Method for repairing a part made of a ceramic matrix composite material

Abstract

The invention relates to a method for repairing a part (10) made of a ceramic matrix composite material, the part having a damaged zone, the method comprising: - a step of machining away at least the damaged zone so as to obtain a recessed repair zone (12); - a step of placing, in the recessed repair zone (12), a repair material (13) comprising at least one ply (15) of composite material preimpregnated with a ceramic matrix precursor; - a step of compacting the one or more plies (15) of composite material preimpregnated with the ceramic matrix precursor in order to increase a fiber volume content; and - a sintering step for consolidating a bond between the repair material (13) and the composite part (10).

Description

DESCRIPTION TITLE: METHOD FOR REPAIRING A PART MADE OF A CERAMIC MATRIX COMPOSITE MATERIAL

[0001] The present invention relates to a method for repairing a part made of a ceramic matrix composite material. The invention finds a particularly advantageous, but not exclusive, application in the repair of ceramic matrix composite (CMC) parts used in an aircraft engine.

[0002] The rear wing components of an aircraft engine are traditionally made from a monolithic metallic material. However, increasing engine temperatures and the obsolescence of certain metallic materials are driving the use of ceramic matrix composite materials, also known as oxide-on-oxide (CMC) materials. These materials consist of a fibrous reinforcement, for example, based on alumina (Al₂O₃), and a matrix, for example, also based on alumina and containing a small proportion of silica (SiO₂).

[0003] These materials offer the advantage of excellent oxidation resistance, good mechanical performance at high temperatures, and low manufacturing costs. They can be used, for example, to create a mixer for a turbofan engine that mixes a primary airflow (hot air) with a secondary airflow (cold air), or an exhaust cone that progressively expands the engine's exhaust gases.

[0004] However, these composite materials are directly exposed to external stresses due to the environment in which an aeronautical structure operates, such as impacts on the tarmac, low-energy impacts (hail, gravel, falling tools, etc.) or during production. These external stresses can damage parts made from composite materials.

[0005] Therefore, there is a need to be able to repair a part made of CMC material with a damaged area in order to restore its aesthetic, aerodynamic, and, where appropriate, structural function.

[0006] The invention aims to effectively address this need by proposing a method for repairing a part made of ceramic matrix composite material having a damaged area comprising: - a machining step to remove at least the damaged area in order to obtain a hollowed-out repair zone, - a step of placing, in the hollowed-out repair area, a repair material comprising at least one ply of composite material pre-impregnated with a ceramic matrix precursor, - a compaction step of the pre-impregnated composite material plies with the ceramic matrix precursor to increase the fiber volume ratio, and - a step of applying a sintering cycle to consolidate a bond between the repair material and the composite material part.

[0007] The invention thus enables the implementation of a simple, rapid, and inexpensive repair process aimed at reducing the downtime of an aircraft with a damaged component. The invention also makes it possible to reduce the scrap rate of parts made from a ceramic matrix composite material, allowing them to be repaired instead of destroyed.

[0008] According to one embodiment of the invention, the step of setting up the repair material includes a draping step of a plurality of plies of composite material pre-impregnated with a ceramic matrix precursor so as to fill the hollowed repair area.

[0009] According to one embodiment of the invention, the step of setting up the repair material includes a step of draping at least one first ply of composite material pre-impregnated with a ceramic matrix precursor against a surface of the hollowed repair area and a step of setting up a filling element made of a sintered ceramic matrix composite material having a shape complementary to the hollowed repair area.

[0010] According to one embodiment of the invention, said process further comprises a step of placing at least one second ply of composite material pre-impregnated with a ceramic matrix precursor so as to encapsulate the filling element between the first ply of composite material pre-impregnated with a ceramic matrix precursor and the second ply of composite material pre-impregnated with a pre-impregnated ceramic matrix precursor.

[0011] According to one embodiment of the invention, the hollowed-out repair area has, in cross-sectional view, a flared shape.

[0012] According to one embodiment of the invention, the material removal step by machining is carried out by a water jet containing an abrasive medium such as sand.

[0013] According to one embodiment of the invention, the composite material part being made up of a plurality of sintered plies, water jet parameters are chosen so as to remove the material ply by ply in the damaged area of ​​the composite material part.

[0014] According to one embodiment of the invention, the compaction step of the ply(ies) includes a step of applying an autoclaving cycle to the assembly "composite material part-repair material".

[0015] The compaction stage may include, prior to the application of the autoclaving cycle: - the application of a technical fabric over the entire "composite material part - repair material" assembly, the technical fabric being arranged to allow the evacuation of a solvent contained in the plies of pre-impregnated composite material and to drain any excess matrix, and / or - the application of a membrane around the entire "composite material part-repair material" assembly.

[0016] According to one embodiment of the invention, the autoclaving cycle is carried out at a temperature between 50°C and 250°C and at a pressure between 0.5 and 2.5 MPa.

[0017] According to one embodiment of the invention, the sintering cycle is carried out at a temperature between 1000°C and 1300°C.

[0018] The invention also relates to a repaired composite material part obtained by implementing the process as previously defined.

[0019] The present invention will be better understood and other features and advantages will become apparent upon reading the following detailed description, which includes embodiments given by way of illustration with reference to the accompanying figures, presented by way of non-limiting examples, which may serve to complete the understanding of the present invention and the explanation of its implementation and, where appropriate, contribute to its definition, on which:

[0020] [Fig. 1 a] [Fig. 1 b] [Fig. 1 c] [Fig. 1 d] [Fig. 1 e] Figures 1 a to 1 e schematically represent the different stages of a first implementation of a repair process for a part made of ceramic matrix composite material according to the invention;

[0021] [Fig. 2] Figure 2 is a map showing material shrinkage thickness curves in micrometers as a function of travel speed and pressure of a water jet used during a machining step of a damaged area of ​​a CMC material part;

[0022] [Fig. 3] Figure 3 is a photograph of a cross-sectional view obtained by X-ray tomography of a repaired area of ​​a part made of CMC material by implementing the process according to the invention of Figures 1a to 1e;

[0023] [Fig. 4a] [Fig. 4b] [Fig. 4c] [Fig. 4d] [Fig. 4e] Figures 4a to 4e schematically represent the different stages of a second implementation of a repair process for a part made of ceramic matrix composite material according to the invention.

[0024] It should be noted that structural and / or functional elements common to the different embodiments may have the same references. Thus, unless otherwise specified, such elements have identical structural, dimensional, and material properties.

[0025] We describe below, with reference to figures 1a to 1e, a first implementation of a repair process for a part made of ceramic matrix composite material 10 having a damaged area 11.

[0026] Part 10, shown in Figure 1a, is made of a material comprising a fibrous reinforcement having ceramic fibers, for example, an oxide such as alumina or mullite or another suitable aluminosilicate or ceramic, and a ceramic matrix, for example, an oxide such as alumina, mullite, and silica. Part 10 consists of a plurality of sintered plies of ceramic matrix composite material stacked one on top of the other and bonded together by an autoclaving step followed by a sintering step. Each sintered ply consists of a thin strip, in particular between 50 and 300 micrometers thick.

[0027] As illustrated in Figure 1b, the process includes a machining removal step of at least the damaged area 11 to obtain a repair cavity 12. The material removal machining step aims to remove the damaged material and obtain the cavity with controlled dimensions.

[0028] During the machining step, material is preferably removed from the damaged area 11 as well as material from an area surrounding the damaged area 11. In other words, to obtain the hollowed-out repair area 12, a larger volume of material is removed than the volume occupied by the damaged area 11.

[0029] The shrinkage thickness can be varied to achieve the desired surface roughness and undulation. Surface preparation of the repair cavity 12 allows for adhesion of a composite material ply impregnated with a ceramic matrix precursor during subsequent assembly phases by autoclaving and co-sintering.

[0030] Preferably, the machining step is carried out using a water jet containing an abrasive medium such as sand. Water jet parameters, in particular pressure, speed, and the particle size of the abrasive medium, are important. abrasive flow rate, distance between nozzle and workpiece and travel speed in damaged area 11 are chosen so as to remove a predetermined thickness of workpiece 10 in damaged area 11.

[0031] The mapping in Figure 2 shows thickness curves E of material removal in micrometers as a function of a speed V of water jet movement and a pressure P of the water jet for a given distance between the water jet nozzle and the part.

[0032] For a given thickness E of material to be removed, the lower the speed V of the water jet, the lower the pressure P, and conversely, the higher the speed V of the water jet, the higher the pressure P.

[0033] Advantageously, the waterjet parameters are chosen to remove material from part 10 layer by layer. Preferably, a number of layers are removed corresponding to the thickness of the damaged area 11 plus an additional layer of sintered material. According to the mapping in Figure 2, to remove a layer with a thickness E of approximately 200 micrometers, the waterjet, for example, has a travel speed V of approximately 12 m / min and a pressure P of approximately 40 MPa. More generally, depending on the layer thickness, the waterjet can have a travel speed V between 6 and 18 m / min and a pressure P between 35 and 60 MPa. Alternatively, by adjusting the waterjet parameters, it is possible to remove any desired thickness in a single machining pass, for example, the total removal thickness.

[0034] As illustrated in Figure 1c, the process includes a draping step, in the hollow repair area 12, of a plurality of plies 15 of composite material pre-impregnated with a ceramic matrix precursor so as to fill the hollow repair area 12. The plies 15 are stacked one on top of the other inside the hollow repair area 12. The set of plies 15 constitutes a repair material 13.

[0035] A 15-ply consists of a thin strip, typically between 50 and 300 micrometers thick, comprising a fibrous reinforcement and a ceramic matrix precursor arranged between and around the fibers fibrous reinforcement. Fibrous reinforcement includes, for example, fibers made of alumina or mullite. Fibrous reinforcement can consist of unidirectional fibers, a two-dimensional fiber mesh with weft and warp fibers, or a three-dimensional fiber mesh with additional fibers extending through the thickness of the fiber reinforcement and providing a bond between the weft and warp fibers.

[0036] The matrix precursor comprises a preceramic resin or a ceramic, for example, alumina-based, a small proportion of silica, and organic components (solvent, plasticizer) to impart flexibility to the 15-ply composite material. As an alternative to alumina, the fibers of the fibrous reinforcement and the matrix precursor can be made from any other type of oxide suitable for the application.

[0037] The 15 plies of composite material pre-impregnated with a ceramic matrix precursor can be draped automatically, notably using an AFP (Automatic Fiber Placement) technique. This technique is particularly well-suited to 15 plies with unidirectional fibers. Alternatively, the 15 plies of composite material pre-impregnated with a ceramic matrix precursor can be draped manually. This technique is particularly well-suited to 15 plies with a two- or three-dimensional fiber mesh.

[0038] As illustrated in Figure 1d, the process includes a step of covering the plies 15 of composite material pre-impregnated with a ceramic matrix precursor. The covering step consists of placing a technical fabric 17 over the stack of plies 15. The technical fabric(s) 17 may be of different types. Their role is to drain any excess matrix initially present in the pre-impregnated fabric, to allow the evacuation of solvents, and ultimately, to reduce the porosity of the intermediate plies 15.

[0039] The process includes a compaction step of 15 plies of pre-impregnated composite material of the ceramic matrix precursor to increase a volumetric fiber content.

[0040] To this end, after placing a membrane 18 around the technical fabric 17 and the assembly "10-ply composite material part 15" placed on a support 19, an autoclaving cycle is applied under a pressure between 0.5 and 2.5 MPa, preferably between 0.5 and 1.5 MPa, and a temperature between 50 and 250°C. The autoclaving cycle lasts between 5 and 10 hours. The autoclaving cycle aims to pre-cure the plies 15 to increase the fiber volume fraction in the plies 15 due to the partial removal of the organic components contained in the ceramic matrix precursor.

[0041] Alternatively, the step of compacting the folds 15 can be implemented using a vacuum bag, for example under a pressure of -0.1 MPa.

[0042] As illustrated in Figure 1e, the process includes a sintering cycle application step to consolidate the bond between the composite material part 10 and the plies 15. The sintering cycle is carried out at a temperature between 1000°C and 1300°C. The sintering cycle has a duration of, for example, between 5 and 30 hours. If the ceramic matrix precursor includes a resin, the sintering cycle may include one or more intermediate steps allowing for pyrolysis and transformation of the resin, typically between 200 and 650°C, preferably between 250 and 550°C.

[0043] Under the influence of heat, the ceramic matrix grains of part 10 and the ceramic matrix grains of the plies 15 bond together, forming the cohesion between part 10 and the plies 15. During the sintering cycle, the organic components of the matrix precursor are burned off. At the end of the process, a repaired part 10 is obtained, having, including in the repaired area, the following composition: ceramic fibers at 45% by volume, ceramic matrix at 35% by volume, and porosity at 20% by volume.

[0044] Figure 3 is a photograph of a cross-sectional view obtained by X-ray tomography of a repaired area of ​​a part made of CMC material by the implementation of the process of Figures 1a to 1e. This photograph highlights the good material condition and the absence of delamination at the bond between the part 10 and the plies 15 after sintering.

[0045] A second implementation of the repair process for a part made of ceramic matrix composite material 10 having a damaged area 11 is described below with reference to figures 4a to 4e (see figure 4a).

[0046] As illustrated in Figure 4b, the process includes a machining removal step of at least the damaged area 11 to obtain a repair cavity 12. This step is analogous to that implemented by water jet in the process of Figures 1a-1e.

[0047] Following this embodiment, the repair recess 12, in cross-section, has a flared shape with a width that increases vertically from the bottom of the repair recess 12. Faces 20 delimiting the repair recess 12 have a sloping shape forming a non-zero angle with respect to the flat bottom of the repair recess 12. This configuration maximizes the contact area between a fold 15 and the surface of the repair area 12. It is of course possible to provide the same shape of the repair area 12 in the context of the first implementation of the process.

[0048] As illustrated in Figure 4c, the process includes a draping step of one or more first plies 15 of pre-impregnated composite material(s) of a ceramic matrix precursor against a surface of the hollow repair area 12 and a placement step of a filling element 16 made of a sintered ceramic matrix composite material having a shape complementary to the hollow repair area 12.

[0049] The process further includes a step of placing one or more second layers 15 of composite material pre-impregnated with a ceramic matrix precursor so as to encapsulate the filling element 16 between the first ply(ies) 15 of composite material pre-impregnated with a ceramic matrix precursor and the second ply(ies) 15 of composite material pre-impregnated with a pre-impregnated ceramic matrix precursor. The first and second ply(ies) 15 thus surround the filler element 16. The assembly formed by the filler element 16 and the two ply(ies) 15 constitutes the repair material 13.

[0050] As illustrated in Figure 4d, the process includes a step of covering the second ply(ies) 15 with composite material pre-impregnated with a ceramic matrix precursor. The covering step consists of placing one or more technical fabrics 17 over the second ply(ies) 15. The technical fabric(s) 17 may be of different types. Their role is to drain any excess matrix initially present in the pre-impregnated fabric, to allow solvent evacuation, and ultimately, to reduce the porosity of the intermediate ply(ies) 15.

[0051] The process includes a compaction step of 15 plies of pre-impregnated composite material of the ceramic matrix precursor to increase a volumetric fiber content.

[0052] After placing a membrane 18 around the technical fabric 17 and the assembly "composite part 10-repair material 13" arranged on a support 19, an autoclaving cycle is applied under a pressure between 0.5 and 2.5 MPa, preferably between 0.5 and 1.5 MPa, and a temperature between 50 and 250°C. The autoclaving cycle lasts between 5 and 10 hours. The autoclaving cycle aims to pre-cure the plies 15 to increase the fiber volume fraction in the plies 15 due to the partial removal of the organic components contained in the ceramic matrix precursor.

[0053] Alternatively, the step of compacting the folds 15 can be implemented using a vacuum bag, for example under a pressure of -0.1 MPa.

[0054] As illustrated in Figure 4e, the process includes a sintering cycle application step to consolidate the bond between the composite material part 10 and the repair material 13. The sintering cycle is carried out at a temperature between 1000°C and 1300°C. The sintering cycle lasts between 5 and 10 hours. If the ceramic matrix precursor includes a resin, the sintering cycle may include one or more intermediate steps allowing for pyrolysis and transformation of the resin, typically between 200 and 650°C, preferably between 250 and 550°C.

[0055] Under the influence of heat, the ceramic matrix grains of part 10 and the ceramic matrix grains of the plies 15 and the filler element 16 bond together, forming cohesion between part 10 and the repair material 13. During the sintering cycle, the organic components of the matrix precursor are burned off. At the end of the process, a repaired part 10 is obtained, having, including in the repaired area, the following composition: ceramic fibers at 45% by volume, ceramic matrix at 35% by volume, and porosity at 20% by volume.

[0056] Alternatively, the machining of the damaged area 11 can be carried out in a more conventional way using a machine tool.

[0057] Of course, the different features, variants and / or embodiments of the present invention can be combined with each other in various ways as long as they are not incompatible or mutually exclusive.

[0058] Furthermore, the invention is not limited to the embodiments described above and provided solely by way of example. It encompasses various modifications, alternative forms, and other variations that a person skilled in the art may envision within the scope of the present invention, and in particular all combinations of the different modes of operation described above, which may be considered separately or in combination.

Claims

DEMANDS 1. A method for repairing a part made of ceramic matrix composite material (10) having a damaged area (11), characterized in that said method comprises: - a machining removal step of at least the damaged area (11) so as to obtain a hollowed-out repair area (12), - a step of placing, in the hollowed-out repair area (12), a repair material (13) comprising at least one ply (15) of composite material pre-impregnated with a ceramic matrix precursor, - a compaction step of the ply(ies) (15) of composite material pre-impregnated with the ceramic matrix precursor to increase the fiber volume ratio, and - a step of applying a sintering cycle to consolidate a bond between the repair material (13) and the composite material part (10), - the step of setting up the repair material (13) comprising a step of draping at least one first ply (15) of composite material pre-impregnated with a ceramic matrix precursor against a surface of the hollowed repair area (12) and a step of setting up a filling element (16) made of a sintered ceramic matrix composite material having a shape complementary to the hollowed repair area (12).

2. Method according to claim 1, characterized in that it further comprises a step of placing at least a second ply (15) of composite material pre-impregnated with a ceramic matrix precursor so as to encapsulate the filling element (16) between the first ply (15) of composite material pre-impregnated with a ceramic matrix precursor and the second ply (15) of composite material pre-impregnated with a pre-impregnated ceramic matrix precursor.

3. Method according to claim 1 or 2, characterized in that the hollowed repair area (12) has, in cross-sectional view, a flared shape.

4. A method according to any one of claims 1 to 3, characterized in that the material removal step by machining is carried out by a water jet containing an abrasive medium such as sand.

5. Method according to claim 4, characterized in that the composite material part (10) being constituted by a plurality of sintered plies, water jet parameters are chosen so as to remove the material ply by ply in the damaged area (11) of the composite material part (10).

6. A method according to any one of claims 1 to 5, characterized in that the compaction step of the ply(15) includes a step of applying an autoclaving cycle to the assembly "part in composite material (10)-repair material (13)".

7. Method according to claim 6, characterized in that the autoclaving cycle is carried out at a temperature between 50°C and 250°C and at a pressure between 0.5 and 2.5 MPa.

8. A process according to any one of claims 1 to 7, characterized in that the sintering cycle is carried out at a temperature between 1000°C and 1300°C.

9. Repaired composite material part obtained by implementing the process defined according to any one of the preceding claims.

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