Method for producing a part made of composite ceramic material, with geometry correction

Abstract

The invention relates to a method for producing a part made of ceramic matrix composite material, the method comprising at least: - a step of forming (10) a fibrous preform; - a step of consolidating (20) the fibrous preform; - an injection step (30), during which a slurry comprising a powder of silicon carbide particles is injected into the fibrous preform, so as to obtain a raw part (100); and - an infiltration step (60), during which a molten silicon-based composition is infiltrated, so as to form a part made of ceramic matrix composite material. The method further comprises: - a step of depositing (40) at least one finishing layer (200) over all or part of the raw part (100); and - a thermocompression step (50), during which the finishing layer (200) is thermocompressed according to a final geometry of the part made of ceramic matrix composite material.

Description

[0001] Description

[0002] METHOD FOR MANUFACTURING A PART MADE OF CERAMIC COMPOSITE MATERIAL WITH GEOMETRY CORRECTION

[0003] Technical Field

[0004] The present invention relates to parts made of ceramic matrix composite material. It relates more particularly to improving the control of the final geometry and surface condition of such parts.

[0005] Prior art

[0006] One field of application of the invention is the production of parts made of ceramic matrix composite material, also known by the acronym "CMC". The ceramic matrix composite material replaces parts made of metallic material, particularly in certain parts of aeronautical engines and in particular in the gas turbines of such engines.

[0007] In the aeronautics field, the use of ceramic matrix composite materials contributes to optimizing aircraft performance, in particular by improving the efficiency of the turbomachine and reducing the overall mass of the turbomachine, significantly reducing harmful emissions to the environment (CO, CO2, NOx, etc.).

[0008] Furthermore, the use of a ceramic matrix composite material instead of metallic materials for parts exposed in service to high temperatures has therefore been recommended, especially since a ceramic matrix composite material has a significantly lower density than the metallic materials for which it is substituted.

[0009] A well-known method for manufacturing parts made of ceramic matrix composite material comprises the following steps: a step of producing a fibrous texture from carbon or silicon carbide SiC fibers, a step of consolidating a fibrous preform produced by chemical vapor infiltration, also referred to by the acronym "CVI" for "Chemical Vapor Infiltration" in English, of silicon carbide SiC, the fibrous texture being maintained in a shaping tool during the CVI, a step of injecting a slip of silicon carbide SiC powder into the fibrous preform, also referred to by the English terms "Slurry Cast" or "Slurry Transfer Molding", and a step of infiltrating the preform with a composition based on molten silicon, so as to form a ceramic matrix, a densification process known under the designation "Ml" process for the English term "Melt Infiltration".

[0010] Such a process is notably described in document FR3047988.

[0011] Although such a process makes it possible to obtain parts with very good mechanical properties, the resulting part exhibits deviations from a theoretical geometry as well as surface roughness which can reduce the performance, particularly aerodynamic, of the part.

[0012] One solution to improve the geometry and surface finish of a ceramic matrix composite part is to form a layer of the matrix material, in this case fiber-free silicon carbide SiC, on the surface of the part, which is machined to compensate for geometry deviations and smooth the surface. However, the thickness of the fiber-free layer must be limited to ensure good thermomechanical resistance.

[0013] In addition, machining operations lead to increased complexity and manufacturing costs.

[0014] However, there is a need to be able to manufacture parts in ceramic matrix composite material which, at the end of the manufacturing process, i.e. directly after the infiltration stage of the Ml preform, already respect the geometric and dimensional tolerances while having a good surface finish compatible with the required performance, in particular aerodynamic performance.

[0015] Statement of the invention

[0016] To this end, the invention proposes a method for manufacturing a part made of ceramic matrix composite material comprising at least: a forming step, during which a fiber preform is formed, a consolidation step, during which a consolidation phase of the fiber preform is carried out, an injection step, during which a slip comprising a powder of silicon carbide particles is injected into the fiber preform, so as to obtain a raw part, and an infiltration step, during which a composition based on molten silicon is infiltrated, so as to form a part made of ceramic matrix composite material.

[0017] In addition, the manufacturing method further comprises a deposition step, during which at least one finishing layer is deposited on all or part of the raw part, and a thermocompression step, during which the finishing layer is thermocompressed according to a final geometry of the part made of ceramic matrix composite material to be manufactured.

[0018] Removal and thermocompression of the finishing layer allow the local formation of excess thicknesses on at least part of the surface of the part in order to correct geometry deviations and obtain a final part as close as possible to the desired theoretical geometry.

[0019] The parts of the final part previously covered by the finishing layer also have a controlled surface condition (smoothness), particularly compatible with applications requiring aerodynamic performance.

[0020] Thanks to the presence of a plasticizer, the finishing layer has a flexibility allowing it to be conformed to different geometries, even complex ones, during the removal stage.

[0021] At the end of manufacturing, after the step of infiltrating the raw part with a molten silicon-based composition, it is no longer necessary to carry out a surface coating deposit and additional machining. Indeed, the geometry deviations have already been corrected and the surface condition is controlled.

[0022] According to another particular aspect of the manufacturing method of the invention, the deposition step and thermocompression step may take place after the slip injection step and before the step of infiltrating the raw part with a molten silicon-based composition. According to another particular aspect of the manufacturing method of the invention, during the deposition step, the finishing layer may comprise a silicon carbide powder dispersed in a support comprising at least one binder and / or one plasticizer.

[0023] According to a particular aspect of the manufacturing method of the invention, the finishing layer is produced by strip casting from a slip comprising at least one silicon carbide powder, a binder, a plasticizer, a dispersant and / or an organic solvent.

[0024] According to another particular aspect of the manufacturing method of the invention, the finishing layer may comprise a percentage of silicon carbide particles of between 30% and 60% by volume.

[0025] According to another particular aspect of the manufacturing method of the invention, the finishing layer may have a thickness less than or equal to 0.5 mm, in particular between 0.2 mm and 0.5 mm.

[0026] According to another particular aspect of the manufacturing method of the invention, during the thermocompression step, the raw part covered with the finishing layer is heated to a temperature higher than the creep temperature of the support of the finishing layer, the finishing layer being compacted on the raw part according to the final geometry of the part made of ceramic matrix composite material to be manufactured. This makes it possible to calibrate the thickness of the finishing layer on the external surface of the raw part.

[0027] The method of manufacturing a part made of ceramic matrix composite material of the invention can be applied in particular to the manufacture of a blade or a gas turbine distributor sector comprising an aerodynamic profile connected to at least one platform.

[0028] Furthermore, the invention also relates to a part made of ceramic matrix composite material obtained from the manufacturing method according to the invention which may be a blade or a gas turbine distributor sector comprising an aerodynamic profile connected to at least one platform.

[0029] In this case, at least one finishing layer is deposited on a surface of a leading edge of an aerodynamic profile and / or on a connection zone connecting an aerodynamic profile to a platform. The leading edge and the connection zone connecting an aerodynamic profile to a platform corresponding to zones regularly exhibiting geometry deviations, the manufacturing method of the invention can advantageously be used to correct such deviations.

[0030] Brief description of the drawings

[0031] The present invention will be better understood and other objects, characteristics and advantages will appear even more clearly on reading the detailed description which follows, comprising embodiments given by way of illustration with reference to the appended figures, presented as non-limiting examples, which may serve to complete the understanding of the present invention and the description of its embodiment and, where appropriate, contribute to its definition, in which:

[0032] [Fig. 1] Figure 1 is a flowchart indicating successive steps of an embodiment of a manufacturing method according to the invention,

[0033] [Fig. 2] Figure 2 is a schematic perspective view showing a deposition of a finishing layer on a raw part in accordance with one embodiment of the invention,

[0034] [Fig. 3] Figure 3 is another schematic perspective view showing a deposition of a finishing layer on a raw part in accordance with another embodiment of the invention,

[0035] [Fig. 4] Figure 4 is a schematic perspective view showing an installation of the raw part of Figures 2 and 3 in a thermocompression tool,

[0036] [Fig. 5] Figure 5 is a schematic perspective view showing the blank of Figures 2 and 3 in the closed thermocompression tool, [Fig. 6] Figure 6 is a sectional view of the tool of Figure 5 along section plane VI-VI, and

[0037] [Fig. 7] Figure 7 is a microscopic photograph of a surface of a part made of ceramic matrix composite material obtained by implementing the manufacturing method according to the invention. Description of the embodiments

[0038] The invention applies generally to any part made of ceramic matrix composite material CMC, that is to say comprising a fibrous reinforcement made, for example, with carbon and / or ceramic fibers and densified by a matrix at least partially made of ceramic.

[0039] The invention applies, in particular but not exclusively, to parts made of SiC / SiC material, that is to say parts formed by a fibrous reinforcement of silicon carbide SiC fibers densified by a matrix at least partially made of silicon carbide SiC.

[0040] Figure 1 is a flowchart indicating successive steps of an embodiment of a manufacturing method according to the invention.

[0041] More particularly, different steps of an example of the manufacturing method according to the invention are represented in figure 1.

[0042] Firstly, the manufacturing method according to the invention comprises a forming step 10, during which a fiber preform, comprising in particular silicon carbide fibers, is formed. The fiber preform is in particular intended to form a fiber reinforcement of a part to be obtained.

[0043] The fibers used may be silicon carbide SiC fibers, in particular supplied under the name “Nicalon”, “Hi-Nicalon” or “Hi-Nicalon-S” by the company Nippon Carbon or “Tyranno SA3” by the company UBE.

[0044] The fiber preform can be obtained by three-dimensional weaving between a plurality of layers of warp yarns and a plurality of layers of weft yarns. The three-dimensional weave produced can be an "interlock" weave, that is to say a weave weave in which each layer of weft yarns links several layers of warp yarns with all the yarns of the same weft column having the same movement in the plane of the weave. Different weaving methods that can be used are described in document WO 2006 / 136755.

[0045] The fiber preform may also be obtained by assembling a plurality of fiber textures. In such a case, the fiber textures may be bonded together, for example, by sewing and / or needling. The fiber textures may respectively be obtained from a layer or a stack of several layers of: one-dimensional (UD) fabric, two-dimensional (2D) fabric, braid, knit, felt, and / or unidirectional (UD) web of yarns or cables or multidirectional (nD) webs obtained by superimposing several UD webs in different directions and bonding the UD webs together, for example, by sewing, by chemical bonding agent or by needling.

[0046] Consequently, the manufacturing method according to the invention comprises a consolidation step 20, during which a consolidation phase comprising at least silicon carbide is formed in the porosity of the fiber preform.

[0047] The consolidation phase can be formed by chemical vapor infiltration with the preform held in a shaping tool.

[0048] The consolidation phase may comprise only silicon carbide. Alternatively, the consolidation phase may comprise, in addition to silicon carbide, a self-healing material.

[0049] A thickness of the consolidation phase deposit may be greater than or equal to 500 nm, for example between 1 pm and 30 pm.

[0050] The thickness of the deposit of the consolidation phase is sufficient to consolidate the fiber preform, that is to say to bind the fibers of the preform together sufficiently so that the preform can be handled while maintaining a determined shape without assistance from holding tools.

[0051] After the consolidation phase, the preform remains porous, the initial porosity being, for example, filled only for a minor part by an interphase and the consolidation phase.

[0052] A following step of the manufacturing method according to the invention consists of an injection step 30, during which a slip comprising a powder of silicon carbide particles is injected into the fiber preform, in particular according to the methods known under the names “Slurry Cast” or “Slurry Transfer Molding” in English.

[0053] At the end of the injection step 30, a raw part 100 is obtained.

[0054] Such a raw part 100 is notably presented in figures 2 and 3 which are respectively schematic perspective views showing a deposit of a finishing layer on the raw part 100 in accordance with different embodiments of the invention.

[0055] More specifically, figures 2 and 3 represent the raw part 100 obtained as described previously and which can be intended to form a turbine distributor sector.

[0056] Thus, the raw part 100 may comprise an aerodynamic profile 110 extending in: a longitudinal direction D L between an internal end 110b and an external end 110a, and in a transverse direction D T between a leading edge 111 and a trailing edge 112.

[0057] The aerodynamic profile 110 may also comprise a lower surface face 113 and an upper surface face 114, connecting the leading edge 111 and the trailing edge 112, on either side of the aerodynamic profile 110.

[0058] The blank 100 may also comprise an inner platform 130 connected to the inner end 110b of the airfoil 110 and an outer platform 120 connected to the outer end 110a of the airfoil 110.

[0059] In accordance with the invention, the manufacturing method according to the invention provides one or more finishing layers deposited on the raw part 100. For this purpose, the manufacturing method according to the invention comprises a deposition step 40, during which at least one finishing layer 200 is deposited on all or part of the raw part 100.

[0060] In the example described here in Figures 2 and 3, a finishing layer 200 is deposited on a portion of the surface of the raw part 100, covering a surface of the aerodynamic profile 110 comprising the leading edge 111, the surfaces of the internal platform 130 and the external platform 120 facing each other, a connection zone 1111 connecting the aerodynamic profile 110 to the external platform 120, and a connection zone 1110 connecting the aerodynamic profile 110 to the internal platform 130.

[0061] In the example described here, the finishing layer 200 is deposited on so-called “functional” areas of the raw part 100, i.e. areas present in a flow stream of an aircraft engine and defining a portion of a flow vein. The finishing layer 200 comprises at least one silicon carbide SiC powder, a binder and / or a plasticizer. In particular, the binder and the plasticizer form a support system for the silicon carbide SiC powder which is dispersed in such a support.

[0062] A finishing strip can be prepared using a tape casting technique, also known as "tape casting". This principle is relatively simple and involves applying a slip to a support by passing a shoe over it.

[0063] A deposit thickness adjustment is ensured by a system consisting of two blades. To produce a finishing layer according to the invention, a slip comprising at least one silicon carbide SiC powder, a binder, a plasticizer, a dispersant and / or an organic solvent are used.

[0064] A slip sheet is thus deposited and then dried in order to remove the solvent, thus making it possible to obtain the finishing layer 200 comprising the support in which the silicon carbide SiC powder is dispersed.

[0065] The backing is both strong enough to be easily handled and soft enough to be easily conformable due to the presence of a plasticizer.

[0066] Removal assistance tools may be used if necessary.

[0067] The topcoat 200 may have a thickness between 0.2 mm and 0.5 mm.

[0068] Additionally, the topcoat may contain between 30% and 60% by volume of silicon carbide SiC powder. The remainder of the topcoat may consist mainly of the binder and plasticizer.

[0069] Silicon carbide SiC powder can be micron sized with an average particle size between 0.5 pm and 3 pm.

[0070] The binder may include, but is not limited to: poly(vinyl alcohol) PVA, poly(vinyl butyral) PVB, high molecular weight polyethylene glycol PEG, and / or glycerol.

[0071] The plasticizer may include: low molecular weight polyethylene glycol PEG, dibutyl phthalate DBP, and / or triethylene glycol TEG.

[0072] The dispersant may in particular be chosen from commercial products, in particular those supplied under the name “Disparlon” by the company Kusumoto Chemicals Ltd, “Disperbyk” by the company BYK, “Triton H-66” by the company The Dow Chemical Company, or “Melioran P312L” by the company Arkema.

[0073] The solvent may in particular be a polar solvent, such as methyl ethyl ketone MEK, acetone, ethanol, propanol, butanol, or more generally all ketones and all alcohols.

[0074] By appropriate flexibility, the finishing layer 200 correctly matches the geometry of the raw part 100, in particular at the leading edge 111 and, on the one hand, the connection zone 1111 connecting the aerodynamic profile 110 and the external platforms 120 and, on the other hand, the connection zone 1110 connecting the aerodynamic profile 110 to the internal platform 130, which correspond to zones having geometry deviations from the theoretical geometry. Consequently, the manufacturing method according to the invention comprises a thermocompression step 50, during which the finishing layer 200, respectively the finishing layers 200, deposited on the raw part 100 is, respectively are thermocompressed, according to a final geometry of the part made of ceramic matrix composite material.

[0075] More specifically, Figure 4 presents a schematic perspective view showing an installation of the raw part 100 of Figures 2 and 3 comprising the finishing layer 200 in a thermocompression tool 500 comprising, according to the embodiment example presented: a first shell 501, comprising a first imprint 511 corresponding in part to the shape and final dimensions of the part to be produced, in particular a blade, and surrounded by a first contact plane 512, and a second shell 502, comprising a second imprint 521 corresponding in part to the shape and final dimensions of the part to be produced, in particular a blade, and surrounded by a second contact plane 522. The second contact plane 522 of the second shell 502 is intended to cooperate with the first contact plane 512 of the first shell 501.

[0076] The first shell 501 and the second shell 502 may in particular be made of a metallic material such as aluminum or steel for example. Once the thermocompression tool 500 is closed as shown in FIG. 5, it is placed in a dedicated device, not shown, making it possible to carry out the thermocompression according to the thermocompression step 50, such as for example a mechanical press, an autoclave and / or an autonomous device equipped with hydraulic cylinders. For this purpose, the dedicated device is equipped with heating means, capable of heating the raw part 100 to a determined temperature.

[0077] According to a variant, the thermocompression tool 500 can itself be equipped with heating means, such as, for example, electrical resistors.

[0078] A typical non-limiting cycle of the thermocompression step 50 comprises at least: a pressure application sub-step, during which a holding pressure, in particular between 1 and 2 MPa, is applied to the first shell 501, to the second shell 502 and / or to the thermocompression tool 500; a heating sub-step, during which the temperature is increased until a creep temperature of the finishing layer(s) 200 is reached, the creep temperature typically being between 45°C and 150°C; a compaction sub-step, during which the raw part 100 and the finishing layer(s) 200 are compressed / compacted, in particular at a pressure between 2 and 15 MPa; and a shutdown sub-step, during which the heating is stopped or reduced and the pressure released, in particular to a value between 1 and 2 MPa.

[0079] Applying pressure to the thermocompression tool 500 maintained at a temperature greater than or equal to the creep temperature of the finishing layer(s) 200 makes it possible to calibrate the thickness of the finishing layer(s) 200. In this way and as illustrated in FIG. 6 as a sectional view of the tool of FIG. 5 along the section plane VI-VI, the finishing layer 200 is perfectly distributed in thickness on an external surface of the raw part 100. The finishing layer 200 thus fills in material gaps and surface irregularities of the raw part 100, which makes it possible to correct the geometry of the raw part 100 to achieve a theoretical geometry of the final part defined by the first impression 511 and the second impression 521, when the first contact plane 512 and the second contact plane 522 are joined.

[0080] After cooling of the thermocompression tool 500, the raw part 100 is demolded.

[0081] At this stage, all parts of the raw part 100 covered by one or more finishing layers 200 are conformed to the theoretical, i.e. final, geometry and dimensions of the part to be manufactured.

[0082] The manufacturing method according to the invention then comprises an infiltration step 60, during which the part, obtained at the end of the thermocompression step 50, is infiltrated with a composition comprising mainly silicon in mass.

[0083] The composition may be molten silicon alone or a molten silicon alloy, which may further contain one or more other elements, such as titanium, molybdenum, boron, iron and / or niobium.

[0084] The mass content of silicon in the molten composition may be greater than or equal to 90%.

[0085] During the infiltration step 60, the organic elements of the finishing layer(s) are converted into carbon residues by a heat treatment preceding the infiltration of the molten silicon by capillary rise in a granular stack of silicon carbide SiC particles. The heat treatment may correspond to the temperature rise of the preform and the silicon-based composition bath or a specific heat treatment carried out under vacuum.

[0086] During infiltration, part of the silicon reacts with the carbon residues, silicon carbide SiC particles and silicon carbide SiC macrocrystals are precipitated while the excess silicon remains as free silicon in the material, as well as in the internal matrix of the ceramic matrix composite material. At the end of the manufacturing method according to the invention, a part made of ceramic matrix composite material is thus obtained, which has a geometry with no deviation, or a minimal deviation, compared to a theoretical geometry of the intended part and a very good surface condition in all the surface areas treated with the finishing layer(s) 200.

[0087] Figure 7 is a microscopic photograph of the surface of the part made of ceramic matrix composite material obtained by implementing the manufacturing method according to the invention.

[0088] More particularly, FIG. 7 shows the surface of the part made of ceramic matrix composite material after the infiltration step 60 with a molten silicon-based composition, the part having been coated with a finishing layer 200 after the slip injection step 30 and before the infiltration step 60 as described previously.

[0089] Figure 7 shows that the ceramic matrix composite material 1000 of the part comprises on the surface, at this stage of densification, a layer 2000 consisting of silicon carbide SiC and silicon Si obtained from the finishing layer 200. The layer 2000 does not show any decohesion with the ceramic matrix composite material 1000, which ensures very good resistance of the assembly.

[0090] The table below shows surface finish values ​​including roughness Ra, maximum ripple profile height Wz and total ripple profile height Wt measured on the surface of the ceramic matrix composite part coated with a 200 topcoat as described previously.

[0091] Three different topcoat 200 formulations were used for these tests, respectively: topcoat #1 comprising 1 |im SiC silicon carbide powder, acrylic binder (DEGALAN M920) and PEG 200; topcoat #2 comprising 1 .8|im SiC silicon carbide powder, acrylic binder (DEGALAN M920) and PEG 200; and topcoat #3 comprising 1 .8|im SiC silicon carbide powder, acrylic binder (DEGALAN M920) and DBP.

[0092] Surface state measurements are carried out after the infiltration step with molten silicon Ml. of the ceramic matrix composite part is considerably improved compared to the surface condition of the same part in raw ceramic matrix composite material, i.e. the surface of which has not been coated with a finishing layer 200.

[0093] The manufacturing method of the invention is particularly advantageous for the manufacture of ceramic matrix composite material parts used in gas turbines of aeronautical engines, such as a blade or a nozzle sector comprising an aerodynamic profile connected to at least one platform. Indeed, the areas at the leading edge and / or at the connecting parts between the aerodynamic profile and the platform(s) often have deviations, in particular lack of material, compared to the theoretical geometry of the final part. Such areas correspond to functional areas of the part, i.e. areas present in a flow stream of the engine.

[0094] The manufacturing method of the invention makes it possible to easily correct the geometry of such areas and thus achieve the specifications targeted for the part in operation.

[0095] The manufacturing method of the invention also makes it possible to improve the surface condition, particularly in smoothing, of such functional areas and, consequently, the aerodynamic performance of the part.

Claims

Claims

1. A method of manufacturing a part made of ceramic matrix composite material comprising at least: a forming step (10), during which a fiber preform is formed, a consolidation step (20), during which a consolidation phase of the fiber preform is carried out, an injection step (30), during which a slip comprising a powder of silicon carbide particles is injected into the fiber preform, so as to obtain a raw part (100), and an infiltration step (60), during which a composition based on molten silicon is infiltrated, so as to form a part made of ceramic matrix composite material, the method further comprising: a deposition step (40), during which at least one finishing layer (200) is deposited on all or part of the raw part (100), and a thermocompression step (50),during which the finishing layer (200) is thermocompressed according to a final geometry of the part made of ceramic matrix composite material, characterized in that the deposition step (40) and the thermocompression step (50) take place after the slip injection step and before the step of infiltrating the raw part with a composition based on molten silicon and in that, during the deposition step (40), the finishing layer (200) comprises a silicon carbide powder dispersed in a support comprising at least one binder and / or one plasticizer.,

2. A manufacturing method according to claim 1, wherein the finishing layer (200) is made by tape casting from of a slip comprising at least one silicon carbide powder, a binder, a plasticizer, a dispersant and / or an organic solvent.

3. A manufacturing method according to any preceding claim, wherein the topcoat (200) comprises a percentage of silicon carbide particles of between 30% and 60% by volume.

4. Manufacturing method according to any one of the preceding claims, in which the finishing layer (200) has, before the thermocompression step, a thickness less than or equal to 0.5 mm, in particular between 0.2 mm and 0.5 mm.

5. Manufacturing method according to any one of the preceding claims, wherein, during the thermocompression step (50), the raw part (100) is heated to a temperature above the creep temperature of the support of the finishing layer, the finishing layer being compacted on the raw part (100) according to the final geometry of the part made of ceramic matrix composite material.

6. A manufacturing method according to any preceding claim, wherein at least one finishing layer (200) is deposited on a surface of a leading edge (111) of the airfoil (110).

7. Manufacturing method according to claim 6, in which the finishing layer (200) is deposited on a connection zone (1 1 10, 1 11 1 ) connecting the aerodynamic profile (1 1 1 ) to the platform (1 10, 120).

8. Part made of ceramic matrix composite material obtained from the manufacturing method according to any one of the preceding claims, characterized in that it constitutes a blade or a gas turbine distributor sector comprising an airfoil (110) connected to at least one platform (110, 120).

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