BLADED ARRANGEMENT FOR TURBOMACH, TURBINE FOR TURBOMACH AND TURBOMACH

Description

technical field

[0001] The present invention relates to the general field of bladed assemblies within turbomachinery, particularly turbomachinery for aircraft propulsion. More specifically, the present invention relates to a bladed assembly for a turbomachine, a turbine for a turbomachine, and a turbomachine.

[0002] It relates more specifically to the damping of vibrations appearing during operation between the platforms of two circumferentially adjacent blades of a bladed assembly. Prior art

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

[0004] At their external radial end, also called the apex, each blade has a transverse element, called a platform, which has the function of externally delimiting the flow path of the gas flow passing through the turbine or, more generally, the relevant part of the turbomachine.

[0005] The platform of such a blade has an upstream edge and a downstream edge oriented perpendicular to the direction of gas flow. These edges are connected to each other by means of two lateral edges forming the circumferential ends of the platform and by which the blade platform comes substantially into contact with the platforms of the two blades of the bladed assembly that are directly circumferentially adjacent to it.

[0006] A known solution for limiting vibrational stresses on such blades involves equipping each of their lateral edges with a complex profile that includes non-axial portions, i.e., portions not parallel to the axis of rotation of the bladed assembly, and incorporating, for example, additional protrusions and indentations. Indeed, to dampen the vibrations to which the blades are subjected during turbine operation, it is known to mount the blades on the disk with a torsional stress around their main axis. At the platform of a particular blade, this torsional stress results in contact between non-axial portions of the blade platform and non-axial portions of the platforms of neighboring blades. The vibrations of the operating blades induce relative slippage at these contact zones, which, coupled with contact pressures, creates frictional damping.This type of configuration is sometimes called "interlock".

[0007] With this type of solution, the inter-platform contact force can evolve during operation, notably due to a natural rotation of the blade around its mean line, and due to a relative movement between circumferentially adjacent platforms leading them to move closer together or further apart.

[0008] In the specific case where the contact force is reduced during operation compared to the standstill situation, it is necessary to amplify this contact force during blade assembly to obtain the desired force level at high speed, where the need to dampen vibrations is most acute. Such an adaptation results in static over-stresses applied to the blades.

[0009] This is particularly detrimental in the case of blades made of composite materials, especially CMC (ceramic matrix composite), due to the relatively low allowable stress levels for this type of material, which make the blades poorly tolerant of pretorsion. Thus, for a CMC blade, significant pretorsion during assembly induces static stresses that are high relative to the allowable level, and sometimes even exceed it.

[0010] Furthermore, this type of material expands considerably less than the metallic materials from which blades have conventionally been manufactured. In the case of blades with platforms configured as described above, this results in a separation of circumferentially adjacent platforms when hot, and therefore a reduction in contact force during operation.

[0011] The applicant has proposed methods for damping turbomachine blade vibrations while overcoming the difficulties related to blade pretorsion described above. The general idea is to avoid cold contact between circumferentially adjacent blades, but rather to generate such contact when hot.

[0012] Document FR2956152 describes an example of such a solution, in which the contact between two circumferentially adjacent platforms is achieved by a damping piece mounted with a certain radial clearance under one of the platforms and bearing under the other platform under the effect of the centrifugal force in operation and thus exerting friction forces.

[0013] The solution presented in this document has a drawback, however, because the damping piece extends partially into the flow path of the gas flow and thus constitutes a cause of disturbances to this gas flow.

[0014] Document FR2955142 describes another example, in which the damping part has ends housed within cavities formed within circumferentially adjacent platforms, also with a certain radial play, so as to come to rest against the wall delimiting each of the cavities under the effect of the centrifugal force in operation.

[0015] Because the cavities extend in the circumferential direction and have a closed cross-section transverse to the circumferential direction, the solution presented in this document allows for frictional dissipation due to the relative movements of the platforms. This solution does not interfere with the gas flow. However, this solution leads to an increase in platform weight due to the relatively large volume of material required to form the cavities and the damping element, which is detrimental to the blade's stability.

[0016] Document FR2955608 describes yet another example, in which the damping element is a sheet with ends housed within cavities formed in circumferentially adjacent platforms, so as to bear against the wall delimiting each cavity under the effect of centrifugal force during operation. This solution also limits the possibility of the two platforms moving apart in the direction of the axis of rotation of the bladed assembly.

[0017] The document JP2001200701A describes the prior state of the art. Description of the invention

[0018] The invention aims to provide a solution for attenuating vibrations of bladed turbomachine assemblies, particularly those whose blades are made of composite material such as a ceramic matrix composite (CMC), allowing in particular to limit as much as possible the mechanical stresses applied to the blades and to avoid in whole or in part the disadvantages mentioned above.

[0019] For this purpose, it proposes a bladed assembly for a turbomachine, comprising a plurality of blades distributed around an axis and each comprising a blade, and a platform formed at a free end of the blade, and in which the platform of each blade has a friction member and an oblong opening in a circumferential direction with respect to the axis, the friction member being engaged through the opening formed in the platform of a circumferentially adjacent blade within the bladed assembly with a clearance allowing at least one displacement of the friction member in the circumferential direction within the opening.

[0020] In preferred embodiments, an angle between a plane of an oblong closed section of the opening and a radial direction with respect to the axis, at the level of said opening, is between 45 degrees and 90 degrees.

[0021] The plane of said closed oblong section is preferably orthogonal to the radial direction.

[0022] In preferred embodiments, said friction element has a geometry of revolution about an axis which extends along said radial direction.

[0023] In preferred embodiments, the platform of each blade has a projecting surface from which said friction member extends radially, and an engagement part which extends circumferentially in projection from a circumferential end of the platform, which is offset radially with respect to the surface, and through which said opening is formed.

[0024] In preferred embodiments, said friction element is made of material with said platform.

[0025] In preferred embodiments, the friction element has an end surface that is level with a surface of the engagement portion. In other words, the end surface of the friction element is flush with said surface of the engagement portion.

[0026] The invention also relates to a turbine for a turbomachine, comprising at least one bladed assembly of the type described above.

[0027] The invention also relates to a turbomachine, comprising at least one turbine of the type described above. Brief description of the drawings

[0028] The invention will be better understood, and other details, advantages, and features thereof will become apparent from the following description, given by way of non-limiting example and with reference to the accompanying drawings, in which: [ Fig. 1 ] is a schematic axial cross-sectional view of a turbomachine; [ Fig. 2] is a partial schematic perspective view of a bladed assembly for a turbomachine according to a preferred embodiment of the invention, viewed radially from the outside, showing in particular the respective blade platforms of the bladed assembly; Fig. 3A ] is a partial schematic perspective view of a blade of the entire bladed assembly of the figure 2 , viewed radially from the inside; [ Fig. 3B ] is a partial schematic perspective view of the dawn of the figure 3A , viewed radially from the outside; [ Fig. 4 ] is a partial schematic perspective and cross-sectional view of the entire winged assembly of the figure 2 , showing two circumferentially adjacent blades of this bladed assembly; Fig. 5 ] is a partial schematic cross-sectional view of the bladed assembly of the figure 2 showing the junction region between two blades and a disk of said bladed assembly.

[0029] Throughout these figures, identical references may designate identical or analogous elements. Detailed presentation of preferred embodiments

[0030] There figure 1 illustrates a turbomachine 10, for example a twin-spool turbofan engine for aircraft, generally comprising a fan 12 for the intake of an airflow F1 which divides downstream of the fan into a primary flow F2 circulating in a primary flow channel, hereinafter referred to as the primary flow PV, and a secondary flow F3 circulating in a secondary flow channel, hereinafter referred to as the secondary flow SV, arranged around the primary flow PV.

[0031] 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 the primary flow PV.

[0032] The respective rotors of the high-pressure compressor and the high-pressure turbine are connected by a shaft called the "high-pressure shaft," while the respective rotors of the low-pressure compressor and the low-pressure turbine are connected by a shaft called the "low-pressure shaft," in a well-known manner. These rotors are mounted to rotate about an axis 28 of the turbomachine.

[0033] Throughout this description, the axial direction X is the direction of axis 28. The radial direction R is at every point a direction orthogonal to and passing through axis 28, and the orthoradial or circumferential direction C is at every point a direction orthogonal to both the radial direction R and axis 28. A transverse plane is a plane orthogonal to axis 28. The terms "internal" and "external" refer respectively to the relative proximity and relative distance of an element from axis 28. Finally, the "upstream" and "downstream" directions are defined with reference to the general direction of gas flow in the primary PV and secondary SV sections of the turbomachine, along the axial direction X.

[0034] THE figures 2-4 illustrate a bladed assembly 40 for a turbomachine rotor, for example intended for use within the low-pressure turbine 22 of the turbomachine 10 of the figure 1or another turbomachine component.

[0035] Such a bladed assembly comprises an annular row of blades 42 distributed around an axis of the bladed assembly which coincides with the axis 28 when the bladed assembly is mounted within a turbomachine, and for example mounted on a corresponding rotor disk.

[0036] Each blade 42 includes, in particular, a blade 44 and a platform 46 formed at the radially external end of the blade 44. With reference to the figure 5 , such a blade 42 also typically includes a platform 43A arranged at the base of the blade 44 and a foot 43B arranged under the platform 43A to allow the blade to be mounted in a disk 41 of the bladed assembly 40.

[0037] The 44 blades of the vanes have an aerodynamic profile that allows them to interact with the gases circulating in the primary PV stream.

[0038] The main function of the platforms 46 is to externally delimit the primary PV stream by means of respective internal surfaces 50 of the platforms, and to limit as much as possible the gas leaks around the bladed assembly 40 in order to maximize the interaction between the gas and the blades 44. For this purpose the platforms 46 typically have ribs 48, commonly called sealing strips, formed in radial projection outwards from respective external surfaces 54 of the platforms, and whose ends are intended to cut corresponding grooves within a ring of abradable material arranged around the bladed assembly 40 within a turbomachine.

[0039] Like the platforms of known winged assemblies, each of the platforms 46 has two circumferential ends 55A, 55B respectively opposite corresponding circumferential ends of the two platforms which are circumferentially adjacent to it.

[0040] The bladed assembly 40 has the particularity of offering a new method of limiting the amplitude of the vibrations of the blades 42, the implementation of which is based on the fact that the platform 46 of each blade 42 has a friction element 70 engaged through an opening 72 formed in the platform of a circumferentially adjacent blade and having an oblong closed section in the circumferential direction C so that the friction element 70 can move in the circumferential direction C within the opening 72 according to the vibrations of the platforms. It is therefore necessary to understand that the platform 46 of each blade includes both a friction element 70, which is engaged in the opening 72 of a subsequent blade, and an opening 72, in which the friction element 70 of a previous blade is engaged, the notions of subsequent blade and previous blade being considered by reference to an arbitrary direction of rotation around the axis 28.Furthermore, according to the general definition of the invention, the opening 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 and hereinafter referred to as the "small dimension SD," in a first direction orthogonal to the circumferential direction C, and the opening 72 is open in a second direction orthogonal to both the circumferential direction C and the first direction. The friction element 70 can move at least in the circumferential direction C and, in the preferred example illustrated, can also move to a certain extent in the first direction. In other words, the friction element 70 is engaged in the opening 72 with clearance in the circumferential direction C and in the first direction.

[0041] The oblong closed section of the opening 72 is defined in a plane P ( figure 3Bmaking an angle with the radial direction R between 45 degrees and 90 degrees, and equal to 90 degrees in the preferred example shown. Thus, in this example, the first direction, that is, the direction of the small dimension SD of the opening, is the axial direction X, while the opening is open along the radial direction R.

[0042] Furthermore, the friction element 70 has a geometry of revolution around an axis 73 which, at least in a nominal orientation of the friction element, extends along the radial direction R.

[0043] In the illustrated example, the friction element 70 is thus a pin of general cylindrical shape having as its axis the aforementioned axis 73.

[0044] Furthermore, the friction element 70 extends radially outwards from the external surface 54 of the platform 46 ( figure 3B), and the platform 46 of each blade 42 includes an engagement part 52 extending circumferentially in projection from one of the circumferential ends 55A of the platform, and through which the opening 72 is formed ( figure 3A ). The engagement part 52 is therefore circumferentially offset with respect to the external surface 54.

[0045] The engagement part 52 is further radially offset outwards relative to the external surface 54.

[0046] For this purpose, the platform 46 of each blade comprises, for example, at the circumferential end 55A, a rim 64 extending radially outward from the external surface 54, from which the engagement portion 52, shaped like a tab or blade, projects circumferentially beyond the circumferential end 55A. It should therefore be understood that the engagement portion 52 of a blade is arranged radially opposite the external surface 54 of the platform 46 of a circumferentially adjacent blade. The engagement portion 52 is thus located outside the primary stream PV, which is preferable in order to avoid disturbing the gas flow within the stream.

[0047] In the illustrated example, the friction element 70 is made of material with the platform 46. Furthermore, the friction element 70 has a radially external end surface 70A which is level with a radially external surface 52A of the engagement part 52 ( figure 4 ).

[0048] In general, a mutual separation of the platforms 46 in the circumferential direction during operation due to vibrations of the blades leads the friction element 70 to move along an edge of the opening 72. This generates friction which dissipates energy and thus limits the amplitude of the vibrations, thereby limiting the stresses in the blades.

[0049] Furthermore, the interaction between the friction element 70 and the opening 72 limits the range of motion of the platforms 46 relative to each other in both the circumferential direction and the aforementioned first direction. Indeed, due to the clearance between the edge of the opening 72 and the friction element 70 in the aforementioned first direction—that is, the direction of the small dimension SD of the opening 72—when the friction element 70 is located at a circumferential end of the opening 72, friction occurs between the friction element and the edge of the opening 72 at that end, as the platforms 46 vibrate in the first direction.

[0050] It should be noted, however, that the opening 72 preferably has a circumferential extent sufficient so that in normal operation, the friction member 70, although likely to approach the circumferential ends of the opening 72, or even to come into contact with the edge of the opening at the level of these ends, does not transmit significant force in the circumferential direction to the circumferentially adjacent platform in which the opening 72 is formed.

[0051] It should be noted that the invention is also applicable to bladed assemblies in which the free ends of the blades are their radially internal ends, for example, in counter-rotating turbines. In this case, the configuration of the platforms is preferably reversed with respect to the radial direction. Thus, the friction element in this case preferably projects radially inward from an internal surface of each platform, and the engagement portion of each platform is preferably offset radially inward relative to said internal surface of the platform.

[0052] The invention can also be applied to static bladed assemblies.

Claims

1. Bladed assembly (40) for a turbomachine, comprising a plurality of blades (42) distributed about an axis (28) and each comprising an airfoil (44), and a platform (46) formed at a free end of the blade (42), characterized in that the platform (46) of each blade (42) includes a friction member (70) and an opening (72) that is oblong in a circumferential direction (C) with respect to the axis (28), the friction member (70) being engaged through the opening (72) formed in the platform (46) of a circumferentially adjacent blade (42) within the bladed assembly (40) with a clearance allowing the friction member (70) to move at least in the circumferential direction (C) within the opening (72).

2. Bladed assembly according to claim 1, wherein an angle between a plane (P) of an oblong closed section of the opening (72) and a radial direction (R) with respect to the axis (28), at said opening (72), is between 45 degrees and 90 degrees.

3. Bladed assembly according to claim 2, wherein the plane (P) of said oblong closed section is orthogonal to the radial direction (R).

4. Bladed assembly according to claim 2 or 3, wherein said friction member (70) has a geometry of revolution about an axis (73) that extends in said radial direction (R).

5. Bladed assembly according to any one of claims 2 to 4, wherein the platform (46) of each blade (42) has a surface (54) from which said friction member (70) protrudes radially, and an engagement part (52) that extends circumferentially protruding from a circumferential end (55A) of the platform, which is radially offset with respect to the surface (54), and through which said opening (72) is formed.

6. Bladed assembly according to claim 5, wherein said friction member (70) is integrally formed with said platform (46).

7. Bladed assembly according to claim 5 or 6, wherein the friction member (70) has an end surface (70A) level with a surface (52A) of the engagement part (52).

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

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