US20120237342A1

US20120237342A1 - Turbine stage in a turbine engine

Info

Publication number: US20120237342A1
Application number: US13/511,266
Authority: US
Inventors: Emmanuel Berche; Vincent Philippot
Original assignee: SNECMA SAS
Current assignee: Safran Aircraft Engines SAS
Priority date: 2009-12-18
Filing date: 2010-12-14
Publication date: 2012-09-20
Also published as: EP2513428A1; FR2954400B1; BR112012010257A2; CA2777370A1; WO2011073570A1; RU2012130351A; FR2954400A1; CN102667066A

Abstract

A turbine stage for a turbine engine includes a bladed wheel rotatable inside a sectorized ring of composite material carried by a casing. Each ring sector has a downstream circumferential rim held to bear radially against an annular tab that is engaged radially in an annular groove of the downstream circumferential rim of the ring with axial clearance, when cold, that is designed to be reduced to zero in operation when hot, and to enable the annular tab of the casing to be clamped axially in leaktight manner in the annular groove of the ring sector.

Description

The present invention relates to a turbine stage in a turbine engine such as a turboprop or a turbojet.
A turbine engine essentially comprises, going from upstream to downstream: a compressor, a combustion chamber, and a turbine, the compressor feeds the combustion chamber with air under pressure, and the turbine receives the hot gas coming from the combustion chamber in order to extract energy therefrom.
Conventionally, a low pressure turbine stage comprises a nozzle constituted by an annular row of stationary vanes extending radially between two annular platforms, namely an inner platform and an outer platform, and a rotor wheel mounted downstream from the nozzle inside a sectorized ring carried by a casing surrounding the turbine stage.
Each ring sector carries a sealing lining on an inside face that co-operates with the outer peripheries of the blades of the rotor wheel, and it includes on an outside face means for fastening to the casing, which means are formed by upstream and downstream circumferential rims. The upstream circumferential rim is engaged axially in an annular groove carried by an upstream annular tab of the casing, and the downstream circumferential rim is clamped radially against a downstream annular tab of the casing by a C-section fastener member engaged axially from downstream onto the downstream circumferential rim and the downstream annular tab.
An annular cavity is defined between the ring and the casing and it extends from upstream to downstream between annular tabs of the casing. The upstream annular tab includes orifices feeding the cavity with air taken from a compression stage of the turbine engine.
Introducing cooling air into a cavity in register with the turbine ring serves to avoid greatly enlarging the clearance at the tips of the blades, i.e. between the radially outer ends of the blades and the sealing lining, in order to minimize the amount of air that passes under pressure outside the zone that is swept by the blades, thereby avoiding penalizing the performance of the turbine.
It is known to make ring sectors out of ceramic matrix composite material in order to take advantage of the good mechanical properties of such material at high temperatures, even though the casing is itself generally made of a metal material. The ring is therefore more rigid than the casing and it possesses a coefficient of thermal expansion that is smaller than that of the metal casing, thus leading to differences of expansion between the ring and the casing.
In patent application FR 09/51446, the Applicant proposes blocking the ring axially against the downstream annular tab by radially engaging complementary shapes provided on the ring and on the annular tab, an annular sealing ring being mounted in an annular groove in a face of the annular tab facing the downstream circumferential rim of the ring and pressed against said rim.
Nevertheless, in operation, each ring sector expands and deforms, taking up a curved shape that is concave in the circumferential direction, with its concave side facing outwards (negative camber phenomenon). Thus, radial spaces are observed to form between the downstream annular tab of the casing and the downstream circumferential rims of the ring sectors.
These radial spaces are such that the annular sealing ring becomes incapable of providing sealing between the downstream circumferential rim and the annular tab of the casing, so leaks of cooling air occur between the downstream circumferential rims of the ring sectors and the downstream annular tab of the casing.
In other embodiments where the ring is not made of composite material, the downstream fastening between the downstream circumferential rim of the ring and the downstream annular tab of the casing is sealed by axially prestressing the downstream annular tab against a radial face of the downstream circumferential rim that is opposite to the fastener member.
Nevertheless, such assembly with axial prestress when cold cannot be envisaged when the ring is made of composite material because of its high stiffness and its small thermal expansion.
A particular object of the invention is to provide a solution to these problems that is simple, inexpensive, and effective, and that makes it possible to avoid the drawbacks of the prior art.
To this end, the invention provides a turbine stage for a turbine engine, the stage comprising a bladed wheel rotatable inside a sectorized ring made of composite material and carried by a casing, each ring sector having a downstream circumferential rim held to bear radially against an annular tab of the casing by a C-shaped fastener, the stage being characterized in that the annular tab of the casing is engaged radially in an annular groove in the downstream circumferential rim of the ring with axial clearance, when cold, that is designed to be reduced to zero in operation, when hot, and to enable leaktight axial clamping of the annular tab of the casing in the annular groove of the ring sector.
According to the invention, the downstream circumferential rim of the ring is sealed in operation by axial clamping of the upstream and downstream ends of the downstream annular tab of the casing in the annular groove as a result of the greater expansion of the casing compared with the composite ring. The concave curvature of the ring and of its downstream circumferential rim is thus compensated by the axial clamping of the annular tab, thereby guaranteeing sealing of the downstream fastening of the ring.
Advantageously, the annular tab of the casing has upstream and downstream radial faces for bearing when hot against radial flanks of the groove. In operation, the radial faces of the annular tab and the radial flanks of the groove conserve their radial shape, thereby ensuring annular contact between the radial faces of the ring and the radial flanks of the groove.
According to another characteristic of the invention, the above-mentioned axial clearance when cold is of the order of one-tenth of a millimeter.
It is also possible to provide an annular sealing ring housed in an annular groove in the face of the annular tab that is pressed against the bottom wall of the annular groove of the ring sector.
Advantageously, the composite material is of the ceramic matrix type and the casing is made of a metal material.
The invention also provides a turbine engine such as an airplane turboprop or turbojet, the engine including a high pressure turbine stage of the above-described type.

The invention can be better understood and other details, advantages, and characteristics of the invention appear on reading the following description made by way of non-limiting example and with reference to the accompanying drawings, in which:

FIG. 1 is a fragmentary diagrammatic view in axial section of a prior art turbine stage;

FIG. 2 is a diagrammatic view in cross-section on section plane A-A shown in FIG. 1;

FIG. 3 is a fragmentary diagrammatic view in axial section of a turbine stage of the invention while cold, the section plane not passing through a fastener member; and

FIG. 4 is a fragmentary diagrammatic view in axial section of a turbine stage of the invention while hot, on a section plane that contains a fastener member.

Reference is made initially to FIG. 1 which shows a portion of a turbine stage 10 in a turbine engine that includes a nozzle stage having a plurality of stationary vanes arranged upstream from a rotary wheel carrying a plurality of blades and rotatable inside a ring 12 carried by an outer casing 14.
The ring 12 is made up of a plurality of substantially cylindrical ring sectors that are circularly juxtaposed, end to end. Each ring sector comprises a cylindrical portion 16 carrying on its inside face a sealing lining 18 of abradable material that cooperates with the outer peripheries of the blades of the rotor wheel. Each ring sector includes two annular tabs, namely an upstream tab 18 and a downstream tab 20 for attaching to the casing 14. The outer end of the upstream annular tab 18 has a circumferential rim 22 extending upstream and engaged axially in an annular groove 24 facing downstream and formed in a radial annular tab 26 of the casing. The outer end of the downstream annular tab 20 of the ring has a circumferential rim 28 facing downstream and clamped radially against a cylindrical portion 30 of an annular tab 32 of the casing 14 by means of a C-section fastener member 32 engaged axially on the downstream circumferential rim 28 and on the cylindrical portion 30 of the downstream annular tab 32 of the casing 14.
Each downstream circumferential rim 20 of a ring sector includes at least one notch in radial alignment with a notch in the cylindrical portion 30 of the downstream annular tab 32 of the casing 14 and of width that is sufficient to enable the fastener member 32 to be engaged axially therein and to enable the ring 12 to be fastened on the casing 14.
An annular cavity 34 is defined between the sectorized ring 12 and the casing 14 being defined upstream by the upstream annular tabs 18, 26 of the ring 12 and of the casing 14, respectively, and downstream by the downstream annular tabs 20, 32 of the ring 12 and of the casing 14, respectively.
The upstream annular tab 26 of the casing 14 has orifices 36 for passing cooling air coming from a space surrounding the combustion chamber, i.e. air that flows between the outer casing of the combustion chamber and the outer wall of the combustion chamber, which wall forms a body of revolution. In order to avoid cooling air leaking between the cylindrical portion 30 of the downstream annular tab 32 of the casing 14 and the downstream circumferential rim 20 of the ring 12, an annular sealing ring 38 is mounted in an annular groove 40 of the inside face of the cylindrical portion 30. This sealing ring 38 is compressed radially in the annular groove 40 and against the downstream circumferential rim 28 of the ring 12. The inside face of the cylindrical portion 30 includes a rib 42 engaged radially in an annular recess in the downstream circumferential rim 28 of the ring 12 in order to prevent the ring 12 from moving axially relative to the casing 14.
The radial face at each circumferential end of each ring sector includes three slots 44, 46, 48, each of which houses a sealing strip. A first slot 44 is formed in the cylindrical portion 16 of the ring 12 and extends over substantially the entire length of the ring 12, being parallel to the longitudinal axis of the ring 12. The other two slots 46 and 48 are oblique, each being formed in a respective one of the upstream and downstream annular tabs 18 and 20 of the ring. The radially inner ends of the two oblique slots 46 and 48 open out into a middle portion of the longitudinal slot 44, and their radial ends open out into the outside faces of the upstream and downstream circumferential rims 22 and 28 respectively. Each sealing strip is inserted half in a slot 44, 46, 48 of one sector with the other half in a corresponding facing slot that is formed in a radial face of an adjacent ring sector.
Nevertheless, as explained above, each sector of the composite ring deforms under the effect of temperature and adopts a concave curved shape with its concave side facing outwards (FIG. 2). The casing 14 is also subjected to deformation and has circumferential undulations. As a result, because of the differential expansions between the ring 12 made of composite material and the casing 14, a radial space R forms between each circumferential rim 28 and the cylindrical portion 30 of a downstream tab 32 of the casing 14, thereby leading to leaks of ventilation air from the annular cavity 34 towards the gas flow passage through the turbine.
The invention serves to remedy this problem and those mentioned above by forming an annular groove 50 in the outside cylindrical face of the downstream circumferential rim 52 of the ring 54, which groove receives radially the downstream cylindrical portion 55 of the downstream annular tab 56 of the casing 14 with axial clearance j, when cold, that is designed to be reduced to zero in operation as a result of the greater expansion of the casing 14 and of its downstream annular tab 56 compared with the expansion of the ring 54 made of composite material (FIG. 3).
The annular groove 50 has two radial annular flanks, an upstream flank 58 and a downstream flank 60. The downstream cylindrical portion 55 of the downstream annular tab 56 of the casing 14 has two radial faces, an upstream face 62 and a downstream face 64. When hot, in operation, the radial faces 62 and 64 of the downstream annular tab 56 of the casing 14 come to bear against the radial flanks 58 and 60 of the groove 50 as a result of the differential expansion between the composite ring 54 and the casing 14, thereby ensuring that the annular tab 56 is clamped axially in the groove 50 and establishes sealing against the ventilation air flowing in the cavity 34. This axial clamping also serves to hold the ring 54 axially relative to the casing 14.
The depth of the groove 50 is selected in such a manner as to be greater than the maximum radial difference R in operation between the inside face 66 of the downstream cylindrical portion 55 of the downstream tab 56 of the casing 14 and the bottom wall 68 of the groove 50, so as to ensure continuous leakproof axial clamping when hot and so as to avoid any axial separation of the ring 54 relative to the casing 14.
A ring sector is assembled by inserting the upstream circumferential rim 22 of the ring 54 in the annular groove of the upstream tab 18 of the casing 14, and then tilting the downstream end of the ring outwards so that the cylindrical portion 55 bears against the bottom wall of the groove 50. The axial clearance j, when cold, serves to make it easier to tilt the ring 54 outwards against the casing 14.
An annular sealing ring 38 is housed in an annular groove 40 of the face 66 of the downstream annular tab 56 of the casing that presses against the bottom wall 68 of the groove 50.
In a manner similar to the prior art, each downstream circumferential rim 52 of a ring sector includes a notch in radial alignment with a notch in the cylindrical portion of the downstream annular tab of the casing to enable the C-section fastener member 32 to be mounted axially.
The inter-sector sealing means are similar to those of the prior art. Nevertheless, it should be observed that in the invention the sloping slot in the downstream annular tab 64 of the ring 54 opens out into the groove 50 in register with the sealing ring 38.
In one particular embodiment of the invention, the axial clearance, when cold, is of the order of 0.1 millimeters.
The ring 54 may be made of a ceramic matrix composite material that withstands well the high temperatures of the kind that exist in a high pressure turbine, and the casing 14 is made of a metal material such as Inco or steel.

Claims

1-6. (canceled)

7. A turbine stage for a turbine engine, the stage comprising:

a bladed wheel rotatable inside a sectorized ring made of composite material and carried by a casing, each ring sector having a downstream circumferential rim held to bear radially against an annular tab of the casing by a C-section fastener,

wherein the annular tab of the casing is engaged radially in an annular groove in the downstream circumferential rim of the ring with axial clearance, when cold, that is designed to be reduced to zero in operation, when hot, and to enable leaktight axial clamping of the annular tab of the casing in the annular groove of the ring sector.

8. The turbine stage according to claim 7, wherein the annular tab of the casing has upstream and downstream radial faces for bearing when hot against radial flanks of the groove.

9. The turbine stage according to claim 7, wherein the axial clearance when cold is of the order of one-tenth of a millimeter.

10. The turbine stage according to claim 7, wherein an annular sealing ring is housed in an annular groove in the face of the annular tab that is pressed against the bottom wall of the annular groove in the ring sector.

11. The turbine stage according to claim 7, wherein the composite material is of the ceramic matrix type and the casing is made of a metal material.

12. A turbine engine such as an airplane turboprop or turbojet, wherein the engine includes the turbine stage according to claim 7.