Blade for turbine of turbine engine and corresponding rotor, turbine and turbine engine

Abstract

The invention relates to a turbine engine rotor having an axis (X), comprising at least two sets of circumferentially adjacent blades (3). Each set of blades comprises: a blade root (31) configured to be mounted in an open groove at the outer periphery of a turbine engine rotor disc (2); an airfoil (33) extending the blade root in a radial direction (R) relative to the axis and having an aerodynamic profile; a radially inner platform (32) separating the airfoil from the blade root; a heel (34) at a free radial end of the blade as an extension of the airfoil, located at an end of the airfoil radially opposite the blade root, and including a radially outer platform. Each heel includes a locking device for locking the radially outer platform relative to another circumferentially adjacent radially outer platform. The locking device comprises two locking walls for locking the radially outer platform, each locking wall being configured to receive a circumferential end of a locking plate (42). The rotor also includes a locking plate, one of the two locking walls of a first of the two blades and one of the two locking walls of a second of the two blades receiving a circumferential end of the locking plate, respectively. The locking walls of the radially outer platform extend obliquely while being arranged obliquely and oppositely with respect to the radial direction. The locking plate includes a beveled circumferential end configured to contact the two locking walls of each radially outer platform. The invention also relates to a turbine engine (1) for an aircraft comprising said turbine. The invention also relates to a turbine engine turbine comprising at least one such rotor.

Description

TECHNICAL FIELD

[0001] The present invention relates to the field of turbine engines.

[0002] More particularly, the present invention relates to a turbine engine blade, particularly suitable for being installed in a turbine, such as a low-pressure turbine.

[0003] The present invention applies to all types of aeronautical turbine engines, in particular turbojet engines and turboprop engines. BACKGROUND

[0004] A turbine engine turbine, such as a low-pressure turbine, extending along a main axis defined in the axial direction, is composed of a plurality of stages, each stage being composed of a rotor wheel and a nozzle. The nozzle of the turbine comprises a plurality of stationary vanes for guiding the flow of gas coming from the combustion chamber.

[0005] The rotor wheel, in turn, comprises a plurality of blades. As described in document FR 3 026 429 A1, a blade comprises a root supported by a radially inner platform of the root, and an airfoil extending radially from the radially inner platform of the root, the airfoil defining a flow channel for the gas flowing along the turbine engine in the axial direction from upstream to downstream when the turbine engine is in operation.

[0006] As described in document FR 3 073 000 A1, a blade can also comprise a foot extending from the airfoil at the opposite end of the root, such that the airfoil is delimited by the radially inner platform of the root and the foot.

[0007] Since these blades are exposed to high-temperature environments in operation, they can be made of composite materials, in particular ceramic matrix composite materials. Indeed, such materials have better high-temperature resistance while being lighter than metal alloys.

[0008] Thus, it is known from the prior art to manufacture turbine engine rotor blades using composite materials, in particular ceramic matrix composite materials. For example, document PCT / FR2009 / 052309 describes a method for manufacturing such a blade.

[0009] It is known that a turbine is subjected to various stresses, in particular vibrations, which can cause relative movements between the different components, which in turn can cause malfunctions.

[0010] A known solution to limit the vibratory stresses of a turbine blade is to introduce a contact between two circumferentially adjacent blades, such that they are fixed in pairs. The vibrations of the blades in operation induce a relative sliding at the contact, which, combined with the contact pressure, makes it possible to generate a vibratory damping by friction.

[0011] In the known implementation, two lateral faces of each foot are provided with specific cutouts, such that the lateral faces of the foot and the lateral faces of the foot of the circumferentially adjacent blade form a complementary shape. Thus, the blades come into contact by angular deformation of the airfoils.

[0012] For example, the top surface of the foot can be provided with a protrusion and the bottom surface with a complementary indentation to cooperate with two circumferentially adjacent blades.

[0013] Furthermore, for blades made of ceramic matrix composite material, the mechanical stresses that they can withstand are relatively low, which means that the blades have a poor resistance to the stresses generated by the contact with circumferentially adjacent blades.

[0014] In addition to this relative resistance, the expansion rate of the ceramic matrix composite material is approximately one third of that of metallic materials, which can lead to a reduction in the contact force in operation due to the non-expansion of the ceramic matrix composite material blade.

[0015] Therefore, it is necessary to find an alternative to provide damping for the blades. SUMMARY

[0016] The present invention aims to at least partially overcome the aforementioned drawbacks of the prior art.

[0017] To this end, the present invention relates to a blade for mounting on a turbine engine rotor having an axis X, the blade comprising:

[0018] • a root configured to be mounted in a groove of a turbine engine rotor disc peripheral opening;

[0019] • an airfoil extending the root in a radial direction R with respect to the axis X and having an aerodynamic profile;

[0020] • a radially inner platform separating the airfoil from the root;

[0021] • a foot at the free radial end of the blade as an extension of the airfoil, at the end of the airfoil radially opposite the root, the foot comprising a radially outer platform.

[0022] According to the invention, the foot comprises a locking device for locking the radially outer platform with respect to the radially outer platform of another circumferentially adjacent blade, the locking device comprising two locking walls for locking the radially outer platform, each locking wall being configured to accommodate a circumferential end of a locking plate.

[0023] Thus, the present invention proposes a solution that at least partially solves certain drawbacks of the prior art.

[0024] In particular, the implementation of such a locking device makes it possible to limit the vibratory stresses and thus dampen the movement of each blade with respect to the circumferentially adjacent blades, while limiting the stresses exerted on the foot and the blade structure.

[0025] Furthermore, this makes it possible to reduce the effects caused by the expansion of the different elements, regardless of the material used to manufacture these elements.

[0026] It should be noted that the locking walls are also referred to as fastening walls or walls hereinafter.

[0027] According to particular aspects of at least one embodiment of the application, the locking walls of the radially outer platform are obliquely extended and are obliquely and oppositely arranged with respect to the radial direction R.

[0028] This makes it possible to use a simple mechanical device without having to modify the functional part of the blade, thus not affecting its normal operation.

[0029] According to particular aspects of at least one embodiment of the application, each fastening wall is oblique with respect to said radial axis R, the angle being between 0° and 60°.

[0030] According to particular aspects of at least one embodiment of the application, the fastening walls are symmetrical with respect to said radial axis R.

[0031] According to particular aspects of at least one embodiment of the application, said radially outer platform has a width in the circumferential direction, said walls extending from a middle portion of said radially outer platform width.

[0032] According to particular aspects of at least one embodiment of the application, said blade foot carries an upstream tab and a downstream tab extending radially outward from the radially outer platform, the locking device being axially arranged between the upstream tab and the downstream tab.

[0033] According to particular aspects of at least one embodiment of the application, the blade is at least partially composed of a ceramic matrix composite material.

[0034] The application also relates to a turbine engine rotor having an axis X comprising at least two circumferentially adjacent blades according to the above-described embodiment, and a locking plate, one of the two locking walls of a first of said two blades and one of the two locking walls of a second of said two blades each housing a circumferential end of the locking plate.

[0035] According to particular aspects of at least one embodiment of the application, the locking plate comprises bevelled circumferential ends configured to come into contact with the two locking walls of the radially outer platform.

[0036] According to particular aspects of at least one embodiment of the application, said bevelled circumferential ends are connected by a flat portion.

[0037] The application also relates to a turbine engine turbine comprising a rotor according to the above-described embodiment.

[0038] The application also relates to a turbine engine for an aircraft comprising a turbine according to the above-described embodiment. BRIEF DESCRIPTION OF DRAWINGS

[0039] The application and its different advantages will be easier to understand through the following description of exemplary, non-limiting embodiments of the application, and the attached drawings in which:

[0040] Figure 1 This is a schematic cross-sectional view of a turbine engine;

[0041] Figure 2 This is a side perspective view of a rotor portion according to an embodiment of the present invention, showing two circumferentially adjacent blades;

[0042] Figure 3 yes Figure 2 A front perspective view of two circumferentially adjacent blades;

[0043] Figure 4 It is based on Figure 2 A perspective view of the blades in the embodiment; and

[0044] Figure 5 It is based on Figure 2 Perspective view of the tenon in the embodiment.

[0045] Detailed Description of Embodiments of the Invention

[0046] It should be noted that this invention is applicable to high-pressure and / or low-pressure turbines of turbine engines, such as... Figure 1 The turbine engine shown.

[0047] The turbine engine 1 extends along axis X and is intended, for example, to be mounted on an aircraft (not shown), such as an airplane or helicopter, for example, under the wing, on the wing, or at the tail of the fuselage.

[0048] Figure 1 The turbine engine 1 shown is a direct-drive twin-shaft turbofan engine. However, this is not a limitation, as turbine engine 1 may not be used for installation on an aircraft, and may be other types of turbojet engines, such as geared turbofans, turboprops, or auxiliary power units (APUs).

[0049] Throughout this description, the axial direction corresponds to the direction of the longitudinal axis X, and the radial direction is perpendicular to and intersects the longitudinal axis X. Similarly, the axial plane is the plane that contains the longitudinal axis X, and the radial plane is the plane that is perpendicular to the longitudinal axis X.

[0050] Similarly, the adjectives “inner” (or “internal”) and “outer” (or “external”) are used relative to the radial direction, so the inner part of an element is closer to the longitudinal axis X in the radial direction than the outer part of the same element.

[0051] Furthermore, unless otherwise stated, the terms "upstream" and "downstream" are used relative to the general direction of gas flow through the turbine engine during operation.

[0052] like Figure 1As shown, the turbine engine 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 constitutes at least partially the rotation axis of the compressor section 12 and of the turbine section 16, which are rotatable relative to a casing 18 of the turbine engine 1 about the longitudinal axis X.

[0053] In operation, a gas flows inside the turbine engine along the longitudinal axis X from upstream to downstream of the turbine engine. The fan 10 sucks in an air flow, a part of which flows in a main flow duct 100, is successively compressed in the compressor section 2, ignited in the combustion chamber 14 and expanded in the turbine section 16, before being expelled from the turbine engine 1. In this way, the turbine engine 1 generates a thrust. In addition, this thrust can be used, for example, to propel an aircraft on which the turbine engine 1 is installed and fixed.

[0054] The compressor and the turbine are notably composed of one or more modules, each module comprising a plurality of stages, each stage comprising a rotor and a stator. The nozzles of the turbine comprise a plurality of blades for guiding the flow coming from the combustion chamber.

[0055] As such, the turbine rotor of the turbine engine comprises a disc carrying, at its periphery, a plurality of blades. These blades are uniformly distributed around the X axis of the turbine engine and are inserted into housings provided at the periphery of the rotor disc.

[0056] Reference will now be made to Figures 2 to 5 A first embodiment of these blades of the application is described.

[0057] In this embodiment, each of these blades is at least partially composed of a ceramic matrix composite material.

[0058] However, other embodiments can be provided in which the blades are made of another material, for example a metal or a metal alloy capable of withstanding the operating temperatures of such a turbine engine.

[0059] Each blade 3 comprises:

[0060] • a blade root 31 configured to fit into a recess (not shown) of an opening of the periphery of the turbine engine rotor disc 2;

[0061] • an airfoil 33 extending the blade root 31 in the radial direction R with respect to the axis X and having an aerodynamic profile;

[0062] • a radially inner platform 32 separating the airfoil 33 from the blade root 31.

[0063] Each of these blades 3 also comprises a blade foot 34, which is an extension of the airfoil 33 to the free radial end of the blade 3, and comprises a radially outer platform 40.

[0064] It is noted that the radially inner platform 32 and the radially outer platform 40 delimit the airfoil and determine the radial width of the flow channel.

[0065] The blade foot has the same circumferential width and is connected edge to edge in the tangential direction in the circumferential direction.

[0066] More particularly, in the illustrated embodiment, the tangential width of the blade foot corresponds to the tangential width of the radially outer platform.

[0067] As Figure 3 and Figure 4 In particular, in this embodiment, the blade foot 34 is provided with two tabs 340, namely an upstream tab and a downstream tab. They project radially outwards from the upstream and downstream ends of the blade foot near the radially outer platform 40, so as to be located outside the flow channel.

[0068] The two tabs are identical here, so they are inclined in the same direction and at the same angle.

[0069] In order to limit the vibratory stresses on the blades, the blade foot of each blade comprises locking means for locking the radially outer platform with respect to another circumferentially adjacent radially outer platform.

[0070] In other words, in the rotor, two circumferentially adjacent blades are held in position with respect to each other at their radially outer platforms by the locking means carried by the blade foot, these locking means being configured to each accommodate a circumferential end of a locking plate 42.

[0071] Thus, according to the application, the blade foot 34 of each blade comprises two locking walls 41a, 41b.

[0072] The locking means are provided at the radially outer platform of the blade foot, i.e. outside the flow channel area, so as not to interfere with the operation of the turbine engine.

[0073] The two locking walls of the radially outer platform are configured to each accommodate a circumferential end of a locking plate 42.

[0074] These walls are here centred with respect to the tangential width of the blade foot 34.

[0075] In other words, the walls 41a, 41b here extend from the centre of the tangential width of the radially outer platform 40.

[0076] In other words, the walls extend from the middle of the width of the radially outer platform, which has a width in the circumferential direction.

[0077] According to other embodiments, the walls can be positioned to extend from a position between 10% and 90% of the width of the radially outer platform, leaving about 10% of the edge at the end of the radially outer platform.

[0078] Whatever the embodiment of the application, such an edge makes it possible to achieve support.

[0079] The two walls 41a, 41b inclined with respect to the radial axis R thus extend obliquely from the radial outer platform 40.

[0080] As illustrated, the two walls 41b extend obliquely here towards the lateral end of the blade foot 34.

[0081] The lateral end refers to the end adjacent to the upstream and downstream ends of the blade foot carrying the tabs. In this way, the locking device is arranged axially between the upstream tab and the downstream tab.

[0082] The walls and the radial outer platform 40 thus delimit housings 410a, 410b.

[0083] In this embodiment, the walls are symmetrical, thus arranged opposite with respect to the radial axis R.

[0084] According to other embodiments, each wall can be inclined with respect to the radial axis R, at an angle between 0° and 60°.

[0085] According to other embodiments, the inclination of the walls can also be different.

[0086] Likewise, according to other embodiments, the two walls can extend different lengths.

[0087] In order to connect two circumferentially adjacent or consecutive blades, i.e. in order to lock the radial outer platform of one of the two blades with respect to the radial outer platform of the other blade, the rotor thus comprises a locking plate 42 for the two adjacent blades.

[0088] More particularly, one of the two locking walls of a first one of the two blades and one of the two locking walls of a second one of the two blades respectively house a circumferential end of the locking plate 42.

[0089] Thus, one of the housings 410a, 410b of a first one of the two blades and one of the two housings 410a, 410b of a second one of the two blades (opposite the housing of the first blade) respectively house a circumferential end of the locking plate 42.

[0090] In this way, the locking plate and the housings form a dovetail mechanical assembly.

[0091] This arrangement allows to dampen the movement of circumferentially adjacent blades with respect to each other.

[0092] In particular, the locking device and the locking plate make it possible to dampen the tangential movement of the blades with respect to each other, the movement of the blade foot causing a frictional energy dissipation with the locking plate.

[0093] In this embodiment, the locking plate is made of a ceramic matrix composite material.

[0094] According to other embodiments, the locking plate can be made of a material such as a metal or a metal alloy.

[0095] As Figure 5 In particular, the locking plate 42 of the illustrated embodiment has a trapezoidal prism shape, with two beveled peripheral ends 421 configured to come into contact with the two locking walls 41a, 41b of the radially outer platform 40, and a flat portion 420 connecting the two beveled peripheral ends.

[0096] More particularly, as Figure 3 In particular, the beveled peripheral ends 421 are arranged to be housed in the housings 410a, 410b.

[0097] The shape of the locking plate is complementary to the shape of the housing, so that an effective locking can be obtained, to dampen the movement of each vane relative to the circumferentially adjacent vanes, in particular the circumferential movement.

[0098] Here, since the walls are symmetrical relative to the radial axis R, the locking plate has a symmetrical shape relative to the central axis of the locking plate.

[0099] In other embodiments, locking plates having irregular shapes, in particular asymmetrical, can be envisaged.

[0100] For example, a locking plate having a rectangular profile can be envisaged, which cooperates with a housing formed by walls of substantially complementary shape.

[0101] A locking plate having a rectangular shape on one side and a triangular shape on the other side can also be envisaged.

[0102] Alternatively, a locking device can be implemented in which the walls form, with the locking plate, a dovetail mechanical assembly system, the walls being provided with protrusions and the tenon having, at its beveled peripheral end, a protrusion of complementary shape able to come into contact with the walls.

Claims

1. A turbine engine rotor having an axis (X) comprising at least two blades (3), each blade (3) comprising: The blade root (31) is constructed to be installed in a groove in the outer peripheral opening of the turbine engine rotor disk (2); An airfoil (33) extends the blade root (31) radially (R) relative to the axis (X) and has an aerodynamic profile; A radial inner platform (32) separates the airfoil (33) from the blade root (31); The blade heel (34), which extends from the airfoil (33) at the free radial end of the blade (3), is located at the radial end of the airfoil opposite to the blade root (31), and the blade heel (34) includes a radially outer platform (40). Each leaf heel (34) includes a locking device for locking the radially outer platform (40) relative to another circumferentially adjacent radially outer platform, the locking device including two locking walls (41a, 41b) for locking the radially outer platform (40), each locking wall being configured to receive the circumferential end of a locking plate (42). The two blades (3) are circumferentially adjacent. The rotor further includes a locking plate (42), wherein one of the two locking walls (41a, 41b) of the first blade and one of the two locking walls (41a, 41b) of the second blade respectively accommodate the circumferential end of the locking plate (42). The locking walls (41a, 41b) of the radial outer platform (40) of each of the two blades (3) extend obliquely, while being oblique and opposite to each other relative to the radial (R), and / or the locking plate (42) includes a chamfered circumferential end (421) configured to contact the two locking walls (41a, 41b) of each radial outer platform (40).

2. The rotor according to claim 1, characterized in that, Each locking wall (41a, 41b) is inclined relative to the radial axis (R) at an angle between 0° and 60°.

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

4. The rotor according to any one of the preceding claims, characterized in that, The radial outer platform (40) has a width in the circumferential direction, and the locking walls (41a, 41b) extend from the middle portion of the width of the radial outer platform (40).

5. The rotor according to any one of the preceding claims, characterized in that, The heel (34) has an upstream tab (340) and a downstream tab (341) extending radially outward from the radial outer platform (40), and the locking device is axially arranged between the upstream tab and the downstream tab (340).

6. The rotor according to any one of the preceding claims, characterized in that, It is at least partially composed of ceramic matrix composites.

7. The rotor according to any one of the preceding claims, characterized in that, The oblique circumferential end (421) is connected by a flat portion (420).

8. A turbine for a turbofan engine, comprising at least one rotor according to any one of claims 1 to 7.

9. A turbine engine (1) for an aircraft, comprising the turbine according to claim 8.

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