Guide vane assembly for turbine of aircraft turbine engine

Abstract

The invention relates to a guide vane assembly (34), in particular for a turbine of an aircraft turbine engine, comprising a plurality of sections (40) arranged circumferentially side-by-side about a main axis, each section (40) comprising an outer annular platform section (52), an inner annular platform section (54), and at least one vane (42), according to the invention, at least one vane (42) has a suction side wall and a pressure side wall connected to each other by a leading edge (BA) and a trailing edge (BF) extending between an outer annular platform section (52) and an inner annular platform section (54), respectively, the trailing edge (BF) comprising at least one recess (60b) comprising a bottom (62) offset relative to a plane (M) passing through the trailing edge (BF) of the vane (42) and perpendicular to the main axis (Z).

Description

Technical Field

[0001] This invention relates to a guide vane assembly for a turbine in an aircraft turbine engine. The invention also relates to a turbine of an aircraft turbine engine equipped with such a guide vane assembly, and to an aircraft turbine engine including such a turbine or such a guide vane assembly. Background Technology

[0002] Figure 1 This is an axial (or longitudinal) cross-sectional view of an aircraft turbine engine 10, which typically includes, from upstream to downstream, a fan 12 and multiple modules such as a low-pressure compressor 14, a subsequent high-pressure compressor 16, a combustion chamber 18, a high-pressure turbine 20, a subsequent low-pressure turbine 22, and a gas injection system 24, with the high-pressure turbine and low-pressure turbine driving the corresponding low-pressure compressor 14 or high-pressure compressor 16.

[0003] The terms “upstream” and “downstream” are defined relative to the normal flow direction of gas in the flow passage of the turbine engine 10, that is, from the upstream fan 12 to the downstream gas injection system 24.

[0004] The main airflow P passes through the high-pressure body and the low-pressure body, and the fan 12 generates a secondary airflow S, which flows in the turbine engine 10 between the outer casing 13 and the housing 11 of the turbine engine 10 in the cold flow passage 15.

[0005] At the outlet of the gas injection system 24, the gas from the main stream P mixes with the secondary stream S to generate thrust, wherein the secondary stream S provides most of the thrust.

[0006] Figure 2 It shows Figure 1 A detailed view of the turbine of the turbine engine 10. More specifically, the high-pressure turbine 20 and low-pressure turbine 22 of the turbine engine 10 typically include one or more stages, each stage comprising: - The fixed blade assembly or annular arrangement 30 or ring-shaped portion of the impeller supported by the turbine housing 11, and -A ring-shaped arrangement 32 of wheel blades or blades mounted to rotate around a central axis of rotation.

[0007] The annular portion of the turbine's fixed blade assembly is commonly referred to as the guide vane assembly 34. The guide vane assembly 34 deflects and accelerates the gas flow from the combustion chamber 18 toward the movable turbine blades at an appropriate angle and speed to drive the turbine rotor and the movable blades to rotate.

[0008] To facilitate assembly and reduce manufacturing costs, the annular portion of the blade fixing device is typically manufactured as a component of corner sections, which are placed adjacent to each other to form the entire annular portion for fixing the blade.

[0009] The guide vane assembly 34 of the high-pressure turbine 20 of the turbine engine 10, and thus the blades constituting the guide vane assembly, are components exposed to very high thermal stress. The guide vane assembly and blades are located at the outlet of the combustion chamber 18 and are therefore subjected to extremely hot gases, which subject them to very high thermal stress. Therefore, cooling of the guide vane assembly 34 (especially the blades) is necessary.

[0010] To ensure effective cooling of the guide vane assembly, a highly thermally efficient cooling system is desired, enabling the dissipation of high heat power using a moderate airflow rate.

[0011] One of the most widely used solutions currently is the use of one or more liners or inserts arranged within the blades of the guide vane assembly. The liners are cooling devices formed from porous metal plates to cool the interior of the blades by the impact of cooling air. Thus, the inner surface of the blades is cooled by jet impact and forced convection between the inserts and the inner surface of the blades.

[0012] The term "cooling air" or "cold air" refers to air whose temperature is lower than that of the gas flowing through the flow channels of a turbine engine, thereby reducing the temperature of components located in the flow channels.

[0013] Document FR-A1-2 899 271 describes such a device for cooling wheel blades.

[0014] The impeller comprises a pressure wall and a suction wall joined together by a leading edge BA facing upstream and a trailing edge BF facing downstream.

[0015] However, particularly on one half of the suction wall and at the trailing edge, these blades of the high-pressure guide vane assembly include thin surfaces of considerable size. These are subjected to high compression, which can lead to buckling or warping damage.

[0016] When a structure is longitudinally compressed, it tends to bend perpendicular to the axis of the applied force due to elastic instability. Therefore, this region includes plates that have undergone significant deformation perpendicular to the direction of the stress.

[0017] Here, one direction of the stress is in the plane of the plate, so the resulting deformation is out-of-plane. This out-of-plane deformation can cause bulging at the center of the plate surrounded by cracks.

[0018] This damage caused premature failure of the relevant guide vane assembly section, so the guide vane assembly section had to be replaced.

[0019] Figure 3A section 40 of a high-pressure guide vane assembly with two vanes 42 is shown, at least one of the vanes 42 having a bulging region 44 in the suction wall of the vane 42.

[0020] In addition, another buckling mode was observed on the blades in the high-pressure guide vane assembly area at the trailing edge BF.

[0021] Figure 4 Another such buckling mode is shown at the trailing edge BF of one of the blades 42 in the region 40 of the guide blade assembly 34.

[0022] More specifically, Figure 4 This phenomenon is shown in region 46 of the trailing edge BF of one blade in the blade 42 of the high-pressure guide vane assembly, section 40.

[0023] Because the trailing edge BF of the blade 42 of the high-pressure guide vane assembly is straight, the trailing edge can be likened to a compressed plate, reinforced at its three edges, with the fourth edge free in the channel. This trailing edge is also subject to expansion stress due to its location in a very hot region.

[0024] The cooler remainder of blade 42 prevents this expansion. Furthermore, because the remainder of blade 42 is larger, the trailing edge BF absorbs most of the expansion stress up to a certain threshold beyond which the stress becomes sufficient to buckle and deform the trailing edge BF.

[0025] Therefore, when the trailing edge BF is in a state of intense compression, the trailing edge seeks to release the compression-related energy through deformation that leads to an irreversible final deformation state.

[0026] Another objective of this invention is to provide a solution for limiting buckling at the trailing edge of a high-pressure guide vane assembly.

[0027] Documents FR-A1-3119210, US-B2-9963974, US-A1-2021 / 115796, US-A1 2011 / 293422, US-A1-2020 / 190984, and US-A1-2009 / 155082 illustrate examples of turbine blades with similar problems. Summary of the Invention

[0028] Therefore, a guide vane assembly is proposed, specifically for use in the turbines of aircraft turbine engines. The guide vane assembly includes multiple sections arranged side-by-side circumferentially around a main axis. Each section includes an outer annular platform section, an inner annular platform section, and at least one hollow vane. The at least one hollow vane has suction walls and pressure walls connected to each other by a leading edge and a trailing edge, which extend between the outer and inner annular platform sections, respectively. The at least one vane includes cooling grooves arranged near the trailing edge, which allow cooling airflow to flow from the interior to the exterior of the vane.

[0029] More specifically, the trailing edge of the blade includes at least one recess, the recess including a bottom offset relative to a plane that passes through the trailing edge of the blade guiding the blade assembly and is perpendicular to the main axis, the depth of the bottom of the recess being between 10% and 100% of the region of the trailing edge located downstream of the groove of the blade.

[0030] Therefore, the guide vane assembly according to the invention ensures reduced deformation of the trailing edge of the guide vane in the designated area. The presence of one or more recesses reduces the stiffness of the trailing edge; in other words, this reduces the load on the vane.

[0031] Therefore, the present invention enables the improvement of the blade performance in the guide vane assembly region when the blades and platform of the guide vane assembly are subjected to thermomechanical and mechanical stresses.

[0032] In this application, the depth of the recess is a dimension measured in a direction perpendicular to the aforementioned plane, and the depth of the recess may be parallel to the main axis. Furthermore, the height of the recess is a dimension measured in the aforementioned plane, preferably radially oriented relative to the main axis.

[0033] The guide vane assembly according to the invention may include one or more of the following features, which may be employed individually or in combination with each other: - The recess has a generally straight profile, and in particular has a bottom that is generally parallel to the plane; -The concave portion has a triangular outline; - The concave part has a concave profile; - The bottom of the recess is defined by its depth, which is greater than the height of the recess; - The bottom of the recess is defined by its depth, which is smaller than the height of the recess; -The trailing edge includes a plurality of recesses arranged along the trailing edge; -The trailing edge includes a single recess along the trailing edge; - Particularly in the region of the blade located downstream of the groove on the trailing edge, a single recess is between 50% and 97% of the radial dimension of the trailing edge extending between the outer and inner annular plateau regions, particularly between 75% and 95%. - The depth of the recess is between 0.6 mm and 6 mm, particularly between 3 mm and 6 mm, and especially between 4.5 mm and 5.7 mm; - The depth of the bottom of the recess is between 50% and 95% of the area of ​​the trailing edge located downstream of the groove of the impeller, especially between 75% and 95%.

[0034] The present invention also relates to a turbine for an aircraft turbine engine, the turbine including at least one guide vane assembly as described above.

[0035] The present invention also relates to an aircraft turbine engine comprising at least one guide vane assembly as described above and / or a turbine as described above.

[0036] Of course, various features, variations and / or embodiments of the present invention can be combined with each other in various ways, as long as they are not incompatible or exclusive. Attached Figure Description

[0037] The invention will be better understood from the following detailed description, and other features and advantages will become apparent. The following detailed description includes embodiments given by way of illustration only with reference to the accompanying drawings, which can be used to enhance the understanding of the invention and the description of embodiments thereof, and, where appropriate, to help limit the invention. In the drawings: Figure 1 This is an axial (or longitudinal) cross-sectional view of a turbine engine to which the present invention can be applied; Figure 2 It shows Figure 1 A detailed view of the turbine of the turbine engine shown; Figure 3 The guide vane assembly section with a bulging region having suction walls is shown; Figure 4 Another buckling mode is shown at the trailing edge of the blade in the guide vane assembly area; Figure 5 A schematic perspective view of a guide vane assembly section according to the prior art is shown; Figure 6 It shows Figure 5 A schematic cross-sectional view of the blades in the guide vane assembly area shown; Figure 7 It shows Figure 5 A schematic diagram of the trailing edge of the blade in the guide vane assembly area; Figure 8 A schematic perspective view of a guide vane assembly section according to a first embodiment of the present invention is shown; Figure 9 Showing from Figure 8 A schematic diagram of the trailing edge of the blade in the guide vane assembly area; Figure 10 A schematic perspective view of a guide vane assembly section according to a second embodiment of the present invention is shown; Figure 11 Showing from Figure 10 A schematic diagram of the trailing edge of the blade in the guide vane assembly area; Figure 12 A schematic perspective view of a guide vane assembly section according to a third embodiment of the present invention is shown; and Figure 13 Showing from Figure 12 A schematic diagram of the trailing edge of the blade in the guide vane assembly area. Detailed Implementation

[0038] This invention is applicable to Figure 1 The turbine engine 10 and its turbine are shown in the diagram. One such turbine is, for example, as shown in the diagram. Figure 2 The high-pressure turbine 20 is shown.

[0039] As already noted, the turbine guide vane assembly 34 is typically divided into multiple sections and is formed by multiple guide vane assembly sections arranged juxtaposed on a circumference centered on the turbine axis (i.e., the axis Z of the turbine engine 10), such as... Figure 1 As shown. Axis Z is also referred to as the main axis of the guide vane assembly 34.

[0040] For convenience, the following description and figures refer to section 40 of the guide vane assembly 34. Therefore, the description of section 40 can be applied to all sections of the guide vane assembly 34.

[0041] Figure 5 A schematic perspective view of section 40 of a commonly used guide vane assembly 34 is shown as an example.

[0042] Typically, this section 40 includes an outer annular platform section 52 and an inner annular platform section 54 relative to the axis Z. The outer annular platform section 52 and the inner annular platform section 54 define an annular flow channel for gas in the turbine between them.

[0043] The terms “lower,” “inner,” or “internal” are used to indicate any element whose location is radially close to axis Z, while the terms “upper,” “outer,” or “external” are used to indicate any element whose location is radially further away from axis Z than the lower location.

[0044] Therefore, the outer annular platform region 52 is radially outside the inner annular platform region 54. The outer annular platform region 52 and the inner annular platform region 54 are coaxial.

[0045] It should be understood that the outer annular platform section 52 and the inner annular platform section 54 are configured to form an outer annular platform and an inner annular platform, respectively, which are part of the guide vane assembly 34.

[0046] The guide vane assembly 34 also includes at least one vane 42 extending between the outer annular platform region 52 and the inner annular platform region 54.

[0047] The blade 42 extends longitudinally in the first direction, roughly along the radial direction R.

[0048] The blade 42 can also be defined by a chord 48, which corresponds to the length of the section, i.e. the distance between the leading edge BA and the trailing edge BF.

[0049] The impeller 42 can be connected to the inner annular platform region 52 and the outer annular platform region 54 by connecting fillets 43a, 43b or connecting portions on the pressure side 56 and suction side 58, from the leading edge BA to the trailing edge BF, around the entire circumference of the impeller 42.

[0050] The impeller 42 is hollow up to a certain height from which cooling air flows out. The cooling air comes from the supply casing (not shown) and, for example, from the outlet of the high-pressure compressor 16.

[0051] Figure 6 It shows Figure 5 A schematic cross-sectional view of the blade 42 of section 40. The blade 42 includes a pressure wall 56 and a suction wall 58 joined together by an upstream leading edge BA and a downstream trailing edge BF.

[0052] Figure 7 It shows the commonly used Figure 5 A schematic diagram of the trailing edge BF of blade 42.

[0053] Figures 8 to 13 Different embodiments of the present invention are shown.

[0054] exist Figures 8 to 13 In the embodiment of the invention shown, the trailing edge BF includes at least one recess 60a, 60b, 60c.

[0055] In this specification, "recess" refers to the cut in the trailing edge BF.

[0056] The recesses 60a, 60b, and 60c also include a bottom 62 offset relative to the plane M, which passes through the trailing edge BF of the blade 42 of the guide vane assembly 34 and is perpendicular to the main axis Z of the guide vane assembly 34 (i.e., the axis of the turbine engine 10).

[0057] The blade 42 may include a cooling device for the blade, which communicates with the outside of the blade 42 via slots 50 arranged near the trailing edge BF. These slots 50 are cooling slots, and their function is to allow cooling airflow to pass through them, specifically from the inside of the blade to the outside.

[0058] Alternatively or additionally, the depth of the recesses 60a, 60b, 60c of the trailing edge BF may be between 10% and 100% of the region A of the trailing edge BF downstream of the groove 50 of the guide vane assembly 34 of the vane 42, particularly between 50% and 95%, particularly between 75% and 95%.

[0059] In particular, the recesses 60a, 60b, and 60c of the trailing edge BF can therefore be offset relative to the plane M by a distance between 10% and 100% of the region A of the trailing edge BF located downstream of the injection groove 50 of the impeller 42, particularly between 50% and 95%, and especially between 75% and 95%.

[0060] This maintains the cooling efficiency of the blades while improving their resistance to thermomechanical and mechanical stresses.

[0061] More specifically, the depth of the recesses 60a, 60b, and 60c in the trailing edge BF can be between 0.6 mm and 6 mm, particularly between 3 mm and 6 mm, and especially between 4.5 mm and 5.7 mm.

[0062] Advantageously, the recesses 60a, 60b, and 60c can have different shapes and / or profiles. The choice of shape and / or profile depends on the stress distribution in the trailing edge BF and is determined in advance during the design phase of the guide vane assembly 34.

[0063] Shape and / or profile can affect the aerodynamics and / or integration of desired components. For example, multiple consecutive recesses along the trailing edge BF enable the maintenance of flow rate in the channels guiding the impeller assembly.

[0064] Furthermore, recesses 60a, 60b, and 60c increase the threshold of the first buckling mode. In other words, this imparts better thermomechanical resistance to the trailing edge BF.

[0065] Figure 8 and Figure 9 A first embodiment according to the present invention is shown. Figure 8A schematic perspective view of the trailing edge BF of the blade 42 of the guide vane assembly 34 in the region 40 and the region 40 of the guide vane assembly 34.

[0066] More specifically, Figure 8 and Figure 9 A first embodiment according to the invention is shown, wherein the trailing edge BF includes at least one recess 60a having a generally straight profile. According to an alternative embodiment, the generally straight-profiled recess 60a is unique. According to the first embodiment of the invention, the bottom 62 of the recess 60a may be generally parallel to the plane M.

[0067] A single recess can be shaped as being between 50% and 97% of the height of the trailing edge BF, particularly between 75% and 95%, taking into account the radial dimension relative to the main axis Z that corresponds to the trailing edge BF extending between the outer annular platform region 52 and the inner annular platform region 54 downstream of the groove of the guide vane assembly 34.

[0068] Figure 10 and Figure 11 A second embodiment according to the present invention is shown. Figure 10 A schematic perspective view of the trailing edge BF of the blade 42 of the guide vane assembly 34 in the region 40 and the region 40 of the guide vane assembly 34.

[0069] More specifically, Figure 10 and Figure 11 A second embodiment according to the invention is shown, wherein the trailing edge BF includes at least one recess 60b having a triangular profile.

[0070] The trailing edge BF may include a plurality of recesses 60b having a triangular profile that extends continuously along the trailing edge BF.

[0071] exist Figure 10 and Figure 11 In the example shown, the trailing edge BF includes two recesses 60b with triangular profiles. Alternatively, not shown, the trailing edge BF may include more than two recesses 60b with triangular profiles. Generally, as described above, the number of recesses 60b can vary depending on the desired mechanical properties.

[0072] Alternatively, the trailing edge BF may include at least two recesses 60b, each with a different triangular profile. Therefore, the first recess 60b may have a different triangular profile than the second recess 60b. In other words, the shape of the triangular profile can differ from one recess 60b to the other.

[0073] In the case of the recess 60b with a triangular profile according to the second embodiment, the bottom 62 of the recess 60b may be defined by a depth p, the dimension of which is less than or equal to the height h of the recess 60b.

[0074] The height h represents the opening of the recess 60b. The height h is preferably oriented radially relative to the main axis Z. Furthermore, the depth p is a dimension extending perpendicular to the height h.

[0075] Alternatively, not shown, the dimension of the depth p is greater than or equal to the height h of the recess 60b with a triangular profile according to the second embodiment.

[0076] In a particular case where the trailing edge BF includes multiple consecutive recesses 60b, the triangular profile has a hollow portion 64 and at least one top 66. The hollow portion 64 may be incorporated into the bottom 62. The top 66 may be located in the plane M.

[0077] Alternatively, the top 66 can be offset relative to the plane M. In this case, the top 66 is located at a distance from the plane M, the value of which (except zero) is strictly less than the depth p of the bottom 62.

[0078] In this embodiment, the recesses 60b with triangular contours can engage such that the top 66 can form a point. Alternatively, the recesses 60b with triangular contours can be spaced apart such that the top 66 can form a plane.

[0079] Figure 12 and Figure 13 A third embodiment according to the present invention is shown. Figure 10 A schematic perspective view of the trailing edge BF of the blade 42 of the guide vane assembly 34 in the region 40 and the region 40 of the guide vane assembly 34.

[0080] More specifically, Figure 12 and Figure 13 A third embodiment according to the invention is shown, wherein the trailing edge BF includes at least one recess 60b having a concave profile.

[0081] The term "concave" refers to a circular shape, such as an arc or an ellipse.

[0082] The contour of the concave portion 60c with a concave profile can also take the form of a "spline" (referring to the shape of a function defined by a polynomial piecewise segment).

[0083] The trailing edge BF may include a plurality of recesses 60b with concave profiles that extend continuously along the trailing edge BF.

[0084] exist Figure 12 and Figure 13In the example shown, the trailing edge BF has three recesses 60c with concave profiles. Alternatively, not shown, the trailing edge BF may include fewer or more than three recesses 60c with concave profiles. Generally, as described above, the number of recesses 60c can vary depending on the desired mechanical properties.

[0085] Alternatively, the trailing edge BF may include at least two recesses 60c, each with a different concave profile. Therefore, the first recess 60c may have a different concave profile than the second recess 60c. In other words, the shape of the concave profile can differ from one recess 60c to the other.

[0086] In the case of a recess 60c having a concave profile according to the third embodiment, the bottom 62 of the recess 60c may be defined by a depth p, the dimension of which is greater than or equal to the height h of the recess 60c.

[0087] The height h represents the opening of the recess 60c. The height h is preferably oriented radially relative to the main axis Z. Furthermore, the depth p is a dimension extending perpendicular to the height h.

[0088] Alternatively, not shown, the dimension of the depth p is less than or equal to the height h of the recess 60c having a concave profile according to the third embodiment.

[0089] In a particular case where the trailing edge BF includes multiple consecutive recesses 60c, the concave profile has a hollow portion 64 and at least one top 66. The hollow portion 64 may be incorporated into the bottom 62. The top 66 may be located in the plane M.

[0090] Alternatively, vertex 66 can be offset relative to plane M. In this case, vertex 66 is located at a distance from plane M that is strictly less than the depth p of bottom 62 (except for zero).

[0091] In this embodiment, the recesses 60c with concave profiles can be spaced apart so that the apex 66 can form a plane. Alternatively, the concave profile recesses 60c can be joined together so that the top 66 can form a sharp tip.

[0092] As described above, the guide vane assembly 34 reduces the deformation of the trailing edge BF of the vane 42 in region 40 of the guide vane assembly 34. The presence of one or more recesses 60a, 60b, 60c makes it possible to reduce the stiffness of the trailing edge BF, in other words, it makes it possible to reduce the force applied to the vane 42.

[0093] Another advantage of the present invention is that the blades 42 of the guide vane assembly 34 are able to resist the thermomechanical and mechanical stresses experienced by the platform and blades 42 of the guide vane assembly 34.

[0094] Another advantage of this invention is that it simplifies design. The guide vane assembly 34 does not require modification of the aerodynamic profile of the vane 42, which ensures component performance while deflecting the load on the component to limit the load on the suction wall. The guide vane assembly 34 does not require modification of the aerodynamic profile of the vane 42, which ensures component performance while deflecting the load on the component to limit the load on the suction wall.

Claims

1. A guide vane assembly (34), particularly for a turbine (20) of an aircraft turbine engine (10), said guide vane assembly comprising a plurality of sections (40) arranged side-by-side circumferentially around a main axis (Z), each section (40) comprising -Outer ring platform section (52). -Inner ring platform section (54), and - At least one hollow impeller (42), the at least one hollow impeller having a suction wall (58) and a pressure wall (56) connected to each other by a leading edge (BA) and a trailing edge (BF), the leading edge and the trailing edge extending between the outer annular platform region (52) and the inner annular platform region (54), respectively, the at least one impeller including a cooling groove (50) arranged near the trailing edge, the cooling groove allowing cooling airflow to flow from the interior of the impeller to the exterior. Its features are, The trailing edge (BF) of the blade (42) includes at least one recess (60a, 60b, 60c) including a bottom (62) offset relative to a plane (M) that passes through the trailing edge (BF) of the blade (42) of the guide blade assembly (34) and is perpendicular to the main axis (Z). The depth of the bottom (62) of the recess is between 10% and 100% of the region (A) of the trailing edge (BF) located downstream of the groove (50) of the blade (42).

2. The guide vane assembly (34) according to claim 1, wherein, The recess (60a) has a generally straight profile, and in particular has a bottom (62) that is generally parallel to the plane (M).

3. The guide vane assembly (34) according to claim 1, wherein, The recess (60b) has a triangular profile.

4. The guide vane assembly (34) according to claim 1, wherein, The recess (60c) has a concave profile.

5. The guide vane assembly (34) according to any one of the preceding claims, wherein, The bottom (62) of the recess (60b, 60c) is defined by a depth (p) that is greater than the height (h) of the recess (60b, 60c).

6. The guide vane assembly (34) according to any one of claims 1 to 4, wherein, The bottom (62) of the recess (60b, 60c) is defined by a depth (p), the dimension of which is smaller than the height (h) of the recess (60b, 60c).

7. The guide vane assembly (34) according to any one of the preceding claims, wherein, The trailing edge (BF) includes a plurality of recesses (60a, 60b, 60c) arranged along the trailing edge (BF).

8. The guide vane assembly (34) according to any one of claims 1 to 6, wherein, The trailing edge (BF) includes a single recess (60a, 60b, 60c) along the trailing edge (BF).

9. The guide vane assembly (34) according to claim 8, wherein, The single recess (60a) is located between 50% and 97% of the radial dimension of the trailing edge (BF) extending between the outer annular platform region (52) and the inner annular platform region (54), particularly between 75% and 95%.

10. The guide vane assembly (34) according to any one of the preceding claims, wherein, The depth of the recesses (60a, 60b, 60c) is between 0.6 mm and 6 mm, particularly between 3 mm and 6 mm, and especially between 4.5 mm and 5.7 mm.

11. The guide vane assembly (34) according to any one of claims 1 to 8, wherein, The depth of the bottom (62) of the recess is between 50% and 95%, particularly between 75% and 95%, of the region (A) of the trailing edge (BF) located downstream of the groove of the blade (42).

12. A turbine (20) for use in an aircraft turbine engine, characterized in that, The turbine includes at least one guide vane assembly according to any one of the preceding claims.

13. An aircraft turbine engine (10), characterized in that, The aircraft turbine engine includes at least one turbine guide vane assembly (34) according to any one of claims 1 to 11 and / or a turbine (20) according to claim 12.

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