Improved Backlash Compensation Turbine Ring Assembly

Abstract

Enhanced Backlash Compensation Turbine Ring Assembly Gas turbine ring assembly, comprising a turbine ring including ring sectors (28) of ceramic matrix composite material, a casing (22) of metal alloy and a sectored cooling device (52) including a stilling chamber (52B) delimited by an air diffuser (54) having holes for diffusing cooling air jets onto a radially external surface (30B) of the turbine ring.The assembly further comprises two metal alloy bushings (60) housed in two cavities (62) located at the circumferential ends of the air diffuser (54), each bushing being in the form of a stepped cylinder with a base (60A) whose lower face (601) bears against the radially external surface (30B) of the turbine ring and a stud surmounting the base (60B) and intended to receive a spring (58) compressed between an upper face (602) of the base and an upper face (620) of the cavity. Figure for the abbreviation: Fig. 2.

Description

Title of the invention: Improved backlash compensation turbine ring assembly technical field

[0001] The present invention relates to the field of gas turbines such as turbines of aircraft engines or helicopter engines, turbines of electric generators and turbochargers. Previous technique

[0002] Climate change is a major concern for many legislative and regulatory bodies worldwide. Indeed, various restrictions on carbon emissions have been, are being, or will be adopted by various states. In particular, an ambitious standard applies both to new types of aircraft and to those currently in operation, requiring the implementation of technological solutions to bring them into compliance with current regulations. Civil aviation has been actively working for several years now to contribute to the fight against climate change.

[0003] Technological research efforts have already led to significant improvements in the environmental performance of aircraft. The Applicant takes into account the factors impacting all phases of design and development in order to obtain aeronautical components and products that are less energy-intensive, more environmentally friendly, and whose integration and use in civil aviation have moderate environmental impacts, with the aim of improving the energy efficiency of these aircraft.

[0004] 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 in order to reduce the environmental footprint of its activity.

[0005] This sustained research and development work focuses on new generations of aircraft turbomachinery, the weight reduction of aircraft, particularly through the materials used and lighter on-board equipment, the development of the use of electrical technologies to provide propulsion, and, as essential complements to technological progress, aviation biofuels.

[0006] Current trends in civil aviation and the framework for future civil aircraft engines imply a requirement for high performance at high temperatures, which necessitates the integration of parts made of ceramic matrix composite (CMC) material. in the high-pressure (HP) turbine. CMC has the advantage of withstanding high temperatures (over 1200°C) and its low density (three times less than metallic bases) allows for a significant weight saving.

[0007] However, integrating these CMC parts into an assembly of metallic parts generates stresses due to the different coefficients of expansion existing between the CMC and the metallic elements. It is therefore necessary to find technologies capable of overcoming these problems by incorporating backlash compensation technologies.

[0008] Currently, the CMC turbine rings developed by the applicant are shaped like an inverted letter PI. Their straight, horizontal base forms an aerodynamic channel that guides the combustion gases and is coated with an abradable material that protects the CMC from external damage. The two tabs of this PI ring allow for radial positioning of the base, thanks to cylindrical metal pins that attach them to metal parts connected to the turbine housing. Due to the greater expansion of the pins compared to the ring tabs, the contact between these two elements must be made with some play (by enlarging the bore diameter of the ring tabs), creating a degree of radial translational freedom that no longer allows the ring to be held correctly in position.

[0009] There is therefore still a need today for a solution which still consists of imposing a zero radial displacement on the turbine ring while mitigating the aforementioned disadvantages of the presence of springs. Description of the invention

[0010] 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. Therefore, the main objective of the present invention is to integrate a CMC ring into a gas turbine by compensating for differential expansion without mechanical overstress while protecting the flexible elements from the high temperatures of the combustion chamber.

[0011] These objectives 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 settling chamber delimited by an air diffuser having holes arranged in a matrix pattern to diffuse cooling air jets by impact. dissement on 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 bushing, preferably two bushings of metallic alloy housed in two dedicated cavities present at the circumferential ends of the air diffuser, each bushing having the form of a stepped cylinder with two stages with a base forming the lower stage, one lower face of which rests 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.

[0012] Thus, this embedded spring system makes it possible to hold the CMC ring in position in a metallic environment without compromising the effectiveness of the springs at high temperatures. The cooling of the ring is no longer hindered by the presence of the spring. The presence of an orifice that passes through each sleeve around which the spring is wound improves this cooling.

[0013] Preferably, the base of each socket defines the diffusion distance between the air diffuser and the radially external surface of the turbine ring.

[0014] Advantageously, the base of each socket is split over almost its entire height, so as to define two wings whose parallel inner faces define an enlarged channel for a circumferential flow of the impact jet.

[0015] Preferably, each of the wings has a chamfer at the level of the radially internal face in contact with the radially external surface of the turbine ring.

[0016] Advantageously, the base of each socket has a flat to prevent its rotation and to allow its guidance in translation.

[0017] Preferably, each bushing has a central bore allowing also to ensure impact cooling on the radially external surface of the turbine ring.

[0018] Advantageously, the housing comprises a central ferrule which extends around the turbine ring and from which upstream and downstream annular flanges extend radially inwards, the upstream annular flange receiving an upstream retaining flange and the downstream annular flange a spacer made of a metal alloy 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 retaining 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.

[0019] Preferably, the air diffuser has an inclined plane to produce an impact jet on the downstream attachment tab in contact with the omega seal.

[0020] Preferably, the sectorized cooling device includes one or more ventilation channels having a flared opening leading into the tranquilization chamber.

[0021] The invention also relates to a turbomachine comprising the aforementioned assembly. Brief description of the drawings

[0022] Other features and advantages of the present invention will become apparent from the description given below, with reference to the accompanying drawings which illustrate an example of an embodiment without being limiting in any way and on which:

[0023] [Fig-1] [Fig.1] is a general view of an aeronautical turbomachine,

[0024] [Fig.2] [Fig.2] shows a cross-section of a turbine ring assembly in accordance with the invention,

[0025] [Fig.3] [Fig.3] is a perspective view of a sleeve receiving a backlash compensation spring for the turbine ring of [Fig.2], and

[0026] [Fig.4] [Fig.4] shows the socket of [Fig.3] in the mounting position. Description of the implementation methods

[0027] In the following description, the terms "upstream" and "downstream" will be used with reference to the axial direction of flow of the gas stream in the turbine and the terms "internal" or "inside" and "external" or "outside" are taken in the direction perpendicular to the axial direction of flow of the gas stream depending on whether the object concerned is in contact with this gas stream or not.

[0028] Figure 1 illustrates, in longitudinal section and by way of example, a twin-spool, twin-flow 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 passageway for the hot primary flow 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 the turbine ring made of ceramic matrix composite (CMC) material juxtaposed over 360°, so as to form a sectorized turbine ring (discontinuous cylindrical with inter-sector clearances) surrounding the rotating blade assemblies of the turbomachine.

[0029] As illustrated in [Fig. 2], each angular sector of the turbine ring 28 has a cross-section substantially in the shape of the inverted Greek letter ir (pi) with an annular base whose radially inner face 30A is coated with a layer of material The abradable surface delimits the aerodynamic hot air stream in the gas turbine and a radially external face 30B from which an upstream mounting lug 32 and a downstream mounting lug 34 extend radially outwards. These two mounting lugs, which extend circumferentially over the entire width of the angular sector of the ring 28, are axially spaced from each other and each has at least one lug 32A, 34A pierced with an axial orifice for receiving 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 housing 22 respectively via a free end of an upstream retaining flange 38 and via the free ends of a plurality of spacer sectors 40.

[0030] The spacer sectors, advantageously made of a metal alloy, each have a substantially inverted V shape, with a first radial portion 40A (forming the first arm of the V) parallel to the upstream retaining flange and a second portion 40B (forming the second arm of the V) extending substantially longitudinally from this first portion to the upstream retaining flange. The spacer sectors are fixed to this upstream retaining flange by axial assembly screws 42 arranged circumferentially (typically two screws per spacer sector). The tightening force of the assembly screws is mainly borne by the spacer sectors, thus preventing the ring sectors from being loaded in compression.

[0031] A second end of the upstream retaining 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 housing 22 and the 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 housing. The turbine casing 22, which forms a ring support structure, is in fact formed of a central ferrule 22C which extends around the turbine ring and from which extend radially, towards the aerodynamic flow of the hot air gas flow, the upstream annular radial flange 22A and the downstream annular radial flange 22B.The upstream retaining flange 38 is secured to the turbine housing 22 by a retaining plate 44 bearing radially on the upstream retaining flange and fixed to the housing by a set of retaining screws 46. In order to allow axial expansion of the CMC ring sectors, a set of flexible omega-type seals 48 is mounted in the first part 40A of the spacer sectors bearing against the downstream attachment lug 34 of these ring sectors.

[0032] In a known manner, the seal between angular ring sectors is ensured by inter-sector tabs (not shown) housed in grooves in the angular ring sectors, which have two pairs of grooves arranged opposite each other in such a way circumferential.

[0033] 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 mounting lugs 32, 34 of the turbine ring, together with the upstream retention flange 38 and the spacer sectors 40, form an annular off-stream cavity 50, in other words, a pressurized cavity external to the aerodynamic hot air stream, in which a sectorized cooling device 52 is mounted, supplying cooling air from upstream, via one or more ventilation channels 52A and a settling chamber 52B of the device, to a sectorized air diffuser 54 (comprising a set of holes arranged in a matrix pattern), enabling impact cooling of the annular base of the ring by means of the pressure difference existing between the off-stream cavity 50 and the aerodynamic vein of hot air.This pressure differential also helps to keep the ring 28 pressed against the upstream retaining flange 38 and the radial parts of the spacer sectors 40A. The ventilation channel(s) 52A, which also pass through the upstream retaining flange 38, are protected by a dust filter 56 mounted at their inlet on this upstream retaining flange. Each ventilation channel(s) 52A has a flared opening leading into the settling chamber 52B, allowing for improved cooling efficiency.

[0034] To control the play between the top 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 pins 36A and downstream pins 36B and the orifices receiving them in the upstream 32A and downstream 34A attachment tabs of the ring sectors respectively.

[0035] According to the invention, and as more specifically shown 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 two-stage stepped cylinder (a first or lower stage 60A and a second or upper stage 60B), has its lower face 601 of the lower cylinder 60A bearing against 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 located at the circumferential ends of the air diffuser, so as not to interfere with the drilling plane of the impact plate 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 turbine's fixed structure by means of 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 machined surface allowing each bushing 60 to slide freely within it. The . The second cylinder 60B of each sleeve, approximately the same diameter as the helical spring 58 and forming a centering pin, receives the spring and thus guides its extension. The first or lower cylinder 60A of each sleeve, which forms its base, is split along almost its entire height by a parallelepiped channel 603. This channel defines two wings 604 for the remaining part of the base, the straight inner faces of which are parallel to the annular flanges 32 and 34. This allows the enlarged channel between the two wings to permit the circulation of impact cooling air (illustrated by the arrow in [Fig. 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 inner face 601 in contact with the radially outer face 30B of the annular base, to minimize heat exchange with the turbine ring.

[0036] Each sleeve 60 also has a central hole 607 allowing impact cooling of the ring surface opposite the sleeve. It is also advantageous to provide a hole in the bottom of the cavity 62 to cool the springs with the cooling air.

[0037] 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.

[0038] Finally, we will note this last characteristic contributing to the improvement of cooling and relating to the air diffuser 54 in the form of an air diffusion plate attached to the radially internal face of the stilling 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 area.

[0039] Thus, the invention makes it possible to address several problems:

[0040] - Maintain the ring pressed against the pins without thermomechanical over-stress: This problem was solved with the use of the spring, whose deformation will compensate for differential expansions.

[0041] - Protecting the spring from high temperatures: Indeed, if the spring were directly In contact with the ring, it would quickly lose its elasticity, especially since this temperature could exceed the material's permissible temperature. The pad is several hundred degrees cooler than the ring.

[0042] - Ensuring cooling of the ring: The hole in the pad allows cooling by impact.

Claims

Demands

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 metal alloy housing (22) forming a ring support structure, and a sectorized cooling device (52) comprising a settling chamber (52B) delimited by a sectorized air diffuser (54) having holes arranged in a matrix pattern for diffusing cooling air jets by impact 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 bushing (60), preferably two bushings (60) made of metal alloy housed in two dedicated cavities (62) present at the circumferential ends of the air diffuser (54),each socket (60) having the form of a two-stage stepped cylinder with a base forming the lower stage (60A), the lower face (601) of which bears against 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. Turbine ring assembly according to claim 1, wherein the base (60A) of each socket (60) defines the diffusion distance between the air diffuser (54) and the radially external surface (30B) of the turbine ring (28).

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

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

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

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

7. Turbine ring assembly according to any one of claims 1 to 6, wherein the housing (22) comprises a central ferrule (22C) extending 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 retaining flange (38) and the downstream annular flange (22B) 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 bracket (32A) being fixed to the upstream retention flange by upstream axial fixing pins (36A) and the downstream attachment bracket (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 bracket (34A).

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

9. Ring assembly according to any one of claims 1 to 8, wherein the sectorized cooling device (52) comprises one or more ventilation channels (52A) having a flared opening leading into the stilling chamber (52B).

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

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