ROTOR BLADE FOR A TURBOMACH

Description

Technical field of the invention

[0001] The present invention relates to a rotor blade for a turbomachine, in particular an aircraft one. The field of the invention is that of turbomachine modules comprising a rotor equipped with several blades, and in particular modules of the turbine or compressor wheel type, mixed and / or centrifugal impellers, axial, mixed and centrifugal compressors, as well as modules comprising inlet guide vanes. Technical background

[0002] The state of the art includes, in particular, documents USA-6027306, FR-A5-2054402. EP-A-2309098, US-A1-2010 / 166566, EP-A2-0916811 and EP-A1-1980753. In the subsonic or supersonic compression stages of turbomachines, blades are used to compress the fluid. Such a turbomachine blade conventionally comprises a blade comprising a lower surface and an upper surface extending from a leading edge to a trailing edge. However, by design, there is a clearance between the blade head and the casing or cover. This clearance allows the fluid to pass from the high pressure side (lower surface) to the low pressure side (upper surface), and is the source of secondary flows. However, such secondary flows disrupt the healthy flow and reduce the performance and compression efficiency of the compression stages.

[0003] There is therefore a need to have a rotor blade for a turbomachine, making it possible to reduce secondary flows due to the operating clearance between the blade head and the casing or cover, while improving the aerodynamic performance of the blade.

[0004] The present invention aims in particular to resolve all or part of the aforementioned problems. Summary of the invention

[0005] The present invention proposes for this purpose a rotor blade for a turbomachine, in particular for an aircraft, comprising a blade comprising a lower surface and an upper surface extending from a leading edge to a trailing edge, the blade comprising an axis of elongation extending substantially along the leading and trailing edges, the blade comprising a radially internal end for connection to a rotor and a free radially external end.

[0006] According to the invention, the blade further comprises at least one series of fins located on said free end, each of these fins comprising a lower surface located on the side of the upper surface of the blade, and an upper surface located on the side of the lower surface of the blade, and is characterized in that each of the fins is twisted around this axis of elongation.

[0007] The invention thus makes it possible, thanks to such a configuration of the fins, to create a pressure gradient opposed to the circulation of the clearance flow which circulates from the pressure face (intrados) of the blade to the depression face (extrados) of the latter. Indeed, such fins whose curvature is reversed with respect to that of the blade act as screens by preventing the fluid from being able to circulate from the pressure face (intrados) to the depression face (extrados) in the clearance zone. This makes it possible to increase the performance (flow rate, rate) and the isentropic efficiency by minimizing the clearance flows at the blade tip. The opposing pressure gradient is created through the choice of the curvature of the fins. In addition, the presence of such fins thus configured makes it possible to obtain an aerodynamic gain by controlling the vortex at the blade tip, which contributes to improving the performance and / or stability of the module comprising the blade.Such an aerodynamic gain becomes essential, for example, when the blades are very small. Finally, the present invention makes it possible to obtain a gain in stability for axial compressors, as well as an increase in the total pressure rate of the rotor.

[0008] 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: the fins of the or each row are arranged one behind the other along a chord of the free end of the blade; the blade comprises a single series of fins; the cumulative lengths of the chords of the fins represent between 80% and 120% of a chord of the free end of the blade; the number of fins of the or each series is between 2 and 10, and preferably between 4 and 8, in particular for an axial compressor; - the number of fins of the or each series is a function of the chord of the blade, in particular for a centrifugal compressor; each of the fins has a height measured along said axis of elongation which represents between 1 and 10%, and preferably between 1 and 5%, of the height of the blade measured along the same axis; and each of the fins is inclined relative to said axis of elongation.

[0009] The invention also relates to a turbomachine module, in particular an aircraft module, this module comprising a rotor comprising an annular row of blades as described above.

[0010] The turbomachine module 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: the turbomachine module is chosen from a turbine or compressor wheel, a mixed or centrifugal impeller, and a fan; and the rotor is surrounded by an annular coating of abradable material which is housed in a groove of an annular casing, this groove being formed in an internal cylindrical surface of the casing which has an internal diameter substantially equal to the diameter of the rotor passing through the free ends of the blades of its vanes.

[0011] The invention also relates to an aircraft turbomachine, comprising at least one blade as described above. Brief description of the figures

[0012] Other characteristics and advantages of the invention will appear during the reading of the detailed description which follows for the understanding of which reference will be made to the appended drawings in which: [ Fig. 1 ] there figure 1 is a schematic perspective view of a turbomachine rotor blade, and in particular of a turbomachine fan; [ Fig.2 ] there figure 2 is a perspective view of a turbomachine module provided with a rotor comprising an annular row of blades according to the invention; [ Fig.3 ] there figure 3 is an enlarged view of part of the module of the figure 2 , showing one of the rotor blades, the blade comprising a series of vanes; [ Fig.4 ] there figure 4 is a top view of the free end of the blade of the figure 3 ; [ Fig.5 ] there Figure 5 is a meridian sectional view of the dawn of the figure 3 ; [ Fig.6] there figure 6 is an enlarged view of part of the dawn of the Figure 5 , showing three of the vanes of the dawn; [ Fig.7 ] there figure 7 is a side view of one of the fins of the Figure 5 ; [ Fig.8 ] there figure 8 is a meridian sectional view of a spinning wheel blade. Detailed description of the invention

[0013] There figure 1 shows a rotor blade 10 of a turbomachine and in particular of a fan of an aircraft turbomachine. This blade 10 comprises a blade 12 which has a leading edge 14 and a trailing edge 16, opposite the leading edge.

[0014] The blade 12 has an aerodynamic profile and comprises a lower surface 15 and an upper surface 17 extending between the leading 14 and trailing 16 edges of the blade. The blade 10 has an elongation axis denoted A. In the present application, the term “the elongation axis of the blade” means the axis which extends substantially along the leading and trailing edges 14, 16 of the blade 12, and in particular between the leading and trailing edges.

[0015] A longitudinal end of the blade 12 is free and the opposite longitudinal end is for example connected to a foot 18 for fixing the blade to a rotor of the turbomachine.

[0016] THE figures 2 and following illustrate embodiments of the invention.

[0017] There figure 2 shows a module 20 provided with a rotor comprising an annular row of blades 10 according to the invention. Preferably, and as illustrated in the figure 6, the module 20 also comprises an annular coating 22 made of abradable material which is housed in a groove 23 of an annular casing 25 (the casing 25 being visible on the figure 2 ). The groove 23 is formed in an internal cylindrical surface of the casing 25 which has an internal diameter D1 substantially equal to the diameter of the rotor passing through the free ends of the blades 12 of its vanes 10. The module 20 is for example a turbine or compressor wheel, a mixed or centrifugal impeller, or even a fan. An axial compressor has an axial inlet and an axial outlet, a mixed impeller has an axial inlet and an inclined outlet, and a centrifugal impeller has an axial inlet and a radial outlet.

[0018] THE Figures 3 and 4show the free end of one of the blades 10 according to the invention. This free end shows the aerodynamic profile of the blade 12. In the present application, the term “chord” means the imaginary straight line which connects the leading edge 14 to the trailing edge 16 of the blade. A chord B can be drawn, as illustrated in the figure 3 , on the free end of the blade 12. Such a chord can however be drawn on any cross-section of the blade made perpendicular to the aforementioned axis of elongation A. The blade 10 comprises at least one series of fins 24 located on the free end of the blade 12. Preferably, the number of fins 24 of the or each series is between 2 and 10, and preferably between 4 and 8. In the particular embodiment illustrated in the figures 3 to 5 , the blade 10 comprises a single series of six vanes 24. Preferably, and as visible on the figures 3 to 5, the fins 24 of each row are arranged one behind the other along the chord B of the free end of the blade 12. The blades 10 can equip an axial compressor, for example of an aircraft turbomachine.

[0019] The fins 24 are preferably attached to the free end of the blade 12, for example via an additive machining process. Additive machining is understood to mean a process comprising a step of depositing material to create a preform of the fins, and a step of machining this preform to create a final shape of the fins.

[0020] As illustrated on the Figures 3 and 4, each of the fins 24 comprises a lower surface 26 located on the side of the upper surface 17 of the blade 12, and an upper surface 28 located on the side of the lower surface 15 of the blade 12. Preferably, the cumulative lengths of the chords of the fins 24 represent between 80% and 120% of the chord B. More preferably, each fin 24 has a height measured along the elongation axis A which represents between 1 and 10%, and preferably between 1 and 5%, of the height of the blade 12 measured along the same axis.

[0021] There figure 6 illustrates a particular embodiment where the fins 24 make their place thanks to the presence of the annular coating 22 made of abradable material which is housed in a groove 23 of the annular casing 25.

[0022] According to a particular embodiment variant illustrated on the figure 7, each fin 24 is inclined relative to the elongation axis A. In this variant, the arrow w indicates the direction of rotation of the blade 10. The fins 24 are inclined towards the side of the blade 10 opposite to the direction of rotation of the latter. According to the invention not shown, each fin 24 is twisted around the elongation axis A.

[0023] In operation, the fins 24 create a pressure gradient opposite to the circulation of the clearance flow which circulates from the intrados 15 of the blade 12 to the extrados 17 of the blade 12. This is due to the fact that the intrados 26 of the fins 24 is turned towards the extrados 17 of the blade 12, and vice versa. Such a configuration of the fins 24 makes it possible to prevent the fluid from being able to circulate from the pressure face (intrados 15) to the depression face (extrados 17) of the blade 12 in the clearance zone. The performances (flow rate, rate) and the isentropic efficiency are thus advantageously increased, while minimizing the clearance flows at the blade tip. The opposing pressure gradient is created through the choice of the curvature of the fins. Furthermore, the presence of such fins 24 thus configured makes it possible to obtain an aerodynamic gain by controlling the vortex at the blade tip, which contributes to improving the performance and / or stability of the module 20 comprising the blade 10.Such an aerodynamic gain becomes, for example, essential when the blades 12 are very small. Finally, the present invention makes it possible to obtain a gain in stability for axial compressors, as well as an increase in the total pressure rate of the rotor.

[0024] There figure 8 illustrates the application of the invention to a centrifugal impeller 30. Each blade 10 of the impeller 30 comprises a number of vanes 24 which may be a function of the chord of this blade.

Claims

1. A rotor vane (10) for a turbomachine, in particular of an aircraft, comprising a blade (12) comprising a pressure face (15) and a suction face (17) extending from a leading edge (14) to a trailing edge (16), the blade (12) comprising an elongation axis (A) extending substantially along the leading (14) and trailing (16) edges, the blade (12) comprising a radially inner end for connection to a rotor and a radially outer free end, the vane (10) further comprising at least one series of fins (24) located on said free end, each of these fins (24) comprising a pressure face (26) located on the side of the suction face (17) of the blade (12), and a suction face (28) located on the side of the pressure face (15) of the blade (12), characterized in that each of the fins (24) is twisted around this elongation axis (A).

2. The vane (10) according to claim 1, wherein the fins (24) of the or each row are arranged one behind the other along a chord (B) of the free end of the blade (12).

3. The vane (10) according to claim 1 or 2, wherein it comprises a single series of fins (24).

4. The vane (10) according to one of the preceding claims, wherein the cumulative lengths of the chords of the fins (24) are between 80% and 120% of a chord (B) of the free end of the blade (12).

5. The vane (10) according to one of the preceding claims, wherein the number of fins (24) in the or each series is between 2 and 10, and preferably between 4 and 8.

6. The vane (10) according to one of the preceding claims, wherein each of the fins (24) has a height measured along said elongation axis (A) which represents between 1 and 10%, and preferably between 1 and 5%, of the height of the blade (12) measured along the same axis (A).

7. The vane (10) according to one of the preceding claims, wherein each of the fins (24) is inclined with respect to said elongation axis (A).

8. A turbomachine module (20), in particular for an aircraft, this module (20) comprising a rotor comprising an annular row of vanes (10) according to one of the preceding claims.

9. The module (20) according to the preceding claim, wherein it is selected from a turbine or compressor wheel, a mixed or centrifugal spinning wheel, and a fan.

10. The module (20) according to claim 8 or 9, wherein the rotor is surrounded by an annular coating (22) of abradable material which is housed in a groove (23) of an annular casing (25), this groove (23) being formed in an inner cylindrical surface of the casing (25) which has an inner diameter (D1) substantially equal to the diameter of the rotor passing through the free ends of the blades (12) of its vanes (10).

11. An aircraft turbomachine, comprising at least one vane (10) according to one of claims 1 to 7.