Gasket for turbine rotor and related turbine and rotor assembly

Abstract

The invention relates to a gasket (2) for a moving blade (3) intended to have a blade root (30) mounted in a slot (90) leading to the periphery of a turbine rotor disc (1) having a longitudinal axis (X), the gasket (2) comprising a base (20) configured to be mounted in the slot (90) and configured to be mounted in contact with the blade root (30). According to the invention, the base (20) has at least one portion, referred to as an elastic portion, which is corrugated such that the elastic portion exerts an outward radial force on the blade root (30) when the blade root (30) rests on the base (20) mounted in the slot (90). In addition, the base (20) has an upstream longitudinal end and a downstream longitudinal end, the upstream longitudinal end and the downstream longitudinal end being configured to axially abut against one of the disc and the blade root (30).

Description

Technical Field

[0001] The present invention relates to the field of turbomachines, such as turboprop engines or turbojets.

[0002] More particularly, the present invention relates to shims for turbine rotors, and rotor assemblies equipped with such turbines. Background Art

[0003] A turbine rotor, such as that implemented in a low-pressure turbine extending about a longitudinal axis, typically comprises one or more rotor disks carrying a plurality of rotor blades on their outer circumference. The rotor blades have a radially inner portion, or blade root, extending from the rotor blades. At their root, the rotor blades are housed in slots that open into the outer circumferential surface of the rotor disk.

[0004] Positioning of the blade roots in the slots of the rotor disk is performed by translation along an axis substantially longitudinal to the axis of the turbine.

[0005] To facilitate the installation of the moving blades on the rotor disk, sufficient clearance must be provided to allow the blade roots to slide easily in the slots. Consequently, once the blade roots are installed in the slots, a relatively large clearance remains, which can be detrimental to the proper operation of the turbine. In practice, the moving blades can move in radial translation, particularly during shutdown, and thus deviate toward the turbine engine axis. This also applies when the turbine is operating but at reduced speed.

[0006] It is therefore known to implement radial retaining means for retaining the moving blades relative to the rotor disk, ensuring maintenance of the radial position of each moving blade in the slot of the rotor disk and preventing the moving blades from falling towards the turbine axis when the turbine is not running or is running at reduced speed.

[0007] Documents FR 2 994 211 and FR 3 072 122 disclose blades of this type that are provided with a shank that limits radial movement of the blade toward the engine's axis of rotation or toward the exterior when the blade is housed in a slot of a rotor disk. Such a shank can particularly comprise a wall that abuts against a tooth formed between two consecutive slots, thereby achieving radial retention of the blade's position.

[0008] However, a disadvantage of such a shank is that it has a shape that proves to be relatively complex to produce and therefore has a substantial impact on the costs and time of producing the moving blade.

[0009] Furthermore, the implementation of such a shank proves to be even more complicated when the moving blade consists at least partially of a ceramic matrix composite material.

[0010] Therefore, there is a need to provide another solution making it possible to ensure the positioning of the moving blades on the rotor disk and, in particular, to ensure the radial retention of the moving blades relative to the rotor disk. Summary of the Invention

[0011] The object of the present invention is to at least partially overcome the above-mentioned disadvantages associated with the prior art.

[0012] To this end, the invention relates to a spacer for a moving blade intended to have a blade root mounted in a slot opening into the periphery of a turbine rotor disk having a longitudinal axis X, the spacer comprising a base configured to be mounted in the slot and in contact with the blade root.

[0013] According to the invention, the base has at least one portion, referred to as an elastic portion, which is corrugated so that when the blade root rests on the base mounted in the slot, it exerts an outward radial force on the blade root. Furthermore, the base has an upstream longitudinal end and a downstream longitudinal end, the upstream and downstream longitudinal ends being configured to abut axially against one of the disk and the blade root.

[0014] The present invention therefore proposes a solution making it possible to at least partially resolve some of the drawbacks of the prior art.

[0015] In particular, by implementing elements on the outside of the moving blades having the function of securing the moving blades relative to the slots of the rotor disk, the geometry of the moving blades is simplified, which makes it possible to save time and material in manufacturing such moving blades.

[0016] Furthermore, the invention enables radial retention of the moving blades, which is applicable to all types of moving blades, regardless of the material from which they are manufactured.

[0017] Finally, the invention enables a solution that is mechanically simple and inexpensive to implement, while at the same time being technically reliable.

[0018] According to a particular aspect of at least one embodiment of the invention, the gasket comprises an upstream tab extending from the upstream longitudinal end of the base and / or a downstream tab extending from the downstream longitudinal end of the base, the upstream tab and the downstream tab forming an axial stop means for stopping the blade relative to the disk.

[0019] This enables a simple and inexpensive implementation of an axial stop in the upstream or downstream direction.

[0020] According to a specific aspect of at least one embodiment of the present invention, the upstream tabs extend in a radial direction of the turbine, and / or the downstream tabs extend in a radial direction of the turbine.

[0021] In other words, according to a specific aspect of at least one embodiment of the present invention, the upstream tab extends in a direction orthogonal to the axial direction of the turbine, and / or the downstream tab extends in a direction orthogonal to the axial direction of the turbine.

[0022] In this way, the axial stop can be held against the rotor disk.

[0023] According to a specific aspect of at least one embodiment of the present invention, the elastic portion extends longitudinally at the center of the base.

[0024] This therefore makes it possible to have radial forces exerted by the spacers on the blade roots which are relatively balanced so that the moving blades do not tilt in the upstream or downstream direction.

[0025] The invention also relates to a rotor assembly for a turbomachine, comprising a rotor disk carrying a plurality of moving blades housed at the blade roots in slots opening into the periphery of the disk, the assembly further comprising a plurality of shims according to one of the preceding embodiments, each of the shims being arranged between the blade root and the bottom of one of the slots.

[0026] According to a particular aspect of at least one embodiment of the present invention, the base has a length measured along the longitudinal axis that is substantially equal to a length of each of the slots measured along the longitudinal axis.

[0027] According to a particular aspect of at least one embodiment of the present invention, the base portion has a single corrugated portion extending from an upstream longitudinal end to a downstream longitudinal end and contacting the blade root at both the upstream and downstream longitudinal ends of the base portion. Furthermore, the base portion contacts the bottom of the slot at a longitudinal center portion.

[0028] According to a particular aspect of at least one embodiment of the present invention, the rotor assembly further comprises a plurality of foils, each foil being disposed between one of the blade roots and one of the shims.

[0029] According to a particular aspect of at least one embodiment of the present invention, at least one of the blades is at least partially composed of a ceramic matrix composite material.

[0030] The invention also relates to an aircraft turbomachine comprising a rotor assembly according to one of the aforementioned embodiments.

[0031] According to a particular aspect of at least one embodiment, the turbine is a turbojet engine. BRIEF DESCRIPTION OF THE DRAWINGS

[0032] The present invention and the various advantages it offers will be more readily understood from the following description of illustrative and non-limiting embodiments of the invention and the accompanying drawings, in which:

[0033] [ Figure 1 ] is a schematic cross-sectional view of a turbine;

[0034] [ Figure 2 ] is a cross-sectional view of a portion of a turbine rotor assembly;

[0035] [ Figure 3 ]yes Figure 2 Detailed view of area A;

[0036] [ Figure 4 ] is a perspective view of a portion of a turbine rotor assembly; and

[0037] [ Figure 5 ] is another cross-sectional view of a gasket for a turbine rotor according to an embodiment of the present invention. DETAILED DESCRIPTION

[0038] It should be noted that the present invention is applicable to any type of turbomachine, and in particular to aircraft turbomachines.

[0039] This turbine 1 extending along an axis X is for example intended to be mounted on an aircraft (not shown), such as an airplane or a helicopter, for example under a wing, on a wing or at the rear of the fuselage.

[0040] Figure 1 The turbine 1 shown is a direct drive twin-shaft turbofan engine. However, this is not restrictive, as the turbine 1 may not be intended to be installed on an aircraft, but may be another type of turbojet engine, such as a geared turbojet engine, a turboprop engine or an auxiliary power unit (APU).

[0041] Throughout the specification, the axial direction corresponds to the direction of the longitudinal axis X, and the radial direction is a direction perpendicular to and intersecting the longitudinal axis X. Similarly, an axial plane is a plane containing the longitudinal axis X, and a radial plane is a plane perpendicular to the longitudinal axis X.

[0042] Similarly, the adjectives "inner" (or "inner") and "outer" (or "outer") are used with reference to the radial direction, such that an inner portion of an element is radially closer to the longitudinal axis X than an outer portion of the same element. Additionally, unless otherwise indicated, the terms "upstream" and "downstream" are used with reference to the overall gas flow direction through the turbine in operation.

[0043] like Figure 1 As shown, the turbomachine 1 comprises, from upstream to downstream, a fan 10, a compressor section 12, a combustion chamber 14, and a turbine section 16. A longitudinal axis X forms the axis of rotation of at least a portion of the compressor section 12 and the turbine section 16, which are rotatable about the longitudinal axis X relative to a casing 18 of the turbomachine 1.

[0044] In operation, fan 10 draws in an air flow, a portion of which circulates in main flow path 100 and is subsequently compressed in compressor section 12, ignited in combustion chamber 14, and expanded in turbine section 16 before being exhausted from turbine 1. In this manner, turbine 1 generates thrust. Furthermore, this thrust can be utilized, for example, by the aircraft on which turbine 1 is mounted and fastened.

[0045] The compressor and the turbine comprise an alternation of moving disks, known as rotor disks, and fixed disks, known as stator disks, the moving disks being mounted rotatably about the longitudinal axis X of the turbine.

[0046] Each rotor disk assembly comprises a plurality of slots 90 cut into its peripheral surface, which slots 90 open onto the periphery of the disk. Each of these slots is intended to house a moving blade, more specifically a blade root 30, so that each rotor disk carries a plurality of moving blades, each of which is intended to have a blade root mounted in one of the slots opening onto the periphery of the rotor disk.

[0047] In the embodiments described below, preferably, the moving blades are at least partially composed of a ceramic matrix composite material.

[0048] Due to the size difference between each slot and each blade root, in particular in order to facilitate the installation of the blade root in the slot and to achieve the translation of the blade root in the slot, the blade root is not completely radially retained in the slot, and there is still at least one space 8 between the bottom of the slot 90 and the blade root 30.

[0049] Therefore, the assembly also comprises a plurality of shims mounted in the slots 90 and arranged between the blade root 30 and the bottom of one of the slots 90 .

[0050] For example, the spacers may be made at least partially of metal or a metal alloy.

[0051] Now refer to Figures 2 to 5 A first embodiment of the present invention will be described. As shown in the figure, a gasket 2 for a turbine 1 rotor includes a base portion 20. This base portion is intended to be received in a space 8 between a slot 90 cut into the outer peripheral surface of a rotor disk 9 and the root portion 30 of a blade 3 received in the slot 90, and to contact the blade root 30 and the disk.

[0052] In the illustrated embodiment, the base 20 has a length, or axial length, measured along the longitudinal axis that is substantially equal to the axial length of each slot 90 .

[0053] In this manner, implementation of the shims does not disrupt the normal operation of the rotor assembly.

[0054] According to the invention, the base 20 has at least one elastic portion which is corrugated in order to exert an outward radial force on the blade root 30 .

[0055] This radial force exerted by the base on the blade root thus makes it possible to ensure that the moving blade is radially retained in the slot and prevent the moving blade from being radially displaced or even sinking into the slot.

[0056] Therefore, when the moving blades are positioned in the slots, this makes it possible to prevent overlapping or separation of the moving blades.

[0057] Furthermore, and in order to enhance and stabilize the radial retention of the moving blade in the slot, the base 20 has an upstream longitudinal end and a downstream longitudinal end configured to abut axially against one of the disk and the blade root 30 .

[0058] Here, the upstream and downstream longitudinal ends abut against the blade root 30 .

[0059] However, an embodiment may be provided in which the upstream and downstream longitudinal ends abut against the disk, for example at an axial face of the disk or at the bottom of the slot.

[0060] An embodiment may also be provided, wherein the upstream longitudinal end axially abuts one of the slot bottom and the blade root, and the downstream longitudinal end axially abuts the other of the slot bottom and the blade root.

[0061] Here, therefore, the base 20 has an elasticity which enables the gasket to deform during installation in order to subsequently exert a thrust, ie a force outwards relative to the turbine.

[0062] More specifically, in this embodiment, in particular, Figure 2As shown, the elastic portion extends at the center of the base 20 so that the spacer provides a radial force that is relatively balanced and prevents the moving blade from tilting in the upstream or downstream direction.

[0063] However, an embodiment may be provided in which the elastic portion is offset towards the upstream longitudinal end or the downstream longitudinal end in order to compensate for the mass or force distribution of the moving blade.

[0064] Thus, here, the elastic portion corresponds to the central portion, e.g. Figure 3 As shown, the central portion is here in contact with the bottom of the slot 90 , while the base 20 is in contact with the blade root 30 .

[0065] As shown in this embodiment, the gasket 2 also comprises axial stop means in the form of an upstream tab 21 and a downstream tab 22 .

[0066] Here, the upstream tab 21 extends from the upstream end of the base 20 and extends in the radial direction of the turbine, more specifically radially inwardly relative to the turbine.

[0067] Furthermore, the gasket 2 comprises a downstream tab 22 extending from the downstream end of the base 20. Here, the downstream tab 22 extends in the radial direction of the turbine, more specifically radially outward relative to the turbine.

[0068] However, the orientation of the tabs depends inter alia on the implementation of the upstream or downstream flange at the rotor disk, and another embodiment may be provided in which the upstream tabs extend radially outwards and the downstream tabs extend radially inwards relative to the turbine.

[0069] Since the axial length of the base 20 is substantially equal to the axial length of each slot 90 , in this embodiment, the upstream tab 21 and the downstream tab 22 are in contact with the rotor disk 9 .

[0070] Here, these tabs form axial stops for the moving blades in the upstream and downstream directions respectively and prevent the moving blades from being axially offset relative to the slots of the rotor disk.

[0071] If especially in Figure 3 and Figure 4 As shown in , the turbine rotor assembly further comprises a plurality of foils 4 .

[0072] Each of these foils 4 is here arranged between one of the blade roots 30 and one of the shims 2. More specifically, each foil 4 is here mounted in contact with the blade root and forms a layer capable of protecting the blade root from wear and deformation.

[0073] The base therefore exerts a radial force on the blade root via the foil positioned between the two components, the blade root thus being in indirect contact with the shim.

[0074] If especially in Figure 2 As shown in FIG, each foil 4 includes a foil base configured to be in contact with the lower face of the blade root 30, and two lateral fins configured to be in contact with the lateral surfaces of the blade root.

[0075] In other words, each foil 4 has a substantially U-shape, wherein two fins are connected by the foil base.

[0076] It should be noted that Figure 2 As shown, the contact surface between the blade root 30 and the foil 4, the connection corners between each fin and the foil base have a rounded shape.

[0077] As shown in the various figures, the base of the foil is in contact with the base 20 of the shim, so that the base 20 of the shim 2 exerts an outward radial force on the blade root 30 and thus on the base of the foil 4 .

[0078] More specifically, in this embodiment, the resilient portion of the base 20 thus extends longitudinally at the centre of the base 20 , the resilient portion of the base 20 exerting an outward radial force on the central portion of the base of the foil 4 .

[0079] In this embodiment, the upstream and downstream longitudinal ends of the base 20 abut against the upstream and downstream longitudinal ends of the substrate of the foil 4 .

Claims

1. A turbine rotor assembly comprising a turbine (1) rotor disk (9) having a longitudinal axis (X), said rotor disk (9) carrying a plurality of moving blades (3), each having a blade root (30) mounted in a slot (90), said slot (90) opening into the outer periphery of said rotor disk (9), said assembly further comprising a plurality of shims (2), each of said shims (2) comprising a base (20) arranged between said blade root (30) and the bottom of one of said slots (90), said base (20) having at least one portion, referred to as an elastic portion, said elastic portion being corrugated so that when said blade root (30) rests on said base (20) mounted in said slot (90) ), the resilient portion exerts an outward radial force on the blade root (30), the base (20) having an upstream longitudinal end and a downstream longitudinal end, the upstream longitudinal end and the downstream longitudinal end being configured to abut axially against one of the disk and the blade root (30), the base (20) having a length measured along the longitudinal axis that is substantially equal to the length of each of the slots (90) measured along the longitudinal axis, the base (20) having a single corrugated portion extending from the upstream longitudinal end to the downstream longitudinal end and contacting the blade root (30) at the upstream longitudinal end and the downstream longitudinal end of the base (20), and wherein, The base (20) also contacts the bottom of the slot (90) at a longitudinal center portion of the base (20).

2. The assembly (2) according to claim 1, characterized in that The spacer comprises an upstream tab (21) extending from the upstream longitudinal end of the base (20) and / or a downstream tab (22) extending from the downstream longitudinal end of the base (20), the upstream tab (21) and the downstream tab (22) forming axial stop means for stopping the blade relative to the disk.

3. Assembly (2) according to the preceding claim, characterized in that The upstream tab (21) extends in the radial direction of the turbine, and / or the downstream tab (22) extends in the radial direction of the turbine.

4. Assembly (2) according to any one of the preceding claims, characterized in that The elastic portion extends longitudinally at the center of the base (20).

5. A turbine rotor assembly according to any one of the preceding claims, characterised in that The turbine rotor assembly further comprises a plurality of foils (4), each foil being arranged between one of the blade roots (30) and one of the shims (2).

6. A turbine rotor assembly according to any one of the preceding claims, characterised in that At least one of the blades (3) is at least partially composed of a ceramic matrix composite material.

7. A turbomachine comprising the rotor assembly according to any one of claims 1 to 6.

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