A ceramic matrix composite and a method of making the same

Abstract

The application discloses a ceramic matrix composite material preparation method, and preparation steps are as follows: continuous carbon fibers are used as reinforcing bodies to prepare a fiber preform; the obtained fiber preform is placed in a high-temperature-resistant vacuum sintering furnace to perform an isothermal chemical vapor deposition process, so that a preform with a pyrolytic carbon interface layer is obtained; the isothermal chemical vapor deposition process is used to preliminarily densify the preform, silicon carbide matrix is introduced, so that a C f / SiC blank is obtained; the isothermal chemical vapor deposition process is used to introduce the pyrolytic carbon matrix into the C f / SiC blank; the pressure impregnation-pyrolysis process is used to further introduce an outer resin carbon matrix into the secondary blank; silicon powder is used as raw material, a silicon vapor infiltration process is used to perform silicon infiltration treatment on the tertiary blank, and finally, a C f / C-SiC composite material is prepared. While the mechanical properties of the composite material are ensured, the preparation period is greatly shortened, the preparation cost is reduced, and the ceramic matrix composite material preparation process is easy to realize industrialization.

Description

Technical Field

[0001] This invention relates to the field of ceramic matrix composite material preparation technology, specifically to a method for preparing a high-performance ceramic matrix composite material for the thermal structure of aircraft, using continuous carbon fiber as reinforcement and carbon and silicon carbide as the dual matrix. Background Technology

[0002] With the rapid development of aviation technology, aircraft are facing higher demands in terms of service life, operational efficiency, fuel consumption, and thrust-to-weight ratio. The highest temperature of the high-temperature alloys used in traditional aircraft thermal structural components is only 1080℃, while the operating temperature of the thermal structures in current advanced aircraft models far exceeds this. Even with advanced cooling technologies, it is impossible to meet the high-temperature operating requirements of aircraft thermal structures.

[0003] Ceramic matrix composites (CMCs) have become a strategic thermal structural material that is the focus of research in many countries due to their numerous advantages, such as high temperature resistance, oxidation resistance, wear resistance, low density, corrosion resistance, and crack insensitivity. Carbon fiber reinforced carbon / silicon carbide dual-matrix composites inherit the excellent properties of carbon fiber reinforced carbon matrix composites, such as high specific strength, high specific modulus, high thermal conductivity, and low coefficient of thermal expansion. When used in thermal structural components of aircraft, they can maximize the reliability of the aircraft while significantly reducing the overall weight.

[0004] The main advantages of CVI (Ceramic Matrix Infiltration) technology in preparing ceramic matrix composites include minimal fiber damage, uniform matrix preparation, and good mechanical properties. However, its industrialization is severely limited by its disadvantages, such as low raw material gas utilization, demanding equipment requirements, and preparation cycles often exceeding hundreds of hours. Vapor-phase silicating is another method for preparing ceramic matrix composites. Due to its fast reaction rate, high preparation efficiency, and low porosity of the finished product, it can be used as a reliable process for preparing hot-end components for aerospace vehicles. However, existing vapor-phase silicating methods for preparing C... f The following technical problems exist with C-SiC composite materials: free silicon during silicon infiltration can easily cause serious damage to carbon fibers and interfaces; and the large amount of residual silicon in the prepared ceramic matrix causes thermal mismatch. Therefore, the prepared C-SiC composite materials typically suffer from these problems. f / C-SiC composites have low mechanical properties, with a typical flexural strength of only 80–120 MPa. Existing vapor-phase silicon infiltration processes have stringent requirements for vacuum levels during the reaction, typically requiring vacuum equipment to maintain a pressure below 50 Pa at high temperatures. Therefore, the low performance of the prepared materials and the extremely stringent equipment requirements greatly limit the application of vapor-phase silicon infiltration in the preparation of aerospace structural ceramic matrix composites. Summary of the Invention

[0005] The purpose of this invention is to overcome the shortcomings of the prior art and provide a method for preparing ceramic matrix composites that, while ensuring the mechanical properties of the composites, significantly shortens the preparation cycle and reduces the preparation cost. This method is a ceramic matrix composite preparation process that is easy to industrialize.

[0006] Another objective of this invention is to provide a ceramic matrix composite material prepared by the above method.

[0007] A method for preparing a ceramic matrix composite material includes the following steps:

[0008] (1) Preform preparation

[0009] Fiber preforms were prepared using continuous carbon fibers as reinforcement.

[0010] (2) Fiber interface coating treatment

[0011] The fiber preform obtained in step (1) is placed in a high-temperature vacuum sintering furnace for isothermal chemical vapor deposition process to obtain a preform with a pyrolytic carbon interface layer.

[0012] (3) Introduction of CVI-SiC layer matrix

[0013] The preform obtained in step (2) was initially densified using an isothermal chemical vapor deposition process to introduce a silicon carbide matrix, resulting in C f / SiC preform;

[0014] (4) Carbon introduction into the matrix of the isolation layer

[0015] The C obtained in step (3) is introduced into the C using an isothermal chemical vapor deposition process. f A pyrolytic carbon matrix was introduced into the SiC preform to obtain a density of 1.2–1.3 g / cm³. 3 Secondary billet;

[0016] (5) Carbon introduction into the outer matrix

[0017] An outer layer of resin carbon matrix was further introduced into the secondary preform obtained in step (4) using a pressure impregnation-pyrolysis process to obtain a preform with a density of 1.4–1.5 g / cm³. 3 The third-stage billet;

[0018] (6) Using silicon powder as raw material, the three-stage preform obtained in step (5) is subjected to silicon infiltration treatment using a vapor phase silicon infiltration process to finally obtain C f / C-SiC composite material, with a density of 1.9~2.2g / cm³. 3 .

[0019] The carbon fiber preform employs weaving methods such as 2.5D needle punching, three-dimensional orthogonal weaving, fine weaving puncture, and ply-lay stitching. The carbon fiber volume fraction in the preform is 30–45%, and the initial density of the preform is 0.4–0.9 g / cm³. 3 ;

[0020] Before step (2), the carbon fiber preform prepared in step (1) is subjected to a degumming pretreatment:

[0021] The woven preform is degummed at high temperature and kept at 900°C for two hours in a nitrogen atmosphere.

[0022] Step (2) specifically involves:

[0023] The carbon fiber preform obtained in step (1) was placed in a high-temperature vacuum sintering furnace and deposited using natural gas as the carbon source gas and nitrogen as the dilution and protective gas. The deposition temperature was 1050-1120℃, the deposition time was 10-15 hours, and the deposition pressure was 3000-6000Pa. The natural gas flow rate was 0.5-1.1L / min, and the nitrogen flow rate was 2-4L / min.

[0024] In step (2), the thickness of the pyrolytic carbon interface is 500–1200 nm.

[0025] Step (3) specifically involves:

[0026] The preform prepared in step (2) is placed in a chemical vapor deposition furnace, with trichloromethylsilane as the silicon and carbon source gas, hydrogen as the carrier gas and dilution gas, and argon as the protective gas, and isothermal chemical vapor deposition is carried out. The deposition temperature is 1000-1100℃, the deposition time is 30-60 hours, and the deposition pressure is 1000-2000Pa.

[0027] The flow rate ratio of the reacting gases is 1:30:30 for trichloromethylsilane, hydrogen, and argon.

[0028] C obtained by introducing silicon carbide matrix f The density of the SiC green body is 0.9–1.1 g / cm³. 3 .

[0029] Step (4) is as follows:

[0030] The C prepared in step (3) f The SiC preform is placed in a high-temperature sintering furnace, with propane as the carbon source gas and nitrogen as the carrier gas, and isothermal chemical vapor deposition is carried out. The deposition temperature is 980-1100℃, the deposition time is 15-20 hours, the pressure is 1000-1500Pa, and the flow ratio of propane to nitrogen is 1:4.

[0031] Step (5) specifically involves:

[0032] Add the secondary preform obtained in step (4) into a pressure impregnation tank, evacuate and maintain pressure for half an hour, pour the preform mixture after heat preservation into the impregnation tank and pressurize it to 5 MPa with inert gas, and maintain pressure for 1 hour; put the impregnated secondary preform into an oven and cure it at 120°C for 4 hours, and then place the secondary preform in a vacuum pyrolysis furnace for vacuum high-temperature pyrolysis at 900°C. The vacuum pyrolysis heating rate is: 8°C per hour below 600°C, and 20°C per hour above 600°C.

[0033] The preparation method of the precursor mixture is as follows:

[0034] A precursor mixture was prepared by thoroughly mixing furan resin or phenolic resin as a precursor, phosphoric acid as a curing agent, and anhydrous ethanol as a solvent. The mass ratio of resin, phosphoric acid, and anhydrous ethanol was 4:1:1 to 4. The prepared precursor mixture was placed in a 60°C oven and kept warm for half an hour to increase its fluidity.

[0035] Step (6) specifically involves:

[0036] The tertiary blank prepared in step (5) is placed in a mold with silicon powder spread on the bottom of the mold. The blank is raised to a certain height so that it does not come into direct contact with the silicon powder. After the mold is sealed, it is placed in a high-temperature sintering furnace and heated under argon protection at a rate of 5-8℃ / min. When the temperature reaches 1450℃, it is held for half an hour. The temperature is then raised to 1600-1800℃ and held for 0.5-2 hours. The pressure during the holding period is 1000-3000Pa. After the pressure holding period, the temperature is raised to 1850℃ and held for 0.5 hours. During the holding period, the vacuum system of the high-temperature sintering furnace works continuously to maintain the pressure at 1000-1500Pa.

[0037] The mold includes a cylindrical body and a top cover;

[0038] The cylinder is made of graphite, and the inner wall of the cylinder is uniformly coated with boron nitride powder-varnish dispersant and then dried at 120°C; a graphite sample support frame is installed at the bottom of the cylinder.

[0039] The mold cover has a three-layer structure. The innermost layer is high-purity flexible graphite paper with a thickness of 0.5-1mm. The next layer is carbon felt, graphite felt, hard carbon felt or hard graphite felt with a thickness of 1-5mm. The outermost layer is isostatic graphite plate.

[0040] The third-level blank is placed in the mold in the following state:

[0041] The mass ratio of the third-stage preform to silicon powder is 1:2 to 1:4; the bottom surface of the third-stage preform is placed 10 to 25 mm away from the upper surface of the silicon powder after being placed on the graphite sample support frame.

[0042] When multiple tertiary blanks are placed on the graphite sample support frame, the edge distance between two adjacent tertiary blanks is greater than 20mm; the area of ​​the bottom surface of the tertiary blank covered by the graphite sample support frame in the mold is less than one-fifth of the bottom area of ​​the tertiary blank.

[0043] Ceramic matrix composite material for thermal structural components of aircraft prepared by the method described in the above technical solution.

[0044] This invention introduces a CVI-SiC matrix, a pyrolytic carbon matrix, and an outer resin carbon matrix before silicon infiltration, forming a multilayer structure that provides strong protection for the inner fibers and interfaces, effectively reducing fiber damage during the silicon infiltration process. The developed special mold and silicon infiltration process significantly reduce the residual silicon content in the matrix, solving the problem of thermal mismatch in the matrix after silicon infiltration. Compared with conventional silicon infiltration methods for preparing ceramic matrix composites, this improves the mechanical properties of the components. The prepared material is suitable for the thermal structure environment of aerospace vehicles and is an important candidate material for high-temperature thermal structures in advanced weaponry.

[0045] The superior effects of this invention are as follows:

[0046] (1) The SiC matrix and the carbon matrix of the isolation layer and the carbon resin of the outer layer prepared by CVI constitute a strong protective measure for the inner fiber and interface, which effectively improves the resistance of carbon fiber and interface to erosion during the high-temperature silicon infiltration process. Compared with the carbon fiber reinforced silicon carbide composite material prepared by the general vapor phase silicon infiltration process, its bending strength is increased by 20-30% and its fracture toughness is increased by 10-15%.

[0047] (2) The process adopted uses vapor phase silicon diffusion to densify the green body at the end. Compared with other vapor phase silicon diffusion methods, it achieves an extremely low free silicon residual rate in the ceramic matrix through special molds and process adjustments. The process has lower requirements for equipment vacuum degree and can complete uniform vapor phase silicon diffusion under relatively high partial pressure.

[0048] (3) Compared with traditional chemical vapor deposition, it greatly reduces the process cost, and the process time is only 1 / 6 of that of chemical vapor deposition under the same density conditions.

[0049] (4) It can withstand temperatures up to 1300℃ and is suitable for the working temperature of the thermal structure of advanced aircraft currently in service. Attached Figure Description

[0050] Figure 1 This is a schematic diagram of the mold and silicon infiltration environment used in the vapor phase silicon infiltration of the present invention.

[0051] Figure 2 C prepared for Embodiment 1 of the present invention f Bending test stress-strain curves of C-SiC composite materials;

[0052] Figure 3 This is a scanning electron microscope image of the pyrolytic carbon interface and SiC isolation layer prepared in Embodiment 2 of the present invention;

[0053] Figure 4 C prepared for Embodiment 2 of the present invention f Bending test stress-strain curves of C-SiC composite materials;

[0054] Figure 5 This is the microstructure of the fiber end face after the sample breaks in Embodiment 2 of the present invention;

[0055] Figure 6 This is a scanning electron microscope image of the carbon microstructure of the isolation layer substrate prepared in Example 3 of the present invention;

[0056] Figure 7 This is a scanning electron microscope image of the outer carbon matrix microstructure prepared in Example 3 of the present invention;

[0057] Figure 8 C prepared in Example 3 of this invention f X-ray diffraction pattern of C-SiC composite material;

[0058] Figure 9 The bending stress-strain curves of the sample prepared in Example 3 of this invention and the sample prepared by the prior art are shown.

[0059] Figure 10 The X-ray diffraction patterns of the sample prepared in Example 3 of this invention and the sample prepared by the existing vapor phase silicate infiltration technique are shown. Detailed Implementation

[0060] The following specific preferred embodiments further describe the present invention, but do not limit the scope of protection of the present invention.

[0061] A method for preparing a ceramic matrix composite material specifically includes the following steps:

[0062] Step 1: Preform Preparation: A fiber preform is prepared using T-700 continuous carbon fiber as reinforcement. The preform employs a 2.5D needle-punching, three-dimensional orthogonal weaving, fine-knitting, and lay-up stitching method. The carbon fiber volume fraction in the preform is 30-45%, and the initial density of the preform is 0.4-0.9 g / cm³. 3 ;

[0063] Step 2: Fiber interface coating treatment: The fiber preform obtained in step (1) is placed in a high-temperature vacuum sintering furnace for isothermal chemical vapor deposition treatment, with natural gas as the carbon source gas and nitrogen as the dilution and protective gas. The deposition temperature is 1050-1120℃, the deposition time is 10-15 hours, and the deposition pressure is 3000-6000Pa; a preform with a pyrolytic carbon interface layer is obtained.

[0064] Step 3: Introduction of CVI-SiC substrate: The preform obtained in step (2) is initially densified using an isothermal chemical vapor deposition process to introduce a silicon carbide substrate, resulting in CVI-SiC layer. f / SiC preform, isothermal chemical vapor deposition process is as follows: trichloromethylsilane is used as silicon and carbon source gas, hydrogen is used as carrier gas and dilution gas, and argon is used as protective gas, isothermal chemical vapor deposition is carried out, the deposition temperature is 1000~1100℃, the deposition time is 30~60 hours, and the deposition pressure is 1000~2000Pa.

[0065] Step 4: Introduction of carbon matrix into the isolation layer: A pyrolytic carbon matrix is ​​introduced into the green body obtained in step (3) using an isothermal chemical vapor deposition process to obtain a green body with a density of 1.2-1.3 g / cm³. 3 Secondary billet;

[0066] Step 5: Introduction of outer layer carbon matrix: An outer layer resin carbon matrix is ​​further introduced into the secondary preform obtained in step (4) using a vacuum impregnation-pyrolysis process to obtain a density of 1.4-1.5 g / cm³. 3 The third-stage billet;

[0067] Step 6: Using silicon powder as raw material, the preform described in step (5) is subjected to silicon infiltration treatment using a reactive vapor phase silicon infiltration process to finally obtain C f / C-SiC composite material, with a density of 1.9-2.2 g / cm³. 3 .

[0068] In step 1, the carbon fiber preform undergoes a degumming pretreatment: the woven preform is degummed at high temperature and kept at 900°C for two hours in a nitrogen atmosphere.

[0069] In step 2, the thickness of the pyrolytic carbon interface is 500–1200 nm.

[0070] In step 2, the natural gas flow rate is 0.5–1.1 L / min, and the nitrogen flow rate is 2–4 L / min.

[0071] The C obtained by introducing the silicon carbide matrix in step 3 f The density of the SiC green body is 0.9–1.1 g / cm³. 3 .

[0072] In step 3, the flow rate ratio of the reacting gases is 1:30:30 for trichloromethylsilane, hydrogen, and argon.

[0073] The method for preparing the carbon matrix of the isolation layer in step 4 is as follows:

[0074] Using propane as the carbon source gas and nitrogen as the carrier gas, isothermal chemical vapor deposition was carried out at a deposition temperature of 980–1100℃, a deposition time of 15–20 hours, a pressure of 1000–1500 Pa, and a propane to nitrogen flow ratio of 1:4.

[0075] The method for preparing the outer layer matrix carbon in step 5 is as follows:

[0076] A precursor mixture was prepared by thoroughly mixing furan resin or phenolic resin as a precursor, phosphoric acid as a curing agent, and anhydrous ethanol as a solvent. The mass ratio of resin, phosphoric acid, and anhydrous ethanol was 4:1:1-4. The mixture was placed in a 60°C oven for half an hour to increase its fluidity. The secondary preform obtained in step 4 was added to a pressure impregnation tank, vacuumed, and pressurized for half an hour. The preform mixture after heat preservation was poured into the impregnation tank and pressurized to 5 MPa with inert gas for 1 hour. The impregnated preform was placed in an oven and cured at 120°C for 4 hours. Subsequently, the preform was placed in a vacuum pyrolysis furnace for high-temperature vacuum pyrolysis at 900°C.

[0077] In step 5, the vacuum pyrolysis heating rate is: 8°C per hour for temperatures below 600°C, and 20°C per hour for temperatures above 600°C.

[0078] The silicon infiltration process in step 6 is as follows:

[0079] The blank obtained in step 5 is placed in a special tooling mold with silicon powder spread flat at the bottom. The blank is raised to a certain height so that it does not come into direct contact with the silicon powder. After the mold is sealed, it is placed in a high-temperature sintering furnace and heated under argon protection at a rate of 5-8℃ / min. When the temperature reaches 1450℃, it is held for half an hour. The temperature is then increased to 1600-1800℃ and held for 0.5-2 hours. The pressure during the holding period is 1000-3000Pa. After the pressure holding period, the temperature is increased to 1850℃ and held for another 0.5 hours. During the holding period, the vacuum system works continuously to maintain the pressure at a low level of 1000-1500Pa.

[0080] In the above steps, the tertiary preform is placed in the mold in the following manner: the ratio of tertiary preform to silicon powder is 1:2 to 1:4 by mass; the bottom surface of the tertiary preform is 10 to 25 mm from the top surface of the silicon powder; when multiple tertiary preforms are present, the edge distance between the tertiary preforms is greater than 20 mm; the area of ​​the bottom surface of the tertiary preform covered by the support is less than one-fifth of the bottom area.

[0081] like Figure 1As shown, the special tooling mold structure in the above steps is as follows: the inner wall of the mold is uniformly coated with boron nitride powder-varnish dispersant and then dried at 120℃. The mold cover has a three-layer structure: the innermost layer is high-purity flexible graphite paper with a thickness of 0.5-1mm; the next layer is carbon felt, graphite felt, hard carbon felt, or hard graphite felt with a thickness of 1-5mm; and the outermost layer is isostatic graphite plate.

[0082] Example 1

[0083] A 2.5D needle-punched preform was prepared using T-700 continuous carbon fiber as raw material. The carbon fiber volume fraction in the preform was 30%, and the initial density of the preform was 0.45 g / cm³. 3 Before use, the preform should be kept at 900℃ in a nitrogen atmosphere for two hours to remove the adhesive.

[0084] The prepared fiber preform was placed in a high-temperature vacuum sintering furnace, and the furnace temperature was raised to 1050℃ under continuous vacuum conditions. Natural gas and nitrogen were then introduced at flow rates of 0.5L / min and 2L / min, respectively. Deposition was carried out for 10 hours, and the furnace pressure was maintained at 3000Pa. The thickness of the pyrolytic carbon interface on the surface of the prepared fiber was 500nm.

[0085] The prepared preform was placed in a chemical vapor deposition furnace, and the furnace temperature was raised to 1000℃. Trichloromethylsilane, hydrogen, and argon were introduced into the furnace in a ratio of 1:30:30, and the furnace pressure was maintained at 1000 Pa. The deposition was carried out for 30 hours, yielding a density of 0.9 g / cm³. 3 C f / SiC preform.

[0086] Prepared C f The SiC preform was placed in a high-temperature sintering furnace, and the furnace temperature was raised to 980℃ under continuous vacuum. Propane and nitrogen were introduced into the sintering furnace at flow rates of 1.5 L / min and 6 L / min, respectively, while maintaining a pressure of 1000–1050 Pa. The sintering process lasted for 15 hours, yielding a preform with a density of 1.2 g / cm³. 3 The secondary blank.

[0087] Furan resin, phosphoric acid, and anhydrous ethanol were mixed in a mass ratio of 4:1:4 and stirred thoroughly for 15 minutes. The precursor mixture was then placed in an oven and heated from room temperature to 60°C, holding at that temperature for 30 minutes. The secondary preform prepared in step 4 was placed in a pressure impregnation tank, evacuated, and held under pressure for 30 minutes. The preform mixture, after being heated to the required temperature, was then added, and argon gas was introduced to a furnace pressure of 5 MPa, holding at that pressure for 1 hour. After the impregnated preform was cured in an oven, it was placed in a pyrolysis furnace and heated under vacuum conditions throughout. The temperature was increased by 8°C per hour from room temperature to 600°C, and by 20°C per hour from 600°C to 900°C. It was then held at 900°C for two hours and cooled in the furnace. The resulting product had a density of 1.4 g / cm³.3 The third-stage billet.

[0088] A specially designed mold was used to evenly distribute silicon powder at its bottom, four times the mass of the tertiary preform. The tertiary preform prepared in step 5 was placed on a graphite sample holder within the mold, with a height difference between the preform and the bottom surface. The bottom surface of the tertiary preform was 15 mm from the top surface of the silicon powder layer, and the portion of the bottom surface of the tertiary preform covered by the holder accounted for one-sixth of the total bottom area. The edge distance between multiple tertiary preforms was 20 mm. After sealing the mold, it was placed in a high-temperature sintering furnace. After evacuation, argon gas was introduced to 1500–3000 Pa, and the temperature was increased at a rate of 6 °C / min. After reaching 1000 °C, the heating rate was increased to 8 °C / min. The temperature was then increased to 1450 °C and held for 30 min. Subsequently, the temperature was increased to 1600 °C and held for 1 hour, followed by further increases to 1850 °C and a holding time of 0.5 hours. During the holding time at 1850 °C, the vacuum pump and opening were at maximum, maintaining a pressure of 1000–1500 Pa. The furnace was then cooled. The final prepared C f The final density of the C-SiC composite material is 2.0 g / cm³. 3 Three C3C samples were prepared for Example 1. f The flexural strength test was performed on the / C-SiC composite material specimens, and the results are as follows: Figure 2 As shown in the figure, the three curves correspond to the test results of the three samples, respectively. Figure 2 It can be seen that the composite material prepared in Example 1 has a room temperature flexural strength of at least 150 MPa, an average flexural strength of 185 MPa, and a fracture toughness of at least 14.2 MPa·m. 1 / 2 .

[0089] Example 2

[0090] A layup-stitched preform was prepared using T-700 continuous carbon fiber as raw material. The carbon fiber volume fraction in the preform was 45%, and the initial density of the preform was 0.88 g / cm³. 3 Before use, the preform should be kept at 900℃ in a nitrogen atmosphere for two hours to remove the adhesive.

[0091] The prepared fiber preform was placed in a high-temperature vacuum sintering furnace, and the furnace temperature was raised to 1100℃ under continuous vacuum conditions. Natural gas and nitrogen were then introduced at flow rates of 1.1 L / min and 4 L / min, respectively. Deposition was carried out for 15 hours, and the furnace pressure was maintained at 5000 Pa. The thickness of the pyrolytic carbon interface on the surface of the prepared fiber was 1200 nm.

[0092] The prepared preform was placed in a chemical vapor deposition furnace, and the furnace temperature was raised to 1100℃. Trichloromethylsilane, hydrogen, and argon were introduced into the furnace in a ratio of 1:30:30, and the furnace pressure was maintained at 2000 Pa. The deposition was carried out for 45 hours, yielding a preform with a density of 1.1 g / cm³. 3 Cf A PyC interface layer and a CVI-SiC isolation layer structure were successfully fabricated using a SiC preform, effectively protecting the inner fiber layers from silicon erosion. Scanning electron microscope images of the pyrolytic carbon interface and the SiC isolation layer are shown below. Figure 3 As shown, where Figure 3 (a) in the figure represents the microstructure of the pyrolytic carbon interface. Figure 3 (b) in the middle is Figure 3 Enlarged view of the selected area in (a), Figure 3 Image (c) shows the microstructure of SiC. Figure 3 (d) is Figure 3 Enlarged view of the selected area in (c).

[0093] Prepared C f The SiC preform was placed in a high-temperature sintering furnace, and the furnace temperature was raised to 1050℃ under continuous vacuum. Propane and nitrogen were introduced into the sintering furnace at flow rates of 1.5 L / min and 6 L / min, respectively, maintaining a pressure of 1400–1500 Pa. The sintering process lasted 18 hours, yielding a preform with a density of 1.3 g / cm³. 3 The blank.

[0094] Furan resin, phosphoric acid, and anhydrous ethanol were mixed in a mass ratio of 4:1:1 and stirred thoroughly for 15 minutes. The precursor mixture was then placed in an oven and heated from room temperature to 60°C, holding at that temperature for 30 minutes. The prepared secondary preform was placed in a pressure impregnation tank, evacuated, and held under pressure for 30 minutes. The heated mixture was then added, and argon gas was introduced to a furnace pressure of 5 MPa, which was held for 1 hour. After the impregnated preform was cured in an oven, it was placed in a pyrolysis furnace and heated under vacuum conditions throughout. The temperature was increased by 8°C per hour from room temperature to 600°C, and by 20°C per hour from 600°C to 900°C. It was then held at 900°C for two hours and cooled in the furnace. The resulting product had a density of 1.5 g / cm³. 3 The third-stage billet.

[0095] A specially designed mold was used to evenly spread silicon powder at a mass twice that of the tertiary preform. The preform prepared in step 5 was placed on a graphite sample holder in the mold, with a height difference between the preform and the bottom surface. The bottom surface of the tertiary preform was 25 mm from the top surface of the silicon powder layer. The portion of the bottom surface of the tertiary preform covered by the holder accounted for one-sixth of the entire bottom area. The edge distance between multiple tertiary preforms was 20 mm. After sealing the mold, it was placed in a high-temperature sintering furnace. After evacuation, argon gas was introduced to 1500–3000 Pa, and then the temperature was increased at a rate of 6 °C / min. After reaching 1000 °C, the heating rate was increased to 8 °C / min. The temperature was increased to 1450 °C and held for 30 min. Then, the temperature was increased to 1800 °C and held for 0.5 hours. The temperature was then increased to 1850 °C and held for 0.5 hours. During the holding at 1850 °C, the vacuum pump and the opening were opened to the maximum, maintaining a pressure of 1000–1500 Pa. The furnace was then cooled. The prepared C fThe final density of the C-SiC composite material is 1.9 g / cm³. 3 Three C3C samples were prepared for Example 2. f The flexural strength test was performed on the / C-SiC composite material specimens, and the results are as follows: Figure 4 As shown in the figure, the three curves correspond to the test results of the three samples prepared in Example 2, respectively. Figure 4 As can be seen from Example 2, the prepared composite material has a room temperature flexural strength of at least 180 MPa, an average flexural strength of 187 MPa, and an average flexural modulus of 19.91 GPa. After the sample breaks, it exhibits... Figure 5 As shown, Figure 5 Image (a) shows the microstructure of the fibers at the end face of the tensile fracture sample. Figure 5 (b) shows the microstructure of the fiber in the cross section of the tensile test specimen. There is obvious fiber pull-out, and the fiber surface is smooth and intact without erosion damage, which proves that the multilayer composite matrix plays a good role in isolation and protection during the vapor phase silica infiltration process.

[0096] Example 3

[0097] T-700 continuous carbon fiber was used as raw material to prepare a fine-braided puncture preform. The carbon fiber volume fraction in the preform was 40%, and the initial density of the preform was 0.73 g / cm³. 3 Before use, the preform should be kept at 900℃ in a nitrogen atmosphere for two hours to remove the adhesive.

[0098] The prepared fiber preform was placed in a high-temperature vacuum sintering furnace, and the furnace temperature was raised to 1080℃ under continuous vacuum. Natural gas and nitrogen were then introduced at flow rates of 0.8 L / min and 3 L / min, respectively. Deposition was carried out for 15 hours, and the furnace pressure was maintained at 3500 Pa. The thickness of the pyrolytic carbon interface on the surface of the prepared fiber was 750 nm.

[0099] The prepared preform was placed in a chemical vapor deposition furnace, and the furnace temperature was raised to 1060℃. Trichloromethylsilane, hydrogen, and argon were introduced into the furnace in a ratio of 1:30:30, and the furnace pressure was maintained at 1300 Pa. The deposition was carried out for 52 hours, yielding a preform with a density of 1.05 g / cm³. 3 C f / SiC preform.

[0100] Prepared C f The SiC preform was placed in a high-temperature sintering furnace, and the furnace temperature was raised to 1000℃ under continuous vacuum. Propane and nitrogen were introduced into the sintering furnace at flow rates of 1.5 L / min and 6 L / min, respectively, maintaining a pressure of 1000–1100 Pa. The sintering process lasted 18 hours, yielding a preform with a density of 1.2 g / cm³. 3 The blank.

[0101] Furan resin, phosphoric acid, and anhydrous ethanol were mixed in a mass ratio of 4:1:2 and stirred thoroughly for 15 minutes. The precursor mixture was then placed in an oven and heated from room temperature to 60°C, holding at that temperature for 30 minutes. The prepared secondary preform was placed in a pressure impregnation tank, evacuated, and held under pressure for 30 minutes. The heated mixture was then added, and argon gas was introduced to a furnace pressure of 5 MPa, which was held for 1 hour. After the impregnated preform was cured in an oven, it was placed in a pyrolysis furnace and heated under vacuum conditions throughout. The temperature was increased by 8°C per hour from room temperature to 600°C, and by 20°C per hour from 600°C to 900°C. It was then held at 900°C for two hours and cooled in the furnace. The resulting product had a density of 1.43 g / cm³. 3 The third-stage billet, Figure 6 and Figure 7 The isolation layer and the outer matrix carbon are respectively prepared. Figure 6 The four sub-figures (a), (b), (c), and (d) in the figure represent the microstructure of the isolation layer prepared in Example 3 at different magnifications. Figure 7 The four sub-figures (a), (b), (c), and (d) in the figure represent the microstructure of the material prepared in Example 3 at different magnifications. As can be seen from the figure, the isolation layer PyC tightly covers the surface of the inner structure and plays a certain protective role in the silicon infiltration process, while the outer matrix carbon is distributed in a blocky and broken manner, and its cracks and pores provide a large number of channels for the subsequent silicon infiltration reaction.

[0102] Three times the weight of the third-stage preform was evenly spread on the bottom of a specially made mold, with a density of 1.43 g / cm³. 3 The tertiary preforms were placed on a graphite sample holder in the mold, with a height difference between the preform and the bottom surface. The bottom surface of the tertiary preform was 18 mm from the top surface of the silicon powder layer. The portion of the bottom surface of the tertiary preform covered by the holder accounted for one-sixth of the entire bottom area. The edge distance between multiple tertiary preforms was 20 mm. After sealing the mold, it was placed in a high-temperature sintering furnace. After evacuation, argon gas was introduced to 1500–3000 Pa, and then the temperature was increased at a rate of 6 °C / min. After reaching 1000 °C, the heating rate was increased to 8 °C / min. The temperature was increased to 1450 °C and held for 30 min. Then, the temperature was increased to 1750 °C and held for 1.5 hours. The temperature was then increased to 1850 °C and held for 0.5 hours. During the holding at 1850 °C, the vacuum pump and the opening were opened to the maximum, maintaining a pressure of 1000–1500 Pa. The furnace was then cooled. The prepared C f The final density of the C-SiC composite material is 2.15 g / cm³. 3 Example 3: Preparation of three C f X-ray diffraction analysis was performed on the C-SiC composite material sample, and the results are as follows: Figure 8 As shown in the figure, the blue, red, and black curves correspond to the test results of the three samples prepared in Example 3, respectively. The green curve is the standard curve. Figure 8As can be seen from Example 3, the prepared composite material has a good 3CSiC matrix structure, and the Si signal is not obvious, which proves that the residual silicon content in the matrix is ​​extremely low.

[0103] Figure 9 This section compares the flexural mechanical properties of the sample prepared in Example 3 with those prepared using existing vapor-phase silicon infiltration technology. The red and blue curves represent the two samples prepared in Example 3, the green curve represents the sample prepared using the existing vapor-phase silicon infiltration process with the applied multilayer matrix of this invention, and the black curve represents the sample prepared using the existing vapor-phase silicon infiltration process. It can be seen that the multilayer structure of CVI-SiC layer, insulating carbon matrix, and outer carbon matrix effectively improves the structural integrity and performance retention of the carbon fiber and PyC interface before and after silicon infiltration. The optimized vapor-phase silicon infiltration process significantly improves the prepared C... f Mechanical strength of C-SiC composite materials.

[0104] Figure 10 The X-ray diffraction analysis results are shown for the sample prepared in Example 3 and the sample prepared by the existing vapor phase silicon infiltration technique. The red and blue curves represent the test results of the two samples prepared in Example 3, while the green curve represents the test results of the sample prepared by the existing vapor phase silicon infiltration technique. It can be seen that the matrix of the material prepared by the existing technique contains a large amount of residual silicon, resulting in a high silicon signal intensity. In contrast, the silicon intensity peak of the sample prepared by the present invention is almost invisible, proving that the matrix contains almost no residual silicon, which also confirms... Figure 9 The mechanical property test results.

Claims

1. A method for preparing a ceramic matrix composite material, characterized in that, Includes the following steps: (1) Preform preparation Carbon fiber preforms were prepared using continuous carbon fibers as reinforcement. (2) Fiber interface coating treatment The carbon fiber preform obtained in step (1) is placed in a high-temperature vacuum sintering furnace for isothermal chemical vapor deposition process to obtain a preform with a pyrolytic carbon interface layer. (3) Introduction of CVI-SiC layer matrix The preform obtained in step (2) was initially densified using an isothermal chemical vapor deposition process to introduce a silicon carbide matrix, resulting in C f / SiC preform; C obtained by introducing silicon carbide matrix f The density of the SiC green body is 0.9~1.1 g / cm³. 3 ; (4) Introduction of carbon into the matrix of the isolation layer The C obtained in step (3) is produced by isothermal chemical vapor deposition process. f A pyrolytic carbon matrix was introduced into the SiC preform to obtain a density of 1.2~1.3 g / cm³. 3 Secondary billet; (5) Introduction of carbon into the outer matrix An outer layer of resin carbon matrix was further introduced into the secondary preform obtained in step (4) using a pressure impregnation-pyrolysis process to obtain a preform with a density of 1.4~1.5 g / cm³. 3 The third-stage billet; (6) Using silicon powder as raw material, the three-stage preform obtained in step (5) is subjected to silicon diffusion treatment using a vapor phase silicon diffusion process to finally obtain C f / C-SiC composite material, with a density of 1.9~2.2 g / cm³. 3； Step (6) specifically involves: The three-stage blank prepared in step (5) is placed in a mold, with silicon powder spread on the bottom of the mold, and the blank is raised to a certain height so as not to directly contact the silicon powder. After the mold is sealed, it is placed in a high-temperature sintering furnace and heated under argon protection at a rate of 5-8℃ / min. When the temperature reaches 1450℃, it is held for half an hour. The temperature is then raised to 1600~1800℃ and held for 0.5-2 hours. The pressure during the holding period is 1000~3000Pa. After the pressure holding period, the temperature is raised to 1850℃ and held for 0.5 hours. At the same time, the vacuum system of the high-temperature sintering furnace works continuously during the holding period to maintain the pressure at 1000~1500Pa. The mold includes a cylindrical body and a top cover; The cylinder is made of graphite, and the inner wall of the cylinder is uniformly coated with boron nitride powder-varnish dispersant and then dried at 120°C; a graphite sample support frame is installed at the bottom of the cylinder. The mold cover has a three-layer structure. The innermost layer is high-purity flexible graphite paper with a thickness of 0.5~1mm. The next layer is carbon felt, graphite felt, hard carbon felt, or hard graphite felt with a thickness of 1~5mm. The outermost layer is an isostatically pressed graphite plate. The three-stage blanks are placed in the mold as follows: The mass ratio of the third-stage preform to silicon powder is 1:2 to 1:4; the bottom surface of the third-stage preform is placed 10 to 25 mm away from the upper surface of the silicon powder after being placed on the graphite sample support frame. When multiple tertiary blanks are placed on the graphite sample support frame, the edge distance between two adjacent tertiary blanks is greater than 20mm; the area of ​​the bottom surface of the tertiary blank covered by the graphite sample support frame in the mold is less than one-fifth of the bottom area of ​​the tertiary blank.

2. The method for preparing a ceramic matrix composite material according to claim 1, characterized in that, The carbon fiber preform employs weaving methods such as 2.5D needle punching, three-dimensional orthogonal weaving, fine weaving puncture, and ply-lay stitching. The carbon fiber volume fraction in the preform is 30-45%, and the initial density of the preform is 0.4-0.9 g / cm³. 3 ; Before step (2), the carbon fiber preform prepared in step (1) is subjected to a degumming pretreatment: The woven preform is degummed at high temperature and kept at 900°C for two hours in a nitrogen atmosphere.

3. The method for preparing a ceramic matrix composite material according to claim 1, characterized in that, Step (2) specifically involves: The carbon fiber preform obtained in step (1) was placed in a high-temperature vacuum sintering furnace and deposited using natural gas as the carbon source gas and nitrogen as the dilution and protective gas. The deposition temperature was 1050~1120℃, the deposition time was 10~15 hours, and the deposition pressure was 3000~6000Pa. The natural gas flow rate was 0.5~1.1L / min, and the nitrogen flow rate was 2~4L / min. In step (2), the thickness of the pyrolytic carbon interface is 500~1200 nm.

4. The method for preparing a ceramic matrix composite material according to claim 1, characterized in that, Step (3) is as follows: The preform prepared in step (2) is placed in a chemical vapor deposition furnace, with trichloromethylsilane as the silicon and carbon source gas, hydrogen as the carrier gas and dilution gas, and argon as the protective gas, and isothermal chemical vapor deposition is carried out. The deposition temperature is 1000~1100℃, the deposition time is 30~60 hours, and the deposition pressure is 1000~2000Pa. The flow rate ratio of the reacting gases is 1:30:30 for trichloromethylsilane, hydrogen, and argon.

5. A method for preparing a ceramic matrix composite material according to any one of claims 1 to 4, characterized in that, Step (4) is as follows: The C prepared in step (3) f The SiC preform is placed in a high-temperature sintering furnace, with propane as the carbon source gas and nitrogen as the carrier gas, and isothermal chemical vapor deposition is carried out. The deposition temperature is 980~1100℃, the deposition time is 15-20 hours, the pressure is 1000~1500Pa, and the flow ratio of propane to nitrogen is 1:

4.

6. The method for preparing a ceramic matrix composite material according to claim 1, characterized in that, Step (5) is as follows: Add the secondary preform obtained in step (4) into a pressure impregnation tank, evacuate and maintain pressure for half an hour, pour the preform mixture after heat preservation into the impregnation tank and pressurize it to 5 MPa with inert gas, and maintain pressure for 1 hour; put the impregnated secondary preform into an oven and cure it at 120°C for 4 hours, and then place the secondary preform in a vacuum pyrolysis furnace for vacuum high-temperature pyrolysis at 900°C. The vacuum pyrolysis heating rate is: 8°C per hour below 600°C, and 20°C per hour above 600°C. The preparation method of the precursor mixture is as follows: A precursor mixture was prepared by thoroughly mixing furan resin or phenolic resin as a precursor, phosphoric acid as a curing agent, and anhydrous ethanol as a solvent. The mass ratio of resin, phosphoric acid, and anhydrous ethanol was 4:1:1~4. The prepared precursor mixture was placed in a 60℃ oven and kept warm for half an hour to increase its fluidity.

7. A ceramic matrix composite material for thermal structural components of aircraft, prepared by the method according to any one of claims 1 to 6.

Images