METHOD FOR MANUFACTURING A BLADE

Abstract

The invention relates to a method for manufacturing a turbine blade for an aircraft turbomachine, the method comprising the following steps: (a) providing a blade comprising upper and lower surfaces connected by a leading edge and a trailing edge, the blade comprising a composite material, (b) forming a protective shield on the blade, characterized in that step (b) comprises the following chronological substeps: (b1) depositing a first layer on the blade by dynamic cold gas spraying, and (b2) depositing a second metallic layer on the first layer by concentrated energy deposition. Abbreviated figure: Figure 5

Description

Title of the invention: METHOD FOR MANUFACTURING A BLADE Technical field of the invention

[0001] 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 metal protective shield arranged on the blade. Technical background

[0002] An aircraft turbomachine typically comprises an engine comprising, from upstream to downstream in the direction of gas flow, a fan rotatable 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.

[0003] The blower allows the suction of an air flow dividing 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.

[0004] The primary flow is compressed within the compressors. The compressed air is then mixed with a fuel and burned within the combustion chamber. The gases resulting from the combustion pass through the turbines and then escape through the nozzle, the cross-section of which allows the acceleration of these gases to generate propulsion.

[0005] Turbomachine components such as the fan, compressors or turbines comprise blades that act on the air flow. For example, compressor blades act to compress the primary air flow and fan blades act to compress the secondary air flow.

[0006] A blade comprises a blade which has an aerodynamic shape and thus has 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, for example, reinforcing fibers 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 blade 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, for example, the leading edge and the first lateral fin extends over the pressure face and the second lateral fin extends over the extrados face of the blade. The protective shield is typically bonded to the blade. In this In effect, the blade further comprises a layer of adhesive arranged between the blade and the protective shield. The layer of adhesive typically comprises a polymeric material such as an epoxy resin.

[0007] A method of manufacturing the aforementioned blade comprises the following steps:

[0008] - manufacturing the blade by molding,

[0009] - manufacturing of the protective shield by forging for example.

[0010] The method then comprises steps of surface treatment of the blade and the protective shield to enable efficient bonding between these two parts. The leading edge is typically subjected to chemical machining while the shield is subjected to sandblasting.

[0011] The method then comprises a step of depositing a layer of glue and then pairing the protective shield on the blade. Finally, a step of polymerizing the layer of glue is generally carried out.

[0012] Such a manufacturing method is not entirely satisfactory. Indeed, the step of manufacturing the protective shield presents challenges because the geometry of the protective shield is complex. This step is therefore tedious. Furthermore, the manufacturing tolerances are low and it is difficult to ensure the repeatability of this step.

[0013] Furthermore, the step of pairing the protective shield which consists of attaching the protective shield to the blade is carried out manually so that, guaranteeing the precision of the assembly of the blade and the repeatability of this step, require perfect mastery of the manufacturing process to limit defects on the protective shield or the blade and therefore the scrap of these parts.

[0014] Furthermore, prior to bonding, although particular attention is paid to the preservation of the protective shield and the blade, these parts may be polluted before they are bonded. Such pollution may lead to detachments or geometric defects also generating scrap of these parts.

[0015] Consequently, there is a need to provide a simple method to implement, for the manufacture of a blade for an aircraft turbomachine, comprising a blade made of composite material, which is resistant to shocks and erosion. Summary of the invention

[0016] To this end, the invention proposes a method of manufacturing a blade for an aircraft turbomachine, the method comprising the following steps:

[0017] (a) providing a blade comprising extrados and intrados faces connected by an edge leading edge and a trailing edge, the blade comprising a composite material,

[0018] (b) forming a protective shield on the blade.

[0019] The method is remarkable in that step (b) comprises the chronological sub-steps following logics:

[0020] (bl) depositing a first layer on the blade by dynamic projection using cold gas, and

[0021] (b2) depositing a second metal layer on the first layer by under-deposition concentrated energy.

[0022] According to the method of the invention, the protective shield is formed directly on the blade in two stages.

[0023] The concentrated energy deposition process, also known by the English acronym DED for "Directed Energy Deposition", makes it possible to form a metallic layer having the mechanical criteria of the protective shield. Indeed, the second metallic layer deposited using such a process has the properties of a forged or cast part.

[0024] Also, such a process makes it possible to produce any type of part with complex geometry.

[0025] However, such a process involves a high temperature thermal cycle ranging beyond the melting temperature of the material of the second layer, typically beyond 1660°. Thus, this process involves melting of the substrate on which it is implemented. However, such a thermal cycle is incompatible with a substrate made of composite material and therefore a blade having a blade made of composite material since the matrix of the composite material would not withstand such temperatures.

[0026] It has therefore been proposed according to the invention to form a first layer by a cold gas dynamic projection process, also known as "cold spray". This process makes it possible to form a first layer without heat input. This makes it possible to preserve the chemical and mechanical integrity of the blade which is made of composite material and to move the zone thermally affected by the concentrated energy deposition process away from the blade.

[0027] The combination of these methods therefore makes it possible to ensure that the isotherms likely to affect the blade do not penetrate sufficiently deeply into the blade, thus making it possible to protect the composite material of the blade from thermal degradation.

[0028] Also, the implementation of the concentrated energy deposition process improves the interface between the blade and the first layer because it enables the diffusion, recrystallization and therefore the relaxation of the constraints of the agglomerated particles forming the first layer to be activated.

[0029] Thanks to the combination of these two processes, it is therefore possible to form a protective shield of complex geometry in a precise and repeatable manner, thus reducing the risks of defects in the protective shield and its scrapping.

[0030] Furthermore, thanks to the combination of these two processes, it is possible to avoid a step of manufacturing the protective shield in favor of a step of forming of the protective shield directly on the blade. This eliminates the need for a step of pairing the shield on the blade, thus simplifying the process of assembling a blade and a protective shield and reducing blade assembly defects.

[0031] The invention may comprise one or more of the following features, taken in isolation from one another or in combination with one another:

[0032] — step (bl) comprises the following chronological sequences:

[0033] projecting a first powder comprising a mixture of metallic and polymeric particles,

[0034] during projection, gradually increase the mass fraction of metal particles in the mixture;

[0035] - the first layer comprises a metallic material, preferably an alloy of titanium or a polymeric material or a mixture of metallic and polymeric materials;

[0036] - the second layer comprises a titanium alloy;

[0037] - the first layer has a first thickness of between 5 μm and 200 μm pm ;

[0038] - the second layer has a second thickness greater than the thickness of the first layer;

[0039] - in step (b), the protective shield is formed on the leading edge and / or the edge of escape;

[0040] - the protective shield has serrations;

[0041] - after step (b), a step (c) of machining the protective shield is carried out;

[0042] - the composite material comprises a polymer matrix and fibers embedded in the matrix;

[0043] - the polymer matrix is ​​chosen from epoxy resins. Brief description of the figures

[0044] 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:

[0045] [Fig.l] is a schematic view in longitudinal section of a half aircraft turbomachine,

[0046] [Fig.2] is a schematic perspective view of a blade equipping the turbomachine of [Fig.l],

[0047] [Fig.2a] is a schematic view of a blade according to a preferred embodiment of the invention and equipping the turbomachine of [Fig.l],

[0048] [Fig.3] is a cross-sectional view of the blade of [Fig.2],

[0049] [Fig.4] is another schematic sectional view of the blade of [Fig.2],

[0050] [Fig.5] is a schematic view of a manufacturing method according to the invention,

[0051] [Fig.6] is a schematic sectional view of the blade during step (b2) of the method according to the invention. Detailed description of the invention

[0052] An example of an aircraft turbomachine 1 according to the invention is shown in [Fig.l]. The turbomachine 1 extends around and along a longitudinal axis A.

[0053] In the present application, the terms "axial", "axially", "radial" and "radially" are defined with respect to the longitudinal axis A.

[0054] The terms “upstream” and “downstream” are defined in relation to the direction of circulation of the gases in the turbomachine 1 along the longitudinal axis A.

[0055] The terms “internal”, “interior”, “external”, “exterior”, “externally” are defined in relation to the distance from the longitudinal axis A along a radial axis perpendicular to the longitudinal axis A.

[0056] 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).

[0057] 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).

[0058] The turbomachine 1 further comprises a rectifier 10. The rectifier 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 [Fig.l], the rectifier 10 is located downstream of the fan 2 and makes it possible to straighten the secondary flow F2.

[0059] 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.

[0060] 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 escape finally through the nozzle whose section 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.

[0061] The fan 2 or the rectifier 10 comprise blades 11. The blades 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.

[0062] The blades 11 extend radially relative to the longitudinal axis A.

[0063] With reference to [Fig.2], each blade 11 comprises a blade 12 and a protective shield 14 arranged on the blade 12.

[0064] 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 the blade 11 has been mounted 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.

[0065] The blade 12 also extends longitudinally along the elongation axis X between a first end and a second end opposite the first end.

[0066] The blade 12 comprises a composite material. The composite material comprises a polymer matrix and a fibrous reinforcement embedded in the matrix. The composite material is for example an organic matrix composite (OMC).

[0067] The matrix is ​​for example a thermoplastic or thermosetting polymer matrix. The thermosetting material is for example chosen from epoxy resins.

[0068] The fiber reinforcement comprises fibers which are, for example, carbon fibers or glass fibers. The fibers are organized, for example, in the form of a fiber preform.

[0069] 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 metallic material. The metallic material is, for example, titanium or a titanium alloy. Preferably, the titanium alloy is grade TA6V.

[0070] The protective shield 14 is intended to protect the blade 12 from external impacts. The protective shield 14 has an elongated dihedral shape. As best seen in [Fig. 3], the protective shield 14 has a cross-section relative to the elongation axis X in the shape of a V or U. The protective shield 14 comprises a first lateral fin 14a and a second fin lateral 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 blade 12 is arranged.

[0071] The first lateral fin 14a has a first longitudinal end, along the elongation axis X, free and the second lateral fin 14b has a second longitudinal end, along the elongation axis X, free which are opposite the core 14j. The longitudinal ends extend along the blade 12. The longitudinal ends extend respectively on the intrados face 12i and the extrados face 12b of the blade 12.

[0072] 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.

[0073] According to a first example illustrated in Figures 2 and 3, the protective shield 14 is arranged on the leading edge 12a. 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.

[0074] As visible in [Fig.2a], the protective shield 14 advantageously has serrations or serrations 14' along the elongation axis X. These serrations 14' make it possible to reduce the noise generated by the fan 2.

[0075] The serrations 14' comprise an alternation of teeth 14a' and hollows 14b' along the elongation axis X. Advantageously, the serrations 14' are formed on the core 14j of the protective shield 14 which connects the first and second lateral fins 14a, 14b.

[0076] With reference to [Fig.4], the protective shield 14 has a first layer 16 and a second metallic layer 18 arranged on the first layer 16.

[0077] The first layer 16 has a first thickness el. This first thickness el is advantageously between 5 μm and 200 μm, in particular between 10 μm and 15 μm.

[0078] According to a first example, the first layer 16 comprises a metallic material, advantageously a titanium alloy. The titanium alloy is preferably the TA6V grade. According to a second example, the first layer 16 comprises a polymeric material, advantageously a thermoplastic. According to a third particularly preferred example, the first layer 16 comprises a mixture of polymeric and metallic materials.

[0079] The first layer 16 is formed by a cold gas dynamic projection process, also known as “cold spray”.

[0080] The second layer 18 has a second thickness e2 greater than the first thickness e1. The second thickness e2 is advantageously between 10 pm and 300 pm, in particular between 20 pm and 300 pm and even more particularly between 20 pm and 80 pm. The second layer 18 comprises a metallic material distinct from or advantageously identical to the metallic material of the first layer 16. The second layer 18 thus comprises a metallic material, advantageously a titanium alloy. The titanium alloy is preferably the TA6V grade.

[0081] The second layer 18 is formed by a concentrated energy deposition process, also known by the English acronym DED for “Directed Energy Deposition”.

[0082] A method of manufacturing the blade 11 according to the invention will now be described with reference to [Fig.5].

[0083] The method comprises the following chronological steps:

[0084] (a) providing the blade 12,

[0085] (b) forming the protective shield 14 on the blade 12 by:

[0086] (bl) depositing the first layer 16 on the blade 12 by dynamic projection by cold gas, then

[0087] (b2) depositing the second layer 18 on the first layer 16 by deposition under energy concentrated, and

[0088] (c) optionally, machine the protective shield 14.

[0089] 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 draping.

[0090] A surface preparation step may be carried out after the blade 12 has been supplied. The surface preparation step may be, for example, chemical machining of the blade 12.

[0091] The step (bl) of dynamic projection by cold gas is carried out with a powder comprising particles. The particles comprise, for example, a metallic or polymeric material chosen in particular from thermoplastics or a mixture of metallic particles and polymeric particles, these materials. The metallic particles advantageously comprise a titanium alloy, in particular grade TA6V. The particles have, for example, a size of between 10 μm and 150 μm.

[0092] Advantageously, step (bl) comprises the following chronological sequences:

[0093] - projecting a first powder comprising a mixture of metallic particles and polymeric in which advantageously the mass fraction of metallic particles in the mixture is greater than or equal to 0, preferably 0.2

[0094] - during projection, gradually increase the mass fraction of particles metallic in the mixture until advantageously obtaining a mass fraction of metallic particles in the mixture equal to 1.

[0095] Such an embodiment of step (bl) makes it possible to ensure a progressive transition and a good mechanical connection with the blade 12.

[0096] The particles are transported by a gas. The gas is, for example, nitrogen or argon.

[0097] With reference to [Fig.6], the concentrated energy deposition step (b2) is carried out as follows:

[0098] - depositing a medium 20 on the first layer 16, and

[0099] - fusing the media 20 when depositing the media 20 on the first layer 16.

[0100] The medium 20 is for example in the form of powder or wire. The medium 20 comprises a metallic material selected from titanium alloys, in particular grade TA6V. The media is for example deposited using a nozzle, for example mounted on an articulated arm. According to another example, the nozzle is fixed and the blade 12 is mounted on a movable support.

[0101] The melting of the medium 20 is carried out by a heat source 22. The heat source 22 is for example a laser beam, an electron beam or plasma. The medium 20 is heated during its deposition to a temperature higher than the melting temperature of the medium 20. The temperature is for example greater than or equal to 1660°.

[0102] During step (b), the molten media 20 F is located on the surface of the first layer 16. This makes it possible to create the second layer 18 after hardening.

[0103] The zone Z adjacent to the media 20 in the molten state F and which is thermally affected by heat (ZAT) is concentrated and contained in the first layer 16. The first layer 16 thus makes it possible to protect the blade 12 from thermal degradation during the deposition of the second layer 18.

[0104] Between steps (bl) and (b2), a sub-step (b') of surface treatment of the first layer 16 can be carried out. The surface treatment is for example sandblasting.

[0105] Machining step (c) may be chemical machining or mechanical machining such as cutting.

[0106] Thanks to the combination of these two methods, it is therefore possible to form the protective shield 14 of complex geometry, in particular with teeth, in a precise and repeatable manner, thus reducing the risks of defects in the protective shield 14 and its scrapping.

[0107] Furthermore, thanks to the combination of these two methods, it is possible to dispense with a step of manufacturing the protective shield 14 in favor of a step of forming the protective shield 14 directly on the blade 12. This also makes it possible to avoid a step of pairing and bonding the protective shield 14 to the blade 12, thus simplifying the assembly process of the blade 11 and the protective shield 14 and reducing assembly defects of the blade 11.

[0108] Thanks to the method of the invention, it is also possible to repair the blade 11 by repeating step (b2).

Claims

Claims

1. A method of manufacturing a blade (11) for an aircraft turbomachine (1), the method comprising the following steps: (a) providing a blade (12) comprising extrados (12e) and intrados (12i) faces connected by a leading edge (12a) and a trailing edge (12b), the blade (12) comprising a composite material, (b) forming a protective shield (14) on the blade (12), characterized in that step (b) comprises the following chronological sub-steps: (bl) depositing a first layer (16) on the blade (12) by dynamic projection by cold gas, and (b2) depositing a second metallic layer (18) on the first layer (16) by deposition under concentrated energy.

2. Method according to the preceding claim, characterized in that the first layer (16) comprises a metallic material, preferably a titanium alloy or a polymeric material or a mixture of metallic and polymeric materials.

3. A method according to any preceding claim, characterized in that the second layer (18) comprises a titanium alloy.

4. Method according to any one of the preceding claims, characterized in that the first layer (16) has a first thickness (el) of between 5 pm and 200 pm.

5. Method according to the preceding claim, characterized in that the second layer (18) has a second thickness (e2) greater than the thickness (el) of the first layer (16).

6. A method according to any one of the preceding claims, characterized in that in step (b), the protective shield (14) is formed on the leading edge (12a) and / or the trailing edge (12b).

7. Method according to any one of the preceding claims, characterized in that the protective shield (14) has serrations (14').

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

9. A method according to any one of the preceding claims, characterized in that the composite material comprises a matrix po-

10. polymer and fibers embedded in the matrix. Method according to the preceding claim, characterized in that the polymer matrix is ​​chosen from epoxy resins.

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