Method for manufacturing a vane

Abstract

The invention relates to a method for manufacturing a vane (11) for an aircraft turbine engine (1), comprising the following chronological steps: (a) providing a blade (12) comprising a suction face (12e) and a pressure face (12i) connected by a leading edge (12a) and a trailing edge (12b), the blade (12) comprising a composite material, (b) arranging a bonding layer (16) on the blade (12), the bonding layer (16) comprising a polymer material, the method being characterised in that it further comprises the following chronological steps performed after step (b): (c) forming a protective shield (14) on the blade (12) by thermal spraying, (d) optionally, heat-treating the bonding layer (16).

Description

[0001] DESCRIPTION

[0002] TITLE: METHOD FOR MANUFACTURING A BLADE

[0003] Technical field of the invention

[0004] The invention relates to the field of methods for manufacturing blades for aircraft turbomachines. The invention relates more particularly to the field of manufacturing blades comprising a blade made of composite material and a protective shield arranged on the blade.

[0005] Technical background

[0006] The state of the art is illustrated by documents US-A1 -2016 / 0305442 and FR-A1 -3008 109.

[0007] An aircraft turbomachine typically comprises an engine comprising, from upstream to downstream in the direction of gas flow, a fan rotating about a longitudinal axis, a low-pressure compressor and a high-pressure compressor, a combustion chamber, a high-pressure turbine and a low-pressure turbine and a gas exhaust nozzle.

[0008] The blower allows the suction of an air flow divided into a primary flow and a secondary flow. The primary flow passes through a primary vein of the turbomachine while the secondary flow is directed towards a secondary vein surrounding the primary vein.

[0009] The primary flow is compressed within the compressors. The compressed air is then mixed with fuel and burned within the combustion chamber. The gases from the combustion pass through the turbines and then escape through the nozzle, whose cross-section allows the acceleration of these gases to generate propulsion.

[0010] Turbomachinery components such as the fan, compressors, or turbines include blades that act on the airflow. For example, compressor blades compress the primary airflow, and fan blades compress the secondary airflow.

[0011] A blade comprises a blade which has an aerodynamic shape and thus comprises a pressure face and an extrados face connected to the pressure face by a leading edge and a trailing edge. To reduce the weight of the blade, the blade is made of a composite material. The composite material of the blade comprises reinforcing fibres embedded in a polymer matrix. In order to protect the blade, in particular a fan blade, from degradation caused by the impact of foreign bodies and from erosion, it is known to cover the leading edge with a protective shield. The protective shield has a dihedral shape comprising a first lateral fin and a second lateral fin connected by a core. The core covers the leading edge and the first lateral fin extends over the pressure face and the second lateral fin extends over the extrados face. The protective shield is typically bonded to the blade.For this purpose, the blade further comprises a bonding layer arranged between the blade and the protective shield. The bonding layer typically comprises a polymeric material such as an epoxy resin.

[0012] A method of manufacturing the aforementioned blade comprises the following steps:

[0013] - manufacturing of the blade,

[0014] - manufacturing of the protective shield.

[0015] The process then includes a step of pairing and bonding the protective shield to the blade.

[0016] This manufacturing process is not entirely satisfactory. Indeed, the protective shield manufacturing step presents challenges because the configuration of the protective shield is complex. This step is therefore tedious. Furthermore, manufacturing tolerances are tight and it is difficult to ensure the repeatability of this step.

[0017] Furthermore, the protective shield matching step, which consists of attaching the protective shield to the blade, is carried out manually, so ensuring the precision of the blade assembly and the repeatability of this step requires perfect mastery of the manufacturing process to limit defects on the protective shield or the blade and therefore the scrap of these parts.

[0018] Therefore, there is a need to provide a simple to implement method for manufacturing a blade for an aircraft turbomachine, comprising a composite material blade, which is resistant to shocks and erosion.

[0019] Summary of the invention

[0020] To this end, the invention proposes a method of manufacturing a blade for an aircraft turbomachine, the method comprising the following chronological steps:

[0021] (a) providing a blade comprising an extrados face and a intrados face connected by a leading edge and a trailing edge, the blade comprising a composite material,

[0022] (b) arranging a bonding layer on the blade, the bonding layer comprising a polymeric material.

[0023] The method is particularly remarkable in that it further comprises the following chronological steps after step (b):

[0024] (c) forming a protective shield on the blade by thermal spraying,

[0025] (d) heat treating the bonding layer.

[0026] The thermal spray process allows for the formation of complex geometric coatings in a precise and repeatable manner, thus reducing the risk of defects in the protective shield and its scrapping.

[0027] Furthermore, thanks to the thermal spraying process, it is possible to dispense with a protective shield manufacturing step in favor of a protective shield forming step directly on the blade. This also makes it possible to dispense with a shield pairing step on the blade, thus simplifying the assembly process of a blade and a protective shield and reducing blade assembly defects. The thermal spraying process also allows the implementation of a variability of materials for the protective shield that is not offered by other shield manufacturing processes.

[0028] The method comprises, after the step of forming the protective shield, a step of heat treatment of the bonding layer. The heat treatment step allows the polymerization of the bonding layer and subsequently, the fixing by gluing of the protective shield on the blade.

[0029] The invention may comprise one or more of the following features, taken in isolation from each other or in combination with each other:

[0030] - thermal projection is carried out with a powder comprising metallic or ceramic particles or a mixture of these;

[0031] - at least 90% of the particles have a size (d90) between 5 pm and 200 pm, advantageously between 10 pm and 150 pm;

[0032] - at most 10% of the particles have a size (d10) of between 0.05 pm and 50 pm, advantageously of between 0.1 pm and 20 pm;

[0033] - the bonding layer has a thickness of between 50 pm and 500 pm, advantageously between 80 pm and 400 pm;

[0034] - step (d) is carried out at a temperature between 100°C and 200°C, advantageously between 140°C and 190°C;

[0035] - in step (d) the blade is placed in a heating device such as an oven, furnace or autoclave;

[0036] - the method comprises between steps (b) and (c) a step (f) of heat pre-treatment of the bonding layer;

[0037] - step (f) is carried out at a temperature less than or equal to 150°C, advantageously less than or equal to 130°C, or advantageously still less than 100°C;

[0038] - the method comprises, after step (d), a step (e) of machining the protective shield; -- step (c) of thermal spraying is carried out along the leading edge and / or the trailing edge of the blade.

[0039] Brief description of the figures

[0040] Other characteristics and advantages will emerge from the following description of a non-limiting embodiment of the invention with reference to the appended drawings in which: Figure 1 is a schematic view in longitudinal section of a half-turbomachine of an aircraft, Figure 2 is a schematic perspective view of a blade equipping the turbomachine of Figure 1, Figure 3 is a cross-sectional view of the blade of Figure 2, Figure 4 is another schematic sectional view of the blade of Figure 2, Figure 5 is a schematic view of a manufacturing method according to the invention, Figure 6 is a schematic view of the manufacturing method according to an advantageous embodiment of the invention.

[0041] Detailed description of the invention

[0042] An example of an aircraft turbomachine 1 according to the invention is shown in Figure 1. The turbomachine 1 extends around and along a longitudinal axis A.

[0043] In this application, the terms "axial", "axially", "radial" and "radially" are defined with respect to the longitudinal axis A.

[0044] The terms “upstream” and “downstream” are defined in relation to the direction of gas flow in the turbomachine 1 along the longitudinal axis A.

[0045] The terms "internal", "interior", "external", "exterior",

[0046] "externally" are defined relative to the distance from the longitudinal axis X along a radial axis perpendicular to the longitudinal axis A.

[0047] The turbomachine 1 extends around a longitudinal axis A. It comprises from upstream to downstream in the direction of flow of the gases F along the longitudinal axis A, a fan 2, at least one compressor such as a low pressure compressor 3 and a high pressure compressor 4, a combustion chamber 5, at least one turbine 6 such as a high pressure turbine and a low pressure turbine, and a nozzle (not shown).

[0048] The rotor of the low pressure turbine is connected to the fan 2 and to the rotor of the low pressure compressor 3 by a low pressure shaft (not shown). The rotor of the high pressure turbine is connected to the rotor of the high pressure compressor 4 by a high pressure shaft (not shown).

[0049] The turbomachine 1 also includes a rectifier 10. The rectifier

[0050] 10 makes 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 1. In the particular example of figure 1, the rectifier 10 is located downstream of the fan 2 and makes it possible to straighten the secondary flow F2.

[0051] The blower 2 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 1 while the secondary flow F2 is directed towards a secondary vein surrounding the primary vein.

[0052] 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 a 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 10, which accelerates the circulation speed of the secondary flow F2 to generate propulsion.

[0053] The blower 2 or the rectifier 10 comprises blades 11. The blades

[0054] 11 equipping the rectifier 10 are known by the English term "Outlet Guide Vane" (OGV). The blades 11 are movable or fixed in rotation about the longitudinal axis A. Typically, the blades 11 of the fan 11 are movable in rotation. The blades 11 of the rectifier 10 are fixed. The blades 11 extend radially relative to the longitudinal axis A.

[0055] Referring to Figure 2, each blade 11 comprises a blade 12 and a protective shield 14 arranged on the blade 12.

[0056] The blade 12 extends along an elongation axis X. The elongation axis X of the blade 12 extends radially relative to the longitudinal axis A of the turbomachine 1 after mounting the blade 11 on the turbomachine 1. The blade 12 has an aerodynamic profile. The blade 12 thus comprises an extrados face 12e and a intrados face 12i connected by a leading edge 12a and a trailing edge 12b. The blade 12 thus extends along a transverse axis Y between the leading edge 12a and the trailing edge 12b.

[0057] The blade 12 also extends longitudinally along the elongation axis X between a first end and a second end opposite the first end.

[0058] The blade 12 comprises a composite material. The composite material comprises a polymer matrix and a fiber reinforcement embedded in the matrix. The composite material is, for example, an organic matrix composite (OMC). The matrix is, for example, a thermoplastic or thermosetting polymer matrix. The thermosetting material is, for example, an epoxy polymer. The fiber reinforcement comprises fibers that are, for example, carbon fibers or glass fibers. The fibers are organized, for example, in the form of a fiber preform.

[0059] The protective shield 14 is advantageously arranged on the leading edge 12a. It advantageously extends along the entire length of the leading edge 12a. The protective shield 14 is advantageously made of a metallic material. The metallic material is, for example, titanium or an alloy such as a steel, for example, a stainless steel or a nickel and cobalt alloy (NiCo), an aluminum, silver, zinc, nickel, copper or a mixture thereof.

[0060] The protective shield 14 is, according to another example, made of ceramic material. It is intended to protect the leading edge 12a from external impacts. The protective shield 14 has an elongated dihedral shape. As best seen in FIG. 3, the protective shield 14 has a V-shaped or U-shaped cross-section. The protective shield 14 comprises a first lateral fin 14a and a second lateral fin 14b connected to the first lateral fin 14a by a core 14j. The first and second lateral fins 14a, 14b define between them a cavity in which the leading edge 12a is arranged.

[0061] The first lateral fin 14a has a first free longitudinal end and the second lateral fin 14b has a second free longitudinal end which are opposite the core 14j. The free longitudinal ends extend along the blade 12. The free longitudinal ends extend respectively on the intrados face 12i and the extrados face 12b of the blade 12.

[0062] Advantageously, the thickness of the protective shield 14 is variable. The thickness of the core 14j is greater than the thicknesses of the first and second lateral fins 14a, 14b. Advantageously, the thickness of the first and second lateral fins 14a, 14b decreases in the direction of the trailing edge 12b of the blade 12. The first and second lateral fins 14a, 14b are tapered in the direction of the trailing edge 12b of the blade 12.

[0063] Alternatively, and in a manner not shown, the protective shield 14 is arranged on the trailing edge 12b. According to yet another alternative not shown, the blade 11 comprises two protective shields 14 arranged respectively on the leading edge 12a and the trailing edge 12b.

[0064] The protective shield 14 is very advantageously formed on the blade 12 by thermal spraying.

[0065] The blade 11 further comprises a bonding layer 16 arranged between the blade

[0066] 12 and the protective shield 14. The bonding layer 16 ensures the adhesion of the protective shield 14 to the blade 12 without damaging the blade 12. The bonding layer 16 is made of polymeric material. The polymeric material is advantageously chosen from thermosetting polymers such as an epoxy, silicone or polyurethane polymer or thermoplastic polymers or a mixture thereof. The epoxy polymer is for example the commercial material Redux® 322 from the company HEXCEL.

[0067] The bonding layer 16 is typically in the form of a film based on one of the materials mentioned. The film is for example the commercial film AF191 or AF3109 from the company 3M. Advantageously, the bonding layer 16 has a thickness of between 50 μm and 500 μm, even more advantageously of between 80 μm and 400 μm.

[0068] A method of manufacturing the blade 11 according to the invention will now be described with reference to FIG. 5.

[0069] The process includes the following chronological steps:

[0070] (a) provide blade 12,

[0071] (b) arranging the bonding layer 16 on the blade 12

[0072] (c) forming the protective shield 14 on the blade 12 by thermal spraying,

[0073] (d) heat treating the bonding layer 16,

[0074] (e) optionally, machine the protective shield 14.

[0075] Step (a) may be carried out by molding such as resin transfer molding known by the acronym RTM for “Resin Transfer Molding” in English or layup.

[0076] Step (b) may be performed by draping the film, or by manually applying the bonding layer material 16.

[0077] The thermal spraying process of step (c) is for example a flame spraying process for example by high velocity oxygen flame (HVOF for “High Velocity Oxy-Fuel” in English), by arc-wire, by suspension, by arc plasma, cold spray or any other applicable thermal spraying process.

[0078] Thermal spraying is carried out with a powder comprising particles of which at least 90% of these particles have a size (d90) of between 5 pm and 200 pm, advantageously of between 10 pm and 150 pm and preferably of which at most 10% of the particles have a size (d10) of between 0.05 pm and 50 pm, advantageously of between 0.1 pm and 20 pm.

[0079] In particular, in the case of atmospheric pressure plasma projection (APS), the powder comprises particles of which at least 90% of these particles have a size (d90) of 110 pm and of which at most 10% of the particles have a size (d10) of 10 pm.

[0080] In the case of suspension plasma projection (SPS), the powder is in the form of a suspension comprising particles of which at least 90% of these particles have a size (d90) of 10 pm and of which at most 10% of the particles have a size (d10) of 0.1 pm.

[0081] Advantageously, the particles are metal particles such as aluminum, silver, zinc, nickel, titanium, copper, an iron alloy, a metal alloy or ceramic particles such as non-oxide particles such as silicon carbide, or oxide particles such as aluminum oxide, zirconium oxide, silicon oxide, or a mixture of the particles.

[0082] The thermal spraying step is carried out on a homogeneous bonding layer 16 which contributes to the good adhesion of the protective shield 14 over the entire surface of the bonding layer 16.

[0083] The thermal spraying step (c) is carried out along the leading edge 12a and / or along the trailing edge 12b of the blade 12.

[0084] Step (d) allows the polymerization of the bonding layer 16 and the consolidation of the interface between the bonding layer 16 and the functional layer formed by the coating 14.

[0085] It can be carried out by heating to a temperature between 100°C and 200°C, advantageously between 120°C and 190°C, even more advantageously between 140°C and 190°C. Step (d) is preferably carried out for a duration greater than 10 min, greater than 15 min, greater than 20 min, greater than 30 min, even more preferably between 45 min and 200 min. The blade 11 is for example placed in a heating device such as an oven, a furnace or an autoclave.

[0086] The machining step (e) may be chemical machining or mechanical machining such as cutting. According to a particularly advantageous embodiment of the invention illustrated in FIG. 6, the method comprises between steps (b) and (c) a step (f) of heat pre-treatment of the bonding layer 16. This step may be carried out by heating the bonding layer 16 to a temperature less than or equal to 150°C for pre-polymerization, advantageously less than or equal to 130°C. This heat pre-treatment step advantageously makes it possible to avoid the loss of the bonding layer 16 during the thermal spraying step (c).

[0087] This same step (f) can be carried out at a lower temperature equal to 100°C to increase only the tack. This step can be carried out in an oven or an autoclave or an oven. This step makes it possible to increase the adhesion strength at the interface of the bonding layer 16 and the protective shield 14.

Claims

CLAIMS 1. Method for manufacturing a blade (11) for an aircraft turbomachine (1), the method comprising the following chronological steps: (a) providing a blade (12) comprising an extrados face (12e) and a intrados face (12i) connected by a leading edge (12a) and a trailing edge (12b), the blade (12) comprising a composite material, (b) arranging a bonding layer (16) on the blade (12), the bonding layer (16) comprising a polymeric material, the method being characterized in that it further comprises the following chronological steps after step (b): (c) forming a protective shield (14) on the blade (12) by thermal spraying, (d) heat treating the bonding layer (16).

2. Method according to the preceding claim, characterized in that the thermal projection is carried out with a powder comprising metallic or ceramic particles or a mixture thereof.

3. Method according to the preceding claim, characterized in that at least 90% of the particles have a size (d90) of between 5 pm and 200 pm, advantageously of between 10 pm and 150 pm.

4. Method according to the preceding claim, characterized in that at most 10% of the particles have a size (d10) of between 0.05 pm and 50 pm, advantageously of between 0.1 pm and 20 pm.

5. Method according to any one of the preceding claims, characterized in that the bonding layer (16) has a thickness of between 50 pm and 500 pm, advantageously between 80 pm and 400 pm.

6. Method according to any one of the preceding claims, characterized in that step (d) is carried out at a temperature between 100°C and 200°C, advantageously between 140°C and 190°C.

7. Method according to any one of the preceding claims, characterized in that in step (d) the blade (11) is placed in a heating device such as an oven, a furnace or an autoclave.

8. Method according to any one of the preceding claims, characterized in that it comprises between steps (b) and (c) a step (f) of heat pre-treatment of the bonding layer (16).

9. Method according to the preceding claim, characterized in that step (f) is carried out at a temperature less than or equal to 150°C, advantageously less than or equal to 130°C, or advantageously still less than 100°C.

10. Method according to any one of the preceding claims, characterized in that it comprises, after step (d), a step (e) of machining the protective shield (14).

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