Aircraft engine silicon carbide ceramic matrix composite part patch repair method

Abstract

The application discloses a kind of aviation engine silicon carbide ceramic matrix composite material part patch repair methods, comprising: using CMC-SiC patch as the structure of patch repair part crack, corrosion pit or fiber crack, surface texture is prepared on the contact surface of CMC-SiC patch and part with pulse laser, brazing material is coated on the contact surface of CMC-SiC patch and part, CMC-SiC patch and part surface are pressed tightly, after vacuum brazing, polishing, cleaning and drying, environmental barrier coating recovery, annealing, the patch repair of aviation engine silicon carbide ceramic matrix composite material part is completed.The application adopts patch technology, reconstructs the force transmission path of crack, corrosion pit or fiber fracture, restores the strength of part, solves the repair problem of relatively wide crack, corrosion pit or fiber fracture and other more serious damage on silicon carbide ceramic matrix composite material part.In addition, specific texture, appearance is prepared on the contact surface of CMC-SiC patch and part with pulse laser, and the optimal combination effect and joint strength of brazing material and patch, part base material are accurately realized.

Description

Technical Field

[0001] This invention belongs to the field of aero-engine repair technology, and more specifically, this invention relates to a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. Background Technology

[0002] Silicon carbide ceramic matrix composites (CMC-SiC), especially silicon carbide ceramic fiber reinforced silicon carbide ceramic matrix composites (SiC). f The application of CMC-SiC in aero engines is increasing, and it has now been applied to hot-end stationary parts such as engine flame tubes, turbine outer rings, guide vanes, and tail nozzles. In order to prevent CMC-SiC materials from corroding with water vapor, oxygen, molten salts and other substances in the environment when operating at high temperatures, an environmental barrier coating (EBC) is often prepared on the surface of CMC-SiC parts. The most common EBC coating at present is the third-generation EBC, namely rare earth silicate system EBC.

[0003] Due to the characteristics of ceramic materials, there are limited repair methods for engine parts made of CMC-SiC that suffer from cracks, corrosion, or other damage, and conventional techniques such as argon arc welding are not applicable.

[0004] When the silicon carbide fiber is largely undamaged and the matrix only has narrow cracks less than 0.2 mm wide, brazing can be used to fill the matrix cracks, closing them and restoring some strength. The general method is as follows: Remove the environmental barrier coating from the part's surface, clean and dry the part; select a suitable powdered Cu-based, Ag-based, or Ni-based active metal brazing filler, add a binder, mix thoroughly into a paste, and apply it to the crack and surrounding area on the part's surface; place the part in a vacuum brazing furnace and perform vacuum brazing according to a specific heating, holding, and cooling curve. The molten brazing filler will enter from the part's surface under capillary action and fill the crack, reacting at a suitable temperature to form a welded joint, achieving the effect of closing the crack and restoring a certain strength; thermal spraying restores the environmental barrier coating, followed by annealing and other post-treatments. In other damage cases, such as wide cracks in the matrix (greater than 0.2 mm) or corrosion pits with large diameters (greater than 0.2 mm) in the matrix material, or cracks and fractures in both the matrix and silicon carbide fibers, parts with this type of damage can only be scrapped prematurely, resulting in an actual service life of the parts that is far shorter than the design value.

[0005] Currently, brazing repair technology is only suitable for repairing CMC-SiC parts with minor damage, that is, when the silicon carbide fibers are basically undamaged and the matrix only has narrow cracks with a width of less than 0.2 mm.

[0006] When the silicon carbide fiber is undamaged, but the matrix has wide cracks (greater than 0.2 mm) or the matrix material has large-diameter corrosion pits, the brazing filler metal cannot effectively fill the cracks or corrosion pits due to the large spacing between the matrix material in the cracks or corrosion pits. This prevents capillary action or the brazing filler metal in the middle of the brazed joint from reacting with the ceramic base material to form a connecting phase. As a result, the brazed area has extremely low strength and cannot meet the repair requirements.

[0007] When both the matrix and the silicon carbide fiber, which is the main load-bearing component, develop cracks or fractures, the force transmission path of the part is disrupted, and existing brazing repair methods cannot restore or create an equivalent force transmission path and load-bearing capacity.

[0008] Furthermore, when using existing brazing techniques to repair cracks less than 0.2 mm wide, the capillary action generated when the filler metal enters the crack is significantly affected by the original state of the crack's inner surface. Given the randomness of the crack's internal morphology and the inability to control or optimize the inner surface condition, optimal capillary action and bonding effects cannot be achieved during brazing. In other brazing cases where the contact surface is accessible, sanding and sandblasting are typically used to optimize the brazing joint surface and improve the brazing effect. However, the surface finish achieved by sanding or sandblasting is relatively coarse and cannot precisely achieve optimal surface roughness and texture, thus failing to obtain the best brazing bond performance. Summary of the Invention

[0009] One object of the present invention is to solve at least the above-mentioned problems and / or defects, and to provide at least the advantages described below.

[0010] To achieve these objectives and other advantages according to the present invention, a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines is provided, comprising: using patching technology, using CMC-SiC patches as patches to repair cracks, corrosion pits or fiber cracks in the parts, preparing surface textures on the contact surface between the CMC-SiC patch and the parts using pulsed laser, applying brazing filler metal to the contact surface between the CMC-SiC patch and the parts, then pressing the CMC-SiC patch and the parts surface together, and completing the patching and repair of the silicon carbide ceramic matrix composite parts for aero-engines after vacuum brazing, grinding, cleaning and drying, environmental barrier coating restoration, and annealing.

[0011] Preferably, the process specifically includes the following steps:

[0012] Step 1: Remove the environmental barrier coating from the surface of the parts;

[0013] Step 2: Fabricate CMC-SiC patches using melt impregnation RMI process, or fabricate CMC-SiC patches using CMC-SiC flat plates through machining.

[0014] Step 3: Using laser texturing technology, surface texture is prepared on the contact surface between the CMC-SiC patch and the part using a pulsed laser;

[0015] Step 4: Place the laser-textured CMC-SiC patch and parts in acetone for ultrasonic cleaning for 10-30 minutes, then dry them.

[0016] Step 5: Select the powder brazing filler metal;

[0017] Step 6: Dry the powdered brazing filler metal, then add the bonding agent. The mass ratio of powdered brazing filler metal to bonding agent is 1:1 to 5:1. Mix them evenly into a paste and apply it to the contact surface between the CMC-SiC patch and the part.

[0018] Step 7: Press the CMC-SiC patch firmly against the surface of the part, and use a weight or positioning clamp to ensure the gap;

[0019] Step 8: Place the parts into a vacuum brazing furnace for vacuum brazing. Heat the parts to the welding temperature at a rising rate, hold them at that temperature for a certain time, and finally cool them down to room temperature at a certain cooling rate or by cooling them with the furnace.

[0020] Step 9: Remove the brazed parts, grind the weld seam smooth, and ultrasonically clean it in acetone solution for 10-30 minutes before drying.

[0021] Step 10: Use atmospheric plasma spraying (APS) equipment to restore the environmental barrier coating on the parts;

[0022] Step 11: Polish the restored environmental barrier coating to make it meet the product roughness requirements;

[0023] Step 12: Anneal the parts to complete the patching and repair.

[0024] Preferably, in step one, the environmental barrier coating on the surface of the part is removed by laser cleaning, sandblasting or machining until the CMC-SiC substrate is exposed.

[0025] Preferably, in step two, the basic design principles for the shape and size of the CMC-SiC patch include: the shape of the CMC-SiC patch is circular or rectangular with rounded corners;

[0026] The thickness of the CMC-SiC patch is not less than the original thickness of the damaged part;

[0027] The direction of the silicon carbide fibers in the CMC-SiC patch is parallel to the direction of the fibers at the repair site of the part.

[0028] The minimum distance L from the edge of the CMC-SiC patch to the damaged edge of the part in a direction parallel to the silicon carbide fiber satisfies the following relationship:

[0029]

[0030] Where σ is the tensile strength of the CMC-SiC patch material, τ is the tensile shear strength of the lap joint, and H is the average thickness of the part at the damaged location; the edge of the CMC-SiC patch is at least 3 mm away from the damaged edge in the direction perpendicular to the fiber.

[0031] The CMC-SiC patch is placed on the damaged area of ​​the part to be repaired, and the gap between the CMC-SiC patch and the surface of the part is ensured to be no more than 0.2 mm.

[0032] Preferably, in step three, the pulsed laser is a femtosecond laser with a wavelength of 0.8–1.06 μm, a single pulse energy of 35 fs, a repetition frequency of 200–1000 Hz, and an energy density of 1.5–2.5 J / cm³. 2 The scanning speed is 50–350 μm / s, and the number of scans is 1–5.

[0033] Preferably, in step four, powdered Cu-based, Ag-based, Ni-based, Ti-Si, or Ni-Si-Ti eutectic solder is selected as the solder.

[0034] Preferably, step seven further includes controlling the solder thickness between the CMC-SiC patch and the component to not exceed 0.2 mm.

[0035] Preferably, in step eight, the welding temperature is 800–1400℃, the holding time is 5–50 min, the heating rate is 5–20℃ / min, and the cooling rate is 3–10℃ / min.

[0036] Preferably, in step ten, the specific method for restoring the environmental barrier coating of the part using an atmospheric plasma spraying (APS) device includes: placing the part to be sprayed on a turntable using a fixture, rotating the part with the turntable during spraying to ensure uniform environmental barrier coating; and preheating the part substrate with a plasma arc before spraying.

[0037] Preferably, in step twelve, the annealing temperature is 800–1300°C, the annealing holding time is half an hour, and the furnace is cooled.

[0038] This invention offers at least the following advantages: It employs patching technology to reconstruct the force transmission path at cracks, corrosion pits, or fiber breaks, restoring part strength and solving the challenge of repairing severe damage such as wide cracks, corrosion pits, or fiber breaks on silicon carbide ceramic matrix composite parts. Furthermore, pulsed lasers are used to prepare specific textures and morphologies on the bonding surfaces of the CMC-SiC patch and the part, precisely achieving optimal bonding and joint strength between the solder and the patch / part base material.

[0039] Other advantages, objectives and features of the present invention will become apparent in part from the following description, and in part from those skilled in the art through study and practice of the invention. Attached Figure Description

[0040] Figure 1 This is a schematic diagram of the CMC-SiC patch in Embodiment 1 of the present invention. Detailed Implementation

[0041] The present invention will now be described in further detail with reference to the accompanying drawings, so that those skilled in the art can implement it based on the description.

[0042] It should be understood that terms such as “having,” “comprising,” and “including” as used herein do not exclude the presence or addition of one or more other elements or combinations thereof.

[0043] Example 1

[0044] A method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines includes the following steps:

[0045] Step 1: Use laser cleaning technology (laser parameters are set based on coating composition and thickness) to remove the original environmental barrier coating, expose the CMC-SiC substrate, and grind the damaged areas smooth, removing burrs, edges, protrusions, etc.

[0046] Step 2: Repair the 0.2mm wide crack in the substrate. Design a CMC-SiC patch based on the shape and size of the crack and corrosion pit, as well as the part's shape, thickness, and silicon carbide fiber extension direction. Fabricate the CMC-SiC patch using the melt infiltration (RMI) process. See the schematic diagram of the CMC-SiC patch. Figure 1 , Figure 1 In the image, 1 represents the original crack in the part, and 2 represents the CMC-SiC patch. The basic design principles for the shape and dimensions of the CMC-SiC patch are as follows:

[0047] CMC-SiC patches are available in circular or rectangular shapes with rounded corners, such as... Figure 1 As shown, the CMC-SiC patch used in this embodiment is a rectangle with rounded corners;

[0048] The thickness of the CMC-SiC patch is not less than the original thickness of the damaged part, and the thickness of the damaged part is 2mm. The thickness of the CMC-SiC patch used in this embodiment is 2mm.

[0049] In CMC-SiC patches, the silicon carbide fibers are aligned parallel to the fiber direction at the repair site.

[0050] The minimum distance between the edge of the CMC-SiC patch and the edge of the damage along the parallel fiber direction. σ is the tensile strength of CMC-SiC material, which is 800MPa; τ is the tensile shear strength of the lap joint, which is 45MPa; H is the average thickness of the damaged part, which is 2mm; therefore, L is calculated to be 36mm. The edge of the CMC-SiC patch should be at least 3mm away from the edge of the damage in the direction perpendicular to the fiber.

[0051] Place the CMC-SiC patch on the damaged area to be repaired, ensuring that the gap between the CMC-SiC patch and the surface of the part is no greater than 0.2 mm.

[0052] Step 3: Laser texturing is employed. This involves using a pulsed laser to create specific surface textures on the contact surface between the CMC-SiC patch and the component. This optimizes the surface microstructure, increases the contact area and wettability of the molten solder during subsequent brazing, and ultimately improves the connection effect and strength between the CMC-SiC patch and the component. The laser texturing parameters need to be selected based on the properties of the solder used. The laser parameters are: pulsed laser wavelength of 1.06 μm, single pulse energy of 35 fs, repetition frequency of 500 Hz, and energy density of 2 J / cm³. 2 The scanning speed was 200 μm / s, and the number of scans was 3.

[0053] Step 4: After laser texturing, the CMC-SiC parts and CMC-SiC patches are placed in acetone solution for ultrasonic cleaning for 15 minutes and then dried.

[0054] Step 5: Select powdered Ag-based solder as the brazing filler metal;

[0055] Step 6: Dry the Ag-based powder solder, then add NiCrobraz II type water-based binder. The mass ratio of Ag-based powder solder to binder is 5:1. Mix evenly to form a paste and apply it to the contact surface between the patch and the part.

[0056] Step 7: Press the CMC-SiC patch firmly against the surface of the part, and use a weight or positioning clamp to ensure the gap, controlling the solder thickness between the CMC-SiC patch and the part to not exceed 0.2mm;

[0057] Step 8: Place the parts into a vacuum brazing furnace for vacuum brazing. The molten brazing filler metal will react with the ceramic base material at a suitable temperature to form a welded joint. The welding temperature is 890℃, which needs to be selected according to the specific brazing filler metal composition and the working temperature requirements of the parts. The holding time is 20 minutes, the heating rate is 10℃ / min, and the cooling rate is furnace cooling.

[0058] Step 9: Remove the brazed parts, grind the weld seam smooth, and ultrasonically clean it in acetone solution for 15 minutes before drying.

[0059] Step 10: Use atmospheric plasma spraying (APS) equipment to restore the environmental barrier coating. Use a fixture to place the parts to be sprayed on a turntable. During spraying, the parts rotate with the turntable to ensure uniform coating. Before spraying, preheat the substrate of the parts with a plasma arc.

[0060] Step 11: Polish the restored environmental barrier coating to meet the roughness requirements of the product.

[0061] Step 12: Perform annealing heat treatment at 800℃, hold for half an hour, and then cool with the furnace to reduce residual stress.

[0062] Step 13: Complete the repair so that the part can continue to be used and avoid premature scrapping.

[0063] Example 2

[0064] This embodiment provides a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. Compared with Embodiment 1, the difference is that in step two, a CMC-SiC plate of the required thickness is selected and then a CMC-SiC patch is made by machining. The process parameters of the remaining steps in this embodiment are the same as those in Embodiment 1.

[0065] Example 3

[0066] This embodiment provides a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. Compared with Embodiment 1, the difference is that powdered Cu-based material is used as the brazing filler in step five, and the brazing temperature is about 800°C. The process parameters of the remaining steps in this embodiment are the same as those in Embodiment 1.

[0067] Example 4

[0068] This embodiment provides a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. Compared with Embodiment 1, the difference is that powdered Ni-based material is used as the brazing filler metal in step five, and the brazing temperature is about 1400℃. The process parameters of the remaining steps in this embodiment are the same as those in Embodiment 1.

[0069] Example 5

[0070] This embodiment provides a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. Compared with Embodiment 1, the difference is that powdered Ti-Si is used as the solder in step five. The process parameters of the remaining steps in this embodiment are the same as those in Embodiment 1.

[0071] Example 6

[0072] This embodiment provides a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. Compared with Embodiment 1, the difference is that powdered Ni-Si-Ti eutectic is used as the solder in step five. The process parameters of the remaining steps in this embodiment are the same as those in Embodiment 1.

[0073] Comparative Example 1

[0074] This comparative example provides a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. Compared with Example 1, the difference is that in step three, laser texturing technology is not used to prepare surface texture on the contact surface between the CMC-SiC patch and the part. Instead, the contact surface between the CMC-SiC patch and the part is polished one by one with 400-grit, 800-grit, and 1000-grit sandpaper, and then vacuum brazing is performed. The process parameters of the remaining steps are the same as those in Example 1.

[0075] Comparative Example 2

[0076] This comparative example provides a method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines. The difference from Example 1 is that in step three, instead of using laser texturing to create surface texture on the contact surface between the CMC-SiC patch and the part, sandblasting is performed on the contact surface, followed by vacuum brazing. The process parameters for the remaining steps are the same as in Example 1.

[0077] In Example 1, the tensile shear strength of the joint after laser roughening and brazing reached 47 MPa, which is 38% higher than the tensile shear strength of 34 MPa of the joint after sanding and brazing in Comparative Example 1.

[0078] The number of devices and processing scale described herein are for the purpose of simplifying the description of the invention. Applications, modifications, and variations of the invention will be readily apparent to those skilled in the art.

[0079] Although embodiments of the present invention have been disclosed above, they are not limited to the applications listed in the specification and embodiments. They can be applied to various fields suitable for the present invention. For those skilled in the art, other modifications can be easily made. Therefore, without departing from the general concept defined by the claims and their equivalents, the present invention is not limited to the specific details and illustrations shown and described herein.

Claims

1. A method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines, characterized in that, include: The patching technique uses CMC-SiC patches to repair cracks, corrosion pits, or fiber cracks in parts. A pulsed laser is used to create a surface texture on the contact surface between the CMC-SiC patch and the part. Solder is applied to the contact surface between the CMC-SiC patch and the part. Then, the CMC-SiC patch and the part surface are pressed together. After vacuum brazing, grinding, cleaning and drying, environmental barrier coating restoration, and annealing, the patching and repair of silicon carbide ceramic matrix composite parts for aero-engines is completed. The direction of the silicon carbide fibers in the CMC-SiC patch is parallel to the direction of the fibers at the repair site of the part. The minimum distance between the edge of the CMC-SiC patch and the edge of the damaged part in a direction parallel to the silicon carbide fiber. L The following relationship must be satisfied: in, σ For the tensile strength of CMC-SiC patch material, τ The tensile shear strength of the lap joint. H The average thickness of the part at the damaged site; the edge of the CMC-SiC patch is at least 3 mm away from the edge of the damage in the direction perpendicular to the fiber; The pulsed laser is a femtosecond laser with a wavelength of 0.8–1.06 μm, a single pulse energy of 35 fs, a repetition rate of 200–1000 Hz, and an energy density of 1.5–2.5 J / cm³. 2 The scanning speed is 50~350μm / s, and the number of scans is 1~5.

2. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 1, characterized in that, Specifically, the following steps are included: Step 1: Remove the environmental barrier coating from the surface of the parts; Step 2: Fabricate CMC-SiC patches using melt impregnation RMI process, or fabricate CMC-SiC patches using CMC-SiC flat plates through machining. Step 3: Using laser texturing technology, surface texture is prepared on the contact surface between the CMC-SiC patch and the part using a pulsed laser; Step 4: Place the laser-textured CMC-SiC patch and parts in acetone for ultrasonic cleaning for 10-30 minutes, then dry them. Step 5: Select the powder brazing filler metal; Step 6: Dry the powdered brazing filler metal, then add the bonding agent. The mass ratio of powdered brazing filler metal to bonding agent is 1:1 to 5:

1. Mix evenly to form a paste, and apply it to the contact surface between the CMC-SiC patch and the component. The bonding agent is one of NiCrobraz II type water-based adhesive, TENSOL No. 6, trichloroethylene, polystyrene, trichloroethylene, ethylene glycol, and glycerol. Step 7: Press the CMC-SiC patch firmly against the surface of the part, and use a weight or positioning clamp to ensure the gap; Step 8: Place the parts into a vacuum brazing furnace for vacuum brazing. Heat the parts to the welding temperature at a rising rate, hold them at that temperature for a certain time, and finally cool them down to room temperature at a certain cooling rate or by cooling them with the furnace. Step 9: Remove the brazed parts, grind the weld seam smooth, and ultrasonically clean it in acetone solution for 10-30 minutes before drying. Step 10: Use atmospheric plasma spraying (APS) equipment to restore the environmental barrier coating on the parts; Step 11: Polish the restored environmental barrier coating to make it meet the product roughness requirements; Step 12: Anneal the parts to complete the patching and repair.

3. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 2, characterized in that, In step one, the environmental barrier coating on the surface of the part is removed by laser cleaning, sandblasting or machining until the CMC-SiC substrate is exposed.

4. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 2, characterized in that, In step two, the basic design principles for the shape and size of the CMC-SiC patch include: the shape of the CMC-SiC patch is circular or rectangular with rounded corners; The thickness of the CMC-SiC patch is not less than the original thickness of the damaged part; The CMC-SiC patch is placed on the damaged area of ​​the part to be repaired, and the gap between the CMC-SiC patch and the surface of the part is ensured to be no more than 0.2 mm.

5. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 2, characterized in that, In step five, powdered Cu-based, Ag-based, Ni-based, Ti-Si, or Ni-Si-Ti eutectic solders are selected.

6. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 2, characterized in that, Step seven also includes controlling the solder thickness between the CMC-SiC patch and the component to not exceed 0.2 mm.

7. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 2, characterized in that, In step eight, the welding temperature is 800~1400℃, the holding time is 5~50min, the heating rate is 5~20℃ / min, and the cooling rate is 3~10℃ / min.

8. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 2, characterized in that, In step ten, the specific method for restoring the environmental barrier coating of the parts using atmospheric plasma spraying APS equipment includes: placing the parts to be sprayed on a turntable using a fixture, rotating the parts with the turntable during spraying to ensure uniform environmental barrier coating; and preheating the part substrate with a plasma arc before spraying.

9. The method for patching and repairing silicon carbide ceramic matrix composite parts for aero-engines as described in claim 2, characterized in that, In step 12, the annealing temperature is 800~1300℃, the annealing holding time is half an hour, and the furnace is cooled.

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