METHOD FOR MANUFACTURING A PART FROM CERAMIC MATRIX COMPOSITE MATERIAL

Abstract

The invention relates to a method (100) for manufacturing a part from a ceramic matrix composite (CMC), the ceramic matrix comprising a polymer binder, ceramic microbeads and a solvent, the method comprising the following steps: - producing (110) preforms in the form of a fabric of fibers or a ply of dry fibers to be impregnated or pre-impregnated with the ceramic matrix, - draping (120) each of the preforms separately so as to form a plurality of components, - thermally curing (130) at least one of the components obtained during the draping step (120) at a temperature greater than or equal to the evaporation temperature of the solvent, - arranging (140) the components so as to form a pre-assembly, having the shape of the part to be manufactured, in which each of the components is in contact with at least one cured component at at least one contact zone, the components being chemically unbonded,and - to perform (160) a sintering of the pre-assembly so as to chemically bond the components together and obtain a single-piece part. Figure for the abbreviation: Figure 2,

Description

Title of the invention: METHOD FOR MANUFACTURED A PART FROM A CERAMIC MATRIX COMPOSITE MATERIAL Technical field of the invention

[0001] The invention relates to the technical field of manufacturing processes for a part made of ceramic matrix composite material, particularly of the oxide-oxide type. "Manufacturing" means any operation or series of operations aimed at producing, creating, or assembling such a part made of ceramic matrix composite material.

[0002] The invention applies to any part made of ceramic matrix composite material, particularly of the oxide-oxide type. That being said, it is of particular interest for complex parts intended to be manufactured by direct assembly of components with a geometric shape simpler than the final part to be produced. Technical background

[0003] Oxide-oxide composite (COX) matrix composite parts are increasingly used in modern turbomachinery, particularly aircraft turbomachinery, due to their low weight and good temperature resistance at high operating temperatures. Conventionally, the oxide-oxide matrix composite is produced by powder sintering, which allows for adjustment of its microporosity and mechanical properties, resulting in a lighter final part. This oxide-oxide matrix composite is used in conjunction with a solution as an impregnation material within a fabric or ply of pre-impregnated fibers to form the composite material. Currently, the most common COX material in aircraft consists of alumina or aluminosilicate-based fibers, for example, alumina oxide fibers, impregnated with a matrix based on alumina oxide powder and solvent.

[0004] As illustrated in [Fig. 1], a prior art method 10 for manufacturing a component for a composite part conventionally comprises a first step 11 of producing fibrous preforms, a second step 12 of manually or automatically draping 2D plies in a mold so as to form the component by superimposing layers of 2D fabric. In a third step 13, the component is then thermally hardened in the presence of a polymer binder and at least one solvent, which allows its shape to be fixed. In a fourth step 14, the component is demolded, and then, in a fifth step 16, sintering is carried out. This fifth step 16 is implemented in two stages. The component is first brought to a medium temperature enabling the polymer binder to polymerize, which This process ensures the geometry of the component during the handling and operations that take place between the third step (13) of curing and the fifth step (16) of pyrolysis of the polymer binder and sintering. The component, in particular the oxide-oxide composite matrix, is then consolidated by sintering the ceramic microbeads that make up its component. This process (10) allows for the fabrication of a single component and is repeated as many times as necessary to manufacture all the components required to form the part to be manufactured. Each component then undergoes an operation to attach it to other components.

[0005] Like the methods of assembling metal components, the various COX components are generally assembled to each other by riveting and / or bolting, which therefore implies making at least one hole 17 in each component after the sintering step 16. However, this complex fastening technique is the cause of a break in the continuity of the composite material, a local weakening of the part, an increased risk of delamination during the assembly of the components or in operation, etc. because it involves drilling the components.

[0006] Furthermore, CMC materials, including COX, have a specific coefficient of thermal expansion (-6 to 8 x 10⁶ K⁻¹), whereas the components used (rivets, bolts) to create this type of assembly are frequently nickel-based materials to ensure adequate temperature resistance. However, the coefficient of thermal expansion of nickel is much higher (-20 x 10⁶ K⁻¹) than that of CMC materials, which implies a high risk of damage to CMC parts subjected to thermal loading.

[0007] More recently, prior art has employed chemical bonding 17' using adhesive films to assemble COX components. However, the availability of materials used in these adhesive films is limited, not to mention the difficulty in finding materials with good temperature resistance. Indeed, the assembly of the components is carried out during the first step and therefore necessarily precedes the sintering step.

[0008] Furthermore, the draping of material to form a component is limited by the complexity of the mold required for components with complex shapes, generally an order of magnitude more complex than the desired component. This necessitates the creation of tooling with movable or removable keys, the ability to demold components after hardening, and the precise adjustment of sealing or temperature ranges for draping and sintering. It is therefore more common to produce a multitude of simple components and then assemble them, which is costly.

[0009] The invention provides a method for manufacturing a part in COX material without drilling, which is simpler to implement and less expensive than the methods known until now. Summary of the invention

[0010] The invention proposes for this purpose a method for manufacturing a part made of ceramic matrix composite (CMC) material, the ceramic matrix comprising a polymer binder, ceramic microbeads and a solvent, the method comprising the following steps:

[0011] - to produce a plurality of preforms in CMC materials, each of the preforms presenting in the form of a fabric of fibers or a ply of fibers impregnated with the ceramic matrix, and impregnating each of the preforms with the ceramic matrix when the fabric or ply of fibers is made of dry fibers

[0012] - drape each of the preforms separately so as to form a plurality of components,

[0013] - thermally harden at least one of the components obtained during the step of draping at a temperature equal to or greater than the solvent's evaporation temperature,

[0014] - arrange the components so as to form a pre-assembly having the shape of the part to be manufactured, wherein each of the components is in contact with at least one hardened component during the hardening step at at least one contact zone, the components being chemically separated from each other, and

[0015] - to perform a sintering of the pre-assembly so as to chemically bond the components between them and thus obtain a one-piece unit.

[0016] The manufacturing process according to the invention solves the aforementioned problems of the prior art. The invention makes it possible to manufacture both simple and complex parts without drilling the material. Indeed, the various components are bonded together by chemical bonds that form between them, particularly in the contact areas, during the sintering stage of the pre-assembly. Thus, the risks of breakage in the continuity of the composite material, local embrittlement of the part, and delamination during the assembly of the components are greatly reduced. This prevents premature wear of the part and therefore allows it to maintain good mechanical strength throughout its service life.

[0017] In addition, the invention does not require an adhesive film to bond the components together, which makes it possible to overcome the constraints related to the use of these films, in particular the limited availability of materials which have both good temperature resistance and which allow the adhesion of the CMC components.

[0018] Although the invention applies to simple parts as well as complex parts, the impact of the invention is particularly notable for parts with complex shapes due to the multiplicity of components to be joined together to obtain the desired shape for manufacturing the part.

[0019] According to various features of the invention which may be taken together or separately: • the polymer binder is capable of transforming into ceramic, said polymer binder being a pre-ceramic binder; • all the components obtained during the draping step are hardened; • the process further includes, subsequent to the arrangement stage and prior to the sintering stage, a stage during which the components of the pre-assembly are hardened jointly; • During the assembly stage, the components are brought into contact with each other by inserting a fold of material at the contact areas, • during the joint hardening stage, a contacting force is exerted simultaneously on the components; • the material fold is made of the same material as the preforms; • The component(s) are simultaneously hardened by heating in autoclave; • the process includes, following the sintering step, a step consisting of carrying out finishing work on the part; • The CMC material is of the oxide-oxide (COX) type, • COX material is based on alumina or aluminosilicate; • the impregnation of the preforms is carried out by injection; • The manufactured part is a part for an aircraft turbomachine. Brief description of the figures

[0020] Other objects, features and advantages of the invention will become more apparent in the following description, made with reference to the accompanying figures, in which:

[0021] - Fig. 1 illustrates a method for manufacturing a part in CMC according to art previous;

[0022] - [Fig. 2] illustrates a method for manufacturing a CMC part according to a first method of embodiment of the present invention;

[0023] - [Fig. 3] illustrates a method for manufacturing a CMC part according to a second method of embodiment of the present invention;

[0024] - Figure 4 illustrates a method for manufacturing a CMC part according to a third embodiment of the present invention.

[0025] In the figures, optional steps are indicated by boxes made of dotted lines. Detailed description of the invention

[0026] With reference to figures 2 to 4, the invention relates to a method 100 for manufacturing a part in ceramic matrix composite (CMC) material, preferably of the oxide-oxide (COX) type.

[0027] COX materials are a particular type of ceramic matrix composite material known by the abbreviation CMC. COX materials typically comprise an oxide-based fiber fabric or fiber ply impregnated with an oxide-based matrix used in conjunction with a solution, in which the matrix and the solution form what is called the impregnation material. The fiber fabric or, as the case may be, the fiber ply conventionally forms what is called the reinforcement of the composite material.

[0028] Parts made with COX materials differ from parts made with CMC materials in their increased lightness and good temperature resistance at high operating temperatures. Reducing the weight of turbomachinery is a constant research topic for aeronautical professionals because it helps reduce aircraft fuel consumption. In the context of this invention, the manufactured part is more particularly a part for an aircraft turbomachine. It is a part that has a simple or complex shape.

[0029] A "complex shape" is defined as any shape whose geometry requires the assembly of at least two components to reconstruct it. Conversely, a "simple shape" is defined as any shape whose geometry requires the assembly of at most one component. Generally, shapes are complex due to a high aspect ratio between their dimensions or the existence of at least one portion exhibiting a non-zero Gaussian curvature. Other factors can also contribute to the geometric complexity of a shape.

[0030] Fig. 2 illustrates a method of manufacturing the invention according to a first embodiment.

[0031] According to this first embodiment, the manufacturing process 100 comprises a first step consisting of producing 110 a plurality of preforms from CMC materials, each of the preforms being in the form of a fabric of fibers or a ply of dry or pre-impregnated fibers of the ceramic matrix. These fibrous preforms are generally basic shapes in one, two, or three dimensions produced by cutting and intended to have the shape of the components constituting the part to be manufactured. Two-dimensional fibrous preforms are produced by stacking of layers. Whatever the shape of the preform, it is therefore likely to evolve during the subsequent stages of the manufacturing process. By way of non-limiting examples, the fibers can be unidirectional (0° fibers), or form a biaxial weave (e.g., 0° and 90° woven fibers), etc.

[0032] According to a particular embodiment, the ply or fabric is made of aluminum oxide fibers, i.e., aluminum oxide (alumina) fibers. Alumina fibers are not only electrically insulating but also very chemically stable, which gives them excellent corrosion resistance. In aircraft turbomachinery, corrosion can be generated by reactants, typically reactants from the combustion chamber. Aluminosilicate fibers are also excellent candidates for making the fiber ply or fabric. The use of alumina oxide or aluminosilicate fibers results in the formation of a COX material part in which the COX material is therefore alumina-based or, as the case may be, aluminosilicate-based.

[0033] The ceramic matrix typically comprises ceramic microbeads, a polymer binder, a solvent, and possibly other additives well known to those skilled in the art. The ceramic microbeads are capable of bonding to one another when the material is heated to a temperature typically used during the sintering operation. At this temperature, i.e., the temperature used during the sintering operation, the polymer binder is pyrolyzed and thus degraded, which also contributes to chemically bonding the different components together. The solvent, for its part, is intimately bound to the polymer binder and allows the ceramic matrix to be conveyed more easily to the surface of the components, and even to the interstices between the material fibers. In addition, the solvent provides a viscosity suitable for rheological finishing of the ceramic matrix.

[0034] Fibers are said to be pre-impregnated with the ceramic matrix when they are previously impregnated with the ceramic matrix. However, as previously indicated, the fibers may not be pre-impregnated; that is, they may be dry. In this case, during the first step 110, each of the dry fibrous preforms 115 is impregnated with the ceramic matrix. Advantageously, this impregnation 115 is carried out by injection, that is, by a combination of injection and infusion of each of the preforms into the ceramic matrix. In practice, the dry fibrous preform is placed in a sealed tool, and then the ceramic matrix is ​​inserted into it until a predetermined fiber-to-ceramic matrix ratio is obtained. This is one example of a method for carrying out the impregnation 115, although the invention is not limited to it.

[0035] The manufacturing process 100 further comprises a step in which each of the preforms is draped separately 120 so as to form a plurality of components. In practice, draping 120 is an operation in which an object is formed by superimposing 2D folds or layers of unidirectional or 2D fabric. This operation can be performed manually or automatically, for example by needle punching, for each of the components to be produced. The components thus formed will be assembled in a subsequent step, according to a predetermined process, to obtain the shape of the part to be manufactured.

[0036] In a third step, at least one of the components obtained in the draping step 120 is thermally hardened. In this respect, this thermal hardening should be carried out at a temperature equal to or greater than the solvent's evaporation temperature, allowing the solvent to transition to a gaseous state and thus be at least partially removed from the components. The thermal hardening 120 is performed at a temperature lower than the temperature(s) required for sintering because solvent evaporation requires only a moderate temperature. This hardening step 130 can be carried out in a suitable curing device, for example, an oven or, more advantageously, an autoclave.

[0037] In a particular embodiment, the ceramic matrix is ​​a powder of alumina and silicon oxide (silica), which is better suited when the fibers of the ply or fiber fabric are alumina fibers.

[0038] At the end of the hardening step 130, the hardened component(s) has a higher hardness index than at the end of the draping step 120. This stiffening facilitates the implementation of the subsequent steps of the process 100 by making the hardened component(s) easier to handle.

[0039] At this stage, it should be clarified that it is not mandatory to harden all the components. Indeed, depending on the part to be manufactured and its complexity, it may only be necessary to harden one or a few components to achieve the subsequent assembly.

[0040] In this regard, the manufacturing process 100 further comprises a fourth step in which the components 140 are arranged to form a pre-assembly having the shape of the part to be manufactured, and in which each component is in contact with at least one component hardened during the hardening step 130 at at least one contact zone. The arrangement allows each component to be positioned at a predetermined location in order to constitute the shape of the part to be manufactured.

[0041] At the end of this arrangement step 140, a pre-assembly is obtained which has the shape of the part to be manufactured. Although in contact with each other at the contact areas, the components are chemically separated from each other, that is to say, the components are not chemically bonded and do not They are not yet fixed together. At this stage, it is still possible to move the components relative to each other.

[0042] According to this first embodiment of the invention, the manufacturing process 100 includes a step in which 160 a sintering of the pre-assembly is carried out so as to chemically bond the components together and thus obtain a monobloc part.

[0043] Sintering 160 is typically carried out at a temperature above the softening temperature of the ceramic microbeads so as to soften the ceramic microbeads in the contact areas and thus obtain a monolithic part in which the components are chemically bonded to each other. Simultaneously, the polymer binder present in the ceramic matrix degrades and promotes chemical contact between the components. Thus, the softened ceramic microbeads and the degraded polymer binder in the contact areas form material bridges that create chemical bonds between the components, similar to what occurs during welding without the addition of material.

[0044] Indeed, the ceramic microbeads and the polymer binder are already present in the ceramic matrix of the components, so no additional material is needed at the contact points to create chemical bonds between the components and secure them during the sintering step 160. This significantly simplifies the manufacturing process 100 according to this first embodiment of the invention compared to known prior art processes. Since all the components are bonded to one another, this step results in a single-piece part in which the components are chemically bonded. "Single-piece" means that the part is a single, continuous piece without any discontinuities or interfaces.

[0045] The process according to the invention makes it possible to manufacture both simple and complex parts without drilling the components. Thus, the risks of breakage in the continuity of the composite material, local weakening of the part due to damage to the components or differential thermal expansion, and delamination during component assembly are eliminated. This prevents premature wear of the part and therefore extends its lifespan. It should also be noted that avoiding the use of complex fasteners, i.e., rivets, bolts, adhesive films, etc., results in a reduction in the mass of the manufactured part.

[0046] Furthermore, the invention also does not require adhesive films to bond the components together, which makes it possible to overcome the constraints related to the use of these films, in particular the limited availability of materials which have both good temperature resistance and which allow the adhesion of the CMC components.

[0047] Although the invention applies to both simple and complex parts, its impact is particularly noticeable for parts with complex shapes because the more components there are to join together to obtain the desired shape, the more complex the prior art processes become due to the increased drilling or bonding of adhesive films required. This is not the case in the process according to this first embodiment of the invention. Indeed, for an equal number of components, the time required to attach the components together is less in the process according to the first embodiment of the invention compared to known processes because there is no drilling and no bonding at each contact point.

[0048] Incidentally, complex finished parts can be produced by direct assembly of simple components. This limits the complexity of the tooling required during draping 120. Furthermore, holding the components in position and shape during arrangement 140 requires relatively simple tooling, even if a subsequent sintering step 160 is required.

[0049] According to a particular embodiment, the polymer binder is advantageously capable of transforming into a ceramic, which facilitates and improves the reliability of the bonding of the components to each other during the sintering operation compared to other types of commercially available polymer binders. Such a polymer binder is called a "pre-ceramic binder" because it transforms into a ceramic at the temperatures used during the sintering process.

[0050] Still with reference to [Fig.2] and according to the first embodiment of the invention, all the components obtained during the draping step 120 can be hardened 130, that is to say that all the components are hardened during the hardening step 130. This is particularly useful for components that are difficult to handle, for example because these components have a particular shape, or because their initial hardness index, when raw, is low or for any other reason.

[0051] The manufacturing process 100 may include, following the sintering step 160, a step consisting of carrying out 170 of the finishing on the manufactured part.

[0052] That being said, the hardening of the components can be carried out in two steps. Figure 3 illustrates a second embodiment of the manufacturing process 100 according to the invention. This embodiment differs from the first embodiment only in the way the components are hardened. In this respect, the manufacturing process 100 according to the second embodiment of the invention further comprises, subsequent to the assembly step 140 and prior to the sintering step 160, a step during which the pre-assembly components 150 are hardened together.

[0053] In this second embodiment, the process 100 therefore includes an additional hardening step 150. This additional hardening step 150 is not carried out under the same conditions as the first hardening step 130 because the components are already arranged relative to each other to take the shape of the part to be manufactured. This not only makes the components even more rigid but also stiffens, and therefore stabilizes, the pre-assembly before proceeding with sintering 160.

[0054] This is particularly advantageous if the operator wishes to achieve the arrangement 140 with components which, although hardened, remain sufficiently flexible to facilitate the arrangement 140. In this case, the components have, at the end of the first hardening step 130, an intermediate hardness index, that is to say, a hardness index located between its initial hardness index, i.e., the hardness index of the component(s) before the first hardening step 130, and its hardness index after the additional hardening step 150. The additional hardening step 150 thus completes what was initiated during the first hardening step 130.

[0055] Alternatively, this can also be advantageous when only one component or part of the components has been hardened during the first hardening step 130. Thus, the additional hardening step makes it possible to harden the components while they are still raw, which ensures a sufficient level of rigidity of the ceramic matrices of all the components and makes it possible to avoid mechanical stresses which could possibly occur due to differences in the rigidity of the components.

[0056] Concurrently with the additional hardening step 150, a contacting force can be applied to the components. This makes it possible to stiffen the pre-assembly to the desired shape, i.e., the shape of the part to be manufactured, by creating an assembly as close as possible to the final shape, and without gaps.

[0057] According to a third embodiment illustrated in [Fig. 4], during the assembly step 140, the components are brought into contact 142 with each other by inserting 144 a layer of material at the contact areas, after all the components have been hardened. These layers of material cannot in any way be considered equivalent to the adhesive films used in prior art processes because they are integrated during the assembly step 140, and therefore during the formation of pre-assemblies.

[0058] This third embodiment differs from the second embodiment only in that the components are brought into contact 142 by inserting 144 a fold of material at the contact areas. This means that in the process according to this embodiment, following the production steps 110 of the plurality of preforms, draping 120 and hardening 130, as defined previously:

[0059] - the arrangement 140 of the components is implemented by placing the components in contact 142 with each other while inserting 144 a fold of material so as to form the pre-assembly having the shape of the part to be manufactured, all the components having been hardened during the first hardening stage 130, the components being chemically separated from each other,

[0060] - the joint hardening of the pre-assembly components,

[0061] - the realization 160 of the sintering.

[0062] The material folds inserted between the previously hardened components have not undergone the first hardening step 130 and are not hardened at the time of their insertion between the components during the assembly step 140. As this latter step is followed by the additional hardening step 150, it is only at this point that the material folds are hardened, enabling them to conform to the shape of the components at the contact areas while stabilizing their shape at said contact areas. In this third embodiment, the "contact areas" are actually bonding areas between the components, since the latter are not directly in contact with each other but via the material folds.

[0063] The use of a material ply is advantageous because no tooling capable of withstanding the sintering temperature during step 160 is required to hold the components in position. Furthermore, the use of a ply allows the hardened component(s) to be brought back into contact with "raw" and still highly solvated material, which promotes the formation of continuity in the matrix and allows for better assembly following sintering.

[0064] Furthermore, the use of a material fold also prevents the physical separation of the components (brought into contact) during the additional hardening step 150 when carried out in an autoclave. Indeed, physical separation of the components can occur due to the geometric shrinkage of the material during autoclave hardening, which can create an irreparable gap between two components during sintering. This can be anticipated by careful design of the individual components and the application of a contacting force during hardening, as mentioned previously.

[0065] Except for step 140, the other steps of process 100 according to the third embodiment are, as indicated above, carried out under the same conditions as those of the second embodiment.

[0066] Still with reference to [Fig.4] and the third embodiment, it is advantageous that the material folds be made of the same material as the preforms. Indeed, this significantly improves mechanical strength at the bonding zones. The material plies can be formed from a fiber fabric or from a fiber ply. If the preforms are in the form of a fiber fabric, then the material plies are advantageously made from the same fiber fabric, that is, a fiber fabric with fibers, and a ceramic matrix comprising a solvent, a polymer binder, and ceramic beads of the same type as those of the preforms. Similarly, if the preforms are in the form of a fiber ply, then the material plies are advantageously made from the same fiber ply, including the same fibers, namely a fiber ply with fibers of the same type as those of the preforms.

[0067] Alternatively, each ply of material may be in the form of a matrix or any other shape allowing its insertion between the components. In this case as well, the matrix may advantageously be of the same nature as the ceramic matrix used in the fiber plies or fiber fabrics.

[0068] The second and third embodiments require additional heat treatment cycles in autoclave to implement the additional hardening step 150, which may impact the organization of production.

[0069] The configurations shown in the cited figures are only possible examples, by no means limiting, of the invention which on the contrary encompasses the variants of designs within the reach of the person skilled in the art.

Claims

Demands

1. A method (100) for manufacturing a part from a ceramic matrix composite (CMC) material, the ceramic matrix comprising a polymer binder, ceramic microbeads, and a solvent, the method (100) comprising the following steps: - producing (110) a plurality of preforms from CMC materials, each of the preforms being in the form of a fabric of fibers or a ply of dry or pre-impregnated fibers of the ceramic matrix, and impregnating (115) each of the preforms with the ceramic matrix when the fabric or ply of fibers is made of dry fibers, - draping (120) each of the preforms separately so as to form a plurality of components, - thermally curing (130) at least one of the components obtained during the draping step (120) at a temperature greater than or equal to the evaporation temperature of the solvent, - arranging (140) the components so as to form a pre-assembly having the shape of the part to be manufactured,in which each component is in contact with at least one hardened component during the hardening step (130) at at least one contact zone, the components being chemically separated from each other, and - perform (160) a sintering of the pre-assembly so as to chemically bond the components together and thus obtain a single-piece part.

2. A method (100) according to claim 1, wherein the polymer binder is capable of transforming into ceramic, said polymer binder being a pre-ceramic binder.

3. A method (100) according to any one of claims 1 or 2, wherein all the components obtained during the draping step (120) are hardened (130).

4. A method (100) according to any one of claims 1 to 3, further comprising, subsequent to the arrangement step (140) and prior to the sintering step (160), a step in which the pre-assembly components (150) are jointly hardened.

5. A method (100) according to claim 4 when it depends on claim 3, wherein, during the arrangement step (140), the components are brought into contact with each other by inserting a fold of material at the contact areas.

6. Method (100) according to claim 5, wherein the material fold is made of the same material as the preforms.

7. Method (100) according to claim 4, wherein, during the joint hardening step (150), a contacting force is simultaneously exerted on the components.

8. A method (100) according to any one of the preceding claims, wherein the component(s) (130) are simultaneously hardened by heating in an autoclave.

9. A method (100) according to any one of the preceding claims, wherein the CMC material is of the oxide-oxide (COX) type, said COX material being based on alumina or aluminosilicate.

10. A method according to any one of the preceding claims, wherein the manufactured part is a part for an aircraft turbomachine.

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