Blade made of composite material for a turbine engine, aircraft turbine engine and method for manufacturing a blade made of composite material for a turbine engine, in particular of an aircraft

Abstract

The invention relates to a blade (7) made of composite material for a turbine engine (10), in particular of an aircraft, the blade (7) comprising an airfoil (70) formed by a fibrous preform embedded in a resin and extending along an axis (A) of elongation, the airfoil (70) comprising a leading edge (73), a trailing edge (74), a pressure side (71) and a suction side (72) connected to the pressure side (71) by the leading edge (73) and the trailing edge (74), the airfoil (70) further comprising at least one serrated element (9) formed by a plurality of teeth (92) and covering at least part of the leading edge (73) and / or the trailing edge (74), wherein the at least one serrated element (9) is made of a thermosetting or thermoplastic material.

Description

[0001] DESCRIPTION

[0002] TITLE :

[0003] BLADE MADE OF COMPOSITE MATERIAL FOR A TURBOMACHINE, AIRCRAFT TURBOMACHINE AND METHOD FOR MANUFACTURING A BLADE MADE OF COMPOSITE MATERIAL FOR A TURBOMACHINE, IN PARTICULAR AN AIRCRAFT TURBOMACHINE

[0004] Technical field

[0005] The invention relates to the field of turbomachines, in particular aircraft turbomachines, and in particular to the propulsion propellers of these turbomachines which comprise blades. More particularly, the invention relates to blades made of composite material for such turbomachines, and more precisely to improving the noise reduction of these blades in operation, as well as a method for manufacturing such blades.

[0006] Technical background

[0007] The state of the art includes in particular documents FR-A1 -3073018, FRAI -3073019, US-A1 -2017 / 226865 and US-A1 -2013 / 156592.

[0008] A propulsive propeller for a turbomachine, particularly an aircraft one, can be shrouded (figure 1) as is the case for a fan for example, or can be unshrouded (figure 2) as is the case for an “open-rotor” type architecture for example.

[0009] A propeller comprises a plurality of blades rotating about a longitudinal axis of the turbomachine. Each blade comprises a blade having a leading edge, a trailing edge, a lower surface and an upper surface connected to the lower surface by the leading edge and the trailing edge.

[0010] Currently, more and more blades are made of composite material to reduce mass while having good mechanical properties. For example, composite blades are made by injecting a resin, for example by liquid resin injection molding (RTM, acronym for the English expression "Resin Transfer Molding") into a fiber preform, formed for example from fibers woven in three dimensions (3D).

[0011] Regardless of the composite or metallic material of the blades, it is known to modify the geometry of these blades to control and / or reduce aerodynamic noise. To do this, the geometry of the leading edge and / or the trailing edge of the blade is modified by adding a serration element (known as "wavy leading / trailing edge", "serrated leading / trailing edge" or "leading / trailing edge serration").

[0012] The principle of reducing noise emissions generated by blades is based on spatially shifting the noise sources distributed along the leading edge or the trailing edge by means of serrations (or otherwise known as undulations or teeth), which may or may not be identical. For this principle to apply, the dimensions (height, length, thickness, etc.) of the serrations must be adapted to the incident aerodynamic field (such as the airflow passing through the blades with more or less turbulence) which varies according to the engine speed of the turbomachine.

[0013] Furthermore, the clamping element must be both easily repairable (since it is visible to passengers) and resistant to HCF (High-Cycle Fatigue) type damage which can generate cracks in the clamping element (and consequently in the blade) during operation. In addition, several other constraints must be taken into account for the production of composite material blades comprising such a clamping element, such as:

[0014] - the clamping element has a complex geometry, in particular with teeth having variations in dimensions (such as its thickness, length, shape, etc.) which can alter and form a misalignment with the composite material (in particular the woven fibers) during shaping in an RTM injection mold,

[0015] - the placement of the clamping element on the leading edge or the trailing edge can be complex, particularly for complex geometries (dimensions and shape) of the aerodynamic profile blades,

[0016] - the teeth of the clamping element have low thicknesses (around 2.5 mm minimum) which can be damaged during handling,

[0017] - the clamping element, typically made of metal (such as titanium), must be compatible with the composite material of the blade (particularly of the three-dimensional woven type), and

[0018] - the clamping element, particularly made of metal, can make the blade heavier and more cumbersome. In this context, it is interesting to propose a solution that can overcome at least one of the aforementioned drawbacks, in particular by optimizing and simplifying the integration of a clamping element in a blade made of composite material, while facilitating the maintenance of this clamping element.

[0019] Summary of the invention

[0020] The present invention provides a simple, effective and economical solution to the aforementioned drawbacks of the prior art.

[0021] To this end, the invention relates to a blade made of composite material for a turbomachine, in particular for an aircraft, said blade comprising a blade formed by a fiber preform embedded in a resin and extending along an elongation axis A, said blade comprising a leading edge, a trailing edge, a lower surface and an upper surface connected to the lower surface by the leading edge and the trailing edge, said blade further comprising at least one serration element formed of several teeth and covering at least part of the leading edge and / or the trailing edge.

[0022] According to the invention, said at least one clamping element is made of a thermosetting or thermoplastic material.

[0023] Thus, this solution allows the aforementioned objective to be achieved. The thermosetting or thermoplastic clamping element allows its integration (for example by gluing) to be reinforced and facilitated at the leading edge and / or trailing edge of the blade.

[0024] Indeed, the thermoset or thermoplastic has mechanical properties compatible with the composite material of the blade. For example, the thermoset and the thermoplastic are more flexible and deformable materials than metal, on the one hand, to simplify the production of the complex geometry of the clamping element, and on the other hand, to help and facilitate its positioning in the composite material blade and its handling in general. Furthermore, the thermoset or the thermoplastic also has good resistance to shocks (especially from external bodies such as birds, stones, etc.) and a low mass compared to metal.

[0025] Finally, the thermoset or thermoplastic, especially with pre-impregnated fibers, facilitates the removal and repair of the clamping element. In this way, the blade according to the invention helps to limit environmental impacts, especially for repair applications.

[0026] The invention therefore has the advantage of being based on a design that is simple to produce and repair, offering very high reliability, and is not very penalizing in terms of cost and size.

[0027] The term "serration element" means a part comprising teeth projecting from at least part of this part. These teeth may have a U and / or V shape in axial section. These teeth may be more or less numerous, more or less large and move along this serration element. The serration element may also be referred to as a tooth or undulation element.

[0028] 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:

[0029] - said at least one clamping element is fixed along the leading edge and / or the trailing edge by a layer of glue;

[0030] - said at least one serration element extends from the leading edge and / or the trailing edge to a part of the intrados and / or the extrados;

[0031] - the blade comprises at least one sacrificial layer comprising, for example, glass fibers, and interposed between the layer of glue and the leading edge and / or the trailing edge;

[0032] - the blade comprises a protective sheath covering at least in part the leading edge, said at least one clamping element at least in part covering this protective sheath; - said at least one clamping element has a length measured along a longitudinal axis B perpendicular to the elongation axis A, which represents between 20 and 100% of a chord length of the blade measured in the same plane;

[0033] - said at least one clamping element has a height measured along the elongation axis A, which represents between 30 and 100% of a total height of the blade measured in the same plane;

[0034] - the fiber preform is produced by weaving fibers in three dimensions or by superimposing several layers of fibers;

[0035] - the thermosetting material comprises an epoxy resin pre-impregnated with carbon fibers or glass fibers, or the thermoplastic material comprises a polyether-ether-ketone, poly-aryl-ether-ketone, poly-ether-imide or epoxy resin;

[0036] - the adhesive layer has a thickness of between 50 and 400 pm, preferably between 100 and 350 pm, the thickness being measured along a cross-section at the elongation axis A;

[0037] -- the glue layer is made of polyester, nylon or epoxy;

[0038] -- the sacrificial layer S comprises glass fibers.

[0039] By the term "sacrificial" is meant a portion (or an area) of this sacrificial layer which is non-functional and therefore configured to be degraded if necessary, in particular during the removal and / or repair of the clamping element, without damaging the composite material of the blade.

[0040] The present invention also relates to an aircraft turbomachine, comprising at least one blade made of composite material according to the invention.

[0041] The turbomachine can be a turbojet, turboprop or aircraft turboshaft engine.

[0042] The present invention further relates to a method for manufacturing a blade made of composite material according to one of the features of the invention, for a turbomachine, in particular an aircraft. The method comprises the following steps:

[0043] (a) providing the blade comprising the blade formed from the fiber preform embedded in resin, the blade comprising the leading edge, the trailing edge, the intrados and the extrados connected to the intrados by the leading edge and the trailing edge,

[0044] (b) producing said at least one clamping element from thermosetting or thermoplastic material,

[0045] (c) mounting and securing said at least one clamping element so as to cover at least a portion of the leading edge and / or the trailing edge.

[0046] The invention may comprise one or more of the following features, taken in isolation from each other or in combination with each other:

[0047] - step (c) comprises gluing (C2) said at least one clamping element onto at least one part of the leading edge and / or the trailing edge;

[0048] - the bonding (C2) is carried out in an autoclave under a vacuum cover, for example at a pressure between 2 and 10 bars and at a temperature between 100 and 200°C;

[0049] - before bonding (C2), step (c) comprises a deposition (ci) of a sacrificial layer on at least one part of the leading edge and / or the trailing edge, so as to place this sacrificial layer between the layer of glue and the leading edge and / or the trailing edge.

[0050] Brief description of the figures

[0051] The present invention will be better understood and other details, characteristics and advantages of the present invention will appear more clearly on reading the description of a non-limiting example which follows, with reference to the appended drawings in which: Figure 1 is a half axial sectional view schematically representing an aircraft turbomachine; Figure 2 is a perspective view schematically representing a first example of a moving blade of the turbomachine of Figure 1 comprising a serration element; Figure 3 is a schematic axial sectional view of the blade of Figure 2; Figure 4 is an enlarged and sectional view along a plane PH schematically representing a part of the blade of Figure 2; Figure 5 is an enlarged and sectional view along the plane PH schematically representing a part of the blade of Figure 4 with a sacrificial layer;Figure 6 is a sectional view along the plane PH schematically representing a second example of a fixed blade of the turbomachine of Figure 1 comprising a serration element; Figure 7 is an enlarged and sectional view along a plane PH schematically representing a part of the blade of Figure 6; Figure 8 is an enlarged and sectional view along the plane PH schematically representing a part of the blade of Figure 7 with a sacrificial layer; Figure 9 is a sectional view along the plane PH schematically representing a part of the blade of Figure 8 with a protective sheath at the leading edge; Figure 10 is a sectional view along a plane PL schematically representing a third example of a blade of the turbomachine of Figure 1; Figure 11 is a block diagram of a method of manufacturing the blade of the invention.;

[0052] Elements having the same functions in different implementations have the same references in the figures.

[0053] Detailed description of the invention

[0054] By convention, in the description below, the terms "longitudinal" and "axial" describe the orientation of structural elements extending in the direction of a longitudinal axis (such as a turbomachine). The terms "radial" or "vertical" describe an orientation of structural elements extending in a direction perpendicular to the longitudinal axis. The terms "inner" and "outer", and "internal" and "external" are used in reference to a positioning relative to the longitudinal axis. Thus, a structural element extending along the longitudinal axis has an inner face facing the longitudinal axis and an outer surface, opposite its inner surface. Similarly, the terms "upstream" and "downstream" are defined in relation to the direction of circulation of the gases in the aircraft propulsion system.

[0055] The invention can be applied in a non-limiting manner to a turbomachine 10, in particular an aircraft turbomachine.

[0056] The turbomachine 10 may be shrouded, which is for example shown in FIG. 1, or may be unshrouded, as is the case with the open-rotor, for example (not illustrated in the figures). The turbomachine 10 may be a turbojet, turboshaft or turboprop.

[0057] The turbomachine 10 extends around a longitudinal axis X. It comprises from upstream to downstream in the direction of flow of the gases F along the longitudinal axis X, a fan 1, at least one compressor (such as a low pressure compressor 2 and / or a high pressure compressor 4), a combustion chamber 5, at least one turbine 6 (such as a high pressure turbine and / or a low pressure turbine) and a nozzle (not shown).

[0058] A rotor of the low-pressure turbine is connected to the fan 1 and to the rotor of the low-pressure compressor 2 by a low-pressure shaft (not shown). A rotor of the high-pressure turbine is connected to the rotor of the high-pressure compressor 4 by a high-pressure shaft (not shown).

[0059] The turbomachine 10 further comprises a rectifier 3. The rectifier 3 may comprise at least one annular row of vanes (in particular stator or fixed) called rectifier or outlet guide vanes (designated by the English term "Outlet Guide Vanes" OGV). These OGV vanes make it possible to straighten the flow at the outlet of a rotor located upstream in order to provide maximum thrust at the outlet of the turbomachine 10. In the particular example of FIG. 1, the rectifier 3 is located downstream of the fan 1 and makes it possible to straighten a secondary flow F2.

[0060] The fan 1 comprises an annular row of fan blades (in particular rotor or mobile) extending around the X axis. The fan 1 allows the suction of an air flow dividing into a primary flow F1 and a secondary flow F2. The primary flow F1 passes through a primary vein of the turbomachine 10 while the secondary flow F2 is directed towards a secondary vein surrounding the primary vein.

[0061] The primary flow F1 is compressed within the low-pressure compressor 3 and then the high-pressure compressor 4. The compressed air is then mixed with fuel and burned within the combustion chamber 5. The gases formed by the combustion pass through the high-pressure turbine and the low-pressure turbine. The gases finally escape through the nozzle, the cross-section of which allows the acceleration of these gases to generate propulsion. The secondary flow F2 passes through the rectifier 3, which accelerates the circulation speed of the secondary flow F2 to generate propulsion.

[0062] The fan 1, the low-pressure compressor 2, the high-pressure compressor 4, the high-pressure turbine, the low-pressure turbine, and the rectifier 3 comprise blades 7. The blades 7 may be mobile (for example the fan blade of the turbomachine 10, shrouded or unshrouded, of FIG. 2) in rotation about the longitudinal axis X, or fixed (the OGV blade 7 of the rectifier 3 of FIG. 6) relative to the axis X. The blades 7 extend substantially radially relative to the axis X.

[0063] In the following description, the invention will be described in the context of its application to a blade 7 made of composite material, in particular with reference to figures 2 to 10. This blade 7 can be mobile from the fan 1 (figure 2) or fixed from the rectifier 3 (figure 6).

[0064] The invention is however not limited to a moving fan blade or a rectifier of a shrouded turbomachine, and can be applied generally to other types of blades made of composite material, such as fixed and / or moving blades of an unshrouded turbomachine.

[0065] With reference to Figures 2, 3, 6 and 10, each blade 7 comprises a blade 70. The blade 70 is formed by a fiber preform embedded in a resin. The blade 70 (and in particular the blade 7) extends, on the one hand, along an elongation axis A (substantially vertical in Figures 2 and 6), and on the other hand, along a longitudinal axis B (substantially horizontal in Figures 2 and 6). This axis A is substantially perpendicular to the axis B. The axis A is substantially perpendicular or inclined to the axis X of the turbomachine 10.

[0066] The blade 70 comprises a lower surface face 71 (hereinafter referred to as lower surface 71) and an upper surface face 72 (hereinafter referred to as upper surface 72). The lower surface 71 and the upper surface 72 extend transversely between a leading edge 73 and a trailing edge 74 of the blade 70.

[0067] The blade 70 may have an aerodynamic profile to form the aerodynamic part of the blade 7. For this, the blade 70 may have a curved profile of variable thickness between the leading edge 73 and its trailing edge 74 of the blade 7.

[0068] In the examples of figures 2 and 6, the blade 70 extends along the axis A between a first end (in particular at the level of a tip 75 of the blade) and a second end opposite the first end.

[0069] In the case of the moving blades 7 of figure 2 or of an unducted turbomachine, these blades 7 may each further comprise a root 76. The root 76 is in particular connected to the second end of the blade 70. It is intended to be fixed to a disk (not shown) for example mobile in rotation around the axis X. The second end is free and configured to form the tip 75 (or a head) of the blade 7.

[0070] In the case of the fixed blades 7 of the shrouded turbomachine of FIG. 6, these blades 7 may each further comprise a first platform 77a and a second opposite platform 77b. The first platform 77a is integral with the first end of the blade 70 and the second platform 77b is integral with the second end.

[0071] As a variant (not shown), of the fixed blades of an unducted turbomachine, the second end of these blades 7 comprises the second platform 77b and the first end is free.

[0072] The blade 70 may have a first height H?, called total, measured substantially along the axis A for example in a plane PH. This plane PH is substantially parallel to the axis A. This first height H? may be between 0.25 and 1.50 meters. Preferably, the first height H? is between 0.60 and 1.30 meters.

[0073] The blade 70 may have a first length L?, called the chord length, measured along the axis B, for example in a plane PL. This plane PL is substantially perpendicular to the axis A.

[0074] The blade 70 comprises at least one serration element 9, 9a, 9b covering at least a portion of the leading edge 73 and / or at least a portion of the trailing edge 74. As mentioned below, the serration element makes it possible to reduce the noise of aerodynamic origin generated by the blade in operation.

[0075] The blade 70 may thus comprise a serration element 9 located either at the leading edge 73 (FIG. 6) or at the trailing edge 74 (FIG. 2), or two serration elements 9 (called first 9a and second 9b serration elements) located, respectively, on the leading edge 73 and the trailing edge 74 (FIG. 10). For example, the serration element 9 may be located on the leading edge 73 when the blade 7 corresponds to a propeller of the unducted turbomachine. The serration element 9 may be on the trailing edge 74 when the blade 7 corresponds to the OGV blade of the ducted turbomachine.

[0076] The serration element 9 comprises teeth 92 (or undulations). For example, the serration element 9 may comprise between five and twenty teeth 92. These teeth 92 may be aligned with respect to each other, in particular along the leading edge 73 and / or the trailing edge 74. These teeth 92 may be aligned with each other along the same plane PH. The teeth 92 may extend partially or over the entire leading edge 73 and / or the trailing edge 74.

[0077] The serration element 9 may extend from the leading edge 73 to a portion of the lower surface 71 and / or the upper surface 72, and / or the serration element 9 may extend from the trailing edge 74 to a portion of the lower surface 71 and / or the upper surface 72. The serration element 9 may thus cover either a portion of the lower surface 71, or a portion of the upper surface 72, or possibly a portion of the lower surfaces 71 and upper surfaces 72. Advantageously, the serration element 9 may extend from a portion of the lower surface 71 and / or upper surface 72 to beyond the leading edge 73 and / or the trailing edge 73. The serration element 9 may have a second height Hg measured substantially along the axis For example, in the PH plane. This second height Hg can be less than or equal to the first height H?. In particular, the second height Hg can be between 30 and 100% of the first height H?.This allows noise reduction to be regulated over part or all of the blade height as required. For example, the second height Hg can be between 50 and 120 cm. More specifically, the second height Hg can be between 60 and 100 cm.

[0078] The clamping element 9 may have a second length Lg measured along the axis B. This second length Lg may be less than or equal to the first length L?. In particular, the second length Lg may be between 20 and 100% of the first length L?. This makes it easier to maintain the mounting of the clamping element 9 on the blade 70. For example, the second length Lg may be between 20 and 50 cm. More particularly, the second length Lg may be between 25 and 40 cm.

[0079] The teeth 92 may be identical or non-identical to each other. For example, the teeth 92 may have second lengths L? that are different from each other.

[0080] Figure 2 illustrates in a non-limiting manner approximately seven teeth 92. These teeth 92 are not identical to each other and have variable lengths (relative to the axis B). In particular, the teeth 92 of Figure 2 have decreasing lengths starting from the apex 75 towards the second end of the blade 70. The clamping element 9 of Figure 2 extends from a portion of the intrados face 71 to beyond the trailing edge 74, while covering a portion of this trailing edge 74.

[0081] Figure 6 illustrates in a non-limiting manner approximately fifteen substantially identical teeth 92 of variable lengths. The gripping element 9 of Figure 6 extends from a portion of the intrados face 71 to beyond the leading edge 73, while covering a portion of this leading edge 73.

[0082] The blade 70 may comprise a polymer matrix (or resin) and a fibrous reinforcement embedded and densified in the matrix to form a fibrous preform. In other words, the fibrous preform is embedded in a resin. By “fibrous preform” is meant an intermediate part for producing the final blade 7. The composite material is for example an organic matrix composite (OMC) or ceramic matrix composite (CMC).

[0083] Organic matrix composites (OMCs) and ceramic matrix composites (CMCs) are used as replacements for metallic parts in certain parts of turbomachines (such as blades). Furthermore, their use contributes to optimizing aircraft performance, particularly by improving the efficiency of the turbomachine and reducing the overall mass of the turbomachine, significantly reducing harmful emissions to the environment (CO, CO2, NOx, etc.).

[0084] The fiber preform (in particular the fiber reinforcement) can be obtained by weaving fibers in three dimensions or by superimposing (or draping) several layers of fibers.

[0085] The fibrous preform may comprise fibers which are, for example, carbon fibers, ceramic fibers (such as silicon carbide), glass fibers, aramid fibers, or a mixture of at least two of these fibers.

[0086] The resin or matrix (once the resin is densified) is, for example, a thermoplastic or thermosetting polymer resin. The thermosetting material can be, for example, based on epoxy, polyepoxide, polyimide, polybismaleimide, polyurethane, polyester or vinylester.

[0087] Furthermore, the blade 7 may optionally comprise a protective sheath 8. This protective sheath 8 at least partially covers the leading edge 73 of the blade. Advantageously, the protective sheath 8 may extend over the entire height (relative to the axis A) and over a portion in length (relative to the axis B) of the intrados 71 and the extrados 72 from the leading edge 73.

[0088] In the present application, the term "protective sheath" means a member of variable shape (for example which has a U or V shape in cross section relative to the axis A) which at least partially covers the leading edge 73 of the blade 7. The protective sheath 8 makes it possible to protect the blade (in particular the blade blade) against external impacts (gravel from a takeoff / landing runway, hailstones, birds, etc.) and against erosion of the blade.

[0089] The protective sheath 8 can be fixed to the leading edge 73 by gluing or by projection of a film, for example metallic.

[0090] The protective sheath 8 may be a metal reinforcement, a metal shield, a metal foil or any other form making it possible to cover at least part of the leading edge of the blade.

[0091] The metallic material of the protective sheath 8 is for example titanium or an alloy such as steel (for example stainless steel) or a nickel and cobalt alloy (NiCo).

[0092] One of the particularities of the invention is that the clamping element 9 is made of a thermoplastic or thermosetting material. This makes it possible in particular to simplify and reinforce the integration of the clamping element 9 on the blade 70 of the vane 7 made of composite material, while facilitating the repair and handling of this clamping element 9.

[0093] The thermosetting or thermoplastic material can be pre-impregnated with a resin.

[0094] The thermosetting material may comprise a carbon fiber or glass fiber pre-impregnated epoxy resin. For example, the carbon fiber pre-impregnated epoxy resin may be of the MTM49-L® or HexPIy® M91 type. The glass fiber pre-impregnated epoxy resin may be of the HexPIy® M26T type.

[0095] The thermoplastic material may comprise a polyether-ether-ketone, poly-aryl-ether-ketone or poly-ether-imide resin.

[0096] With reference to figures 2 to 10, the clamping element 9 can be fixed along the leading edge 73 and / or the trailing edge 74 (possibly on a part of the intrados 71 and / or the extrados 42) by a layer of adhesive C. This makes it possible to facilitate and simplify the integration (and / or the maintenance or repair) of the clamping element 9 on the blade 70. Furthermore, the use of the layer of adhesive C makes it possible to control the mechanical properties of the clamping element 9 to be bonded to the blade 70 upstream (i.e. during the manufacture of the blade 7), such as the volume ratio of TVF fibers, sufficient porosity, and / or a reduction in the formation of resin clumps within the clamping element 9.

[0097] The adhesive layer C may extend over at least part of the intrados 71 and / or the extrados 72 up to the leading edge 73 and / or the trailing edge 74.

[0098] In a non-limiting manner, the layer of glue C can be located on the blade 7 according to at least one of the following different possibilities:

[0099] - cover part of the intrados 71 and the trailing edge 74 (figure 6),

[0100] - cover part of the intrados 71 and the leading edge 73,

[0101] - cover both a part of the intrados 71 and the leading edges 73 and trailing edges 74,

[0102] - cover part of the extrados 72 and the leading edge 73 (figure 7),

[0103] - cover part of the extrados 72 and the trailing edge 74,

[0104] - cover both a portion of the extrados 72 and the leading edges 73 and trailing edges 74, and / or

[0105] - cover both, on the one hand, a part of the intrados 71 and the extrados 72, and on the other hand, a part of the leading edges 73 and trailing edges 74.

[0106] Advantageously, the adhesive layer C may have a first free external surface portion C? and a second external surface portion C9 covered with the clamping element 9. The first external surface portion C7 makes it possible to ensure perfect continuity between the surfaces of the blade 7 (in particular the intrados 71 and / or the extrados 72) and the clamping element 9. The first external surface portion C7 may have a third length Lc7 less than a fourth length L of the second external surface portion C9. These third Lc7 and fourth L lengths are measured along the axis B. For example, the third length Lc7 may have between 5 and 25% of the fourth length Leg.

[0107] The adhesive layer C may have a first thickness Ec measured along a cross-section at the axis A. This first thickness Ec may be between 50 and 400 μm. Preferably, the first thickness Ec may be between 100 and 350 μm. This first thickness Ec may vary along the leading edge and / or the trailing edge to adapt to the aerodynamic shape of the blade 70.

[0108] The adhesive layer C may be made of polyester, nylon or epoxy. The adhesive layer C made of polyester or nylon may be reinforced, for example with fibers. Preferably, the adhesive layer C may comprise fibers chosen from nylon, polyester and glass fibers. These fibers allow the adhesive layer to better withstand breaking stresses that may be applied to the clamping element 9 during operation.

[0109] Advantageously, the blade 70 may comprise at least one sacrificial layer S which may be interposed between the adhesive layer C and the leading edge 73 (with or without protective sheath 8) and / or the trailing edge 74. The blade 70 may comprise several sacrificial layers S.

[0110] The sacrificial layer S may extend over at least a portion of the intrados 71 and / or the extrados 72 up to the leading edge 73 and / or the trailing edge 74. In this configuration, the adhesive layer C may cover at least a portion or all of the sacrificial layer S. This sacrificial layer S makes it possible to protect the blade 70 during a maintenance operation on the blade 7 (such as removal of the clamping element 9 for example to repair it and / or stripping of the adhesive layer C for example to deposit a new layer of adhesive C and bonding of a new clamping element 9, etc.). Thus, the sacrificial layer S may be at least partially removed during one of the aforementioned maintenance operation examples without damaging the blade 70. The sacrificial layer S may also form a visual indicator for an operator to carry out one of the maintenance operation examples.

[0111] The sacrificial layer S may comprise glass fibers.

[0112] The sacrificial layer S may have a second thickness Es measured along a cross-section to the axis A. This second thickness Es may be greater than or equal to the first thickness Ec. For example, the second thickness Es may be between 0.1 and 1 mm. Preferably, this second thickness Es may be between 0.15 and 0.5 mm. In this way, the sacrificial layer is not too thin to be able to provide sufficient minimum protection and is not too thick to limit the size of the blade. The present application will now describe the different possibilities of the blade 7 with the clamping element 9 with reference to FIGS. 2 to 10.

[0113] Figures 2 to 5 illustrate a first example of the blade 7 as described above, in which the clamping element 9 (or the first clamping element 9a) can be fixed by the adhesive layer C for example only along the trailing edge 74 of the blade 70 of this blade. With reference to Figure 4, the adhesive layer C covers a portion of the intrados 71 up to the trailing edge 74, and the clamping element 9 covers a portion of the adhesive layer C (in particular that of the second external surface portion Cg) up to and beyond the trailing edge 74.

[0114] Alternatively, the clamping element 9 may also be fixed along the upper surface 72 of the blade 70 of FIG. 4 at the trailing edge 74. According to another variant, another clamping element 9 may cover a portion of the leading edge 73 of the blade 70 of FIG. 4, and possibly the lower surface 71 and / or the upper surface 72.

[0115] As previously described, the glue layer C comprises the first C? and second Cg outer surface portions, wherein the third length L? is less than the fourth length Lg.

[0116] In the example of Figure 5, the sacrificial layer S can be interposed between the adhesive layer C and the trailing edge 74 of the blade 70.

[0117] Figures 6 to 9 illustrate a second embodiment of the blade 7 which differs from the blade 7 of the first embodiment by the position of the clamping element 9. In fact, the clamping element 9 is fixed by the layer of glue C for example only along the leading edge 73 of the blade 70. In particular, the adhesive layer C covers a portion of the lower surface 71 up to the leading edge 73, and the clamping element 9 covers a portion of the adhesive layer C (in particular that of the second external surface portion Cg) up to and beyond the leading edge 73. As a variant, the clamping element 9 can also be fixed along the upper surface 72 of the blade 70 of FIGS. 6 and 7 at the leading edge 73. According to another variant, another clamping element 9 can cover a portion of the trailing edge 74 of the blade 70 of these FIGS. 6 and 7, and possibly the lower surface 71 and / or the upper surface 72.

[0118] In the example of Figure 8, the sacrificial layer S can be interposed between at least a portion of the adhesive layer C and the leading edge 73 of the blade 70.

[0119] In particular, the sacrificial layer S can be interposed between at least a part of the adhesive layer C and the protective sheath 8 at least partially covering the leading edge 73, as illustrated schematically in FIG. 9.

[0120] Figure 10 illustrates a third embodiment of the blade 7 which differs from the blade 7 of the first and second embodiments by the clamping element 9. Indeed, a first clamping element 9a is fixed along the trailing edge 74 of the blade 70 and a second clamping element 9b is fixed along the leading edge 73 of this blade 70. In particular, the first clamping element 9 covers a portion of the extrados 72 up to and beyond the trailing edge 74, and the second clamping element 9 covers a portion of the intrados 71 up to and beyond the leading edge 73.

[0121] Alternatively, the first clamping element 9a may also be fixed along the intrados 71 of the blade 70 of FIG. 10 at the trailing edge 74, and the second clamping element 9b may also be fixed along the extrados 72.

[0122] According to another variant not shown in Figure 10, the blade 7 may comprise the protective sheath 8 covering at least in part the leading edge 73. In this configuration, the protective sheath 8 is covered at least in part by the second clamping element 9b, so that this protective sheath 8 is arranged between the second clamping element 9b (in particular at the level of the adhesive layer C or the sacrificial layer S) and the leading edge 73. The present application will now describe a method for manufacturing the blade 7 as described above, the successive steps of the method of which are for example summarized in Figure 11. The optional steps are shown in dotted lines in Figure 11.

[0123] According to the invention, the method comprises the following steps:

[0124] (a) providing the blade 7 comprising the blade 70 formed from the fiber preform embedded in resin, the blade 70 comprising the leading edge 73, the trailing edge 74, the intrados 71 and the extrados 72 connected to the intrados 71 by the leading edge 73 and the trailing edge 74,

[0125] (b) making the at least one clamping element 9 from thermosetting or thermoplastic material,

[0126] (c) mounting and fixing the at least one clamping element 9 so as to cover at least a portion of the leading edge 73 and / or the trailing edge 74.

[0127] In step (a), the composite material blade 7 can be made by any type of composite material process by injection or impregnation and polymerization of resin in the fiber preform, and in particular by the RTM resin injection molding technique. The fiber preform of the blade 70 can be obtained by three-dimensional fiber weaving or can be obtained by superimposing several layers / folds of fibers (i.e. in the form of multilayers).

[0128] Conventionally, the fibers, whether woven in three dimensions or draped, are inserted into a cavity of a manufacturing mold, then the resin is injected into the mold. Depending on the desired production rate, the resin is polymerized at room temperature or by heating.

[0129] In step (b), the clamping element 9 can be produced by any type of method for producing thermosetting or thermoplastic material, such as by molding, additive manufacturing, etc. In step (c), the clamping element 9 can be fixed along the leading edge 73 and / or the trailing edge 74, and possibly partly on the intrados 71 and / or the extrados 72.

[0130] Step (c) may comprise bonding (C2) of the clamping element 9 to at least one portion of the leading edge 73 and / or the trailing edge 74. For this, the bonding may be carried out by depositing a layer of adhesive C on at least one portion of the leading edge and / or the trailing edge, and possibly the intrados and the extrados. The bonding according to the invention thus makes it possible to avoid any reworking of the clamping element 9 and / or the blade 7 by machining, since the clamping element 9 can be geometrically controlled before bonding to the blade 70.

[0131] Bonding (C2) can be carried out in an autoclave under a vacuum cover, for example at a pressure between 2 and 10 bars and at a temperature between 100 and 200°C. For example, this temperature can be between 120 and 180°C.

[0132] Advantageously, before the bonding (C2), step (c) may comprise a deposition (ci) of the sacrificial layer S on at least one part of the leading edge 73 and / or the trailing edge 74, so as to place this sacrificial layer S between the adhesive layer C and the leading edge 73 and / or the trailing edge 74. Thus, one or more sacrificial layers S may be deposited on a part of the leading edge 73 and / or the trailing edge 74, and possibly the intrados 71 and / or the extrados 72.

Claims

CLAIMS 1. Blade (7) made of composite material for a turbomachine (10), in particular for an aircraft, said blade (7) comprising a blade (70) formed by a fiber preform embedded in a resin and extending along an axis (A) of elongation, said blade (70) comprising a leading edge (73), a trailing edge (74), a lower surface (71) and an upper surface (72) connected to the lower surface (71) by the leading edge (73) and the trailing edge (74), the blade (70) further comprising at least one serration element (9) formed of several teeth (92) and covering at least a portion of the leading edge (73) and / or the trailing edge (74), characterized in that said at least one serration element (9) is made of a thermosetting or thermoplastic material.

2. Blade made of composite material according to claim 1, characterized in that said at least one clamping element (9) is fixed along the leading edge (73) and / or the trailing edge (74) by a layer of glue (C).

3. Blade made of material according to claim 1 or 2, characterized in that said at least one serration element (9) extends from the leading edge (73) and / or the trailing edge (74) to a part of the intrados (71) and / or the extrados (72).

4. Blade made of composite material according to claim 2 or 3, characterized in that it comprises at least one sacrificial layer (S) comprising for example glass fibers, and interposed between the layer of glue (C) and the leading edge (73) and / or the trailing edge (74).

5. Blade made of composite material according to any one of the preceding claims, characterized in that it comprises a protective sheath (8) covering at least in part the leading edge (73), said at least one clamping element (9) at least in part covers this protective sheath (8).

6. Blade made of composite material according to any one of the preceding claims, characterized in that said at least one clamping element (9) has a length (Lg) measured along a longitudinal axis (B) perpendicular to the elongation axis (A), which represents between 20 and 100% of a length (L?) of chord of the blade (70) measured in the same plane (PL).

7. Blade made of composite material according to any one of the preceding claims, characterized in that said at least one serration element (9) has a height (Hg) measured along the elongation axis (A), which represents between 30 and 100% of a total height (H?) of the blade (70) measured in the same plane PH.

8. Blade made of composite material according to any one of the preceding claims, characterized in that the fiber preform is produced by weaving fibers in three dimensions or by superimposing several layers of fibers.

9. A composite material blade according to any one of the preceding claims, characterized in that the thermosetting material comprises an epoxy resin pre-impregnated with carbon fibers or glass fibers, or the thermoplastic material comprises a polyether-ether-ketone, poly-aryl-ether-ketone, poly-ether-imide or epoxy resin.

10. Blade made of composite material according to any one of claims 2 to 9, characterized in that the adhesive layer (C) has a thickness (Ec) of between 50 and 400 pm, preferably between 100 and 350 pm, the thickness (Ec) being measured along a cross-section to the axis of elongation (A).

11. Aircraft turbomachine (10) comprising at least one blade (7) made of composite material according to any one of the preceding claims.

12. Method for manufacturing a blade (7) made of composite material according to any one of claims 1 to 10, for a turbomachine (10), in particular an aircraft, the method comprising the following steps: (a) providing the blade (7) comprising the blade (70) formed from the fiber preform embedded in resin, the blade (70) comprising the leading edge (73), the trailing edge (74), the intrados (71) and the extrados (72) connected to the intrados (71) by the leading edge (73) and the trailing edge (74), (b) producing said at least one clamping element (9) from thermosetting or thermoplastic material, (c) mounting and fixing said at least one clamping element (9) so as to cover at least a portion of the leading edge (73) and / or the trailing edge (74).

13. Manufacturing method according to the preceding claim, characterized in that step (c) comprises gluing (C2) said at least one clamping element (9) on at least one part of the leading edge (73) and / or the trailing edge (74).

14. Manufacturing method according to the preceding claim, characterized in that the bonding (C2) is carried out in an autoclave under a vacuum cover, for example at a pressure of between 2 and 10 bars and at a temperature of between 100 and 200°C.

15. Manufacturing method according to claim 13 or 14, characterized in that before the bonding (C2), step (c) comprises a deposition (ci) of a sacrificial layer (S) on at least one part of the leading edge (73) and / or the trailing edge (74), so as to place this sacrificial layer (S) between the layer of glue (C) and the leading edge (73) and / or the trailing edge (74).

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