Turbomachine rotor, turbomachine turbine and turbomachine for an aircraft

Abstract

The invention relates to a turbomachine rotor having an axis (X) comprising at least two circumferentially adjacent sets of blades (3). Each set of blades comprises: - a root (31) configured to be mounted in a groove which opens onto the outer periphery of a disc (2) of the turbomachine rotor; - a blade (33) extending the root in a radial direction (R) with respect to the axis and having an aerodynamic profile; - a radially inner platform (32) separating the blade from the root; - a shroud (34) which extends in the extension of the blade at a free radial end of the set of blades at the end of the blade that is radially opposite the root (31), the shroud (34) comprising a radially outer platform (40). The shroud comprises means for locking the radially outer platform with respect to another circumferentially adjacent radially outer platform. The locking means comprise two locking flanges (41a, 41b) of the radially outer platform which are configured each to receive a circumferential end of a locking plate (42). The rotor further comprises a locking plate, one of the two locking flanges of the first of the two sets of blades and one of the two locking flanges of the second of the two sets of blades each accepting a circumferential end of the locking plate. The locking flanges of the radially outer platform extend obliquely and are inclined in a direction opposite to the radial direction. The locking plate comprises bevelled circumferential ends which are configured to come into contact with the two locking flanges of the radially outer platform. The invention also relates to a turbomachine (1) for an aircraft comprising such a turbine. The invention also relates to a turbomachine turbine comprising at least one such rotor.

Description

[0001] DESCRIPTION

[0002] TITLE: TURBOMACHINE ROTOR, TURBOMACHINE TURBINE AND AIRCRAFT TURBOMACHINE

[0003] Field of invention

[0004] The present invention relates to the field of turbomachines.

[0005] More particularly, the invention relates to a turbomachine blade intended to equip in particular a turbine, and for example a low pressure turbine.

[0006] The invention relates to all types of aircraft turbomachine, in particular turbojets and turboprops.

[0007] State of the prior art

[0008] A turbomachine turbine extending along a main axis defining an axial direction, for example a low-pressure turbine, consists of a plurality of stages each formed by a moving wheel and a nozzle. The nozzle of a turbine comprises a plurality of fixed blades intended to straighten the gas flow from the combustion chamber.

[0009] For its part, the moving wheel comprises a plurality of blades. A blade comprises, as described for example in document FR 3 026429 A1, a root surmounted by a radially internal platform, and a blade extending radially from the radially internal platform of the root and delimiting a vein in which the gases of the turbomachine flow in operation, from upstream to downstream along the axial direction of the turbomachine.

[0010] A blading may also comprise, as described in document FR 3 073 000 A1, a heel extending at one end of the blade opposite the blade root, such that the blade is delimited by the radially internal platform of the root and by the heel.

[0011] Since these blades are exposed to high temperatures during operation, they can be made from composite materials, particularly ceramic matrix composite materials. In fact, compared to metal alloys, such materials can offer improved resistance to high temperatures while representing a lower mass.

[0012] It is thus known from the prior art to produce blades for turbomachine rotors from composite material, and in particular from ceramic matrix composite material. A method for manufacturing such a blade is for example described in document PCT / FR2009 / 052309.

[0013] The turbine is, as is known, subject to numerous constraints, particularly vibration, which can cause movements between the different parts and therefore a malfunction. A known solution to limit the vibration constraints on the blades of a turbine consists of introducing a contact between two circumferentially adjacent blades so as to maintain them in a pair-by-pair position. The vibrations of the blades during operation induce relative sliding at the level of this contact, which, coupled with the contact pressures, create vibration damping by friction.

[0014] Implementations are known in which the two lateral faces of each of the heels have a specific cutout such that a lateral face of the heel has a shape complementary to a lateral face of a heel of a circumferentially adjacent blade. Such contact is thus generated on these blades by means of an angular deformation of the blade.

[0015] For example, a heel may have a protrusion on the extrados side, and a complementary impression on the intrados side so as to cooperate with the two circumferentially adjacent blades.

[0016] Furthermore, and in the case of blades made of ceramic matrix composite material, the permissible mechanical stresses are relatively low, which makes the blades not very tolerant of the stresses generated by contact between circumferentially adjacent blades.

[0017] To this relatively low tolerance, we can add the fact that ceramic matrix composite materials expand approximately three times less than metallic materials, which can lead to a reduction in contact force during operation due to the non-expansion of ceramic matrix composite material blades.

[0018] There is therefore a need to find an alternative solution to provide damping to the blades.

[0019] Statement of the invention

[0020] The invention aims to remedy at least in part the drawbacks mentioned above relating to the techniques of the prior art.

[0021] To this end, the invention relates to a blading intended to be mounted on a turbomachine rotor of axis X, the blading comprising: a root configured to be mounted in a groove which opens at the external periphery of a disk of the turbomachine rotor; a blade extending said root, in a radial direction R with respect to the axis X, and having an aerodynamic profile; a radially internal platform separating said blade from said root; a heel which extends in the extension of said blade at a free radial end of said blading, at the end of the blade radially opposite the root, said heel comprising a radially external platform,

[0022] According to the invention, said heel comprises means for locking the radially external platform with respect to another circumferentially adjacent radially external platform, the locking means comprising two locking walls of the radially external platform configured to each receive a circumferential end of a locking plate.

[0023] Thus, the invention provides an approach for at least partially overcoming some of the drawbacks of the prior art.

[0024] In particular, the implementation of such locking means makes it possible to limit the vibrational stresses and therefore to dampen the movement of each of the blades relative to the circumferentially adjacent blades while limiting the stresses applied to the structure of the heel and the blade.

[0025] In addition, this helps to reduce the effects due to the expansion of the different elements regardless of the materials from which these elements are made.

[0026] According to a particular aspect of at least one embodiment of the invention, the locking walls of the radially external platform extend obliquely while being inclined and opposite with respect to the radial direction R.

[0027] This makes it possible to implement mechanically simple means that do not require modification of the functional part of a blade so as not to disrupt its proper functioning. According to a particular aspect of at least one embodiment of the invention, the fixing walls are each inclined obliquely relative to said radial axis R at an angle of between 0° and 60°.

[0028] According to a particular aspect of at least one embodiment of the invention, the fixing walls are symmetrical with respect to said radial axis R.

[0029] According to a particular aspect of at least one embodiment of the invention, said radially external platform has a width in the circumferential direction, said walls extending from a median portion of the width of said radially external platform. According to a particular aspect of at least one embodiment of the invention, said heel carries an upstream lip and a downstream lip which extend radially projecting from the radially external platform, the locking means being arranged axially between the upstream and downstream lips.

[0030] According to a particular aspect of at least one embodiment of the invention, the blading is at least partially composed of ceramic matrix composite material. The invention also relates to an X-axis turbomachine rotor comprising at least two bladings according to one of the aforementioned embodiments, circumferentially adjacent, and a locking plate, one of said two locking walls of the first of said two bladings and one of said two locking walls of the second of said two bladings each receiving a circumferential end of the locking plate.

[0031] According to a particular aspect of at least one embodiment of the invention, the locking plate comprises beveled circumferential ends configured to come into contact with said two locking walls of the radially external platform.

[0032] According to a particular aspect of at least one embodiment of the invention, said beveled circumferential ends are connected by a flat portion.

[0033] The invention also relates to a turbomachine turbine comprising a rotor according to the aforementioned embodiment.

[0034] The invention also relates to a turbomachine for aircraft comprising a turbine according to the aforementioned embodiment.

[0035] Presentation of figures

[0036] The invention, as well as the various advantages that it presents, will be more easily understood in the light of the following description of an illustrative and non-limiting embodiment thereof, and of the appended drawings among which:

[0037] [Fig. 1] is a schematic sectional view of a turbomachine;

[0038] [Fig. 2] is a side perspective view of a rotor portion according to one embodiment of the invention, two circumferentially adjacent blades being illustrated;

[0039] [Fig. 3] is a front perspective view of the two circumferentially adjacent blades of Fig. 2;

[0040] [Fig. 4] is a perspective view of a blade according to the embodiment of Figure 2, and [Fig. 5] is a perspective view of a tenon according to the embodiment of Figure 2.

[0041] Detailed description of an embodiment of the invention

[0042] It should be noted that the invention applies to high-pressure and / or low-pressure turbines of a turbomachine, such as the turbomachine shown in Figure 1.

[0043] This turbomachine 1, which extends 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 the wing of the aircraft, on the wing or even at the rear of the fuselage of the aircraft.

[0044] The turbomachine 1 illustrated in Figure 1 is a twin-spool, twin-flow, direct-drive turbojet engine. This is not, however, limiting since the turbomachine 1 may not be intended to be mounted on an aircraft, may be another type of turbojet engine, such as a geared turbojet engine, a turboprop engine or an auxiliary power unit (also called "APU", for "Auxiliary Power Unit").

[0045] Throughout the description, an axial direction corresponds to the direction of the longitudinal axis X and a radial direction is a direction perpendicular to the longitudinal axis X 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.

[0046] Similarly, the adjectives "inner" (or "internal") and outer ("or external") are used with reference to a radial direction so that the inner part of an element is, in a radial direction, closer to the longitudinal axis X than the outer part of the same element.

[0047] Additionally, unless otherwise specified, the terms "upstream" and "downstream" are used in reference to the overall direction of gas flow through the operating turbomachine.

[0048] As seen in Figure 1, the turbomachine 1 comprises, from upstream to downstream, a fan 10, a compressor section 12, a combustion chamber 14 and a turbine section 16. The 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 capable of being driven in rotation about the longitudinal axis X relative to a casing 18 of the turbomachine 1. In operation, the gases circulate inside the turbomachine from upstream to downstream of the turbomachine, along the longitudinal axis X. The fan 10 sucks in an air flow of which a portion, circulating within a primary vein 100 is, successively, compressed within the compressor section 12, ignited within the combustion chamber 14, and expanded within the turbine section 16 before to be ejected from the turbomachine 1. In this way, the turbomachine 1 generates thrust.This thrust can also, for example, be used to benefit the aircraft on which the turbomachine 1 is attached and fixed.

[0049] Compressors and turbines are composed in particular of one or more modules, each of these modules comprising a plurality of stages each comprising a rotor and a stator. The distributor of a turbine comprises a plurality of blades intended to straighten the flow of gas from the combustion chamber.

[0050] For its part, the turbomachine turbine rotor comprises a disc carrying at its periphery a plurality of blades. These blades are regularly distributed around the X axis of the turbomachine and are inserted into housings provided on the periphery of the rotor disc.

[0051] We now present, in relation to figures 2 to 5, a first embodiment of the invention of these blades.

[0052] In this embodiment, each of these blades is at least partially composed of ceramic matrix composite material. However, embodiments could be provided in which the blades are made of another material, such as a metal or a metal alloy, capable of withstanding the operating temperatures of such a turbomachine.

[0053] Each of the blades 3 comprises: a root 31 configured to fit into a groove (not shown) which opens onto the outer periphery of the turbomachine turbine rotor disk 2; a blade 33 extending the root 31 in a radial direction R relative to the axis X and having an aerodynamic profile, and a radially internal platform 32 separating the blade 33 from the root 31.

[0054] Each of these blades 3 also comprises a heel 34 which extends in the extension of the blade 33 at a free radial end of the blade 3, and which comprises a radially external platform 40.

[0055] It should be noted that the radially internal platform 32 as well as the radially external platform 40 delimit the blade and define the radial width of the vein.

[0056] The blade heels have an identical circumferential width and are joined edge to edge circumferentially in a tangential direction.

[0057] More particularly, in the illustrated embodiment, the tangential width of the heel corresponds to the tangential width of the radially outer platform.

[0058] As illustrated in particular in Figures 3 and 4, the heel 34 carries, in this embodiment, two wipers 340, namely an upstream wiper and a downstream wiper. They extend radially projecting outwards from the radially external platform 40, close to the upstream and downstream ends of the heel, so as to be outside the vein.

[0059] These two lips are identical here so that they are inclined in the same direction and at the same angle.

[0060] To limit the vibrational stresses on the blades, the heel of each blade includes means for locking the radially external platform with respect to another circumferentially adjacent radially external platform.

[0061] In other words, in the rotor, two circumferentially adjacent blades are held in position relative to each other at their radially outer platform by locking means carried by the heel, these locking means being configured to each receive a circumferential end of a locking plate 42.

[0062] Therefore, according to the invention, the heel 34 of each blade comprises two locking walls 41a, 41b. The locking means are arranged at the level of the radially external platform of the heel, that is to say outside the zone of the vein so as not to disturb the operation of the turbomachine.

[0063] These two radially outer platform locking walls are each configured to receive a circumferential end of a locking plate 42.

[0064] These walls are here centered in relation to said tangential width of the heel 34.

[0065] In other words, the walls 41a, 41b extend here from a center of the platform radially external 40 relative to the tangential width.

[0066] In other words, the walls extend from a middle portion of the width of the radially external platform, this radially external platform having a width in the circumferential direction.

[0067] In other embodiments, the walls could be positioned to extend from a position between 10% and 90% of the width of the radially outer platform, so as to leave a rim of approximately 10% at the ends of the radially outer platform.

[0068] Such a rim provides a possibility of support regardless of the embodiment of the invention.

[0069] From the radially external platform 40 there therefore extend obliquely two walls 41a, 41b inclined relative to a radial axis R.

[0070] As illustrated, the two walls 41b here extend obliquely towards the lateral ends of the heel 34.

[0071] By lateral end is meant the ends adjacent to the upstream and downstream ends of the heel carrying the wipers. In this way, the locking means are arranged axially between the upstream and downstream wipers.

[0072] The low walls and the radially external platform 40 thus delimit housings 410a, 410b.

[0073] In this embodiment, the walls are symmetrical and therefore opposite with respect to the radial axis R.

[0074] According to other embodiments, the walls could each be inclined obliquely relative to the radial axis R at an angle between 0° and 60°.

[0075] According to still other embodiments, the inclination of the walls could also be different.

[0076] Similarly, according to other embodiments, the two walls could extend over a different length.

[0077] To connect two circumferentially adjacent or successive blades, that is to say to lock the radially external platform of one of the two blades with respect to the radially external platform of the other of the two blades, the rotor therefore comprises a locking plate 42 for two adjacent blades.

[0078] More particularly, one of the two locking walls of the first of the two blades and one of the two locking walls of the second of the two blades each receive a circumferential end of the locking plate 42.

[0079] Thus, one of the housings 410a, 410b of the first of the two blades and one of the two housings 410a, 410b of the second of the two blades, opposite the housing of the first of the two blades, each receive a circumferential end of the locking plate 42.

[0080] In this way, the locking plate and the housings form a mechanical dovetail assembly.

[0081] Such an assembly makes it possible to dampen the movement of the circumferentially adjacent blades relative to each other.

[0082] In particular, the locking means and the locking plate make it possible to dampen a tangential movement of the blades relative to each other, the movements of a heel of a blade generating energy dissipation by friction with the locking plate. In this embodiment, the locking plate is made from a ceramic matrix composite material.

[0083] In other embodiments, the locking plate may be made from a material such as a metal or a metal alloy.

[0084] As can be seen in particular in Figure 5, the locking plate 42 of the illustrated embodiment takes the form of a trapezoidal prism with two beveled circumferential ends 421 configured to come into contact with the two locking walls 41a, 41b of the radially external platform 40, and a flat portion 420 connecting these two beveled circumferential ends.

[0085] More particularly, and as can be seen in particular in Figure 3, the beveled circumferential ends 421 are designed to be housed in the housings 410a, 410b.

[0086] The shape of the locking plate which is complementary to the shapes of the housings makes it possible to obtain effective locking to dampen a movement of each blade relative to a circumferentially adjacent blade, and in particular a movement in a circumferential direction.

[0087] Here, the locking plate has a symmetrical shape with respect to a central axis of this locking plate because the walls are symmetrical with respect to the radial axis R.

[0088] In other embodiments, it would be possible to implement a locking plate having a non-regular shape, and in particular not having any symmetry. For example, it would be possible to implement a locking plate having a rectangular profile, cooperating with housings formed by walls of substantially complementary shape.

[0089] Locking plates could also be implemented with a rectangular shape on one side and a triangular shape on the other.

[0090] It would also be possible to implement locking means in which the walls form, with the locking plate, a mechanical dovetail assembly system with the walls which carry protuberances and the tenon which has, on its bevelled circumferential ends capable of being in contact with the walls, protuberances of complementary shape.

Claims

CLAIMS

1. Turbomachine rotor of axis (X) characterized in that it comprises at least two blades (3), each blade (3) comprising: a root (31) configured to be mounted in a groove which opens at the external periphery of a disk (2) of the turbomachine rotor; a blade (33) extending said root (31), in a radial direction (R) with respect to the axis (X), and having an aerodynamic profile; a radially internal platform (32) separating said blade (33) from said root (31); a heel (34) which extends in the extension of said blade (33) at a free radial end of said blading (3), at the end of the blade radially opposite the root (31), said heel (34) comprising a radially external platform (40), characterized in that said heel (34) comprises means for locking the radially external platform (40) with respect to another circumferentially adjacent radially external platform,the locking means comprising two locking walls (41a, 41b) of the radially external platform (40) configured to each receive a circumferential end of a locking plate (42). said two blades (3) being circumferentially adjacent, said rotor further comprising a locking plate (42), one of said two locking walls (41a, 41b) of the first of said two blades and one of said two locking walls (41a, 41b) of the second of said two blades each receiving a circumferential end of the locking plate (42), the locking walls (41a, 41b) of the radially outer platform (40) extending obliquely while being inclined and opposite with respect to the radial direction (R), and / or the locking plate (42) comprising beveled circumferential ends (421) configured to come into contact with said two locking walls (41a,41b) of the radially external platform (40).,

2. Rotor according to claim 1, characterized in that the fixing walls (41a, 41b) are each inclined obliquely relative to said radial axis (R) at an angle between 0° and 60°.

3. Rotor according to claim 1 or 2, characterized in that the fixing walls (41a, 41b) are symmetrical with respect to said radial axis (R).

4. Rotor according to one of the preceding claims, characterized in that said radially external platform (40) has a width in the circumferential direction, said walls (41a, 41b) extending from a median portion of the width of said radially external platform (40).

5. Rotor according to one of the preceding claims, characterized in that said heel (34) carries an upstream lip (340) and a downstream lip (341) which extend radially projecting from the radially external platform (40), the locking means being arranged axially between the upstream and downstream lips (340).

6. Rotor according to one of the preceding claims, characterized in that it is at least partially composed of ceramic matrix composite material.

7. Rotor according to one of the preceding claims, characterized in that said beveled circumferential ends (421) are connected by a flat portion (420).

8. Turbomachine turbine comprising at least one rotor according to one of claims 1 to 7.

9. Turbomachine (1) for aircraft comprising a turbine according to claim 8.

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