Turbine ring assembly with improved gap compensation

Abstract

A gas turbine shroud assembly comprises a turbine shroud comprising shroud segments (28) made of ceramic matrix composite, a casing (22) made of a metal alloy and a segmented cooling device (52) comprising a plenum (52B) delimited by an air diffuser (54) having holes for diffusing cooling air jets onto a radially outer side 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 bushing is in the form of a stepped cylinder with a base (60A) and a head (60B) placed on top of the base, the lower face (601) of the base (60A) bearing on the radially outer side surface (30B) of the turbine shroud. The head is intended to receive a spring (58) compressed between the upper face (602) of the base and the upper face (620) of the cavity.

Description

Technical Field

[0001] This invention relates to the field of gas turbines, such as turbines for aircraft or helicopter engines, generators, and turbines for turbochargers. Background Technology

[0002] Climate change is a major concern for many legislative and regulatory bodies around the world. In fact, countries have already implemented, are implementing, or will implement various restrictions on carbon emissions. In particular, ambitious standards apply to both new and currently in-service aircraft, requiring the implementation of technological solutions to ensure compliance with existing regulations. Civil aviation has been actively involved in addressing climate change for many years.

[0003] Technological research has yielded significant improvements in the environmental performance of aircraft. The applicant considers the influencing factors at each stage of design and development to obtain more energy-efficient and environmentally friendly aviation components and products, whose integration and use in civil aviation have moderate environmental consequences, aiming to improve the energy efficiency of aircraft.

[0004] Therefore, the applicant continues to reduce climate impact and minimize greenhouse gas emissions by employing sound development and manufacturing methods and processes, thereby reducing the environmental footprint of the activity.

[0005] These ongoing research and development efforts focus on next-generation aircraft turbines, aircraft weight reduction, particularly through the use of materials and lightweight airborne equipment, the development of electrical technologies to ensure propulsion, and aviation biofuels as an important complement to technological advancements.

[0006] Current trends in civil aviation and the framework for future civil aircraft engines involve the requirement for high performance at high temperatures, which necessitates the integration of ceramic matrix composite (CMC) components into high-pressure (HP) turbines. CMCs do indeed offer the advantage of withstanding high temperatures (above 1200°C), and their low density (three times lighter than a metal matrix) allows for significant mass gains.

[0007] However, integrating these ceramic matrix composite components into metal component assemblies can lead to stress due to the different coefficients of thermal expansion between the ceramic matrix composites and the metal components. Therefore, it is necessary to find a technology that can overcome these problems through backlash compensation techniques.

[0008] Currently, the applicant's ceramic matrix composite turbine ring has an inverted π shape, with its straight and horizontal base forming an aerodynamic flow path that allows combustion gases to be guided. An abrasive material is deposited on this base to protect the ceramic matrix composite from external erosion. The two legs of this π-shaped ring allow for radial positioning of the base thanks to cylindrical metal pins that allow them to be attached to metal components connected to the turbine housing. Because the pins have a greater expansion than the ring legs, the contact between these two elements must be made with a gap (by increasing the bore diameter of the ring legs), resulting in a radial translational degree of freedom that no longer allows the ring to be properly held in place.

[0009] Therefore, a solution is still needed that still involves applying zero radial displacement to the turbine ring while overcoming the aforementioned drawbacks of the presence of springs. Summary of the Invention

[0010] Therefore, this invention is the result of technical research aimed at significantly improving aircraft performance, and in this sense, contributing to reducing the environmental impact of these aircraft. The main objective of this invention is to integrate ceramic matrix composite rings into gas turbines by compensating for differential expansion without mechanical overstress while protecting flexible components from the high temperatures of the combustion chamber.

[0011] These objectives are achieved by a gas turbine ring assembly comprising, about a longitudinal axis of rotation: a turbine ring comprising multiple ring segments made of ceramic matrix composite material arranged end-to-end circumferentially around the axis of rotation; a metal alloy housing forming a ring support structure; and a segmented cooling device comprising a pressurization chamber defined by an air diffuser having orifices arranged in a matrix to cool an air jet by impingement diffusion on the radially outer surface of the turbine ring, the air diffuser being separated from the radially outer surface by a diffusion distance. The assembly is characterized by further comprising at least one bushing, preferably two bushings, made of a metal alloy, housed in two dedicated cavities present at the circumferential ends of the air diffuser, each bushing having the form of a two-stage stepped cylinder with a collar forming the lower stage and a head positioned above the collar to form the upper stage, the lower part of the collar being supported on the radially outer surface of the turbine ring, the head intended to receive a spring configured to compress between the collar and the cavity.

[0012] Therefore, this embedded spring system allows ceramic matrix composite rings to be held in place in metallic environments without compromising the spring's effectiveness at high temperatures. Cooling of the rings is no longer hindered by the presence of the springs. The presence of orifices through each bushing, around which the springs wrap, further enhances this cooling.

[0013] Preferably, the collar of each bushing defines the diffusion distance between the air diffuser and the radially outer surface of the turbine ring.

[0014] Advantageously, the collar of each bushing is slit along almost its entire height, thereby defining two wing sections whose parallel inner surfaces define a widened channel for circumferential flow of the impinging jet.

[0015] Preferably, each wing has a chamfer at the level where its radially inner surface contacts the radially outer surface of the turbine ring.

[0016] Advantageously, each bushing's collar has a flat section to prevent its rotation and allow its translational guidance.

[0017] Preferably, each bushing includes a central bore, which also allows cooling to be ensured by impact on the radially outer surface of the turbine ring.

[0018] Advantageously, the housing includes a central shroud extending around the turbine ring, and an upstream annular jaw and a downstream annular jaw extending radially inward from the central shroud. The upstream annular jaw receives an upstream retaining flange, and the downstream annular jaw receives a spacer made of a metal alloy. The spacer includes a plurality of spacer segments arranged circumferentially end-to-end around the axis of rotation. Each annular segment of the turbine ring includes an upstream hook leg and a downstream hook leg extending radially inward from the annular base of the turbine ring. The upstream hook leg is secured to the upstream retaining flange by an upstream axial retaining pin, and the downstream hook leg is secured to the spacer segment by a downstream axial retaining pin. An Ω-shaped seal is mounted in the spacer segment, abutting against the downstream hook leg.

[0019] Preferably, the air diffuser has an inclined plane to generate an impinging jet on the downstream hook leg that contacts the Ω-shaped seal.

[0020] Preferably, the segmented cooling device includes one or more ventilation channels, each including a flared opening leading to the pressurization chamber.

[0021] The present invention also relates to a turbine that includes the aforementioned components. Attached Figure Description

[0022] Other features and advantages of the invention will become apparent from the following description with reference to the accompanying drawings, which illustrate an exemplary embodiment without any limitations, in which: [ Figure 1 ] Figure 1 This is an overall view of an aircraft turbine. [ Figure 2 ] Figure 2 A cross-sectional view of the turbine ring assembly according to the present invention is shown. [ Figure 3 ] Figure 3 It is receiving Figure 2 A three-dimensional view of the bushing of the turbine ring clearance compensation spring, and [ Figure 4 ] Figure 4 Showing the installation position Figure 3 The bushing. Detailed Implementation

[0023] In the following description, the terms “upstream” and “downstream” will be used with reference to the axial flow direction of the airflow in the turbine, while the terms “inner” or “internal” and “outer” or “external” will be used in a direction perpendicular to the axial flow direction of the airflow, depending on whether the object in question is in contact with the airflow.

[0024] Figure 1 The example illustrates a longitudinal section of a dual-rotor turbofan aircraft engine 10, which, from upstream to downstream, includes: 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 components, except for the fan, are housed in a casing 22, which forms an inner channel duct for the main thermal flow path of the turbine. Guide vanes 24, installed between the casing 22 and the nacelle 26, guide the secondary flow of the turbine. The casing 22, made of metallic material, constitutes a support structure for angular turbine ring segments made of ceramic matrix composite (CMC) material juxtaposed in a 360° configuration, thereby forming segmented turbine rings (discontinuous cylindrical sections with intersegment gaps) surrounding the rotating blade assembly of the turbine.

[0025] like Figure 2 As illustrated, each angular turbine ring segment 28 has a cross-section generally shaped like an inverted Greek letter π (pi), with an annular base having a radially inner surface 30A and a radially outer surface 30B. The radially inner surface 30A is coated with a wearable material layer that defines the aerodynamic hot airflow path in the gas turbine. An upstream hook leg 32 and a downstream hook leg 34 extend radially outward from the radially outer surface 30B. These two hook legs extend circumferentially over the entire width of the angular ring segment 28, are axially spaced from each other, and each has at least one lug 32A, 34A with lug perforations intended to receive axial orifices of axial retaining pins 36A, 36B. Advantageously, there are four upstream and downstream pins per ring segment (two per leg), ensuring radial connection to the housing 22 via the free end of the upstream retaining flange 38 and via the free ends of a plurality of spacer segments 40, respectively.

[0026] The spacer segments are advantageously made of a metal alloy and each has a generally inverted V shape, with a first radial portion 40A (forming a first arm of the V) parallel to the upstream retaining flange and a second portion 40B (forming a second arm of the V) extending generally longitudinally from the first portion to the upstream retaining flange. The spacer segments are secured to the upstream retaining flange by circumferentially arranged axial assembly screws 42 (typically two screws per spacer segment). The clamping force of the assembly screws is primarily borne by the spacer segments, thereby making it possible to avoid loading the ring segments under compression.

[0027] The second end of the upstream retaining flange 38, opposite its free end, terminates at a hook portion 38A extending downstream and intended to engage with the upstream annular radial flange 22A of the turbine housing 22. Similarly, the second end of the spacer section 40, opposite its free end (forming the junction between the two arms of the V), terminates at a hook portion 40C also extending downstream and intended to engage with the downstream annular radial clamp 22B of the turbine housing. The turbine housing 22, constituting the ring support structure, is indeed formed by a central shroud 22C extending around the turbine ring, with the upstream annular radial clamp 22A and the downstream annular radial clamp 22B extending from the central shroud towards the aerodynamic flow path of the hot air flow. The upstream retaining flange 38 is secured to the turbine housing 22 by a retaining piece 44 that abuts radially against the upstream retaining flange and is fixed to the housing by a set of retaining screws 46. To allow for axial expansion of the ceramic matrix composite ring sections, a set of flexible seals 48 of the Ω-shaped seal type are installed in the first part 40A of the spacer section, abutting against the downstream hook leg 34 of these ring sections.

[0028] As is known, the seal between the angular ring segments is ensured by inter-segment tabs (not shown) housed in the grooves of the angular ring segments, which have pairs of grooves arranged circumferentially facing each other.

[0029] Once the ring sections are aligned (with the upstream retaining flange and spacer section in place), the outer surface 30B of the annular base and the radial hook legs 32, 34 of the turbine ring, together with the upstream retaining flange 38 and spacer section 40, form an annular cavity 50 located outside the flow path. In other words, it is a pressurized cavity located outside the hot air aerodynamic flow path. A segmented cooling device 52 is installed in this pressurized cavity, supplying cooling air from upstream to the segmented air diffuser 54 (including a set of orifices arranged in a matrix) via one or more ventilation channels 52A and a pressurization chamber 52B of the device, ensuring cooling of the annular base of the ring due to the pressure difference between the cavity 50 outside the flow path and the aerodynamic hot air flow path. This pressure difference also allows the retaining ring 28 to press against the radial portion of the upstream retaining flange 38 and spacer section 40A. The ventilation channels 52A, which also pass through the upstream retaining flange 38, are protected by dust filters 56 installed at their inlets on the upstream retaining flange. Each of the (one or more) ventilation ducts 52A has a flared opening leading to the pressurization chamber 52B, allowing for better cooling efficiency.

[0030] In order to control the clearance between the top of the turbine high-pressure blades and the turbine ring during the flight cycle of the aircraft, radial springs are provided to compensate for the installation clearance between the upstream pin 36A and the downstream pin 36B and the orifices that receive them in the upstream hook leg 32A and the downstream hook leg 34A of the ring section, respectively.

[0031] According to the present invention, and as Figure 3 and Figure 4More specifically, for each annular segment 28, at least one bushing 60 made of a high-temperature resistant metal alloy has a two-stage stepped cylindrical form (a first or lower stage 60A and a second or upper stage 60B), the lower cylinder 60A's lower surface 601 supported on the radially outer surface 30B of the annular base of the turbine ring. Preferably, two bushings 60 are used for each annular segment 28. One or more bushings are housed within a cooling device 52 in two cavities 62 located at the circumferential ends of the air diffuser, so as not to interfere with the drilled plane of the impact plate or the impact jet, and thus to provide uniform cooling of the turbine ring. One or more bushings are connected to the remaining fixed structure of the turbine by means of a helical spring 58 compressed between the upper surface 602 of the lower cylinder 60A and the upper surface 620 of the cavity 62, the cavity 62 having a machined surface that allows each bushing 60 to slide freely therein. The second cylinder 60B of each bushing is approximately the same diameter as the coil spring 58 and forms a centering head that allows the spring to be received and thus guides its elongation. The first or lower cylinder 60A of each bushing, constituting its collar, is slit along almost its entire height by a parallelepiped channel 603, which defines two wings 604 for the remainder of the collar. The straight inner surfaces of the wings are parallel to the annular jaws 32, 34, such that this widened channel between the two wings allows shock cooling air to circulate (by... Figure 4 (Indicated by the middle arrow). The collar defines an air gap with the turbine ring to ensure a controlled impingement jet. The first cylinder 60A also has a flat section 605 that allows the translation of the blocking bushing relative to the cooling device 52 to be guided, and each wing 604 has a chamfer 606 at the level where the inner side 601 contacts the radially outer side 30B of the annular base to minimize heat exchange with the turbine ring.

[0032] Each bushing 60 also has a central drilled hole 607, allowing cooling by impact on the bushing-facing surface of the ring. It is also advantageous to provide a drilled hole at the bottom of the cavity 62 to cool the spring with cooling air.

[0033] During operation, as the components expand, this expansion is compensated by the deformation of the springs, which allows the components to remain in place without applying overstress to them.

[0034] Finally, note the last feature involved in improving cooling, which involves an air diffuser 54, which is in the form of an air diffuser plate fixed to the radially inner side of the pressurization chamber 52B. This air diffuser plate has an inclined plane 54A to generate an impinging jet on the downstream hook leg 34 in axial contact with the Ω-shaped seal 48, thereby limiting the temperature in this sensitive area.

[0035] Therefore, this invention allows for the solution of several problems: - Keeping the ring pressed against the pin without thermomechanical overstress: This problem is solved by using a spring, whose deformation will compensate for differential expansion.

[0036] - Protect the spring from high temperatures: Indeed, if the spring is in direct contact with the ring, it will quickly lose its elasticity, especially since the temperature may exceed the material's tolerance. Contact with the pad is several hundred degrees cooler than with the ring.

[0037] - Ensure ring cooling: Drilling holes in the pads allow for cooling through impact.

Claims

1. A gas turbine ring assembly, comprising: A turbine ring (28) surrounding a longitudinal axis of rotation (X), the turbine ring (28) comprising multiple annular segments made of ceramic matrix composite material, the annular segments being arranged end-to-end circumferentially around the axis of rotation (X); a metal alloy housing (22) forming a ring support structure; and a segmented cooling device (52) comprising a pressurization chamber (52B) defined by a segmented air diffuser (54) having orifices arranged in a matrix to cool air jets by impingement diffusion on the radially outer surface (30B) of the turbine ring (28), the air diffuser (54) and the radially outer surface (30B) 30B) Separating the diffusion distance, the component is characterized in that it further includes at least one bushing (60), preferably two bushings (60), the bushings (60) being made of a metal alloy and housed in two dedicated cavities (62) present at the circumferential end of the air diffuser (54), each bushing (60) having the form of a two-stage stepped cylinder with a collar forming the lower stage (60A) and a head placed above the collar to form the upper stage (60B), the lower part (601) of the collar being supported on the radially outer surface (30B) of the turbine ring (28), the head being intended to receive a spring (58) configured to compress between the collar (60A) and the cavity (62).

2. The turbine ring assembly according to claim 1, characterized in that, The collar (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. The turbine ring assembly according to claim 2, characterized in that, The collar (60A) of each bushing (60) is slit along almost its entire height, thereby defining two wings (604), the parallel inner surfaces of which define a widened channel (603) that allows circumferential flow of the impinging jet.

4. The turbine ring assembly according to claim 3, characterized in that, Each of the wing portions (604) has a radially inner surface (601) with a chamfer (606), and the radially inner surface (601) of each of the wing portions (604) contacts the radially outer surface (30B) of the turbine ring (28).

5. The turbine ring assembly according to any one of claims 1 to 4, characterized in that, The collar (60A) of each bushing (60) has a flat section (605) to prevent rotation and allow translational guidance.

6. The turbine ring assembly according to any one of claims 1 to 5, characterized in that, Each of the bushings (60) includes a central bore (607) that allows cooling to be ensured by impact on the radially outer surface (30B) of the turbine ring (28).

7. The turbine ring assembly according to any one of claims 1 to 6, characterized in that, The housing (22) includes a central shroud (22C) extending around the turbine ring (28), and an upstream annular jaw (22A) and a downstream annular jaw (22B) extending radially inward from the central shroud (22C). The upstream annular jaw (22A) receives an upstream retaining flange (38), and the downstream annular jaw (22B) receives spacers made of a metal alloy, the spacers comprising a plurality of spacer segments (40) arranged end-to-end circumferentially around a rotation axis (X). The turbine ring (28) Each annular segment includes an upstream hook leg (32A) and a downstream hook leg (34A) extending radially inward from the annular base (30) of the turbine ring (28). The upstream hook leg (32A) is fixed to the upstream retaining flange by an upstream axial retaining pin (36A), and the downstream hook leg (34A) is fixed to the spacer segment (40) by a downstream axial retaining pin (36B). An Ω-shaped seal (48) is installed in the spacer segment (40) and abuts against the downstream hook leg (34A).

8. The turbine ring assembly according to claim 7, characterized in that, The air diffuser (54) has an inclined plane (54A) to generate an impingement jet on the downstream hook leg (34A) in contact with the Ω-shaped seal (48).

9. The ring assembly according to any one of claims 1 to 8, characterized in that, The segmented cooling device (52) includes one or more ventilation channels (52A), each ventilation channel (52A) having a flared opening leading to the pressurization chamber (52B).

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

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