Ultrahigh-temperature ceramic modified C / C composite material for aerospace power and preparation method of ultrahigh-temperature ceramic modified C / C composite material

Abstract

The invention discloses an ultra-high-temperature ceramic modified C / C composite material for aerospace power and a preparation method thereof. The oxidation ablation resistance of the C / C composite material is improved by introducing a ZrC and SiBCN multiphase ultra-high-temperature ceramic matrix into a C / C green body. The C / C-ZrC-SiBCN ultrahigh-temperature ceramic-based composite material is prepared by the following steps: carrying out high-temperature heat treatment on a carbon fiber preform, carrying out chemical vapor-phase oriented deposition on a carbon matrix process, carrying out densification by a high-pressure complex-phase precursor uniform mixing, dipping and co-cracking process, and carrying out post-high-temperature heat treatment until the material density reaches 2.3-2.5 g / cm < 3 >. The specific preparation method comprises the following steps: firstly, carrying out high-temperature heat treatment on a carbon fiber preform, then carrying out PyC interface deposition and matrix carbon preparation by adopting a chemical vapor deposition process, then preparing a ZrC ceramic matrix and a SiBCN matrix, and then carrying out post-high-temperature heat treatment on the formed composite material to obtain the ZrC / SiBCN composite material. Finally, SiC surface coating is conducted through chemical vapor deposition, and the C / C-ZrC-SiBCN ultrahigh-temperature ceramic-based composite material for aerospace power is obtained. The C / C-ZrC-SiBCN ultrahigh-temperature ceramic-based composite material has the advantages of being resistant to ultrahigh temperature, low in expansion coefficient, excellent in oxidation resistance and ablation performance and the like.

Description

Technical Field

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

[0002] Hypersonic aerospace vehicles, represented by near-space vehicles and highly maneuverable long-range strategic missiles, can effectively improve the high maneuverability of weapon systems and their long-range precision strike capabilities, and have become an important development direction in the military aerospace field. Carbon / carbon (C / C) composites have the characteristics of high strength, ablation resistance, and excellent high-temperature mechanical properties, and have become one of the most influential systems in thermal protection materials. However, their poor oxidation resistance makes it difficult to cope with the long-duration, reusable aerobic service environment of the new generation of hypersonic vehicles. By introducing ultra-high temperature ceramic phases, the oxidation and ablation properties of C / C composites are improved, broadening their application range in the field of thermal protection, and have become 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 composites, ZrC precursor impregnation liquid and SiBCN precursor impregnation liquid are usually used to densify the material by alternating impregnation and pyrolysis. However, due to the uneven content distribution of ZrC and SiBCN multiphase matrix inside the carbon fiber preform during the alternating impregnation 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 cannot synergistically play an effective antioxidant protection for carbon fiber. In addition, different precursors are prone to cause surface sealing of the material during long-term alternating and repeated impregnation, 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 pyrolysis process. In order to avoid ZrO2 and carbon fiber (C f ) or pyrolytic carbon interface (PyC) reaction causes the reduction of mechanical properties. Therefore, an appropriate amount of carbon matrix is ​​required to undergo carbothermal reduction reaction with ZrO2 to ensure the generation of ZrC with a controllable content. Summary of the invention

[0004] In view of 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 fibers, uniform distribution of 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, and adopts a chemical vapor phase directional carbon matrix deposition process to force the generated carbon matrix to quickly and uniformly diffuse to various regions inside the preform, thereby providing sufficient carbon matrix for the subsequent ZrO2 carbothermal reduction reaction, thereby avoiding the C f The ZrC matrix and SiBCN multiphase matrix are evenly distributed by the high-pressure multiphase precursor uniform mixing impregnation co-cracking process, and the high-performance content of the composite material can be controlled, which greatly improves the densification efficiency and final density. In addition, the ZrC and SiBCN multiphase ultra-high temperature ceramic matrix prepared by low-temperature pyrolysis of the precursor avoids the damage to the fiber caused by the high-temperature reaction infiltration process. The residual stress between the components of the material is eliminated by the post-high-temperature heat treatment process, which improves 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: According to the harsh and complex service environment that the thermal structure of aerospace power engines needs to face, such as ultra-high temperature (>2000°C), great thermal gradient and thermal stress, and severe erosion of a large number of solid particles such as Al2O3, the present invention adopts a preform woven by 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 carbon fiber preform: First, blow away the dust on the surface of the carbon fiber preform with compressed air, then put it into an ultrasonic cleaning machine for ultrasonic cleaning for 30min to 60min, and then put it into an oven at 100℃ to 120℃ for drying for 2h. Perform high-temperature heat treatment at 2300℃ to 2400℃ in a high-temperature furnace for 1h to 2h, introduce argon as a protective gas, and maintain the pressure at 3000Pa to 5000Pa.

[0010] Step 3: Preparation of PyC interface: First, a PyC interface with 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 carbon matrix graphite column fixture for directional deposition is placed at each gas outlet in the deposition furnace, and then the carbon fiber preform is placed in the column, so that the gas can flow from the inside of the column to the outside, so as to achieve the purpose of uniform deposition in the 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 erosion damage to Cf and PyC.

[0012] Step 5: Preparation of ZrC ceramic matrix and SiBCN matrix: The C / C body is densified into an ultra-high temperature ceramic matrix by using a high-pressure multiphase precursor uniformly mixed impregnation co-cracking process. 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: SiC coating is deposited on the surface of C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material by chemical vapor deposition method for surface sealing, and 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 is: placing the fiber preform in step 2 in a chemical deposition device, using natural gas as a carbon source and nitrogen as a diluent gas for pyrolytic carbon deposition. The deposition process parameters are: deposition temperature of 950°C to 980°C, natural gas flow rate of 10L / h to 20L / h, nitrogen flow rate of 10L / h to 20L / h, deposition time of 10h to 15h, and deposition pressure of 1000Pa to 1500Pa.

[0016] Preferably, in step 4, the preparation method of the pyrolytic carbon matrix is: placing the preform in step 3 in a chemical deposition device, using natural gas as a carbon source and nitrogen as a diluent gas for pyrolytic carbon deposition. The deposition process parameters are: 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 preparation method of the ZrC ceramic matrix and the SiBCN matrix is ​​as follows: first, a mixed impregnation solution of polycarbon zirconium and polyborosilazane is prepared. The conversion rate of the ZrC and SiBCN matrices after cracking is obtained by thermogravimetric analysis based on the thickness of the PyC interface, the density of the pyrolyzed carbon matrix, and the polycarbon zirconium and polyborosilazane, and the polycarbon zirconium and polyborosilazane are accurately proportioned and stirred and mixed, wherein xylene is used as a solvent, wherein the mass ratio of xylene to the total mass of polycarbon zirconium and polyborosilazane is 1:1, and the stirring time during the preparation of the mixed impregnation solution of polycarbon zirconium 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 polycarbon zirconium and polyborosilazane is added to the device, nitrogen is introduced to a pressure of 3MPa to 5MPa, and pressurized impregnation is performed for 2h. After the pressure impregnation is completed, it is placed in an oven for curing for 3 hours. The curing temperature is 250℃~300℃. After curing, 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 insulation time is 1h~3h, and nitrogen is introduced as a protective gas. Repeat the above high-pressure multiphase precursor uniform mixing impregnation and co-cracking process until the C / C body density reaches 2.0~2.3g / cm 3 The C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material was obtained. In the previous study, thermogravimetric analysis was performed on polycarbon zirconane and polyborosilazane, respectively, in order to determine how much mass of ZrC matrix can be converted from polycarbon zirconane at a cracking temperature of 1400℃~1600℃, and then calculate the ratio of ZrC mass to polycarbon zirconane mass to obtain the ZrC conversion rate; similarly, the mass of SiBCN matrix that can be converted from polyborosilazane at a cracking temperature of 1400℃~1600℃ was measured, and then the SiBCN matrix conversion rate of polyborosilazane can also be obtained; according to the matrix conversion rates of the two substances, the mass of polycarbon zirconane and polyborosilazane that need to be added during the preparation process can be calculated in reverse, so as to ensure that the mass ratio of ZrC ceramic matrix to SiBCN matrix content in the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material is 7:3~6:4.

[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 seven, the SiC surface coating preparation method is: placing the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material in a chemical deposition device, 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 25h to 40h, and deposition pressure of 1000Pa to 2000Pa.

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

[0021] The present invention adopts a combination of specific carbon fiber reinforcement and ceramic matrix according to the characteristics of ultra-high temperature (>2000℃), extremely large thermal gradient and thermal stress, and severe scouring of large amounts of solid particles such as Al2O3 faced by ultra-high temperature composite materials during the service of near-space aircraft, and realizes the preparation of high-performance composite materials 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, precise ratio and 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 gas flow direction is not controlled, the deposition uniformity and deposition efficiency of the carbon matrix will be seriously affected. Therefore, the 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 carbon fiber porous body in the graphite column fixture, the gas is forced to flow from the graphite column fixture from the inside to the outside, so that the generated carbon matrix can be 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 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] The high-pressure multiphase precursor uniform mixing impregnation co-cracking process is adopted to achieve uniform distribution of ZrC matrix and SiBCN multiphase matrix, realize the controllable preparation of high-performance composite materials, greatly improve the densification efficiency and final density, and shorten the preparation cycle. In the preparation process of C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composites, due to the different thermal expansion coefficients of carbon fiber, PyC interface, carbon matrix, ZrC matrix and SiBCN matrix, internal stress will be formed inside the material during repeated heating and cooling. In order to eliminate the residual thermal stress between the components, the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite needs to be post-high temperature heat treatment. The residual stress between the components of the material is eliminated through post-high temperature heat treatment, and the ZrC and SiBCN multiphase ultra-high temperature ceramic matrix prepared by low-temperature cracking of the precursor avoids the damage to the fiber caused by the high-temperature reaction infiltration process, thereby improving the overall oxidation resistance and mechanical properties of the material. The temperature of low temperature cracking 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, which greatly improves the toughness of the material, improves 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 invention is simple, easy to operate, and has good process applicability. The prepared ultra-high temperature composite material has excellent mechanical properties, antioxidant properties, and anti-ablation properties, and has broad application prospects in hot end components of aerospace power ultra-high-speed aircraft. BRIEF DESCRIPTION OF THE DRAWINGS

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

[0028] Figure 2 The present invention is a flow chart for preparing an ultra-high temperature ceramic modified C / C composite material for aerospace power.

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

[0030] Figure 4 This is the internal microscopic morphology of the C / C-ZrC-SiBCN ultra-high temperature ceramic-based 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 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 hot end component of the aerospace power. 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 1h to 2h, 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 pyrolyze the carbon matrix. The deposition process parameters are: natural gas is used as the carbon source and nitrogen is used as the diluent gas for pyrolysis carbon deposition; the deposition process parameters are: 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 blank is obtained after pyrolysis carbon matrix deposition. The density of the C / C blank is 1.2~1.3g / cm 3During the deposition process, a graphite material column (referred to as material 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 material column, so that the gas 2 can flow from the inside of the material column to the outside, so as to achieve 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 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 7:3-6:4, and the amounts of polycarbon zirconium and polyborosilazane are calculated and weighed according to the ZrC conversion rate in the polycarbon zirconium and the SiBCN matrix conversion rate in the polyborosilazane obtained by the previous thermogravimetric analysis, and then the weighed polycarbon zirconium and polyborosilazane are added into a xylene solvent and stirred to obtain a mixed impregnation solution of polycarbon zirconium and polyborosilazane, wherein the mass ratio of xylene to the total mass of polycarbon zirconium and polyborosilazane is 1:1, and the polycarbon zirconium 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 blank is subjected to high-pressure impregnation; the C / C blank is placed in an impregnation device, and then vacuum is drawn to less than 100Pa, the mixed impregnation liquid of polycarbon zirconium ane and polyborosilazane is added to the impregnation device, nitrogen is introduced to the pressure of 3MPa to 5MPa, and the pressure impregnation is carried out 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°C to 300°C; 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°C to 1600°C, the heating rate is 5 to 10°C / min, the insulation time is 1h to 3h, and nitrogen is introduced as a protective gas; the above high-pressure multiphase precursor uniform mixing impregnation co-cracking process is repeated for many times until the density of the C / C blank reaches 2.0 to 2.3g / cm 3 A C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material is obtained;

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

[0043] The post-heat treatment temperature process parameters are: 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 ultra-high temperature ceramic matrix composite material by chemical vapor deposition to seal the surface, and finally a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material with a density of 2.3-2.5 g / cm 3 .

[0046] The present invention will be further described below in conjunction with 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 fiber. 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 preform: The carbon fiber preform is first blown with compressed air to remove dust from the surface, then placed in an ultrasonic cleaning machine for ultrasonic cleaning for 30 minutes, and then placed in an oven for drying at 120°C for 2 hours. A high-temperature heat treatment is performed at 2300°C in a high-temperature furnace for 2 hours, with argon gas introduced as a protective gas at a pressure of 3000Pa.

[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 polycarbon zirconium 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 polycarbon zirconium 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 polycarbon zirconium and polyborosilazane was added. Add it to the impregnation equipment, introduce nitrogen to the pressure of 3MPa, and perform pressure impregnation for 2h; after the pressure impregnation is completed, put it into the oven for curing for 3h; the curing temperature is 250℃; after the curing is completed, put it into the cracking furnace for cracking; the cracking temperature is 1400℃, the heating rate is 5℃ / min, the insulation time is 1h, and nitrogen is introduced as the protective gas. After 8 times of impregnation and cracking, the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material with a density of 1.95g / cm 3 .

[0053] (7) Preparation of SiC surface coating: The SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultra-high 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, the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material was obtained.

[0054] The C / C-ZrC-SiBCN ultra-high 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 153.9MPa. The composite material was ablated at 2100℃ for 100s, and the linear ablation rate was 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 pull-out and low mechanical properties.

[0055] Embodiment 1:

[0056] (1) Preparation of carbon fiber preform: The carbon fiber preform is woven in a three-dimensional orthogonal manner using T1000 carbon fiber. 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 preform: The carbon fiber preform is first blown with compressed air to remove dust from the surface, then placed in an ultrasonic cleaning machine for ultrasonic cleaning for 30 minutes, and then placed in an oven for drying at 120°C for 2 hours. A high-temperature heat treatment is performed at 2300°C in a high-temperature furnace for 2 hours, with argon gas introduced as a protective gas at a pressure of 3000Pa.

[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 polycarbon zirconium 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 polycarbon zirconium 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 to impregnate the mixed impregnation solution of polycarbon zirconium and polyborosilazane. Add it to the impregnation equipment, introduce nitrogen to the pressure of 3MPa, and perform pressure impregnation for 2h; after the pressure impregnation is completed, put it into the oven for curing for 3h; the curing temperature is 250℃; after the curing is completed, put it into the cracking furnace for cracking; the cracking temperature is 1400℃, the heating rate is 5℃ / min, the insulation time is 1h, and nitrogen is introduced as the protective gas. After 8 times of impregnation and cracking, a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material with a density of 2.0g / cm 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: The SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultra-high 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, the C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material was obtained.

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

[0064] Embodiment 2:

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

[0066] (2) High-temperature heat treatment of carbon fiber preform: The carbon fiber preform is first blown with compressed air to remove dust from the surface, then placed in an ultrasonic cleaning machine for ultrasonic cleaning for 30 minutes, and then placed in an oven for drying at 120°C for 2 hours. A high-temperature heat treatment is performed at 2300°C in a high-temperature furnace for 2 hours, with argon gas introduced as a protective gas at a pressure of 3000Pa.

[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 a directional deposition carbon matrix graphite column tooling 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 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, a mixed impregnation solution of polycarbon zirconium and polyborosilazane in xylene was prepared, wherein the mass ratio of ZrC matrix to SiBCN matrix was 7:3, the mass ratio of xylene to the total mass of polycarbon zirconium and polyborosilazane was 1:1, and the stirring time was 8 hours; then, the C / C body was subjected to high-pressure impregnation; the C / C body was placed in an impregnation device, and then vacuumed to 40 Pa, and the mixed impregnation solution of polycarbon zirconium and polyborosilazane was added The mixture was put into the impregnation equipment, nitrogen was introduced to the pressure of 5MPa, and the pressure impregnation was carried out for 2h; after the pressure impregnation was completed, it was placed in an oven for curing for 3h; the curing temperature was 300℃; after the curing was completed, it was placed in a cracking furnace for cracking; the cracking temperature was 1600℃, the heating rate was 5℃ / min, the insulation time was 2h, nitrogen was introduced as the protective gas, and after 10 times of impregnation and cracking, a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material was obtained with a density of 2.15g / cm 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: The SiC coating was deposited on the surface of the C / C-ZrC-SiBCN ultra-high 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, the C / C-ZrC-SiBCN ultra-high 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 233.7MPa. The composite material was ablated at 2100℃ for 100s, and the linear ablation rate was 0.37μm / s. Figure 4 From the internal microstructure, it can be seen 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 power, characterized in that: 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 hot end component of the aerospace power. 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%. (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 1h to 2h, argon is introduced as a protective gas, and the pressure is 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 by chemical deposition to obtain a carbon fiber porous body; the thickness of the PyC interface is 600nm to 800nm; (4) Pyrolytic carbon matrix deposition The carbon fiber porous body is subjected to chemical vapor phase directional deposition to pyrolyze the carbon matrix. The deposition process parameters are: natural gas is used as the carbon source and nitrogen is used as the diluent gas for pyrolysis carbon deposition; the deposition process parameters are: 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 blank is obtained after pyrolysis carbon matrix deposition. The density of the C / C blank is 1.2~1.3g / cm 3 ; (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 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 7:3-6:4, and the amounts of polycarbon zirconium and polyborosilazane are calculated and weighed according to the ZrC conversion rate in the polycarbon zirconium and the SiBCN matrix conversion rate in the polyborosilazane obtained by the previous thermogravimetric analysis, and then the weighed polycarbon zirconium and polyborosilazane are added into a xylene solvent and stirred to obtain a mixed impregnation solution of polycarbon zirconium and polyborosilazane, wherein the mass ratio of xylene to the total mass of polycarbon zirconium and polyborosilazane is 1:1, and the polycarbon zirconium 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 blank is subjected to high-pressure impregnation; the C / C blank is placed in an impregnation device, and then vacuum is drawn to less than 100Pa, the mixed impregnation liquid of polycarbon zirconium ane and polyborosilazane is added to the impregnation device, nitrogen is introduced to the pressure of 3MPa to 5MPa, and the pressure impregnation is carried out 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°C to 300°C; 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°C to 1600°C, the heating rate is 5 to 10°C / min, the insulation time is 1h to 3h, and nitrogen is introduced as a protective gas; the above high-pressure multiphase precursor uniform mixing impregnation co-cracking process is repeated for many times until the density of the C / C blank reaches 2.0 to 2.3g / cm 3 A C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material is obtained; (6) Post-high temperature heat treatment The post-heat treatment temperature process parameters are: 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 ultra-high temperature ceramic matrix composite material by chemical vapor deposition to seal the surface, and finally a C / C-ZrC-SiBCN ultra-high temperature ceramic matrix composite material with a density of 2.3-2.5 g / cm 3 .

2. The method for preparing the ultra-high temperature ceramic modified C / C composite material for aerospace power according to claim 1, characterized in that: In step (1), carbon fiber is firstly 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 to 60min, and then placed in an oven at 100°C to 120°C for drying for 2h.

3. The method for preparing the ultra-high temperature ceramic modified C / C composite material for aerospace power 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 to 980°C, the natural gas flow rate is 10L / h to 20L / h, the nitrogen flow rate is 10L / h to 20L / h, the deposition time is 10h to 15h, and the deposition pressure is 1000Pa to 1500Pa.

4. The method for preparing the ultra-high temperature ceramic modified C / C composite material for aerospace power according to claim 1, characterized in that: In step (4), the pyrolytic carbon matrix deposition process is to place the carbon fiber porous body in step three in a graphite material column tooling of a directional deposition carbon matrix in a chemical deposition furnace chamber, wherein the graphite material column tooling is a cylindrical tooling with only one air inlet on the bottom and evenly distributed small air outlets on the top.

5. The method for preparing the ultra-high temperature ceramic modified C / C composite material for aerospace power according to claim 1, characterized in that: In step (7), the SiC 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: deposition temperature of 1000°C to 1200°C, flow ratio of trichloromethylsilane: hydrogen: argon of 1:20:30, deposition time of 25h to 40h, and deposition pressure of 1000Pa to 2000Pa.

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

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