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 sequential steps: (a) providing a body (16) made of composite material comprising fibers embedded in a polymer matrix, (c) depositing a coating (17) onto the body (16), characterized in that step (c) comprises the following substeps: (d) depositing a bonding layer (18) onto the body (16) by magnetron sputtering in a high-power pulse regime, and (e) depositing a functional layer (19) onto the bonding layer (18) by thermal spraying. (Shorthand figure: Figure 2)

Description

Title of the invention: METHOD FOR MANUFACTURING A PART FOR A PROPULSION ASSEMBLY Technical field of the invention

[0001] 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. Technical background

[0002] A propulsion assembly 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.

[0003] The blower allows the intake of an airflow that splits into a primary airflow and a secondary airflow. The primary airflow passes through a primary channel of the turbomachine, while the secondary airflow is directed towards a secondary channel surrounding the primary channel.

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

[0005] The propulsion assembly comprises parts having a body made of composite material and a coating arranged on the body and having a functional layer.

[0006] For example, the turbomachine includes blades that allow an action to be exerted on the airflow. For example, compressor blades allow the primary airflow to be compressed and blower blades allow the secondary airflow to be compressed.

[0007] A blade comprises a body, such as a blade, which has 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 weight of the blade, the blade is made of a composite material. The composite material of the blade comprises reinforcing fibers embedded in a polymer matrix, in particular an epoxy matrix. In order to protect the 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. forming a shock-resistant and wear-resistant coating. The coating includes a functional shock-resistant and wear-resistant layer, usually metallic.

[0008] The coating is typically glued to the blade.

[0009] A method for manufacturing the aforementioned blade comprises the following steps:

[0010] - manufacturing of the blade,

[0011] - manufacturing of the coating,

[0012] - matching of the coating to the blade, and

[0013] - bonding of the coating to the blade.

[0014] Such a manufacturing process does not provide complete satisfaction. Indeed, the coating manufacturing step presents challenges because the coating configuration is complex. This step is therefore tedious. Furthermore, manufacturing tolerances are tight, and it is difficult to ensure the repeatability of this step.

[0015] Furthermore, the coating matching step, which consists of attaching the coating to the blade, is carried out manually, so that guaranteeing the precision of the blade assembly and the repeatability of this step requires perfect control of the manufacturing process to limit defects on the coating or the blade and therefore the scrap of these parts.

[0016] In this context, thermal spray coating was considered. This process eliminates the need for coating manufacturing and matching steps. However, since the blade is made of a composite material based on an epoxy matrix, it exhibits low wettability and lacks functional groups to ensure adhesion of the thermally sprayed coating. Therefore, a significant limitation in implementing such a thermally sprayed coating is the sensitivity of the blade's carbon fibers to the impact of particles projected during the thermal spraying process.

[0017] These issues concern all aircraft parts, in particular turbomachinery and nacelles, which include a body made of composite material, which it is necessary to equip with a coating having a functional layer, of the anti-shock, anti-wear, anti-icing or even anti-lightning type.

[0018] Therefore, there is a need to provide a method for manufacturing a part for an aircraft comprising a body made of composite material and a coating, which is compatible with the deposition of a functional layer by thermal spraying. Summary of the invention

[0019] To this end, the invention proposes a method for manufacturing a part for an aircraft propulsion assembly, the method comprising the following chronological steps:

[0020] (a) provide a body made of composite material comprising fibers embedded in a polymer matrix,

[0021] (c) deposit a coating on the body.

[0022] The process is remarkable in that step (c) comprises the following substeps:

[0023] (d) depositing a bonding layer on the body by sputtering magnetron in high-power pulse mode, and

[0024] (e) depositing a functional layer onto the bonding layer by projection thermal.

[0025] 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.

[0026] This functional layer is deposited by thermal spraying. The thermal spraying process makes it possible to form coatings with complex geometries precisely and repeatably, thus reducing the risk of coating defects and its disposal.

[0027] 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 on the body, thus simplifying the part assembly process.

[0028] 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.

[0029] The bonding layer promotes adhesion of the functional layer to the body and protects the body during thermal spray deposition of the functional layer. Thanks to the bonding layer, this thermal spraying can be carried out without damaging the body.

[0030] 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".

[0031] Magnetron sputtering deposition in high-power pulse regime makes it possible to protect the body from high temperature and impacts which could degrade the composite material of the body.

[0032] The invention may comprise one or more of the following features, taken individually or in combination with each other:

[0033] — the functional layer comprises a WCCo material deposited by HVOF and the The bonding layer advantageously comprises titanium deposited by HIPIMS.

[0034] - between steps (a) and (c), a step (b) of surface treatment of the body,

[0035] - 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,

[0036] - the bonding layer comprises a metallic material,

[0037] - the bonding layer has a thickness between 1 pm and 20 pm, of Preferably between 1 pm and 10 pm,

[0038] - step (d) comprises the following substeps:

[0039] (dl) depositing a first bonding sublayer on the body by magnetron sputtering in high-power pulse mode, and

[0040] (d2) deposit a second bonding sublayer on the first sublayer of magnetron sputtering bonding in high-power pulse regime,

[0041] - the second bonding sublayer has a thickness between 5 pm and 15 pm,

[0042] - the functional layer has a thickness between 30 pm and 2000 pm,

[0043] - step (e) comprises the following substeps:

[0044] (el) depositing a first functional sub-layer onto the bonding layer by thermal spraying, and

[0045] (e2) deposit a second functional sublayer on the first sublayer functional by thermal projection

[0046] - the first functional sublayer has a thickness of between 40 pm and 120 pm. Brief description of the figures

[0047] 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:

[0048] [Fig. 1] is a schematic longitudinal cross-sectional view of a propulsion assembly to which the invention can be applied,

[0049] [Fig.2] is a schematic cross-sectional view of a part according to the invention and equipping the propulsion assembly of [Fig.1],

[0050] [Fig. 3] is a schematic cross-sectional view of a part according to an embodiment of the invention,

[0051] [Fig.4] is a schematic cross-sectional view of a part according to an embodiment of the invention,

[0052] [Fig. 5] is a schematic cross-sectional view of a part according to an embodiment of the invention,

[0053] Figure 6 is a schematic representation of the manufacturing process according to the invention,

[0054] [Fig.7] is a schematic representation of an installation that can be implemented during the manufacturing process of the invention,

[0055] [Fig.8] is a schematic representation of a device that can be implemented in step (c) of the manufacturing process according to the invention. Detailed description of the invention

[0056] An example of an aircraft propulsion assembly 1' is schematically illustrated in [Fig.1].

[0057] The propulsion assembly 1' extends around and along a longitudinal axis A.

[0058] In the present application, the terms “axial”, “axially”, “radial” and “ radially » are defined with respect to the longitudinal axis A.

[0059] The terms "upstream", "downstream" are defined with respect to the direction of gas flow in the propulsion assembly 1' along the longitudinal axis A.

[0060] The terms "internal", "interior", "external", "outside", "externally" are defined with respect to the distance from the longitudinal axis A along a radial axis perpendicular to the longitudinal axis A.

[0061] The propulsion assembly 1' comprises a turbomachine 1 and optionally a nacelle surrounding the turbomachine 1.

[0062] 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).

[0063] The rotor of the low-pressure turbine is connected to the blower 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).

[0064] The turbomachine 1 also includes a rectifier 10. The rectifier 10 allows the flow at the outlet of an upstream rotor to be rectified in order to provide maximum thrust at the outlet of the turbomachine 1. In the particular example of [Fig.1], the rectifier 10 is located downstream of the fan 2 and allows the secondary air flow F2 to be rectified.

[0065] The blower 2 allows the intake of an airflow that splits into a primary airflow Fl and a secondary airflow F2. The primary airflow Fl passes through a vein primary of turbomachine 1 while the secondary airflow F2 is directed towards a secondary vein surrounding the primary vein.

[0066] The primary airflow Fl 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 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 these gases to be accelerated to generate propulsion. The secondary airflow F2 passes through the rectifier 10, which accelerates the circulation speed of the secondary airflow F2 to generate propulsion.

[0067] The blower 2 and / or the straightener 10 each comprise blades 11 which are carried 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.

[0068] The blades 11 extend radially with respect to the longitudinal axis A and are regularly distributed around the longitudinal axis A.

[0069] Each blade 11 comprises a blade 12 and a foot 13.

[0070] 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.

[0071] The 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 in the matrix.

[0072] The matrix is ​​a polymeric matrix. The matrix polymer comprises a thermoplastic or thermosetting polymer. The thermosetting polymer is, for example, an epoxy polymer.

[0073] The fibrous reinforcement comprises fibers such as carbon, polyester, polypropylene, aramid, or glass fibers. The fibers may be short or long. Preferably, the fibers are woven. The fibers are woven according to a three-dimensional fibrous preform.

[0074] In a particularly preferred manner, the composite material comprises an epoxy matrix and carbon fibers embedded in the epoxy matrix.

[0075] 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.

[0076] Each vane 11 may include a protective shield 14 which is advantageously arranged on the leading edge 12a. It advantageously extends along the entire length of the leading edge 12a.

[0077] 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 between them in which the leading edge 12a is arranged.

[0078] 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. The free longitudinal ends extend respectively over the lower surface 12i and the upper surface of the blade 12.

[0079] Alternatively, and 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 on the trailing edge 12b.

[0080] The nacelle la is annular and centered on the longitudinal axis A. It extends around the turbomachine 1. The nacelle la comprises at least one cowling 1b including a panel le and a coating arranged on the panel le. The panel le comprises a composite material which may be identical to the composite material of the blade 12. The coating may be a lightning-resistant coating, for example.

[0081] With reference to figures 2 to 5, the propulsion assembly 1' comprises at least one part 15. The part 15 is, according to a first example, the blade 11 or, according to a second example, the hood 1 b of the nacelle la.

[0082] The part 15 includes a body 16 comprising a composite material and a coating 17 arranged on the body 16.

[0083] The composite material is, for example, an organic matrix composite (OMC). The composite material comprises a polymer matrix and a fibrous reinforcement embedded in the matrix.

[0084] The matrix is ​​a polymeric matrix. The matrix polymer comprises a thermoplastic or thermosetting polymer. The thermosetting polymer is, for example, an epoxy polymer.

[0085] The fibrous reinforcement comprises fibers such as carbon, polyester, polypropylene, aramid, or glass fibers. The fibers may be short or long. Preferably, the fibers are woven. The fibers are woven according to a three-dimensional fibrous preform.

[0086] In a particularly preferred manner, the composite material comprises an epoxy matrix and carbon fibers embedded in the epoxy matrix.

[0087] The body 16 comprises a surface having a roughness Ra of between 0.5 pm and 8 pm, preferably between 0.5 pm and 5 pm. Such roughness promotes the adhesion of the coating 17.

[0088] The body 16 is, according to the first example, the blade 12 or the foot 12 or, according to the second example, the panel le of the hood 1b of the nacelle la.

[0089] Coating 17 is for example an anti-shock and / or anti-erosion, anti-wear, anti-lightning or even anti-frost coating.

[0090] 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 le of the hood 1b of the nacelle la.

[0091] The coating 17 comprises a bonding layer 18 and a functional layer 19.

[0092] The bonding layer 18 is located between the body 16 and the functional layer 19. The bonding layer 18 has a thickness el between 1 pm and 20 pm, preferably between 1 pm and 10 pm.

[0093] 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.

[0094] The bonding layer 18 and the body 16 have an adhesion force between 10 MPa and 80 MPa.

[0095] The bonding layer 18 helps to promote the adhesion of the functional layer 19 to the body 16.

[0096] The bonding layer 18 is deposited by magnetron sputtering in high-power impulse mode, also known by the English acronym HIPIMS for "high-power impulse magnetron sputtering".

[0097] Magnetron sputtering deposition in high-power pulse regime makes it possible to protect the body 16 from high temperatures and impacts which could degrade the composite material of the body 16.

[0098] According to an advantageous embodiment illustrated in [Fig.3], the bonding layer 18 is multilayered. It comprises a first bonding sublayer 18a and a second bonding sublayer 18b.

[0099] The first bonding subshell 18a is located between the body 16 and the second bonding subshell 18b.

[0100] 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.

[0101] According to an advantageous embodiment illustrated in [Fig.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.

[0102] The second bonding sublayer 18b and / or the additional bonding sublayer 18n limit the spalling and / or erosion of the first sublayer 18a during the deposition of the functional layer 19. They also 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 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.

[0103] The functional layer 19 has a thickness e2 between 30 pm and 2000 pm.

[0104] The functional layer 19 comprises a material selected from the list presented in Table 1.

[0105] [Table 1] functional layer material 19 Functional layer material 19 WC-Co Cr3C2-NiCr CrO2 SiCLAhCL Al2O3-TiO2 Ni-SiC Cr Mo Co. Co

[0106] 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-velocity flame spraying in powder and liquid stages, also known by the English acronym HVOF for "High Velocity Oxygen Fuel" or high-velocity flame spraying in liquid stages, 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 arc wire or by the HVAF process for "High Velocity Air Fuel" or by a combination of these processes.

[0107] Preferably, the functional layer 19 comprises a WCCo material deposited by HVOF and the bonding layer 18 advantageously comprises titanium deposited by HIPIMS.

[0108] The functional layer 19 is advantageously multilayered. With reference to [Fig.4], the functional layer 19 advantageously comprises a first functional sublayer 19a and a second functional sublayer 19.

[0109] 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.

[0110] The first functional sublayer 19a limits the spalling and / or erosion of the bonding layer 18 caused by the impact of the particles projected during thermal spray deposition, thanks to the ductile properties of this first functional sublayer 19a. It also reduces the residual stresses generated by the difference in coefficient of thermal expansion between the bonding layer 18 and the functional layer 19.

[0111] According to an advantageous embodiment illustrated in [Fig.5], the functional layer 19 can 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.

[0112] A method for manufacturing part 15 according to the invention will now be described with reference to [Fig.6].

[0113] The process comprises the following chronological steps:

[0114] (a) provide body 16,

[0115] (b) optionally, perform a surface treatment on body 16, and

[0116] (c) deposit the coating 17 onto the body 16.

[0117] Step (a) may include the following substeps:

[0118] (aO) produce a three-dimensional fibrous preform of body 16, and

[0119] (al) impregnate the fibrous preform with a resin.

[0120] Step (aO) can be carried out by weaving.

[0121] Step (al) can be carried out by molding such as resin transfer molding known by the acronym RTM for "Resin Transfer Molding" in English or draping.

[0122] Step (b) of surface treatment of the body can be carried out by mechanical treatment for example of the sandblasting or sanding type, or by laser treatment, or by chemical treatment or even by plasma treatment.

[0123] At the end of this step (b), the surface of the body 16 has a roughness Ra between 0.5 pm and 8 pm, preferably between 0.5 pm and 5 pm.

[0124] According to the invention, step (c) of coating deposition 17 comprises the following substeps:

[0125] (d) deposit the bonding layer 18 onto the body 16 by sputtering magnetron in high-power pulse mode, and

[0126] (e) deposit the functional layer 19 onto the bonding layer 18 by projection thermal.

[0127] According to one embodiment, step (d) comprises the following substeps:

[0128] (dl) deposit the first bonding undercoat 18a onto the body 16 by spraying cathode ray magnetron in high-power pulse mode,

[0129] (d2) deposit the second 18b bonding subshell onto the first subshell magnetron sputtering bonding in high-power pulse regime, and

[0130] (d3) optionally, deposit the additional 18n bonding sublayer on the second bonding sublayer 18b by magnetron sputtering in high power pulse regime.

[0131] High-power impulse magnetron sputtering is also known by the English acronym HIPIMS for High-power impulse magnetron sputtering.

[0132] The magnetron sputtering process in high-power pulse mode comprises the following substeps:

[0133] - introducing a gas into a vacuum chamber,

[0134] - ionized the gas around a target,

[0135] - to impact the target with ions from the ionized gas in order to detach atoms from the target,

[0136] - deposit the atoms of the target onto the body.

[0137] During magnetron sputtering in high pulse regime power, atoms of the bonding layer material 18 are vaporized by magnetron sputtering in high power pulse regime, using gas.

[0138] The gas comprises, for example, argon, nitrogen, oxygen, carbon dioxide and / or noble gas. The gas is ionizable.

[0139] The gas is ionized in pulsed regime by applying high-power pulses to a magnetron cathode.

[0140] The pulses comprise an alternation of positive and negative voltage. The positive voltage is between 0 V and 50 V and the negative voltage is between -2000 V and 0 V. Each pulse has a duration between 1 ps and 500 ps.

[0141] The pulses have, for example, a frequency between 100 Hz and 10 kHz.

[0142] Such a magnetron sputtering process in high-power pulse regime 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.

[0143] 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 implantations that promote chemical adhesion and to surface morphology modifications that promote mechanical anchoring, thereby 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, which also contributes to the adhesion of the coating 17.

[0144] According to one embodiment, step (e) comprises the following substeps:

[0145] (el) deposit the first functional sublayer 19a on the bonding layer 18 by thermal projection,

[0146] (e2) deposit the second functional sublayer 19b onto the first sublayer functional 19a by thermal projection, and

[0147] (e3) optionally, deposit the additional functional sublayer 19n on the second functional undercoat 19b by thermal spraying.

[0148] 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 (dl0) between 0.05 pm and 50 pm, advantageously between 0.1 pm and 20 pm.

[0149] The coating 17 obtained has a dense nanostructured microstructure without defects and with strong adhesion to the body 16.

[0150] With reference to [Fig.6], the manufacturing process for part 15 can be implemented in a part 15 manufacturing installation 100. The installation 100 comprises a first station 200 for manufacturing the body 16 and a second station 300 for manufacturing the coating 17.

[0151] The first station 200 includes, for example, a weaving device 202 and a molding device 204.

[0152] The second station 300 includes a high-power pulse magnetron sputtering device 302 and a thermal spraying device 304.

[0153] With reference to [Fig.8], the high-power pulsed magnetron sputtering device 302 comprises a vacuum enclosure 305, a gas inlet 306 and a gas outlet 307.

[0154] 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.

[0155] The gas is introduced into the chamber 305 and ionized. The magnetic field in the chamber accelerates and confines the I ions of the gas around the sputtering target 309.

[0156] 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. Examples

[0157] Preferred examples of the invention will now be described.

[0158] [Example 1]

[0159] According to a first example, the coating 17 is an anti-shock and anti-erosion coating.

[0160] Part 15 is the awl 11. According to this first example, the body 16 is the blade 12.

[0161] According to this first example, the body 16 has a roughness Ra of 3.5 pm after corundum sandblasting.

[0162] 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.

[0163] [Table 2] shock and erosion resistant coating Bonding layer 1 8 or first bonding sublayer 1 8a Material Process Thickness (pm) Ti HIPIMS 10 Working layer!! e 19 or first working sublayer 19a WC-Co HVOF 1000 Cr3C2-NiCr HVOF CrO2 APS SiO2Al2O3 APS Al2O3-TiO2 APS Ni-SiC HVOF Cr HVOF Mo HVOF Co HVOF Ni HVOF NiCr HVOF Ni-SiC HVOF MoAINi HVOF NiCrW HVOF CoCrW HVOF FeCrNbBWC HVOF FeCrC HVOF Second functional underlay 19b optional Al Arc wire 80 Second connecting underlay 18b opt ional Al HIPIMS 15

[0164] [Example 2]

[0165] According to a second example, the coating 17 is an anti-wear coating.

[0166] 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.

[0167] According to this second example, body 16 has a roughness Ra of 3.5 pm after corundum sandblasting.

[0168] The coating material 17 and the coating deposition process are shown in Table 3.

[0169] [Table 3] wear-resistant coating Bonding layer 1 8 or first bonding sublayer 1 8a Material Process Thickness (pm) Ti HIPIMS 10 Diaper works!! e 19 or first functional sub-layer 19a WC-Co HVOF 1000 Cr3C2-NiCr HVOF CrO2 APS A12O3 APS SiO2Al2O3 APS Al2O3-TiO2 APS Ni-SiC HVOF Cr HVOF Mo HVOF Co HVOF Ni HVOF NiCr HVOF Ni-SiC HVOF MoAINi HVOF FeCrC HVOF NiCrW HVOF CoCrW HVOF FeCrNbBWC HVOF Optional second functional sub-layer 19b Al Arc wire 80 Optional second connecting sub-layer 18b Al HIPIMS 15

[0170] [Example 3]

[0171] According to a third example, coating 17 is a lightning-resistant coating.

[0172] Part 15 is, for example, the cover 1b. According to this first example, the body 16 is for example the sign.

[0173] According to this second example, body 16 has a roughness Ra of 3.5 pm after corundum sandblasting.

[0174] The coating material 17 and the coating deposition process are shown in Table 4.

[0175] [Table 4] lightning protection coating Bonding layer 18 or first bonding sublayer 18a Material Process Thickness (µm) Cu HIPIMS 10 Functional layer 19 or first functional sublayer 19a Cu Cold spray 50 CuSn Cold spray Sb Cold spray Ag Cold spray Au Cold spray Zn Cold spray Al Cold spray Second functional sublayer 19b optional Cu Wire arc 80 Second bonding sublayer 18b optional Al HIPIMS 15

[0176] [Example 4]

[0177] According to a fourth example, coating 17 is an anti-frost coating.

[0178] Part 15 is the awl 11. According to this first example, the body 16 is the blade 12.

[0179] According to this second example, body 16 has a roughness Ra of 3.5 pm after corundum blasting.

[0180] The coating material 17 and the coating deposition process are shown in Table 5.

[0181] [Table 5] anti-frost coating Bonding layer 18 or first bonding sublayer 18a Material Process Thickness (pm) Ti HIPIMS 10 Functional layer 19 or first functional sublayer 19a TiO2 APS / SPS 200 Al2O3-TiO2 APS / SPS Yb2O3 APS / SPS CuZn3 Cold spray AuZn3 Cold Spray SiO2 APS / SPS PPS-PTFE / SiO2 APS / SPS Second functional sub-coat 19b optional Al Wire Arc 80 Second bonding sub-coat 18b optional Al HIPIMS 15

Claims

Demands

1. A 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, the) of composite material comprising fibers embedded in a polymer matrix, (c) depositing a coating (17, 14) on the body (16, 12, the), characterized in that step (c) comprises the following substeps: (d) depositing a bonding layer (18) on the body (16, 12, the) by magnetron sputtering in high-power pulse mode, and (e) depositing 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, the).

3. A method according to the preceding claim, characterized in that at the end of step (b), the body (16, 12, le) 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 of between 1 pm and 20 pm, preferably 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) depositing a first bonding sublayer (18a) on the body (16, 12, 1e) by magnetron sputtering in high-power pulse mode, and (d2) depositing a second bonding sublayer (18b) on the first bonding sublayer (18a) by magnetron sputtering in high-power pulse mode.

7. A 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) depositing a first functional sublayer (19a) onto the bonding layer (18) by thermal spraying, and (e2) depositing a second functional sublayer (19b) onto the first functional sublayer (19a) by thermal spraying.

10. A 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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