EP3315721A1 - Leading-edge reinforcement of a turbine engine blade - Google Patents

Abstract

The invention relates to a turbomachine blade, comprising an aerodynamic surface (28) extending in a first direction between a leading edge (8a) and a trailing edge, and in a second direction substantially perpendicular to the first direction between a root and a top (8d) of the blade, and a leading edge reinforcement (32) comprising a fin (32a) partially covering the aerodynamic surface (28) of the blade, characterized in that, the fin (32a) has a radially outer edge (34) arranged in the vicinity of the apex (8d) of the blade and extending between the leading edge (8a) and the trailing edge, this edge (34) radially outer comprising an upstream point (34a) flush with the apex (8d) of the blade at the edge (8a) of attack and a point (34b) said downstream spaced from the vertex (8d) of the blade.

Description

The present invention relates to a turbomachine blade, and more particularly to a leading edge reinforcement of this blade.

By dawn means here both the blades and the blades of the turbomachines.

In order to increase the resistance of the blades to the Foreign Object Damages (FOD) in the airflow, that is to say to foreign bodies such as birds or hailstones, these include a Leading edge reinforcement whose role is to protect the leading edge from deterioration during an impact with a FOD and to distribute the impact force a large area of the dawn.

A reinforcement of the blade leading edge conventionally comprises an extrados vane at least partially covering the aerodynamic surface of the upper surface of the vane and a vane fin at least partially covering the aerodynamic surface of the lower surface of the vane. dawn, these two fins being joined by a leading edge of the reinforcement.

When the blade is movable relative to the axis of the turbomachine, it turns its surface intrados forward, that is to say that the air comes into contact with the surface of the lower surface thus creating an overpressure on the underside surface and a depression on its extrados surface.

The impact of a FOD on the leading edge reinforcement tends to cause a detachment of the upper portion of the intrados vane. Beyond a certain mass of the FOD, the force of the impacts is greater on the reinforcement, which also causes a detachment of the upper portion of the extrados fin. The overpressure generated on the intrados surface tends to limit the detachment of the intrados vane to the intrados surface. On the other hand, the combination of the centrifugal force, more important at the top of the blade than at the bottom, with the depression generated on the extrados surface, tends to favor the separation of the extrados vane.

When the blade is a blade of a fan mounted in an external fairing bearing an internal layer of abradable facing the blades, the detachment of the extrados blade causes damage to the inner layer of abradable. Indeed, the extrados vane protrudes from the upper surface of the blade and penetrates into the abradable inner layer which creates a groove in the abradable inner layer. It is then necessary to immobilize the turbomachine to replace both the blade whose leading edge reinforcement has taken off and the inner layer of abradable. Such immobilization generates a significant cost resulting from the lack of exploitation of the turbomachine that it is important to reduce or even eliminate.

The invention aims in particular to provide a simple, effective and economical solution to this problem.

For this purpose, the invention proposes, firstly, a turbomachine blade extending along a longitudinal axis, comprising an aerodynamic surface which 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 foot and a top of the blade, and a leading edge reinforcement comprising a fin partly covering the aerodynamic surface of the blade, characterized in that, the fin has a radially outer edge arranged in the vicinity of the apex of the blade and extending between the leading edge and the trailing edge, this radially outer edge comprising an upstream point flush with the top of the blade at the edge attack and a downstream point separated from the top of the dawn.

The spacing of the downstream point from the upper edge of the extrados vane makes it possible to limit the penetration of the vane into the abradable inner layer of the turbomachine, in case of detachment from the downstream point of the vane since this It is then removed from the abradable because of its remoteness at the mounting of the dawn apex.

In a particular embodiment of the invention, the upstream point is located at the upstream end of the upper edge, that is to say at the leading edge of the blade and the downstream point is located at the downstream end of the radially outer edge of the fin.

In the reference of the turbomachine, it can thus be considered that the downstream point is spaced radially inwardly of the blade tip.

Advantageously, the aerodynamic surface is an extrados surface, and the fin is an extrados fin, the extrados portion of the reinforcement being more particularly subject to detachment, increased separation in particular by the centrifugal force for a movable blade.

Advantageously, the radially outer edge of the fin comprises an intermediate point located between the upstream point and the downstream point and defining with the upstream point a first portion of the radially outer edge, flush with the top of the blade and, with the downstream point a second portion of the radially outer edge progressively diverging from the apex of the blade towards the vanishing point.

The separation into two portions offers a good compromise between limiting penetration of the fin in the abradable inner layer in case of separation of the fin, and good distribution of forces in case of impact of a FOD on the reinforcement leading edge.

The intermediate point may be arranged longitudinally at an equal distance from the upstream point and the downstream point.

This protects the dawn all the way up since the first portion is flush with the top of the dawn.

Preferably, the second portion of the radially outer edge of the extrados vane is curved convex. This particular shape makes it easier to manufacture the reinforcement and also to limit the creation of disturbances in the flow of the air flow.

• L est la longueur de l'ailette avant optimisation, c'est-à-dire entre le point amont et un point extrême fictif correspondant à la symétrie du point amont par rapport à l'axe médian sensiblement perpendiculaire à l'axe longitudinal de la turbomachine, et passant au moins par le centre du sommet de l'ailette, et
• α est l'angle mesuré entre une ligne passant par le point amont et le point intermédiaire du bord radialement extérieur et une tangente au bord radialement extérieur, parallèle à l'axe longitudinal et passant par le point intermédiaire.

Advantageously, the intermediate point and the vanishing point are separated from one another by a distance, measured along a median longitudinal axis of the fin, between 0 and sinα × L ÷ 4 or :

• L is the length of the fin before optimization, that is to say between the upstream point and a fictive extreme point corresponding to the symmetry of the upstream point relative to the median axis substantially perpendicular to the longitudinal axis of the turbomachine, and passing at least through the center of the top of the fin, and
• α 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.

This distance also offers a good compromise between limiting penetration of the fin in the abradable inner layer in case of separation of the fin, and good distribution of forces in case of impact of an FOD on the edge reinforcement. attack.

Preferably, the reinforcement comprises a vane fin partly covering an aerodynamic surface of intrados of the blade.

This underside fin also helps protect the aerodynamic surface of intrados of the blade against the FOD.

To ensure good protection of the blade, the leading edge reinforcement is made of a metallic material.

The invention proposes, secondly, an assembly comprising a central disc on which are mounted a plurality of blades as previously described, said vanes being evenly distributed around the periphery of the central disc, and extending substantially radially to the disc central.

The invention proposes, thirdly, a turbomachine comprising an assembly as previously described.

• la figure 1 est une vue schématique d'une turbomachine comprenant un ensemble ayant une pluralité d'aubes;
• la figure 2 est une vue en perspective d'une aube selon l'invention, en particulier une aube de soufflante, cette aube portant un renfort de bord d'attaque limitant la dégradation de la couche interne d'abradable de la turbomachine ;
• la figure 3 est une vue en section de l'aube selon le plan de section III - III de la figure 2 ;
• la figure 4 est une vue de détail d'une portion supérieure de l'aube selon l'encart IV de la figure 2, et
• la figure 5 est une vue de détail à échelle agrandie du détail V de la figure 4.

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

• the figure 1 is a schematic view of a turbomachine comprising an assembly having a plurality of blades;
• the figure 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 inner abradable layer of the turbomachine;
• the figure 3 is a sectional view of the dawn according to the section III - III plan of the figure 2 ;
• the figure 4 is a detail view of an upper portion of dawn according to inset IV of the figure 2 , and
• the figure 5 is an enlarged detail view of detail V of the figure 4 .

On the figure 1 , a turbomachine 2 having an assembly 4 comprising a central disc 6 rotating about a longitudinal axis A of the turbomachine 2, and on which is mounted a plurality of blades 8. The blades 8 are regularly distributed around the periphery 6a of the disc 6 central, and extending substantially radially to the disc 6 central. In this case, the assembly 4 is the fan of the turbomachine 2, and the blades 8 are the fan blades.

Conventionally, the turbomachine 2 also comprises, upstream to downstream, and downstream of the blower, 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. In addition, for its attachment to the aircraft, the turbomachine 2 comprises hooking means 22, in this case two, each carried by an intermediate blower housing 24 carrying an internal layer 24a abradable (visible on the figure 4 ), and a turbine casing 26.

In the remainder of this description, the term radial (e) is understood to mean any direction substantially perpendicular to the axis A of the turbomachine 2, upstream the side by which the air reaches a part of the turbomachine 2, and downstream the side by which the air moves away from said part of the turbomachine 2. The air flow direction is represented in FIG. figure 2 by the arrow F.

By blade 8, is meant here both the blades 8 movable (for example the rotor blades) and the blades (for example the stator vanes) of the turbomachines 2.

Dawn 8, illustrated in perspective on the figure 2 and in section on the figure 3 comprises an aerodynamic surface 28 of an upper surface and an aerofoil surface 30 which extends 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 moves in an XY plane following the section taken in a radial direction along the Z axis which forms with the X and Y axes an orthonormal reference on the figure 2 . In a second direction substantially perpendicular to the first direction, the aerodynamic surface 28 of the upper surface and the aerodynamic surface 30 of the undersides extend between a foot 8c and an apex 8d of the blade 8.

The blade 8 also comprises a leading edge reinforcement 32 comprising an extrados fin 32a partly covering the aerodynamic surface 28 of the extrados of the substantially radial blade 8, and a fin 29b of a lower surface covering part of the airfoil. aerodynamic surface of the underside of the blade 8. These two fins 32a, 32b have, as visible in FIG. figure 3 , a section that is refined from upstream to downstream.

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

As we see on the figure 2 , the reinforcement 32 of edge 8a of attack of the blade 8 extends substantially from the foot 8c of the blade 8, to its top 8d.

The leading edge reinforcement 32 is preferably made of a highly resistant metallic material, such as, for example, a titanium alloy.

The detail view of the figure 4 highlights a particularity of the fin 32a of extrados of the leading edge reinforcement 32. Indeed, the extrados vane 32a has a radially outer (also upper) edge 34 arranged in the vicinity of the blade crown 8d and which extends from the leading edge 8a to the edge 8b ( figure 2 ) leak. This radially outer edge 34 comprises an upstream point 34a which is flush with the apex 8d of the blade 8 at the leading edge 8a and a downstream point 34b which is spaced from the apex 8d of the blade 8. The term "superior" according to the orientation of the figure 4 . In other words, the radially outer edge 34 is disposed radially outwardly with respect to the axis A of the turbomachine 2.

It will be understood that the upstream point 34a is arranged on the side of the leading edge 8a of the blade 8 and the downstream point 34b is arranged on the side of the trailing edge 8b of the blade 8 in the direction of air flow. F ( figure 2 ) on the blade 8 from the leading edge 8a to the trailing edge 8b.

In addition, the upper radially outer edge 34 of the extrados 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 flush with the top 8d of the blade 8 and, with the point 34b downstream, a second portion 38 of the upper edge progressively diverging from the top 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.

In one aspect, the intermediate point 34c is arranged equidistant from the upstream point 34a and the downstream point 34b, in an axial direction parallel to the longitudinal axis A. However, intermediate point 34c could be closer to point 34a upstream or point 34b downstream.

We have shown on the figure 5 , a fictitious extreme point 34e corresponding to the symmetry of the point 34a upstream with respect to a median axis M substantially perpendicular to the axis A of the turbomachine 2, and passing at least through the center of the top of the fin 32a of extrados . This 34th point fictitious extreme corresponds to an extreme point of the fin 32a extrados before optimization of the latter.

Advantageously, this point 34th extreme makes it possible to define the progressive spacing of the point 34b downstream from the vertex 8d of the blade 8.

The spacing of the second portion 38 of the radially outer edge 34 of the extrados fin 32a is preferably convexly curved. In other words, the second portion 38 has a substantially curved shape that moves away from the top 8d of the blade 8 in the direction of the foot 8c ( figure 2 ) of the latter, and this upstream downstream.

However, according to alternative embodiments not shown in the figures, the second portion 38 of the radially outer edge 34 of the extrados vane 32a could be rectilinear or, conversely, comprise an alternation of bumps and depressions.

• L est la longueur de l'ailette 32a avant optimisation, c'est-à-dire entre le point 34a amont et le point 34e fictif, et
• α est l'angle mesuré entre une ligne passant par le point 34a amont et le point 34c intermédiaire du bord 34 radialement extérieur et une tangente T audit bord 34 radialement extérieur, parallèle à l'axe longitudinal A de la turbomachine 2 et passant par le point 34 c intermédiaire.

According to a preferred embodiment represented at figure 5 the intermediate point 34c and the downstream point 34b are separated from each other by a distance H1 measured along the median longitudinal axis M, that is to say in the radial direction Z, H1 being between 0 and sinα × L ÷ 4 where:

• L is the length of the fin 32a before optimization, that is to say between the point 34a upstream and the point 34e fictitious, and
• α is the angle measured between a line passing through the point 34a upstream and the intermediate point 34c of the radially outer edge 34 and a tangent T to said radially outer edge 34, parallel to the longitudinal axis A of the turbomachine 2 and passing through the point 34c intermediate.

The distance L, the tangent T and the angle α are illustrated on the figure 5 .

Thus, in the event of an impact of an FOD on the leading edge reinforcement 32, if the upper winglet 32a comes off, it will not come into contact with the abradable inner layer 24a carried by the housing 24 intermediate blower. Therefore, it will only be necessary to repair the blade 8 which has been impacted (or the blades 8 impacted), which is simpler, faster and less expensive than the complete immobilization of the turbomachine 2 for the replacement of the impacted blade 8 (or impacted blades 8) and the intermediate fan casing 24 and its abradable internal layer 24a.

For the sake of simplicity of manufacture of the leading edge reinforcement 32, the lower surface fin 32b also comprises an upper edge having an upstream point flush with the apex 8d of the blade 8 and a downstream point remote from the upstream point and spaced from the vertex 8d of the blade 8, that is to say radially distant internally.

The upper edge of the lower surface fin 32b may also comprise an intermediate point situated between the point of attack and the vanishing point and defining with the point of attack a first portion of the upper edge, flush with the top 8d of the dawn 8 and, with the vanishing point, a second portion of the upper edge progressively diverging from the apex 8d of the blade 8 towards the foot 8c.

However, the shapes and dimensions of the portions of the vane fin 32b are reduced with respect to the shapes and dimensions of the portions 36, 38 of the upper edge 34 of the extrados vane 32a.

Thus, an asymmetrical reinforcement 32 will be obtained.

Claims (10)

A turbomachine blade (8) extending along a longitudinal axis (A), comprising an aerodynamic surface (28, 30) extending in a first direction between a leading edge (8a) and an edge (8b) of leak, and in a second direction substantially perpendicular to the first direction between a foot (8c) and a vertex (8d) of the blade (8), and a leading edge reinforcement (32) comprising a fin (32a, 32b) partially covering the aerodynamic surface (28, 30) of the blade (8), characterized in that the fin (32a, 32b) has a radially outer edge (34) arranged in the vicinity of the apex (8d) blade (8) and extending between the leading edge (8a) and the trailing edge (8b), this radially outer edge (34) comprising an upstream point (34a) flush with the apex (8d) of the blade (8) at the leading edge and a point (34b) downstream radially distant from the apex (8d) of the blade (8). A blade (8) according to claim 1, wherein the aerodynamic surface is an extrados surface (28), and the fin is an extrados fin (32a). A blade (8) according to claim 1 or 2, wherein the radially outer edge (34) of the fin (32a, 32b) comprises an intermediate point (34c) between the upstream point (34a) and the point (34b). downstream and defining with the point (34a) upstream a first portion (36) of the edge (34) radially outer, flush with the apex (8d) of the blade (8) and, with the point (34b) downstream, a second portion (38) radially outer edge (34) progressively diverging from the apex (8d) of the blade (8) towards the point (34b) downstream. A blade (8) according to claim 3, wherein the intermediate point (34c) is arranged longitudinally equidistant from the upstream point (34a) and the downstream point (34b). A blade (8) according to any one of claims 3 to 4, wherein the second (38) portion of the radially outer edge (34) of the fin (32a, 32b) is convexly curved. Blade (8) according to one of claims 3 to 5, wherein the point (34c) intermediate and the point (34b) downstream are separated from each other by a distance (H1), measured along the a median longitudinal axis (M) of the fin, between 0 and sinα × L ÷ 4
or : - L is the length of the fin (32a) before optimization, that is to say between the point (34a) upstream and a point (34e) fictitious extreme corresponding to the symmetry of the point (34a) upstream with respect to the central axis (M) substantially perpendicular to the longitudinal axis (A) of the turbomachine (2), and passing at least through the center of the top of said fin, and - α is the angle measured between a line passing through the point (34a) upstream and the intermediate point (34c) of the radially outer edge (34) and a tangent (T) at the radially outer edge (34), parallel to the longitudinal axis (A) and passing through the point (34c) intermediate. A blade (8) according to any one of the preceding claims, wherein the leading edge reinforcement (32) comprises a lower surface fin (32b) partly covering an aerofoil surface (30) of the underside of the dawn (8). A blade (8) according to any one of the preceding claims, wherein the leading edge reinforcement (32) is made of a metallic material. Assembly (4) comprising a central disc (6) on which are mounted a plurality of vanes (8) according to any one of the preceding claims, said vanes (8) being evenly distributed around the periphery (6a) of the disc ( 6) and extending substantially radially with respect to the central disk (6). Turbomachine (2) comprising an assembly (4) according to claim 9.

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