Method for manufacturing a part for a propulsion assembly

Abstract

The invention relates to a method for manufacturing a part (15) for an aircraft propulsion assembly, the method comprising the following chronological steps: (a) providing a body (16) made of composite material comprising fibres embedded in a polymer matrix, (c) depositing a coating (17) on the body (16), characterised in that step (c) comprises the following sub-steps: (d) depositing a bonding layer (18) on the body (16) by high-power pulse magnetron cathode sputtering, and (e) depositing a functional layer (19) on the bonding layer (18) by thermal spraying.

Description

[0001] DESCRIPTION

[0002] TITLE: METHOD FOR MANUFACTURING A PART FOR A PROPULSIVE ASSEMBLY

[0003] Technical field of the invention

[0004] The invention relates to the field of manufacturing processes for parts for an aircraft propulsion system. More particularly, the invention relates to the field of manufacturing parts comprising a body made of composite material and a coating.

[0005] Technical background

[0006] A propulsion unit typically comprises a turbomachine and optionally a nacelle surrounding the turbomachine. Each turbomachine typically comprises, from upstream to downstream in the direction of gas flow, a movable 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.

[0007] The blower draws in an airflow that splits into a primary airflow and a secondary airflow. The primary airflow passes through a primary duct of the turbomachine, while the secondary airflow is directed towards a secondary duct surrounding the primary duct.

[0008] The primary airflow is compressed within the compressors. The compressed air is then mixed with fuel and burned in the combustion chamber. The combustion gases pass through the turbines and then escape through the nozzle, whose cross-section allows these gases to be accelerated to generate propulsion.

[0009] The propulsion assembly comprises components with a composite body and a coating applied to the body, which includes a functional layer. For example, the turbomachine includes blades that influence the airflow. For instance, compressor blades compress the primary airflow, and fan blades compress the secondary airflow.

[0010] A blade comprises a body, such as a blade with an aerodynamic shape, and thus includes an upper and lower surface connected to the lower surface by a leading edge and a trailing edge. To reduce the blade's weight, it is made of a composite material. The composite material of the blade includes reinforcing fibers embedded in a polymer matrix, particularly an epoxy matrix. To protect the blade from degradation caused by the impact of foreign objects and from erosion, it is common practice to coat the leading edge with a protective shield forming an impact-resistant and wear-resistant coating. This coating includes a functional impact-resistant and wear-resistant layer, usually metallic.

[0011] The coating is typically glued to the blade.

[0012] A manufacturing process for the aforementioned blade comprises the following steps:

[0013] - manufacturing of the blade,

[0014] - coating manufacturing,

[0015] - matching the coating to the blade, and

[0016] - bonding the coating to the blade.

[0017] This manufacturing process is not entirely satisfactory. The coating fabrication stage presents challenges due to the complex coating configuration, making it a tedious process. Furthermore, manufacturing tolerances are tight, and ensuring repeatability is difficult.

[0018] Furthermore, the coating application step, which involves attaching the coating to the blade, is performed manually. Therefore, ensuring the precision of the blade assembly and the repeatability of this step requires perfect mastery of the manufacturing process to minimize defects in the coating or the blade, and thus the scrapping of these parts. In this context, thermal spraying of the coating was considered. This process eliminates the manufacturing and coating application steps. However, since the blade is made of a composite material based on an epoxy matrix, it exhibits low wettability and lacks functional groups that would allow for adhesion of the thermally sprayed coating. Consequently, a significant limitation in implementing such a thermally sprayed coating is the sensitivity of the blade's carbon fibers to the impact of the particles projected during the thermal spraying process.

[0019] These issues concern all aircraft parts, including turbomachinery and nacelles, which include a body made of composite material, which needs to be equipped with a coating having a functional layer, such as anti-shock, anti-wear, anti-icing or even anti-lightning.

[0020] Therefore, there is a need to provide a manufacturing process for an aircraft part comprising a composite material body and a coating, which is compatible with the deposition of a functional layer by thermal spraying.

[0021] Summary of the invention

[0022] To this end, the invention proposes a method for manufacturing a part for an aircraft propulsion assembly, the method comprising the following chronological steps:

[0023] (a) provide a body made of composite material comprising fibres embedded in a polymer matrix,

[0024] (c) apply a coating to the body.

[0025] The process is remarkable in that step (c) comprises the following substeps:

[0026] (d) deposit a bonding layer on the body by magnetron sputtering in high-power pulse regime, and (e) deposit a functional layer on the bonding layer by thermal spraying, the bonding layer having a thickness of between 1 pm and 20 pm.

[0027] The coating thus comprises a bonding layer and a functional layer. The functional layer protects the part, for example, from wear, erosion, impacts, frost, or any attack related to its environment.

[0028] This functional layer is deposited by thermal spraying. The thermal spraying process allows for the precise and repeatable formation of coatings with complex geometries, thus reducing the risk of coating defects and their disposal.

[0029] Furthermore, thanks to the thermal spraying process, it is possible to eliminate a coating manufacturing step in favor of a coating formation step directly onto the body. This also eliminates the need for a coating matching step, thus simplifying the part assembly process.

[0030] The thermal spraying process also allows for the implementation of a variability of materials for the coating which is not offered by other coating manufacturing processes.

[0031] The bonding layer promotes adhesion of the functional layer to the body and protects the body during thermal spraying of the functional layer. Thanks to the bonding layer, this thermal spraying can be performed without damaging the body.

[0032] According to the invention, the bonding layer is deposited by magnetron sputtering in high-power impulse mode, also known by the English acronym HIPIMS for "high-power impulse magnetron sputtering".

[0033] Magnetron sputtering deposition in high-power pulse mode protects the body from high temperatures and impacts that could degrade the composite material. The invention may include one or more of the following features, taken individually or in combination:

[0034] -- the functional layer comprises a WCCo material deposited by HVOF and the bonding layer advantageously comprises titanium deposited by HIPIMS,

[0035] - between steps (a) and (c), a step (b) of surface treatment of the body,

[0036] - at the end of step (b), the body has a surface with a roughness Ra between 0.5 pm and 8 pm, preferably between 0.5 pm and 5 pm,

[0037] - the bonding layer comprises a metallic material,

[0038] - the bonding layer has a thickness between 1 pm and 10 pm,

[0039] - Step (d) includes the following sub-steps:

[0040] (d1) deposit a first bonding sublayer onto the body by magnetron sputtering in a high-power pulse regime, and

[0041] (d2) deposit a second bonding sublayer on the first bonding sublayer by magnetron sputtering in high-power pulse regime,

[0042] - the second bonding sublayer has a thickness between 5 pm and 15 pm,

[0043] - the functional layer has a thickness between 30 pm and 2000 pm,

[0044] - Step (e) includes the following sub-steps:

[0045] (e1) deposit a first functional sublayer onto the bonding layer by thermal spraying, and

[0046] (e2) deposit a second functional sub-layer onto the first functional sub-layer by thermal spraying,

[0047] - The first functional sublayer has a thickness between 40 µm and 120 µm. Brief description of the figures

[0048] Other features and advantages will become apparent from the following description of a non-limiting embodiment of the invention with reference to the accompanying drawings, in which: Figure 1 is a schematic longitudinal cross-sectional view of a propulsion assembly to which the invention can be applied; Figure 2 is a schematic cross-sectional view of a part according to the invention and equipping the propulsion assembly of Figure 1; Figure 3 is a schematic cross-sectional view of a part according to an embodiment of the invention; Figure 4 is a schematic cross-sectional view of a part according to an embodiment of the invention; Figure 5 is a schematic cross-sectional view of a part according to an embodiment of the invention; Figure 6 is a schematic representation of the manufacturing process according to the invention; Figure 7 is a schematic representation of an installation that can be implemented during the manufacturing process of the invention.Figure 8 is a schematic representation of a device that can be implemented in step (c) of the manufacturing process according to the invention.

[0049] Detailed description of the invention

[0050] An example of an aircraft propulsion assembly 1' is schematically illustrated in Figure 1.

[0051] The propulsion assembly 1' extends around and along a longitudinal axis A.

[0052] In this application, the terms "axial", "axially", "radial" and "radially" are defined with respect to the longitudinal axis A.

[0053] The terms "upstream" and "downstream" are defined in relation to the direction of gas flow in the propulsion assembly 1' along the longitudinal axis A. The terms "internal," "inside," "external," and "outside" are defined in relation to the direction of gas flow in the propulsion unit 1' along the longitudinal axis A.

[0054] "Externally" are defined with respect to the distance from the longitudinal axis A along a radial axis perpendicular to the longitudinal axis A.

[0055] The propulsion assembly 1' comprises a turbomachine 1 and optionally a nacelle 1a surrounding the turbomachine 1.

[0056] The turbomachine 1 extends around and along the longitudinal axis A. It comprises, from upstream to downstream in the direction of gas flow F along the longitudinal axis A, a blower 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 low-pressure turbine rotor is connected to the blower 2 and the low-pressure compressor rotor 3 by a low-pressure shaft (not shown). The high-pressure turbine rotor is connected to the high-pressure compressor rotor 4 by a high-pressure shaft (not shown).

[0058] The turbomachine 1 also includes a rectifier 10. The rectifier 10 rectifies the flow at the outlet of an upstream rotor 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 rectifies the secondary air flow F2.

[0059] The blower 2 allows the intake of an airflow which splits into a primary airflow F1 and a secondary airflow F2. The primary airflow F1 passes through a primary channel of the turbomachine 1 while the secondary airflow F2 is directed towards a secondary channel surrounding the primary channel.

[0060] The primary airflow 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 in the combustion chamber 5. The gases formed by combustion pass through the high-pressure turbine and the low-pressure turbine. Finally, the gases escape through the nozzle, the cross-section of which accelerates these gases to generate propulsion. The secondary airflow F2 passes through the rectifier 10, which accelerates the flow velocity of the secondary airflow F2 to generate propulsion.

[0061] The blower 2 and / or the straightener 10 each comprise blades 11 which are supported by an annular disk and centered on the longitudinal axis A. The blades 11 equipping the straightener 10 are known by the English term "Outlet Guide Vane" (OGV). The blades 11 are either movable or fixed in rotation about the longitudinal axis A. Typically, the blades 11 of the blower 11 are movable in rotation and the blades 11 of the straightener 10 are fixed in rotation about the longitudinal axis A.

[0062] The blades 11 extend radially with respect to the longitudinal axis A and are regularly distributed around the longitudinal axis A.

[0063] Each blade 11 comprises a blade 12 and a foot 13.

[0064] The blade 12 extends along an aspect ratio axis X. The aspect ratio axis X of the blade 12 extends radially with respect to the longitudinal axis A of the turbomachine 1 after the blade 11 is mounted on the turbomachine 1. The blade 12 has an aerodynamic profile. The blade 12 thus comprises an upper surface and an lower surface 12i connected by a leading edge 12a and a trailing edge 12b. The blade 12 thus extends along a transverse axis between the leading edge 12a and the trailing edge 12b.

[0065] Blade 12 comprises a composite material. The composite material is, for example, an organic matrix composite (OMC). The composite material comprises a polymer matrix and a fibrous reinforcement embedded within the matrix.

[0066] The matrix is ​​a polymer matrix. The matrix polymer comprises a thermoplastic or thermosetting polymer. The thermosetting polymer is, for example, an epoxy polymer.

[0067] Fiber reinforcement comprises fibers such as carbon fibers, polyester, polypropylene, aramid, or glass fibers. These fibers can be short or long. Preferably, the fibers are woven into a three-dimensional fiber preform.

[0068] In a particularly preferred manner, the composite material comprises an epoxy matrix and carbon fibers embedded in the epoxy matrix.

[0069] The foot 13 allows the blade 12 to be anchored to the disc. It may comprise a composite material identical to the composite material of the blade 12. It forms a single piece with the blade 12.

[0070] Each dawn 11 may include a protective shield 14 which is advantageously arranged on the leading edge 12a. It extends advantageously all along the leading edge 12a.

[0071] The protective shield 14 is designed to protect the leading edge 12a from external impacts and erosion. The protective shield 14 has an elongated, dihedral shape. The protective shield 14 has a V- or U-shaped cross-section. The protective shield 14 comprises a first lateral fin and a second lateral fin connected to the first lateral fin by a web. The first and second lateral fins define a cavity in which the leading edge 12a is arranged.

[0072] The first lateral fin has a first free longitudinal end, and the second lateral fin has a second free longitudinal end, both of which are opposite the web. The free longitudinal ends extend along the blade 12. Specifically, they extend over the lower surface (intrados) 12i and the upper surface (extrados) of the blade 12.

[0073] Alternatively, and not shown, the protective shield 14 is arranged on the trailing edge 12b. According to yet another alternative, not shown, the vane 11 comprises two protective shields 14 arranged respectively on the leading edge 12a and on the trailing edge 12b.

[0074] The nacelle 1a is annular and centered on the longitudinal axis A. It extends around the turbomachine 1. The nacelle 1a comprises at least one cowling 1b including a panel 1c and a covering arranged on the panel 1c. The panel 1c comprises a composite material which may be identical to the composite material of the blade 12. The covering may be a lightning-resistant covering, for example.

[0075] With reference to figures 2 to 5, the propulsion assembly 1' comprises at least one part 15. Part 15 is, according to a first example, the blade 11 or, according to a second example, the hood 1b of the nacelle 1a.

[0076] Part 15 includes a body 16 comprising a composite material and a coating 17 arranged on the body 16.

[0077] The composite material is, for example, an organic matrix composite (OMC). The composite material comprises a polymer matrix and a fibrous reinforcement embedded within the matrix.

[0078] The matrix is ​​a polymer matrix. The matrix polymer comprises a thermoplastic or thermosetting polymer. The thermosetting polymer is, for example, an epoxy polymer.

[0079] Fiber reinforcement comprises fibers such as carbon fibers, polyester, polypropylene, aramid, or glass fibers. These fibers can be short or long. Preferably, the fibers are woven into a three-dimensional fiber preform.

[0080] In a particularly preferred manner, the composite material comprises an epoxy matrix and carbon fibers embedded in the epoxy matrix.

[0081] The body 16 comprises a surface having a roughness Ra between 0.5 µm and 8 µm, preferably between 0.5 µm and 5 µm. Such roughness promotes the adhesion of the coating 17.

[0082] The body 16 is, according to the first example, the blade 12 or the foot 12, or according to the second example, the panel 1c of the hood 1b of the nacelle 1a.

[0083] Coating 17, for example, is an anti-shock and / or anti-erosion, anti-wear, anti-lightning or even anti-frost coating.

[0084] The coating 17 is, according to the first example, the protective shield 14 of the blade 11 or, according to the second example, the coating of the panel 1c of the hood 1b of the nacelle 1a. The coating 17 comprises a bonding layer 18 and a functional layer 19.

[0085] The bonding layer 18 is located between the body 16 and the functional layer 19. The bonding layer 18 has a thickness e1 between 1 pm and 20 pm, preferably between 1 pm and 10 pm.

[0086] The bonding layer 18 comprises a metallic material. For example, it comprises a metal selected from chromium, titanium, nickel, aluminum, zirconium, molybdenum, magnesium, copper, tin, or an alloy thereof. The alloy may be a binary alloy comprising two metals selected from these, a ternary alloy comprising three metals selected from these, or a multi-component alloy of the high-entropy alloy type comprising at least five metals selected from these.

[0087] The bonding layer 18 and the body 16 exhibit an adhesion strength between 10 MPa and 80 MPa.

[0088] The bonding layer 18 helps to promote the adhesion of the functional layer 19 to the body 16.

[0089] The bonding layer 18 is deposited by high-power impulse magnetron sputtering.

[0090] Magnetron sputtering deposition in high-power pulse regime helps to protect body 16 from high temperatures and impacts that could degrade the composite material of body 16.

[0091] According to an advantageous embodiment illustrated in Figure 3, the bonding layer 18 is multilayered. It comprises a first bonding sublayer 18a and a second bonding sublayer 18b.

[0092] The first bonding subshell 18a is located between the body 16 and the second bonding subshell 18b.

[0093] The second bonding sublayer 18b is located between the first bonding sublayer 18a and the functional layer 19. The second bonding sublayer 18b has a thickness between 5 pm and 15 pm, in particular between 5 pm and 9 pm.

[0094] According to an advantageous embodiment illustrated in Figure 5, the bonding layer 18 may further comprise at least n additional bonding sublayers 18n, n being greater than or equal to 1. The additional bonding sublayer(s) 18n are located between the second bonding sublayer 18b and the functional layer 19.

[0095] The second bonding sublayer 18b and / or the additional bonding sublayer 18n help to limit the spalling and / or erosion of the first sublayer 18a during the deposition of the functional layer 19. They also help to reduce the residual stresses generated by the difference in coefficients of thermal expansion between the bonding layer 18 and the functional layer 19. They also help to dissipate the kinetic energy of the particles projected during the deposition of the functional layer 19, which the first sublayer 18a would not be able to absorb alone.

[0096] The functional layer 19 has a thickness e2 between 30 pm and 2000 pm.

[0097] The functional layer 19 comprises a material chosen from the list presented in Table 1.

[0098] [Table 1] Functional layer material 19

[0099] The functional layer 19 is deposited by thermal spraying. Thermal spraying can be atmospheric pressure plasma spraying, also known by the English acronym APS "Atmospheric Plasma Spraying" or suspension plasma spraying, also known by the English acronym SPS for "Suspension Plasma Spraying", or solution plasma spraying, also known by the English acronym SPPS for "Solution Precursor Plasma Spraying", or high-speed flame spraying in powder and liquid modes, also known by the English acronym HVOF for "High Velocity Oxygen Fuel" or high-speed flame spraying in liquid mode, also known by the English acronym HVSFS for "High Velocity Suspension Flame Spray", or by cold spraying, also known by the English acronym CS for "cold-spray", or by wire arc or by the HVAF process for "High Velocity Air Fuel" or by a combination of these processes.

[0100] Preferably, the functional layer 19 comprises a WCCo material deposited by HVOF and the bonding layer 18 advantageously comprises titanium deposited by HIPIMS.

[0101] The functional layer 19 is advantageously multilayered. With reference to Figure 4, the functional layer 19 advantageously comprises a first functional sublayer 19a and a second functional sublayer 19.

[0102] The first functional sublayer 19a is located between the bonding layer 18 and the second functional sublayer 19b. Advantageously, the first functional sublayer 19a has a thickness between 40 pm and 120 pm.

[0103] The first functional sublayer 19a limits spalling and / or erosion of the bonding layer 18 caused by the impact of particles projected during thermal spray deposition, thanks to the ductile properties of this first functional sublayer 19a. It also reduces residual stresses generated by the difference in coefficient of thermal expansion between the bonding layer 18 and the functional layer 19.

[0104] According to an advantageous embodiment illustrated in Figure 5, the functional layer 19 may further comprise at least n additional functional sublayers 98n, n being greater than or equal to 1. The second functional sublayer 19b is located between the additional functional sublayer(s) 19n and the first functional sublayer 19a. A manufacturing method for the part 15 according to the invention will now be described with reference to Figure 6.

[0105] The process includes the following chronological steps:

[0106] (a) provide body 16,

[0107] (b) optionally, perform a surface treatment on body 16, and

[0108] (c) place the coating 17 onto the body 16.

[0109] Step (a) may include the following substeps:

[0110] (aO) produce a three-dimensional fibrous preform of body 16, and

[0111] (a1) impregnate the fibrous preform with a resin.

[0112] Step (aO) can be performed by weaving.

[0113] Step (a1) can be carried out by molding such as resin transfer molding known by the acronym RTM for "Resin Transfer Molding" in English or draping.

[0114] Step (b) of surface treatment of the body can be carried out by mechanical treatment for example of the type sandblasting or sanding, or by laser treatment, or by chemical treatment or even by plasma treatment.

[0115] At the end of this step (b), the surface of body 16 has a roughness Ra between 0.5 pm and 8 pm, preferably between 0.5 pm and 5 pm.

[0116] According to the invention, step (c) of coating deposition 17 comprises the following sub-steps:

[0117] (d) deposit the bonding layer 18 onto the body 16 by magnetron sputtering in high-power pulse mode, and

[0118] (e) deposit the functional layer 19 onto the bonding layer 18 by thermal spraying.

[0119] According to one embodiment, step (d) comprises the following substeps:

[0120] (d1) deposit the first bonding sublayer 18a onto the body 16 by magnetron sputtering in high-power pulse mode, (d2) deposit the second bonding sublayer 18b onto the first bonding sublayer by magnetron sputtering in high-power pulse mode, and

[0121] (d3) Optionally, deposit the additional bonding subshell 18n onto the second bonding subshell 18b by magnetron sputtering in high-power pulse regime.

[0122] High-power impulse magnetron sputtering is also known by the English acronym HIPIMS.

[0123] The high-power pulsed magnetron sputtering process comprises the following sub-steps:

[0124] - to introduce a gas into a vacuum chamber,

[0125] - ionized the gas around a target,

[0126] - to impact the target with ions from the ionized gas to detach atoms from the target,

[0127] - deposit the atoms of the target onto the body.

[0128] During high-power pulse magnetron sputtering, atoms of the bonding layer material 18 are vaporized by high-power pulse magnetron sputtering, using gas.

[0129] The gas includes, for example, argon, nitrogen, oxygen, carbon dioxide, and / or noble gases. The gas is ionizable.

[0130] The gas is ionized in pulsed mode by applying high-power pulses to a magnetron cathode.

[0131] The pulses consist of alternating positive and negative voltages. The positive voltage ranges from 0 V to 50 V, and the negative voltage ranges from -2000 V to 0 V. Each pulse has a duration between 1 ps and 500 ps.

[0132] The pulses, for example, have a frequency between 100 Hz and 10 kHz. Such a magnetron sputtering process in high-power pulse mode allows the deposition of the bond layer 18 without degradation of the body 16. Indeed, this process uses short current pulses, which allows the use of high power over short periods, thus resulting in very high ionization of the body 16 to be sputtered without overheating it.

[0133] The ionized particles reach the body 16 with high energy. This results in denser bonding sublayers 18a, 18b, 18n due to the higher mobility of the ionized species. Furthermore, bombarding the body 16 with energetic ions can lead to ion implantation that promotes chemical adhesion and to surface morphology modifications that promote mechanical anchoring, thus improving the adhesion of the deposited bonding layer 18. In addition, the discharge in this process also produces hot electrons at the beginning of the pulse, which cause surface modifications once absorbed by the body 16, such as local heating that also contributes to the adhesion of the coating 17. In one embodiment, step (e) comprises the following substeps:

[0134] (e1) deposit the first functional sublayer 19a onto the bonding layer 18 by thermal spraying,

[0135] (e2) deposit the second functional sublayer 19b onto the first functional sublayer 19a by thermal spraying, and

[0136] (e3) Optionally, deposit the additional functional sublayer 19n onto the second functional sublayer 19b by thermal spraying.

[0137] Thermal spraying is carried out with a powder of the material of the functional layer 19. The powder comprises particles of which at least 90% of these particles have a size (d90) between 5 pm and 200 pm, advantageously between 10 pm and 150 pm and preferably of which at most 10% of the particles have a size (d10) between 0.05 pm and 50 pm, advantageously between 0.1 pm and 20 pm.

[0138] The resulting coating 17 exhibits a dense nanostructured microstructure without defects and with strong adhesion to the body 16.

[0139] With reference to figure 6, the manufacturing process of part 15 can be implemented in a part 15 manufacturing installation 100. The installation 100 includes a first station 200 for manufacturing the body 16 and a second station 300 for manufacturing the coating 17.

[0140] The first station 200 includes, for example, a weaving device 202 and a molding device 204.

[0141] The second station 300 includes a high-power pulse magnetron sputtering device 302 and a thermal spraying device 304.

[0142] With reference to Figure 8, the high-power pulsed magnetron sputtering device 302 comprises a vacuum chamber 305, a gas inlet 306 and a gas outlet 307.

[0143] The high-power pulse magnetron sputtering device 302 further includes at least one magnetron cathode 308 disposed in the enclosure 305, a sputtering target 309 disposed in the enclosure 305 and a frame 310 for fixing the body 16 in the enclosure 305.

[0144] The gas is introduced into chamber 305 and ionized. The magnetic field in the chamber accelerates and confines the I ions of the gas around the sputtering target 309.

[0145] The impact of the ions on the sputtering target 309 generates an energy transfer and causes the detachment of atoms A from the sputtering target 309 which are returned to the body 16 thus forming the coating 17.

[0146] [Examples]

[0147] Preferred examples of the invention will now be described. [Example 1]

[0148] According to a first example, coating 17 is a shock-resistant and erosion-resistant coating.

[0149] Part 15 is the awl 11. According to this first example, body 16 is the blade 12.

[0150] According to this first example, body 16 has a roughness Ra of 3.5 pm after corundum blasting.

[0151] The material of the bonding layer 18 and the functional layer 19, the thickness of each of the layers 18, 19 and the deposition process of each of the layers 18, 19 is presented in Table 2.

[0152] [Table 2] shock and erosion resistant coating

[0153] [Example 2]

[0154] According to a second example, coating 17 is an anti-wear coating.

[0155] Part 15 is the blade 11 or the hood 1b. According to this first example, the body 16 is the foot 13 or the panel 1c.

[0156] According to this second example, body 16 has a roughness Ra of 3.5 pm after corundum blasting.

[0157] The coating material 17 and the coating deposition process are shown in Table 3.

[0158] [Table 3] wear-resistant coating

[0159] [Example 3]

[0160] According to a third example, coating 17 is a lightning protection coating.

[0161] Part 15 is, for example, the hood 1b. According to this first example, the body 16 is, for example, the panel 1c.

[0162] According to this second example, body 16 has a roughness Ra of 3.5 pm after corundum blasting.

[0163] The coating material 17 and the coating deposition process are shown in Table 4. [Table 4] Lightning protection coating

[0164] [Example 4]

[0165] According to a fourth example, coating 17 is an anti-icing coating. Part 15 is blade 11. According to this first example, body 16 is blade 12.

[0166] According to this second example, body 16 has a roughness Ra of 3.5 pm after corundum blasting.

[0167] The coating material 17 and the coating deposition process are shown in Table 5.

[0168] [Table 5] anti-frost coating

Claims

DEMANDS 1. Method for manufacturing a part (15, 11, 1b) for an aircraft propulsion assembly (1'), the method comprising the following chronological steps: (a) providing a body (16, 12, 1c) of composite material comprising fibers embedded in a polymer matrix, (c) depositing a coating (17, 14) on the body (16, 12, 1c), characterized in that step (c) comprises the following substeps: (d) depositing a bonding layer (18) onto the body (16, 12, 1c) by magnetron sputtering in high-power pulse mode, the bonding layer (18) having a thickness between 1 pm and 20 pm, and (e) deposit a functional layer (19) on the bonding layer (18) by thermal spraying.

2. Method according to the preceding claim, characterized in that it comprises between steps (a) and (c), a step (b) of surface treatment of the body (16, 12, 1c).

3. Method according to the preceding claim, characterized in that at the end of step (b), the body (16, 12, 1c) has a surface having a roughness Ra between 0.5 pm and 8 pm, preferably between 0.5 pm and 5 pm.

4. A method according to any one of the preceding claims, characterized in that the bonding layer (18) comprises a metallic material.

5. A method according to any one of the preceding claims, characterized in that the bonding layer (18) has a thickness between 1 pm and 10 pm.

6. A method according to any one of the preceding claims, characterized in that step (d) comprises the following substeps: (d1) deposit a first bonding sublayer (18a) onto the body (16, 12, 1c) by magnetron sputtering in high-power pulse mode, and (d2) deposit a second bonding subshell (18b) on the first bonding subshell (18a) by magnetron sputtering in high-power pulse regime.

7. Method according to the preceding claim, characterized in that the second bonding sublayer (18b) has a thickness between 5 pm and 15 pm.

8. A method according to any one of the preceding claims, characterized in that the functional layer (19) has a thickness between 30 pm and 2000 pm.

9. A method according to any one of the preceding claims, characterized in that step (e) comprises the following substeps: (e1) deposit a first functional sublayer (19a) onto the bonding layer (18) by thermal spraying, and (e2) deposit a second functional sublayer (19b) on the first functional sublayer (19a) by thermal spraying.

10. Method according to the preceding claim, characterized in that the first functional sublayer (19a) has a thickness between 40 pm and 120 pm.

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