Carbon-silicon carbide honeycomb sandwich structure material as well as preparation method and application thereof

Abstract

The invention relates to the technical field of composite materials, in particular to a carbon-silicon carbide honeycomb sandwich structure material and a preparation method and application thereof. The carbon-silicon carbide honeycomb sandwich structure material provided by the invention has excellent high-temperature stability and mechanical properties, on the one hand, light weight is realized, the overall structure efficiency is improved, on the other hand, functional diversification is realized, the material has light weight, high bearing capacity and high temperature resistance, the density is 1.8-2.0 g / cm < 3 >, and the compression strength in an air atmosphere at 1200 DEG C is greater than or equal to 10 MPa. The carbon-silicon carbide honeycomb sandwich structure material provided by the invention meets the multi-target requirements of light weight, high bearing capacity and high temperature resistance, and can be widely applied to hypersonic thermal protection structures, such as wing rudders, end heads or throat liners.

Description

Technical Field

[0001] This invention relates to the field of composite materials technology, and in particular to a carbon-silicon carbide honeycomb sandwich structure material, its preparation method, and its application. Background Technology

[0002] To achieve high reliability and repeatability in the actual operation of hypersonic vehicles, their thermal structures and thermal protection components require excellent performance, such as lightweight, oxidation resistance, ablation resistance, corrosion and wear resistance, and good high-temperature mechanical properties. This places stringent requirements on the design and manufacture of thermal protection systems. Carbon-silicon carbide composite materials possess excellent properties such as lightweight, high strength, high temperature resistance, wear resistance, low coefficient of thermal expansion, and good thermal stability, making them one of the candidate materials for the aforementioned thermal structures and thermal protection components.

[0003] Traditional carbon-silicon carbide composite materials are typically used in bulk or sheet form for critical components of hypersonic aircraft, such as the nose, wing surfaces, and throat liners. This results in low material utilization, particularly in large-scale applications where overall structural efficiency is low, hindering lightweight requirements. Furthermore, traditional carbon-silicon carbide composite materials have limited functionality and cannot meet the multi-objective requirements of lightweight, high load-bearing capacity, and high-temperature resistance for critical thermal structural components. Summary of the Invention

[0004] In view of this, the present invention provides a carbon-silicon carbide honeycomb sandwich structure material, its preparation method and application. The carbon-silicon carbide honeycomb sandwich structure material provided by the present invention achieves the integration of lightweight, high load-bearing capacity and high temperature resistance.

[0005] This invention provides a carbon-silicon carbide honeycomb sandwich structure material, comprising a matrix and a composite layer covering the surface of the matrix; the matrix comprises a honeycomb core and panels covering the upper and lower surfaces of the honeycomb core; the matrix is ​​made of carbon fiber woven material; the composite layer comprises carbon and silicon carbide.

[0006] Preferably, the parameters of the honeycomb core include: a honeycomb cell side length of 3~16mm, a honeycomb cell wall thickness of 0.2~0.4mm, and a honeycomb height of 10~44mm.

[0007] Preferably, the porosity of the honeycomb core is less than 5%, and the compressive strength is not less than 8 MPa.

[0008] Preferably, the thickness of the panel is 1.5~3mm.

[0009] The present invention also provides a method for preparing the carbon-silicon carbide honeycomb sandwich structure material described above, comprising the following steps: (1) The carbon cloth is layered and woven into a planar fiber layer. The interlayer fibers are sewn into the two adjacent carbon cloth layers to obtain a planar preform with a semi-closed honeycomb cell. Then, a core mold is inserted into the planar preform and locked using a shape control fixture to obtain a honeycomb core preform. (2) The honeycomb core preform is sequentially impregnated with resin, cured and deposited with pyrolytic carbon to obtain a honeycomb core; (3) After shaping the panel preform, pyrolytic carbon is deposited to obtain the panel; (4) The honeycomb core and the two panels are impregnated with polycarbosilane, co-cured and pyrolyzed in sequence, and the polycarbosilane impregnation, co-curing and pyrolysis are repeated 8 to 12 times to obtain the carbon-silicon carbide honeycomb sandwich structure material. There is no requirement for the time order of steps (1) to (2) and step (3).

[0010] Preferably, the raw material of the carbon cloth includes one or more of T300 fiber, T700 fiber, T80 fiber, M40 fiber, and M55 fiber; the areal density of the raw material of the carbon cloth is 200 g / m³. 3 The weaving angle of the carbon cloth is +45° / -45°; the warp and weft density of the planar fiber layer is 6 threads / cm × 6 threads / cm or 7 threads / cm × 7 threads / cm.

[0011] Preferably, the curing conditions include: increasing the temperature from room temperature to 100 degrees Celsius at a rate of 5°C / min, holding at 100 degrees Celsius for 2 hours, increasing the temperature from 100 degrees Celsius to 120 degrees Celsius at a rate of 1°C / min, holding at 120 degrees Celsius for 2 hours, increasing the temperature from 120 degrees Celsius to 160 degrees Celsius at a rate of 2°C / min, holding at 160 degrees Celsius for 2 hours, and then cooling in the furnace.

[0012] Preferably, the co-curing molding includes a first curing stage and a second curing stage performed sequentially; the temperature of the first curing stage is 120 degrees Celsius, the pressure is 0.3~0.5 MPa, and the holding time is 180 minutes; the temperature of the second curing stage is 150 degrees Celsius, the pressure is 0.3~0.5 MPa, and the holding time is 180 minutes.

[0013] Preferably, the pyrolysis is carried out in a protective atmosphere; the pyrolysis temperature is 1100~1300 degrees Celsius, and the holding time is 2 hours; after each pyrolysis, the product is further subjected to heat treatment; the heat treatment temperature is 1800~2000 degrees Celsius, and the holding time is 2 hours.

[0014] The present invention also provides the application of the carbon-silicon carbide honeycomb sandwich structure material described in the above-described scheme or the carbon-silicon carbide honeycomb sandwich structure material obtained by the preparation method described in the above-described scheme in the field of aircraft.

[0015] This invention provides a carbon-silicon carbide honeycomb sandwich structure material. The carbon-silicon carbide honeycomb sandwich structure material provided by this invention possesses excellent high-temperature stability and mechanical properties. On the one hand, it achieves lightweighting and improves overall structural efficiency; on the other hand, it achieves functional versatility, combining lightweight, high load-bearing capacity, and high-temperature resistance, with a density of 1.8~2.0 g / cm³. 3 The compressive strength in air at 1200℃ is ≥10MPa.

[0016] This invention also provides a method for preparing the carbon-silicon carbide honeycomb sandwich structure material described above. This invention utilizes an interlayer interlocking method of planar fibers to achieve overall honeycomb weaving. Compared to traditional adhesive-bonded honeycombs, this effectively solves the problem of W-direction node cracking in honeycombs, exhibiting excellent load-bearing and high-temperature resistance. This invention utilizes chemical vapor infiltration (CVI) deposition of pyrolytic carbon interfaces to fill the small pores within the fiber bundles in the matrix, while simultaneously reducing shrinkage deformation during subsequent precursor impregnation pyrolysis (PIP), thus improving the bonding strength between carbon and the matrix. During precursor impregnation pyrolysis, polycarbosilane is impregnated to fill the large pores between fiber bundles in the matrix, significantly reducing material porosity, optimizing the material interface, protecting the carbon fibers in the matrix, and improving the material's high-temperature stability and mechanical properties. This invention integrates the panel and honeycomb core through co-curing molding, and through multiple impregnation, curing, and pyrolysis processes, ensures the connection strength between the honeycomb core and the panel, improving the shear and bending resistance of the honeycomb sandwich structure. Furthermore, the preparation method provided by this invention effectively shortens the preparation cycle and improves densification efficiency.

[0017] This invention also provides the application of the carbon-silicon carbide honeycomb sandwich structure material described in the above-described scheme or the carbon-silicon carbide honeycomb sandwich structure material prepared by the above-described scheme in the field of aircraft. The carbon-silicon carbide honeycomb sandwich structure material provided by this invention meets the multi-objective requirements of lightweight, high load-bearing capacity and high temperature resistance, and can be widely used in hypersonic thermal protection structures, such as wing rudders, nose cones or throat liners. Attached Figure Description

[0018] To more clearly illustrate the technical solutions of this invention, the accompanying drawings used in the embodiments of this invention or in the prior art are briefly described below. For those skilled in the art, other drawings can be derived from the following drawings without creative effort, and all such drawings are within the protection scope of this invention.

[0019] Figure 1 Flowchart for the preparation of carbon-silicon carbide honeycomb sandwich structure material; Figure 2 This is a schematic diagram illustrating the interlayer interlocking fibers of the honeycomb core preform of the present invention. Figure 3 A schematic diagram of a hexagonal core mold being implanted into honeycomb fabric; Figure 4 This is a schematic diagram of the assembly of the panel and the honeycomb core; Figure 5 The diagram shows the high-temperature mechanical properties of carbon-silicon carbide honeycomb sandwich structure materials. Detailed Implementation

[0020] This invention provides a carbon-silicon carbide honeycomb sandwich structure material, comprising a matrix and a composite layer covering the surface of the matrix; the matrix comprises a honeycomb core and panels covering the upper and lower surfaces of the honeycomb core; the matrix is ​​made of carbon fiber woven material; the composite layer comprises carbon and silicon carbide.

[0021] In this invention, the parameters of the honeycomb core preferably include: the side length of the honeycomb core is 3~16mm, more preferably 5~9mm, the wall thickness of the honeycomb core is 0.2~0.4mm, more preferably 0.3mm, and the honeycomb height is 10~44mm, more preferably 20~36.5mm.

[0022] In this invention, the porosity of the honeycomb core is preferably less than 5%, and the compressive strength is preferably not less than 8 MPa.

[0023] In this invention, the thickness of the panel is preferably 1.5~3mm, more preferably 2mm.

[0024] The present invention also provides a method for preparing the carbon-silicon carbide honeycomb sandwich structure material described above, comprising the following steps: (1) The carbon cloth is layered and woven into a planar fiber layer. The interlayer fibers are sewn into the two adjacent carbon cloth layers to obtain a planar preform with a semi-closed honeycomb cell. Then, a core mold is inserted into the planar preform and locked using a shape control fixture to obtain a honeycomb core preform. (2) The honeycomb core preform is sequentially impregnated with resin, cured and deposited with pyrolytic carbon to obtain a honeycomb core; (3) After shaping the panel preform, pyrolytic carbon is deposited to obtain the panel; (4) The honeycomb core and the two panels are impregnated with polycarbosilane, co-cured and pyrolyzed in sequence, and the polycarbosilane impregnation, co-curing and pyrolysis are repeated 8 to 12 times to obtain the carbon-silicon carbide honeycomb sandwich structure material. There is no requirement for the time order of steps (1) to (2) and step (3).

[0025] The preparation process of the carbon-silicon carbide honeycomb sandwich structure material in this invention is as follows: Figure 1As shown. In this invention, carbon cloth is woven into planar fiber layers, and the interlayer fibers are sewn into adjacent carbon cloth layers to obtain a planar preform with semi-closed honeycomb cells. Then, a core mold is inserted into the planar preform and locked using a shape control fixture to obtain a honeycomb core preform.

[0026] In this invention, the raw material of the carbon cloth preferably includes one or more of T300 fiber, T700 fiber, T80 fiber, M40 fiber, and M55 fiber, and more preferably T300-3K plain weave fabric; the areal density of the raw material of the carbon cloth is preferably 200 g / m³. 3 The preferred weaving angle of the carbon cloth is +45° / -45°.

[0027] In this invention, the warp and weft density of the planar fiber layer is preferably 6 threads / cm × 6 threads / cm or 7 threads / cm × 7 threads / cm.

[0028] In this invention, the interlayer fibers preferably include one or more of T300 fibers, T700 fibers, T80 fibers, M40 fibers, and M55 fibers.

[0029] In this invention, the core mold is preferably a hexagonal core mold.

[0030] In this invention, the weaving process strictly controls the yarn spacing according to the core mold size and process parameters. After obtaining a planar preform with semi-closed honeycomb cells, the core mold is implanted to form a uniformly arranged hexagonal honeycomb. At the same time, shape control fixtures are used around the perimeter to lock the overall size of the honeycomb, thereby ensuring the dimensional stability of the honeycomb core preform.

[0031] After obtaining the honeycomb core preform, the present invention sequentially impregnates the honeycomb core preform with resin, cures it, and deposits pyrolytic carbon (denoted as the first pyrolytic carbon deposition) to obtain the honeycomb core. In the present invention, the impregnation solution used for resin impregnation is preferably a phenolic resin solution; the solvent of the impregnation solution is preferably ethanol; the concentration of the impregnation solution is preferably 10~40wt%, more preferably 15~25wt%.

[0032] In this invention, the resin impregnation equipment preferably includes a pressure vessel; the resin impregnation is preferably carried out in a protective atmosphere; the protective atmosphere is preferably argon; the resin impregnation pressure is preferably 0.3~0.5MPa, more preferably 0.4MPa, and the pressure holding time is preferably 4~5 hours; the resin impregnation temperature is preferably room temperature. After resin impregnation is completed, the product is removed and prepared for subsequent steps.

[0033] In this invention, the curing equipment preferably includes an oven; the curing conditions preferably include: heating from room temperature to 100 degrees Celsius at a rate of 5°C / min, holding at 100 degrees Celsius for 2 hours, then heating from 100 degrees Celsius to 120 degrees Celsius at a rate of 1°C / min, holding at 120 degrees Celsius for 2 hours, then heating from 120 degrees Celsius to 160 degrees Celsius at a rate of 2°C / min, holding at 160 degrees Celsius for 2 hours, followed by oven cooling. This invention, through gradient heating, can avoid resin foaming, control uniform resin diffusion, and ensure complete resin curing. Through the above operations, this invention shapes the honeycomb core preform.

[0034] In this invention, the shaping tooling and core mold are preferably removed after curing.

[0035] In this invention, the temperature of the first deposited pyrolytic carbon is preferably 900-1100 degrees Celsius, more preferably 1000 degrees Celsius, the holding time is preferably 40-60 hours, more preferably 50 hours, the thickness of the pyrolytic carbon is preferably 100-500 nanometers, more preferably 300-400 nanometers; the carbon source gas for the first deposited pyrolytic carbon is preferably propylene, the total system pressure is preferably 3-5 kPa, and the carrier gas is preferably argon.

[0036] The present invention involves depositing pyrolytic carbon (referred to as the second pyrolytic carbon deposition) after shaping the panel preform to obtain the panel.

[0037] In this invention, the weaving method of the panel preform is preferably 2.5D weaving, 3D weaving or needle punching; the raw material of the panel preform preferably includes one or more of T300 fiber, T700 fiber, T80 fiber, M40 fiber and M55 fiber, more preferably T300-3K plain weave fabric.

[0038] In this invention, the shaping is preferably performed using a graphite mold clamping method.

[0039] In this invention, the temperature of the second deposited pyrolytic carbon is preferably 900-1100 degrees Celsius, more preferably 1000 degrees Celsius, the holding time is preferably 30-60 hours, more preferably 40-50 hours, and the thickness of the pyrolytic carbon is preferably 100-500 nanometers, more preferably 300-400 nanometers; the carbon source gas for the first deposited pyrolytic carbon is preferably propylene, the total system pressure is preferably 3-5 kPa, and the carrier gas is preferably argon.

[0040] In this invention, the process of pre-depositing pyrolytic carbon preferably includes heat treatment of the panel preform; the heat treatment temperature is preferably 2000 degrees Celsius, and the holding time is preferably 2 hours. Through the above heat treatment, this invention removes moisture from the panel preform and improves the graphitization degree of the carbon fibers in the panel preform.

[0041] In this invention, the second deposition of pyrolytic carbon preferably includes surface treatment of the resulting product; the surface treatment is preferably grinding or milling. This invention, through surface treatment, ensures better integration with the honeycomb core.

[0042] After obtaining the honeycomb core and the panel, the present invention sequentially impregnates the honeycomb core and the two panels with polycarbosilane, co-cures and heats the material, repeating the polycarbosilane impregnation, co-curing and heat treatment process 8 to 12 times to obtain the carbon-silicon carbide honeycomb sandwich structure material.

[0043] In this invention, the parameters for the polycarbosilane impregnation preferably include: a vacuum degree of 0.1~1 Pa, more preferably 0.5~0.8 Pa, and a pressure holding time of 45 min.

[0044] In this invention, the impregnation solution used for polycarbosilane impregnation is preferably a polycarbosilane precursor solution; the viscosity of the impregnation solution is preferably 100 mPa. The concentration of the polycarbosilane precursor solution is preferably 30 wt%; the organic solvent preferably includes polycarbosilane and an organic solvent; the organic solvent is preferably xylene; the molar mass of the polycarbosilane is preferably 1000~1500 g / mol; the polycarbosilane preferably includes one or more of vinyl polycarbosilane and low viscosity / liquid polycarbosilane.

[0045] In this invention, the co-curing process preferably includes bringing the honeycomb core and the two panels into close contact (integration) in the order of panels, honeycomb core and panels before the co-curing process; the close contact preferably uses a metal mold; the pressure of the close contact is preferably 10 MPa.

[0046] In this invention, the co-curing molding is preferably carried out in an autoclave; the co-curing molding preferably includes a first curing stage and a second curing stage performed sequentially; the temperature of the first curing stage is preferably 120 degrees Celsius, the pressure is preferably 0.3~0.5 MPa, and the holding time is preferably 180 minutes; the temperature of the second curing stage is preferably 150 degrees Celsius, the pressure is preferably 0.3~0.5 MPa, and the holding time is preferably 180 minutes. After co-curing molding, the product is removed and prepared for subsequent steps.

[0047] In this invention, the pyrolysis is preferably carried out in a sintering furnace; the pyrolysis is preferably carried out in a protective atmosphere; the protective atmosphere is preferably nitrogen; the pyrolysis temperature is preferably 1100~1300 degrees Celsius, more preferably 1200 degrees Celsius, and the holding time is preferably 2 hours.

[0048] In this invention, after each pyrolysis, it is preferable to further subject the resulting product to heat treatment; the temperature of the heat treatment is preferably 1800~2000 degrees Celsius, more preferably 1900 degrees Celsius, and the holding time is preferably 2 hours. This invention alleviates residual thermal stress through heat treatment.

[0049] In this invention, the polycarbosilane impregnation, co-curing, and pyrolysis are repeated 8 to 12 times, preferably 10 times. By repeating the above steps, this invention improves the density of the material.

[0050] The present invention also provides the application of the carbon-silicon carbide honeycomb sandwich structure material described in the above-described scheme or the carbon-silicon carbide honeycomb sandwich structure material obtained by the preparation method described in the above-described scheme in the field of aircraft.

[0051] The carbon-silicon carbide honeycomb sandwich structure material provided by this invention meets multiple objectives such as lightweight, high load-bearing capacity and high temperature resistance, and can be widely used in hypersonic thermal protection structures, such as wing rudders, nose cones or throat liners.

[0052] To further illustrate the present invention, the technical solutions of the present invention will be clearly and completely described below with reference to the accompanying drawings and embodiments.

[0053] Example 1 This embodiment prepares a carbon-silicon carbide honeycomb sandwich structure material with dimensions of 300mm×300mm×43mm, panel thickness of 1.5mm, honeycomb height of 36.5mm, honeycomb core cell side length of 5mm, and honeycomb core cell wall thickness of 0.2mm. The porosity of the honeycomb core is less than 5%, and the compressive strength is ≥8MPa. The specific steps are as follows: (1) such as Figures 2-3 As shown, after being cut to a specific size from T300-3K plain weave fabric, the areal density is 200g / m². 3 The fabric, with a warp and weft density of 6 threads / cm x 6 threads / cm, is cut into carbon fiber fabric with a weaving angle of +45° / -45° according to size requirements. Figure 2 The layered stitching shown yields a planar prefabricated structure with semi-closed honeycomb cells. The honeycomb wall fibers are oriented at ±45°. Then, according to... Figure 3 The core mold is inserted into the semi-closed honeycomb cells to form a uniformly arranged hexagonal honeycomb. At the same time, the overall honeycomb size is locked in place by the shape control fixture around the perimeter.

[0054] (2) The hexagonal honeycomb tooling in step (1) is impregnated with phenolic resin with an ethanol content of 25wt%. The curing process is as follows: the temperature is increased from room temperature to 100 degrees Celsius at 5℃ / min and kept at that temperature for 2 hours; then the temperature is increased to 120 degrees Celsius at 1℃ / min and kept at that temperature for 2 hours; then the temperature is increased to 160 degrees Celsius at 2℃ / min and kept at that temperature for 2 hours. Finally, the temperature is cooled with the furnace. After cooling to room temperature, the tooling is removed and the core mold is knocked off to obtain the resin-based honeycomb core.

[0055] (3) Insert hexagonal graphite fixtures around the resin-based honeycomb core in step (2), deposit pyrolytic carbon in a deposition furnace, the deposition temperature is 1000 degrees Celsius, the carbon source gas is propylene, the total system pressure is 3~5 kPa, the carrier gas is argon, the deposition time is 50 hours, and the thickness of the pyrolytic carbon interface is 300 nanometers.

[0056] (4) Select T300-3K plain weave fabric, cut it to the predetermined size, and then weave it into a panel preform using the 2.5D weaving method. The thickness of the carbon fiber preform is 1.5mm. Then, the preform is subjected to high-temperature heat treatment at a temperature of 2000 degrees Celsius for 2 hours.

[0057] (5) Place the carbon fiber fabric from step (4) into a deposition furnace to deposit a pyrolytic carbon matrix. The deposition temperature is 1000 degrees Celsius, the carbon source gas is propylene, the total system pressure is 3~5 kPa, the carrier gas is argon, the deposition time is 50 hours, and the thickness of the pyrolytic carbon interface is 300 nanometers.

[0058] (6) Place the panels and honeycomb cores from steps (5) and (3) into vacuum bags respectively, vacuum impregnate them with the polycarbosilane precursor solution, remove them, and assemble the honeycomb core and the upper and lower panels together. Use a metal mold to ensure tight contact between the honeycomb core and the upper and lower panels, such as... Figure 4 As shown. The concentration of the polycarbosilane precursor solution used was 30 wt%, and the molecular weight of the polycarbosilane was 1000~1500 g / mol. It was then placed in an autoclave and cured at 120°C for 180 minutes under a pressure of 0.3 MPa, followed by curing at 150°C for another 180 minutes. The cured sandwich structure was then removed and pyrolyzed in a sintering furnace at 1200°C for 2 hours under a N2 atmosphere. After each round of pyrolysis, it was heat-treated at 1800°C for 2 hours. This step was repeated 10 times to obtain the carbon-silicon carbide honeycomb sandwich structure material.

[0059] Example 2 This embodiment prepares a carbon-silicon carbide honeycomb sandwich structure material with dimensions of 500mm×500mm×50mm, panel thickness of 3mm, honeycomb height of 44mm, honeycomb core cell side length of 9mm, and honeycomb core cell wall thickness of 0.4mm. The specific steps are as follows: (1) such as Figures 2-3As shown, T700 plain weave fabric with an areal density of 200 g / m² is selected. 3 The fabric, with a warp and weft density of 7 threads / cm × 7 threads / cm, is cut into carbon fiber fabric with a weaving angle of 0 / 90° according to size requirements. Figure 2 The layers are stitched together in pairs, with a stitch spacing of 6 mm in the height direction and 9 mm in the L direction. The stitches are made of T300-3K carbon fiber, resulting in a planar prefabricated structure with a semi-closed honeycomb structure. The honeycomb wall fiber orientation is 0 / 90°. Then, according to... Figure 3 The core mold is inserted into the semi-closed honeycomb cells to form a uniformly arranged hexagonal honeycomb. At the same time, the overall honeycomb size is locked in place by the shape control fixture around the perimeter.

[0060] (2) The hexagonal honeycomb tooling in step (1) is impregnated with a phenolic resin ethanol solution with a concentration of 40wt%. The curing process is as follows: the temperature is raised from room temperature to 100 degrees Celsius at a rate of 5℃ / min and held for 2 hours. Then the temperature is raised to 120 degrees Celsius at a rate of 1℃ / min and held for 2 hours. Then the temperature is raised to 160 degrees Celsius at a rate of 2℃ / min and held for 2 hours. Finally, the temperature is cooled with the furnace. After cooling to room temperature, the tooling is removed and the core mold is knocked off to obtain the resin-based honeycomb core.

[0061] (3) Insert hexagonal graphite fixtures around the resin-based honeycomb core in step (2), deposit pyrolytic carbon in a deposition furnace, the deposition temperature is 900 degrees Celsius, the carbon source gas is propylene, the total system pressure is 3~5 kPa, the carrier gas is argon, the deposition time is 40 hours, and the thickness of the pyrolytic carbon interface is 200 nanometers.

[0062] (4) Select T700 plain weave fabric, cut it to the predetermined size, and then use needle punching process to connect it into carbon fiber fabric to obtain panel preform. The thickness of the panel preform is 3mm. Then, the panel preform is subjected to high temperature heat treatment at a temperature of 2000 degrees Celsius for 2 hours.

[0063] (5) The carbon fiber fabric from step (4) is placed in a deposition furnace to deposit pyrolytic carbon. The deposition temperature is 900 degrees Celsius, the carbon source gas is propylene, the total system pressure is 3~5 kPa, the carrier gas is argon, the deposition time is 30 hours, and the thickness of the pyrolytic carbon interface is 200 nanometers.

[0064] (6) Place the panels and honeycomb cores from steps (5) and (3) into vacuum bags respectively, vacuum impregnate them with the polycarbosilane precursor solution, remove them, and assemble the honeycomb core and the upper and lower panels together. Use a metal mold to ensure tight contact between the honeycomb core and the upper and lower panels, such as... Figure 4As shown. The concentration of the polycarbosilane precursor solution used was 30 wt%, and the molecular weight was 1000~1500 g / mol. It was then placed in an autoclave and cured at 120°C for 180 minutes under a pressure of 0.3 MPa, followed by curing at 150°C for another 180 minutes. The cured sandwich structure was then removed and pyrolyzed in a sintering furnace at 1200°C for 2 hours under a N2 atmosphere. After each round of pyrolysis, it was heat-treated at 2000°C for 2 hours. This process was repeated 8 times to obtain the carbon-silicon carbide honeycomb sandwich structure material.

[0065] Test Example 1 The out-of-plane compressive properties of the carbon-silicon carbide honeycomb sandwich structure material prepared in Example 1 were tested at 1200 degrees Celsius. In this test example, three samples were prepared under the same conditions according to the preparation method of Example 1. The test results are as follows: Figure 5 As shown. According to Figure 5 It can be seen that the strength values ​​of the three samples are all greater than 10 MPa, and they show good consistency.

[0066] The embodiments of the present invention have been described above; however, these embodiments are for illustrative purposes only and are not intended to limit the scope of the invention. All other embodiments obtained by those skilled in the art based on the above embodiments of the present invention without inventive effort are within the protection scope of the present invention.

Claims

1. A carbon-silicon carbide honeycomb sandwich structure material, characterized in that, Includes a matrix and a composite layer covering the surface of the matrix; The substrate includes a honeycomb core and a panel covering the upper and lower surfaces of the honeycomb core; The substrate is made of carbon fiber woven material; The composite layer comprises carbon and silicon carbide.

2. The carbon-silicon carbide honeycomb sandwich structure material according to claim 1, characterized in that, The parameters of the honeycomb core include: the side length of the honeycomb core is 3~16mm, the wall thickness of the honeycomb core is 0.2~0.4mm, and the honeycomb height is 10~44mm.

3. The carbon-silicon carbide honeycomb sandwich structure material according to claim 1 or 2, characterized in that, The porosity of the honeycomb core is less than 5%, and the compressive strength is not less than 8 MPa.

4. The carbon-silicon carbide honeycomb sandwich structure material according to claim 1, characterized in that, The thickness of the panel is 1.5~3mm.

5. The method for preparing the carbon-silicon carbide honeycomb sandwich structure material according to any one of claims 1 to 4, characterized in that, Includes the following steps: (1) The carbon cloth is layered and woven into a planar fiber layer. The interlayer fibers are sewn into the two adjacent carbon cloth layers to obtain a planar preform with a semi-closed honeycomb cell. Then, a core mold is inserted into the planar preform and locked using a shape control fixture to obtain a honeycomb core preform. (2) The honeycomb core preform is sequentially impregnated with resin, cured and deposited with pyrolytic carbon to obtain a honeycomb core; (3) After shaping the panel preform, pyrolytic carbon is deposited to obtain the panel; (4) The honeycomb core and the two panels are impregnated with polycarbosilane, co-cured and pyrolyzed in sequence, and the polycarbosilane impregnation, co-curing and pyrolysis are repeated 8 to 12 times to obtain the carbon-silicon carbide honeycomb sandwich structure material. There is no requirement for the time order of steps (1) to (2) and step (3).

6. The preparation method according to claim 5, characterized in that, The raw materials for the carbon cloth include one or more of T300 fiber, T700 fiber, T80 fiber, M40 fiber and M55 fiber. The areal density of the raw material for the carbon cloth is 200 g / m³. 3 ; The weaving angle of the carbon cloth is +45° / -45°; The warp and weft density of the planar fiber layer is 6 fibers / cm × 6 fibers / cm or 7 fibers / cm × 7 fibers / cm.

7. The preparation method according to claim 5, characterized in that, The curing conditions include: increasing the temperature from room temperature to 100 degrees Celsius at a rate of 5°C / min, holding at 100 degrees Celsius for 2 hours, increasing the temperature from 100 degrees Celsius to 120 degrees Celsius at a rate of 1°C / min, holding at 120 degrees Celsius for 2 hours, increasing the temperature from 120 degrees Celsius to 160 degrees Celsius at a rate of 2°C / min, holding at 160 degrees Celsius for 2 hours, and then cooling in the furnace.

8. The preparation method according to claim 5, characterized in that, The co-curing molding includes a first curing stage and a second curing stage performed sequentially. The temperature of the first curing stage is 120 degrees Celsius, the pressure is 0.3~0.5 MPa, and the holding time is 180 minutes; The temperature of the second curing stage is 150 degrees Celsius, the pressure is 0.3~0.5 MPa, and the holding time is 180 minutes.

9. The preparation method according to claim 5, characterized in that, The pyrolysis is carried out in a protective atmosphere; The pyrolysis temperature is 1100~1300 degrees Celsius, and the holding time is 2 hours; Each pyrolysis is followed by heat treatment of the resulting product; The heat treatment temperature is 1800~2000 degrees Celsius, and the holding time is 2 hours.

10. The application of the carbon-silicon carbide honeycomb sandwich structure material according to any one of claims 1 to 4 or the carbon-silicon carbide honeycomb sandwich structure material obtained by the preparation method according to any one of claims 5 to 9 in the field of aircraft.

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