondary air system. However, this significantly impacts the turbomachine's performance.

[0005] Furthermore, applicant's document FR-A1-2 841 591 describes turbine stages comprising a plurality of radially aligned blades facing labyrinths. Moreover, applicant's document FR-A1-3 039 589 proposes an internal distributor shell which includes, on at least one radially internal portion, a component forming an additional obstacle in the purge zone. Summary of the invention

[0006] The present invention aims to provide sealing means compatible with a reduction in the axial distance of the purge zone.

[0007] To this end, the invention relates to a turbomachine stage, in particular for an aircraft, this stage having a main axis and comprising: - a bladed wheel comprising a disc with blades on its periphery, these blades comprising platforms forming upstream an annular spoiler extending around the axis and oriented axially upstream, - a distributor comprising an annular ferrule extending around the axis and blades extending radially outwards from this ferrule. According to the invention, the stage also includes at least one first sealing ring extending around the axis and arranged radially inside said ferrule, this first sealing ring being located axially upstream and opposite said spoiler and being made of an abradable material so as to be able to wear away by friction with the spoiler.

[0008] Thus, thanks to the invention, the axial distance between the impeller and the distributor is reduced without the platform damaging the internal ferrule. Indeed, the presence of the abradable sealing ring protects the ferrule from any friction with the platform's spoiler. Furthermore, by reducing the axial distance between the impeller and the distributor, the pressure in the purge zone increases, which facilitates the management of cooling flows to the parts in this area. In fact, bringing the impeller and the distributor closer together limits the risk of hot air being re-injected from the flow stream into the purge zone and avoids increasing the flow rate of the secondary air system's purge.

[0009] The turbomachine stage may also have one or more of the following characteristics The following risks, taken alone or in combination with each other: - said first sealing ring is made of a honeycomb-type cellular material, - said ferrule is connected to the outer periphery of a radial wall whose inner periphery is connected to a second sealing ring independent of said first sealing ring; the radial wall may be connected to a cylindrical wall supporting the second ring or may be fixed by means such as a screw and nut to a support for this second ring,

[0010] - said first sealing ring is fixed on a downstream radial face of said wall radial, - said first sealing ring has a generally rectangular or square cross-section, - said first sealing ring comprises a plurality of chimneys oriented radially with respect to said axis and distributed around this axis, these chimneys being configured to allow passage of purge flow; preferably including in a rotor-stator contact situation between the moving wheel and the distributor, thus avoiding blockage of purge flow,

[0011] — the chimneys are inclined in an azimuthal direction to put the purge air in rotation and allow for better adaptation to the air's rotation within the vein, - The number of chimneys is between 2 and 50, and preferably between 10 and 50; this number may depend on the size of the turbomachine considered and may have an impact on the homogeneity of ventilation, - said first sealing ring has an internal diameter smaller than the diameter of said spoiler, and an external diameter larger than the diameter of said spoiler, - each of the chimneys opens onto an internal cylindrical surface of said first sealing ring, which has said internal diameter, and onto an external cylindrical surface of said first sealing ring, which has said external diameter, - said ferrule has a generally frustoconical shape and comprises at a downstream end a diameter greater than the external diameter of said first sealing ring, and at an upstream end a diameter less than the external diameter of said first sealing ring, — said chimneys are arranged near the downstream radial face of the radial wall, and — the annular beak includes, at an upstream end, an annular edge comprising saw teeth and / or having a flared shape.

[0012] The invention further relates to a turbomachine, in particular for aircraft, comprising at least one stage according to one of the preceding characteristics. Brief description of the figures

[0013] Other features and advantages of the invention will become apparent upon reading the detailed description that follows, for an understanding of which reference should be made to the accompanying drawings in which:

[0014] [Fig-1] The [Fig.1] is a partial schematic half-view in axial section of a turbomachine turbine comprising a plurality of stages. [Fig.2] Fig.2 is an enlarged view of the turbine shown in Fig.1.

[0015] [Fig.3] The [Fig.3] schematically represents a stage comprising a sealing ring according to an embodiment of the invention.

[0016] [Fig.4] Fig.4 schematically represents a stage comprising a sealing ring in contact with a spoiler according to another embodiment of the invention.

[0017] [Fig.5] Fig.5 schematically represents an axial section of the sealing ring according to an embodiment of the invention. [Fig.6] Fig.6 illustrates the circulation of a purge flow in a stage comprising a sealing ring without contact with a spoiler according to one embodiment.

[0018] [Fig.7] The [Fig.7] illustrates the circulation of a purge flow in a stage through the chimneys of a sealing ring in contact with a spoiler according to one embodiment. Detailed description of the invention

[0019] By convention in this application, the terms "inside" and "outside," and "internal" and "external," are defined radially with respect to an axis XX of the turbomachine. Thus, a cylinder extending along axis XX has an inner face facing the axis of the turbomachine and an outer surface opposite its inner surface. "Axial" or "axially" means any direction parallel to the axis X, and "radially" or "radial" means any direction perpendicular to the axis XX. Similarly, the terms "upstream" and "downstream" are defined with respect to the direction of airflow in the turbomachine.

[0020] Figure 1 shows a portion of a low-pressure turbine 1 of an aircraft turbomachine, whether for a civil or military transport aircraft. The turbine comprises a series of stages. Each stage includes a main shaft AA coaxial with the axis of the turbomachine. A stage includes a movable runner 2A and a distributor 2B. The runner 2A has a disk 3 which carries, on its periphery, a plurality of individual blades 4. Each of these blades 4 has a blade 5 comprising a free outer end 5A and an inner end 5B opposite the outer end 5A. The inner end 5B of each blade 5 is connected to a platform 6 which itself connects each blade 5 to a midwall or strut 7 arranged radially ([Fig.2]). The strut 7 extends radially inwards by a foot 8 which is engaged in a groove of the disc 3. The housing grooves of the feet 8 of the blades 4 define teeth between them which are surrounded by the platforms 6 of the blades 4. The wheel 2A is connected to the turbine shaft by means of a drive cone 9.

[0021] As shown in Figures 1 and 2, the platform 6 forms, in axial projection, an upstream spoiler 6A oriented towards the upstream side of stage 1 and a downstream spoiler 6B oriented towards the downstream side of the stage. The upstream spoiler 6A is annular in shape and extends around the main axis AA along a diameter DI A to the shaft. The downstream spoiler 6B also has an annular shape extending around the main axis AA along a diameter DIB to the turbine shaft.

[0022] Furthermore, the distributor 2B comprises a plurality of fixed blades 10 arranged circumferentially around the main axis AA. Each of the fixed blades 10 has a vane 11. Each of the vanes 11 has a radially external end by which it is fixed to the outer casing 11A of the turbine. The vanes 11 are joined together at their radially internal ends by an annular ferrule 12 extending around the main axis AA such that the vanes 11 extend radially outwards from the annular ferrule 12.

[0023] In a preferred embodiment, the annular ferrule 12 has a frustoconical shape, the radially external surface 12C of which delimits a portion of the primary flow path VP of the combustion gases through the turbine 1. As shown in particular in Figures 3 and 4, the frustoconical shape of the ferrule 12 comprises a downstream end 12B and an upstream end 12A. The upstream end 12A has an upstream diameter D2A which is smaller than the downstream diameter D2B of the downstream end 12B.

[0024] Furthermore, rotating parts such as the disc 3 of the movable wheel 2A are cooled by an airflow also called the purge flow F, represented in Figures 3, 4, 6, and 7 by arrow F. This purge flow is taken from the airflow exiting the compressors and circulates in the secondary air circuit, which includes, in particular, the purge zone ZP. The term "purge zone ZP" refers to the area located axially between the wheel 2A and the distributor 2B, which bores one stage, and radially inside the annular shell 12 and the platform 6. In the purge zone ZP, the purge flow F flows radially from the inside to the outside until it reaches the primary stream VP.

[0025] According to one embodiment of the invention, the ferrule 12 is connected to the outer periphery 13A of a radial wall 13 which includes a downstream radial face 13C ([Fig. 3]). The inner periphery 13B of this radial wall 13 is mounted between a downstream flange 14B and an upstream flange 14A which are held together by a bolt. The bolt also carries a flange 16 extending radially outwards. The flange 16 comprises An annular rim 16B is oriented downstream along the main axis AA. This annular rim 16B is staggered with the upstream spoiler 6A. The shape of the annular rim 16B of each of the distributors 2B can vary depending on the application and the stage considered within the turbine 1. Furthermore, it is possible that the annular shell 12 may lack an annular rim 16B. The presence of upstream spoilers 6A and annular rims 16B creates baffles to limit the passage of combustion gases circulating radially in the primary flow VP from the outside to the inside.

[0026] Alternatively, the inner periphery 13B of this radial wall 13 is connected to a cylindrical wall 14 which may carry a sealing ring 15 ([Fig. 2]). The radial wall 13 also carries an annular rim 16B oriented downstream along the main axis AA so as to be staggered with the upstream spoiler 6A. The shape of the annular rim 16B may vary depending on the application and the stage considered within the turbine 1. Furthermore, it is possible that the shell 12 may lack an annular rim 16B. The presence of upstream spoilers 6A and annular rims 16B form baffles to limit the passage of combustion gases circulating in the primary stream VP radially from the outside to the inside.

[0027] According to a preferred embodiment, the stage also includes at least one first sealing ring 17 extending around the main axis AA. This sealing ring 17 is arranged radially inside the annular ferrule 12 between the downstream end 12B of the frustoconical shape of the ferrule 12 and the annular rim 16B. As shown in particular in Figures 3 and 4, the sealing ring 17 is fixed to the downstream radial face 13C of the radial wall 13 and extends axially downstream and opposite the upstream spoiler 6A of the platform 6.

[0028] According to a preferred embodiment, the sealing ring 17 is made of an abradable material such that it wears down during friction with the upstream spoiler 6A during the rotation of the wheel 2A around the main axis AA. The abradable material from which the sealing ring is made is resistant to temperatures typical of turbomachine operation, for example, temperatures on the order of 1000 degrees Celsius. The sealing ring 17 is made of a honeycomb-type cellular material. The honeycomb cells are preferably oriented axially because in the radial direction they are likely to allow airflow. Alternatively, the sealing ring 17 is made of a polymeric material.

[0029] As shown in Figures 3, 4, 6 and 7, the sealing ring 17 has a generally square or rectangular cross-section. It comprises a downstream face 17A opposite the upstream spoiler 6A and extending axially over an axial length comparable to the axial length of the annular rim 16B relative to the wall radial 13. The sealing ring 17 also includes an external cylindrical surface 17B of the sealing ring that is opposite the downstream end 12B of the frustoconical shape of the ferrule 12. The diameter of the external cylindrical surface 17B defines an external diameter D3A of the sealing ring 17. This external diameter D3A is greater than the diameter DI A of the upstream spoiler 6A of the platform 6. The sealing ring 17 also has an internal cylindrical surface 17C arranged radially opposite the annular rim 16B such that the internal diameter D3B of the sealing ring 17, defined by the diameter of the internal cylindrical surface 17C, is less than the diameter DI A of the upstream spoiler 6A. The downstream radial face 17A of the sealing ring 17 is centered on the upstream spoiler 6A.

[0030] Furthermore, in a preferred embodiment, the upstream diameter D2A of the upstream end of the ferrule 12 is less than the external diameter D3A of the sealing ring 17. The downstream diameter D2B of the downstream end of the ferrule 12 is greater than the external diameter D3A of the sealing ring 17.

[0031] As shown in Figures 3 and 6, when the upstream spoiler 6A is not in contact with the downstream radial face 17A of the sealing ring 17, the purge flow F circulates radially from the inside to the outside through the baffles formed by the spoilers and the annular rims to the primary stream VP. It is then possible that some of the combustion gases may flow into the primary stream VP.

[0032] In order to limit the passage of combustion gases into the purge zone ZP during operating transients, the wheel 2A and the distributor 2B are brought closer together axially so that the upstream spoiler 6A is in contact with the sealing ring 17, as shown in Figures 4 and 7. During rotations of the movable wheel around the main axis AA, the upstream spoiler 6A rubs against the downstream face 17C of the sealing ring 17 without damaging the ferrule 12. The sealing ring 17 is, in fact, made of an abradable honeycomb-type material capable of absorbing this friction.

[0033] According to a preferred embodiment, the sealing ring 17 comprises a plurality of chimneys 18 which are oriented radially with respect to the main axis AA. As shown in Figures 4, 5 and 7, the chimneys are distributed circumferentially around the main axis AA so as to allow passage of purge flow when the spoiler is in contact with the sealing ring 17.

[0034] The chimneys 18 can also be inclined in an azimuthal or tangential direction with respect to the axis AA to set the purge air in rotation and allow better adaptation to the gyration of the air in the vein.

[0035] Each of the chimneys 18 has a cylindrical shape, an external face 18A of which opens onto the external cylindrical surface 17B of the sealing ring 17 and one of the inner faces 18B opens onto the inner cylindrical surface 17C of the sealing ring 17 so that each of the chimneys 18 passes through it. The chimneys 18 are aligned in a radial plane which is close to the downstream radial face 13C of the radial wall 13. By keeping the chimneys away from the downstream face 17C of the sealing ring 17, the integrity of the chimneys is preserved when the upstream spoiler 6A rubs against the sealing ring 17.

[0036] The closer the chimneys 18 are to the face, the greater the risk that, during interaction phases, the chimney will be opened by the spoiler. Air, on the other hand, has easy access to the chimney because there is no significant baffle. Conversely, the greater this distance, the lower the risk of perforation, but the more difficulty the purge air may have in reaching the chimney. A compromise must therefore be found between the robustness of the design and the pressure drop in the purge circuit.

[0037] Indeed, as shown in [Fig.7], in this configuration, the cavity forming the purge zone is closed so that the purge flow can only be evacuated into the primary vein by passing through the chimneys 18.

[0038] According to a preferred embodiment, the number of chimneys 18 arranged around the main axis AA is between 2 and 50. Preferably this number is between 10 and 50.

[0039] According to a particular embodiment, the upstream annular spoiler 6A comprises, at an upstream end, an annular edge. This annular edge may comprise an azimuthal and / or radial arrangement of teeth. In one variant, the annular edge of the spoiler also has a flared, elephant-foot-type shape.

[0040] Thus, the invention has the advantage of reducing the permeability of the purge zone to combustion gases circulating in the primary stream VP by bringing the moving wheel 2A and the distributor 2B closer together. This configuration increases the static pressure of the secondary air circuit supply so that the purge flow is carried out through the stacks 18 without combustion gases passing from the primary stream to the purge zone.

Claims

Demands

1. A turbomachine stage, particularly for aircraft, this stage having a main axis (AA) and comprising: - a bladed wheel (2A) having a disk (3) carrying blades (4) around its periphery, these blades (4) having platforms (6) forming upstream an annular spoiler (6A) extending around the axis (AA) and oriented axially upstream, - a distributor (2B) having an annular ferrule (12) extending around the axis (AA) and blades (11) extending radially outwards from this annular ferrule (12), - at least one first sealing ring (17) extending around the axis (AA) and arranged radially inside said ferrule (12), this first sealing ring (17) being located axially upstream and opposite said spoiler (6A) and being made of a material abradable so that it can wear down by friction with the spoiler (6A),characterized in that said first sealing ring (17) comprises a plurality of chimneys (18) oriented radially with respect to said axis (AA) and distributed around this axis (AA), these chimneys (18) being configured to allow passages of purge flow (F).

2. Turbomachine stage according to the preceding claim, characterized in that said first sealing ring (17) is made of honeycomb-type alveolar material.

3. Turbomachine stage according to any one of the preceding claims, characterized in that said shell (12) is connected to the outer periphery of a radial wall (13) whose inner periphery (13B) is connected to a second sealing ring (15) independent of said first sealing ring (17).

4. Turbomachine stage according to the preceding claim, characterized in that said first sealing ring (17) is fixed on a downstream radial face (13C) of said radial wall (13).

5. Turbomachine stage, according to any one of the preceding claims, characterized in that said first sealing ring (17) has in cross-section a generally rectangular or square shape.

6. Turbomachine stage according to any one of the preceding claims, ca- characterized in that the number of chimneys (18) is between 2 and 50 and preferably between 10 and 50.

7. Turbomachine stage according to any one of the preceding claims, characterized in that said first sealing ring (17) has an internal diameter (D3B) less than the diameter (DIA) of said spoiler (6A), and an external diameter (D3A) greater than the diameter (DIA) of said spoiler (6A).

8. Turbomachine stage according to claim 7, characterized in that each of the chimneys (18) opens onto an internal cylindrical surface (17C) of said first sealing ring (17), which has said internal diameter (D3B), and onto an external cylindrical surface (17B) of said first sealing ring (17), which has said external diameter (D3A).

9. Turbomachine stage according to claim 7 or 8, characterized in that said ferrule (12) has a generally frustoconical shape and comprises at a downstream end (12B) a diameter (D2B) greater than the external diameter (D3A) of said first sealing ring (17), and at an upstream end (12A) a diameter (D2A) less than the external diameter (D3A) of said first sealing ring (17).

10. Turbomachine, in particular aircraft, comprising at least one stage according to any one of the preceding claims.

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