Method for patching and repairing silicon carbide ceramic matrix composite part of aero-engine

Abstract

The invention discloses a method for patching and repairing a silicon carbide ceramic matrix composite part of an aero-engine, which comprises the following steps: using a CMC-SiC patch as a structure for patching and repairing part cracks, corrosion pits or fiber cracks, preparing surface textures on the contact surface of the CMC-SiC patch and the part by using pulse laser, coating brazing filler metal on the contact surface of the CMC-SiC patch and the part, and preparing the surface textures on the contact surface of the CMC-SiC patch and the part. And the CMC-SiC patch and the surface of the part are pressed tightly, and patching and repairing of the silicon carbide ceramic matrix composite part of the aero-engine are completed after vacuum brazing, polishing, cleaning and drying, environmental barrier coating recovery and annealing. According to the repairing method, a patching technology is adopted, a force transmission path at cracks, corrosion pits or fiber fractures is rebuilt, the strength of the part is recovered, and the repairing problem of relatively serious damage such as relatively wide cracks, corrosion pits or fiber fractures on the silicon carbide ceramic-based composite part is solved. In addition, specific textures and shapes are prepared on the binding face of the CMC-SiC patch and the part through pulse laser, and the optimal combination effect and joint strength of the brazing filler metal, the patch and part base metal are accurately achieved.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft engine repair, and more specifically, relates to a method for patching and repairing a silicon carbide ceramic-based composite material part of an aircraft engine. Background Art

[0002] Silicon carbide ceramic matrix composites (CMC-SiC), especially silicon carbide ceramic fiber reinforced silicon carbide ceramic matrix composites SiC f / SiC is increasingly used in aero-engines, and has been used in 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 working at high temperatures, an environmental barrier coating (EBC) is often prepared on the surface of CMC-SiC parts. Currently, the most common EBC coating is the third-generation EBC, namely the 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 when they are damaged by cracks, corrosion, etc. Conventional argon arc welding and other technologies are not applicable.

[0004] When the silicon carbide fiber is basically intact and the matrix only has narrow cracks with a width of less than 0.2mm, brazing can be used to fill the matrix cracks, close the cracks and restore some strength. The general method is: remove the environmental barrier coating on the surface of the parts, clean the parts and dry them; select suitable powdered Cu-based, Ag-based or Ni-based active metal brazing filler metals, then add a connector, mix them evenly into a paste, and apply them to the cracks on the surface of the parts and nearby areas; put the parts into a vacuum brazing furnace and perform vacuum brazing according to a certain heating, insulation and cooling curve. The molten brazing filler metal will enter from the surface of the parts under capillary action and fill the cracks inside, and react at the appropriate temperature to form a welded joint, so as to close the cracks and restore a certain strength; thermal spray to restore the environmental barrier coating, and perform post-treatment such as annealing. In other damage cases, such as wide matrix cracks (greater than 0.2mm) or corrosion pits with large diameters (more than 0.2mm) formed in the matrix material, or cracks and fractures in both the matrix and the silicon carbide fiber, such damaged parts can only be scrapped in advance, making the actual service time of the parts far lower than the design value.

[0005] At present, brazing repair technology is only suitable for repairing slightly damaged CMC-SiC parts, that is, the silicon carbide fibers are basically intact and the matrix only has narrow cracks with a width of less than 0.2 mm.

[0006] When the silicon carbide fiber is not damaged, but there are wide cracks (greater than 0.2mm) in the matrix or corrosion pits with large diameters (more than 0.2mm) are formed in the matrix material, after the brazing filler metal is melted, the distance between the matrix materials in the cracks and corrosion pits is too large to produce capillary action, or the brazing filler metal in the middle of the brazing seam is too far away from the ceramic base material to react with each other to produce a connecting phase, and ultimately the cracks or corrosion pits cannot be effectively filled, and the strength of the brazing area is extremely low, which cannot meet the repair requirements.

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

[0008] In addition, when the existing brazing technology repairs cracks with a width of less than 0.2mm, the capillary effect generated when the brazing material enters the crack is greatly affected by the original state of the inner surface of the crack, and has a certain randomness according to the internal morphology of the crack. In addition, it is impossible to control and optimize the state of the inner surface of the crack, so it is impossible to produce a better capillary effect and connection effect during brazing. In other brazing situations where the contact surface is accessible, sandpaper grinding and sand blowing are usually used to optimize the brazing connection surface and improve the brazing effect. However, the surface state formed by sandpaper grinding or sand blowing is relatively rough, and it is impossible to accurately achieve the optimal surface roughness and surface texture, so it is impossible to obtain the optimal brazing bonding performance. Summary of the invention

[0009] An object of the present invention is to solve at least the above problems and / or disadvantages and to provide at least the advantages which will be described hereinafter.

[0010] In order to achieve these purposes and other advantages according to the present invention, a method for patching and repairing a silicon carbide ceramic matrix composite part of an aircraft engine is provided, comprising: adopting a patching technology, using a CMC-SiC patch as a structure for patching and repairing cracks, corrosion pits or fiber cracks in a part, and using a pulsed laser to prepare a surface texture on the contact surface between the CMC-SiC patch and the part, applying a brazing material on the contact surface between the CMC-SiC patch and the part, and then pressing the CMC-SiC patch and the surface of the part, and completing the patching and repair of the silicon carbide ceramic matrix composite part of the aircraft engine after vacuum brazing, grinding, cleaning and drying, environmental barrier coating restoration, and annealing.

[0011] Preferably, the method specifically comprises the following steps:

[0012] Step 1: Remove the environmental barrier coating on the surface of the part;

[0013] Step 2: using a melt infusion RMI process to make a CMC-SiC patch, or using a CMC-SiC flat plate to make a CMC-SiC patch by machining;

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

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

[0016] Step 5: Select powder solder;

[0017] Step 6: Dry the powder solder, then add the connecting agent, the mass ratio of the powder solder to the connecting agent is 1:1 to 5:1, mix evenly into a paste, and apply it on the contact surface between the CMC-SiC patch and the part;

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

[0019] Step 8: Place the parts in a vacuum brazing furnace for vacuum brazing, heat them to the welding temperature at a heating rate, then keep them warm for a certain period of time, and finally cool them down to room temperature at a certain cooling rate or by cooling them with the furnace;

[0020] Step 9: Take out the brazed parts, grind the welds to make them smooth, ultrasonically clean them in acetone solution for 10 to 30 minutes, and then dry them;

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

[0022] Step 11: Grind the restored environmental barrier coating to make the environmental barrier coating meet the roughness requirements of the composite product;

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

[0024] Preferably, in the step 1, the environmental barrier coating on the surface of the part is removed by laser cleaning, sandblasting or mechanical processing until the CMC-SiC matrix is ​​exposed.

[0025] Preferably, in the step 2, the basic design principles of 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 fiber direction of the repaired part;

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

[0029]

[0030] Wherein, 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 damaged part; 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 at the damaged part 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 greater than 0.2 mm.

[0032] Preferably, in step 3, the pulse laser is a femtosecond laser, the pulse laser wavelength is 0.8-1.06 μm, the single pulse energy is 35 fs, the repetition frequency is 200-1000 Hz, and the energy density is 1.5-2.5 J / cm 2 , scanning speed is 50~350μm / s, and the number of scans is 1~5 times.

[0033] Preferably, in step 4, powdered Cu-based, Ag-based, Ni-based, Ti-Si or Ni-Si-Ti eutectic is selected as the solder.

[0034] Preferably, the step seven further comprises: 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° C., the holding time is 5-50 min, the heating rate is 5-20° C. / min, and the cooling rate is 3-10° C. / min.

[0036] Preferably, in step ten, the specific method of using atmospheric plasma spraying APS equipment to restore the environmental barrier coating of parts includes: using a fixture to place the parts to be sprayed on a turntable, and the parts rotate with the turntable during spraying to ensure that the environmental barrier coating is uniform; and preheating the part substrate by a plasma arc before spraying.

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

[0038] The present invention has at least the following beneficial effects: The present invention adopts patching technology to rebuild the force transmission path of cracks, corrosion pits or fiber breaks, restore the strength of parts, and solve the problem of repairing wider cracks, corrosion pits or fiber breaks on silicon carbide ceramic-based composite parts. In addition, pulsed laser is used to prepare specific textures and morphologies on the CMC-SiC patch and part bonding surface, accurately achieving the optimal bonding effect and joint strength of the brazing material, patch, and part base material.

[0039] Other advantages, objectives and features of the present invention will be embodied in part through the following description, and in part will be understood by those skilled in the art through study and practice of the present invention. BRIEF DESCRIPTION OF THE DRAWINGS

[0040] Figure 1 Schematic diagram of the CMC-SiC patch in Example 1 of the present invention. DETAILED DESCRIPTION

[0041] The present invention is further described in detail below in conjunction with the accompanying drawings so that those skilled in the art can implement the invention with reference to the description.

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

[0043] Example 1

[0044] A method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine comprises 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 polish the damaged area to remove burrs, flanges, protrusions, etc.

[0046] Step 2: Repair the crack size of the substrate with a width of 0.2 mm. Design the CMC-SiC patch according to the shape and size of the crack and corrosion pit, the shape and thickness of the part at that location, and the extension direction of the silicon carbide fiber; use the melt infiltration RMI process to make the CMC-SiC patch. See the schematic diagram of the CMC-SiC patch for details. Figure 1 , Figure 1 Figure 1 is the original crack of the part, and Figure 2 is the CMC-SiC patch. The basic design principles for the shape and size of the CMC-SiC patch are as follows:

[0047] The CMC-SiC patch shape is circular or rectangular 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. The thickness of the damaged part is 2 mm. The thickness of the CMC-SiC patch used in this embodiment is 2 mm.

[0049] The direction of silicon carbide fibers in the CMC-SiC patch is parallel to the fiber direction of the repaired part

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

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

[0052] Step 3: Use laser texturing technology, that is, use pulsed laser to prepare specific surface texture on the contact surface between CMC-SiC patch and part, optimize the surface microstructure, increase the contact area and wettability of molten brazing material during subsequent brazing between CMC-SiC patch and part surface, and thus improve the connection effect and connection strength between CMC-SiC patch and part. Laser texturing parameters need to be selected according to the performance of the brazing material used. The laser parameters are: pulsed laser wavelength is 1.06μm, single pulse energy is 35fs, repetition frequency is 500Hz, and energy density is 2J / cm 2 , scanning speed 200 μm / s, scanning times 3 times;

[0053] Step 4: The laser-textured CMC-SiC parts and CMC-SiC patches are ultrasonically cleaned in an acetone solution for 15 minutes and then dried;

[0054] Step 5: Select powdered Ag as solder;

[0055] Step 6: Dry the Ag-based solder powder, then add NiCrobraz II water-based bonding agent, the mass ratio of Ag-based solder powder to bonding agent is 5:1, mix evenly into a paste, and apply it on the contact surface between the patch and the part;

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

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

[0058] Step 9: Take out the brazed parts, grind the welds to make them smooth, ultrasonically clean them in acetone solution for 15 minutes, and then dry them;

[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 part substrate with a plasma arc.

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

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

[0062] Step 13: Complete the repair so that the parts can continue to be used, avoiding premature scrapping of the parts.

[0063] Example 2

[0064] This embodiment provides a method for patching and repairing silicon carbide ceramic matrix composite parts of aircraft engines. Compared with Example 1, the difference is that in step 2, a CMC-SiC flat plate of the required thickness is purchased, and then the CMC-SiC patch is made by mechanical processing. The process parameters of the remaining steps of this embodiment are the same as those of Example 1.

[0065] Example 3

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

[0067] Example 4

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

[0069] Example 5

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

[0071] Example 6

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

[0073] Comparative Example 1

[0074] This comparative example provides a method for patching and repairing a silicon carbide ceramic matrix composite part of an aircraft engine. Compared with Example 1, the difference is that in step three, laser roughening technology is not used to prepare the surface texture on the contact surface between the CMC-SiC patch and the part. Instead, 400 mesh, 800 mesh, and 1000 mesh sandpaper are used at a time to polish the contact surface between the CMC-SiC patch and the part, 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 a silicon carbide ceramic matrix composite part of an aircraft engine. Compared with Example 1, the difference is that in step 3, laser texturing technology is not used to prepare a surface texture on the contact surface between the CMC-SiC patch and the part, but sand blowing is performed on the contact surface between the CMC-SiC patch and the part, and then vacuum brazing is performed. The process parameters of the remaining steps are the same as those of Example 1.

[0077] The tensile shear strength of the joint brazed after laser roughening in Example 1 reached 47 MPa, which was 38% higher than the tensile shear strength of 34 MPa of the joint brazed after sandpaper polishing in Comparative Example 1.

[0078] The number of devices and processing scales described here are used to simplify the description of the present invention. Applications, modifications and variations of the present invention will be obvious to those skilled in the art.

[0079] Although the embodiments of the present invention have been disclosed as above, they are not limited to the applications listed in the specification and the implementation modes, and they can be fully applied to various fields suitable for the present invention. For those familiar with the art, additional modifications can be easily implemented. Therefore, without departing from the general concept defined by the claims and the scope of equivalents, the present invention is not limited to the specific details and the illustrations shown and described herein.

Claims

1. A method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine, characterized in that: include: Patching technology is adopted, and CMC-SiC patches are used as the structure for patching and repairing cracks, corrosion pits or fiber cracks in parts. The surface texture is prepared on the contact surface between the CMC-SiC patch and the part with a pulsed laser. The brazing material is applied on the contact surface between the CMC-SiC patch and the part, and then the CMC-SiC patch and the surface of the part are pressed tightly. After vacuum brazing, grinding, cleaning and drying, environmental barrier coating restoration and annealing, the patching and repair of silicon carbide ceramic-based composite parts of aircraft engines are completed.

2. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 1, characterized in that: The specific steps include: Step 1: Remove the environmental barrier coating on the surface of the part; Step 2: using a melt infusion RMI process to make a CMC-SiC patch, or using a CMC-SiC flat plate to make a CMC-SiC patch by machining; Step 3: Using laser texturing technology, a pulsed laser is used to prepare a surface texture on the contact surface between the CMC-SiC patch and the part; Step 4: Place the laser-textured CMC-SiC patch and parts in acetone for ultrasonic cleaning for 10 to 30 minutes and then dry; Step 5: Select powder solder; Step 6: Dry the powder solder, then add a connecting agent, the mass ratio of the powder solder to the connecting agent is 1:1 to 5:1, mix evenly into a paste, and apply it on the contact surface between the CMC-SiC patch and the part; the connecting 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 against the surface of the part and use a heavy object or a positioning clamp to ensure the gap; Step 8: Place the parts in a vacuum brazing furnace for vacuum brazing, heat them to the welding temperature at a heating rate, then keep them warm for a certain period of time, and finally cool them down to room temperature at a certain cooling rate or by cooling them with the furnace; Step 9: Take out the brazed parts, grind the welds to make them smooth, ultrasonically clean them in acetone solution for 10 to 30 minutes, and then dry them; Step 10: Use atmospheric plasma spraying APS equipment to restore the environmental barrier coating of the parts; Step 11: Grind the restored environmental barrier coating to make the environmental barrier coating meet the roughness requirements of the composite product; Step 12: Anneal the parts to complete the patching and repair of the parts.

3. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 2, characterized in that: In the step 1, the environmental barrier coating on the surface of the part is removed by laser cleaning, sandblasting or mechanical processing until the CMC-SiC matrix is ​​exposed.

4. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 2, characterized in that: In the step 2, the basic design principles of 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 direction of the silicon carbide fibers in the CMC-SiC patch is parallel to the fiber direction of the repaired part; The minimum distance L between the edge of the CMC-SiC patch and the damaged edge of the part in a direction parallel to the silicon carbide fiber satisfies the following relationship: Wherein, 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 damaged part; the edge of the CMC-SiC patch is at least 3 mm away from the damaged edge in the direction perpendicular to the fiber; The CMC-SiC patch is placed at the damaged part 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 greater than 0.2 mm.

5. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 2, characterized in that: In step 3, the pulse laser is a femtosecond laser with a pulse laser wavelength of 0.8 to 1.06 μm, a single pulse energy of 35 fs, a repetition frequency of 200 to 1000 Hz, and an energy density of 1.5 to 2.5 J / cm 2 , scanning speed is 50~350μm / s, and the number of scans is 1~5 times.

6. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine as claimed in claim 2, characterized in that: In the step 4, powdered Cu-based, Ag-based, Ni-based, Ti-Si or Ni-Si-Ti eutectic is selected as the solder.

7. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine as claimed in claim 2, characterized in that: The step seven also includes: controlling the solder thickness between the CMC-SiC patch and the component to not exceed 0.2 mm.

8. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine as claimed in claim 2, characterized in that: In the step eight, the welding temperature is 800-1400° C., the holding time is 5-50 min, the heating rate is 5-20° C. / min, and the cooling rate is 3-10° C. / min.

9. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine as claimed in claim 2, characterized in that: In the step ten, the specific method of using atmospheric plasma spraying APS equipment to restore the environmental barrier coating of the parts includes: using a fixture to place the parts to be sprayed on a turntable, and the parts rotate with the turntable during spraying to ensure that the environmental barrier coating is uniform; before spraying, the part substrate is preheated by a plasma arc.

10. The method for patching and repairing a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 2, characterized in that: In the step 12, the annealing temperature is 800-1300° C., the annealing holding time is half an hour, and the furnace is cooled.

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