ROTOR BLADE FOR AN AIRCRAFT TURBOMACHINE

Abstract

A rotor blade (100) for an aircraft turbomachine (10), this blade (100) comprising two platforms, respectively inner (102) and outer (104), between which extends a blade (106) along an axis of extension (A), the blade (106) comprising leading (108) and trailing (110) edges and lower (112) and upper (114) surfaces extending between the leading (108) and trailing (110) edges, the inner platform (102) being connected to a foot (116), the outer platform (104) being connected to at least one sealing flap (118) projecting from the outer platform (104), characterized in that the flap or flaps (118) comprise a cutting edge (130) and a cutting face (132) extending from the cutting edge (130). Figure for the abbreviation: Figure 4

Description

Title of the invention: Rotor blade for an aircraft turbomachine Technical field of the invention

[0001] The present invention relates in particular to a rotor blade for an aircraft turbomachine. Technical background

[0002] An aircraft turbomachine 10, such as that shown in [Fig.l], generally comprises at least one compressor 12, 14, an annular combustion chamber 16 and at least one turbine 18, 20. The air which enters the compressor 12, 14 is compressed then mixed with fuel and burned in the combustion chamber 16. The combustion gases are then expanded in the turbine 18, 20, which makes it possible to drive the rotor of the turbine in rotation which in turn drives the rotor of the compressor.

[0003] There are several turbomachine technologies, such as turboprops or turbojets, which comprise for example a propulsion propeller 22 located upstream of the turbomachine 10 and which is also driven by the rotor of the turbine, this propeller being able to be shrouded or unshrouded.

[0004] There are also several turbomachine configurations, such as single- or multi-spool turbomachine. In a twin-spool turbomachine as illustrated in [Fig.l], the turbomachine 10 comprises a low-pressure LP spool and a high-pressure HP spool. The low-pressure LP spool comprises a low-pressure compressor rotor 12 connected by a low-pressure shaft 24 to a low-pressure turbine rotor 18. The high-pressure spool 14 comprises a high-pressure compressor rotor connected by a high-pressure shaft 26 to a high-pressure turbine rotor 20. In the direction of gas flow in the turbomachine, from upstream to downstream (see arrows), the twin-spool turbomachine 10 then comprises the low-pressure compressor 12, the high-pressure compressor 14, the combustion chamber 16, the high-pressure turbine 18 and the low-pressure turbine 20.

[0005] A turbomachine compressor 12, 14 comprises one or more compression stages which extend around and along the same axis X, and which each comprise a rotor wheel 28 and a stator blade 30 (also called a rectifier blade - [Fig.2]).

[0006] Similarly, a turbomachine turbine comprises one or more expansion stages which extend around and along the same axis X, and which each comprise a rotor wheel 28 and a stator blade 30 (also called a distributor blade - [Fig.2]).

[0007] A rotor wheel 28, for a compressor or a turbine, generally comprises a disk 32 carrying blades 34 at its periphery. The disk 32 is centered on the axis X of rotation of the wheel 28 and the blades 34 are distributed around this axis X and mounted at the periphery of the disk 32.

[0008] A rotor blade 34 comprises two platforms, respectively internal 36 and external 38, between which a blade 40 extends. The blade 40 comprises leading edges 42 and trailing edges 44 and intrados 46 and extrados 48 faces which extend between the leading and trailing edges. The leading edge 42 is located upstream and the trailing edge 44 is located downstream with reference to the gas flow (see arrows).

[0009] The internal platform 36 is connected to a foot (not shown) which allows it to be mounted on the disc 32. The external platform 38 is connected to at least one sealing lip 50 which projects from the external platform 38 and which is intended to cooperate by friction with an annular coating 52 made of abradable material. This coating 52 surrounds the rotor wheel 28 and is carried by a casing 54 which also carries the stator vane 30 located downstream of the wheel 28.

[0010] Similarly, the stator blade 30 comprises vanes 56 distributed around the X-axis. Each of these vanes 56 extends between internal 58 and external 60 platforms. The external platforms 60 of the vanes 56 are generally fixed to the housing 54 by hooking, and the internal platforms 58 of the vanes carry scuffs 62 that cooperate by friction with an annular coating 64 made of abradable material. This coating 64 is carried by the rotor, of which the wheel 28 is a part.

[0011] The blades 34 and 56 have radial orientations relative to the axis X, that is to say that they each comprise an elongation axis A which extends substantially radially relative to the axis X. The platforms 36, 38, 58, 60 form annular assemblies which extend around the axis X and are generally sectorized to facilitate the mounting of the wheel 28 and the blading 30. The wipers 50, 62 also form annular assemblies which are also sectorized. The wipers 50 extend radially outwards and the wipers 62 extend radially inwards.

[0012] The cooperation of the wipers 50, 62 with the coatings 52, 62 makes it possible to form seals by chicane effect and thus to limit the passage of gas in operation between the rotor and the stator, respectively at the external periphery of the wheel 28 and the internal periphery of the blading 30.

[0013] Originally, each of the coatings 52, 62 has a cylindrical or stepped shape, that is to say it comprises one or more cylindrical surfaces. During a running-in phase of the turbomachine, the wipers 50, 62 rub on the coatings and form annular grooves therein.

[0014] In some cases, the temperatures reached and the consumption of games during and after operation can cause a rotor blocking phenomenon, called “rotor lock”, resulting from a deformation of the wiper / abradable coating assembly which can lead to friction / diffusion welding of these elements.

[0015] The invention provides a solution to this problem, which is simple, effective and economical. Summary of the invention

[0016] The invention relates to a rotor blade for an aircraft turbomachine, this blade comprising two platforms, respectively internal and external, between which a blade extends along an axis of elongation, the blade comprising leading and trailing edges and intrados and extrados faces which extend between the leading and trailing edges, the internal platform being connected to a root, the external platform being connected to at least one sealing lip which projects from the external platform, the or each lip comprising two faces situated respectively on the leading edge side and on the trailing edge side, these faces being connected together by an upper edge situated on the side opposite the blade, and two lateral edges situated respectively on the side of the intrados and extrados faces of the blade, characterized in that the or each lip comprises at the junction between the upper edge and one of the lateral edges,a cutting edge and a cutting face extending from the cutting edge along a portion of that side edge.

[0017] The distinctive feature of the blade according to the invention is that its cutting edge(s) is / are designed as a cutting tool. In other words, a cutting function, and not merely a friction function, is added to the blade's cutting edge(s). It is thus understood that the blade's cutting edge is capable of machining an abradable material by cutting, and not just by friction. It is further understood that this facilitates the machining of the abradable material and, in particular, reduces the forces required for this machining. The risk of overheating and welding of the cutting edges to the abradable material is reduced, thereby eliminating or limiting the risk of the aforementioned blockage phenomenon.

[0018] The blade according to the invention may comprise one or more of the following characteristics, taken in isolation from one another, or in combination with one another: • said cutting face forms a cutting angle between 95 and 150° with a first plane perpendicular to the elongation axis and passing through said upper edge, this angle being measured in a second plane perpendicular to this first plane and parallel to the lick; • the or each lip further comprises a relief face which extends from the cutting edge along a portion of the upper edge; • said draft face forms a draft angle between 1 and 20° with a first plane perpendicular to the axis of elongation and passing through said upper edge, this angle being measured in a second plane perpendicular to this first plane and parallel to the lick; • the or each lip further comprises an undercut face which extends from the undercut face along another portion of the upper edge; • said cutting face is located in a notch of said side edge; • said relief face is located in a notch of said upper edge; • the or each lick is covered with a hardening coating, such as nitriding, and for example a titanium or boron nitride, and in particular the edge and the cutting face are covered with this coating; • the or each lip is formed by a ceramic insert which is mounted on the external platform, preferably by a male-female fitting, for example of the dovetail type; • the external platform includes a hole for mounting an element for locking the insert with respect to the platform.

[0019] The present invention also relates to a turbine for an aircraft turbomachine, comprising several blades as described above, which are arranged next to each other and around a common axis of rotation of these blades, the turbine further comprising an abradable annular coating which surrounds the blades and in which the blade tips can cooperate by friction. Brief description of the figures

[0020] Other characteristics and advantages will emerge from the following description of a non-limiting embodiment of the invention with reference to the appended drawings in which:

[0021] [Fig-1] [Fig.l] is a schematic axial sectional view of a turbomachine aircraft,

[0022] [Fig.2] [Fig.2] is a semi-schematic axial sectional view of a compressor or turbine stage of an aircraft turbomachine,

[0023] [Fig.3] [Fig.3] is a schematic perspective view of a rotor blade according to the invention,

[0024] [Fig.4] [Fig.4] is an enlarged view of part of [Fig.3],

[0025] [Fig.5] [Fig.5] is a larger scale view of the external platform and the licks of dawn of the [Fig.3],

[0026] [Fig.6] [Fig.6] is an enlarged view of part of [Fig.5],

[0027] [Fig.7] [Fig.7] is a larger scale view of one of the dawn licks of the [Fig.3],

[0028] [Fig.8] [Fig.8] is a schematic view of a portion of the blade of [Fig.3], as well as a portion of an abradable coating which surrounds this blade,

[0029] [Fig.9] [Fig.9] is a schematic view similar to that of [Fig.8] and shows several blades and the abradable coating surrounding these blades,

[0030] [Fig. 10] [Fig. 10] is a schematic perspective view of an alternative embodiment of a rotor blade according to the invention, and

[0031] [Fig. 11] [Fig. 11] is an enlarged view of part of [Fig.3]. Detailed description of the invention

[0032] Figures 1 and 2 have already been described in the above.

[0033] [Fig. 3] illustrates a first embodiment of a rotor blade 100 according to the invention for an aircraft turbomachine 10, such as that illustrated in [Fig. 1].

[0034] The blade 100 is similar to the blade 34 of [Fig.2] and comprises two platforms, respectively internal 102 and external 104, between which a blade 106 extends along an axis of elongation A.

[0035] The blade 106 has leading edges 108 and trailing edges 110 and intrados 112 and extrados 114 faces which extend between the leading edges 108 and trailing edges 110.

[0036] The internal platform 102 is connected to a foot 116, and the external platform 104 is connected to at least one sealing lip 118 which projects from the external platform 102. The number of lips 118 is here two in the example shown. The or each lip 118 extends substantially parallel to the axis A.

[0037] Each wiper 118 comprises two faces 120, 122 located respectively on the side of the leading edge 108 and on the side of the trailing edge 110, that is to say on the upstream side and on the downstream side. Each wiper 118 thus comprises an upstream face 120 and a downstream face 122. These faces 120, 122 can be substantially parallel to the axis A.

[0038] These faces 120, 122 are connected together by an upper edge 124 located on the side opposite the blade 106, and two lateral edges 126, 128 located respectively on the side of the intrados 112 and extrados 114 faces of the blade 106.

[0039] Depending on the direction of rotation of the blade 100 around its axis of rotation (which is the X axis in FIGS. 1 and 2), one of these edges, namely the edge 126 in the example shown, is an upstream edge (with reference to this direction of rotation and no longer with reference to the flow of the air stream), and the other of these edges, namely the edge 128 in the example shown, is a downstream edge.

[0040] The particularity of the blade 100 is that at least one of its wipers 118 (and preferably both) comprises (each comprise) at the junction between the upper edge 124 and one of the lateral edges 126, 128, a cutting edge 130 and a cutting face 132 which extends from the cutting edge 130 along a part of this edge lateral 126, 128.

[0041] The cutting edge 130 is intended to cut the abradable coating that surrounds it and is therefore located upstream with reference to the direction of rotation of the blade 100. It is therefore understood that the cutting edge 130 is located on the edge 126 in the example shown.

[0042] The cutting edge 130 has a generally rectilinear shape and extends in the example shown over the entire thickness of the lip 118 or at least of its upper edge 124.

[0043] The cutting face 132 forms a cutting angle α between 95 and 150° with a first plane PI perpendicular to the elongation axis A and passing through said upper edge 124. This angle α is measured in a second plane P2 perpendicular to the first plane PI and parallel to the slit 118 (cf. [Fig.6]).

[0044] The cutting face 132 is preferably planar. It has a height or dimension along the axis A which represents for example between 10 and 30% of the height or dimension along the axis A of the lip 118.

[0045] [Fig.7] shows that the cutting face 132 can be inclined relative to a plane P3 passing through the axis X of rotation of the blade and through the lateral edge 126 of the wiper 118, on which the cutting face 132 is located. The angle of inclination is noted y and is measured in the plane PI for example.

[0046] The wiper 118 further includes a relief face 134 that extends from the cutting edge 130 along a portion of the upper edge 124.

[0047] The clearance face 134 forms a clearance angle [3 of between 1 and 20° with the first plane PI perpendicular to the elongation axis A and passing through the upper edge 124. This angle [3 is measured in the second plane P2 perpendicular to the first plane PI and parallel to the lip 118 (see [Fig.6]).

[0048] The relief face 134 is preferably planar. It has a length or dimension along the lip, in a direction transverse to the axis A, which represents for example between 2 and 10% of the length or dimension of the lip 118 in the same direction.

[0049] The lick 118 further includes an undercut face 136 which extends from the undercut face 134 along another part of the upper edge 124.

[0050] The undercut face 136 is preferably flat.

[0051] The cutting face 132 can be located in a notch 138 of the lateral edge 126, as in the example shown ([Fig.6]).

[0052] The relief face 136 may be located in a notch 140 of the upper edge, as in the example shown ([Fig.6]).

[0053] Advantageously, the wiper 118 is covered with a hardening coating 142, such as nitriding, and for example a titanium or boron nitride. In particular, it is the edge 130 and the cutting face 132 at least, or even also the clearance faces 134 and undercut faces 136, which are covered with this coating 142. (see figures 5-7).

[0054] Figures 8 and 9 show the cooperation of the wiper 118 and in particular of the cutting edge 130 and the cutting face 132 with the coating 52 of abradable material, as illustrated in [Fig.2]. The cutting of the abradable material by the wiper 118, which forms a cutting tool, makes it possible to reduce the forces at the interface between these elements in operation and thus to limit the risk of “rotor lock”, as mentioned above.

[0055] In the alternative embodiment of Figures 10 and 11, at least one of the wipers 118 (and preferably both) is formed (are each formed) by a ceramic insert 144 which is mounted on the external platform 104, preferably by a male-female fitting, for example of the dovetail type.

[0056] Y denotes the axis of insertion of the insert 144 into the corresponding housing of the platform 104. The dovetail has an elongated shape along this axis Y. The platform 104 preferably includes an opening 146 that leads into the housing of the insert 144 and is intended to receive a locking or securing element for the insert 144 along the axis Y, in order to prevent it from coming out of its housing. This element is, for example, a pin, a screw, or a rivet.

Claims

Claims

1. Rotor blade (100) for an aircraft turbomachine (10), this blade (100) comprising two platforms, respectively internal (102) and external (104), between which a blade (106) extends along an elongation axis (A), the blade (106) comprising leading (108) and trailing (110) edges and intrados (112) and extrados (114) faces which extend between the leading (108) and trailing (110) edges, the internal platform (102) being connected to a root (116), the external platform (104) being connected to at least one sealing lip (118) which projects from the external platform (104), the or each lip (118) comprising two faces (120, 122) located respectively on the side of the leading edge (108) and on the trailing edge side (110), these faces (120, 122) being connected together by an upper edge (124) located on the side opposite the blade (106), and two lateral edges (124,126) located respectively on the side of the intrados (102) and extrados (104) faces of the blade (106), characterized in that the or each wiper (118) comprises at the junction between the upper edge (124) and one of the lateral edges (126), a cutting edge (130) and a cutting face (132) which extends from the cutting edge (130) along a part of this lateral edge (126).,

2. Blade (100) according to claim 1, wherein said cutting face (132) forms a cutting angle (a) of between 95 and 150° with a first plane (PI) perpendicular to the axis of elongation (A) and passing through said upper edge (124), this angle (a) being measured in a second plane (P2) perpendicular to this first plane (PI) and parallel to the lip (118).

3. A blade (100) according to claim 1 or 2, wherein the or each wiper (118) further comprises a relief face (134) which extends from the cutting edge (130) along a portion of the upper edge (124).

4. Blade (100) according to claim 3, in which said clearance face (134) forms a clearance angle (|3) of between 1 and 20° with a first plane (PI) perpendicular to the axis of elongation and passing through said upper edge (124), this angle (|3) being measured in a second plane (P2) perpendicular to this first plane (PI) and parallel to the lip (118).

5. A blade (100) according to claim 3 or 4, wherein the or each lip (118) further comprises an undercut face (136) which extends from the relief face (134) along another portion of the upper edge (124).

6. A blade (100) according to any preceding claim, wherein said cutting face (132) is located in a notch (138) of said side edge (126).

7. A blade (100) according to one of claims 3 to 5, wherein said relief face (134) is located in a notch (140) of said upper edge (124).

8. Blade (100) according to one of the preceding claims, in which the or each wiper (118) is covered with a hardening coating (142), such as nitriding, and in particular the edge (130) and the cutting face (132) are covered with this coating (142).

9. Blade (100) according to one of the preceding claims, in which the or each wiper (118) is formed by a ceramic insert (144) which is mounted on the external platform (104), preferably by a male-female fitting, for example of the dovetail type.

10. Blade (100) according to the preceding claim, in which the external platform (104) comprises an orifice (146) for mounting an element for locking the insert (144) with respect to the platform (104).

11. Turbine for an aircraft turbomachine (10), comprising several blades (100) according to one of the preceding claims, which are arranged next to each other and around a common axis (X) of rotation of these blades (100), the turbine (18, 20) further comprising an abradable annular coating (52) which surrounds the blades (100) and in which the wipers (118) of the blades (100) can cooperate by friction.

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