A turbine nozzle segment for an aircraft turbine engine having improved robustness

Abstract

The present invention relates to a turbine nozzle segment (40) for an aircraft turbine engine, comprising a first platform (44a), a plurality of blades (42) projecting from the first platform along a first direction (S1) in the radial direction (R) of the segment, and a first end member (50a) projecting from the first platform (44a) along a second direction (S2) opposite to the first direction (S1) in the radial direction, the first end member (50a) extending along the circumferential direction (C) of the segment. According to the invention, the segment includes at least one first mechanical unloading groove (66a) formed through the first end member (50a), each first groove (66a) having a radial groove bottom (68a) and a radial slot (70a) spaced apart from the radial groove bottom (68a) in the second direction (S2).

Description

Technical Field

[0001] This invention relates to an aircraft turbine engine, preferably a turbojet engine or a turboprop engine.

[0002] More specifically, the present invention relates to the design of turbine nozzles for aircraft turbine engines, such nozzles being known, for example, from document FR 2 825 748 A1. Background Technology

[0003] Turbine engine nozzles for aircraft typically consist of multiple nozzle segments arranged end-to-end in the circumferential direction. Each segment usually comprises two platforms with blades arranged between them, circumferentially spaced apart from each other within the segment. Each of the two platforms has a radial surface that defines the main passageway of the turbine engine, enabling the main stream to be guided through the turbine and originating from the combustion chamber.

[0004] Furthermore, on the side radially opposite the radial surface defining the main channel, each platform carries one or more protruding end members. Each protruding end member extends circumferentially, typically covering the entire length of the sector. It may be, for example, a tab for attaching the nozzle sector to the turbine housing, or a support leg, for example, for supporting a wearable sealing member.

[0005] Within the nozzle sector, each platform and its protruding end members together form a radial end structure, which is subjected to high thermal stress. In fact, on the side of the radial surface defining the main channel, this surface is directly exposed to the high temperatures from the main channel, while on the opposite side of the radial end structure, the temperature generally appears much lower. This latter finding is explained not only by the radial distance of the main channel but also by the possible accessibility of one or more cooling air chambers.

[0006] As a result, each radial end structure exhibits a high temperature gradient in the radial direction, which can have a detrimental effect on the nozzle sector, known as the "bimetallic effect." This is characterized by differential thermal expansion between the two sides of each radial end structure. On the side defining the radial surface of the main channel, the elongation of the structure is more pronounced than on the opposite side (referred to as the cold side). The bimetallic effect then manifests itself as high mechanical stress within the radial end structure, which tends to deform by bending. Since the blades in the nozzle sector are connected to the two radial end structures, they also experience high mechanical stress, resisting deformation of these radial end structures. Due to the deformable nature of these two structures, the blades located at the circumferential ends of the sector primarily bear tensile forces, while the blades located more centrally in the sector essentially bear compressive forces between the two radial end structures.

[0007] These stresses on the blades in the nozzle sector significantly affect the sector's service life, especially due to the risk of cracks forming on the blades located at the circumferential ends.

[0008] To address this issue, materials with greater resistance to fatigue and / or creep can be used to enhance the mechanical robustness of the nozzle segments. However, this solution is particularly expensive. Another solution involves manufacturing nozzle segments with a reduced circumferential amplitude to limit the aforementioned bimetallic effect. However, this solution remains detrimental in terms of nozzle assembly time and cost, and also increases the leakage cross-section between a greater number of segments.

[0009] Therefore, there is a need to optimize turbine nozzle sector sections to enhance their mechanical robustness while controlling assembly and manufacturing costs. Summary of the Invention

[0010] To meet the above requirements, the present invention first relates to a turbine nozzle sector of an aircraft turbine engine, including a first platform, a plurality of blades protruding from the first platform along the radial direction of the sector in a first direction, and a first end member protruding from the first platform along the radial direction in a second direction opposite to the first direction, the first end member extending along the circumferential direction of the sector.

[0011] According to the present invention, the nozzle segment includes at least one first mechanical unloading groove formed through the first end member, each first groove having a radial groove bottom and a radial groove opening spaced apart from the radial groove bottom in a second direction.

[0012] Due to the presence of one or more mechanical unloading grooves, the first end member is divided into multiple circumferential sections, thereby imparting a degree of deformation flexibility. This flexibility on the coldest side is used to accommodate thermal expansion on the hottest side of the first platform. Advantageously, this allows for a significant reduction in the aforementioned bimetallic effect. Consequently, the deformation resisted by the sector blades is reduced, resulting in these blades experiencing less mechanical stress. As a result, they can have an extended service life and require less maintenance / replacement of the nozzle sector.

[0013] In other words, the presence of the mechanical unloading grooves allows for the limitation of bending variations between the hot and cold portions of the radial end structure formed by the first platform and each associated first end member. This results in lower stress on the blades of the nozzle sector, advantageously improving their mechanical strength.

[0014] Furthermore, this solution enables further standardization of the behavior and stress of all blades within the same nozzle segment. Moreover, the number of blades per segment can be increased, advantageously reducing the number of segments assembled end-to-end in the circumferential direction. This results in reduced assembly time and cost, and improved aerodynamic performance due to a decrease in the number of leakage sections per ring. Additionally, when the end members are designed to support the turbine's casing elements, contact forces are better distributed across the casing elements, which has the beneficial effect of reducing the risk of clogs and leaks at the interfaces. Due to the more uniform force distribution across the casing elements, wear on the casing elements is reduced, and the risk of segment tipping caused by such wear is also decreased. By reducing the segment swaying effect due to wear on the casing elements, the nozzle segment blades maintain optimal positioning relative to the main flow, leading to improved turbine performance.

[0015] Finally, this solution allows for the use of less expensive materials to manufacture the nozzle sector, and also provides a mass gain. Therefore, this invention is the result of technical research aimed at significantly improving aircraft performance and, in this respect, contributing to reducing the environmental impact (decarbonization) of these aircraft.

[0016] Preferably, the present invention provides at least one of the following optional technical features, implemented individually or in combination.

[0017] Preferably, each first groove is axially open on both sides of the first end member.

[0018] Preferably, the radial depth of each first groove is greater than at least 20% of the maximum radial length of the first end member, more preferably greater than at least 50% or even 70% of the maximum radial length, while preferably remaining strictly less than the maximum radial length.

[0019] Preferably, the first end member is one of the following elements:

[0020] • Protrusions used for attaching nozzle segments;

[0021] • Support leg, preferably supporting sealing member. Alternatively, the support leg can be replaced by a leg connected to another structural mechanical component (e.g., flange clamp).

[0022] Preferably, the first end member includes a web protruding from the first platform and a flange supported by the web and spaced at a distance from the first platform, each first groove extending radially through the entire flange and preferably through at least a portion of the web.

[0023] Preferably, the circumferential width of the first groove narrows from the radial groove bottom to the radial groove opening.

[0024] Preferably, the nozzle segment includes at least one sealing lug that seals at least a portion of the first unloading groove. This advantageously reduces leakage through the unloading groove.

[0025] Preferably, the nozzle segment includes a second platform, blades projecting from the second platform in a second radial direction, and a second end member projecting from the second platform in a first direction, the second end member extending in the circumferential direction of the segment. Furthermore, the segment includes at least one second mechanical unloading groove formed through the second end member, each second groove having a radial groove bottom and a radial opening spaced apart from the radial groove bottom in the first direction. In this regard, it should be clearly noted that all the optional features previously described for the first end member and its first unloading groove are applicable to the second end member and its second unloading groove.

[0026] The present invention also relates to a turbine nozzle for an aircraft turbine engine, comprising at least one such nozzle segment, and preferably comprising a plurality of such segments arranged end-to-end in the circumferential direction.

[0027] Finally, the object of the present invention is an aircraft turbine engine comprising at least one such turbine nozzle for a low-pressure or high-pressure turbine.

[0028] Other advantages and features of the invention will become apparent from the following non-limiting detailed description. Attached Figure Description

[0029] The following detailed description refers to the accompanying drawings, in which:

[0030] Figure 1 This is a schematic axial cross-sectional view of a turbojet engine according to the present invention;

[0031] Figure 2 This is an axial half-section view of a portion of the turbine stage of the turbine engine shown in the previous figure, which includes a nozzle in the form of a preferred embodiment of the present invention;

[0032] Figure 3 It is formed Figure 2 An enlarged perspective view of a nozzle sector, representing a portion of the nozzle shown.

[0033] Figure 4 yes Figure 3 The cross-sectional view of a portion of the sector shown in the diagram, more specifically, shows the first mechanical unloading groove;

[0034] Figure 5 yes Figure 3 The cross-sectional view of a portion of the sector shown in the diagram, more specifically illustrating the second mechanical unloading groove;

[0035] Figure 6 yes Figure 3 The perspective view of a portion of the sector shown in the figure, more specifically, shows the first mechanical unloading groove sealed by the sealing lug;

[0036] Figure 7yes Figure 6 The cross-sectional view of the sector shown shows the cross-sectional plane passing through the sealing lug;

[0037] Figure 8 yes Figure 3 The perspective view of a portion of the sector shown in the figure, more specifically, shows another first mechanical unloading groove sealed by a sealing lug;

[0038] Figure 9 yes Figure 8 The cross-sectional view of the sector shown in the figure shows the cross-sectional plane passing through the sealing lug. Detailed Implementation

[0039] The accompanying drawings include reference frames L, R, and C, which define longitudinal, radial, and circumferential directions that are orthogonal to each other, corresponding to those directions of the aircraft turbojet engine 1 according to the invention.

[0040] Figure 1 An aircraft turbojet engine 1 is shown, preferably with a dual-rotor, turbofan type design. However, other types of turbojet engines are also possible, such as turboprop engines, in addition to turbojet engines.

[0041] In the following text, the terms "upstream" and "downstream" are defined relative to the primary direction D1 through which gas flows through the turbojet engine 1 when it is operating in direct thrust mode. Direction D1 is parallel to the longitudinal direction L and also parallel to the longitudinal axis A1 of the turbojet engine around which its various components extend. In this case, from upstream to downstream of the turbojet engine 1 are the fan 4, the low-pressure compressor 5, the high-pressure compressor 6, the combustion chamber 7, the high-pressure turbine 8, and the low-pressure turbine 9.

[0042] During operation of the turbojet engine 1, airflow 10 enters the turbojet engine 1 through air inlet 3, passes through fan 4, and then splits into a central main flow 10A and a secondary flow 10B. The main flow 10A flows through the main passage 11A, which is a gas passage that passes through compressors 5 and 6, combustion chamber 7, and turbines 8 and 9. The secondary flow 10B then flows into the secondary passage 11B surrounding the main passage 11A, also known as the primary passage or aerodynamic flow path.

[0043] In a manner known per se, a turbine, such as a high-pressure turbine 8 or a low-pressure turbine 9, comprises one or more stages 12. Each stage 12 includes a stator assembly 12A, which includes a nozzle 29, also referred to as a bladed stator ring, and a movable wheel 12B arranged directly downstream of the assembly 12A. One such stage 12 is shown in... Figure 2 middle.

[0044] Now for reference Figures 2 to 5The turbine nozzle 29 will be described here, which is preferably integrated into the high-pressure turbine 8 and centered around axis A1. The annular nozzle 29 is passed through its center by one or more drive shafts 30, which are also centered around axis A1.

[0045] The nozzle 29 is made of a plurality of nozzle segments 40, which are arranged end-to-end in the circumferential direction C. Each segment 40 corresponds to a corner sector of the ring and includes two platforms with blades 42 arranged between the platforms, circumferentially spaced from each other within the segment. A segment 40 may, for example, include between two and ten blades 42. Only one of the segments 40 will be described below, but it should be noted that other segments 40 of the ring may have the same or similar design.

[0046] The two platforms of sector 40 correspond to the radially outer platform, referred to as the first platform 44a, and the radially inner platform, referred to as the second platform 44b. An opposite arrangement is also possible without departing from the scope of the invention.

[0047] Therefore, blades 42 protrude from the first platform 44a along the radial direction R of the sector in a first direction S1, while these same blades 42 protrude from the second platform 44b along the radial direction R in a second direction S2 opposite to the first direction S1. Here, the first direction S1 extends radially inward from the outside in the direction R, while the second direction S2 is the opposite.

[0048] The first platform 44a includes two surfaces opposite each other in the radial direction R, namely a first radially defining surface 46a of the main channel 11A oriented in direction S1, and a first outer surface 48a of the sector oriented in direction S2. Similarly, the second platform 44b includes two surfaces opposite each other in the radial direction R, namely a second radially defining surface 46b of the main channel 11A oriented in direction S2, and a second outer surface 48b of the sector oriented in direction S1.

[0049] The nozzle sector 40 includes at least one first end member 50a at its radial end, which protrudes from a first outer surface 48a along a second direction S2 in the direction R. Each first end member 50a preferably extends along the circumferential direction C of the sector 40, more preferably along its sector. Here, these are two tabs 50a for attaching the nozzle sector 40 to one or more turbine casing elements 52, such as... Figure 2As shown. These attachment tabs 50a do indeed conventionally engage with associated hooks provided on the turbine casing element 52. The upstream attachment tab 50a has an upstream hook shape, while another downstream attachment tab 50a has a downstream hook shape. Between these tabs 50a, for example, a cooling air cavity 54 is defined, which is radially inwardly defined by a first outer surface 48a of a first platform 44a. It should be noted that this first platform 44a, together with the two attachment tabs 50a, forms a first radial end structure, which is subjected to differential thermal expansion for the reasons described above. In this respect, it should be noted that the attachment tabs 50a may not be hook-shaped, but have a simpler shape for pressing their ends against the casing.

[0050] Similarly, the nozzle segment 40 includes at least one second end member 50b at its radial end, which protrudes from the second outer surface 48b along the first direction S1. Each second end member 50b preferably extends along the circumferential direction C of the segment 40, more preferably along the segment. Here, it is a leg 50b supporting the wearable seal 56 of the labyrinth seal, as... Figure 2 As shown, a support leg 50b extends radially through the free space defined between two axial spoilers, which are respectively mounted on the wheels 12B of two movable turbines located on either side of the nozzle 29. Furthermore, at least a portion of the leg 50b is received in a cooling air chamber 60 and / or participates in defining one or more chambers of this type. Alternatively, the support leg 50b may be replaced by a leg connected to another structural component, for example, by being inserted into a flange clamp.

[0051] It should be noted that the second platform 44b together with the support leg 50b forms the second radial end structure, which is also subject to differential thermal expansion for the reasons mentioned above.

[0052] like Figure 4 As best seen, each attachment tab 50a includes a web 62a projecting from the first outer surface 48a of the first platform 44a along direction S2. The web 62a is radially oriented, or inclined relative to the radial direction. Furthermore, each attachment tab 50a includes a flange 64a supported by the web 62a and spaced at a distance from the platform, and has an axial or substantially axial orientation. The web 62a and flange 64a together form a generally L-shaped cross-section member, and the angle between the web and flange can be approximately 90° or greater, for example, approximately 100° to 150°, as in the case of the upstream member 50a.

[0053] Similarly, such as Figure 5As best seen in the image, the support leg 50b includes a web 62b projecting from the second outer surface 48b of the second platform 44b along direction S1. The web 62b has a radial or substantially radial orientation. Furthermore, the support leg 50b may also include a flange 64b supported by the web 62b and spaced at a distance from the platform, and having an axial or substantially axial orientation. The web 62b and flange 64b together form a generally T-shaped cross-section member, and the angle between the web and the flange may be approximately 90° or different values.

[0054] In order to avoid or greatly limit the aforementioned bimetallic effect on the two radial end structures of the nozzle sector 40, one of the distinguishing features of the present invention is the manufacture of one or more mechanical unloading grooves.

[0055] More precisely, such as Figure 3 As can be seen, each attachment tab 50a is passed through at least one first mechanical unloading groove 66a, preferably through a plurality of grooves 66a spaced apart from each other in the circumferential direction. Each first groove 66a has a radial groove bottom 68a and a radial groove opening 70a spaced apart from the radial groove bottom 68a in the second direction S2.

[0056] Each first groove 66a is a through groove because it is axially open on both sides of the attachment tab 50a on which the groove is formed. Here, each groove 66a radially passes through the entire flange 64a, and preferably at least through at least a portion of the web 62a, with the radial bottom 68a of the groove thus located on the web 62a.

[0057] The first groove 66a disposed on one of the two attachment tabs 50a can be offset in the circumferential direction relative to the first groove 66a disposed on the other of the two attachment tabs.

[0058] Similarly, the support leg 50b can be passed through at least one second mechanical unloading groove 66b, preferably through a plurality of grooves 66b spaced apart from each other in the circumferential direction. Each second groove 66b has a radial groove bottom 68b and a radial slot 70b spaced apart from the radial groove bottom 68b in the first direction S1.

[0059] Each second groove 66b is a through groove because it is axially open on both sides of the support leg 50b that forms the groove. Here, each groove 66b radially passes through the entire flange 64b, and preferably at least through a portion of the web 62b, with the radial bottom 68b of the groove thus located on the web 62b.

[0060] Regardless of which slot 66a, 66b is involved, its radial depth Pr is greater than at least 20% of the maximum radial length Lrmax of the associated end member 50a, 50b, more preferably greater than at least 50% or even 70% of the maximum radial length Lrmax. However, the radial slot depth Pr, measured in the radial direction R between the slot bottom 68a, 68b and the point in the radial direction R furthest from the bottom of the slot 70a, 70b, preferably remains strictly less than the maximum radial length Lrmax, which is also measured in the direction R from the outer platform surface 48a, 48b.

[0061] It should be noted that the radial depth Pr of the grooves 66a, 66b is preferably greater than or equal to the thickness of the flanges 64a, 64b, when such flanges are provided in a generally axial or substantially axial orientation. In other words, each groove 66a, 66b extends deep enough to at least pass through the flanges 64a, 64b, and preferably also through at least a portion of the webs 62a, 62b.

[0062] The number of slots 66a and 66b in each component 50a and 50b can vary, for example, from two to ten, depending on the circumferential length of the sector 40.

[0063] The shapes of slots 66a and 66b can be adjusted according to requirements and constraints. For example, as... Figure 3 As shown above, the circumferential width Lc of the first groove 66a gradually narrows from the radial groove bottom 68a to the radial groove opening 70a. This principle can be applied to some or all of the mechanical unloading grooves 66a and 66b.

[0064] As previously mentioned, the presence of mechanical unloading grooves 66a and 66b particularly enables them to limit the bending variation between the cold and hot portions of each of the two radial end structures formed by platforms 44a and 44b and their associated end members 50a and 50b. This results in lower stress on the blades of the nozzle sector, advantageously improved mechanical strength, and extended service life. In fact, due to the presence of the mechanical unloading grooves, each end member 50a and 50b is divided into multiple circumferential portions, thus imparting them a degree of deformation flexibility. This flexibility on the colder side is used to accommodate the thermal expansion of the hotter side of platforms 44a and 44b. This significantly reduces the aforementioned bimetallic effect. Furthermore, the solution according to the invention allows for further standardization of the behavior and stress of all blades 42 of the same sector 40. Therefore, the number of blades 42 in each nozzle sector 40 can be increased, which reduces the number of sectors assembled end-to-end in the circumferential direction C. This reduces assembly time and cost, while also improving aerodynamic performance due to the reduced number of leakage sections per ring.

[0065] Furthermore, since the first end member 50a is supported on the turbine casing element 52, the contact forces are better distributed on these elements 52, advantageously leading to a reduction in the risk of clogging and leakage at the interface with the nozzle segment 40. Due to the more uniform force distribution on the casing element 52, its wear is reduced, and the risk of segment rollover caused by such wear is also lowered. By reducing the tilting effect of the nozzle segment 40, the blades 42 of the nozzle segment 40 maintain an optimal position relative to the main flow, meaning improved turbine performance.

[0066] Now for reference Figure 6 and Figure 7 The diagram shows an upstream attachment tab 50a, wherein one or more of the unloading grooves 66a are at least partially sealed by sealing lugs to prevent leakage through these grooves.

[0067] In the illustrated embodiment, each groove 66a is provided with two sealing lugs 72, 74. Lug 72 is used to seal the portion of groove 66a passing through the web 62a, while the other lug 74 is used to seal the portion of groove 66a passing through the flange 64a. Each of these lugs 72, 74 is guided and held in a track-shaped groove 76 at its two opposite circumferential ends, the track-shaped groove 76 being disposed in opposite circumferential sides of groove 66a.

[0068] Here, for example, lug 72 first slides from the outside of flange 64a into its associated track 76, and then, once it reaches the bottom of the track, another lug 74 slides from the outside of web 62a into its associated track 76, also until the bottom of its track. In the final position, lug 74 extends sufficiently to cover lug 72, thus preventing it from sliding off its track 76. Therefore, lug 72 does not require other means of holding it in place, while lug 74 can be held in its track 76 by one or more weld points 78, one of which is schematically shown in [illustration missing]. Figure 7 middle.

[0069] Lugs 72 and 74 can be made of any material chosen so that the lugs do not affect the flexibility of the radial end structure once installed. They have a thinner thickness than their associated members 50a and 50b and are used solely for sealing purposes to limit leakage through the grooves 66a and 66b.

[0070] Figure 8 and Figure 9 Another embodiment is shown on the downstream attachment tab 50a, which has a similar principle to the two lugs 72, 74. The lug 74 is no longer held in its track 76 by a weld point, but by a stop member 80 outside the sector 40, such as formed by an adjacent sleeve or any other stator element.

[0071] Other technical solutions for retaining the lugs can be considered, such as folding the lugs. Furthermore, it should be noted that the principle of sealing the lugs can also be applied to the groove 66b of the support leg 50b without departing from the scope of the invention.

[0072] Of course, those skilled in the art can make various modifications to the invention just described, which are provided as non-limiting examples and within the scope of the appended claims.

Claims

1. A turbine nozzle sector (40) for an aircraft turbine engine, comprising a first platform (44a), a plurality of blades (42) projecting from the first platform along a first direction (S1) in the radial direction (R) of the sector, and a first end member (50a) projecting from the first platform (44a) along a second direction (S2) opposite to the first direction (S1) in the radial direction, the first end member (50a) extending along the circumferential direction (C) of the sector, characterized in that, It includes at least one first mechanical unloading groove (66a) formed through the first end member (50a), each first groove (66a) having a radial groove bottom (68a) and a radial groove opening (70a) spaced apart from the radial groove bottom (68a) in a second direction (S2).

2. The nozzle segment according to claim 1, characterized in that, Each first groove (66a) is axially open on both sides of the first end member (50a).

3. The nozzle segment according to claim 1 or 2, characterized in that, The radial depth (Pr) of each first groove (66a) is at least 20% greater than the maximum radial length (Lrmax) of the first end member (50a), more preferably at least 50% or even 70% greater than the maximum radial length (Lrmax), while preferably remaining strictly less than the maximum radial length (Lrmax).

4. The nozzle sector according to any one of the preceding claims, characterized in that, The first end member is one of the following elements: - A tab (50a) for attaching nozzle sector segments; - Support leg (50b), preferably supporting sealing member (56).

5. The nozzle segment according to any one of the preceding claims, characterized in that, The first end member (50a) includes a web (62a) protruding from the first platform (44a) and a flange (64a) supported by the web and spaced at a distance from the first platform (44a), each first groove (66a) radially passing through the entire flange (64a) and preferably passing through at least a portion of the web (62a).

6. The nozzle sector according to any one of the preceding claims, characterized in that, The circumferential width (Lr) of the first groove (66a) gradually narrows from the radial groove bottom (68a) to the radial groove opening (70a).

7. The nozzle sector according to any one of the preceding claims, characterized in that, It includes at least one sealing lug (72, 74) that seals at least a portion of the first unloading groove (66a).

8. The nozzle sector according to any one of the preceding claims, characterized in that, It includes a second platform (44b), a blade (42) protruding from the second platform in a second direction (S2) along the radial direction (R), and a second end member (50b) protruding from the second platform (44b) in a first direction (S1), the second end member (50b) extending along the circumferential direction (C) of the fan segment, and is characterized in that it further includes at least one second mechanical unloading groove (66b) formed through the second end member (50b), each second groove (66b) having a radial groove bottom (68b) and a radial slot (70b) spaced apart from the radial groove bottom (68b) in the first direction (S1).

9. A turbine nozzle (29) for an aircraft turbine engine, comprising at least one nozzle sector (40) according to any one of the preceding claims.

10. An aircraft turbine engine (1) comprising at least one turbine nozzle (29) according to the preceding claim.

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