Blade assembly for turbine, turbine for turbine, and turbine

Abstract

A blade assembly (40) for a turbomachine includes a plurality of blades (42) distributed about an axis (28), each blade including an airfoil (44) and a platform (46) formed at a free end of the blade (42). The platform (46) of each blade comprises a friction member (70) and an opening (72) that is elongated in a circumferential direction (C) relative to the axis (28). The friction members (70) are engaged by openings (72) formed in the platforms (46) of circumferentially adjacent blades (42) within the blade assembly (40), the openings having gaps allowing the friction members (70) to move within the openings (72) at least in the circumferential direction (C). During operation, mutual separation of the platforms (46) in the circumferential direction due to vibration of the blade causes the friction member (70) to rub against the edge of the opening (72), thereby dissipating energy.

Description

Technical Field

[0001] The present invention relates to the general field of turbomachines, in particular blade assemblies in turbomachines intended for aircraft propulsion.

[0002] More specifically, the invention relates to damping of vibrations occurring during operation between platforms of two adjacent circumferential blades of a blade assembly. Background Art

[0003] A blade assembly for a turbine engine, for example a movable blade assembly of a low-pressure turbine stage of a turbojet engine, comprises a disk on which the blades are mounted.

[0004] At their outer radial ends, also called apex, the blades each have a transverse element called platform which has in particular the function of externally delimiting a flow channel for the gas flow through the turbine or more generally the associated parts of a turbomachine.

[0005] The platform of such a blade comprises an upstream edge and a downstream edge oriented perpendicularly to the flow direction of the gas flow. These edges are connected together via two side edges, which form the circumferential ends of the platform and, as a result, the platform of the blade is substantially in contact with the platforms of the two blades of the blade assembly, which are directly adjacent to it in the circumferential direction.

[0006] A known solution for limiting the vibration stresses on such blades consists in providing each of these side edges with a complex profile comprising a non-axial portion, that is to say, not parallel to the axis of rotation of the blade assembly, and comprising, for example, projections and complementary recesses. In practice, in order to damp the vibrations to which the blades are subjected during operation of the turbine, it is known to mount the blades on a disk with a torsional stress around the main axis of the disk. At the platform of a particular blade, this torsional stress causes the non-axial portion of the platform of the blade to be placed in contact with the non-axial portion of the platform of an adjacent blade. The vibrations of the blades during operation cause relative sliding at these contact areas, which, in combination with the contact pressure, produces friction damping. This type of construction is sometimes referred to as "interlocking".

[0007] By means of a solution of this type, the inter-platform contact forces may vary during operation, in particular due to the natural rotation of the airfoil about its centre line and due to the relative movement between circumferentially adjacent platforms, causing the latter to move closer or further apart.

[0008] In this particular case, where the contact force is reduced during operation relative to the stopped situation, it is necessary to amplify this contact force when mounting the blade in order to obtain the required level of force at high speeds, where the need to damp vibrations is most felt. This adaptation results in static overstress applied to the blade.

[0009] This is particularly detrimental in the case of blades made of composite materials, in particular ceramic matrix composites (CMCs), due to the relatively low permissible stress levels of this type of material, which makes the blades rather intolerant of pre-twist. Thus, for blades made of CMCs, a strong pre-twist upon installation causes high static stresses relative to the permissible level, and sometimes even above the permissible level.

[0010] Furthermore, this type of material expands significantly less than the metal material from which the blade is conventionally made. In the case of a blade having platforms configured as described above, this leads to a separation of circumferentially adjacent platforms when hot and thus to a reduction in contact forces during operation.

[0011] The applicant proposes means for damping turbine blade vibrations while overcoming the problems explained above with respect to the pre-twist of the blades.The general idea is thus to have no cold contact between circumferentially adjacent blades, but to create such a hot contact.

[0012] Document FR2956152 describes an embodiment of a solution of this type in which the contact between two circumferentially adjacent platforms is implemented by a damper component mounted with a certain radial clearance under one of the platforms which, during operation, bears under the action of the centrifugal force and thus exerts a friction force.

[0013] However, the solution presented in this document has disadvantages due to the fact that the damper component extends partially in the flow duct of the air flow and therefore constitutes the cause of a disturbance of this air flow.

[0014] Document FR2955142 describes another embodiment in which the damper member has an end recessed in a cavity formed in circumferentially adjacent platforms, also with a certain radial clearance, so as to bear during operation, under the action of centrifugal forces, against the wall delimiting each of the cavities.

[0015] The solution presented in this document makes it possible to introduce friction dissipation due to the relative movement of the platform, due to the fact that the cavity extends in the circumferential direction and has a closed section transverse to the circumferential direction. This solution does not intrude into the airflow. On the other hand, due to the relatively large volume of material required to form the cavity and the damper components, this solution leads to a heavier platform, which is detrimental to the strength of the airfoil.

[0016] Document FR 2 955 608 also describes another embodiment in which the damper member is a sheet having an end recessed in a cavity formed in circumferentially adjacent platforms so as to bear, during operation, under the action of centrifugal forces, against the wall delimiting each of the cavities. This solution therefore also makes it possible to limit the possibility of mutually separating the two platforms in the direction of the axis of rotation of the blade assembly. Summary of the invention

[0017] The object of the present invention is to provide a solution for damping the vibrations of a blade assembly of a turbomachine, in particular those whose blades are made of a composite material such as a ceramic matrix composite (CMC), in particular making it possible to limit the mechanical stresses applied to the blades as best as possible and avoiding all or some of the above-mentioned disadvantages.

[0018] To this end, the present invention proposes a blade assembly for a turbine, which includes a plurality of blades distributed around an axis, and each blade includes an airfoil, and a platform formed at the free end of the blade, and wherein the platform of each blade includes a friction member and an opening, which is elongated in a circumferential direction relative to the axis, and the friction member is engaged by an opening formed in the platform of circumferentially adjacent blades in the blade assembly, and the opening has a gap allowing the friction member to move in the opening at least in the circumferential direction.

[0019] In a preferred embodiment, at the opening, the angle between the plane of the elongated closed section of the opening and the radial direction relative to the axis is between 45 degrees and 90 degrees.

[0020] The plane of the elongated closed section is preferably orthogonal to the radial direction.

[0021] In a preferred embodiment, the friction member has a geometry of revolution about an axis extending in the radial direction.

[0022] In a preferred embodiment, the platform of each blade has a surface from which the friction member radially projects, and an engagement portion extending circumferentially from a circumferential end of the platform, radially offset relative to the surface and through which the opening is formed.

[0023] In a preferred embodiment, the friction member is formed integrally with the platform.

[0024] In a preferred embodiment, the friction member has an end surface that is flush with the surface of the engagement portion. In other words, the end surface of the friction member is flush with the surface of the engagement portion.

[0025] The invention also relates to a turbine for a turbomachine, comprising at least one blade assembly of the above-described type.

[0026] The invention also relates to a turbomachine comprising at least one turbine of the type described above. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] The invention will be better understood and other details, advantages and features will become apparent on reading the following description, provided by way of non-limiting example, and with reference to the accompanying drawings, in which:

[0028] Figure 1 is a schematic axial cross-sectional view of a turbine;

[0029] Figure 2 is a schematic partial perspective view of a blade assembly for a turbine according to a preferred embodiment of the present invention, viewed radially from the outside, particularly showing various platforms of blades of the blade assembly;

[0030] Figure 3A Seen radially from the inside Figure 2 A schematic partial perspective view of a blade of a blade assembly;

[0031] Figure 3B Seen radially from the outside Figure 3A A schematic partial perspective view of a blade;

[0032] Figure 4 yes Figure 2 Schematic partial perspective and cross-sectional views of a blade assembly showing two circumferentially adjacent blades of the blade assembly;

[0033] Figure 5 yes Figure 2 Schematic partial cross-sectional view of a blade assembly showing the connection area between two blades of a disk of the blade assembly.

[0034] Throughout the drawings, the same reference numbers may refer to the same or similar elements. DETAILED DESCRIPTION

[0035] Figure 1 A turbomachine 10 is shown, for example a twin-shaft turbofan engine for an aircraft, which generally comprises a fan 12 intended to suck in an air flow F1 which, downstream of the fan, is divided into a main flow F2 circulating in a primary flow channel, hereinafter referred to as primary duct PV, and a secondary flow F3 circulating in a secondary flow channel, hereinafter referred to as secondary duct SV, arranged around the primary duct PV.

[0036] The turbomachine generally comprises a low-pressure compressor 14 , a high-pressure compressor 16 , a combustion chamber 18 , a high-pressure turbine 20 and a low-pressure turbine 22 , which together define a main duct PV.

[0037] The corresponding rotors of the high-pressure compressor and high-pressure turbine are connected by a shaft called "high-pressure shaft", while the corresponding rotors of the low-pressure compressor and low-pressure turbine are connected in a well-known manner by a shaft called "low-pressure shaft". These rotors are rotatably mounted about an axis 28 of the turbine.

[0038] Throughout this description, the axial direction X is the direction of the axis 28. The radial direction R is a direction orthogonal to the axis 28 at any point and passing through the axis 28, and the orthogonal radial direction or circumferential direction C is a direction orthogonal to the radial direction R and the axis 28 at any point. A transverse plane is a plane orthogonal to the axis 28. The terms "inside" and "outside" refer to the relative proximity and relative distance of an element with respect to the axis 28, respectively. Finally, according to the axial direction X, with reference to the general direction of the flow of gas in the primary duct PV and the secondary duct SV of the turbine, the "upstream" and "downstream" directions are defined.

[0039] Figure 2-4 A blade assembly 40 for a turbine rotor is shown, such as that intended for use Figure 1 The low-pressure turbine 22 of the turbine 10 or another turbine component.

[0040] Such a blade assembly comprises an annular row of blades 42 distributed around an axis of the blade assembly which axis merges with the axis 28 when the blade assembly is mounted in a turbomachine and, for example, on a corresponding rotor disk.

[0041] Each blade 42 comprises in particular an airfoil 44 and a platform 46 formed at the outer radial end of the airfoil 44. Figure 5 Such a blade 42 generally also includes a platform 43A arranged at the base of the airfoil 44 and a foot 43B arranged below the platform 43A, so as to enable the blade to be mounted in the disk 41 of the blade assembly 40 .

[0042] The airfoils 44 of the blades have an aerodynamic profile enabling them to interact with the gas circulating in the main duct PV.

[0043] The main function of the platform 46 is to delimit the main duct PV externally through its respective inner surface 50 and to limit the gas leakage around the blade assembly 40 as best as possible, so as to maximize the interaction between the gas and the airfoil 44. To this end, the platform 46 generally comprises a rib 48, generally called a blade edge, formed so as to project radially outwards from its respective outer surface 54 and whose ends are intended to hollow out corresponding grooves in a ring made of abradable material arranged around the blade assembly 40 in the turbine.

[0044] Similar to known blade assembly platforms, each platform 46 has two circumferential ends 55A, 55B facing respectively toward corresponding circumferential ends of two platforms circumferentially adjacent thereto.

[0045] The blade assembly 40 has a specific feature that provides a new mode for limiting the vibration amplitude of the blade 42, which is achieved based on the fact that the platform 46 of each blade 42 includes a friction member 70 that is engaged by an opening 72 formed in the platform of the circumferentially adjacent blade, and the opening has an elongated closed section in the circumferential direction C, so that the friction member 70 can move in the opening 72 in the circumferential direction C according to the vibration of the platform. It should therefore be understood that the platform 46 of each blade includes a friction member 70 engaged in the opening 72 of the subsequent blade, and an opening 72 in which the friction member 70 of the preceding blade is engaged, with reference to any rotation direction around the axis 28, considering the concepts of the subsequent blade and the preceding blade. In addition, according to the general definition of the invention, the opening thus has a first dimension in the circumferential direction C, hereinafter referred to as the "large dimension LD", a second dimension smaller than the first dimension in a first direction orthogonal to the circumferential direction C, and hereinafter referred to as the "small dimension SD", and the opening 72 is open along the second direction orthogonal to the circumferential direction C and the first direction. The friction member 70 is movable at least in the circumferential direction C and in the preferred embodiment shown can also be moved to a certain extent in the first direction. In other words, the friction member 70 is engaged in the opening 72 with play in the circumferential direction C and in the first direction.

[0046] The elongated closed section of the opening 72 is defined in the plane P ( Figure 3B ), the plane forms an angle between 45 and 90 degrees with the radial direction R, and in the preferred embodiment shown is equal to 90 degrees. Thus, in this embodiment, the first direction, i.e. the direction of the small dimension SD of the opening is the axial direction X, while the opening is open in the radial direction R.

[0047] Furthermore, the friction member 70 has a geometry of revolution about an axis 73 extending in a radial direction R at least in one nominal orientation of the friction member.

[0048] In the embodiment shown, the friction member 70 is therefore a pin of generally cylindrical shape having, for the axis, the aforementioned axis 73 .

[0049] In addition, the friction member 70 protrudes radially outward from the outer surface 54 of the platform 46 ( Figure 3B ), and the platform 46 of each blade 42 includes a joint portion 52 extending circumferentially from one of the circumferential ends 55A of the platform, and an opening 72 is formed through the joint portion ( Figure 3A). Thus, the engagement portion 52 is circumferentially offset relative to the outer surface 54.

[0050] The engagement portion 52 is further offset radially outwardly relative to the outer surface 54 .

[0051] To this end, the platform 46 of each blade comprises, for example, at the circumferential end 55A, a rim 64 which extends protruding radially outwards relative to the outer surface 54 and from which a joining portion 52 in the form of a lug or blade protrudes circumferentially beyond the circumferential end 55A. It must therefore be understood that the joining portion 52 of a blade is arranged radially facing the outer surface 54 of the platform 46 of the circumferentially adjacent blade. The joining portion 52 is thus located outside the primary duct PV, which is preferred to avoid disturbing the airflow inside the duct.

[0052] In the illustrated embodiment, the friction member 70 is formed integrally with the platform 46. In addition, the friction member 70 has a radially outer end surface 70A that is flush with the radially outer surface 52A of the engagement portion 52 ( Figure 4 ).

[0053] Normally, during operation, due to vibrations of the blade, the mutual separation of the platforms 46 in the circumferential direction causes the friction members 70 to move along the edges of the openings 72. This generates friction with the nature of dissipating energy and thus limits the amplitude of the vibrations, which makes it possible to limit the stresses in the blade.

[0054] Furthermore, the cooperation between the friction member 70 and the opening 72 makes it possible to limit the gap amplitude of the platforms 46 relative to each other in the circumferential direction and in the aforementioned first direction. In fact, due to the gap provided between the edge of the opening 72 and the friction member 70 in the aforementioned first direction (i.e., the direction of the small dimension SD of the opening 72), when the friction member 70 is at the circumferential end of the opening 72, according to the vibration of the platform 46 in the first direction, friction between the friction member 70 and the edge of the opening 72 occurs at the end.

[0055] However, it should be noted that the opening 72 preferably has sufficient circumferential extension so that during normal operation, the friction member 70, although it may move closer to the circumferential ends of the opening 72, or even contact the edges of the opening at these ends, will not transmit significant forces in the circumferential direction of the adjacent circumferential platforms forming the opening 72 in question.

[0056] It should be noted that the invention is of course also applicable to blade assemblies in which the free end of the blade is the radially inner end of the latter, for example for a counter-rotating turbine. In this case, the configuration of the platforms is preferably reversed with respect to the radial direction. Thus, in this case, the friction members preferably protrude radially inwardly from the inner surface of each platform, and the engagement portion of each platform is preferably offset radially inwardly with respect to said inner surface of the platform.

[0057] The present invention may also be applied to static blade assemblies.

Claims

1. A blade assembly (40) for a turbine, comprising a plurality of blades (42) distributed around an axis (28), and each blade comprising an airfoil (44) and a platform (46) formed at a free end of the blade (42), and wherein: The platform (46) of each blade (42) includes a friction member (70) and an opening (72) that is elongated in the circumferential direction (C) relative to the axis (28), the friction member (70) being engaged through the opening (72) formed in the platform (46) of the circumferentially adjacent blade (42) in the blade assembly (40), the opening having a gap that allows the friction member (70) to move at least in the circumferential direction (C) within the opening (72).

2. The blade assembly according to claim 1, wherein: The angle between the plane (P) of the elongated closed section of the opening (72) and the radial direction (R) relative to the axis (28) at the opening (72) is between 45 degrees and 90 degrees.

3. The blade assembly according to claim 2, wherein: The plane (P) of the elongated closed section is orthogonal to the radial direction (R).

4. The blade assembly according to claim 2 or 3, wherein: The friction member (70) has a geometry that rotates about an axis (73) extending in the radial direction (R).

5. The blade assembly according to any one of claims 2 to 4, wherein: The platform (46) of each blade (42) has a surface (54) and an engagement portion (52), the friction member (70) radially protruding from the surface, the engagement portion extending circumferentially from a circumferential end (55A) of the platform, the engagement portion being radially offset relative to the surface (54) and the opening (72) being formed through the engagement portion.

6. The blade assembly according to claim 5, wherein: The friction member (70) is formed integrally with the platform (46).

7. The blade assembly according to claim 5 or 6, wherein: The friction member (70) has an end surface (70A) flush with a surface (52A) of the engagement portion (52).

8. A turbine (22) for a turbomachine, comprising at least one blade assembly (40) according to any one of claims 1 to 7.

9. Turbine (10) comprising at least one turbine (22) according to claim 8.

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