A kind of ultra-high temperature ceramic modified C / C composite material for aerospace power and its preparation method

Abstract

A C / C composite material modified with ultra-high temperature ceramics for aerospace propulsion and its preparation method improves the oxidation and ablation resistance of the C / C composite material by introducing a multiphase ultra-high temperature ceramic matrix of ZrC and SiBCN into the C / C blank. The C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material is prepared by subjecting a carbon fiber preform to high-temperature heat treatment, chemical vapor phase directional deposition of a carbon matrix process, and high-pressure multiphase precursor uniform mixing, impregnation, and co-cracking densification, followed by a post-high-temperature heat treatment process until the material density reaches 2.3g / cm 3 ~2.5g / cm 3 The specific preparation method is as follows: first, the carbon fiber preform is subjected to high-temperature heat treatment, and then the PyC interface deposition and matrix carbon preparation are carried out by chemical vapor deposition process, and then the ZrC ceramic matrix and SiBCN matrix are prepared, and then the formed composite material is subjected to post-high-temperature heat treatment, and finally the SiC surface coating is carried out by chemical vapor deposition to obtain a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material for aerospace propulsion, which has the advantages of ultra-high temperature resistance, low expansion coefficient, excellent oxidation resistance and ablation performance.

Description

Technical Field

[0001] The present invention relates to an ultra-high temperature ceramic modified C / C composite material for aerospace propulsion and a preparation method thereof, belonging to the field of ultra-high temperature ceramic-based composite material preparation. Background Art

[0002] Hypersonic aerospace vehicles, represented by near-space vehicles and highly maneuverable, long-range strategic missiles, can effectively enhance the high maneuverability of weapon systems and their long-range precision strike capabilities, and have become a key development direction in the military aerospace field. Carbon / carbon (C / C) composites, characterized by high strength, ablation resistance, and excellent high-temperature mechanical properties, have become one of the most influential systems in thermal protection materials. However, their poor oxidation resistance makes them difficult to withstand the long-duration, reusable aerobic service environments of the new generation of hypersonic vehicles. By introducing an ultra-high-temperature ceramic phase, the oxidation and ablation resistance of C / C composites has been improved, broadening their application in the field of thermal protection and making them a new material for thermal protection of aerospace powered vehicles.

[0003] In the preparation process of C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite materials, ZrC precursor impregnation solution and SiBCN precursor impregnation solution are usually used to densify the material by alternating impregnation and cracking. However, due to the uneven distribution of the content of ZrC and SiBCN multiphase matrix inside the carbon fiber preform during the alternating impregnation process of different precursors, the thickness of the multilayer matrix is ​​uncontrollable, resulting in the formation of ZrC matrix and SiC, BN and Si3N4 matrix in the SiBCN multiphase matrix unable to synergistically play an effective antioxidant protection for the carbon fiber. Moreover, different precursors are prone to cause surface sealing of the material during the long-term alternating and repeated impregnation process, and the internal pores cannot be densified in the later stage, which prolongs the preparation cycle and reduces the material performance. Moreover, the ZrC precursor will first produce ZrO2 when it changes from organic to inorganic during the cracking process. In order to avoid the reaction between ZrO2 and carbon fiber (C f ) or the pyrolytic carbon interface (PyC) causes a decrease in mechanical properties. Therefore, an appropriate amount of carbon matrix is ​​required to undergo carbothermal reduction reaction with ZrO2 to ensure the generation of a controllable content of ZrC. Summary of the Invention

[0004] In response to the problems existing in the preparation process of the above-mentioned C / C-ZrC-SiBCN ultra-high temperature ceramic-based composite materials, such as low final material density, long preparation cycle, damage to carbon fiber and PyC by ZrC precursor during cracking, uneven distribution of ZrC matrix and SiBCN multiphase ceramics, and uncontrollable matrix content, the present invention provides a method for preparing an ultra-high temperature ceramic-modified C / C composite material for aerospace propulsion, which has a short preparation cycle, little damage to the fibers, uniform distribution of the ceramic matrix, controllable content, and excellent anti-oxidation and ablation properties.

[0005] The present invention provides a preparation method for modifying a C / C composite material by using a ZrC and SiBCN multiphase ultra-high temperature ceramic matrix. The chemical vapor phase directional carbon matrix deposition process is used to force the generated carbon matrix to quickly and uniformly diffuse into various areas inside the preform, thereby providing sufficient carbon matrix for the subsequent ZrO2 carbothermal reduction reaction, thereby avoiding the damage to C f The ZrC and SiBCN multiphase matrix is ​​uniformly distributed using a high-pressure, multiphase precursor impregnation and co-cracking process, enabling the controlled preparation of high-performance composite materials. This significantly improves densification efficiency and final density. Furthermore, the ZrC and SiBCN multiphase ultra-high-temperature ceramic matrix prepared by low-temperature cracking of the precursor avoids fiber damage caused by the high-temperature reactive infiltration process. A post-high-temperature heat treatment eliminates residual stresses between the material components, improving the overall performance of the material.

[0006] The present invention also aims to provide an ultra-high temperature ceramic modified C / C composite material for aerospace propulsion prepared by the above method.

[0007] The technical solution of the present invention comprises the following steps:

[0008] Step 1: Preparation of carbon fiber preform: Due to the harsh and complex service environment that aerospace power engine thermal structures are subject to, such as ultra-high temperature (>2000°C), large thermal gradients and thermal stresses, and severe erosion of large amounts of solid particles such as Al2O3, the present invention uses a preform woven with T1000 carbon fiber in a three-dimensional orthogonal manner, wherein the spacing between two adjacent layers of carbon fiber cloth is 2mm×2mm, and the carbon fiber content in the carbon fiber preform is 40% to 45%;

[0009] Step 2: High-temperature heat treatment of the carbon fiber preform: First, use compressed air to blow away surface dust from the carbon fiber preform. Then, place it in an ultrasonic cleaner for 30 to 60 minutes, and then place it in an oven to dry at 100 to 120 degrees Celsius for 2 hours. Heat treatment is then performed in a high-temperature furnace at 2300 to 2400 degrees Celsius for 1 to 2 hours, using argon as a protective gas at a pressure of 3000 to 5000 Pa.

[0010] Step 3: Preparation of PyC interface: First, a PyC interface of a specific thickness is deposited on the heat-treated fiber preform using a chemical deposition method.

[0011] Step 4: Preparation of pyrolytic carbon matrix: Through the chemical vapor phase directional deposition carbon matrix process, the deposition gas flow field is precisely controlled to force the deposition gas to quickly and directionally diffuse to various areas inside the preform, thereby efficiently and controllably depositing the pyrolytic carbon matrix on the preform fiber surface. After deposition, the C / C green body density is 1.2-1.3 g / cm 3During the deposition process, a graphite column fixture for directional deposition of the carbon matrix is ​​placed at each gas outlet in the deposition furnace. The carbon fiber preform is then placed within the column, forcing the gas to flow from the inside of the column to the outside, thereby achieving uniform deposition within the preform. This allows the generated carbon matrix to quickly and uniformly diffuse into various areas within the carbon fiber preform, providing sufficient carbon matrix for the subsequent ZrO2 carbothermal reduction reaction and avoiding erosion damage to Cf and PyC.

[0012] Step 5: Preparation of ZrC ceramic matrix and SiBCN matrix: Using high-pressure multiphase precursor uniform mixing impregnation co-cracking process, the C / C green body is densified into ultra-high temperature ceramic matrix. After repeated impregnation and cracking, a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material with a density of 2.0-2.3 g / cm 3 .

[0013] Step 6. Post-high temperature heat treatment: Due to the differences in thermal expansion coefficients of carbon fiber, PyC interface, carbon matrix, ZrC ceramic matrix and SiBCN matrix, in order to release the residual stress between the components, the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material needs to be post-high temperature heat treated in a nitrogen atmosphere.

[0014] Step 7: Preparation of SiC surface coating: Chemical vapor deposition is used to deposit SiC coating on the surface of C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material for surface sealing. Finally, C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material is obtained with a density of 2.3-2.5 g / cm 3 .

[0015] Preferably, in step 3, the PyC interface preparation method comprises placing the fiber preform from step 2 in a chemical deposition apparatus and performing pyrolytic carbon deposition using natural gas as the carbon source and nitrogen as the diluent gas. Deposition process parameters are: deposition temperature of 950°C to 980°C, natural gas flow rate of 10 L / h to 20 L / h, nitrogen flow rate of 10 L / h to 20 L / h, deposition time of 10 to 15 hours, and deposition pressure of 1000 Pa to 1500 Pa.

[0016] Preferably, in step 4, the preparation method of the pyrolytic carbon matrix is ​​as follows: the preform in step 3 is placed in a chemical deposition device, and pyrolytic carbon deposition is performed using natural gas as the carbon source and nitrogen as the diluent gas. The deposition process parameters are as follows: deposition temperature of 1000°C to 1100°C, natural gas flow rate of 30L / h to 50L / h, nitrogen flow rate of 30L / h to 50L / h, deposition time of 20 to 30h, deposition pressure of 2000Pa to 3000Pa, and the material density of the pyrolytic carbon matrix after deposition is 1.2 to 1.3g / cm 3 .

[0017] Preferably, in step five, the ZrC ceramic matrix and SiBCN matrix preparation method are as follows: first, a mixed impregnation solution of polycarbonazirane and polyborosilazane is prepared. The conversion rate of the ZrC and SiBCN matrices after cracking is obtained by thermogravimetric analysis based on the PyC interface thickness, the density of the pyrolyzed carbon matrix, and the polycarbonazirane and polyborosilazane. The polycarbonazirane and polyborosilazane are precisely proportioned and then stirred and mixed, wherein xylene is used as a solvent, wherein the mass ratio of xylene to the total mass of polycarbonazirane and polyborosilazane is 1:1. The stirring time during the preparation of the mixed impregnation solution of polycarbonazirane and polyborosilazane is 6 to 8 hours; then, the C / C blank is subjected to high-pressure impregnation. The C / C blank is placed in an impregnation device, and then vacuumed to <100Pa. The mixed impregnation solution of polycarbonazirane and polyborosilazane is added to the device, and nitrogen is introduced to a pressure of 3MPa to 5MPa, and pressurized impregnation is performed for 2 hours. After the pressure impregnation is completed, it is placed in an oven for curing for 3 hours. The curing temperature is 250℃~300℃. After the curing is completed, it is placed in a cracking furnace for cracking to obtain ZrC and SiBCN ceramic matrices. The cracking temperature is 1400℃~1600℃, the heating rate is 5~10℃ / min, the holding time is 1h~3h, and nitrogen is introduced as a protective gas. Repeat the above high-pressure multiphase precursor uniform mixing impregnation co-cracking process until the C / C green body density reaches 2.0~2.3g / cm 3 The C / C-ZrC-SiBCN ultrahigh-temperature ceramic matrix composite was obtained. In previous studies, thermogravimetric analysis of polycarbozircane and polyborosilazane was performed to determine the mass of ZrC matrix that could be converted from polycarbozircane at a cracking temperature of 1400°C to 1600°C. The ZrC conversion rate was then calculated by calculating the ratio of the ZrC mass to the mass of polycarbozircane. Similarly, the mass of SiBCN matrix that could be converted from polyborosilazane at a cracking temperature of 1400°C to 1600°C was measured, and the SiBCN matrix conversion rate of polyborosilazane was similarly determined. Based on the matrix conversion rates of the two substances, the required mass of polycarbozircane and polyborosilazane, respectively, could be calculated during the formulation process to ensure that the mass ratio of the ZrC ceramic matrix to the SiBCN matrix in the C / C-ZrC-SiBCN ultrahigh-temperature ceramic matrix composite was 7:3 to 6:4.

[0018] Preferably, in step six, the post-high temperature heat treatment method is: placing the C / C-ZrC-SiBCN ultra-high temperature ceramic-based composite material into a high temperature furnace, the heat treatment temperature is 1600℃~1800℃, the heat treatment time is 2-3h, nitrogen is introduced as a protective gas, the pressure is maintained at 3000Pa~5000Pa, the heating rate is 60℃ / h~80℃ / h, and the cooling rate is controlled at 50℃ / h~80℃ / h during the cooling process.

[0019] Preferably, in step 7, the SiC surface coating preparation method comprises placing the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material in a chemical deposition apparatus, using trichloromethylsilane as a precursor, hydrogen as a carrier gas and a reaction gas, and argon as a diluent gas to deposit SiC. The deposition process parameters are: deposition temperature of 1000°C to 1200°C, trichloromethylsilane:hydrogen:argon flow ratio of 1:20:30, deposition time of 25 hours to 40 hours, and deposition pressure of 1000 Pa to 2000 Pa.

[0020] An ultra-high temperature ceramic modified C / C composite material for aerospace propulsion is prepared using the method described in the above technical steps.

[0021] The present invention adopts a combination of specific carbon fiber reinforcement and ceramic matrix based on the characteristics of ultra-high temperature (>2000℃), extremely large thermal gradients and thermal stresses, and severe erosion of large amounts of solid particles such as Al2O3 faced by ultra-high temperature composite materials during the service of near-space aircraft. The preparation of high-performance composite materials is achieved through a unique chemical vapor phase directional deposition process and a high-pressure multiphase precursor uniform mixing, impregnation and co-cracking combined process. The expected performance improvement can be achieved through their specific combination and precise ratio as well as the control of specific parameters such as temperature and pressure in the preparation process.

[0022] The present invention innovatively adopts a chemical vapor phase directional deposition carbon matrix process, a high-pressure multiphase precursor uniform mixing impregnation co-cracking process and a post-high temperature heat treatment process to prepare a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material.

[0023] Since there are multiple air inlets evenly distributed in the vapor deposition furnace, if the direction of gas flow is not controlled, the deposition uniformity and efficiency of the carbon matrix will be seriously affected. Therefore, a chemical vapor phase directional deposition carbon matrix process is adopted. By placing a graphite column fixture at each air inlet, and then placing the carbon fiber preform and the carbon fiber porous body in the graphite column fixture, the gas is forced to flow directionally from the inside to the outside of the graphite column fixture, so that the generated carbon matrix is ​​quickly and uniformly diffused to various areas inside the carbon fiber preform, thereby achieving efficient and controllable deposition of pyrolytic carbon matrix in the carbon preform and the carbon fiber porous body. After the reaction gas is introduced, a pressure difference will be formed inside and outside the graphite column fixture, forcing the reaction gas to diffuse evenly inside the carbon fiber porous body, thereby achieving the purpose of rapid and uniform densification of the carbon fiber porous body and shortening the preparation cycle. The chemical vapor phase directional deposition carbon matrix process is used to force the generated carbon matrix to quickly and uniformly diffuse to various areas inside the preform, providing sufficient carbon matrix for the subsequent ZrO2 carbon thermal reduction reaction, avoiding the C f and erosion damage of PyC.

[0024] A high-pressure, multiphase precursor-based, uniformly mixed, impregnated, and co-pyrolysis process achieves uniform distribution of the ZrC matrix and SiBCN multiphase matrix, enabling the controlled preparation of high-performance composite materials. This significantly improves the densification efficiency and final density, while shortening the preparation cycle. During the preparation of C / C-ZrC-SiBCN ultrahigh-temperature ceramic-based composites, internal stresses form within the material during repeated heating and cooling due to the different thermal expansion coefficients of the carbon fibers, the PyC interface, the carbon matrix, the ZrC matrix, and the SiBCN matrix. To eliminate residual thermal stresses between the components, the C / C-ZrC-SiBCN ultrahigh-temperature ceramic-based composite requires post-high-temperature heat treatment. This post-high-temperature heat treatment eliminates residual stresses between the material components. Furthermore, the ZrC and SiBCN multiphase ultrahigh-temperature ceramic matrix prepared by low-temperature pyrolysis of the precursor avoids fiber damage caused by the high-temperature reactive infiltration process, thereby improving the material's overall oxidation and ablation resistance and mechanical properties. The temperature of low-temperature pyrolysis is 1400℃~1600℃, compared with other preparation processes such as reactive infiltration process which requires a temperature of above 1900℃.

[0025] The ultra-high temperature ceramic-based composite material prepared by the present invention has a ZrC matrix and a SiBCN multiphase ultra-high temperature ceramic structure. When the material is subjected to stress, cracks will deflect between different interfaces and the multiphase matrix, greatly improving the toughness of the material, the mechanical properties of the material, and the oxidation resistance and ablation resistance of the C / C composite material.

[0026] The preparation process of the present invention is simple, easy to operate, and has good process applicability. The prepared ultra-high temperature composite material has excellent mechanical properties, oxidation resistance and ablation resistance, and has broad application prospects in the hot end components of aerospace power ultra-high-speed aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 This is a structural diagram of the directional deposition carbon matrix graphite column tooling of the present invention.

[0028] Figure 2 This is a flow chart for preparing an ultra-high temperature ceramic modified C / C composite material for aerospace propulsion according to the present invention.

[0029] Figure 3 This is a fracture morphology of the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material prepared in comparative example.

[0030] Figure 4 This is a microscopic morphology of the interior of the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material prepared in Example 2. DETAILED DESCRIPTION

[0031] In order to make the purpose, technical solutions and advantages of the present invention clearer, the technical solutions of the present invention will be clearly and completely described in conjunction with the embodiments and drawings. It is obvious that the described embodiments are only part of the embodiments of the present invention, rather than all the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention.

[0032] Preparation method of ultra-high temperature ceramic modified C / C composite material for aerospace power (such as Figure 2 ), including the following steps:

[0033] (1) Preparation of carbon fiber preform

[0034] The carbon fiber preform is a preform made of T1000 carbon fiber three-dimensional orthogonal weaving according to the structure of the aerospace power hot end component. The volume density of the carbon fiber preform is 0.85g / cm 3 ~0.95g / cm 3 The fiber content in the carbon fiber preform is 40% to 45%.

[0035] (2) High temperature heat treatment

[0036] The carbon fiber preform is subjected to a high temperature heat treatment at 2300°C to 2400°C in a high temperature furnace for 1 hour to 2 hours, argon is introduced as a protective gas, and the pressure is maintained at 3000Pa to 5000Pa;

[0037] (3) Preparation of PyC interface

[0038] A PyC interface is deposited on a carbon fiber preform after high-temperature heat treatment by chemical deposition to obtain a carbon fiber porous body; the thickness of the PyC interface is 600nm to 800nm;

[0039] (4) Pyrolytic carbon matrix deposition

[0040] The carbon fiber porous body is subjected to chemical vapor phase directional deposition to form a pyrolytic carbon matrix. The deposition process parameters are as follows: natural gas is used as the carbon source and nitrogen is used as the diluent gas for pyrolytic carbon deposition; the deposition process parameters are as follows: deposition temperature is 1000℃~1100℃, natural gas flow rate is 30L / h~50L / h, nitrogen flow rate is 30L / h~50L / h, deposition time is 20~30h, deposition pressure is 2000Pa~3000Pa, and a C / C green body is obtained after pyrolytic carbon matrix deposition. The density of the C / C green body is 1.2~1.3g / cm 3During the deposition process, a graphite column (referred to as column) 4 for directional deposition of carbon matrix is ​​placed at each gas outlet 1 in the deposition furnace chamber 5, and then the carbon fiber preform 3 is placed in the column, thereby forcing the gas 2 to flow from the inside of the column to the outside, thereby achieving the purpose of uniform deposition in the carbon fiber preform. The generated carbon matrix is ​​quickly and uniformly diffused to various areas inside the carbon fiber preform, providing sufficient carbon matrix for the subsequent ZrO2 carbothermal reduction reaction, avoiding the carbon matrix. f and PyC erosion damage. The graphite column is a cylindrical tooling with only one air inlet on the bottom and evenly distributed small air outlets on the top. (such as Figure 1 )

[0041] (5) A high-pressure multiphase precursor uniform mixing impregnation co-cracking process is adopted to densify the ZrC-SiBCN ultra-high temperature ceramic matrix of the C / C green body, and a method for preparing a mixed impregnation solution of polycarbon zirconium alkane and polyborosilazane is prepared: the mass ratio of the ZrC matrix and the SiBCN matrix in the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material is 7:3 to 6:4, and the respective amounts of polycarbon zirconium alkane and polyborosilazane are calculated and weighed according to the ZrC conversion rate in polycarbon zirconium alkane and the SiBCN matrix conversion rate in polyborosilazane obtained by the previous thermogravimetric analysis, and then the weighed polycarbon zirconium alkane and polyborosilazane are added to the xylene solvent and stirred to obtain a mixed impregnation solution of polycarbon zirconium alkane and polyborosilazane, wherein the mass ratio of xylene to the total mass of polycarbon zirconium alkane and polyborosilazane is 1:1, and the polycarbon zirconium alkane and polyborosilazane are prepared. The stirring time during the process of the mixed impregnation liquid of alkane is 6 to 8 hours; then the C / C green body is subjected to high-pressure impregnation; the C / C green body is placed in an impregnation device, and then vacuumed to <100Pa, the mixed impregnation liquid of polycarbonazirane and polyborosilazane is added to the impregnation device, nitrogen is introduced to the pressure of 3MPa to 5MPa, and pressure impregnation is performed for 2h; after the pressure impregnation is completed, it is placed in an oven for curing for 3h; the curing temperature is 250℃ to 300℃; after the curing is completed, it is placed in a cracking furnace for cracking to obtain ZrC and SiBCN ceramic matrix; the cracking temperature is 1400℃ to 1600℃, the heating rate is 5 to 10℃ / min, the holding time is 1h to 3h, and nitrogen is introduced as a protective gas; the above high-pressure multiphase precursor uniform mixing impregnation and co-cracking process is repeated multiple times until the density of the C / C green body reaches 2.0 to 2.3g / cm 3 Obtain C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material;

[0042] (6) Post-high temperature heat treatment

[0043] The post-heat treatment temperature process parameters are as follows: heat treatment temperature 1600℃~1800℃, heat treatment time 2h~3h, nitrogen is introduced as protective gas, pressure is maintained at 3000Pa~5000Pa, heating rate is 60℃ / h~80℃ / h, and cooling rate is controlled at 50℃ / h~80℃ / h during cooling process;

[0044] (7) Preparation of SiC surface coating

[0045] The SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material by chemical vapor deposition to seal the surface. Finally, the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material was obtained with a density of 2.3-2.5 g / cm 3 .

[0046] The present invention will be further described below with reference to comparative examples and embodiments.

[0047] Comparative Example 1

[0048] This comparative example prepares a C / ZrC-SiC composite material, and the preparation method includes the following specific steps:

[0049] (1) Preparation of carbon fiber preform: The carbon fiber preform is woven in a three-dimensional orthogonal manner using T1000 carbon fibers. The preform volume density is 0.9 g / cm 3 , the fiber content in the carbon fiber preform is 40%.

[0050] (2) High-temperature heat treatment of carbon fiber preforms: The carbon fiber preforms were first blown with compressed air to remove surface dust, then ultrasonically cleaned in an ultrasonic cleaner for 30 min, and then dried in an oven at 120°C for 2 h. Heat treatment was then performed in a high-temperature furnace at 2300°C for 2 h, with argon as the protective gas at a pressure of 3000 Pa.

[0051] (3) Preparation of PyC interface: The PyC interface was deposited on the heat-treated carbon fiber preform by chemical deposition to obtain a carbon fiber porous body. The deposition parameters were as follows: deposition temperature of 950 °C, natural gas flow rate of 10 L / h, nitrogen flow rate of 15 L / h, deposition time of 15 h, deposition pressure of 1500 Pa, and the thickness of the PyC interface after deposition was 700 nm.

[0052] (4) Preparation of ZrC ceramic matrix and SiBCN matrix: First, a mixed impregnation solution of polycarbonazirane and polyborosilazane in xylene was prepared, wherein the mass ratio of ZrC matrix to SiBCN matrix was 6:4, the mass ratio of xylene to the total mass of polycarbonazirane and polyborosilazane was 1:1, and the stirring time was 6 hours; then, the C / C blank was subjected to high-pressure impregnation; the C / C blank was placed in an impregnation device, and then vacuumed to 50 Pa, and the mixed impregnation solution of polycarbonazirane and polyborosilazane was added. The product was added to the impregnation equipment, nitrogen was introduced to the pressure of 3MPa, and pressure impregnation was performed for 2h. After the pressure impregnation was completed, it was placed in an oven for curing for 3h. The curing temperature was 250℃. After the curing was completed, it was placed in a cracking furnace for cracking. The cracking temperature was 1400℃, the heating rate was 5℃ / min, the holding time was 1h, and nitrogen was introduced as a protective gas. After 8 impregnation and cracking, a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material with a density of 1.95g / cm was obtained. 3 .

[0053] (7) Preparation of SiC Surface Coating: A SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material by chemical vapor deposition for surface sealing. The deposition parameters were as follows: deposition temperature of 1000°C, trichloromethylsilane: hydrogen: argon flow ratio of 1:20:30, deposition time of 30 h, and deposition pressure of 1100 Pa. Finally, a C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material was obtained.

[0054] The C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material prepared in this comparative example has a mass of 2.1 g / cm 3 The room temperature flexural strength is only 153.9 MPa. When the composite material is ablated at 2100℃ for 100s, the linear ablation rate is 0.64μm / s. Figure 3 The fracture morphology shows that due to the lack of introduction of sufficient carbon matrix and post-high-temperature heat treatment, the interface between carbon fiber and PyC is severely corroded during the cracking process. The bond between the matrix and the interface is strong and the internal stress is not released, resulting in useless long fiber pullout and low mechanical properties.

[0055] Example 1:

[0056] (1) Preparation of carbon fiber preform: The carbon fiber preform is woven in a three-dimensional orthogonal manner using T1000 carbon fibers. The preform volume density is 0.92 g / cm 3 , the fiber content in the carbon fiber preform is 40%.

[0057] (2) High-temperature heat treatment of carbon fiber preforms: The carbon fiber preforms were first blown with compressed air to remove surface dust, then ultrasonically cleaned in an ultrasonic cleaner for 30 min, and then dried in an oven at 120°C for 2 h. Heat treatment was then performed in a high-temperature furnace at 2300°C for 2 h, with argon as the protective gas at a pressure of 3000 Pa.

[0058] (3) Preparation of PyC interface: The PyC interface was deposited on the heat-treated carbon fiber preform by chemical deposition to obtain a carbon fiber porous body. The deposition parameters were as follows: deposition temperature of 950 °C, natural gas flow rate of 10 L / h, nitrogen flow rate of 10 L / h, deposition time of 10 h, deposition pressure of 1000 Pa, and the thickness of the PyC interface after deposition was 600 nm.

[0059] (4) Preparation of pyrolytic carbon matrix: The carbon fiber porous body was placed in a directional deposition carbon matrix graphite column in a chemical deposition furnace chamber, and the pyrolytic carbon matrix was subjected to chemical vapor phase directional deposition. The deposition parameters were: deposition temperature of 1050 °C, natural gas flow rate of 30 L / h, nitrogen flow rate of 30 L / h, deposition time of 20 h, deposition pressure of 2000 Pa, and the density of the C / C blank after deposition was 1.2 g / cm 3 .

[0060] (5) Preparation of ZrC ceramic matrix and SiBCN matrix: First, a mixed impregnation solution of polycarbonazirane and polyborosilazane in xylene was prepared, wherein the mass ratio of the ZrC matrix to the SiBCN matrix was 6:4, the mass ratio of xylene to the total mass of polycarbonazirane and polyborosilazane was 1:1, and the stirring time was 6 hours; then, the C / C blank was subjected to high-pressure impregnation; the C / C blank was placed in an impregnation device, and then vacuumed to 60 Pa, and the mixed impregnation solution of polycarbonazirane and polyborosilazane was added. The mixture was added to the impregnation equipment, nitrogen was introduced to the pressure of 3 MPa, and pressure impregnation was performed for 2 hours. After the pressure impregnation was completed, the mixture was placed in an oven for curing for 3 hours at a curing temperature of 250°C. After the curing was completed, the mixture was placed in a cracking furnace for cracking. The cracking temperature was 1400°C, the heating rate was 5°C / min, the holding time was 1 hour, and nitrogen was introduced as a protective gas. After 8 impregnation and cracking cycles, a C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material with a density of 2.0 g / cm was obtained. 3 .

[0061] (6) Post-high temperature heat treatment: The C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material was subjected to post-high temperature heat treatment in a nitrogen atmosphere. The heat treatment temperature was 1600°C, the heat treatment time was 2 h, the pressure was maintained at 3000 Pa, the heating rate was 60°C / h, and the cooling rate was controlled at 50°C / h during the cooling process.

[0062] (7) Preparation of SiC Surface Coating: A SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material by chemical vapor deposition for surface sealing. The deposition parameters were as follows: deposition temperature of 1000°C, trichloromethylsilane: hydrogen: argon flow ratio of 1:20:30, deposition time of 25 h, and deposition pressure of 1000 Pa. Finally, a C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material was obtained.

[0063] The C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material prepared in this embodiment has a mass of 2.3 g / cm 3 The room temperature flexural strength is 223.7 MPa. The composite material is ablated at 2100℃ for 100s, and the linear ablation rate is 0.45μm / s.

[0064] Example 2:

[0065] (1) Preparation of carbon fiber preform: The carbon fiber preform is woven in a three-dimensional orthogonal manner using T1000 carbon fibers. The preform volume density is 0.95 g / cm 3 , the fiber content in the fiber preform is 45%.

[0066] (2) High-temperature heat treatment of carbon fiber preforms: The carbon fiber preforms were first blown with compressed air to remove surface dust, then ultrasonically cleaned in an ultrasonic cleaner for 30 min, and then dried in an oven at 120°C for 2 h. Heat treatment was then performed in a high-temperature furnace at 2300°C for 2 h, with argon as the protective gas at a pressure of 3000 Pa.

[0067] (3) Preparation of PyC interface: The PyC interface was deposited on the heat-treated carbon fiber preform by chemical deposition to obtain a carbon fiber porous body. The deposition parameters were as follows: deposition temperature of 980 °C, natural gas flow rate of 15 L / h, nitrogen flow rate of 15 L / h, deposition time of 15 h, deposition pressure of 1500 Pa, and the thickness of the PyC interface after deposition was 800 nm.

[0068] (4) Preparation of pyrolytic carbon matrix: The carbon fiber porous body was placed in the directional deposition carbon matrix graphite column tooling in the chemical deposition furnace chamber, and the pyrolytic carbon matrix was subjected to chemical vapor phase directional deposition. The deposition parameters were: deposition temperature of 1100 °C, natural gas flow rate of 40 L / h, nitrogen flow rate of 40 L / h, deposition time of 30 h, deposition pressure of 3000 Pa, and the density of the C / C blank after deposition was 1.28 g / cm 3 .

[0069] (5) Preparation of ZrC ceramic matrix and SiBCN matrix: First, prepare a mixed impregnation solution of polycarbonazirane and polyborosilazane in xylene, wherein the mass ratio of ZrC matrix to SiBCN matrix is ​​7:3, the mass ratio of xylene to the total mass of polycarbonazirane and polyborosilazane is 1:1, and the stirring time is 8 hours; then, high-pressure impregnation is performed on the C / C blank; the C / C blank is placed in an impregnation device, and then vacuumed to 40 Pa, and the mixed impregnation solution of polycarbonazirane and polyborosilazane is added. The product was placed in an impregnation device, nitrogen was introduced to a pressure of 5 MPa, and pressure impregnation was performed for 2 hours. After the pressure impregnation was completed, the product was placed in an oven for curing for 3 hours at a curing temperature of 300°C. After the curing was completed, the product was placed in a cracking furnace for cracking. The cracking temperature was 1600°C, the heating rate was 5°C / min, the holding time was 2 hours, and nitrogen was introduced as a protective gas. After 10 impregnation and cracking cycles, a C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material with a density of 2.15 g / cm was obtained. 3 .

[0070] (6) Post-high temperature heat treatment: The C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material was subjected to post-high temperature heat treatment in a nitrogen atmosphere. The heat treatment temperature was 1800°C, the heat treatment time was 2 h, the pressure was maintained at 5000 Pa, the heating rate was 80°C / h, and the cooling rate was controlled at 80°C / h during the cooling process.

[0071] (7) Preparation of SiC Surface Coating: A SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material by chemical vapor deposition for surface sealing. The deposition parameters were as follows: deposition temperature of 1100°C, trichloromethylsilane: hydrogen: argon flow ratio of 1:20:30, deposition time of 30 h, and deposition pressure of 1000 Pa. Finally, a C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material was obtained.

[0072] The C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material prepared in this embodiment has a mass of 2.5 g / cm 3 The room temperature flexural strength is 233.7 MPa. The composite material is ablated at 2100℃ for 100s, and the linear ablation rate is 0.37μm / s. Figure 4 The internal microstructure shows that the ZrC matrix and SiBCN matrix with silvery white metallic luster are uniformly dispersed inside the preform. The matrix density is high and the mechanical properties and ablation properties are excellent compared with the control example.

Claims

1. A method for preparing an ultra-high temperature ceramic modified C / C composite material for aerospace propulsion, characterized by: The following steps are involved: (1) Preparation of carbon fiber preform The carbon fiber preform is a preform made of T1000 carbon fiber three-dimensional orthogonal weaving according to the structure of the aerospace power hot end component. The volume density of the carbon fiber preform is 0.85g / cm 3 ~0.95g / cm 3 , the fiber content in the carbon fiber preform is 40%~45%; (2) High temperature heat treatment The carbon fiber preform is subjected to a high-temperature heat treatment at 2300°C to 2400°C in a high-temperature furnace for 1 hour to 2 hours, with argon gas introduced as a protective gas and the pressure maintained at 3000Pa to 5000Pa; (3) Preparation of PyC interface A PyC interface is deposited on a carbon fiber preform after high-temperature heat treatment using a chemical deposition method to obtain a carbon fiber porous body; the thickness of the PyC interface is 600nm~800nm; (4) Pyrolytic carbon matrix deposition The carbon fiber porous body is subjected to chemical vapor phase directional deposition to form a pyrolytic carbon matrix. The deposition process parameters are as follows: natural gas is used as the carbon source and nitrogen is used as the diluent gas for pyrolytic carbon deposition. The deposition process parameters are as follows: deposition temperature of 1000℃~1100℃, natural gas flow rate of 30L / h~50L / h, nitrogen flow rate of 30L / h~50L / h, deposition time of 20~30h, deposition pressure of 2000Pa~3000Pa, and the C / C green body obtained after pyrolysis carbon matrix deposition. The density of the C / C green body is 1.2~1.3g / cm 3 The pyrolytic carbon matrix deposition process is to place the carbon fiber porous body in step (3) in a graphite column fixture for directional deposition of carbon matrix in a chemical deposition furnace chamber, wherein the graphite column fixture is a cylindrical fixture with only one air inlet on the bottom and evenly distributed small air outlets on the top; (5) A high-pressure multiphase precursor uniform mixing impregnation co-cracking process is adopted to densify the ZrC-SiBCN ultra-high temperature ceramic matrix of the C / C green body, and a method for preparing a mixed impregnation solution of polycarbon zirconium alkane and polyborosilazane is provided: the mass ratio of the ZrC matrix and the SiBCN matrix in the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material is 7:3~6:4, and the amount of polycarbon zirconium alkane and polyborosilazane is calculated and weighed according to the ZrC conversion rate in polycarbon zirconium alkane and the SiBCN matrix conversion rate in polyborosilazane obtained by the previous thermogravimetric analysis, and then the weighed polycarbon zirconium alkane and polyborosilazane are added to the xylene solvent and stirred to obtain a mixed impregnation solution of polycarbon zirconium alkane and polyborosilazane, wherein the mass ratio of xylene to the total mass of polycarbon zirconium alkane and polyborosilazane is 1:1, and the mixed impregnation solution of polycarbon zirconium alkane and polyborosilazane is prepared. The stirring time during the impregnation process of the mixed impregnation liquid of polycarbonate zirconium alkane is 6 to 8 hours; then the C / C green body is subjected to high-pressure impregnation; the C / C green body is placed in an impregnation device, and then vacuumed to <100Pa, the mixed impregnation liquid of polycarbonate zirconium alkane and polyborosilazane is added to the impregnation device, nitrogen is introduced to the pressure of 3MPa to 5MPa, and pressure impregnation is performed for 2h; after the pressure impregnation is completed, it is placed in an oven for curing for 3h; the curing temperature is 250℃ to 300℃; after the curing is completed, it is placed in a cracking furnace for cracking to obtain ZrC and SiBCN ceramic matrices; the cracking temperature is 1400℃ to 1600℃, the heating rate is 5 to 10℃ / min, the holding time is 1h to 3h, and nitrogen is introduced as a protective gas; the above high-pressure multiphase precursor uniform mixing impregnation and co-cracking process is repeated multiple times until the density of the C / C green body reaches 2.0 to 2.3g / cm 3 Obtain C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material; (6) Post-high temperature heat treatment The post-heat treatment temperature process parameters are as follows: heat treatment temperature 1600℃~1800℃, heat treatment time 2h~3h, nitrogen is introduced as protective gas, pressure is maintained at 3000Pa~5000Pa, heating rate is 60℃ / h~80℃ / h, and cooling rate is controlled at 50℃ / h~80℃ / h during cooling process; (7) Preparation of SiC surface coating The SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material by chemical vapor deposition to seal the surface. Finally, the C / C-ZrC-SiBCN ultrahigh temperature ceramic matrix composite material was obtained with a density of 2.3~2.5g / cm 3 .

2. The method for preparing the ultra-high temperature ceramic modified C / C composite material for aerospace propulsion according to claim 1, characterized in that: In step (1), carbon fiber is first woven into carbon fiber cloth, and the carbon fiber cloth is stacked layer by layer at 0° and 90°, and then carbon fiber is introduced in the vertical direction to finally form a carbon fiber preform, and the spacing between two adjacent layers of carbon fiber cloth is 2mm×2mm; before high-temperature heat treatment in step (2), the carbon fiber preform is first blown off the surface dust with compressed air, and then placed in an ultrasonic cleaning machine for ultrasonic cleaning for 30min~60min, and then placed in an oven at 100℃~120℃ for drying for 2h.

3. The method for preparing the ultra-high temperature ceramic modified C / C composite material for aerospace propulsion according to claim 1, characterized in that: In step (3), the process used for the PyC interface deposition is: natural gas as the carbon source, nitrogen as the dilution gas, the deposition temperature is 950°C~980°C, the natural gas flow rate is 10L / h~20L / h, the nitrogen flow rate is 10L / h~20L / h, the deposition time is 10h~15h, and the deposition pressure is 1000Pa~1500Pa.

4. The method for preparing the ultra-high temperature ceramic modified C / C composite material for aerospace propulsion according to claim 1, characterized in that: In step (7), the SiC surface coating preparation process is to place the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material in a chemical deposition device, use trichloromethylsilane as a precursor, hydrogen as a carrier gas and a reaction gas, and argon as a diluent gas to deposit SiC; The deposition process parameters are as follows: deposition temperature of 1000°C~1200°C, flow ratio of trichloromethylsilane: hydrogen: argon of 1:20:30, deposition time of 25h~40h, and deposition pressure of 1000Pa~2000Pa.

5. An ultra-high temperature ceramic modified C / C composite material for aerospace propulsion prepared by the method according to any one of claims 1 to 4.

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