Turbine shroud ring assembly with improved load pathway
Patent Information
- Authority / Receiving Office
- EP · EP
- Patent Type
- Applications
- Current Assignee / Owner
- Filing Date
- 2024-05-30
- Publication Date
- 2026-04-08
AI Technical Summary
Current gas turbine designs face challenges in cooling turbine components due to the inefficiency of using fresh air bypassing the combustion chamber, leading to increased temperature demands and the need for ceramic matrix composite materials, which require precise positioning and assembly to avoid loading and leaking, resulting in complex and costly assembly processes and potential leaks.
A turbine ring assembly featuring a ceramic matrix composite ring supported by a metal alloy casing with a diffuser and spacer sectors, utilizing radial bolted connections and ventilation holes for cooling air supply, simplifying assembly and reducing mass and leaks by eliminating separate flange and spacer parts and axial bolted connections.
This design reduces the number of parts, simplifies assembly, minimizes mass and cost, and enhances sealing, achieving better aerodynamic control and robustness by eliminating multiaxial stress and complex machining requirements.
Smart Images
Figure FR2024050692_05122024_PF_FP_ABST
Abstract
Description
[0001] Description
[0002] Title of the invention: Turbine ring assembly with improved force path
[0003] Technical Field
[0004] The present invention relates to the field of gas turbines for aircraft (jet and propulsion engines, helicopter engines), gas turbines for electric generators and gas turbines for turbochargers.
[0005] Prior art
[0006] 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 to both new aircraft types and those currently in operation, requiring the implementation of technological solutions to comply with current regulations. Civil aviation has been mobilizing for several years now to contribute to the fight against climate change.
[0007] 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.
[0008] 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 the minimum possible in order to reduce the environmental footprint of its activity. This sustained research and development work covers 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.
[0009] The current trend in civil aeronautics is towards an increase in the temperature of the gases leaving the combustion chamber, leading to a rise in the temperature of the turbine components and therefore to an increasingly important need to cool these parts. However, the fresh air which allows this cooling is taken from the outlet of the high-pressure compressor and will therefore bypass the combustion chamber without contributing to the combustion, and therefore to the efficiency of the turbomachine.
[0010] This is why the use of ceramic matrix composite (CMC) materials on the hottest turbine parts currently reduces the need for cooling them because these CMC materials perform well at high temperatures (e.g. 1700°C). They also allow for weight savings because they are lighter than the metal alloys traditionally used for these hottest parts due to their low density (three times lower than metal bases).
[0011] However, the integration of a CMC ring into a set of metal parts subjected to high temperature requires the design of a force path that does not load the CMC elements, which are more fragile than the metal elements and which expand differently from the other parts, while maintaining them deterministically in position so as to ensure the performance of the turbomachine.
[0012] The state of the art typically consists of a set of CMC ring sectors in the shape of the inverted letter PI. Each ring has an annular base that forms an aerodynamic vein and is held in position by two flanges that extend radially on either side of the annular base. These two flanges are held to the rest of the stator by means of four pins, two upstream connecting the flange to a flange, and two downstream connecting the flange to the spacer. A diffuser bringing cooling air is positioned so as to cool the radially external face of the annular flange, and is assembled with the flange and the spacer by means of axial bolted connections (two fasteners per sector) that are sensitive to vibrations and that cause leak openings that reduce the performance of the turbomachine.
[0013] To route the cooling air to the annular base, the flange is drilled and the diffuser contains a recess / air duct. The drilling of the flange and the recess of the diffuser must be positioned opposite each other to allow the air to pass through while minimizing pressure losses, which requires particularly precise centering of the diffuser on the flange. Another precision centering is necessary to mount the spacer with the diffuser and the flange so that the three parts can be assembled by the bolted connection. The centering of the diffuser is achieved by means of a circular support, which implies that the diffuser is in the force path from the distributor to the housing.
[0014] With this technology, the screws carry additional mass (mass of the screws + additional mass in the spacer for tapping) and are in the force path with multiaxial stresses (tension + shear). In addition, the presence of two separate parts (flange + spacer) increases the risk of cooling air leaks between the different parts.
[0015] The positioning of the CMC ring is also complex because for the vein to be straight the axes of the upstream and downstream pins must be parallel, this therefore requires tight machining tolerance and additional controls.
[0016] Statement of the invention
[0017] 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. To this end, the main purpose of the present invention is therefore to reduce the number of parts in contact and to simplify their assembly, to reduce cooling air leaks, to reduce mass and cost and to improve the force path.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 circumferentially end to end around the axis of rotation, a diffuser comprising a plurality of diffuser sectors arranged circumferentially end to end around the axis of rotation, and a metal alloy casing forming a ring support structure, in which the casing comprises a central shell which extends around the turbine ring and from which extend radially inwards an upstream annular flange and a downstream annular flange between which is retained a metal alloy spacer comprising a plurality of spacer sectors, characterized in that each of the ring sectors is supported by a diffuser sector fixed by a radial bolted connection to a spacer sector.
[0018] Thus, the attachment of the turbine ring to the diffuser alone, combined with the single radial fixing of the assembly to a single-piece spacer, allows, in addition to reducing mass and cost, to reduce sealing leaks while simplifying the assembly's mounting on the turbine casing.
[0019] Preferably, each spacer sector comprises at least one through ventilation hole intended to be in fluid communication with at least one corresponding ventilation hole made in the diffuser sector so as to allow, through these ventilation holes, a supply of cooling air into a plenum of a diffuser sector.
[0020] Advantageously, each spacer sector comprises a flat portion arranged opposite a flat portion of a diffuser sector so as to allow centering of the radial bolted connection.
[0021] Preferably, each diffuser sector is fixed to a ring sector by four axial fixing pins, including two upstream fixing pins each passing through a respective ear located at the opposite circumferential ends of an upstream tab of the ring sector and two downstream fixing pins each passing through a respective ear located at one of the two opposite circumferential ends of a downstream tab of the ring sector. Advantageously, the ring assembly further comprises a seal configured to ensure a seal between the upstream annular flange of the casing and each spacer sector.
[0022] Preferably, each of the ventilation holes of a spacer sector is divergent and comprises a dust filter mounted at the inlet of the ventilation hole on an upstream radial end wall of each spacer sector.
[0023] Advantageously, each spacer sector comprises an upstream hook and a downstream hook, each oriented upstream and mounted in an axial groove of the upstream and downstream annular flange of the casing.
[0024] Preferably, the turbine ring comprises a seal mounted in an axial groove of the spacer sector so as to ensure a seal between the downstream leg of the ring sector and the downstream wall of the spacer sector.
[0025] Advantageously, the ring sectors have two-by-two grooves arranged circumferentially opposite each other, and inter-sector tongues being housed in these grooves so as to extend between these grooves.
[0026] The invention also relates to a turbomachine comprising a turbine ring assembly as mentioned above.
[0027] Brief description of the drawings
[0028] 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:
[0029] [Fig. 1] Figure 1 is a general view of a turbomachine,
[0030] [Fig. 2] Figure 2 is a perspective view of a turbine ring sector and its diffuser,
[0031] [Fig. 3] Figure 3 shows in cross-section a turbine ring assembly according to the invention,
[0032] [Fig. 4] Figure 4 is a perspective view of a spacer sector of the turbine ring of Figure 3, and [Fig. 5] Figure 5 is a top view of the diffuser of Figure 2.
[0033] Description of the embodiments
[0034] 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 perpendicular direction depending on whether the object concerned is in contact or not with this gas flow.
[0035] 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 arranged circumferentially over 360°, so as to form a sectored turbine ring (discontinuous cylindrical with inter-sector clearances) surrounding the sets of rotating blades of the turbomachine.
[0036] 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 lug 32 and a downstream lug 34 extend radially outward. These two lugs which extend in the circumferential direction over the entire width of the ring angular sector 28 are axially spaced from each other and for example each have two ears 32A, 32B; 34A, 34B each pierced with an axial orifice intended to receive an axial fixing pin (only the downstream axial fixing pins 35A and 35B are illustrated). As illustrated, the upstream and downstream pawns are advantageously four in number (two per leg) per ring sector.The holes in the pins are advantageously made by machining in a single pass so as to facilitate obtaining coaxiality of the upstream and downstream pins.
[0037] In a known manner, the seal between angular ring sectors is ensured by inter-sector tabs (not shown) housed in grooves 37 of the angular ring sectors which have two-by-two grooves arranged opposite each other circumferentially.
[0038] As illustrated in Figure 3, the angular ring sectors 28 are supported by the casing 22 successively through spacer sectors 36 and diffuser sectors (sectored diffusers 38). The spacer sectors and diffuser sectors are assembled end to end circumferentially. More specifically, the casing 22 comprises a central shroud which extends around the turbine ring and from which extend radially inward an upstream annular flange 42 provided with an upstream annular groove 42A which extends axially and opens upstream and a downstream annular flange 44 provided with a downstream annular groove 44A which extends axially and also opens upstream. These upstream and downstream annular grooves are intended to receive upstream 46 and downstream 48 hooks oriented upstream in the form of a ring sector of the spacer sector 36.
[0039] The spacer sector, advantageously made of metal alloy and which is illustrated in perspective in Figure 4, is in the form of a single-piece part of substantially H shape with oblique bar (or inverted N), with an upstream radial wall 36A (forming the first arm of the H) whose external end projects downstream to form the upstream hook 46, a downstream radial wall 36B (forming the second arm of the H) whose external end projects downstream to form the downstream hook 48 and a central wall 36C (forming the oblique bar of the H) connecting the upstream and downstream radial walls.The inner end of the upstream radial wall 36A comprises a support 50 extending axially downstream to support the upstream lug 32 of the ring sector 28 and the inner end of the downstream radial wall 36B comprises an axial groove 52 extending axially and opening upstream and intended to receive a flexible support 54, typically machined from a metal alloy having an S or W shape, to support the downstream lug 34 of the ring sector 28 and thus make it possible to compensate for the different axial expansions of the turbine ring and the metal parts. The sealing downstream and upstream of the ring is controlled by a seal (typically an Omega seal 56) (see figure 3) mounted between the upstream annular flange 42 and the upstream radial wall 36A of the spacer sector.
[0040] On the upstream wall 36A of the spacer sector, at the inlet of at least one ventilation hole 64A, 64B advantageously diverging and ensuring the circulation of the cooling air through the spacer sector 36, a dust filter 57 is mounted,
[0041] Let us return to Figure 3 where the sectored diffuser 38 comprises on its two side walls several blind cavities intended to receive the axial fixing pins advantageously coaxial, without this being an obligation, however, these holes may not be coaxial for integration needs. The sectored diffuser 38 also conventionally comprises an internal plenum chamber 58 closed by a radially internal diffusion plate 60 pierced with a plurality of air ejection orifices. On an opposite, radially external wall, a flat 38A is machined intended to cooperate with a corresponding flat 36D of the spacer sector 36, the spacer sector and the sectored diffuser being pierced with coaxial orifices 66 to receive a radial bolted connection 68, typically one assembly screw per spacer sector.Thus, with the invention, the conventional axial bolted connection with two screws is replaced by a radial bolted connection with a single screw which connects the ring-diffuser assembly to the spacer. The spacer is thus improved to absorb the forces of the high-pressure distributor (HPD) and transmit them to the casing without loading the turbine ring, the bolted connection working only in traction and being outside the flow of forces from the distributor which now passes through the oblique bar of the H. The fixing is thus less subject to vibrations. This configuration also makes it possible to reduce the number of parts, therefore limiting the cost and leaks as well as saving weight.
[0042] Furthermore, with the invention, the drillings of the pin housings on the spacer made in the prior art are eliminated and by holding the pins and the ring by the diffuser, better control of the aerodynamic vein is ensured because the drillings of the pins are made by machining in a single pass, and therefore coaxiality is easier to obtain.
[0043] Figure 5 illustrates, with the preceding figures 3 and 4, the passage of the cooling air through the sectored diffuser. This fluid introduced by the dust filter 57 mounted on the upstream wall 36A of the spacer sector and passing through the ventilation holes 64A, 64B drilled in this spacer sector 36 is then routed through at least one advantageously divergent ventilation hole (for example 62A, 62B) corresponding to the sectored diffuser 38, to reach the plenum chamber 58 before impacting the internal face 30A of the annular base through the diffusion plate 60.
[0044] It can be noted that the invention allows for simplified assembly because it now only requires a single centering, which is that of the diffuser with the spacer, carried out using only two flat supports Pl, P2. Assembly is thus carried out as follows:
[0045] (1): Positioning of the ring 28 on the diffuser 38,
[0046] (2): Fixing of the assembly thus obtained by means of axial pins 35A, 35B,
[0047] (3): Positioning of the assembly thus fixed on the spacer 36 and fixing by the radial assembly screw 68, and
[0048] (4): Attachment of the assembly to the annular flanges of the casing 42, 44.
[0049] With the invention, the integration of a CMC ring in a modern engine is improved. A weight saving is obtained by eliminating the fixing screws between the parts, a reduction in thicknesses by more direct force paths, and better robustness by an absence of bolted connections under multiaxial stress. Machining and assembly are simplified and the leaks existing between the parts in the prior art are eliminated.
Claims
Claims
1. A gas turbine ring assembly, comprising, around a longitudinal axis of rotation, a turbine ring comprising a plurality of ring sectors (28) of ceramic matrix composite material arranged circumferentially end to end around the axis of rotation, a diffuser comprising a plurality of diffuser sectors (38) arranged circumferentially end to end around the axis of rotation, and a casing (22) of metal alloy forming a ring support structure, in which the casing (22) comprises a central shroud which extends around the turbine ring (28) and from which extend radially inward an upstream annular flange (42) and a downstream annular flange (44) between which is retained a spacer of metal alloy comprising a plurality of spacer sectors (36),characterized in that each of the ring sectors (28) is supported by a diffuser sector (38) fixed by a radial bolted connection (68) to a spacer sector (36), characterized in that each diffuser sector (38) is fixed to a ring sector (28) by four axial fixing pins (35A, 35B), including two upstream fixing pins each passing respectively through an ear (32A, 32B) located at one of the two opposite circumferential ends of an upstream tab (32) of the ring sector (28) and two downstream fixing pins each passing respectively through an ear (34A, 34B) located at one of the two opposite circumferential ends of a downstream tab (34) of the ring sector (28).,
2. A turbine ring assembly according to claim 1, wherein each spacer sector (36) comprises at least one through ventilation hole (64A, 64B) intended to be in fluid communication with at least one corresponding ventilation hole (62A, 62B) made in the diffuser sector (38) so as to allow, through these ventilation holes, a supply of cooling air into a plenum (58) of a diffuser sector (38).
3. A turbine ring assembly according to claim 1 or claim 2, wherein each spacer sector (36) comprises a flat (36D) arranged opposite a flat (38A) of a diffuser sector (38) so as to allow centering of the radial bolted connection (68).
4. A turbine ring assembly according to any one of claims 1 to 3, comprising a seal (56) configured to provide a seal between the upstream annular flange (42) of the casing (22) and each spacer sector (38).
5. A turbine ring assembly according to any one of claims 1 to 4, wherein each of the ventilation holes (64A) of a spacer sector (36) is divergent and comprises a dust filter (57) mounted at the inlet of the ventilation hole (64A) on an upstream end radial wall (36A) of each spacer sector (36).
6. A turbine ring assembly according to any one of claims 1 to 5, wherein each spacer sector (36) comprises an upstream hook (46) oriented upstream and mounted in an axial groove (42A) of the upstream annular flange (42) of the casing (22) and a downstream hook (48) oriented upstream and mounted in an axial groove (44A) of the downstream annular flange (44) of the casing (22).
7. Turbine ring assembly according to claim 6, comprising a seal (54) mounted in an axial groove (52) of the spacer sector (36) so as to ensure a seal between the downstream leg of the ring sector (28) and the downstream wall (36B) of the spacer sector (36).
8. A turbine ring assembly according to any one of claims 1 to 7, wherein the ring sectors (28) have two-by-two grooves (37) arranged circumferentially opposite each other, inter-sector tabs being housed in these grooves (37) so as to extend between these grooves (37).
9. An aeronautical turbomachine comprising a turbine ring assembly according to any one of claims 1 to 8.