GUIDE BOWL ARRANGEMENT FOR AN AIRCRAFT TURBINE ENGINE COMPRESSOR

Description

technical field

[0001] The present invention relates to an aircraft turbomachine compressor. Previous art

[0002] It is known, for example from document EP2799721B1, to fix the stator blades of a rectifier assembly of an aircraft turbomachine compressor to a ferrule located radially on the outside, called the outer ferrule. This document also describes auxiliary blades, which are elements located between the stator blades, and having a radial height of between 10% and 50% of the radial height of the stator blades.

[0003] US document 3,778,184 A describes a compressor in which blade damping is achieved by surrounding one end of the blade with a damping material such as steel wool or metallic felt held in contact with the shroud.

[0004] Document EP 2 093 383 A1 describes a compressor in which the stator blades are fixed to the inner shell. Document FR2 950 116 A1 describes a compressor rectifier stage in which each blade head is mounted in an opening in the outer shell by means of fastening means. Summary of the invention

[0005] The outer shell is subjected to significant mechanical stresses, particularly in turbomachinery architectures where it lies in the main thrust path. Part of these mechanical stresses originates from the stator blades, which are attached to this outer shell.

[0006] One object of the present invention is to reduce mechanical stresses in an aircraft turbomachine.

[0007] To this end, the invention proposes a rectifier assembly for an aircraft turbomachine compressor, comprising: an inner ferrule, an outer ferrule, and stator blades, in which the stator blades are fixed only to the inner ferrule and are in non-immobilizing mechanical contact with the outer ferrule; in which the outer ferrule includes a groove receiving radially external ends of the stator blades; characterized in that the groove extends axially to a downstream end of the outer ferrule.

[0008] In the invention, the stator blades are fixed only to the inner ferrule, thus avoiding areas of concentrated mechanical stress on the outer ferrule. The contact with the outer ferrule is non-immobilizing, meaning that it does not involve immobilizing the stator blades relative to the outer ferrule. The terms "free mechanical contact" or "non-fixing mechanical contact" could be used instead of "non-immobilizing mechanical contact." In other words, no element of the outer end of the blade immobilizes it on the outer ferrule. Such contact prevents the transmission of force between the blade and the outer ferrule via the outer radial end of the blade, while also preventing air leaks between the outer ferrule and the outer radial end of the blade.

[0009] Furthermore, in the invention, the groove in the outer ferrule, which extends axially to a downstream end of the outer ferrule, makes the mounting of the blades particularly easy.

[0010] In the prior art, the inner shell is attached to the other turbomachine components via the stator blades and the outer shell; therefore, a person skilled in the art would not think to eliminate the attachment to the outer shell. In the invention, the inner shell is designed to be attached to the other turbomachine components by other means. These means are preferably more rigid than in the prior art (generally supported by constrained supports). The force transmission chain (turbomachine / inner shell / blade) is thus more rigid than in the prior art.

[0011] It is interesting to note that, in the invention, it is the blades, which have mechanical contact (direct or indirect) with each of the two shells, that are fixed to the inner shell, and not auxiliary blades as described in EP2799721B1. Indeed, the latter only have mechanical contact with one of the two shells. Furthermore, they complement the stator blades to prevent flow stall on the stator blades; their function is not to replace the stator blades.

[0012] In one embodiment, the stator blades are welded to the inner shell. The welding provides a particularly strong attachment. Other attachment methods, such as bolting and / or riveting, are possible, while remaining within the scope of the invention.

[0013] In one embodiment, the outer shell comprises a sealing element made of a flexible material in contact with the radially external ends of the stator blades. The sealing element prevents leakage between the radially external ends of the stator blades and the outer shell. The flexible material preferably has a Young's modulus of less than 10 GPa. The flexible material may be, for example, silicone. The sealing element is preferably at least partially located within the groove. The sealing element may comprise several disjointed parts while remaining within the scope of the invention.

[0014] In one embodiment, the sealing element is located, at least partially, at a radially external position relative to the radially external ends of the stator blades and extends, at least partially, axially along the radially external ends of the stator blades. The radially external ends of the stator blades can slide on the sealing element while remaining in contact with it.

[0015] In one embodiment, the sealing element comprises a gasket. The gasket is preferably located at an upstream end or a downstream end of the groove. The radially external ends abut against it.

[0016] In one embodiment, the stator blades comprise, at their radially external end, a platform extending downstream. A sealing element in the form of a gasket is particularly advantageous in this case.

[0017] In one embodiment, the inner ferrule is a single piece. In another embodiment, the inner ferrule is made up of a plurality of sectors forming a ring.

[0018] The invention further proposes an aircraft turbomachine comprising a first compressor having a rectifier assembly according to an embodiment of the invention. The first compressor can be, for example, the low-pressure compressor or the high-pressure compressor of the turbomachine. In an aircraft turbomachine comprising the invention, the relative positioning of the outer shell with respect to the inner shell is achieved not by the blade but by one or more elements of the turbomachine external to the rectifier assembly.

[0019] The invention is particularly suited to a turbomachine comprising a reducer between the shaft and the fan, because the presence of the latter generates particularly significant mechanical forces on the outer shell.

[0020] In one embodiment, the turbomachine includes a second compressor, downstream of the first compressor. In this specific embodiment, it is the upstream compressor of the two that includes the rectifier assembly according to the invention.

[0021] In one embodiment, the stator blades attached only to the inner shell and in non-immobilizing mechanical contact with the outer shell are the stator blades furthest downstream of the first compressor. This allows the inner shell to be fixed downstream of the first compressor more easily than if the inner shell to which the stator blades are attached were axially in the middle of the first compressor.

[0022] In one embodiment, the turbomachine comprises an intermediate support housing located, preferably directly, downstream of the first compressor, the inner shell being fixed to the intermediate support housing or being integral with the intermediate support housing. This makes the attachment of the inner shell particularly easy and robust. The invention also relates to an assembly comprising the intermediate support housing and the rectifier assembly.

[0023] According to one embodiment, the outer ferrule is fixed to the intermediate support housing.

[0024] The invention further proposes an aircraft comprising a turbomachine according to the invention.

[0025] The invention further proposes a method for manufacturing a rectifier assembly, comprising the steps of: fix the stator blades to the inner ferrule, position the stator blades relative to the outer ferrule, and create a non-immobilizing mechanical contact between the stator blades and the outer ferrule, preferably by forming a sealing element at the junction between the stator blades and the outer ferrule. Brief description of the figures

[0026] Other features and advantages of the invention will become apparent upon reading the detailed description that follows, for an understanding of which reference should be made to the attached figures, among which: there figure 1 is an axial cross-section of a turbomachine according to an embodiment of the invention, the figure 2 illustrates a stator blade according to one embodiment of the invention, the figures 3a to 3c illustrate three embodiments of the invention, the figure 4is a flowchart of a manufacturing process for a rectifier assembly according to an embodiment of the invention, the figures 5a to 5d illustrative steps of this process in the case of an annular external ferrule, and the figure 6 is the equivalent of the figure 5b in the case of an external half-shell ferrule. Embodiments of the invention

[0027] The present invention is described with particular embodiments and references to figures, but the invention is not limited by them. The drawings or figures described are schematic only and are not limiting. Furthermore, the functions described can be performed by structures other than those described in this document.

[0028] In the context of this document, the terms "first" and "second" serve only to differentiate the different elements and do not imply any order between these elements.

[0029] In the figures, identical or analogous elements may bear the same references.

[0030] There figure 1This illustrates an aircraft turbomachine 100 that may include a rectifier assembly 1 according to the invention. It may also be called a "stator assembly". The aircraft turbomachine 100 is, for example, a twin-flow axial turbomachine comprising successively along the drive axis X, a fan 110, a first compressor 120 (or low-pressure compressor), a second compressor 130 (or high-pressure compressor), a combustion chamber 160, a high-pressure turbine 140, and a low-pressure turbine 150. In operation, the mechanical power of the low-pressure turbine 150 and the high-pressure turbine 140 is transmitted respectively via shafts 101 and 102 to the low-pressure compressor 120 and the high-pressure compressor 130, as well as to the fan 110 via a reduction gear 111 interposed at the shaft 101.The blower 110 generates a primary flow 106 through the aircraft turbomachine 100 in a primary aerodynamic channel and a secondary flow 107 externally around the compressors 120, 130 and the turbines 140, 150.

[0031] The first compressor 120 is provided with at least one row of rotor blades 122 followed directly downstream by a row of stator blades 10, each row of stator blades 10 forming a rectifier assembly 1. The invention can be applied to any one or more of the rectifier assemblies of the first compressor 120, and in particular to the rectifier assembly furthest downstream of the first compressor 120.

[0032] The aircraft turbomachine 100 includes an inlet support housing 181 that extends around the inlet of the primary flow (through which the primary flow 106 passes), downstream of the fan 110. The aircraft turbomachine 100 also includes an intermediate support housing 40 that extends circumferentially between the first 120 and second 130 compressors. This intermediate support housing 40 includes an annular sleeve, preferably with a gooseneck profile, that delimits the primary aerodynamic flow between the first 120 and second 130 compressors. It is preferably equipped with structural arms 184 extending radially through the primary flow.

[0033] There figure 2The figure illustrates a stator blade 10 of a rectifier assembly 1 according to an embodiment of the invention. The stator blade 10 is fixed, preferably by a weld 11, at its radially internal end 12, to an internal ferrule 20. The fixing between the stator blade 10 and the internal ferrule 20 prevents any relative movement. The stator blade 10 is in non-immobilizing mechanical contact, for example via a sealing element 31, at its radially external end 13, with an external ferrule 30. In one embodiment of the invention, the sealing element 31 is located, at least partially, in a groove 35, preferably circumferential, in the external ferrule 30. The groove 35 preferentially receives the radially external ends 13 of all the stator blades 10 of the rectifier assembly 1.

[0034] THE figures 3a to 3cillustrate three embodiments of the invention, which differ, on the one hand, in the attachment of the inner ferrule 20 to the intermediate support housing 40, and, on the other hand, in the mechanical coupling between the stator blade 10 and the outer ferrule 30. Those skilled in the art will understand that all ways of attaching the inner ferrule 20 to the intermediate support housing 40 are compatible with all mechanical couplings between the stator blade 10 and the outer ferrule 30.

[0035] As illustrated in the figure 3aThe groove 35 extends to the downstream end 32 of the outer ferrule 30. It is filled with a flexible material in contact with radially external ends 13 of the stator blades 10, and which forms the sealing element 31. This is located at a radially external position relative to the radially external ends 13 of the stator blades 10 and extends axially along the radially external ends 13 of the stator blades 10. In addition, the downstream end 22 of the inner ferrule 20 is fixed to the intermediate support housing 40 by fastening means 52, for example screws.

[0036] As illustrated in the figure 3bThe groove 35 extends to the downstream end 32 of the outer ferrule 30. It is filled with a flexible material in contact with the radially external ends 13 of the stator blades 10, and which forms the sealing element 31. This is located at a radially external position relative to the radially external ends 13 of the stator blades 10 and extends axially along the radially external ends 13 of the stator blades 10. In addition, the downstream end 22 of the inner ferrule 20 is integral with the intermediate support housing 40.

[0037] As illustrated in the figure 3cThe groove 35 extends to the downstream end 32 of the outer ferrule 30. A seal 60, for example an O-ring, is located at an upstream end 37 of the groove 35. The upstream end of the stator blade 10 abuts against it. It forms the sealing element 31. Furthermore, the stator blade 10 includes, at its radially external end 13, a platform 15 extending downstream and abutting against the outer ferrule 30. Moreover, the downstream end 22 of the inner ferrule 20 is integral with the intermediate support housing 40.

[0038] In the three embodiments illustrated in figures 3a-3cThe downstream end 32 of the outer shell 30 is fixed to the intermediate support housing 40 by means of fasteners 51, for example, screws. Furthermore, the downstream end 22 of the inner shell 20 is fixed to the intermediate support housing 40 or is integral with it. Consequently, in these three embodiments, the positioning of the inner shell 20 relative to the outer shell 30 does not stress the junction between the stator blades 10 and the outer shell 30, as this junction allows for relative movement. The positioning of the inner shell 20 relative to the outer shell 30, which bears the structural and operating loads of the turbomachine, is ensured by the junction of the inner shell 20 with respect to the intermediate support housing 40, and of the intermediate support housing 40 with respect to the outer shell 30.

[0039] THE figures 4 , 5a to 5d and 6illustrate certain steps of a process 200 for manufacturing a rectifier assembly 1 according to the invention, and its assembly with the intermediate support housing 40.

[0040] A block of metal, for example titanium, 201 is machined 202 to form the inner ferrule 20, preferably with holes 301 for fastening means 52. The inner ferrule 20 is then fixed 203 to the stator blades 10 ( figure 5a ). Inserts 302 (visible figures 5b And 6 in particular) are preferentially inserted into holes 301.

[0041] Then the stator blades 10 and the outer ferrule 30 are positioned 204 so as to leave a space between them which will be filled with a suitable material for non-immobilizing mechanical contact ( figures 5b And 6This positioning is preferably such that the downstream end 32 of the outer ferrule 30, and the downstream end 22 of the inner ferrule 20 are located lower than the stator blades 10 and rest on a support tool 304. The abradable 303 of the outer ferrule 30 is then preferably located higher than the stator blades 10. The figure 5b Diagram shows a lifting tool 305 for lifting the outer annular ferrule 30. The arrow 306 of the figure 6 indicates that the radial flange of the external half-shell ferrule 30 is located higher.

[0042] The material suitable for non-immobilizing mechanical contact is then deposited 205 at the junction between the stator blades 10 and the outer shell 30, for example using a mold 307, which is preferably such that the said material does not adhere to it. The mold 307 can be fixed to the support tooling 304. This yields a rectifier assembly 1, which is inverted and assembled 206 to the intermediate support housing 40. The fastening means 51 may include screws 51a and nuts 51b.

[0043] The present invention has been described in relation to specific embodiments, which are purely illustrative and should not be considered limiting. In general, the present invention is not limited to the examples illustrated and / or described above. The use of the verbs "to include," "to comprise," or any other variant thereof, as well as their conjugations, does not in any way preclude the presence of elements other than those mentioned. The use of the indefinite article "a," "an," or the definite article "the," "a," or "it" to introduce an element does not preclude the presence of a plurality of such elements. The reference numbers in the claims do not limit their scope.

Claims

1. A stator vane assembly (1) for a compressor (120, 130) of an aircraft turbine engine (100), comprising: - an internal shroud (20), - an external shroud (30), and - stator vanes (10), wherein the stator vanes (10) are attached only to the internal shroud (20) and are in non-immobilizing mechanical contact with the external shroud (30); wherein the external shroud (30) comprises a groove (35) receiving radially external ends (13) of the stator vanes (10); characterized in that the groove (35) extends axially to a downstream end (32) of the external shroud (30).

2. The stator vane assembly according to claim 1, wherein the stator vanes (10) are welded to the internal shroud (20).

3. The stator vane assembly according to any of the preceding claims, wherein the external shroud (30) comprises a sealing element (31, 60) of a flexible material in contact with radially external ends (13) of the stator vanes (10).

4. The stator vane assembly according to the preceding claim, wherein the sealing element (31, 60) is located, at least partly, at a radially external position relative to the radially external ends (13) of the stator vanes (10) and at least partly extends axially along the radially external ends (13) of the stator vanes (10).

5. The stator vane assembly according to claim 3 or 4, wherein the sealing element (31, 60) comprises a seal (60).

6. The stator vane assembly according to any of the preceding claims, wherein the internal shroud (20) is in one piece or is made up of a plurality of sectors forming a ring.

7. An aircraft turbine engine (100) comprising a first compressor (120, 130) comprising a stator vane assembly (1) according to any of the preceding claims.

8. The aircraft turbine engine according to the preceding claim, comprising a second compressor (130), downstream of the first compressor (120).

9. The aircraft turbine engine according to claim 7 or 8, wherein the stator vanes (10) of said stator vane assembly (1) are the stator vanes furthest downstream of the first compressor (120).

10. The aircraft turbine engine according to the preceding claim, comprising an intermediate support casing (40) located, preferably directly, downstream of the first compressor (120), the internal shroud (20) being attached to the intermediate support casing (40) or being in one piece with the intermediate support casing (40).

11. The aircraft turbine engine according to the preceding claim, wherein the external shroud (30) is attached to the intermediate support casing (40).

12. An aircraft comprising a turbine engine according to any of claims 7 to 11.