Combustion module for a turbomachine

Abstract

The invention relates to a combustion module (1) for a turbomachine (10), comprising: - an annular casing (3) extending around a longitudinal axis (X), - an annular combustion chamber (2) located inside the casing (3) and comprising a coaxial annular inner wall (4) and an annular outer wall (5) connected to one another by an annular bottom cavity (6), the walls (4, 5) and the bottom cavity (6) being constituted by at least two annular enclosures (50, 60) made of a ceramic matrix composite and comprising annular rims (52, 62) nested one inside the other and axially pressed against one another, the module (1) further comprising: - anti-disengagement means (8) configured to maintain the annular rims (52, 62) axially pressed against one another, these means (8) being carried by the combustion chamber (2) and / or the casing (3), - and a fuel injector (7) carried by the casing (3) and engaged in a bore (54) of one of the enclosures (50), said anti-disengagement means (8) comprising a gasket (86) mounted around the injector (7) and radially clamped against the enclosure, and comprising a lug (862) configured to cooperate with the other enclosure (60) by pressing against and / or hooking onto it.

Description

Technical Field

[0001] The present invention relates to a combustion module for a turbine, and more specifically, to the configuration and installation of the ceramic matrix composite (CMC) wall of the combustion chamber of the combustion module. Background Technology

[0002] Prior art includes, in particular, documents US-A1-2018 / 238232, JP-B2-3331826 and US-A1-2010 / 139283.

[0003] Generally speaking, in particular, the turbine of an aircraft includes a gas generator, which includes one or more compressors (e.g., a low-pressure compressor and a high-pressure compressor), and the one or more compressors are arranged upstream of the combustion module.

[0004] Reference Figure 1 Specifically, the combustion module 1 of the turbine 10 of the aircraft has a longitudinal axis X, which can coincide with the longitudinal axis of the turbine 10. Module 1 includes a combustion chamber 2, which has an annular shape. In the example, the combustion chamber 2 extends about the axis X and is surrounded by an annular shell 3, which also extends about the axis X.

[0005] Combustion chamber 2 is defined by a coaxial annular inner wall 4 and an annular outer wall 5, which are connected by a bottom chamber 6. The outer wall 5 is attached to a housing 3, which carries an annular row of fuel injectors 7 to supply fuel to combustion chamber 2. In this example, each injector 7 passes through an axis A perpendicular to axis X.

[0006] A portion of the compressed air from centrifugal compressor 90 enters combustion chamber 2 via an annular diffuser 92 and mixes with fuel supplied by injector 7. Another portion of this air circulates around the combustion chamber, such as... Figure 1 As indicated by the arrows in the diagram, combustion of the air / fuel mixture is initiated in combustion chamber 2 by an ignition device (not shown) and combustion is produced in the bottom chamber 6.

[0007] Figure 1 A recirculation combustion chamber 2 is shown, with a bottom chamber 6 located downstream of it. Furthermore, the annular inner wall 4 and annular outer wall 5 of the combustion chamber 2 extend radially toward the interior of the module 1 (relative to axis X) via bends 4a and 5a to supply combustion gases from the combustion chamber 2 to the turbine stator 94. The bends 4a and 5a include an inner bend 4a connected to the inner wall 4 and an outer bend 5a connected to the outer wall 5.

[0008] It should be noted that a direct-flow (i.e., normal-flow) combustion chamber includes an upstream bottom chamber and a downstream outlet that leads to the turbine stator of the turbine.

[0009] Figure 1 A recirculation combustion chamber is shown, in which the bottom chamber is located on the downstream side, and a bend is provided at the outlet of the combustion chamber to allow the combustion gases to be redirected into the turbine stator.

[0010] Within the framework of combustion module design and integration, the combustion chamber architecture can be divided into multiple parts to facilitate, in particular, the manufacture and / or operability of the combustion module. This leads to the challenge of assembling separately manufactured components in extremely hot and mechanically constrained environments.

[0011] Combustion chambers are typically made of metal, which allows individually manufactured components to be assembled via mechanical connections (such as bolts), welding, or brazing.

[0012] Combustion chambers can also be made of composite materials (such as CMC ceramic matrix composites), which include components having a connection known as a “hybrid” (i.e., a component made of a metallic material connected to a component made of a composite material, the component made of the metallic material and the component made of the composite material being able to withstand different expansions) and a connection between two components made of composite materials.

[0013] The use of composite materials is particularly advantageous in the turbine field because these materials are relatively lightweight and have better temperature resistance, which allows them to save cooling air or operate at higher temperatures.

[0014] In the case of combustion chambers made of CMC material, the assembly of components made of CMC material is typically carried out using bolt-type attachments. This bolt assembly is relatively cumbersome and complex to implement in the environment of the combustion module, which involves mechanical, thermomechanical, and chemical stresses. For example, it is necessary to ensure reliable tightening at all operating points and to install anti-rotation devices for the bolts. These difficulties in assembling CMC material components often affect the performance or integrity of the combustion chamber (and therefore the performance or integrity of the turbine).

[0015] In this context, it would be useful to at least partially overcome the aforementioned drawbacks by proposing a simpler solution that assembles at least two components made of CMC-type composite materials in the turbine's combustion module. Summary of the Invention

[0016] Therefore, the present invention proposes a combustion module for turbines, particularly for turbines in aircraft, the combustion module comprising:

[0017] - An annular housing that extends about the longitudinal axis X.

[0018] - An annular combustion chamber, located within the housing, includes a coaxial annular inner wall and an annular outer wall connected to each other via an annular bottom chamber. The walls and the bottom chamber are formed by at least two annular shells made of ceramic matrix composite material, and the at least two annular shells include annular edges, one of which is fitted together inside the other and axially abuts against each other.

[0019] According to the invention, the combustion module further includes anti-detachment devices configured to keep the annular edges axially abutting each other, and these devices are carried by the combustion chamber and / or housing.

[0020] The anti-detachment device according to the invention has several advantages. In particular, the anti-detachment device enables the CMC material casing (forming the annular wall and bottom chamber) to be positioned and assembled together in a simple and efficient manner, while preventing the casing from separating during operation.

[0021] During the hot operation phase of the combustion module (e.g., above 1000°C), a pressure difference exists between the outside and inside of the combustion chamber, which keeps the two assembled components (i.e., the CMC material enclosure) in contact.

[0022] In situations where the pressure difference is insufficient to maintain component contact, such as when the turbine is stationary, the anti-disengagement device prevents the connection from being broken.

[0023] This type of connection is small in size and requires little or no adjustment in the combustion module.

[0024] In general, using pressure differentials to maintain component contact, supplemented by anti-detachment devices, is an alternative solution to rigid connections (such as bolts) and non-removable connections (such as welding or brazing).

[0025] The combustion chamber according to the invention may include one or more of the following features, either individually or in combination:

[0026] - The module includes a first enclosure and a second enclosure, the first enclosure defining an inner wall and a bottom chamber, and the second enclosure defining an outer wall, the edges being located at the junction of the outer wall and the bottom chamber;

[0027] - The annular edge of the first shell includes a serrated shape;

[0028] - The module also includes a fuel injector carried by a housing and engaged in a hole in one of the housings, the anti-disengagement device being formed by a gasket mounted around the injector and radially fastened to the housing, and the gasket including a lug configured to engage with another housing in the housing by abutment and / or hooking.

[0029] - Each washer in the washers is inserted between the housing and the nut, which is used to fasten the washer to the housing;

[0030] - Each washer in the series has a coefficient of thermal expansion, which is sufficient to compensate for the expansion gap between the fastening nut and the CMC housing;

[0031] - The nut is tightened onto the body of the tubular sleeve, which is mounted around the injector. The sleeve includes an annular abutment collar on the circumferential edge of the orifice in which the injector engages.

[0032] - Each washer in the washers includes two lugs, preferably oriented in a generally parallel direction;

[0033] - Each lug has a shape complementary to the shape of the edge of the other enclosure, and / or each has a hook shape for engaging in an opening in the other enclosure to prevent relative displacement of the two elements along the axial orientation;

[0034] - The anti-detachment device is a stop member that may be formed protrudingly on the turbine housing and includes a free end configured to abut against the combustion chamber, particularly against the bottom chamber;

[0035] - Each of the components is in the form of an elongated arm that can be tilted relative to the engine axis, and the free end of the arm includes an abutment surface on the combustion chamber that extends in a plane substantially perpendicular to the surface of the bottom chamber;

[0036] - The number of abutting components is three to eight, and the abutting components are distributed circumferentially around the axis X;

[0037] - The module includes a gap between each of the abutting members (particularly the free end of the arm) and the combustion chamber, for example, the gap being on the order of millimeters;

[0038] -The abutment component is made of composite material, metal, or metal alloy material;

[0039] -The abutment component and the shell are a single piece;

[0040] -The combustion chamber is counter-current;

[0041] - The housing carries an annular row of fuel injectors, which are distributed at an angle around axis X;

[0042] - Each injector in the injector system extends along axis A, which is perpendicular to axis X.

[0043] - The first and second covers extend upstream through a bend that extends radially toward the interior of the module;

[0044] - Each of the abutting members is configured to be installed with a gap or with a preload;

[0045] - The free end of the abutment member is made of a material different from that of the abutment member to ensure, for example, chemical compatibility or thermal compatibility.

[0046] Advantageously, the anti-detachment device can be formed by both a washer and an abutment member, as described in at least one of the special features of the invention.

[0047] The present invention also relates to a turbine, particularly a turbine for an aircraft, the turbine comprising the combustion module as described above.

[0048] The present invention also relates to an aircraft comprising a fuselage and powered by at least one turbine, the at least one turbine comprising a combustion module as described above. Attached Figure Description

[0049] The invention will be better understood from the following description, which is by way of non-limiting example and with reference to the accompanying drawings, and other details, features and advantages of the invention will become clearer, in which:

[0050] [ Figure 1 ] Figure 1 It is a schematic half-view of the axial cross-section of a combustion module including a recirculation combustion chamber of a turbine according to the prior art;

[0051] [ Figure 2 ] Figure 2 This is a partial schematic half-view of an axial cross-section of a first embodiment of the combustion module assembly according to the present invention;

[0052] [ Figure 3 ] Figure 3 yes Figure 2 A schematic perspective view of the anti-detachment device in the middle;

[0053] [ Figure 4 ] Figure 4 It is along Figure 3 A schematic cross-sectional view of the plane CC;

[0054] [ Figure 5] Figure 5 It is along Figure 3 A schematic cross-sectional view of the plane DD;

[0055] [ Figure 6 ] Figure 6 This is a schematic perspective view of the anti-detachment device according to a second embodiment of the present invention;

[0056] [ Figure 7 ] Figure 7 This is a partial schematic diagram of the axial cross-section of the anti-detachment device according to a third embodiment of the present invention; and

[0057] [ Figure 8 ] Figure 8 This is a schematic half-view of the axial cross-section of a fourth embodiment of an anti-detachment device for a combustion module according to the present invention. Detailed Implementation

[0058] By convention, in the following description, the terms "longitudinal" and "axial" refer to the orientation of a structural element extending along a longitudinal axis (such as the longitudinal axis of a combustion module). The terms "radial" or "vertical" refer to the orientation of a structural element extending in a direction perpendicular to the longitudinal axis. The terms "inner" and "outer," as well as "internal" and "external," refer to positioning relative to the longitudinal axis. Thus, a structural element extending along a longitudinal axis includes an inner surface facing the longitudinal axis and an outer surface opposite to the inner surface of the structural element. The terms "upstream" and "downstream" are defined according to the orientation of gas flow within the turbine.

[0059] Figure 1 As described above, and shown is the combustion module 1 of the turbine 10 according to the background art.

[0060] Figures 2 to 8 Several embodiments of the combustion module 1 according to the present invention are shown.

[0061] In the following description, the invention is generally applied to turbine 10, and particularly to turbines of aircraft, such as turbojet engines or turboprop engines.

[0062] Turbine 10 typically includes: a compressor module including at least one compressor; a turbine module including at least one turbine; and a combustion module 1 inserted between the compressor module and the turbine module.

[0063] As described above, the combustion module 1 includes an annular housing 3 that extends about a longitudinal axis X and surrounds an annular combustion chamber 2. This axis X may coincide with the longitudinal axis of the turbine 10 (such as the axis of rotation of the rotor). The combustion chamber 2 and the housing 3 extend about the axis X. The combustion chamber 2 may extend parallel to or at an angle relative to the axis X.

[0064] Figure 2 Combustion chamber 2 is schematically shown, comprising an annular inner wall 4 and an annular outer wall 5. These walls 4 and 5 are coaxial and connected via an annular bottom chamber 6. Combustion chamber 2 is counter-flow type. Therefore, the bottom chamber 6 is arranged downstream of the turbine, with combustion chamber 2 opening towards the turbine module and then upstream of the turbine. Alternatively, when the combustion chamber is direct-flow type, the bottom chamber 6 is arranged upstream of the turbine. The bottom chamber 6 may include a transverse annular wall that passes approximately through a first plane P1 perpendicular to the axis X.

[0065] An outer wall 5 is attached to a housing 3, which carries an annular row of fuel injectors 7, the injectors being angled about an axis X to supply fuel to the combustion chamber 2. Specifically, the outer wall 5 includes annular holes 54 extending about the axis X. Each hole 54 has an inner diameter D. 54 Each of the holes 54 includes a circumferential edge 56. Each of the holes 54 is adapted to receive a fuel injector 7. Figure 2 In the example shown, the injector 7 passes through axis A, which is approximately perpendicular to axis X. Alternatively, the orifice 54 and the injector 7 may be tilted or parallel to axis X.

[0066] In this invention, the combustion chamber 2 is made of CMC ceramic matrix composite material. The inner wall 4, the outer wall 5, and the bottom chamber 6 are formed by at least two annular shells 50 and 60 made of CMC ceramic matrix composite material.

[0067] exist Figure 2 In the example shown, the annular inner wall 4 and the bottom chamber 6 are formed as a single integral piece by a first cover 60. In the example, the first cover 60 has a generally inverted "S" shape, in which one of the two rings generally corresponds to the bottom chamber, while the other ring generally corresponds to the inner wall 4 and the inner bend 4a of the combustion chamber. The annular outer wall 5 is formed as a single integral piece by a second cover 50. In the example, the second cover 50 has a generally "C" shape, in which the upper portion generally corresponds to the outer wall 5, while the lower portion generally corresponds to the outer bend 5a.

[0068] In the example, the first casing 50 and the second casing 60 are connected to each other on the downstream side. These casings 50 and 60 continue to extend on the upstream side with bends 4a and 5a to lead to the distributor 94 of the turbine module, the bends extending radially toward the interior of the module 1 (relative to the axis X).

[0069] The first casing 50 and the second casing 60 each include an annular edge 62 referred to as the inner annular edge and an annular edge 52 referred to as the outer annular edge. The annular edges 52 and 62 are joined together, one inside the other, particularly at the junction between the outer wall 5 and the bottom chamber 6. Specifically, the inner edge 62 abuts radially (or substantially radially) against the outer edge 52. "Radial abutment" refers to the abutting force exerted by the inner edge 62 on the cylindrical surface of the outer edge 52 in a transverse plane (relative to the axis X). "Substantially radial abutment" refers to the abutting force exerted by the inner edge 62 along an inclined plane (relative to the axis X), particularly when the combustion chamber 2 is inclined relative to the axis X, on the truncated conical surface of the outer edge 52.

[0070] Furthermore, the edge 62 of the first casing 60 (corresponding to the bottom chamber 6 in this example) particularly near the injector 7 may include a serrated shape (not shown in the figures). This allows the combustion chamber 2, in particular, to be axially compacted. The serrated shape may be made of a series of protruding or recessed arcuate segments (such as corrugations).

[0071] One of the special features of this invention is that the combustion module 1 includes an anti-detachment device 8, which is supported by the combustion chamber 2 (e.g., Figures 2 to 7 (as shown), or carried by the housing 3 of module 1 (as shown). Figure 8 (As shown). Each device in device 8 is used to keep the annular edges 52, 62 axially (or substantially axially) abutting each other, or to maintain a limited gap between the annular edges. In this way, the casings 50, 60 of the combustion chamber 2 are prevented from separating during operation and when at rest (thus preventing connection with the outer wall 5 and the bottom chamber 6), while allowing for simple and efficient assembly and disassembly of device 8 within the available overall dimensions of the combustion module 1.

[0072] "Axial contact" refers to the contact or abrasion exerted by the outer edge 52 along the axis X on the cylindrical surface of the inner edge 62. "Approximately axial contact" refers to the contact or abrasion exerted by the outer edge 52 along an inclined plane (relative to the axis X), particularly when the combustion chamber 2 is inclined relative to the axis X, on the truncated conical surface of the inner edge 62; or conversely, the contact exerted by the inner edge 62 along an inclined plane (relative to the axis X) on the truncated conical surface of the outer edge 52.

[0073] According to a first embodiment of the present invention, such as Figures 2 to 5As shown, the anti-detachment device 8 can be mounted on the combustion chamber 2. In particular, in these example figures, the device 8 is mounted around the fuel injector 7.

[0074] Each of the devices 8 has a washer 86, which is configured to surround the injector 7.

[0075] Reference Figure 3 Therefore, washer 86 has a generally annular and flat shape. Washer 86 includes a first central opening 860 and a lug 862. Opening 860 has an inner diameter D. 860 .

[0076] The gasket 86 may include one or more lugs 862. The number and size (shape, length, thickness, etc.) of the lugs 862 of each gasket 86 may vary depending on the size and material of the components constituting the combustion module 1. Figure 3 In the example shown, the washer 86 includes two lugs 862, which are oriented parallel to each other. Each of these lugs 862 has a hook shape.

[0077] The thickness of each washer (86) determines the degree to which the washer helps maintain the fastening force, especially during the turbine's flight phase.

[0078] Gasket 86 may be made of CMC type composite material or metal alloy. Preferably, gasket 86 is made of stainless steel (e.g., A286 type). The advantage of stainless steel A286 is that it is compatible with the thermal environment of combustion chamber 2 and has a high coefficient of thermal expansion to optimally maintain the connection between components under the extreme operating temperatures of combustion module 1.

[0079] The washer 86 is configured to be radially (i.e., substantially perpendicular to axis X) against the second cover 50 of the outer wall 5 by means of the fastening nut 84 and the sleeve 82.

[0080] Reference Figure 5 The sleeve 82 has a tubular shape and therefore includes a second central opening 820 and a tube body 822. The second opening 820 has an inner diameter D. 820 The tube body 822 may include external threads.

[0081] The sleeve 82 also includes an annular collar 824. The collar 824 and the tube body 822 can be defined by a tubular portion 826. Figure 5 In the example shown, the annular collar 824 includes a first side portion 824a with a flat shape and a second side portion 824b with a truncated conical shape. The second side portion 824b tapers gradually toward the tubular portion 826.

[0082] During hot operation, the truncated conical shape of the second side 824b of the annular collar 824 of the sleeve 82 can expand. This may result in displacement relative to the housing 50. This truncated conical shape of the second side 824b can return to radial displacement relative to axial displacement, thereby securing the connection between the sleeve 82, nut 84, and housing 50. The washer 86 may be made of a material with a coefficient of thermal expansion that allows for compensation of expansion gaps between the housing 50 and nut 84.

[0083] In addition, Figure 5 In the example shown, collar 824 has an outer diameter D 824 The outer diameter is larger than the outer diameter D of the tubular portion 826 and the tube body 822. 826 D 822 The outer diameter D of the tubular portion 826 826 The outer diameter D of the tube body is greater than 822. 822 .

[0084] Now refer to Figures 3 to 5 The assembly of the gasket 86 of the first embodiment on the combustion chamber 2 is described, particularly around the injector 7 and on the edges 52, 62 of the casings 50, 60.

[0085] For this purpose, the sleeve 82 is specifically installed around the injector 7 via a second opening 820. Therefore, this second opening 820 extends around the axis A of the injector 7. Therefore, in Figure 5 In the example shown, the inner diameter D of the second opening 820 of the sleeve 820 It is roughly the same as the outer diameter D7 of the injector 7.

[0086] Next, the annular edge 52 of the second casing 50 includes a hole 54 into which the injector 7 is intended to engage. The edge 52 is specifically mounted around the collar 824 of the sleeve 82 via the hole 54. This causes the circumferential edge 56 of the hole 54 to abut against the second side 824b of the collar 824. Figure 5 In the example shown, the circumferential edge 56 abuts against the second side portion 824b at a position inclined relative to axes X and A. Furthermore, the outer diameter D of the tubular portion 826 of the sleeve 82... 826 The inner diameter D of the hole 54 of the second cover 50 54 They are roughly the same.

[0087] Then, the washer 86 is specifically installed around the sleeve 82 through a first opening 860. This first opening 860 also extends around axis A. Figure 5 In the example shown, the inner diameter D of the first opening 860 of the washer is... 860 Therefore, the outer diameter D of the tubular portion 826 of the sleeve... 826 And the inner diameter D of the hole 54 in the second cover 5054 Broadly the same. Furthermore, in this example, the annular edge 62 of the first cover 60 includes an opening 64 that complements the shape of the lug 862 of the washer 86. This allows the hook-shaped lug 862 to engage with the opening 64 of the first cover 60, as... Figure 4 As shown.

[0088] Finally, nut 84 is tightened specifically around the body 822 of sleeve 82 through the third opening 840 of nut 84. This ensures that the edge 52 of washer 86 and second seal 50 remain on sleeve 82. Figure 5 In the example shown, the inner diameter D of the third opening 840 840 The outer diameter D of the tube body 822 of the sleeve 82 822 The same applies. The third opening 840 may include an internal thread that is complementary to the external thread of the body 822 of the sleeve 82.

[0089] Figure 6 An anti-detachment device and its components are shown in an anti-detachment device 8 according to a second embodiment. In the second embodiment, the anti-detachment device 8 may also be installed on the combustion chamber 2.

[0090] The difference between the anti-detachment device 8 of the second embodiment and the device 8 of the first embodiment lies in the lug 862 of the washer 86 and the edge 62 of the first cover 60.

[0091] Reference Figure 6 The lug 862 of the washer 86 has an elongated and curved shape. The edge 62 of the first cover 60 includes a boss 66 that passes substantially through the second plane P2. The plane P2 is substantially perpendicular to the axis X of the module 1 and is located upstream of the first plane P1 of the bottom chamber 6. Alternatively, the edge 62 may include a margin of the cover 60 at the plane P2.

[0092] exist Figure 6 In this configuration, lug 862 and boss 66 have complementary shapes. This allows lug 862 to be attached to boss 66 by direct abutment, such that lug 862, radially fastened to edge 52, abuts against boss 66 on edge 62. Therefore, this configuration also allows the edges 52 and 62 of the covers 50 and 60 to remain axially abutting against each other.

[0093] Preferably, the lug 862 is elongated, such that the free and curved end of the lug faces a plane P1 that roughly corresponds to the transverse wall of the bottom chamber 6. In particular, this allows for compensation of axial displacement of the components at the edges 52, 62 during operation of the combustion module 1.

[0094] Furthermore, the elongated shape of lug 862 allows for flexibility in the connection between lug 862 and boss 66. In this way, the assembly between the shells 50 and 60 (of the walls 5, 6 and the bottom chamber 6) is rigid, especially with permanent contact between lug 862 and boss 66, and with very small or no assembly gap between the edges 52 and 62 of the shells 50 and 60.

[0095] This second embodiment has the particular advantage of avoiding complex processing on the edge 62 of the first cover 60.

[0096] Figure 7 An anti-detachment device and its components are shown in an anti-detachment device 8 according to a third embodiment, wherein these devices 8 may also be installed on the combustion chamber 2.

[0097] The difference between the anti-detachment device 8 of the third embodiment and the device 8 of the first embodiment lies in the lug 862 of the washer 86 and the edge 62 of the first cover 60.

[0098] Reference Figure 7 The edge 62 of the first cover 60 includes at least one shoulder 68 to form an axial abutment surface for one of the lugs 862 for the washer 86. Preferably, the shoulder 68 is formed on the allowance of the cover 60 located approximately in plane P2. The lugs 862 and the shoulder 68 may have complementary shapes.

[0099] exist Figure 7 In this configuration, lug 862 includes a curved free end 864 that engages in the shoulder 68 of edge 62. In this way, lug 862 is also attached to the shoulder 68 by direct axial abutment. Therefore, this configuration also allows edges 52 and 62 to remain together.

[0100] This third embodiment is a simple alternative solution implemented to prevent the separation of the connection of the covers 50 and 60.

[0101] According to the fourth embodiment of the present invention, such as Figure 8 As shown, the anti-detachment device 8 can be carried by the housing 3. Specifically, in Figure 8 In this example, device 8 and housing 3 are a single piece (i.e., made of one material).

[0102] Advantageously, these devices 8 are abutment members that can be protruded onto the housing 3. These abutment members may include a free end 880 adapted to abut against the combustion chamber 2, particularly against the bottom chamber 6. This configuration makes it possible to prevent any displacement of the bottom chamber 6 relative to the outer wall 5.

[0103] These abutment components can be made of rigid or flexible materials. For example, abutment components can be made of composite materials, metals, or metal alloys.

[0104] Preferably, the abutment member can be made of the same material as the housing 3, but has a reduced first thickness compared to the second thickness of the housing 3. In particular, this allows the abutment member to be flexible while still being able to apply sufficient pressure to the combustion chamber 2 to keep the edges 52, 62 axially abutting each other.

[0105] When the abutment member is made of a flexible material, the free end 880 can make direct contact with the bottom chamber 6. When the abutment member is made of a rigid or flexible material, an installation clearance and / or a clearance to compensate for expansion during operation can be added between the abutment member and the bottom chamber. These installation clearances and expansion clearances can vary depending on the material or size (such as thickness) used for the enclosure and the abutment member. For example, the installation clearance and / or expansion compensation clearance can be on the order of millimeters.

[0106] The abutment component and the shell 3 can be made as a single piece (i.e., made of one material).

[0107] exist Figure 8 In this configuration, the abutment member is in the form of an arm 88. The arm 88 has an elongated shape along the direction of the bottom chamber 6. The end 880 of the arm 88 is tiltable relative to the axis X. This end 880 includes an abutment surface 882 on the bottom chamber 6. Figure 8 In the example shown, the abutting surface 882 passes approximately through the first plane P1.

[0108] The housing 3 may include three to eight arms 88 distributed circumferentially around axis X, such that airflow from the diffuser 92 of the compressor module can flow into module 1.

[0109] Furthermore, the free end 880 of arm 88 may include a heat-protective coating. This allows for a reduction in temperature, particularly of arm 88. The coating can be combined with a material different from the material used to manufacture the abutment member. This allows for chemical or thermal compatibility between the abutment member and the combustion chamber.

[0110] The fourth embodiment of device 8 has a particular advantage in that it facilitates the static indeterminateness of the assembled edges 52, 62 (even if the edges are immovable), and thus limits movement and / or deformation.

[0111] According to another embodiment (not shown in the accompanying drawings), Figure 8 The abutment member of the fourth embodiment can be with Figures 2 to 7 The washer is engaged in at least one embodiment of the embodiments. In fact, the device 8 can be formed by both the washer 86 and the abutment member.

[0112] The anti-detachment device equipped on the combustion module of the present invention is advantageous, particularly for the following reasons:

[0113] - The anti-detachment device strengthens the assembly between components made of CMC-type composite materials in the combustion module.

[0114] - The anti-detachment device eliminates the separation of components assembled in the combustion module during operation.

[0115] - A method for detachably assembling and removing the anti-detachment device on the combustion chamber was proposed.

[0116] - An alternative using bolt-type attachments was proposed.

[0117] - The anti-detachment device limits the maintenance costs of the combustion chamber, and

[0118] - The anti-detachment device is easily adapted to existing turbines.

[0119] In summary, the proposed solution is simple, effective, and economical to implement and assemble on turbines and aircraft, while ensuring the safe assembly and disassembly of the combustion chamber (made of composite material components) in the turbine.

Claims

1. A combustion module (1) for a turbomachine, comprising: - an annular casing (3) extending around a longitudinal axis (X), - an annular combustion chamber (2) located inside the annular casing (3) and comprising coaxial annular walls, an inner wall (4) and an outer wall (5) connected to each other by an annular bottom cavity (6), the inner wall (4), the outer wall (5) and the annular bottom cavity (6) being formed by at least two annular enclosures made of ceramic matrix composite and comprising annular rims (52, 62) assembled one inside the other and axially abutted against each other, characterized in that the combustion module (1) further comprises anti- disengagement means (8) configured to maintain the annular rims (52, 62) axially abutted against each other, the anti-disengagement means (8) being carried by the annular combustion chamber (2) and / or by the annular casing (3), the at least two annular enclosures comprising a first enclosure (60) and a second enclosure (50), the combustion module (1) further comprising a fuel injector (7) carried by the annular casing (3) and engaged in a bore (54) of the second enclosure (50), the anti-disengagement means (8) comprising a gasket (86) mounted around the fuel injector (7) and radially fastened on the second enclosure (50), and comprising lugs (862) configured to cooperate by abutting and / or hooking with the first enclosure (60).

2. The combustion module of claim 1, wherein, the first enclosure (60) defining the inner wall (4) and the annular bottom cavity (6), the second enclosure (50) defining the outer wall (5), the annular rims (52, 62) being located at the connection of the outer wall (5) with the annular bottom cavity (6).

3. The combustion module of claim 1 or 2, wherein, the annular rim of the first enclosure (60) comprises a scalloped shape.

4. The combustion module of claim 1 or 2, wherein, each of the gaskets (86) is interposed between the second enclosure (50) and a nut (84) for fastening the gasket on the second enclosure (50).

5. The combustion module of claim 4, wherein, the nut (84) is screwed onto a tube body (822) of a tubular sleeve (82) mounted around the fuel injector (7), the tubular sleeve (82) comprising an annular abutment collar (824) on a circumferential edge (56) of the bore (54) in which the fuel injector (7) is engaged.

6. The combustion module of claim 1 or 2, wherein, each of the gaskets (86) comprises two lugs (862).

7. The combustion module of claim 1 or 2, wherein, the lugs (862) each have a shape complementary to the shape of the annular rim of the first enclosure (60) and / or each have a hook shape for engagement in an opening (64) in the first enclosure (60).

8. The combustion module of claim 1 or 2, wherein, the gaskets (86) are made of stainless steel.

9. The combustion module of claim 1 or 2, wherein, the annular combustion chamber (2) is of the reverse-flow type.

10. The combustion module of claim 1, wherein, the turbomachine is a turbomachine for an aircraft.

11. The combustion module of claim 6, wherein, The two lugs are oriented in a substantially parallel direction.

12. The combustion module of claim 8, wherein, The stainless steel is of type A286.

13. A turbomachine (10) comprising a combustion module (1) according to any one of claims 1 to 12.

14. The turbomachinery (10) of claim 13, characterized in that, The turbomachine is a turbomachine for an aircraft.

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