Gas turbine shroud assembly and aeronautical turbine engine

Abstract

A gas turbine shroud assembly comprises a turbine shroud comprising shroud sectors (28) made of ceramic matrix composite material, a casing (22) made of metal alloy and a sectorized cooling device (52) comprising a plenum chamber (52B) delimited by an air diffuser (54) having holes for diffusing jets of cooling air over a radially outer surface (30B) of the turbine shroud. The assembly further comprises two metal alloy bushings (60) housed in two cavities (62) present at the circumferential ends of the air diffuser (54). Each of the bushings is in the form of a stepped cylinder with a base (60A), a lower face (601) of which bears on the radially outer surface (30B) of the turbine shroud, and a stud surmounting the base (60B). The stud is intended to receive a spring (58) compressed between an upper face (602) of the base and an upper face (620) of the cavity.

Description

[0001] Description

[0002] Title of the invention:

[0003] IMPROVED BACKLASH ADJUSTMENT TURBINE RING ASSEMBLY GAS TURBINE RING ASSEMBLY AND AEROSPACE TURBOMACHINERY

[0004] Technical Field

[0005] The present invention relates to the field of gas turbines such as turbines for aircraft engines or helicopter engines, turbines for electric generators and turbochargers.

[0006] Prior art

[0007] Climate change is a major concern for many legislative and regulatory bodies around the world. Indeed, various carbon emission restrictions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies both to new aircraft types and to those currently in operation, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been mobilizing for several years now to contribute to the fight against climate change.

[0008] Technological research efforts have already led to significant improvements in the environmental performance of aircraft. The Applicant takes into consideration the impact factors in all phases of design and development to obtain less energy-intensive, more environmentally friendly aeronautical components and products whose integration and use in civil aviation have moderate environmental impacts with the aim of improving the energy efficiency of these aircraft.

[0009] Consequently, the Applicant is constantly working to reduce its climate impact by using methods and operating virtuous development and manufacturing processes that minimize greenhouse gas emissions to reduce the environmental footprint of its activity. This sustained research and development work focuses on new generations of aircraft turbomachines, the weight reduction of aircraft, in particular through the materials used and lighter onboard equipment, the development of the use of electric technologies to provide propulsion, and, as essential complements to technological progress, aeronautical biofuels.

[0010] The current trend in civil aeronautics and the framework of future civil aircraft engines imply the requirement for high performance at high temperatures, which requires the integration of ceramic matrix composite (CMC) parts in the high pressure (HP) turbine. CMC has the advantage of withstanding high temperatures (over 1200°C) and its low density (three times lower than metal bases) allows for significant weight savings.

[0011] However, the integration of these CMC parts into a set of metal parts generates constraints due to the different expansion coefficients existing between CMC and metal elements. It is therefore necessary to find technologies capable of overcoming these problems by integrating clearance compensation technologies.

[0012] Currently, the CMC turbine rings developed by the applicant have the shape of the inverted letter PI, the base of which, rectilinear and horizontal, forms an aerodynamic vein which allows the combustion gases to be guided, and on which is deposited an abradable material which protects the CMC from external aggressions. The two legs of this PI ring allow the base to be positioned radially, thanks to cylindrical metal pins which allow them to be attached to the metal parts connected to the turbine casing. Due to the greater expansion of the pins compared to the ring legs, the contact between these two elements must be made with clearance (by widening the drilling diameter of the ring legs), releasing a degree of freedom of radial translation which no longer allows the ring to be correctly held in position.

[0013] There is therefore still a need today for a solution which always consists of imposing zero radial displacement on the turbine ring while overcoming the aforementioned drawbacks of the presence of springs. Description of the invention

[0014] To this end, the invention is the result of technological research aimed at significantly improving aircraft performance and, in this sense, contributes to reducing the environmental impact of these aircraft. For this purpose, the present invention therefore has the main aim of integrating a CMC ring into a gas turbine by compensating for differential expansions without mechanical over-stress while protecting the flexible elements from the high temperatures of the combustion chamber.

[0015] These aims are achieved by a gas turbine ring assembly, comprising around a longitudinal axis of rotation, a turbine ring comprising a plurality of ring sectors made of ceramic matrix composite material arranged end to end circumferentially around the axis of rotation, a metal alloy casing forming a ring support structure, and a sectorized cooling device comprising a plenum chamber delimited by an air diffuser having holes arranged in a matrix manner for diffusing by impact cooling air jets onto a radially external surface of the turbine ring from which it is separated by a diffusion distance, assembly characterized in that it further comprises at least one sleeve, preferably two metal alloy sleeves housed in two dedicated cavities present at the circumferential ends of the air diffuser,each socket having the shape of a two-stage stepped cylinder with a base forming the lower stage, a lower face of which bears on the radially external surface of the turbine ring and a stud surmounting the base to form the upper stage intended to receive a spring configured to be compressed between the base and the cavity.,

[0016] Thus, this embedded spring system allows the CMC ring to be held in position in a metallic environment without impairing the effectiveness of the springs at high temperatures. Cooling of the ring is no longer hindered by the presence of the spring. The presence of an orifice passing through each bushing around which the spring is wound improves this cooling. Preferably, the base of each bushing defines the diffusion distance between the air diffuser and the radially external surface of the turbine ring.

[0017] Advantageously, the base of each socket is split over almost its entire height, so as to define two wings whose parallel internal faces define an enlarged channel for circumferential flow of the impact jet.

[0018] Preferably, each of the wings has a chamfer at the radially inner face in contact with the radially outer surface of the turbine ring.

[0019] Advantageously, the base of each socket has a flat to prevent its rotation and allow its translational guidance.

[0020] Preferably, each sleeve has a central bore to also provide impact cooling on the radially outer surface of the turbine ring.

[0021] Advantageously, the casing comprises a central shroud which extends around the turbine ring and from which upstream and downstream annular flanges extend radially inwards, the upstream annular flange receiving an upstream retention flange and the downstream annular flange receiving a metal alloy spacer comprising a plurality of spacer sectors arranged end to end circumferentially around the axis of rotation, each ring sector of the turbine ring comprising upstream and downstream attachment lugs extending radially inwards from an annular base of the turbine ring, the upstream attachment lug being fixed to the upstream retention flange by upstream axial fixing pins and the downstream attachment lug to the spacer sector by downstream axial fixing pins, an omega seal being mounted in the spacer sectors bearing against the downstream attachment lug.

[0022] Preferably, the air diffuser has an inclined plane to produce an impact jet on the downstream attachment lug in contact with the omega seal.

[0023] Preferably, the sectorized cooling device comprises one or more ventilation channels comprising a flared opening opening into the plenum chamber. The invention also relates to a turbomachine comprising the aforementioned assembly.

[0024] Brief description of the drawings

[0025] Other characteristics and advantages of the present invention will emerge from the description given below, with reference to the appended drawings which illustrate an exemplary embodiment thereof without any limiting character and in which:

[0026] [Fig. 1] Figure 1 is a general view of an aeronautical turbomachine,

[0027] [Fig. 2] Figure 2 shows in cross-section a turbine ring assembly according to the invention,

[0028] [Fig. 3] Figure 3 is a perspective view of a bushing receiving a play-compensating spring of the turbine ring of Figure 2, and

[0029] [Fig. 4] Figure 4 shows the socket of Figure 3 in the mounting position.

[0030] Description of the embodiments

[0031] In the remainder of the description, the terms "upstream" and "downstream" will be used with reference to the axial direction of flow of the gas flow in the turbine and the terms "internal" or "interior" and "external" or "exterior" are taken in the direction perpendicular to the axial direction of flow of the gas flow depending on whether the object concerned is in contact or not with this gas flow.

[0032] Figure 1 illustrates in longitudinal section and by way of example a dual-flow, dual-spool aircraft turbomachine 10 comprising from upstream to downstream: a fan 12, a first low-pressure compressor 14 and a second high-pressure compressor 16, a combustion chamber 18 and a gas turbine 20, these elements with the exception of the fan being mounted in a casing 22 forming an internal passage duct for the hot primary flow stream of the turbomachine. Guide vanes 24 mounted between the casing 22 and a nacelle 26 serve to channel the secondary flow of the turbomachine. The casing 22 made of metallic material constitutes the support structure for angular sectors of turbine ring made of ceramic matrix composite (CMC) material juxtaposed over 360°, so as to form a sectored turbine ring (discontinuous cylindrical with intersector clearances) surrounding the sets of rotating blades of the turbomachine.

[0033] As illustrated in Figure 2, each turbine ring angular sector 28 has a section substantially in the shape of the inverted Greek letter n (pi) with an annular base whose radially inner face 30A coated with a layer of abradable material delimits the aerodynamic vein of hot air in the gas turbine and a radially outer face 30B from which an upstream attachment lug 32 and a downstream attachment lug 34 extend radially outward. These two attachment lugs which extend in the circumferential direction over the entire width of the ring angular sector 28 are axially spaced from one another and each have at least one ear 32A, 34A pierced with an axial orifice intended to receive an axial fixing pin 36A, 36B.The upstream and downstream pins, advantageously four in number (two per lug) per ring sector, provide the radial connection to the casing 22 respectively via a free end of an upstream retention flange 38 and via free ends of a plurality of spacer sectors 40.

[0034] The spacer sectors, advantageously made of metal alloy, each have a substantially inverted V shape, with a first part 40A (forming the first arm of the V) radially parallel to the upstream retention flange and a second part 40B (forming the second arm of the V) extending substantially longitudinally from this first part to the upstream retention flange, the spacer sectors being fixed to this upstream retention flange by axial assembly screws 42 arranged circumferentially (typically two screws per spacer sector). The tightening force of the assembly screws is taken up mainly by the spacer sectors, thus making it possible not to load the ring sectors in compression.

[0035] A second end of the upstream retention flange 38, opposite its free end, is terminated by a hook 38A extending downstream and intended to engage with an upstream annular radial flange 22A of the turbine casing 22 and second ends of the spacer sectors 40, opposite their free ends (and forming the junction between the two arms of the V) are terminated by a hook 40C also extending downstream and intended to engage with a downstream annular radial flange 22B of the turbine casing. The turbine casing 22 constituting a ring support structure is in fact formed of a central shroud 22C which extends around the turbine ring and from which extend radially, towards the aerodynamic flow path of the hot air gas flow, the upstream annular radial flange 22A and the downstream annular radial flange 22B.The upstream retention flange 38 is secured to the turbine casing 22 by a retaining plate 44 bearing radially on the upstream retention flange and fixed to the casing by a set of retaining screws 46. In order to allow axial expansion of the CMC ring sectors, a set of flexible seals of the omega seal type 48 is mounted in the first part 40A of the spacer sectors bearing against the downstream attachment lug 34 of these ring sectors.

[0036] In a known manner, the seal between angular ring sectors is ensured by inter-sector tabs (not shown) housed in grooves of the angular ring sectors which have two-by-two grooves arranged opposite each other circumferentially.

[0037] Once the ring sectors are juxtaposed (upstream retention flange and spacer sectors in place), the external face 30B of the annular base and the radial attachment lugs 32, 34 of the turbine ring form, with the upstream retention flange 38 and the spacer sectors 40, an off-stream annular cavity 50, in other words a pressurized cavity external to the aerodynamic hot air stream, in which is mounted a sectored cooling device 52 bringing cooling air from upstream, via one or more ventilation channels 52A and a plenum 52B of the device, to a sectored air diffuser 54 (comprising a set of holes arranged in a matrix manner) making it possible to ensure impact cooling of the annular base of the ring thanks to the pressure difference existing between the off-stream cavity 50 and the aerodynamic vein of hot air.This pressure differential also makes it possible to keep the ring 28 pressed against the upstream retention flange 38 and the radial parts of the spacer sectors 40A. The ventilation channel(s) 52A which also pass through the upstream retention flange 38 are protected by a dust filter 56 mounted at their inlet on this upstream retention flange. The ventilation channel(s) 52A each have a flared opening opening into the plenum chamber 52B allowing better cooling efficiency.

[0038] To control the clearance between the tip of the high-pressure moving blade of the turbine and the turbine ring during the flight cycle of the aircraft, radial springs are provided to compensate for the mounting gap between the upstream 36A and downstream 36B pins and the holes receiving them in the upstream 32A and downstream 34A attachment lugs of the ring sectors respectively.

[0039] According to the invention, and as shown more precisely in Figures 3 and 4, for each ring sector 28, at least one sleeve 60 made of a high-temperature resistant metal alloy, having the shape of a stepped cylinder with two stages (a first or lower stage 60A and a second or upper stage 60B) has its lower face 601 of the lower cylinder 60A bearing on the radially external face 30B of the annular base of the turbine ring. Preferably, two sleeves 60 are used for each ring sector 28. The sleeve(s) are housed inside the cooling device 52 in two cavities 62 present at the circumferential ends of the air diffuser, so as not to interfere with the plane of holes in the impact sheet or with the impact jets, and thus to have uniform cooling of the turbine ring.The bushing(s) are connected to the rest of the fixed structure of the turbine by means of the helical springs 58 compressed between the upper face 602 of the lower cylinder 60A and the upper face 620 of the cavity 62, the latter having a machining allowing each bushing 60 to slide freely therein. The second cylinder 60B of each bushing substantially the same diameter as the helical spring 58 and forming a centering stud allows it to be received and therefore to guide its elongation. The first cylinder or lower cylinder 60A of each bushing which constitutes the base is split over almost its entire height by a parallelepiped channel 603 which defines for the remaining part of the base two wings 604 whose rectilinear internal faces are parallel to the annular flanges 32, 34 so that this widened channel between the two wings can allow the cooling air to circulate by impact (illustrated by the arrow in Figure 4).This base defines the air gap with the turbine ring to ensure a controlled impact jet. The first cylinder 60A also has a flat 605 allowing its translational guidance by blocking the sleeve in rotation, relative to the cooling device 52, and each of the wings 604 has a chamfer 606 at the level of the internal face 601 in contact with the radially external face 30B of the annular base, to minimize the heat exchange with the turbine ring.

[0040] Each bushing 60 also has a central bore 607 for cooling by impact on the surface of the ring opposite the bushing. It is also advantageous to provide a bore at the bottom of the cavity 62 for cooling the springs with cooling air.

[0041] In operation, when the parts expand, this expansion is compensated by the deformation of the springs, allowing the parts to be held in position without over-stressing them.

[0042] Finally, this last characteristic will be noted, which contributes to improving cooling and relates to the air diffuser 54 in the form of an air diffusion plate secured to the radially internal face of the plenum chamber 52B which has an inclined plane 54A to produce an impact jet on the downstream attachment lug 34 axially in contact with the omega seal 48 so as to limit the temperature in this sensitive zone.

[0043] Thus, the invention makes it possible to respond to several problems:

[0044] - Keep the ring pressed on the pins without thermomechanical over-stress: This problem was resolved with the use of the spring whose deformation will compensate for differential expansions.

[0045] - Protect the spring from high temperatures: Indeed, if the spring were in direct contact with the ring, it would quickly lose its elasticity, especially since this temperature could exceed the permissible material. The contact with the pad is a few hundred degrees colder than the ring.

[0046] - Ensure ring cooling: The drilling in the pad allows cooling by impact.

Claims

Claims

1. Gas turbine ring assembly, comprising around a longitudinal axis of rotation (X), a turbine ring (28) comprising a plurality of ring sectors made of ceramic matrix composite material arranged end to end circumferentially around the axis of rotation (X), a casing (22) made of metal alloy forming a ring support structure, and a sectored cooling device (52) comprising a plenum chamber (52B) delimited by a sectored air diffuser (54) having holes arranged in a matrix manner for diffusing by impact cooling air jets onto a radially external surface (30B) of the turbine ring (28) from which it is separated by a diffusion distance, assembly characterized in that it further comprises at least one sleeve (60), preferably two sleeves (60) made of metal alloy housed in two dedicated cavities (62) present at the circumferential ends of the air diffuser (54),each sleeve (60) having the shape of a two-stage stepped cylinder with a base forming the lower stage (60A) of which a lower face (601) comes to bear on the radially external surface (30B) of the turbine ring (28) and a stud surmounting the base to form the upper stage (60B) intended to receive a spring (58) configured to be compressed between the base (60A) and the cavity (62).,

2. The turbine ring assembly of claim 1, wherein the base (60A) of each bushing (60) defines the diffusion distance between the air diffuser (54) and the radially outer surface (30B) of the turbine ring (28).

3. A turbine ring assembly according to claim 2, wherein the base (60A) of each sleeve (60) is split over almost its entire height, so as to define two wings (604) whose parallel internal faces define an enlarged channel (603) allowing circumferential flow of the impact jet.

4. A turbine ring assembly according to claim 3, wherein each of the wings (604) has a radially inner face (601) which has a chamfer (606), the radially inner face (601) of each of the wings (604) being in contact with the radially outer surface (30B) of the turbine ring (28).

5. Turbine ring assembly according to any one of claims 1 to 4, in which the base (60A) of each sleeve (60) has a flat (605) to prevent its rotation and allow its translational guidance.

6. A turbine ring assembly according to any one of claims 1 to 5, wherein each bushing (60) has a central bore (607) for providing impact cooling to the radially outer surface (30B) of the turbine ring (28).

7. A turbine ring assembly according to any one of claims 1 to 6, wherein the casing (22) comprises a central shroud (22C) which extends around the turbine ring (28) and from which upstream and downstream annular flanges (22A, 22B) extend radially inward, the upstream annular flange (22A) receiving an upstream retention flange (38) and the downstream annular flange (22B) receiving a metal alloy spacer comprising a plurality of spacer sectors (40) arranged end to end circumferentially around the axis of rotation (X), each ring sector of the turbine ring (28) comprising upstream (32A) and downstream (34A) attachment lugs extending radially inward from an annular base (30) of the turbine ring (28),the upstream attachment lug (32A) being fixed to the upstream retention flange by upstream axial fixing pins (36A) and the downstream attachment lug (34A) to the spacer sector (40) by downstream axial fixing pins (36B), an omega seal (48) being mounted in the spacer sectors (40) bearing against the downstream attachment lug (34A).,

8. A turbine ring assembly according to claim 7, wherein the air diffuser (54) has an inclined plane (54A) for producing an impact jet on the downstream attachment lug (34A) in contact with the omega seal (48).

9. A ring assembly according to any one of claims 1 to 8, wherein the sectored cooling device (52) comprises one or more ventilation channels (52A) having a flared opening opening into the plenum (52B).

10. An aeronautical turbomachine comprising a turbine ring assembly according to any one of claims 1 to 9.

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