Photocuring molding method for ceramic matrix composite

Abstract

The present invention relates to the technical field of ceramic matrix composite preparation, and specifically relates to a photocuring molding method for a ceramic matrix composite. The method uses an ultraviolet-light-curable ceramic precursor to fabricate a protective layer on the surface of an impregnated composite, and controls the volatilization rate of an impregnation matrix in the composite by controlling the thickness of the protective layer. The main function of the protective layer is to reduce the volatilization of small molecules in the impregnation matrix during high-temperature pyrolysis of the composite, thereby improving the curing efficiency of the composite, simplifying the curing process, and effectively improving the preparation efficiency of the ceramic matrix composite.

Description

A ceramic matrix composite light curing forming method

[0001] The present application claims priority to the Chinese patent application No. 202410848724.5, filed on June 27, 2024, and entitled "A ceramic matrix composite light curing forming method", the content of which is incorporated herein by reference in its entirety. TECHNICAL FIELD

[0002] The present application relates to the technical field of ceramic matrix composite manufacturing, and particularly to a ceramic matrix composite light curing forming method. BACKGROUND

[0003] Continuous fiber reinforced ceramic matrix composites have great potential as high-temperature structural materials in the fields of aerospace, space industry and energy industry due to their good mechanical properties, high temperature resistance, strong oxidation and ablation resistance, and many other advantages. Continuous fiber reinforced ceramic matrix composites have become the preferred material solution for the thermal structure of the new generation of near space vehicles.

[0004] Currently, ceramic matrix composites are mostly manufactured using PIP technology. Under certain temperature and pressure, carbon fiber preforms are used as the framework, and organic precursor solution is used for impregnation. After cross-linking and curing, high-temperature pyrolysis is performed under inert gas protection to convert the precursor into a ceramic matrix. PIP technology has low cost, simple equipment requirements, and the precursor molecules can be designed, allowing the manufacture of large and complex shaped components. However, due to the overflow of small molecules during the traditional precursor pyrolysis process, and the increase in the density of the pyrolysis products, the volume of the precursor shrinks after pyrolysis, forming pores and defects in the material manufacturing process, which affects the performance of the ceramic matrix composite and becomes a major challenge in the manufacture of ceramic matrix composites using PIP. In order to obtain a dense composite material, several cycles of "impregnation-curing-pyrolysis" process are required to obtain a dense ceramic matrix composite. After the impregnation process, metal tooling is added to the surface of the composite material for high-temperature curing treatment, and after the curing process is completed, graphite tooling is replaced for high-temperature pyrolysis. In this process, the curing process is time-consuming and complex.

[0005] SUMMARY

[0006] The present application provides a ceramic matrix composite light curing forming method, which aims to provide an alternative forming scheme for ceramic matrix composites, which can overcome the shortcomings of existing ceramic matrix composite curing process technology.

[0007] In a first aspect, a ceramic matrix composite light curing forming method is provided, comprising:

[0008] The impregnation process is as follows: the fiber preform is immersed in an immersion barrel containing a ceramic precursor impregnation matrix solution, after immersion, the preform is taken out of the immersion barrel, and the excess impregnation matrix on the surface is scraped off;

[0009] The ultraviolet light curing process is as follows: the prepared ultraviolet light precursor solution is sprayed on the surface of the impregnated composite material, ultraviolet light is irradiated on the surface of the composite material to form a protective layer on the surface of the composite material, and finally the composite material is placed in a graphite tooling;

[0010] The high-temperature pyrolysis process is as follows: the graphite tooling containing the composite material is loaded into a high-temperature furnace, and high-temperature pyrolysis is carried out according to the high-temperature pyrolysis system, and finally a ceramic matrix composite material is obtained.

[0011] In combination with the first aspect, in some implementations of the first aspect, the ultraviolet light curing process satisfies at least one of the following:

[0012] The ultraviolet irradiation is performed at a distance of 0.1 cm to 10 cm from the surface of the composite material;

[0013] The intensity of the ultraviolet light is 30-80 mW / cm 3 ;

[0014] The irradiation time is 2-30 min;

[0015] The irradiation angle is -45° to +45°;

[0016] The number of times of spraying the ultraviolet light precursor solution is controlled to be 3-20 times.

[0017] In combination with the first aspect, in some implementations of the first aspect, the thickness of the protective layer obtained by ultraviolet light curing is 30-600 μm.

[0018] In combination with the first aspect, in some implementations of the first aspect, the ceramic matrix composite material obtained finally is used for a heat insulation component of an aircraft, and the thickness of the protective layer is 30-150 μm.

[0019] In combination with the first aspect, in some implementations of the first aspect, the ceramic matrix composite material obtained finally is used for a heat-resistant structural component of an aircraft, and the thickness of the protective layer is 150-600 μm.

[0020] In combination with the first aspect, in some implementations of the first aspect, the impregnation matrix solution includes at least one of the following organic precursor solutions: polycarbosilane, polyborosilazane, polysilazane.

[0021] In combination with the first aspect, in some implementations of the first aspect, the fiber used in the fiber preform includes at least one of the following: carbon fiber, silicon carbide fiber, Al-containing ceramic fiber, Hf-containing ceramic fiber, Zr-containing ceramic fiber, Ti-containing ceramic fiber.

[0022] With reference to the first aspect, in some embodiments of the first aspect, the ultraviolet light precursor is a resin containing acrylic monomers.

[0023] With reference to the first aspect, in some embodiments of the first aspect, the ultraviolet light precursor comprises at least one of: a SiC ceramic precursor containing acrylic groups, a SiBCN ceramic precursor containing acrylic groups, an acrylic resin.

[0024] With reference to the first aspect, in some embodiments of the first aspect, the high-temperature pyrolysis procedure comprises two stages, in the first stage, the composite material is heated at a heating rate of 2-15 ℃ / min to 150-350 ℃, and held for 1-4 h, in the second stage, the composite material is heated at a heating rate of 5-10 ℃ / min to 800-1100 ℃.

[0025] With reference to the first aspect, in some embodiments of the first aspect, the manufacturing steps of the ceramic matrix composite material comprise a plurality of process cycles, the plurality of process cycles comprises a process cycle A and a process cycle B, the process cycle A is an impregnation process-ultraviolet light curing process-high-temperature pyrolysis process, the process cycle B is an impregnation process-thermal curing process-high-temperature pyrolysis process; the process cycle B is performed before the process cycle A.

[0026] With reference to the first aspect, in some embodiments of the first aspect, the thermal curing process comprises: placing the impregnated composite material into a metal tooling, placing it in an oven for curing at 300-400 ℃ for 2-4 h, and after the thermal curing process is completed, the composite material is taken out of the oven and reloaded into a graphite tooling.

[0027] With reference to the first aspect, in some embodiments of the first aspect, the process cycle A is performed after the process cycle B is completed, and the process cycle A is used in the stage where the weight gain rate of the impregnated matrix is less than 10%.

[0028] With reference to the first aspect, in some embodiments of the first aspect, the number of times of performing the process cycle B is 2-4 times; the number of times of performing the process cycle A is 1-4 times.

[0029] With reference to the first aspect, in some embodiments of the first aspect, after the process cycle B is completed, the density of the composite material is greater than 1.3 g / cm 3 .

[0030] With reference to the first aspect, in some embodiments of the first aspect, after the process cycle B is completed, the final porosity of the composite material is 5-30%.

[0031] In a second aspect, a ceramic matrix composite is provided, which is prepared by the method according to any one of the implementation manners of the first aspect.

[0032] In a third aspect, an aircraft is provided, which comprises the ceramic matrix composite according to any one of the implementation manners of the second aspect, and the ceramic matrix composite is applied to a thermal insulation part or a thermal protection structure part of the aircraft.

[0033] Compared with the prior art, the scheme provided in the application has at least the following beneficial technical effects:

[0034] (1) The application simplifies the curing process, and only a protective layer is manufactured on the surface of the composite material by ultraviolet light curing of the precursor, so that overflow of the impregnated matrix is prevented by the protective layer.

[0035] (2) The protective layer is manufactured by ultraviolet light curing, which reduces the volatilization of small molecules in the impregnated matrix during high-temperature pyrolysis of the composite material, and improves the manufacturing efficiency of the composite material.

[0036] (3) The protective layer manufactured by the application can be cured by ultraviolet light, has low weight loss rate, simple curing process, short curing time, and low energy consumption.

[0037] (4) The protective layer manufactured by the application can effectively control the thickness of the protective layer by adjusting the type and spraying times of the precursor, so as to control the volatilization rate of small molecules in the impregnated matrix. BRIEF DESCRIPTION OF DRAWINGS

[0038] FIG. 1 is a TG diagram of the cured protective layer.

[0039] FIG. 2 is an SEM diagram of the cured protective layer.

[0040] FIG. 3 is a diagram of the correspondence between the irradiation intensity and the curing degree of the light-cured precursor. DETAILED DESCRIPTION

[0041] The application will be described in further detail below with reference to the drawings and specific embodiments.

[0042] The application relates to a ceramic matrix composite light-curing forming process, which comprises an impregnation process, an ultraviolet light curing process and a high-temperature pyrolysis process. The impregnation process, the ultraviolet light curing process and the high-temperature pyrolysis process can be cyclically performed to obtain a dense ceramic matrix composite.

[0043] (1) Impregnation process

[0044] The fiber preform is immersed in an immersion barrel containing a ceramic precursor impregnated matrix solution, and after immersion is completed, the preform is taken out of the immersion barrel, and the excess impregnated matrix on the surface is scraped off.

[0045] (2) UV curing process

[0046] The prepared UV precursor solution is sprayed on the surface of the impregnated composite material, and then a certain intensity of UV light is irradiated on the surface of the composite material for a certain time according to a certain UV curing system. Finally, the composite material is placed in a graphite tool.

[0047] The UV-cured ceramic precursor undergoes crosslinking and curing reaction under UV irradiation, and a cured product with a certain degree of crosslinking is formed, that is, a protective layer is formed on the surface of the composite material. The cured product is basically not melted at 300°C (as shown in FIG. 1, the cured product obtained by irradiating under UV light with an intensity of 30 W / cm 3 for 5 min has a weight loss of <6 wt% before 300°C). In addition, after using the UV-cured precursor to manufacture the protective layer, the volatilization of small molecules of the precursor can be effectively prevented, so that the precursor ceramic conversion rate is improved, and the densification efficiency of the composite material is improved. The protective layer is shown in the interval in FIG. 2.

[0048] (3) High-temperature pyrolysis process

[0049] The prepared tool is loaded into a high-temperature furnace, and high-temperature pyrolysis is performed according to a certain high-temperature pyrolysis system, and finally a ceramic matrix composite material is obtained.

[0050] In some embodiments, the UV curing process satisfies at least one of the following: UV irradiation is performed at a distance of 0.1 cm-10 cm from the surface of the composite material; the intensity of the UV light is 30-80 mW / cm 3 ; the irradiation time is 2-30 min; the irradiation angle is -45°~+45°, and the irradiation angle of 0° is vertical irradiation.

[0051] Different UV irradiation intensities result in different degrees of crosslinking and curing of the protective layer. The higher the intensity of the UV light, the higher the curing degree of the protective layer, and the protective layer is close to a densified film in the temperature range of 200-300°C. If the curing degree of the protective layer is low, the protective layer is a porous film in the temperature range of 200-300°C. The present application can control the curing degree of the protective layer by controlling the irradiation intensity or irradiation time of the protective layer. The specific corresponding relationship is shown in FIG. 2.

[0052] In some embodiments, the thickness of the protective layer obtained by UV curing is 30-600 μm.

[0053] Different application environments require different porosity of composite materials, and the thermal insulation parts of aircrafts often use porous ceramic matrix composites, and the heat-proof structural parts often use densified ceramic matrix composites. In view of the different densification requirements of the ceramic matrix composites, the volatilization rate of the small molecules of the impregnated matrix can be controlled by adjusting the thickness of the protective layer, and then the densification degree of the composite material can be finally adjusted by adjusting the thickness of the protective layer. When the thickness of the protective layer is 30-150 μm, the porous structure is formed at 200-300 ℃, and the volatilization amount of the small molecules of the impregnated matrix is >10%. When the thickness of the protective layer is >150 μm, the volatilization amount of the impregnated matrix is <5 wt%, and the densification degree of the composite material can be finally adjusted by adjusting the thickness of the protective layer, and the porosity of the composite material changes in the range of 5-30%. The spraying times are controlled in the range of 3-20 times according to the requirements, and the SEM diagram of the protective layer is shown in Fig. 2.

[0054] In some embodiments, the impregnated matrix includes at least one of a polycarbosilane, a polyborosilazane, a polysilazane, and other organic precursor solutions.

[0055] In some embodiments, the fibers used in the fiber preform include at least one of carbon fibers, silicon carbide fibers, aluminum-containing ceramic fibers, and other refractory metal (Hf, Zr, Ti, etc.) containing ceramic fibers.

[0056] In some embodiments, the ultraviolet light curing resin includes at least one of an SiC ceramic precursor containing an acrylic group, an SiBCN ceramic precursor containing an acrylic group, an acrylic resin, or other resins containing an acrylic monomer.

[0057] In some embodiments, the high-temperature pyrolysis system mainly includes two stages. In the first stage, the composite material is heated to 150-350 ℃ at a heating rate of 2-15 ℃ / min, and is kept for 1-4 h. In the second stage, the composite material is heated to 800-1100 ℃ at a heating rate of 5-10 ℃ / min.

[0058] In some other embodiments provided by the present application, the manufacturing process of the ceramic matrix composite material can further include a heat curing process. That is, the manufacturing steps of the ceramic matrix composite material can include multiple process cycles, and the multiple process cycles include a process cycle A and a process cycle B. The process cycle A is an impregnation process-ultraviolet light curing process-high temperature pyrolysis process, and the process cycle B is an impregnation process-heat curing process-high temperature pyrolysis process. The process cycle B can be executed before the process cycle A. In some embodiments, the number of times of execution of the process cycle B can be 2-4 times. After the process cycle B is executed completely, the process cycle A can be executed. The number of times of execution of the process cycle A can be 1-4 times.

[0059] The following describes a manufacturing step of a ceramic matrix composite material, which can include 6 cycles of composite processes, as follows.

[0060] First to third round composite process:

[0061] (1) Impregnation process

[0062] The fiber preform is immersed in an immersion barrel containing a ceramic precursor impregnation matrix solution, soaked at room temperature for 4h, and after soaking, the preform is taken out of the immersion barrel and the excess impregnation matrix on the surface is scraped off.

[0063] (2) Thermal curing process

[0064] The impregnated composite material is placed in a metal tooling and placed in an oven for curing at 300-400°C for 2-4h, and after the thermal curing process is completed, the composite material is taken out of the oven and reloaded into graphite tooling.

[0065] (3) High temperature pyrolysis process

[0066] The prepared tooling is loaded into a high temperature furnace and high temperature pyrolysis is carried out according to a certain high temperature pyrolysis schedule, and finally a ceramic matrix composite material is obtained.

[0067] The first to third round composite process mainly adopts the "impregnation-thermal curing-high temperature pyrolysis" process route, the main reason is that the early stage density of the composite material is low, the weight gain rate of the composite material after impregnation process is >15%, and during high temperature pyrolysis, the volatilization amount of small molecules of the precursor is large, which will cause the protective layer made of ultraviolet light curing precursor to crack, and the protective effect is relatively limited, while the thermal curing process can realize the curing of the impregnation matrix, and after the first three rounds of composite process, the density of the composite material is greater than 1.3g / cm 3 .

[0068] Fourth to sixth round composite process:

[0069] (1) Impregnation process

[0070] The fiber preform is immersed in an immersion barrel containing a ceramic precursor impregnation matrix solution, soaked at room temperature for 4h, and after soaking, the preform is taken out of the immersion barrel and the excess impregnation matrix on the surface is scraped off.

[0071] (2) Ultraviolet light curing process

[0072] The prepared ultraviolet light precursor solution is sprayed on the surface of the impregnated composite material, and then according to a certain ultraviolet light curing schedule, a certain intensity of ultraviolet light is irradiated on the surface of the composite material for a certain time, and finally the composite material is placed in graphite tooling. The specific implementation of the ultraviolet light curing process can refer to the foregoing.

[0073] (3) High temperature pyrolysis process

[0074] The prepared tooling is loaded into a high-temperature furnace, and high-temperature pyrolysis is carried out according to a certain high-temperature pyrolysis system, and finally a ceramic matrix composite material is obtained. The final porosity of the composite material is 8%. The composite material can be used for heat-resistant structural parts of an aircraft, for example.

[0075] The 4th-6th round composite process mainly adopts the process route of "dipping-ultraviolet light curing-high temperature pyrolysis". Since the weight gain rate of the matrix in the last three rounds is less than 10%, after the protective layer is made by using the ultraviolet light curing precursor, the volatilization of small molecules of the precursor can be effectively prevented, the small molecules can be fully cross-linked and cured in the protective layer, the ceramic conversion rate of the precursor is improved, and the densification efficiency of the composite material is improved.

[0076] Example 1

[0077] A piece of 200mm×200mm fiber preform is placed in a solid polycarbosilane toluene solution dipping matrix, and soaked at room temperature for 4h. After the dipping is completed, the composite material is taken out. The composite material is loaded into a graphite tooling and placed into an oven, and the temperature is raised to 400℃ at a heating rate of 10℃ / min, and held for 2h. After natural cooling, it is taken out, reloaded into the graphite tooling, and the temperature is raised to 1000℃ at a heating rate of 10℃ / min for high-temperature treatment. This process is repeated 3 times.

[0078] An ultraviolet light curable SiC ceramic precursor is used as a spraying slurry, and the prepared slurry is sprayed on the surface of the composite material by using a spraying machine. The composite material is placed under a ultraviolet light lamp at a distance of 1cm, the light intensity of the ultraviolet light is 40mW / cm 3 , irradiation is carried out for 5min, and the spraying is repeated 3 times. After the protective layer is cured, the composite material is finally placed in a graphite tooling, loaded into a high-temperature pyrolysis furnace, and the temperature is raised to 180℃ at a heating rate of 3℃ / min and held for 2h. Then the temperature is raised to 1000℃ at a heating rate of 10℃ / min for high-temperature treatment. This process is only carried out once, and the final porosity of the composite material is 28.2%. The composite material prepared in Example 1 can be used for heat insulation parts of an aircraft, for example.

[0079] Example 2

[0080] A piece of 200mm×200mm fiber preform is placed in a solid polycarbosilane toluene solution dipping matrix, and soaked at room temperature for 4h. After the dipping is completed, the composite material is taken out. The composite material is loaded into a graphite tooling and placed into an oven, and the temperature is raised to 400℃ at a heating rate of 10℃ / min, and held for 2h. After natural cooling, it is taken out, reloaded into the graphite tooling, and the temperature is raised to 1000℃ at a heating rate of 10℃ / min for high-temperature treatment. This process is repeated 3 times.

[0081] The prepared slurry is sprayed on the surface of the composite material by using a spraying machine, the composite material is placed under a UV light lamp at a distance of 5 cm, the light intensity of the UV light is 80 mW / cm 3 , irradiation is performed for 20 min, the spraying is repeated for 10 times, after the protective layer is cured, finally the composite material is placed in a graphite tool, is loaded into a high temperature pyrolysis furnace, the temperature is increased to 180°C at a heating rate of 2°C / min and is kept for 2 h, then the temperature is increased to 1000°C at a heating rate of 5°C / min for high temperature treatment, the process is repeated for 2 times, and the final porosity of the composite material is 15.3%. The composite material prepared in Example 2 can be used in, for example, a heat insulation part or a heat-proof structural part of an aircraft.

[0082] Although the present application is disclosed with preferred embodiments, it is not intended to limit the present application, any person skilled in the art can make possible changes and modifications without departing from the spirit and scope of the present application, therefore, the protection scope of the present application should be defined by the scope of the claims of the present application.

Claims

1. A method of photocuring molding of a ceramic matrix composite material, characterized by, The application relates to a manufacturing method of a ceramic matrix composite material. The method comprises the following steps: an impregnation process: a fiber preform is immersed into an immersion barrel filled with a ceramic precursor impregnation matrix solution, after immersion, the preform is taken out of the immersion barrel, and the surface of the preform is scraped to remove the excess impregnation matrix; an ultraviolet light curing process: a prepared ultraviolet light precursor solution is sprayed on the surface of the impregnated composite material, ultraviolet light is irradiated on the surface of the composite material to form a protective layer on the surface of the composite material, and finally the composite material is placed in a graphite tooling; and a high-temperature pyrolysis process: the graphite tooling containing the composite material is loaded into a high-temperature furnace, high-temperature pyrolysis is carried out according to a high-temperature pyrolysis system, and finally a ceramic matrix composite material is obtained. The ultraviolet light curing process satisfies at least one of the following conditions: ultraviolet irradiation is carried out at a position of 0.1 cm-10 cm on the surface of the composite material; the irradiation time is 2-30 min; the irradiation angle is -45 DEG ~ +45 DEG; and the spraying times of the ultraviolet light precursor solution are controlled to be 3-20 times. The thickness of the protective layer obtained through the ultraviolet light curing process is 30-600 mu m.

2. The method of claim 1, wherein, The ceramic matrix composite material obtained finally is used for a heat insulation component of an aircraft, and the thickness of the protective layer is 30-150 mu m. The ceramic matrix composite material obtained finally is used for a heat insulation component of an aircraft, and the thickness of the protective layer is 30-150 mu m. The intensity of the UV light is 30-80 mW / cm 3 ; The ceramic matrix composite material obtained finally is used for a heat insulation component of an aircraft, and the thickness of the protective layer is 30-150 mu m. The impregnation matrix solution comprises at least one of the following organic precursor solutions: polycarbosilane, polysilazane, polysilazane. The fiber used in the fiber preform comprises at least one of the following: carbon fiber, silicon carbide fiber, Al-containing ceramic fiber, Hf-containing ceramic fiber, Zr-containing ceramic fiber, Ti-containing ceramic fiber.

3. The method according to claim 1 or 2, characterized in that, The ultraviolet light precursor is a resin containing an acrylic monomer.

4. The method of claim 3, wherein, The ultraviolet light precursor comprises at least one of the following: SiC ceramic precursor containing an acrylic group, SiBCN ceramic precursor containing an acrylic group, and acrylic resin.

5. The method of claim 3, wherein, The high-temperature pyrolysis system comprises two stages, in the first stage, the composite material is heated to 150-350 DEG C at a heating rate of 2-15 DEG C / min, and is kept at the temperature for 1-4 h, and in the second stage, the composite material is heated to 800-1100 DEG C at a heating rate of 5-10 DEG C / min.

6. The method according to any one of claims 1 to 5, characterized in that, The manufacturing steps of the ceramic matrix composite material comprise multiple process cycles, the multiple process cycles comprise process cycle A and process cycle B, the process cycle A is an impregnation process-ultraviolet light curing process-high-temperature pyrolysis process, and the process cycle B is an impregnation process-thermal curing process-high-temperature pyrolysis process.

7. The method according to any one of claims 1 to 6, characterized in that, The process cycle B is executed before the process cycle A.

8. The method according to any one of claims 1 to 7, characterized in that, The thermal curing process comprises the following steps: the impregnated composite material is placed into a metal tooling, and is placed into an oven to be cured at 300-400 DEG C for 2-4 h, after the thermal curing process is completed, the composite material is taken out of the oven, and is reloaded into the graphite tooling.

9. The method according to any one of claims 1 to 8, characterized in that, After the process cycle B is completely executed, the process cycle A is executed, and the process cycle A is used in the stage of an impregnation matrix weight gain rate <10%.

10. The method according to any one of claims 1 to 9, characterized in that, The execution times of the process cycle B are 2-4 times, and the execution times of the process cycle A are 1-4 times.

11. The method according to any one of claims 1 to 10, characterized in that, After the process cycle B is completely executed, the final porosity of the composite material is 5-30%. ​ 12. The method of claim 11, wherein, ​ 13. The method according to claim 11 or 12, characterized in that, ​ 14. The method of claim 13, wherein, ​ 15. The method according to any one of claims 11 to 14, characterized in that, After the process cycle B is completed, the density of the composite material is greater than 1.3 g / cm 3 .

16. The method according to any one of claims 11 to 15, characterized in that, ​ 17. A ceramic matrix composite material, characterized by, The ceramic matrix composite is prepared by the method according to any one of claims 1 to 16.

18. An aircraft, characterized in that The aircraft comprises the ceramic matrix composite according to claim 17, which is applied to a heat-insulating part or a heat-resistant structural part of the aircraft.

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