ROTOR BLADE FOR A TURBOMACHINE

Abstract

The present invention relates to a rotor blade (2) for an aircraft turbomachine (10), the blade comprising an aerodynamic blade (3) extending along an axis (A) from a root to a tail (5), the blade comprising lower (32) and upper (34) surfaces connected by leading (36) and trailing (38) edges, the tail comprising a platform (50) comprising lateral edges (55, 56) situated respectively on the lower and upper surfaces, each of the lateral edges comprising a wear-resistant coating (60, 70) comprising a contact face (60, 70) adapted to come into contact with another contact face (62, 72) of a wear-resistant coating (6, 7) of another rotor blade (2), wherein the contact face of the wear-resistant coating of each of the lateral edges comprises teeth (8, 86, 87) which are configured to cooperate by interlocking with complementary teeth on the other contact surface. Figure for the abbreviation: Fig. 6

Description

Title of the invention: ROTOR BLADE FOR A TURBOMACHINE Technical field of the invention

[0001] The present application relates to the general field of aircraft turbomachinery. More particularly, the invention relates to a rotor blade for an aircraft turbomachine, as well as a rotor comprising such a blade. Technical background

[0002] Conventionally, an axial turbine of a turbomachine consists of a succession of axial stages (along the axis of gas flow) arranged one behind the other. Each stage comprises a rotating bladed wheel forming a rotor and a bladed distributor forming a stator. The rotor is rotated opposite the corresponding distributor.

[0003] The rotor is conventionally made up of an annular disk centered on an axis of the turbomachine, on which a plurality of blades are fixed.

[0004] Traditionally and with reference to figures 1 to 4, a rotor blade 2 1 comprises an aerodynamic blade 3 extending along an axis A of lengthening (or otherwise stacking) from a foot 4 to a heel 5.

[0005] The blade 3 comprises intrados faces 32 and extrados faces 34 connected by leading edges 36 and trailing edges 38.

[0006] When several blades 2 are fixed to the disk, their flanges 5 are arranged circumferentially edge to edge so as to form a circumferential ring delimiting a surface of revolution around the axis of the turbomachine ([Fig.4]). This ring has, in particular, the function of delimiting the outer surface of a channel in which a flow of gas circulates.

[0007] The heel 5 includes a platform 50 having an upstream spoiler 53 and a downstream spoiler 54. The platform 50 includes opposing lateral edges 55, 56 configured to cooperate by interlocking with complementary lateral edges 55, 56 of adjacent rotor blades 2.

[0008] Each of the lateral edges 55, 56 comprises a wear-resistant coating 60, 70 having a contact face 62, 72 adapted to come into contact with another contact face 62, 72 of a wear-resistant coating 60, 70 of another rotor blade 2. The contact height between contact faces 62, 72 of adjacent blades 2 is generally on the order of 1 to 2 mm.

[0009] Once the blades are mounted in a ring on the disc, the wear-resistant coatings allow, during operation, the dissipation of the vibrational energy of a natural mode of The blade is damped by contact between the blades. Such damping is used in particular to prevent failure due to vibration fatigue.

[0010] In order to dampen the vibrations to which the blades 2 are subjected during operation and to provide rigidity to the assembly, the blades 2 are mounted on their disk with a torsional stress around a torsional axis positioned relative to the axis A. The geometry of the flanges 5 is such that each blade 2 is subjected to torsional stress by bearing on the neighboring blades 2 at the lateral edges 55, 56. These lateral edges 55, 56 therefore define inter-blade contact surfaces Z and are the site of significant friction during the operation of the turbomachine 10. [Fig. 2] illustrates an example of this inter-blade contact surface Z.

[0011] The role of the lateral edges, which are held in contact by the torsional stress applied to the blade, is to dampen the bending vibration mode to prevent blade breakage during dynamic operation due to resonance. However, during operation, such a rotor blade can be subject to interference from other parts and can therefore wear through friction during blade movement, and / or the contact height may be insufficient to hold the interlocking blades together. These factors can cause loss of contact between the lateral edges.

[0012] Such a loss of contact and the resulting play between the heels can cause the blade to dislocate, i.e. the lateral edges of the adjacent blades which should be in contact are no longer in contact, with in particular an axial and / or radial displacement of the blade heel.

[0013] For example, the blade 2 can disengage from the rotor 1 in a radial direction DR (i.e., with respect to axis A) and / or a lateral direction DL (for example, along a plane transverse to axis A). The possible DR and DL directions of blade 2 disengagement are illustrated in [Fig. 3].

[0014] Figure 4 illustrates another possible example of the blades 2d disengaging and overlapping in the rotor 1 during an abnormally hot start-up of the turbomachine. Indeed, during start-up, the rotor blades can abruptly penetrate an abradable material of an external casing surrounding the rotor, which generates a torque. Under the effect of this torque, the blades can disengage and then overlap.

[0015] The disengagement of the blade (or blades) during operation may also be due to an accumulation of dimensional tolerances at the level of the blade foot or disc, the blade wall thicknesses, the blade mass, and / or the different modes of vibration at the level of the blade roots.

[0016] This disengagement can lead to an increase in rotor blade vibrations, and possibly premature deterioration of the disengaged blade or blades.

[0017] This disengagement can also lead to the removal and dismantling of the turbomachine to reposition the bearings correctly. If such removal and dismantling are not carried out, there may be a loss of damping or even a risk of dynamic failure, which in the worst-case scenario could cause an engine shutdown in flight or partial or total turbine failure in flight. Besides the risk of failure, there may also be the consequence of the parts being unrepairable and having to be scrapped during maintenance or even before, which negatively impacts costs.

[0018] Thus, there is a need to prevent and stop the disengagement of rotor blades during operation for an aircraft turbomachine. Summary of the invention

[0019] The present invention proposes a simple, effective and economical solution to the aforementioned disadvantages of the prior art.

[0020] For this purpose, the invention relates to a rotor blade for an aircraft turbomachine, said blade comprising an aerodynamic blade extending along an axis A of elongation from a root to a tail, the blade comprising intrados and extrados faces connected by leading and trailing edges, said tail comprising a platform comprising an upstream spoiler located on the leading edge side, a downstream spoiler located on the trailing edge side, and lateral edges located respectively on the intrados and extrados faces, each of the lateral edges comprising an anti-wear coating comprising a contact face adapted to come into contact with another contact face of an anti-wear coating of another rotor blade.

[0021] According to the invention, the contact face of the wear-resistant coating of each of the lateral edges comprises teeth which are configured to cooperate by interlocking with complementary teeth of said other contact face.

[0022] Thus, this solution makes it possible to achieve the aforementioned objective. In general, the invention makes it possible to effectively prevent the disengagement (or in other words, the separation or disengagement) of the rotor blade during operation, particularly with respect to the other blades of this rotor. To this end, the invention proposes integrating teeth on the contact faces of the wear-resistant coatings that are complementary to those of another contact face of another rotor blade, so that these teeth cooperate by interlocking. This makes it possible to strengthen the retention, particularly in a radial, axial, and / or lateral direction (relative to (the blade extension axis), of rotor blades nested one inside the other at their roots. Furthermore, the teeth significantly increase the bearing surface of each lateral edge of the blade. Thus, the lateral edges of the blade are mutually locked with those of one or more other blades in operation.

[0023] Consequently, the invention makes it possible to significantly reduce (or even eliminate) the risk of rotor blade breakage during operation and premature deterioration of this blade, and thus increase the lifespan of the rotor blade.

[0024] The rotor blade according to the invention may comprise one or more of the following features, taken individually or in combination with each other:

[0025] - a first of the lateral edges comprises a first wear-resistant coating comprising a first contact face with first teeth, and a second of the lateral edges includes a second wear-resistant coating comprising a second contact face with second teeth which are complementary to the first teeth;

[0026] - said teeth on each contact face are distributed in a matrix manner in rows and columns of teeth;

[0027] - said teeth have in cross-section a shape chosen from a rectangle, a square, a polygon, a parabola, a groove, a chevron and a helix;

[0028] - said teeth are staggered between the rows and / or between the columns;

[0029] - each of the anti-wear coatings is formed in a single piece with its teeth;

[0030] - each of the wear-resistant coatings is made of a cobalt-based alloy and / or tungsten;

[0031] - the number of teeth on each contact face is between two and twenty;

[0032] - the teeth on each contact face are separated by hollows;

[0033] - the teeth have bearing faces extending in the same plane P;

[0034] — each of the first teeth of the first contact face has a size (such as a length, width or thickness) identical to that of the second teeth of the second contact face;

[0035] — each of the first teeth of the first contact face has a size (such as a length, width or thickness) different from that of the second teeth of the second contact face;

[0036] — the size of each of the first teeth of the first contact face is inferior to that of the second teeth of the second contact face.

[0037] The present invention also relates to a rotor (or in other words a moving wheel) for an aircraft turbomachine, comprising a disc carrying on its periphery an annular row of blades as described previously.

[0038] The rotor can be a rotor of a high-pressure turbine of the turbomachine, and / or a rotor of a low-pressure turbine of the turbomachine.

[0039] The present invention also relates to an aircraft turbomachine, comprising at least one rotor blade as described above or a rotor as described above.

[0040] The turbomachine can be a turbojet, a turboprop or a turbomotor. Brief description of the figures

[0041] The invention will be better understood and other details, features and advantages of the invention will become more apparent upon reading the following description, given by way of non-limiting example and with reference to the accompanying drawings in which:

[0042] [Fig.1] is a schematic perspective and partial view of a rotor blade according to the prior art for an aircraft turbomachine;

[0043] [Fig.2] is a schematic perspective and partial view of a rotor comprising two rotor blades of [Fig.1];

[0044] [Fig.3] is a schematic view similar to [Fig.2] illustrating directions of rotor blade disengagement in operation;

[0045] [Fig.4] is another schematic perspective and partial view illustrating a disengagement of the rotor blades;

[0046] [Fig.5] is a schematic half axial cross-sectional view of an aircraft turbomachine according to the invention;

[0047] [Fig.6] is a schematic perspective view of a rotor blade of the turbomachine of [Fig.5];

[0048] [Fig.7] is a schematic perspective and partial view of a first example of the rotor blade of [Fig.6] from the intrados side;

[0049] [Fig.8] is another schematic perspective and partial view of the rotor blade of [Fig.6] from the extrados side;

[0050] [Fig.9] is an enlarged schematic view of teeth of an anti-wear coating of the rotor blade of [Fig.7];

[0051] [Fig.10] is a schematic perspective and partial view of a rotor comprising two rotor blades of [Fig.7] or 8;

[0052] [Fig. 1 1] is an enlarged schematic view of the interlocking teeth of two rotor blades of [Fig. 10];

[0053] [Fig. 12] is a schematic perspective and partial view of a second example of the rotor blade of [Fig. 6];

[0054] [Fig. 13] is a schematic view of an assembly of the teeth of the wear-resistant coatings of two rotor blades of [Fig. 12].

[0055] Elements having the same functions in the different implementations have the same references in the figures. Detailed description of the invention

[0056] Generally, in this application, the terms "longitudinal" and "axial" refer to the orientation of structural elements extending along a longitudinal axis (such as that of the turbomachine). The term "radial" refers to the orientation of structural elements extending in a direction perpendicular to the longitudinal axis. The terms "interior" and "exterior," and "internal" and "external," are used with reference to positioning relative to the longitudinal axis. Thus, a structural element extending along the longitudinal axis has an interior surface facing the longitudinal axis and an exterior surface opposite its interior surface.

[0057] Furthermore, the terms "upstream" and "downstream" can be defined in relation to the direction of air flow in the turbomachine.

[0058] Figures 1 to 4 have been described in the technical background of this application and illustrate a part of the rotor blade 2 according to the prior art.

[0059] The invention can be applied in a non-limiting way to an aircraft turbomachine 10, such as a turbojet, a turboprop or a turbomotor.

[0060] Fig. 5 illustrates an example of an aircraft turbomachine 1 extending along a longitudinal X axis.

[0061] The turbomachine 10 may include, from upstream to downstream, a blower 11 (or in other words a propeller), at least one compressor (such as a low pressure compressor 12 and a high pressure compressor 13), a combustion chamber 14, at least one turbine (such as a high pressure turbine 15 and a low pressure turbine 16), and optionally a gas exhaust nozzle 16.

[0062] The blower 11 allows the intake of an airflow F which divides into a primary flow Fl and a secondary flow F2. The primary flow Fl passes through a primary channel of the turbomachine engine 1 while the secondary flow F2 is directed towards a secondary channel extending around the primary channel.

[0063] Typically, the primary flow Fl is compressed within the low-pressure compressor 12 and then the high-pressure compressor 13. The compressed air is then mixed with a fuel and burned within the combustion chamber 14. The gases formed by the combustion pass through the high-pressure turbines 15 and low-pressure turbines 16. The gases finally escape through the nozzle 17, the cross-section of which allows the acceleration of these gases to generate propulsion.

[0064] In the example of [Fig. 5], the blower 11 is, for example, enclosed. It is surrounded by an annular housing centered on the X-axis. The housing is, for example, surrounded by a nacelle (not shown in [Fig.5]) of the turbomachine 10. Alternatively, the turbomachine 10 may include an unfaired fan 11.

[0065] The turbomachine 10 may include at least one rotor 1 or at least one rotor blade 2.

[0066] The rotor 1 may include at least one rotor blade 2. Advantageously, the rotor 1 may include a disk carrying on its periphery an annular row of blades 2 extending, for example, around the X axis.

[0067] The rotor 1 can be mobile around an axis of rotation coaxial with the X axis of the turbomachine 10.

[0068] The rotor 1 can be a rotor of the low pressure turbine 16.

[0069] Such a rotor blade 1 is illustrated by way of example in [Fig.6].

[0070] The blade 2 can be intended to equip the low pressure turbine 16 and / or the high pressure turbine 15.

[0071] The rotor blade 2 can be made of metal or metal alloy. By way of example, the blade 2 can be made of nickel alloy or titanium aluminide (TiAl).

[0072] The blade 1 comprises an aerodynamic blade 2 extending along an axis A of lengthening (or otherwise stacking) from a foot 4 to a heel 5. The axis A may be substantially perpendicular to the axis X of turbomachine 10.

[0073] The blade 3 comprises a leading edge 36, a trailing edge 38, an intrados face 32 and an extrados face 34. The intrados faces 32 and extrados faces 34 are connected to each other by the leading edge 36 and trailing edge 38.

[0074] In the example of [Fig.6], foot 4 can extend inwards and heel 5 can extend outwards relative to axis A.

[0075] The foot 4 of the blade 2 may include a bulb 40 configured to be engaged in a complementary cavity made in the rotor disc 1.

[0076] Figures 7 and 8 illustrate an example of a blade heel 5 of the blade 2. The blade heel 5 has a platform 50. This platform 50 includes an upstream spoiler 23 located on the leading edge side 36 and a downstream spoiler 24 located on the trailing edge side 38. The upstream spoilers 53 and downstream spoilers 54 can extend outwards (for example along an axis perpendicular to axis A) from the platform 50.

[0077] The heel 5 may include an upstream scraper 51 and a downstream scraper 52. These scrapers 51, 52 may extend radially outwards from the platform 50 (i.e., along axis A). The upstream scrapers 51 and downstream scrapers 52 may be designed to cooperate by complementary shape with an annular block of abradable material (for example, a honeycomb structure) fixed to an external housing surrounding the rotor 1, to form a labyrinth-type seal.

[0078] The platform 50 includes lateral edges (or in other words flanks) 55, 56 located, respectively, on the side of the intrados 32 and extrados 34 faces. These lateral edges 55, 56 can be configured to cooperate by interlocking with complementary lateral edges 56, 57 of adjacent blades 2 of the rotor 1.

[0079] In particular, the platform 50 may include a first lateral edge 55 and a second lateral edge 56. The first lateral edge 55 may be located on the side of the intrados face 32, and the second lateral edge 56 may be located on the side of the extrados face 34.

[0080] Each of the lateral edges 55, 56 has an anti-wear coating 60, 70. In particular, the first lateral edge 56 may include a first anti-wear coating 60, and the second lateral edge 56 may include a second anti-wear coating 70.

[0081] Advantageously, each of the lateral edges 55, 56 may include a portion 6, 7 provided with the corresponding wear-resistant coating 60, 70. For example, the first lateral edge 55 may have a first portion 6 provided with the first wear-resistant coating 60; and the second lateral edge 56 may have a second portion 7 provided with the second wear-resistant coating 70.

[0082] Each of the lateral edges 55, 56 can have a substantially "V" shaped profile and is thus defined as the portion 6, 7 provided with the wear-resistant coating 60, 70.

[0083] The lateral edges 55, 56 may each have at least one part, such as the portions 6, 7 provided with the wear-resistant coatings 60, 70, which may be parallel to each other.

[0084] Preferably, the first 60 and second 70 wear-resistant coatings can be made of the same material.

[0085] Each of the wear-resistant coatings 60 and 70 can be made of a cobalt and / or tungsten-based alloy. Preferably, the first 60 and second 70 wear-resistant coatings can be made of a cobalt-based alloy, this type of material being, for example, marketed under the brand name Stellite®.

[0086] Each of the wear-resistant coatings 60, 70 has a contact face 62, 72. Each contact face 62, 72 is able to come into contact with another contact face 62, 72 of a wear-resistant coating 60, 70 of another blade 2 of the rotor 1. In particular, the first wear-resistant coating 60 may include a first contact face 62, and the second wear-resistant coating 70 may include a second contact face 72.

[0087] Each of the contact faces 62, 72 can have a contact surface greater than 6 mm2. In particular, the contact surfaces of the contact faces 62, 72 are identical.

[0088] Each of the contact faces 62, 72 can have a width 162, 172 (expressed along the axis A) and a length L62, L72 (expressed along a plane inclined with respect to the axis A passing in particular through the portion 6, 7).

[0089] According to one of the features of the invention, the contact face 62, 72 of the wear-resistant coating 60, 70 of each of the lateral edges 55, 56 comprises teeth 8, 86, 87. These teeth 8, 86, 87 are configured to cooperate by interlocking with complementary teeth 8, 86, 87 of the other contact face 62, 72 of the other rotor blade 2. As mentioned above, this makes it possible to limit radial, axial, and / or lateral displacements and thus prevent the blade 2 from disengaging during operation.

[0090] In particular, the first contact face 62 may include first teeth 86, and the second contact face 72 may include second teeth 87 which are complementary to the first teeth 86, in particular to the first contact face 62 of the other rotor blade 2.

[0091] Each of the first teeth 86 of the first contact face 62 can have a size (such as a length, a width or a thickness) identical to that of the second teeth 87 of the second contact face 72.

[0092] Alternatively, the size of each of the first teeth 86 of the first contact face 62 may be different from that of the second teeth 87 of the second contact face 72. This may strengthen the interlocking between the teeth and limit the disengagement of the blade during operation.

[0093] In particular, the size of each of the first teeth 86 of the first contact face 62 can be smaller than that of the second teeth 87 of the second contact face 72. This makes it possible to limit the addition of stresses at the first contact face 62 of the heel, and to rebalance the stresses between the first 62 and second 72 contact faces. Indeed, this first contact face 62 located on the side of the intrados face 32 can be subjected to a greater mechanical load, since it can be subjected to more bending stresses during operation compared to the second contact face 72 located on the side of the extrados face 34.

[0094] Each of the anti-wear coatings 60, 70 can be formed in a monobloc fashion (i.e. from material) with these teeth 8, 86, 87.

[0095] The number of teeth 8, 86, 87 on each contact face 62, 72 can be between two and twenty. The teeth 86, 87 in the example of Figures 7 and 8 are, without limitation, twelve in number.

[0096] The teeth 8, 86, 87 may have bearing faces 860, 870 extending in the same plane P, particularly when the teeth 8, 86, 87 of at least two adjacent blades 2 are fitted together. This plane P may pass through the corresponding lateral edge 55, 56 and be inclined with respect to the axis A. The faces Supports 860, 870 can be configured to bear against hollows 85, 88 of the contact faces 62, 72 of the other blade 2. This allows a plane-to-plane contact to be formed between two adjacent nested blades 2, without generating local stresses.

[0097] The teeth 8, 86, 87 can have in section (in particular planar with respect to plane P) a shape chosen from among a rectangle, a square, a polygon, a parabola, a groove, a chevron and a helix.

[0098] In particular, the first 86 and / or second 87 teeth may present in section one of the aforementioned shapes.

[0099] The cross-sectional shape of each of the first teeth 86 of the blade 2 may be identical or different from that of each of the second teeth 87 of the same blade 2.

[0100] Advantageously, the teeth 8, 86, 87 of each contact face 62, 72 can be separated by hollows 85, 88.

[0101] The recesses 85, 88 may have the same length and / or width as the teeth 86, 87. Alternatively, the length and / or width of the recesses 85, 88 may be different from that of the teeth 86, 87. In particular, the length and / or width of the recesses 85, 88 may be greater than that of the teeth 86, 87. This allows the blade 2 to have more freedom of movement during operation.

[0102] The teeth 8, 86, 87 of each contact face 62, 72 can be arranged in a matrix pattern in rows and columns 80 of teeth 8, 86, 87 (Figures 7, 8). This allows for the formation of a complex grid-like contact (or, in other words, an interlocking) between the teeth of at least two adjacent rotor blades. This complex contact facilitates the interlocking of adjacent blades during assembly, while limiting their disengagement during operation.

[0103] Advantageously, the teeth 8, 86, 87 can be staggered between the rows and / or between the columns 80.

[0104] The teeth 8, 86, 87 can be produced by machining the wear-resistant coatings 60, 70, such as by die-sinking electrical discharge machining (EDM) or by laser. Alternatively, the teeth 8, 86, 87 can be produced by casting or by additive manufacturing.

[0105] Figures 7 to 13 illustrate several examples of the rotor blade 2 according to the invention.

[0106] With reference to Figures 7 to 11, the present application will now describe the rotor blade 2 according to a first example, and the rotor 1 comprising such a blade 2.

[0107] The blade 2 comprises the first lateral edge 55 having the first wear-resistant coating 60 having the first contact face 62 with the first teeth 86, and the second lateral edge 56 having the second wear-resistant coating 70 having the second face of contact 72 with the second teeth 87 which are complementary to the first teeth 86.

[0108] Each of the first 60 and second 70 wear-resistant coatings is formed in a single piece with the corresponding first 86 and second 87 teeth.

[0109] The first 860 and second 870 bearing faces of the first 86 and second 87 teeth extend in the same plane P, in particular when the first 86 and second 87 teeth of at least two adjacent blades 2 are nested with each other.

[0110] The first 86 and second 87 teeth have a rectangular shape in section.

[0111] The first 86 and second 87 teeth are distributed in a matrix-like fashion in rows and columns of 80 teeth. These teeth 86, 87 are arranged in a staggered pattern between the rows and columns of 80 teeth.

[0112] The first 86 and second 87 teeth are twelve in number. In particular, each column 80 of teeth of the first 62 and second 72 contact faces can comprise two teeth 86, 87 which are separated from each other by two hollows 85 ([Fig.9]).

[0113] Figures 10 and 11 illustrate a cyclic (or in other words circumferentially around the X axis) interlocking of two rotor blades 2, the second teeth 87 of the second contact face 72 of one of the two blades 2 cooperating by interlocking with the first teeth 86 of the first contact face 62 of the other of the two blades 2.

[0114] Figures 12 and 13 illustrate a second example of the rotor blade 2 according to the invention. This blade 2 of the second example differs from the blade 2 of the first example by the teeth 8, 86, 87. Indeed, the first 86 and / or second 87 teeth have a fluted (or in other words, gear-like) cross-section.

Claims

Demands

1. Rotor blade (2) for an aircraft turbomachine (10), said blade (2) comprising an aerodynamic blade (3) extending along an aspect ratio axis (A) from a root (4) to a tail (5), the blade (3) comprising lower (32) and upper (34) surfaces connected by leading (36) and trailing (38) edges, said tail (5) comprising a platform (50) comprising an upstream spoiler (53) located on the leading edge (36) side, a downstream spoiler (54) located on the trailing edge (38) side, and side edges (55, 56) located respectively on the lower (32) and upper (34) surfaces, each of the side edges (55, 56) comprising a wear-resistant coating (60, 70) comprising a contact face (60, 70) adapted to come into contact with another face of contact (62, 72) of a wear-resistant coating (6, 7) of another rotor blade (2), characterized in that the contact face (62, 72) of the wear-resistant coating (60, 70) of each of the lateral edges (55, 56) comprises teeth (8,86, 87) which are configured to cooperate by interlocking with complementary teeth (8, 86, 87) of said other contact face (62, 72).

2. Rotor blade according to claim 1, characterized in that a first (55) of the lateral edges (55, 56) comprises a first anti-wear coating (60) having a first contact face (62) with first teeth (86), and a second (56) of the lateral edges (55, 56) comprises a second anti-wear coating (70) having a second contact face (72) with second teeth (87) which are complementary to the first teeth (86).

3. Rotor blade according to claim 1 or 2, characterized in that said teeth (8, 86, 87) of each contact face (62, 72) are distributed in a matrix manner in rows and columns of teeth (8, 86, 87).

4. Rotor blade according to any one of the preceding claims, characterized in that said teeth (8, 86, 87) have in section a shape selected from a rectangle, a square, a polygon, a parabola, a groove, a chevron and a helix.

5. Rotor blade according to claim 3 or 4, characterized in that said teeth (8, 86, 87) are staggered between the rows and / or between the columns.

6. Rotor blade according to any one of the preceding claims, characterized in that each of the wear-resistant coatings (60, 70) is formed in a single piece with its teeth (8, 86, 87).

7. Rotor blade according to any one of the preceding claims, characterized in that each of the wear-resistant coatings (60, 70) is made of cobalt and / or tungsten-based alloy.

8. Rotor blade according to any one of the preceding claims, characterized in that the number of teeth (8, 86, 87) on each contact face (62, 72) is between two and twenty.

9. Rotor blade according to any one of the preceding claims, characterized in that the teeth (8, 86, 87) of each contact face (62, 72) are separated by grooves (85, 88).

10. Rotor blade according to any one of the preceding claims, characterized in that the teeth (8, 86, 87) have bearing faces (860, 870) extending in the same plane (P).

11. Rotor (1) for an aircraft turbomachine (10), comprising a disc carrying on its periphery an annular row of blades (2) according to any one of the preceding claims.

12. Aircraft turbomachine (10), comprising at least one rotor blade (2) according to any one of claims 1 to 10 or at least one rotor (1) according to claim 11.

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