Method for repairing crack damage of silicon carbide ceramic matrix composite part of aero-engine

Abstract

The invention discloses an aero-engine silicon carbide ceramic matrix composite part crack damage repairing method which comprises the following steps: removing an environmental barrier coating on the surface of a part, cleaning and drying; powdery active metal brazing filler metal and a connecting agent are selected to be mixed and then coated on cracks and nearby areas on the surface of the part; vacuum brazing, weld joint polishing, cleaning and drying are conducted; preparing a bonding layer and a middle layer of the environmental barrier coating by using atmospheric plasma spraying equipment, and preparing a surface layer of the environmental barrier coating by using a plasma-physical vapor deposition method; and the environmental barrier coating is polished, annealing treatment is conducted, and crack damage repairing of the silicon carbide ceramic matrix composite part of the aero-engine is completed. According to the vacuum brazing process, the strength of the base material can be effectively recovered by 50%-60%, the usable time of the base material is prolonged, parts can continue to serve, and the parts are prevented from being scrapped in advance; in addition, according to the EBC coating preparation technology, the porosity of the coating can be reduced, and therefore the service time of the repaired part is further prolonged.

Description

Technical Field

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

[0002] Compared with the nickel-based high-temperature alloy materials that are currently widely used, silicon carbide ceramic-based composites (CMC-SiC) can withstand higher temperatures. More importantly, their density is only 1 / 4 to 1 / 3 of that of high-temperature alloys. They can greatly reduce the weight of the engine structure while ensuring high-temperature performance. They have now been used in hot-end stationary parts such as engine flame tubes, turbine outer rings, guide vanes, and tail nozzles.

[0003] In order to prevent CMC-SiC materials from corroding with water vapor, oxygen, molten salt 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. Typical CMC-SiC parts with EBC coatings are as follows: Figure 1 As shown. The damage forms of CMC-SiC parts with EBC coating include coating damage and substrate damage. After long-term high-temperature use, the EBC coating is prone to cracks or even peeling. The CMC-SiC ceramic substrate will lose coating protection and suffer from the combined effects of engine gas thermal shock, thermal alternating stress, corrosive atmosphere and other factors, resulting in cracks and other damage. If not repaired, the part cannot continue to be used. CMC-SiC materials and engine parts made of CMC-SiC are expensive, but there is still a lack of suitable repair methods. After the substrate of the part is damaged, it can only be scrapped in advance, making the actual service time of the part far lower than the design value, causing huge material waste and economic losses.

[0004] The reason why cracks in CMC-SiC materials are difficult to repair is that the melting temperature of ceramic materials is high and the diffusion coefficient of elements is low. It is difficult to form high-strength connection joints like metals by heating up and melting-cooling and solidifying (such as argon arc welding) or element mutual diffusion (diffusion welding). If high-energy beams such as lasers are used to irradiate CMC-SiC ceramic-based composites, too low energy will have no effect on CMC-SiC, and too high energy will easily cause the SiC matrix material to sublimate and volatilize directly in an instant, which will cause greater damage to the parts.

[0005] In addition, when repairing the cracks in the CMC-SiC part matrix, the removal of the original EBC coating on the surface of the part must also be considered, otherwise it will affect the crack repair on the matrix material CMC-SiC. After the cracks in the CMC-SiC matrix material are repaired, the previously removed EBC coating must be restored to continue to protect the CMC-SiC matrix from damage by the corrosive atmosphere in the environment during the subsequent work of the part. The EBC coating includes a bonding layer (such as silicon Si), an intermediate layer (such as mullite) and a surface layer (such as ytterbium silicate Yb 2 SiO 5 ), currently most of them are prepared by atmospheric plasma spraying APS and other means. Due to the principle of the equipment itself, a certain amount of pores are inevitably present in the coating, which weakens the protective effect of the EBC coating. Summary of the invention

[0006] 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.

[0007] In order to achieve these purposes and other advantages of the present invention, a method for repairing crack damage in silicon carbide ceramic-based composite parts of aircraft engines is provided, comprising: removing the environmental barrier coating on the surface of the part, cleaning and drying; selecting powdered active metal brazing filler metal and a connecting agent to mix and then coating the surface cracks and nearby areas of the part; vacuum brazing, grinding the weld, cleaning and drying; using atmospheric plasma spraying equipment to prepare the bonding layer and intermediate layer of the environmental barrier coating, and using a plasma-physical vapor deposition method to prepare the surface layer of the environmental barrier coating; grinding the environmental barrier coating and performing annealing treatment to complete the crack damage repair of the silicon carbide ceramic-based composite parts of the aircraft engine.

[0008] Preferably, the method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aero-engine specifically comprises the following steps:

[0009] Step 1: Use laser cleaning technology to remove the environmental barrier coating on the surface of the part to expose the part substrate;

[0010] Step 2: Place the parts in an acetone solution for ultrasonic cleaning for 10 to 30 minutes and then dry;

[0011] Step 3: Dry the powdered active metal brazing filler metal, add the bonding agent, mix evenly into a paste, and apply it on the cracks on the surface of the parts and around 10 to 30 mm along the crack direction. 2 Area region;

[0012] Step 4: Place the parts into a vacuum brazing furnace for vacuum brazing;

[0013] Step 5: 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;

[0014] Step 6: Using atmospheric plasma spraying equipment to prepare the bonding layer and the intermediate layer of the environmental barrier coating;

[0015] Step 7: preparing a surface layer of the environmental barrier coating by using a plasma-physical vapor deposition method;

[0016] Step 8: Grind the restored environmental barrier coating to make it meet the roughness requirements of the product;

[0017] Step 9: Anneal the parts to complete the repair.

[0018] Preferably, in step one, the specific process parameters of laser cleaning are: the laser is a pulsed nanosecond laser, the laser wavelength is 1.064 μm, the power is 300-500 W, the spot diameter is 0.5-2 mm, the single pulse width is 50-150 ns, the repetition frequency is 10-30 kHz, the scanning speed is 300-500 mm / s, and the overlap rate is 95-99%.

[0019] Preferably, in the step three, the powdered active metal brazing filler metal is one of Ag-Cu-Ti brazing filler metal, Ni-based brazing filler metal, and Cu-based brazing filler metal; the connecting agent is one of NiCrobrazⅡ type water-based adhesive, TENSOL No. 6 and trichloroethylene, polystyrene and trichloroethylene, ethylene glycol, and propylene glycol; and the mass ratio of the powdered active metal brazing filler metal to the connecting agent is 3 to 6:1.

[0020] Preferably, in the step four, the specific conditions of vacuum brazing are: heating to 200-220°C at a heating rate of 5-10°C / min and keeping warm for 10-20 min, heating to 750-1350°C at a heating rate of 8-10°C / min and keeping warm for 10-20 min, heating to 800-1400°C at a heating rate of 5-10°C / min and keeping warm for 20-30 min, cooling to 600-700°C at a cooling rate of 3-5°C / min, and opening the furnace after cooling to below 70°C.

[0021] Preferably, in step six, the bonding layer is a silicon layer with a thickness of 20 to 50 μm; the intermediate layer is a mullite layer with a thickness of 30 to 80 μm; when preparing the coating, a fixture is used to place the parts to be sprayed on a turntable of the atmospheric plasma spraying equipment. During spraying, the parts rotate with the turntable to ensure uniform coating, and the part substrate is preheated by a plasma arc before spraying.

[0022] Preferably, the atmospheric plasma spraying process parameters for preparing the bonding layer are: argon flow rate 40-60L / min, hydrogen flow rate 4-6L / min, current 500-600A, spraying distance 100-110mm, power 26-27kW, powder feeding rate 4-6r / min, carrier gas flow rate 8-10L / min, turntable speed 40-70r / min, substrate preheating 550-650°C; the atmospheric plasma spraying process parameters for preparing the intermediate layer are: argon flow rate 35-45L / min, hydrogen flow rate 8-12L / min, current 600-700A, spraying distance 100-110mm, power 33-34kW, powder feeding rate 2-4r / min, carrier gas flow rate 8-10L / min, turntable speed 40-70r / min, substrate preheating 750-850°C.

[0023] Preferably, in step seven, the surface layer is Yb 2 SiO 5 layer with a thickness of 50 to 120 μm; when preparing the coating, use a fixture to place the part to be sprayed on the turntable in the vacuum tank of the plasma-physical vapor deposition equipment. During spraying, the part rotates with the turntable. The pressure in the vacuum tank is first pumped to 100 to 200 Pa, and then filled with argon to 3000 to 5000 Pa. The part substrate is preheated by a plasma arc before spraying.

[0024] Preferably, the plasma-physical vapor deposition process parameters for preparing the surface layer are: argon flow rate 50-70 L / min, hydrogen flow rate 25-35 L / min, current 2000-3000 A, spraying distance 1000-1100 mm, power 35-45 kW, powder feeding rate 4-6 r / min, carrier gas flow rate 8-10 L / min, turntable speed 40-70 r / min, substrate preheating 550-650 ° C.

[0025] Preferably, in step nine, the annealing treatment temperature is 800-1000° C., the heating rate is 3-10° C. / min, the holding time is 20-40 min, and the furnace is cooled.

[0026] The present invention at least has the following beneficial effects: the repair method of the present invention is aimed at the matrix crack damage of the silicon carbide ceramic-based composite material parts of the aircraft engine with a width of less than 0.2 mm and a length of less than 10 mm, and a suitable metal brazing filler and a connecting agent are selected for vacuum brazing. The molten brazing filler will fill the inside of the crack under the capillary action, and react under the appropriate temperature conditions to form a welding joint, which can effectively restore 50-60% of the strength of the matrix material, prolong the usable time of the matrix material, so that the parts can continue to serve, and avoid the early scrapping of the parts; in addition, when preparing the EBC coating, the present invention uses atmospheric plasma spraying (APS) equipment to prepare the bonding layer (silicon Si) and the middle layer (mullite mullite), and uses a plasma-physical vapor deposition (PS-PVD) method to prepare the surface layer (ytterbium silicate Yb 2 SiO 5 ), compared with the EBC coating prepared by conventional atmospheric plasma spraying (porosity of about 10%), the porosity of the EBC coating prepared by the present invention can be reduced to about 1%. The reduced porosity of the EBC coating can greatly improve the coating performance, slow down the occurrence of cracks or peeling, and provide better protection for the CMC-SiC substrate, thereby delaying the cracking or corrosion of the substrate material, and further improving the service life of the repaired parts.

[0027] 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

[0028] Figure 1 It is a turbine guide vane made of silicon carbide ceramic matrix composite material. DETAILED DESCRIPTION

[0029] 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.

[0030] 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.

[0031] Figure 1 A typical silicon carbide ceramic matrix composite part with EBC coating is shown, namely an aircraft engine turbine guide vane.

[0032] Example 1

[0033] A method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aero-engine comprises the following steps:

[0034] Step 1: The substrate of the silicon carbide ceramic-based composite material part has a crack with a width of 0.1 mm and a length of 8 mm. The original environmental barrier coating is removed by laser cleaning technology to expose the CMC-SiC substrate. The specific process parameters of laser cleaning are as follows: the laser is a pulsed nanosecond laser, the laser wavelength is 1.064 μm, the power is 400 W, the spot diameter is 1 mm, the single pulse width is 100 ns, the repetition frequency is 20 kHz, the scanning speed is 400 mm / s, and the overlap rate is 98%;

[0035] Step 2: Place the parts in an acetone solution for ultrasonic cleaning for 15 minutes and then dry;

[0036] Step 3: Dry the Ag-Cu-Ti powder solder, then add NiCrobrazⅡ water-based adhesive. The mass ratio of Ag-Cu-Ti powder solder to NiCrobrazⅡ water-based adhesive is 5:1. Mix them evenly into a paste and apply it to the crack on the surface of the part and about 20mm near the crack with a fine brush along the direction of the crack. 2 For large areas, apply twice with a coating thickness of ~0.2mm;

[0037] Step 4: Place the parts into a vacuum brazing furnace for vacuum brazing. The molten brazing filler metal will fill the cracks under capillary action and react at a suitable temperature to form a welded joint. The Ti in the molten brazing filler metal first reacts with the SiC matrix to generate TiC and Si. The generated TiC grains nucleate on the surface of the SiC matrix and grow toward the liquid phase. As the brazing proceeds, the TiC and Si generated by the reaction continue to react with the Ti in the brazing filler metal to generate TiC. 3 SiC 2 Compound, i.e. Si+TiC+Ti→Ti 3 SiC 2 , finally, TiC and Ti 3 SiC 2 The mixed reaction layer composed of the compounds fills the cracks and is tightly connected to the CMC-SiC matrix. The specific conditions of vacuum brazing are: heating to 200°C at a heating rate of 5°C / min and keeping warm for 10 minutes, heating to 750°C at a heating rate of 8°C / min and keeping warm for 10 minutes, heating to 900°C at a heating rate of 8°C / min and keeping warm for 20 minutes, cooling to 600°C at a cooling rate of 3°C / min, and then opening the furnace after cooling to below 70°C. After testing, the parts recovered 56% of the strength of the matrix material, extending the usable time of the matrix material, allowing the parts to continue to serve and avoiding premature scrapping of the parts.

[0038] Step 5: 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;

[0039] Step 6: Use atmospheric plasma spraying (APS) equipment to prepare a bonding layer (silicon Si) and an intermediate layer (mullite mullite). 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, the part substrate is preheated by a plasma arc. The specific spraying process is shown in Table 1. The thickness of the bonding layer is ~30 μm, and the thickness of the intermediate layer is ~40 μm.

[0040] Table 1

[0041]

[0042] Step 7: Prepare the surface layer (ytterbium silicate Yb by plasma-physical vapor deposition (PS-PVD) method 2 SiO 5 ), the parts to be sprayed are placed on the turntable in the PS-PVD vacuum tank through a fixture, and the parts rotate with the turntable during spraying; the pressure in the PS-PVD vacuum tank is first pumped to 150Pa, and then filled with argon gas to 4000Pa. The part substrate is preheated by a plasma arc before spraying. The specific spraying process is shown in Table 2, and the surface layer thickness is ~100μm; after testing, the porosity of the prepared EBC coating is 1%;

[0043] Table 2

[0044]

[0045] Step 8: Grind the restored EBC coating to meet the roughness requirements of the product;

[0046] Step nine: perform annealing heat treatment at 850°C, with a heating rate of 5°C / min and an annealing holding time of 30min. Cool with the furnace to reduce residual stress and complete the repair.

[0047] Example 2

[0048] The present embodiment provides a method for repairing crack damage of silicon carbide ceramic-based composite parts of aircraft engines, wherein Ni-based powder brazing material is used in step three, and the specific conditions of vacuum brazing in step four are: heating to 200°C at a heating rate of 5°C / min and keeping warm for 10 minutes, heating to 1350°C at a heating rate of 8°C / min and keeping warm for 10 minutes, heating to 1400°C at a heating rate of 5°C / min and keeping warm for 20 minutes, cooling to 600°C at a cooling rate of 3°C / min, and cooling to below 70°C in the furnace before opening the furnace; the remaining steps are the same as in Example 1; after testing, the part recovers 51% of the strength of the matrix material.

[0049] Example 3

[0050] The present embodiment provides a method for repairing crack damage of silicon carbide ceramic-based composite parts of aircraft engines, wherein Cu-based powder brazing material is used in step three, and the specific conditions of vacuum brazing in step four are: heating to 200°C at a heating rate of 5°C / min and keeping warm for 10 minutes, heating to 750°C at a heating rate of 8°C / min and keeping warm for 10 minutes, heating to 800°C at a heating rate of 5°C / min and keeping warm for 20 minutes, cooling to 600°C at a cooling rate of 3°C / min, and cooling to below 70°C before opening the furnace; the remaining steps are the same as those in Example 1; after testing, the parts recovered about 50% of the strength of the matrix material.

[0051] Comparative Example 1

[0052] In this comparative example, the conventional atmospheric plasma spraying (APS) method was used to prepare the EBC coating, including a bonding layer (silicon Si), an intermediate layer (mullite) and a surface layer (ytterbium silicate Yb 2 SiO 5 ), the specific spraying process is shown in Table 3, and the remaining steps are the same as those in Example 1; after testing, the porosity of the prepared EBC coating is 10%.

[0053] Table 3

[0054]

[0055] The present invention selects powdered Ag-Cu-Ti solder, Cu-based or Ni-based active metal solder, which contains active elements such as Ti, Zr, V, Ta, Hf, Nb, Cr, etc.; it can be seen from Examples 1 to 3 that the Ag-Cu-Ti solder has the best connection effect, and the parts recover 56% of the strength of the matrix material, extending the usable time of the matrix material, allowing the parts to continue to serve, and avoiding premature scrapping of the parts.

[0056] It can be seen from Example 1 and Comparative Example 1 that the porosity of the EBC coating prepared by the atmospheric plasma spraying method used in the prior art is about 10%, while in the present invention, when preparing the EBC coating, atmospheric plasma spraying equipment is used to prepare the bonding layer and the intermediate layer, and the surface layer is prepared by combining the plasma-physical vapor deposition method. The porosity of the obtained EBC coating can be reduced to about 1%. The reduced porosity of the EBC coating can greatly improve the coating performance, slow down the occurrence of cracks or peeling, and provide better protection for the CMC-SiC substrate, thereby delaying the cracking or corrosion of the substrate material, and further improving the service life of the repaired parts.

[0057] 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 repairing crack damage of silicon carbide ceramic matrix composite parts of an aircraft engine, characterized in that: include: Remove environmental barrier coating on parts surface, clean and dry; Select powdered active metal brazing filler metal and connecting agent, mix them and apply them to the cracks on the surface of the parts and the surrounding areas; Vacuum brazing, welding seam grinding, cleaning and drying; using atmospheric plasma spraying equipment to prepare the bonding layer and intermediate layer of the environmental barrier coating, and using plasma-physical vapor deposition method to prepare the surface layer of the environmental barrier coating; The environmental barrier coating is polished and annealed to complete the crack damage repair of the silicon carbide ceramic matrix composite parts of the aircraft engine.

2. The method for repairing crack damage of 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: Use laser cleaning technology to remove the environmental barrier coating on the surface of the part to expose the part substrate; Step 2: Place the parts in an acetone solution for ultrasonic cleaning for 10 to 30 minutes and then dry; Step 3: Dry the powdered active metal brazing filler metal, add the bonding agent, mix evenly into a paste, and apply it on the cracks on the surface of the parts and around 10 to 30 mm along the crack direction. 2 Area region; Step 4: Place the parts into a vacuum brazing furnace for vacuum brazing; Step 5: 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 6: Using atmospheric plasma spraying equipment to prepare the bonding layer and the intermediate layer of the environmental barrier coating; Step 7: preparing a surface layer of the environmental barrier coating by using a plasma-physical vapor deposition method; Step 8: Grind the restored environmental barrier coating to make it meet the roughness requirements of the product; Step 9: Anneal the parts to complete the repair.

3. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 2, characterized in that: In the step 1, the specific process parameters of laser cleaning are: the laser is a pulsed nanosecond laser, the laser wavelength is 1.064 μm, the power is 300-500 W, the spot diameter is 0.5-2 mm, the single pulse width is 50-150 ns, the repetition frequency is 10-30 kHz, the scanning speed is 300-500 mm / s, and the overlap rate is 95-99%.

4. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 2, characterized in that: In the step three, the powdered active metal solder is one of Ag-Cu-Ti solder, Ni-based solder, and Cu-based solder; the connecting agent is one of NiCrobrazⅡ type water-based adhesive, TENSOL No.6 and trichloroethylene, polystyrene and trichloroethylene, ethylene glycol, and propylene glycol; the mass ratio of the powdered active metal solder to the connecting agent is 3 to 6:

1.

5. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine as claimed in claim 2, characterized in that: In the step 4, the specific conditions of vacuum brazing are: heating to 200-220°C at a heating rate of 5-10°C / min and keeping warm for 10-20min, heating to 750-1350°C at a heating rate of 8-10°C / min and keeping warm for 10-20min, heating to 800-1400°C at a heating rate of 5-10°C / min and keeping warm for 20-30min, cooling to 600-700°C at a cooling rate of 3-5°C / min, and opening the furnace after cooling to below 70°C.

6. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine as claimed in claim 2, characterized in that: In the step six, the bonding layer is a silicon layer with a thickness of 20 to 50 μm; the middle layer is a mullite layer with a thickness of 30 to 80 μm; when preparing the coating, a fixture is used to place the parts to be sprayed on a turntable of the atmospheric plasma spraying equipment. During spraying, the parts rotate with the turntable to ensure uniform coating, and the part substrate is preheated by a plasma arc before spraying.

7. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 6, characterized in that: The atmospheric plasma spraying process parameters for preparing the bonding layer are: argon flow rate 40-60L / min, hydrogen flow rate 4-6L / min, current 500-600A, spraying distance 100-110mm, power 26-27kW, powder feeding rate 4-6r / min, carrier gas flow rate 8-10L / min, turntable speed 40-70r / min, substrate preheating 550-650°C; the atmospheric plasma spraying process parameters for preparing the intermediate layer are: argon flow rate 35-45L / min, hydrogen flow rate 8-12L / min, current 600-700A, spraying distance 100-110mm, power 33-34kW, powder feeding rate 2-4r / min, carrier gas flow rate 8-10L / min, turntable speed 40-70r / min, substrate preheating 750-850°C.

8. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine as claimed in claim 2, characterized in that: In the step seven, the surface layer is a Yb2SiO5 layer with a thickness of 50 to 120 μm. When preparing the coating, a fixture is used to place the part to be sprayed on a turntable in a vacuum tank of a plasma-physical vapor deposition equipment. During spraying, the part rotates with the turntable. The pressure in the vacuum tank is first pumped to 100 to 200 Pa, and then filled with argon gas to 3000 to 5000 Pa. Before spraying, the part substrate is preheated by a plasma arc.

9. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 8, characterized in that: The plasma-physical vapor deposition process parameters for preparing the surface layer are: argon flow rate 50-70 L / min, hydrogen flow rate 25-35 L / min, current 2000-3000 A, spraying distance 1000-1100 mm, power 35-45 kW, powder feeding rate 4-6 r / min, carrier gas flow rate 8-10 L / min, turntable speed 40-70 r / min, substrate preheating 550-650°C.

10. The method for repairing crack damage of a silicon carbide ceramic matrix composite material part of an aircraft engine according to claim 2, characterized in that: In the step nine, the annealing temperature is 800-1000° C., the heating rate is 3-10° C. / min, the holding time is 20-40 min, and the furnace is cooled.

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