Stator vane assembly for a compressor of an aircraft turbine engine

Abstract

The invention relates to a stator blade assembly (1) for a compressor (120, 130) of an aircraft turbine engine (100). The assembly comprises an inner shroud (20), an outer shroud (30) and stator blades (10) which are attached only to the inner shroud (20) and in non-fixed mechanical contact with the outer shroud (30).

Description

Technical Field

[0001] This invention relates to a turbine engine compressor for aircraft. Background Technology

[0002] For example, as known from document EP2799721B1, the stator blades of the stator blade assembly (or rectifier assembly) of an aircraft turbine engine compressor can be attached to a radially external shroud (referred to as the external shroud). The document also describes auxiliary blades (or blades) that are elements located between the stator blades, and that the radial height of the auxiliary blades is between 10% and 50% of the radial height of the stator blades (or stator blades).

[0003] Document US 3 778 184 A describes a compressor in which blade damping is achieved by surrounding one end of the blade with a damping material of the type of steel wool or metal felt that remains in contact with the fairing.

[0004] Document EP 2 093 383 A1 describes a compressor in which stator blades are attached to an internal shroud. Summary of the Invention

[0005] The outer casing is subjected to considerable mechanical stress, especially in turbine engine structures where it is located in the main thrust path. Some of these mechanical stresses originate from the stator blades attached to the outer casing.

[0006] One objective of this invention is to reduce mechanical stress in aircraft turbine engines.

[0007] Therefore, the present invention provides a stator blade assembly (or rectifier assembly) for a compressor of an aircraft turbine engine, the stator blade assembly (or rectifier assembly) comprising:

[0008] • Inner protective cover

[0009] • External protective cover, and

[0010] • Stator blades (or stator blades)

[0011] The stator blades are only attached to the inner cover and are not in fixed mechanical contact with the outer cover;

[0012] The outer cover includes a groove for receiving the radially outer end of the stator blade;

[0013] The feature is that the groove extends axially to the downstream end of the outer cover.

[0014] In this invention, the stator blades are attached only to the inner shroud, which avoids the formation of areas where mechanical stress concentrates on the outer shroud. The contact between the stator blades and the outer shroud is non-fixed; that is, the stator blades are not involved in fixing the stator blades relative to the outer shroud. "Free mechanical contact" or "non-attached mechanical contact" can be used instead of "non-fixed mechanical contact." In other words, there are no elements on the outer end of the blades to fix the blades to the outer shroud. This contact avoids the transmission of force between the blades and the outer shroud through the radially outer ends of the blades, and also avoids air leakage between the outer shroud and the radially outer ends of the blades.

[0015] Furthermore, in this invention, the groove in the outer casing makes it particularly easy to install the blades, and the groove extends axially to the downstream end of the outer casing.

[0016] In the prior art, the inner shield is attached to other components of the turbine engine via the stator blades and the outer shield, so those skilled in the art would not conceive of removing the attachment to the outer shield. In this invention, the inner shield is designed to be attached to other components of the turbine engine via other means. These means are preferably more rigid than those in the prior art (typically constrained supports). Therefore, the force transmission chain (turbine engine / inner shield / blades) is more rigid than the force transmission chain in the prior art.

[0017] It is noteworthy that, in this invention, the blades (attached to the inner shroud) are in mechanical contact (directly or indirectly) with each of the two shrouds, not with the auxiliary blades as described in EP2799721B1. In fact, the auxiliary blades are in mechanical contact with only one of the two shrouds. Furthermore, the auxiliary blades supplement the stator blades to prevent flux stall on the stator blades: the auxiliary blades are not intended to replace the stator blades.

[0018] In one embodiment, the stator blades are welded to the inner shroud. The weld makes the attachment particularly secure. Other attachment methods (e.g., bolted connections and / or riveting) are possible, but still within the scope of this invention.

[0019] In one embodiment, the outer shroud includes a sealing element of flexible material that contacts the radially outer end of the stator blade. The sealing element prevents leakage between the radially outer end of the stator blade and the outer shroud. The flexible material preferably has a Young's modulus of less than 10 GPa. For example, the flexible material can be silicone resin. The sealing element is preferably at least partially recessed. The sealing element may comprise multiple separate portions, while remaining within the scope of the invention.

[0020] According to one embodiment, the sealing element is at least partially located radially outward relative to the radially outer end of the stator blade and extends axially at least partially along the radially outer end of the stator blade. The radially outer end of the stator blade can slide on the sealing element while maintaining contact with it.

[0021] In one embodiment, the sealing element includes a seal. The seal is preferably located at the upstream or downstream end of the groove. The radially outer end abuts against the seal.

[0022] In one embodiment, the stator blade includes a downstream-extending platform at its radially outer end. In this case, a sealing element in the form of a seal is particularly advantageous.

[0023] In one embodiment, the inner shield is integral. In another embodiment, the inner shield is composed of a plurality of sector-shaped elements forming a ring.

[0024] The present invention also proposes an aircraft turbine engine including a first compressor comprising a stator blade assembly according to an embodiment of the invention. The first compressor may be, for example, a low-pressure compressor or a high-pressure compressor of a turbine engine. In the aircraft turbine engine including the present invention, the relative positioning of the outer casing relative to the inner casing is not achieved using blades, but rather through one or more elements of the turbine engine located outside the stator blade assembly.

[0025] This invention is particularly applicable to turbine engines that include a gearbox located between the shaft and the fan, because the presence of the gearbox generates particularly high mechanical stress on the outer casing.

[0026] In one embodiment, the turbine engine includes a second compressor located downstream of the first compressor. In this particular embodiment, a stator blade assembly according to the invention is located further upstream of both compressors.

[0027] In one embodiment, the stator blades that are only attached to the inner shroud and not in fixed mechanical contact with the outer shroud are the downstream stator blades of the first compressor. This makes it easier to attach the inner shroud downstream of the first compressor compared to the inner shroud with the attached stator blades being axially located in the middle of the first compressor.

[0028] According to one embodiment, the turbine engine includes an intermediate support housing preferably located directly downstream of a first compressor, with an inner shroud attached to or integral with the intermediate support housing. This makes attachment of the inner shroud particularly easy and secure. The invention also relates to an assembly comprising an intermediate support housing and a stator blade assembly.

[0029] In one embodiment, the outer protective cover is attached to the intermediate support housing.

[0030] The present invention also proposes an aircraft comprising a turbine engine according to the invention.

[0031] The present invention also proposes a method for manufacturing a stator blade assembly, the method comprising the following steps:

[0032] • Attach the stator blades to the internal protective cover.

[0033] • Positioning the stator blades relative to the outer casing, and

[0034] • Preferably, a non-fixed mechanical contact is formed between the stator blades and the outer casing by forming a sealing element at the joint between the stator blades and the outer casing. Attached Figure Description

[0035] Further features and advantages of the invention will become apparent from the following detailed description, which, for the purpose of understanding, is illustrated in the accompanying drawings, in which:

[0036] - Figure 1 This is an axial cross-sectional view of a turbine engine according to an embodiment of the present invention.

[0037] - Figure 2 A stator blade according to an embodiment of the present invention is shown.

[0038] - Figures 3a to 3c Three embodiments of the present invention are shown.

[0039] - Figure 4 This is a flowchart of a method for manufacturing a stator blade assembly according to an embodiment of the present invention.

[0040] - Figures 5a to 5d The steps for the annular outer shield in this method are shown, and

[0041] - Figure 6 Equivalent to Figure 5b The semi-shell-type external protective cover. Detailed Implementation

[0042] The invention has been described with reference to specific embodiments and the accompanying drawings, but the invention is not limited thereto. The described figures or drawings are merely illustrative and not limiting. Furthermore, the described functions can be performed by structures other than those described herein.

[0043] In this document, the terms “first” and “second” are used only to distinguish various elements and do not imply the order of these elements.

[0044] In the accompanying drawings, the same or similar elements may have the same reference numerals.

[0045] Figure 1 An aircraft turbine engine 100 is shown, which may include a stator-blade assembly 1 according to the invention. This stator-blade assembly may also be referred to as a "stator assembly". The aircraft turbine engine 100 is, for example, a twin-axial-flow turbine engine that sequentially includes, along the engine axis X, a fan 110, a first compressor 120 (or a low-pressure compressor), a second compressor 130 (or a 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 via shafts 101 and 102 to the low-pressure compressor 120 and the high-pressure compressor 130, respectively, and to the fan 110 by means of a gearbox 111 located at shaft 101. The fan 110 enables the generation of a main flow 106 passing through the main aerodynamic duct of the aircraft turbine engine 100 and a secondary flow 107 surrounding the compressors 120, 130, and turbines 140, 150 externally.

[0046] The first compressor 120 is equipped with at least one row of rotor blades 122, and directly downstream of the rotor blades is a row of stator blades 10, each row of stator blades 10 forming a stator blade assembly 1. The present invention can be applied to any or all stator blade assemblies of the first compressor 120, particularly to the most downstream stator blade assembly of the first compressor 120.

[0047] The aircraft turbine engine 100 includes an inlet support housing 181 that extends downstream of the fan 110 around the inlet of a main duct (through which the main flow 106 passes). The aircraft turbine engine 100 also includes an intermediate support housing 40 extending circumferentially between a first compressor 120 and a second compressor 130. The intermediate support housing 40 includes an annular sleeve, preferably having a gooseneck profile, defining a main aerodynamic duct between the first compressor 120 and the second compressor 130. The intermediate support housing is preferably equipped with a structural arm 184 extending radially across the main duct.

[0048] Figure 2 A stator blade 10 of a stator blade assembly 1 according to an embodiment of the present invention is shown. The stator blade 10 is preferably attached to an inner shroud 20 at its radially inner end portion 12 via a weld 11. The attachment between the stator blade 10 and the inner shroud 20 prevents any relative movement. The stator blade 10 is in non-fixed mechanical contact with an outer shroud 30 at its radially outer end portion 13, for example via a sealing element 31. In one embodiment of the invention, the sealing element 31 is at least partially located in a recess 35 in the outer shroud 30, the recess preferably being circumferential. The recess 35 preferably receives the radially outer ends 13 of all the stator blades 10 of the stator blade assembly 1.

[0049] Figures 3a to 3c Three embodiments of the invention are illustrated, differing in one aspect in the attachment of the inner shield 20 to the intermediate support housing 40, and in another aspect in the mechanical connection between the stator blade 10 and the outer shield 30. Those skilled in the art will understand that all methods of attaching the inner shield 20 to the intermediate support housing 40 are compatible with all mechanical connections between the stator blade 10 and the outer shield 30.

[0050] like Figure 3a As shown, the groove 35 extends to the downstream end 32 of the outer shroud 30. The groove is filled with a flexible material that contacts the radially outer end 13 of the stator blade 10 and forms a sealing element 31. This sealing element is located radially outward relative to the radially outer end 13 of the stator blade 10 and extends axially along the radially outer end 13 of the stator blade 10. Furthermore, the downstream end 22 of the inner shroud 20 is attached to the intermediate support housing 40 by an attachment device 52 (e.g., a screw).

[0051] like Figure 3b As shown, the groove 35 extends to the downstream end 32 of the outer shroud 30. The groove is filled with a flexible material that contacts the radially outer end 13 of the stator blade 10 and forms a sealing element 31. The sealing element is located radially outward relative to the radially outer end 13 of the stator blade 10 and extends axially along the radially outer end 13 of the stator blade 10. Furthermore, the downstream end 22 of the inner shroud 20 is integral with the intermediate support housing 40.

[0052] like Figure 3c As shown, the groove 35 extends to the downstream end 32 of the outer shroud 30. A seal 60 (e.g., an O-ring) is located at the upstream end 37 of the groove 35. The upstream end of the stator blade 10 abuts against the seal. The seal forms a sealing element 31. Furthermore, the stator blade 10 includes a platform 15 at its radially outer end 13, which extends downstream and abuts against the outer shroud 30. Additionally, the downstream end 22 of the inner shroud 20 is integral with the intermediate support housing 40.

[0053] exist Figures 3a to 3cIn the three embodiments shown, the downstream end 32 of the outer shield 30 is attached to the intermediate support housing 40 by an attachment device 51 (e.g., screws). Furthermore, the downstream end 22 of the inner shield 20 is attached to or integral with the intermediate support housing 40. Therefore, in these three embodiments, the positioning of the inner shield 20 relative to the outer shield 30 does not exert stress on the joint between the stator blade 10 and the outer shield 30, as this joint allows for relative displacement. The inner shield 20 is positioned relative to the outer shield 30, and the outer shield absorbs the structural and operating forces of the turbine engine by engaging the inner shield 20 relative to the intermediate support housing 40 and engaging the inner shield to the intermediate support housing 40 relative to the outer shield 30.

[0054] Figure 4 , Figures 5a to 5d as well as Figure 6 Some steps of a method 200 for manufacturing a stator blade assembly 1 according to the invention and for assembling the stator blade assembly with an intermediate support housing 40 are shown.

[0055] A metal (e.g., titanium) block 201 is processed 202 to form an inner shield 20, which preferably has holes 301 for attaching the attachment device 52. The inner shield 20 is then attached 203 to the stator blade 10. Figure 5a Insert 302 (especially in) Figure 5b and Figure 6 (As can be seen in the image) Preferably inserted into hole 301.

[0056] The stator blades 10 and the outer shroud 30 are then positioned 204 to leave a space between the stator blades and the outer shroud, which will be used to fill the space with a suitable material for non-fixed mechanical contact. Figure 5b and Figure 6 The positioning preferably positions the downstream end 32 of the outer cover 30 and the downstream end 22 of the inner cover 20 lower than the stator blade 10 and rests on the support tool 304. Then, the wear-resistant portion 303 of the outer cover 30 is preferably positioned higher than the stator blade 10. Figure 5b A schematic diagram of a lifting tool 305 that enables the lifting of the annular outer protective cover 30 is shown. Figure 6 Arrow 306 in the diagram shows that the radial flange of the semi-shell outer cover 30 is located at a higher position.

[0057] Then, for example, a mold 307 is used to deposit 205 a material suitable for non-fixed mechanical contact at the joint between the stator blade 10 and the outer casing 30, preferably such that the material does not adhere to the mold. The mold 307 can be attached to a support tool 304. The result is the formation of a stator blade assembly 1, which is flipped and assembled 206 to the intermediate support housing 40. The attachment device 51 may include screws 51a and nuts 51b.

[0058] The invention has been described above with reference to specific embodiments, which are illustrative and should not be considered limiting. In general, the invention is not limited to the examples shown and / or described above. The use of the verbs “comprising,” “including,” or any other variations thereof, and their inflectional variations, does not exclude the presence of elements other than those described. The use of the indefinite articles “a,” “an,” or the definite article “the” to introduce an element does not exclude the presence of a plurality of such elements. Reference numerals in the claims do not limit the scope of the claims.

Claims

1. An aircraft turbine engine (100), the aircraft turbine engine including a first compressor (120), the first compressor including a stator blade assembly (1), the stator blade assembly (1) including: • Inner shield (20) • External protective cover (30), and • Stator blades (10). The stator blade (10) is only attached to the inner cover (20) and is not in fixed mechanical contact with the outer cover (30); The outer cover (30) includes a groove (35) that receives the radially outer end (13) of the stator blade (10). The groove (35) is characterized in that it extends axially to the downstream end (32) of the outer cover (30).

2. The aircraft turbine engine (100) according to claim 1, wherein, The stator blades (10) are welded to the inner cover (20).

3. The aircraft turbine engine (100) according to claim 1 or 2, wherein, The outer cover (30) includes sealing elements (31, 60) of flexible material that contact the radially outer end (13) of the stator blade (10).

4. The aircraft turbine engine (100) according to claim 3, wherein, The sealing element (31, 60) is located at least partially in a radially outer position relative to the radially outer end (13) of the stator blade (10), and extends at least partially axially along the radially outer end (13) of the stator blade (10).

5. The aircraft turbine engine (100) according to claim 3, wherein, The sealing elements (31, 60) include seals.

6. The aircraft turbine engine (100) according to claim 1 or 2, wherein, The inner shield (20) is integral or consists of multiple sector-shaped elements forming a ring.

7. The aircraft turbine engine according to claim 1, wherein the aircraft turbine engine includes a second compressor (130) located downstream of the first compressor (120).

8. The aircraft turbine engine according to claim 1 or 2, wherein, The stator blade (10) in the stator blade assembly (1) is the stator blade located at the downstream end of the first compressor (120).

9. The aircraft turbine engine according to claim 8, wherein the aircraft turbine engine includes an intermediate support housing (40), and the inner shield (20) is attached to or integral with the intermediate support housing (40).

10. The aircraft turbine engine according to claim 9, wherein, The outer protective cover (30) is attached to the intermediate support housing (40).

11. The aircraft turbine engine according to claim 10, wherein, The intermediate support housing (40) is located directly downstream of the first compressor (120).

12. An aircraft comprising an aircraft turbine engine according to any one of claims 1 to 11.

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