Reinforcement for the leading edge of a turbine engine blade

Abstract

A turbine engine blade comprising an aerodynamic surface that extends in a first direction between a leading edge and a trailing edge, and in a second direction substantially perpendicular to the first direction between a root and a tip of the blade, and a leading-edge reinforcement comprising a fin partly covering the aerodynamic surface of the blade, characterised in that the fin has a radially outer edge arranged in the vicinity of the tip of the blade and extending between the leading edge and the trailing edge, this radially outer edge comprising an upstream point fitting flush with the tip of the blade at the leading edge and a so-called downstream point separated from the tip of the blade.

Claims

1. A turbine engine blade extending along a longitudinal axis, comprising an aerodynamic surface that extends in a first direction between a leading edge and a trailing edge, and in a second direction substantially perpendicular to the first direction between a root and a tip of the blade, and a leading-edge reinforcement comprising a fin partly covering the aerodynamic surface of the blade, wherein the fin has a radially outer edge arranged in the vicinity of the tip of the blade and extending between the leading edge and the trailing edge, the radially outer edge comprising an upstream point fitting flush with the tip of the blade at the leading edge and a downstream point spaced radially from the tip of the blade.

2. The turbine engine blade of claim 1, wherein the aerodynamic surface is a suction-face surface, and the fin is a suction-face fin.

3. The turbine engine blade of claim 1, wherein the radially outer edge of the fin comprises an intermediate point situated between the upstream point and the downstream point and defining with the upstream point a first portion of the radially outer edge, fitting flush with the tip of the blade and, with the downstream point, a second portion of the radially outer edge separating gradually from the tip of the blade in the direction of the downstream point.

4. The turbine engine blade of claim 3, in which the intermediate point is arranged longitudinally at equal distances from the upstream point and from the downstream point.

5. The turbine engine blade of claim 3, wherein the second portion of the radially outer edge of the fin is curved and convex.

6. The turbine engine blade of claim 3, wherein the intermediate point and the downstream point are separated from each other by a distance, measured along a median longitudinal axis of the fin, between 0 and sin L4, where: L is a length of the fin before optimization, between the upstream point and a fictive extreme point corresponding to a symmetry of the upstream point with respect to the median longitudinal axis substantially perpendicular to the longitudinal axis of the turbine engine, and passing at least through the center of the tip of the fin, and is the angle measured between a line passing through the upstream point and the intermediate point of the radially outer edge and a tangent to the radially outer edge, parallel to the longitudinal axis and passing through the intermediate point.

7. The turbine engine blade of claim 1, wherein the leading-edge reinforcement comprises a pressure-face fin partly covering an aerodynamic pressure-face surface of the blade.

8. The turbine engine blade of claim 1, wherein the leading-edge reinforcement is produced from a metallic material.

9. An assembly comprising a central disc on which a plurality of turbine engine blades according to claim 1 are mounted, said blades being evenly distributed around a periphery of the central disc, and extending substantially radially with respect to the central disc.

10. A turbine engine comprising the assembly of claim 9.

Description

BRIEF DESCRIPTION OF THE DRAWINGS

(1) The invention will be understood better and other details, features and advantages of the invention will emerge from a reading of the following description given by way of non-limitative example with reference to the accompanying drawings, in which:

(2) FIG. 1 is a schematic view of a turbine engine comprising an assembly having a plurality of blades;

(3) FIG. 2 is a perspective view of a blade according to the invention, in particular a fan blade, this blade carrying a leading-edge reinforcement limiting the degradation of the internal abradable layer of the turbine engine;

(4) FIG. 3 is a view in cross section of the blade along the cross-sectional plane III-III in FIG. 2;

(5) FIG. 4 is a detail view of a top portion of a blade in accordance with the inset IV in FIG. 2, and

(6) FIG. 5 is a detail view to an enlarged scale of the detail V in FIG. 4.

DETAILED DESCRIPTION

(7) FIG. 1 shows a turbine engine 2 having an assembly 4 comprising a central disc 6 rotatable about a longitudinal axis A of the turbine engine 2, and on which a plurality blades 8 are mounted. The blades 8 are evenly distributed around the periphery 6a of the central disc 6, and extending substantially radially to the central disc 6. In the present case, the assembly 4 is the fan of the turbine engine 2, and the blades 8 are the fan blades.

(8) Conventionally, the turbine engine 2 also comprises, from upstream to downstream, and downstream of the fan, a low-pressure compressor 10, a high pressure compressor 12, a combustion chamber 14, a high-pressure turbine 16, a low-pressure turbine 18 and an exhaust casing 20. Furthermore, for attachment thereof to the aeroplane, the turbine engine 2 comprises attachment means 22, in this case two, each carried by an intermediate fan casing 24 carrying an internal abradable layer 24a (visible in FIG. 4), and a turbine casing 26.

(9) In the remainder of this description, the term radial means any direction substantially perpendicular to the axis A of the turbine engine 2, the term upstream the side by means of which the air reaches a part of the turbine engine 2, and the term downstream the side through which the air moves away from said part of the turbine engine 2. The airflow direction is depicted in FIG. 2 by the arrow F.

(10) Blade 8 means here both the moving blades (for example the rotor blades) and the fixed blades (for example the stator blades) of the turbine engines 2.

(11) The blade 8, illustrated in perspective in FIG. 2 and in cross section in FIG. 3, comprises an aerodynamic suction-face surface 28 and an aerodynamic pressure-face surface 30 that extend in a first direction between a leading edge 8a and a trailing edge 8b of the blade 8. The blade 8 of a fan being twisted, the first direction changes in a plane XY along the cross section taken in a radial direction along the axis Z, which forms with axes X and Y an orthonormal reference frame in FIG. 2. In a second direction substantially perpendicular to the first direction, the aerodynamic suction-face surface 28 and the aerodynamic pressure-face surface 30 extend between a root 8c and a tip 8d of the blade 8.

(12) The blade 8 also comprises a leading-edge reinforcement 32 comprising a suction-face fin 32a partly covering the aerodynamic suction-face surface 28 of the substantially radial blade 8, and a pressure-face fin 32b partly covering the aerodynamic pressure-face surface 30 of the blade 8. These two fins 32a, 32b have, as can be seen in FIG. 3, a cross section that becomes thinner from upstream to downstream.

(13) The two fins 32a, 32b are joined by a leading edge 32c that covers the leading edge 8a of the blade 8 and, in cross section, has thickness greater than the maximum thickness of the fins 32a, 32b.

(14) As can be seen in FIG. 2, the reinforcement 32 of the leading edge 8a of the blade 8 extends substantially from the root 8c of the blade 8 as far as its tip 8d.

(15) The leading-edge reinforcement 32 is preferably produced from a high-strength metallic material, such as for example a titanium alloy.

(16) The detail view in FIG. 4 shows a particularity of the suction-face fin 32a of the leading-edge reinforcement 32. Indeed, the suction-face fin 32a has a radially outer edge 34 (also referred to as the top edge) arranged in the vicinity of the tip 8d of the blade and which extends from the leading edge 8a to the trailing edge 8b (FIG. 2). This radially outer edge 34 comprises an upstream point 34a that fits flush with the tip 8d of the blade 8 at the leading edge 8a and a downstream point 34b that is spaced from the tip 8d of the blade 8. The term upper extends according to, the orientation in FIG. 4. In other words the radially outer edge 34 is disposed radially externally with respect to the axis A of the turbine engine 2.

(17) It should be understood that the upstream point 34a is arranged on the same side as the leading edge 8a of the blade 8 and the downstream point 34b is arranged on the same side as the trailing edge 8b of the blade 8 in the direction F of airflow (FIG. 2) on the blade 8 from the leading edge 8a to the trailing edge 8b.

(18) Furthermore, the upper radially outer edge 34 of the suction-face fin 32a comprises an intermediate point 34c situated between the upstream point 34a and the downstream point 34b and defining with the upstream point 34a a first portion 36 of the radially outer edge, fitting flush with the tip 8d of the blade 8 and, with the downstream point 34b, a second portion 38 of the upper edge moving away gradually from the tip 8d of the blade 8. The connection of the first portion 36 of the radially outer edge 34 with the second portion 38 of the upper edge is substantially tangential.

(19) According to one aspect, the intermediate point 34c is arranged at equal distances from the upstream point 34a and the downstream point 34b, in an axial direction parallel to the longitudinal axis A. However, the intermediate point 34c could be closer to the upstream point 34a or to the downstream point 34b.

(20) FIG. 5 shows a fictive extreme point 34e corresponding to the symmetry of the upstream point 34a with respect to a median axis M substantially perpendicular to the axis A of the turbine engine 2, and passing at least through the centre of the tip of the suction-face fin 32a. This fictive extreme point 34e corresponds to an extreme point of the suction-face fin 32a before optimisation thereof.

(21) Advantageously, this extreme point 34e makes it possible to define the gradual separation of the downstream point 34b with respect to the tip 8d of the blade 8.

(22) The spacing of second portion 38 of the radially outer edge 34 of the suction-face fin 32a is preferably curved and convex. In other words, the second portion 38 has substantially a curved shape that spaces continuously from the tip 8d of the blade 8 in the direction of the root 8c (FIG. 2) thereof, and this from upstream to downstream.

(23) However, according to variant embodiments not shown in the figures, the second portion 38 of the radially outer edge 34 of the suction-face fin 32a could be rectilinear or on the other hand comprise an alternation of protrusions and hollows.

(24) According to a preferred embodiment shown in FIG. 5, the intermediate point 34c and the downstream point 34b are separated from each other by a distance H1 measured along the longitudinal median axis M, that is to say in the radial direction Z, H1 being between 0 and sin L4

(25) where: L is the length of the fin 32a before optimisation, that is to say between the upstream point 34a and the fictive point 34e, and is the angle measured between a line passing through the upstream point 34a and the intermediate point 34c on the radially outer edge 34 and a tangent T to said radially outer edge 34, parallel to the longitudinal axis A of the turbine engine 2 and passing through the intermediate point 34c.

(26) The distance L, the tangent T and the angle are illustrated in FIG. 5.

(27) Thus, in the event of impact of an FOD on the leading-edge reinforcement 32, if the suction-face fin 32a detaches, it will not come into contact with the internal abradable layer 24a carried by the intermediate fan casing 24. Consequently it will be necessary only to repair the blade 8 that has been impacted (or the impacted blades 8), which is simpler, quicker and less expensive that complete immobilisation of the turbine engine 2 for replacing the impacted blade 8 (or impacted blades 8) of the intermediate fan casing 24 and its internal abradable layer 24a.

(28) For reasons of simplicity in manufacture of the reinforcement 32 of the leading edge, the pressure-face fin 32b also comprises a top edge having an upstream point fitting flush with the tip 8d of the blade 8 and a downstream point distant from the upstream point and spaced from the tip 8d of the blade 8, that is to say radially distant internally.

(29) The top edge of the pressure-face fin 32b may also comprise an intermediate point situated between the leading point and the trailing point and defining with the leading point a first portion of the top edge, fitting flush with the tip 8d of the blade 8 and, with the trailing point, a second portion of the top edge spacing gradually from the tip 8d of the blade 8 in the direction of the root 8c.

(30) However, the forms and dimensions of the portions of the pressure-face 32b are smaller compared with the forms and dimensions of the portions 36, 38 of the top edge 34 of the suction-face fin 32a.

(31) Thus an asymmetric reinforcement 32 will be obtained.