High-temperature structure bearing / anti-ablation integrated ceramic matrix composite material and preparation method thereof

Abstract

The invention relates to a high-temperature structure bearing / anti-ablation integrated ceramic matrix composite and a preparation method thereof.The preparation method of the composite comprises the steps that a pyrolytic carbon interface layer is deposited on the fiber surface of a carbon fiber preform through a chemical vapor deposition method, and a low-density C / C base material is obtained; the density of the carbon fiber preform is gradually increased from the surface to the inside; the preparation method comprises the following steps: carrying out first vacuum impregnation, first pressure impregnation, first curing and first cracking on a low-density C / C base material by adopting phenolic resin to obtain a porous C / C intermediate; carrying out reactive infiltration on the porous C / C intermediate by adopting an alloy infiltration agent to generate a multi-element ultra-high-temperature ceramic matrix, so as to obtain a multi-element ultra-high-temperature ceramic matrix composite material; and preparing an anti-oxidation ceramic protective layer on the near surface of the multi-element ultrahigh-temperature ceramic-based composite material to obtain the high-temperature structure bearing / anti-ablation integrated ceramic-based composite material. The ceramic-based composite material prepared by the preparation method has good high-temperature structure bearing performance and oxidation ablation resistance.

Description

Technical Field

[0001] The present invention relates to the technical field of ultra-high temperature ceramic-based composite materials, and in particular to a high-temperature structural load-bearing / ablation-resistant integrated ceramic-based composite material and a preparation method thereof. Background Art

[0002] The wing rudder is a key component of high-speed aircraft and plays an important role in the flight attitude control of the aircraft. When a high-speed aircraft flies for a long time in the atmosphere, the service environment of the front edge of the rudder is harsh, its radiation equilibrium temperature reaches above 2000℃, and it bears loads of dozens or even hundreds of kN. Therefore, very high requirements are placed on the lightweight, high load-bearing and long-term anti-oxidation and ablation capabilities of the wing rudder structural materials. However, the comprehensive performance of existing ceramic-based composite materials is still far from the service requirements of high-speed aircraft wing rudder structural parts. It is difficult to take into account good high-temperature mechanics and anti-oxidation and ablation resistance, and it does not have the function of integrating structural load-bearing and anti-ablation. Summary of the invention

[0003] In response to one or more technical problems existing in the prior art, the present invention provides a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material and a preparation method thereof. The preparation method provided by the present invention can prepare a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material, which has both good high-temperature structural load-bearing performance and anti-oxidation and ablation functions, and can meet the requirements of high-speed aircraft wing and rudder structural materials for high structural load-bearing and good anti-oxidation and ablation performance, and overcomes the problem that existing ceramic-based composite materials are difficult to take into account both high-temperature structural load-bearing and anti-oxidation and ablation functions.

[0004] In a first aspect, the present invention provides a method for preparing a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material, the preparation method comprising:

[0005] S1. Depositing a pyrolytic carbon interface layer on the fiber surface of a carbon fiber preform by chemical vapor deposition to obtain a low-density C / C substrate; wherein the density of the carbon fiber preform increases from the surface to the inside;

[0006] S2. A low-density C / C substrate is subjected to a first vacuum impregnation, a first pressure impregnation, a first curing, and a first pyrolysis to obtain a porous C / C intermediate using a phenolic resin;

[0007] S3. Using an alloy infiltration agent to react and infiltrate the porous C / C intermediate to generate a multi-element ultra-high temperature ceramic matrix to obtain a multi-element ultra-high temperature ceramic matrix composite material;

[0008] S4. Prepare an anti-oxidation ceramic protective layer near the surface of the multi-component ultra-high temperature ceramic-based composite material to obtain a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material.

[0009] Preferably, the increasing gradient of the density from the surface to the interior of the carbon fiber preform is 0.05 to 0.1 g / cm 3 ;

[0010] The surface density of the carbon fiber preform is 0.30-0.35 g / cm 3 ;

[0011] The maximum density inside the carbon fiber preform is 0.60-0.65 g / cm 3 ; and / or

[0012] Compared with the carbon fiber preform, the density of the low-density C / C substrate is increased by 60 to 100%.

[0013] Preferably, before depositing the pyrolytic carbon interface layer on the fiber surface of the carbon fiber preform by chemical vapor deposition, the method further includes subjecting the carbon fiber preform to high temperature treatment; the high temperature heat treatment temperature is 1000-1800° C., and the time is 2.0-3.0 hours; and / or

[0014] The chemical vapor deposition method uses propylene as a carbon source and nitrogen as a carrier, and performs deposition at 800-1100°C.

[0015] Preferably, the vacuum degree of the first vacuum impregnation is -0.15 to -0.08 MPa, the temperature is room temperature, and the time is 0.5 to 2.0 h;

[0016] The first pressure impregnation has a pressure of 0.5 to 2.0 MPa, a temperature of room temperature, and a time of 0.5 to 2.0 h;

[0017] The first curing is performed at a temperature of 100 to 350° C., a pressure of 0.5 to 2.5 MPa, and a time of 0.5 to 2.0 h; and / or

[0018] The temperature of the first cracking is 700-1000° C., and the time is 2.0-4.0 hours.

[0019] Preferably, the density of the porous C / C intermediate is increased by 20-50% compared to the low-density C / C substrate.

[0020] Preferably, the alloying agent is one of a silicon-zirconium binary alloy and a silicon-hafnium binary alloy; preferably, the atomic weight fraction of zirconium in the silicon-zirconium alloy is 15-25%; the atomic weight fraction of hafnium in the silicon-hafnium alloy is 15-31%; the particle size of the alloying agent is 1-10 μm; and / or

[0021] The temperature of the reaction infiltration is 1600-1800° C., and the time of the reaction infiltration is 1.0-2.0 h.

[0022] Preferably, an anti-oxidation ceramic protective layer is prepared near the surface of the multi-element ultra-high temperature ceramic-based composite material, comprising:

[0023] The multi-component ultra-high temperature ceramic-based composite material is subjected to a second vacuum impregnation, a second pressure impregnation, a second curing and a second cracking by using a first ceramic precursor to form a ceramic transition layer near the surface of the multi-component ultra-high temperature ceramic-based composite material; the first ceramic precursor is a silicon-based ceramic precursor;

[0024] The composite material forming the silicon-based ceramic transition layer near the surface is subjected to a third vacuum impregnation, a third pressure impregnation, a third curing and a third cracking by using a second ceramic precursor to form a ceramic surface layer, thereby obtaining an anti-oxidation ceramic protective layer; the second ceramic precursor is a mixed ceramic precursor including a silicon-based ceramic precursor and an ultra-high temperature ceramic precursor;

[0025] Preferably, the silicon-based ceramic precursor is polycarbosilane; and / or

[0026] The ultra-high temperature ceramic precursor is at least one of a ZrC ceramic precursor and a HfC ceramic precursor.

[0027] Preferably, the vacuum degree of the second vacuum impregnation and the third vacuum impregnation is -0.15 to -0.08 MPa, the temperature is room temperature, and the time is 0.5 to 2.0 h;

[0028] The pressure of the second pressure impregnation and the third pressure impregnation is 0.5-2.0 MPa, the temperature is room temperature, and the time is 0.5-2.0 h;

[0029] The temperature of the second curing and the third curing is 100-350° C., the pressure is 0.5-2.5 MPa, and the time is 0.5-2.0 h;

[0030] The temperature of the second cracking is 700-1000°C and the time is 2.0-4.0h; and / or

[0031] The temperature of the third cracking is 1400-1600° C., and the time is 2.0-4.0 hours.

[0032] In a second aspect, the present invention provides a high-temperature structural load-bearing / ablation-resistant integrated ceramic-based composite material, which is prepared by the preparation method described in the first aspect.

[0033] Compared with the prior art, the present invention has at least the following beneficial effects:

[0034] The present invention first selects a carbon fiber preform whose density increases from the surface to the inside (a carbon fiber preform with a gradient structure with increasing density) and deposits a pyrolytic carbon interface layer on the surface of the carbon fiber, and combines it with phenolic resin impregnation, curing and cracking to prepare a porous C / C intermediate; then, a multi-component ultra-high temperature ceramic matrix is ​​in situ generated in the porous C / C intermediate by a reactive melt infiltration method to obtain a multi-component ultra-high temperature ceramic-based composite material with a gradient structure having a high internal carbon fiber volume fraction and a high surface ceramic matrix content, which has good load-bearing capacity and anti-ablation performance; finally, a high-temperature resistant and anti-oxidation ceramic protective layer is prepared on the near-surface of the multi-component ultra-high temperature ceramic-based composite material to provide anti-oxidation protection for the composite material, thereby further improving the high-temperature anti-oxidation and ablation resistance of the composite material.

[0035] The preparation method provided by the present invention can prepare a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material, which has both good high-temperature structural load-bearing performance and anti-oxidation and ablation functions, and can meet the requirements of high-speed aircraft wing and rudder structural materials for high structural load-bearing and good anti-oxidation and ablation performance, overcoming the problem that existing ceramic-based composite materials are difficult to strike a balance between good high-temperature structural load-bearing and anti-oxidation and ablation resistance. DETAILED DESCRIPTION

[0036] In order to make the purpose, technical solutions and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below. Obviously, the described embodiments are part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments in the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention. In the description of the embodiments of the present invention, unless otherwise clearly specified and limited, the terms "first", "second", "third", etc. are used only for descriptive purposes and cannot be understood as indicating or implying relative importance.

[0037] In a first aspect, the present invention provides a method for preparing a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material, the preparation method comprising:

[0038] S1. Depositing a pyrolytic carbon interface layer on the fiber surface of a carbon fiber preform by chemical vapor deposition to obtain a low-density C / C substrate; wherein the density of the carbon fiber preform increases from the surface to the inside;

[0039] S2. A low-density C / C substrate is subjected to a first vacuum impregnation, a first pressure impregnation, a first curing, and a first pyrolysis to obtain a porous C / C intermediate using a phenolic resin;

[0040] S3. Using an alloy infiltration agent to react and infiltrate the porous C / C intermediate to generate a multi-element ultra-high temperature ceramic matrix to obtain a multi-element ultra-high temperature ceramic matrix composite material;

[0041] S4. Prepare an anti-oxidation ceramic protective layer near the surface of the multi-component ultra-high temperature ceramic-based composite material to obtain a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material.

[0042] It should be noted that the term “near the surface” is used above because in the process of preparing the antioxidant ceramic protective layer, some of it will penetrate into the interior of the multi-component ultra-high temperature ceramic-based composite material, not just on the surface.

[0043] The present invention first selects a carbon fiber preform whose density increases from the surface to the inside (a carbon fiber preform with a gradient structure with increasing density) and deposits a pyrolytic carbon interface layer on the surface of the carbon fiber, and combines it with phenolic resin impregnation and cracking to prepare a porous C / C intermediate; then, a multi-component ultra-high temperature ceramic matrix is ​​in situ generated in the porous C / C intermediate by a reactive melt infiltration method to obtain a multi-component ultra-high temperature ceramic-based composite material with a gradient structure having a high internal carbon fiber volume fraction and a high surface ceramic matrix content, which has good load-bearing capacity and anti-ablation performance; finally, a high-temperature resistant and anti-oxidation ceramic protective layer is prepared on the near-surface of the multi-component ultra-high temperature ceramic-based composite material to provide anti-oxidation protection for the composite material, thereby further improving the high-temperature anti-oxidation and ablation resistance of the composite material.

[0044] The preparation method provided by the present invention can prepare high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite materials, which have both good high-temperature structural load-bearing performance and anti-oxidation and ablation resistance functions, and can meet the requirements of high-speed aircraft wing and rudder structural materials for high structural load-bearing and good anti-oxidation and ablation resistance, overcoming the problem that existing ceramic-based composite materials are difficult to balance high-temperature structural load-bearing and anti-oxidation and ablation resistance.

[0045] According to some preferred embodiments, the increasing gradient of the density from the surface to the interior of the carbon fiber preform is 0.05 to 0.1 g / cm 3 (For example, it can be 0.05 g / cm 3 , 0.06g / cm 3 , 0.07g / cm 3 , 0.08g / cm 3 , 0.09g / cm 3 or 0.1g / cm 3 ).

[0046] According to some preferred embodiments, the surface density of the carbon fiber preform is 0.30-0.35 g / cm 3 (For example, it can be 0.30 g / cm 3 , 0.31g / cm 3 , 0.32g / cm 3 , 0.33g / cm 3 , 0.34g / cm 3 or 0.35g / cm3 ).

[0047] According to some preferred embodiments, the maximum density inside the carbon fiber preform is 0.60-0.65 g / cm 3 (For example, it can be 0.60g / cm 3 、0.61g / cm 3 , 0.62g / cm 3 , 0.63g / cm 3 , 0.64g / cm 3 or 0.65g / cm 3 ).

[0048] According to some preferred embodiments, the density of the low-density C / C substrate is increased by 60-100% (e.g., 60%, 65%, 70%, 75%, 80%, 85%, 90%, 95% or 100%) compared to the carbon fiber preform.

[0049] According to some preferred embodiments, before depositing a pyrolytic carbon interface layer on the fiber surface of the carbon fiber preform by chemical vapor deposition, the carbon fiber preform is also subjected to high-temperature treatment; the high-temperature heat treatment temperature is 1000-1800°C (for example, it can be 1000°C, 1100°C, 1200°C, 1300°C, 1400°C, 1500°C, 1600°C, 1700°C or 1800°C), and the time is 2.0-3.0h (for example, it can be 2.0h, 2.2h, 2.4h, 2.6h, 2.8h or 3.0h).

[0050] According to some preferred embodiments, the chemical vapor deposition method uses propylene as a carbon source and nitrogen as a carrier, and performs deposition at 800-1100°C (for example, 800°C, 850°C, 900°C, 950°C, 1000°C, 1050°C or 1100°C).

[0051] According to some preferred embodiments, the vacuum degree of the first vacuum impregnation is -0.15 to -0.08 MPa, the temperature of the first vacuum impregnation is room temperature, and the time is 0.5 to 2.0 h.

[0052] According to some preferred embodiments, the pressure of the first pressure impregnation is 0.5-2.0 MPa, the temperature is room temperature, and the time is 0.5-2.0 h.

[0053] According to some preferred embodiments, the first curing is performed at a temperature of 100 to 350° C., a pressure of 0.5 to 2.5 MPa, and a time of 0.5 to 2.0 h.

[0054] The temperature of the first cracking is 700-1000°C (for example, 700°C, 750°C, 800°C, 850°C, 900°C, 950°C or 1000°C), and the time is 2.0-4.0h.

[0055] According to some preferred embodiments, the density of the porous C / C intermediate is increased by 20-50% (e.g., 20%, 25%, 30%, 35%, 40%, 45% or 50%) compared to the low-density C / C substrate.

[0056] According to some preferred embodiments, the alloy infiltrant is one of a silicon-zirconium binary alloy and a silicon-hafnium binary alloy; preferably, the atomic weight fraction of zirconium in the silicon-zirconium alloy is 15-25% (for example, it can be 15%, 16%, 17%, 18%, 19%, 20%, 21%, 22%, 23%, 24% or 25%); the atomic weight fraction of hafnium in the silicon-hafnium alloy is 15-31% (for example, it can be 15%, 16%, 18%, 20%, 22%, 24%, 25%, 26%, 28%, 30% or 31%); the particle size of the alloy infiltrant is 1-10 μm.

[0057] According to some preferred embodiments, the temperature of the reactive infiltration is 1600-1800°C (for example, 1600°C, 1650°C, 1700°C, 1750°C or 1800°C), and the time of the reactive infiltration is 1.0-2.0h.

[0058] According to some preferred embodiments, an anti-oxidation ceramic protective layer is prepared near the surface of a multi-element ultra-high temperature ceramic-based composite material, comprising:

[0059] The multi-component ultra-high temperature ceramic-based composite material is subjected to a second vacuum impregnation, a second pressure impregnation, a second curing and a second cracking by using a first ceramic precursor to form a ceramic transition layer near the surface of the multi-component ultra-high temperature ceramic-based composite material; the first ceramic precursor is a silicon-based ceramic precursor;

[0060] A second ceramic precursor is used to perform a third vacuum impregnation, a third pressure impregnation, a third curing and a third cracking on the composite material that forms a silicon-based ceramic transition layer near the surface to form a ceramic surface layer, thereby obtaining an antioxidant ceramic protective layer; the second ceramic precursor is a mixed ceramic precursor including a silicon-based ceramic precursor and an ultra-high temperature ceramic precursor.

[0061] According to some preferred embodiments, the second ceramic precursor is a mixed ceramic precursor including 25-50 wt. % of a silicon-based ceramic precursor and 50-75 wt. % of an ultra-high temperature ceramic precursor by weight fraction.

[0062] According to some preferred embodiments, the silicon-based ceramic precursor is polycarbosilane.

[0063] According to some preferred embodiments, the ultra-high temperature ceramic precursor is at least one of a ZrC ceramic precursor and a HfC ceramic precursor.

[0064] The present invention first uses a silicon-based ceramic precursor to prepare a transition layer near the surface of a multi-element ultra-high temperature ceramic-based composite material through an impregnation, curing and cracking process, and then uses a mixture of a silicon-based ceramic precursor and an ultra-high temperature ceramic precursor to form a ceramic surface layer through an impregnation, curing and cracking process. The transition layer has good compatibility with the multi-element ultra-high temperature ceramic-based composite material and the ceramic surface layer, and can firmly combine the multi-element ultra-high temperature ceramic-based composite material and the ceramic surface layer, thereby enhancing the adhesion and durability of the ceramic surface layer.

[0065] According to some preferred embodiments, the vacuum degree of the second vacuum impregnation and the third vacuum impregnation is -0.15 to -0.08 MPa, the temperature is room temperature, and the time is 0.5 to 2.0 h.

[0066] According to some preferred embodiments, the pressure of the second pressure impregnation and the third pressure impregnation is 0.5-2.0 MPa, the temperature is room temperature, and the time is 0.5-2.0 h.

[0067] According to some preferred embodiments, the temperature of the second curing and the third curing is 100-350°C (for example, 100°C, 150°C, 200°C, 250°C, 300°C or 350°C), the pressure is 0.5-2.5 MPa, and the time is 0.5-2.0 h.

[0068] According to some preferred embodiments, the temperature of the second cracking is 700-1000°C (for example, 700°C, 750°C, 800°C, 850°C, 900°C, 950°C or 1000°C) and the time is 2.0-4.0h.

[0069] According to some preferred embodiments, the temperature of the third cracking is 1400-1600°C (for example, 1400°C, 1450°C, 1500°C, 1550°C or 1600°C), and the time is 2.0-4.0h.

[0070] In a second aspect, the present invention provides a high-temperature structural load-bearing / ablation-resistant integrated ceramic-based composite material, which is prepared by the preparation method described in the first aspect.

[0071] In order to more clearly illustrate the technical solutions and advantages of the present invention, the present invention is further described below in conjunction with the examples. The present invention does not specifically limit the sources of the reagents used in the examples and comparative examples, and the reagents can be directly purchased or synthesized by themselves.

[0072] Example 1

[0073] S1. Preparation of low-density C / C substrate: Design and weave variable-density carbon fiber needle-punched preforms. By adjusting the fiber distribution of the carbon fiber preforms, the density of the carbon fiber preforms gradually increases from the surface to the inside. The surface density of the carbon fiber preforms is 0.35 g / cm 3 The maximum density inside the carbon fiber preform is 0.65g / cm 3 The density gradient is 0.1 g / cm 3 After the carbon fiber preform was heat treated at 1600°C for 2.0h, a pyrolytic carbon interface layer was deposited on the fiber surface of the carbon fiber preform by chemical vapor deposition (CVI) process with propylene as carbon source and nitrogen as carrier gas at a deposition temperature of 1000°C to obtain a low-density C / C substrate (the density increment of the low-density C / C substrate was 80% compared with the carbon fiber preform).

[0074] S2. A porous C / C intermediate is prepared by a precursor impregnation and pyrolysis (PIP) process: a phenolic resin is used as a precursor to sequentially perform the first vacuum impregnation on a low-density C / C substrate (the vacuum degree of the first vacuum impregnation is -0.1 MPa, the temperature is room temperature, and the time is 1.5 h), the first pressure impregnation (the pressure of the first pressure impregnation is 2.0 MPa, the temperature is room temperature, and the time is 1.5 h), the first curing (the temperature of the first curing is 240°C, the pressure is 1.5 MPa, and the time is 2.0 h) and the first pyrolysis (the temperature of the first pyrolysis is 800°C, and the time is 2.0 h) to obtain a porous C / C intermediate (the density increase of the porous C / C intermediate compared to the low-density C / C substrate is 35%).

[0075] S3. Generate a multi-component ultra-high temperature ceramic matrix by using a reactive infiltration method: Use a silicon-zirconium alloy infiltrant (the atomic weight fraction of zirconium in the silicon-zirconium alloy infiltrant is 18%, and the particle size of the silicon-hafnium alloy infiltrant powder is 3-8 μm) to perform reactive infiltration (the reactive infiltration temperature is 1800°C and the time is 1.5 h) on the porous C / C intermediate to in situ generate a multi-component ultra-high temperature ceramic matrix to obtain a multi-component ultra-high temperature ceramic-based composite material.

[0076] S4. The precursor impregnation pyrolysis (PIP) process is used to prepare an antioxidant ceramic protective layer on the near surface of the multi-component ultra-high temperature ceramic-based composite material: first, liquid polycarbosilane is used to perform three rounds of second vacuum impregnation (the vacuum degree of the second vacuum impregnation is -0.1MPa, the temperature is room temperature, and the time is 1.5h), second pressure impregnation (the pressure of the second pressure impregnation is 2.0MPa, the temperature is room temperature, and the time is 1.5h), second curing (the temperature of the second curing is 240℃, the pressure is 1.5MPa, and the time is 2.0h) and second pyrolysis (the temperature of the second pyrolysis is 1000℃, and the time is 2 .0h), and then use a mixed solution including 75wt.% HfC ceramic precursor and 25wt.% liquid polycarbosilane to carry out three rounds of third vacuum impregnation (the vacuum degree of the third vacuum impregnation is -0.1MPa, the temperature is room temperature, and the time is 1.5h), third pressure impregnation (the pressure of the third pressure impregnation is 2.0MPa, the temperature is room temperature, and the time is 1.5h), third curing (the temperature of the third curing is 240℃, the pressure is 1.5MPa, and the time is 2.0h) and third pyrolysis (the temperature of the third pyrolysis is 1500℃, and the time is 2.0h) to obtain a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material.

[0077] Example 2

[0078] S1. Preparation of low-density C / C substrate: Design and weave variable-density carbon fiber needle-punched preforms. By adjusting the fiber distribution of the carbon fiber preforms, the density of the carbon fiber preforms gradually increases from the surface to the inside. The surface density of the carbon fiber preforms is 0.35 g / cm 3 The maximum density inside the carbon fiber preform is 0.65g / cm 3 The density gradient is 0.05 g / cm 3 After the carbon fiber preform was heat treated at 1600°C for 2.0h, a pyrolytic carbon interface layer was deposited on the fiber surface of the carbon fiber preform by chemical vapor deposition (CVI) process with propylene as carbon source and nitrogen as carrier gas at a deposition temperature of 1000°C to obtain a low-density C / C substrate (the density increment of the low-density C / C substrate was 95% compared with the carbon fiber preform).

[0079] S2. A porous C / C intermediate is prepared by a precursor impregnation and pyrolysis (PIP) process: a phenolic resin is used as a precursor to sequentially perform the first vacuum impregnation on a low-density C / C substrate (the vacuum degree of the first vacuum impregnation is -0.1 MPa, the temperature is room temperature, and the time is 1.5 h), the first pressure impregnation (the pressure of the first pressure impregnation is 2.0 MPa, the temperature is room temperature, and the time is 1.5 h), the first curing (the temperature of the first curing is 240°C, the pressure is 1.5 MPa, and the time is 2.0 h) and the first pyrolysis (the temperature of the first pyrolysis is 800°C, and the time is 2.0 h) to obtain a porous C / C intermediate (the density increase of the porous C / C intermediate compared to the low-density C / C substrate is 35%).

[0080] S3. Generate a multi-component ultra-high temperature ceramic matrix by using a reactive infiltration method: a silicon-hafnium alloy infiltrant (the atomic weight fraction of hafnium in the silicon-hafnium alloy infiltrant is 19%, and the particle size of the silicon-hafnium alloy infiltrant powder is 3 to 8 μm) is used to perform reactive infiltration (the reactive infiltration temperature is 1750°C and the time is 1.5 h) in a porous C / C intermediate to in situ generate a multi-component ultra-high temperature ceramic matrix, and obtain a multi-component ultra-high temperature ceramic matrix composite material;

[0081] S4. The precursor impregnation pyrolysis (PIP) process is used to prepare an antioxidant ceramic protective layer on the near surface of the multi-component ultra-high temperature ceramic-based composite material: first, liquid polycarbosilane is used to perform three rounds of second vacuum impregnation (the vacuum degree of the second vacuum impregnation is -0.1MPa, the temperature is room temperature, and the time is 1.5h), second pressure impregnation (the pressure of the second pressure impregnation is 2.0MPa, the temperature is room temperature, and the time is 1.5h), second curing (the temperature of the second curing is 240℃, the pressure is 1.5MPa, and the time is 2.0h;) and second pyrolysis (the temperature of the second pyrolysis is 1000℃, and the time is 2 .0h), and then use a mixed solution including 75wt.% HfC ceramic precursor and 25wt.% liquid polycarbosilane to carry out three rounds of third vacuum impregnation (the vacuum degree of the third vacuum impregnation is -0.1MPa, the temperature is room temperature, and the time is 1.5h), third pressure impregnation (the pressure of the third pressure impregnation is 2.0MPa, the temperature is room temperature, and the time is 1.5h), third curing (the temperature of the third curing is 240℃, the pressure is 1.5MPa, and the time is 2.0h;) and third pyrolysis (the temperature of the third pyrolysis is 1500℃, and the time is 2.0h) to obtain a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material.

[0082] Comparative Example 1

[0083] The carbon fiber preform is basically the same as Example 2, except that the carbon fiber preform is a common carbon fiber preform without density gradient, and the density of the carbon fiber preform is 0.6 g / cm 3 .

[0084] Comparative Example 2

[0085] It is basically the same as Example 2, except that step S4 is not included.

[0086] The testing methods for the performance data of the composite materials obtained in the examples and comparative examples of the present invention are as follows. The relevant performance data are shown in Table 1.

[0087] Mechanical properties

[0088] The room temperature tensile strength test adopts the GJB / 8736-2015 continuous fiber reinforced ceramic matrix composite material room temperature tensile performance test method.

[0089] The room temperature flexural strength test adopts the Q / SB 513-2013 continuous fiber reinforced ceramic matrix composite room temperature flexural performance test method.

[0090] Anti-oxidation and ablation performance

[0091] The oxyacetylene ablation test adopted the GJB323B-2018 ablation material ablation test method. During the ablation test, the ablation surface temperature of the test sample was 2000°C.

[0092] Table 1. Performance data of composite materials obtained from the examples of the present invention and comparative examples

[0093]

[0094] It should be noted that since Comparative Example 1 uses a fiber preform without a density gradient, the composite material obtained also has no density gradient. The “-” in Table 1 indicates that the performance of each part (surface and interior) of the composite material is similar, and the surface and interior are not distinguished.

[0095] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for preparing a high-temperature structural load-bearing / anti-ablation integrated ceramic matrix composite material, characterized in that: The preparation method comprises: S1. Depositing a pyrolytic carbon interface layer on the fiber surface of a carbon fiber preform by chemical vapor deposition to obtain a low-density C / C substrate; wherein the density of the carbon fiber preform increases from the surface to the inside; S2. A low-density C / C substrate is subjected to a first vacuum impregnation, a first pressure impregnation, a first curing, and a first pyrolysis to obtain a porous C / C intermediate using a phenolic resin; S3. Using an alloy infiltration agent to react and infiltrate the porous C / C intermediate to generate a multi-element ultra-high temperature ceramic matrix to obtain a multi-element ultra-high temperature ceramic matrix composite material; S4. Prepare an anti-oxidation ceramic protective layer near the surface of the multi-component ultra-high temperature ceramic-based composite material to obtain a high-temperature structural load-bearing / anti-ablation integrated ceramic-based composite material.

2. The preparation method according to claim 1, characterized in that: The carbon fiber preform has a density increasing gradient from the surface to the interior of 0.05 to 0.1 g / cm 3 ; The surface density of the carbon fiber preform is 0.30-0.35 g / cm 3 ; The maximum density inside the carbon fiber preform is 0.60-0.65 g / cm 3 ; and / or Compared with the carbon fiber preform, the density of the low-density C / C substrate is increased by 60 to 100%.

3. The preparation method according to claim 1, characterized in that: Before depositing the pyrolytic carbon interface layer on the fiber surface of the carbon fiber preform by chemical vapor deposition, the carbon fiber preform is subjected to high temperature treatment; the high temperature heat treatment temperature is 1000-1800°C and the time is 2.0-3.0h; and / or The chemical vapor deposition method uses propylene as a carbon source and nitrogen as a carrier, and performs deposition at 800-1100°C.

4. The preparation method according to claim 1, characterized in that: The vacuum degree of the first vacuum impregnation is -0.15 to -0.08 MPa, the temperature is room temperature, and the time is 0.5 to 2.0 h; The first pressure impregnation has a pressure of 0.5 to 2.0 MPa, a temperature of room temperature, and a time of 0.5 to 2.0 h; The first curing is performed at a temperature of 100 to 350° C., a pressure of 0.5 to 2.5 MPa, and a time of 0.5 to 2.0 h; and / or The temperature of the first cracking is 700-1000° C., and the time is 2.0-4.0 hours.

5. The preparation method according to claim 4, characterized in that: Compared with the low-density C / C substrate, the density of the porous C / C intermediate is increased by 20-50%.

6. The preparation method according to claim 1, characterized in that: The alloying agent is one of a silicon-zirconium binary alloy and a silicon-hafnium binary alloy; preferably, the atomic weight fraction of zirconium in the silicon-zirconium alloy is 15-25%; the atomic weight fraction of hafnium in the silicon-hafnium alloy is 15-31%; the particle size of the alloying agent is 1-10 μm; and / or The temperature of the reaction infiltration is 1600-1800° C., and the time of the reaction infiltration is 1.0-2.0 h.

7. The preparation method according to claim 1, characterized in that: Preparation of an anti-oxidation ceramic protective layer near the surface of a multi-element ultra-high temperature ceramic matrix composite material, including: The multi-component ultra-high temperature ceramic-based composite material is subjected to a second vacuum impregnation, a second pressure impregnation, a second curing and a second cracking by using a first ceramic precursor to form a ceramic transition layer near the surface of the multi-component ultra-high temperature ceramic-based composite material; the first ceramic precursor is a silicon-based ceramic precursor; A second ceramic precursor is used to perform a third vacuum impregnation, a third pressure impregnation, a third curing and a third cracking on the composite material that forms a silicon-based ceramic transition layer near the surface to form a ceramic surface layer, thereby obtaining an antioxidant ceramic protective layer; the second ceramic precursor is a mixed ceramic precursor including a silicon-based ceramic precursor and an ultra-high temperature ceramic precursor.

8. The preparation method according to claim 7, characterized in that: The silicon-based ceramic precursor is polycarbosilane; and / or The ultra-high temperature ceramic precursor is at least one of a ZrC ceramic precursor and a HfC ceramic precursor.

9. The preparation method according to claim 7, characterized in that: The vacuum degree of the second vacuum impregnation and the third vacuum impregnation is -0.15 to -0.08 MPa, the temperature is room temperature, and the time is 0.5 to 2.0 h; The pressure of the second pressure impregnation and the third pressure impregnation is 0.5-2.0 MPa, the temperature is room temperature, and the time is 0.5-2.0 h; The temperature of the second curing and the third curing is 100-350° C., the pressure is 0.5-2.5 MPa, and the time is 0.5-2.0 h; The temperature of the second cracking is 700-1000°C and the time is 2.0-4.0h; and / or The temperature of the third cracking is 1400-1600° C., and the time is 2.0-4.0 hours.

10. A high-temperature structural load-bearing / anti-ablation integrated ceramic matrix composite material, characterized in that: The method is prepared by the method according to any one of claims 1 to 9.

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