A high-temperature resistant lightweight composite material, its preparation method and application

Abstract

A high-temperature resistant lightweight composite material, its preparation method, and its applications are disclosed, relating to the field of composite material technology. This invention solves the problems of existing materials where it is difficult to simultaneously achieve both high temperature resistance and lightweighting, low structural efficiency due to the separation of heat protection and load-bearing design, insufficient temperature resistance and toughness of the sandwich layer, and difficulties and high costs associated with large-size molding. The material comprises an integrated co-cured outer heat-resistant layer, an intermediate transition layer, and an inner load-bearing layer. The outer heat-resistant layer uses a modified epoxy / cyanate copolymer resin as the matrix and a high-silica glass fiber and carbon fiber blended fabric as reinforcement, with fumed silica and carbon nanotube fillers incorporated into the matrix resin. The intermediate transition layer is polymethacrylamide foam. The inner load-bearing layer uses an epoxy-modified cyanate ester resin as the matrix and a carbon fiber biaxial plain weave fabric as reinforcement. This invention can be used in rocket fairings, hypersonic vehicle skins, spacecraft reentry capsule thermal protection structures, hypersonic missile shells, aircraft engine nacelles, or high-speed train front fairings.

Description

Technical Field

[0001] This invention relates to the field of composite material technology, specifically to a high-temperature resistant lightweight composite material, its preparation method, and its application. Background Technology

[0002] The rocket fairing is a key structural component of a launch vehicle, located at the rocket's front end. Its main function is to protect satellites, spacecraft, and other payloads from damage caused by aerodynamic heating, aerodynamic loads, and acoustic vibration during atmospheric flight. After the rocket exits the atmosphere, the fairing must reliably separate. As space missions continue to evolve towards deep space exploration, hypersonic flight, and reusability, the service environment faced by fairings is becoming increasingly harsh. Especially during high-speed flight, the outer surface of the fairing must withstand severe aerodynamic heating, with local temperatures reaching over 350°C or even higher. At the same time, to improve the rocket's carrying capacity, the fairing structure must achieve a lightweight design. To ensure the safety of the internal payload, the materials used must also possess high specific stiffness, high specific strength, and good impact toughness.

[0003] Currently, traditional rocket fairings mostly use aluminum alloy honeycomb sandwich structures or carbon fiber / epoxy resin composite materials. Among them, aluminum alloy structures have the problem of high density and limited heat resistance limit. Their long-term heat resistance temperature is only about 150~200℃, requiring the use of an additional thick heat protection layer. While conventional carbon fiber / epoxy resin composite materials have excellent specific strength, the temperature resistance of the matrix resin is poor. The long-term operating temperature is generally no more than 150℃. At higher temperature environments, the mechanical properties will drop sharply, making it difficult to meet the high heat load requirements of high Mach number flight or reusable conditions.

[0004] Through research and analysis of existing technologies, it is evident that composite materials used in rocket fairings still face numerous unresolved technical challenges. Achieving both high temperature resistance and lightweight properties is difficult, failing to meet the demands of high Mach number flights. Current composite fairings primarily utilize epoxy resin-based carbon fiber panels, with a long-term operating temperature of only approximately 120-150°C. When flight Mach numbers increase or during repeated use, aerodynamic heating can cause the panel temperature to exceed 180°C, leading to thermo-oxidative aging of the epoxy resin and a sharp decline in its mechanical properties, posing a risk of structural failure. Even though some high-temperature resistant resins, such as polyimide and bismaleimide, can withstand temperatures of 250-350°C, their high density, typically exceeding 1.4 g / cm³, presents significant challenges. 3Furthermore, insufficient interfacial bonding after composite with carbon fiber leads to low interlaminar shear strength, limiting its application in main load-bearing structures. Existing heat protection / load-bearing separation designs result in low structural efficiency. Current technologies generally employ a two-layer structure of "external heat protection layer + composite load-bearing layer." The heat protection layer often uses low-density ablative materials or cork, physically bonded or glued to the composite panel. The significant difference in their thermal expansion coefficients makes them prone to interfacial debonding, warping, and even peeling under combined thermal stress, severely impacting the structural reliability. The temperature resistance and toughness of the sandwich layer are also difficult to balance. While PMI foam core material commonly used in fairings possesses lightweight and high rigidity, the long-term operating temperature of conventional PMI foam does not exceed 180℃, and it is prone to core material crushing under impact loads. While aramid paper honeycomb core material has slightly better temperature resistance, it suffers from inherent defects such as moisture absorption and weak interfacial bonding with the panel. In addition, existing high-temperature resistant composite material systems mostly require high-temperature and high-pressure curing and molding, which places high demands on production equipment and consumes a lot of energy. It is difficult to manufacture large-size, high-curvature fairing integral components, and the manufacturing cost is high. The molding of large-size complex components is also quite difficult.

[0005] Therefore, it is necessary to develop a material that has a long-term operating temperature of ≥220℃, a short-term temperature resistance of ≥350℃ / 300S, and a density of ≤0.95g / cm³. 3 The high-temperature resistance and lightweight properties of new fairing composite materials, along with high mechanical properties and good process adaptability, have become an important research direction in the field of aerospace materials. Summary of the Invention

[0006] To address the complex technical challenges of existing rocket fairing materials, such as the difficulty in balancing temperature resistance and lightweight design, low structural efficiency due to the separation of heat protection and load-bearing functions, insufficient temperature resistance and toughness of the sandwich layer, and difficulties and high costs in molding large-sized components, this invention proposes a high-temperature resistant lightweight composite material, its preparation method, and its applications.

[0007] The specific technical solution of the present invention is as follows: A high-temperature resistant lightweight composite material includes an outer heat-insulating layer, an intermediate transition layer, and an inner load-bearing layer, wherein the three-layer structure is an integrated co-curing molding structure; The outer heat-insulating layer uses modified epoxy / cyanate copolymer resin as the matrix resin and high-silica glass fiber and carbon fiber blended fabric as the reinforcement. The matrix resin is also compounded with fumed silica and carbon nanotube synergistic reinforcing fillers. The intermediate transition layer is polymethacrylimide foam; The inner bearing layer uses epoxy-modified cyanate ester resin as the matrix resin and T700 grade and / or T800 grade carbon fiber biaxial plain weave fabric as the reinforcement. The outer heat-insulating layer has a total thickness of 0.5~1.5mm and has 2~6 layers; the intermediate transition layer has a thickness of 5~20mm; and the inner load-bearing layer has a thickness of 0.8~2.0mm and has 4~8 layers.

[0008] Preferably, the polymethacrylimide foam adopts a density gradient distribution with a denser outer layer and a looser inner layer.

[0009] Preferably, the modified epoxy / cyanate copolymer resin is obtained by copolymerizing epoxy resin and cyanate resin at a mass ratio of 30:70 to 50:50; the modified epoxy / cyanate copolymer resin contains 5 to 15 wt% toughening agent, which is carboxyl-terminated nitrile rubber or core-shell rubber.

[0010] Preferably, the epoxy resin is a bisphenol A type epoxy resin or a bisphenol F type epoxy resin.

[0011] Preferably, the mass ratio of high silica glass fiber to carbon fiber in the high silica glass fiber and carbon fiber blended fabric is 50:50 to 70:30; The amount of fumed silica added in the outer heat-insulating layer is 2-8 wt% of the total mass of the matrix resin, and the amount of carbon nanotubes added is 0.5-2 wt% of the total mass of the matrix resin.

[0012] Preferably, the blended fabric has a plain or twill weave structure and an areal density of 200~400 g / m². 2 The carbon fiber is T700 or T800 grade carbon fiber, and the high silica glass fiber has a SiO2 content of ≥96%.

[0013] Preferably, the surface of the polymethacrylimide foam is treated with plasma or coated with a coupling agent primer.

[0014] Preferably, the epoxy-modified cyanate ester resin is obtained by blending cyanate ester resin and epoxy resin at a mass ratio of 60:40 to 80:20; the epoxy-modified cyanate ester resin contains 10 to 20 wt% of a thermoplastic toughening agent, wherein the thermoplastic toughening agent is polyethersulfone or polyetherimide.

[0015] Preferably, the carbon fiber fabric of the inner supporting layer is a plain weave fabric or an 8-end satin weave fabric made of T700SC-12K or T800SC-12K carbon fiber, with an areal density of 200~400g / m². 2 The fiber volume content is 55%~65%.

[0016] This invention also provides a method for preparing a high-temperature resistant lightweight composite material, which employs a vacuum-assisted resin transfer molding combined with a co-curing process, comprising the following steps: S1. Coat the mold surface with a release agent, and lay out the dry fibers of the outer heat-insulating layer, the high-temperature resistant polymethacrylimide foam core material, and the dry fibers of the inner load-bearing layer in the design order, lay out the flow guiding system and seal the vacuum system. S2. After vacuuming and pressure testing the vacuum system, a two-component injection system is used to inject modified epoxy / cyanate ester copolymer resin into the dry fibers of the outer heat-insulating layer and epoxy modified cyanate ester resin into the dry fibers of the inner load-bearing layer to complete fiber impregnation. S3. After the glue injection is completed, close the glue injection port and maintain vacuum. Apply a pressure of 0.2~0.4MPa and use a stepped temperature increase regime for co-curing. After curing, cool down and demold. S4. Trim, machine and non-destructive test the demolded components to obtain the finished product.

[0017] Preferably, in step S1, the vacuum degree of the vacuum system is ≤100Pa; the polymethacrylimide foam core material is pre-processed according to the mold shape, and the foam surface is subjected to plasma treatment or coated with coupling agent primer before laying.

[0018] Preferably, the modified epoxy / cyanate ester copolymer resin solution is prepared as follows: epoxy resin and cyanate ester resin are added to a reaction vessel in proportion, heated to 80~100℃ and stirred for 30 min; fumed silica is added and dispersed at high speed of 2000~3000 rpm for 30 min; pre-dispersed carbon nanotubes are added and dispersion is continued for 30 min; the temperature is lowered to below 60℃, curing agent and toughening agent are added, stirred evenly and then vacuum degassed.

[0019] Preferably, the preparation method of the epoxy modified cyanate ester resin solution is as follows: mix cyanate ester resin and epoxy resin in proportion, heat to 100~120℃ and stir until transparent; add thermoplastic toughening agent and stir to dissolve for 2~3 hours; cool to 80℃, add curing agent, stir evenly and then vacuum degas.

[0020] Preferably, the injection temperature in step S2 is 60~80℃.

[0021] The stepped temperature curing process in step S3 is as follows: first, heat to 120℃ and hold for 2 hours, then heat to 150℃ and hold for 2 hours, then heat to 180℃ and hold for 4 hours, and finally heat to 200℃ and hold for 2 hours for post-curing; after curing, cool down to below 60℃ to demold.

[0022] The present invention also provides an application of the above-mentioned high-temperature resistant lightweight composite material in rocket fairings, hypersonic vehicle skins, spacecraft reentry capsule thermal protection structures, hypersonic missile shells, aero-engine nacelles, or high-speed train front fairings.

[0023] Compared with the prior art, the specific beneficial effects of the present invention are as follows: 1. This invention uses modified epoxy / cyanate copolymer resin as the outer heat shield matrix and high-toughness epoxy modified cyanate resin as the inner load-bearing layer matrix. Combined with the high heat resistance of the cyanate resin itself and the synergistic effect of epoxy toughening, the long-term service temperature of the outer heat shield is not lower than 220℃, and it can withstand aerodynamic heating of more than 350℃ / 300s for a short period of time. The glass transition temperature of the inner structural layer is not lower than 180℃. This invention significantly improves the problem of high-temperature thermo-oxidative aging and mechanical property decay of the resin, and can stably adapt to the harsh thermal load environment of high Mach number flight and reusable rockets.

[0024] 2. This invention designs the outer heat-insulating layer as a hybrid reinforced structure of high-silica glass fiber and carbon fiber. The high-silica glass fiber provides ablation resistance and heat insulation, while the carbon fiber bears the mechanical load, enabling the single-layer structure to have both heat-insulating and load-bearing capabilities, eliminating the traditional separate design of the outer heat-insulating layer and the load-bearing layer. The three-layer structure adopts an integrated co-curing molding process, eliminating the adhesive interface between the heat-insulating layer and the load-bearing layer, avoiding interface debonding, warping, and peeling failure caused by differences in thermal expansion coefficients, and significantly improving the overall structural integrity and service safety.

[0025] 3. This invention uses low-density, high-temperature resistant PMI foam as the intermediate layer, combined with a high-specific-strength carbon fiber reinforcement system, so that the overall density of the composite material is no higher than 0.95 g / cm³, which reduces weight by 30%~50% compared with traditional aluminum alloy honeycomb sandwich structures and by 15%~25% compared with conventional carbon fiber / epoxy resin sandwich structures. The inner load-bearing layer is made of T700 / T800 grade carbon fiber and high-toughness resin composite, with a tensile strength of no less than 800 MPa, an interlaminar shear strength of no less than 25 MPa, and an impact compressive strength of no less than 200 MPa. While achieving lightweighting, it ensures high specific stiffness, high specific strength and excellent impact toughness, directly improving the effective launch efficiency of rockets.

[0026] 4. This invention selects high-temperature resistant PMI foam with a long-term operating temperature of not less than 200℃ as the core material. The temperature resistance level matches the resin system of the panel, avoiding the softening and instability of the core material at high temperatures. PMI foam itself has excellent energy absorption characteristics. Combined with the toughening of the outer layer of nanofiller and the toughening of the inner layer of thermoplastic resin, it can effectively inhibit crack propagation and resist aerodynamic loads and acoustic vibration impacts. It solves the defects of conventional PMI foam such as insufficient temperature resistance, easy moisture absorption of aramid paper honeycomb and weak interfacial bonding, and improves the stability and integrity of the structure in complex mechanical environments.

[0027] 5. This invention employs a vacuum-assisted resin transfer molding combined with co-curing process, which can complete fiber impregnation and curing under medium temperature and pressure conditions, reducing reliance on ultra-large autoclaves and enabling one-time integral molding of large-diameter, high-curvature fairing components with diameters of over 5 meters. The integrated co-curing process eliminates secondary bonding and multiple machining processes, reducing material costs by approximately 30% compared to traditional polyimide-based composite materials, lowering overall manufacturing costs by 35% to 45%, and reducing production energy consumption by over 50%, significantly improving process adaptability and economic efficiency.

[0028] 6. The composite material of this invention has comprehensive advantages such as high temperature resistance, low density, high specific strength, high toughness, ablation resistance, and low thermal conductivity. It is not only suitable for rocket fairings, but can also be extended to hypersonic vehicle skins and control surfaces, spacecraft reentry capsule thermal protection structures, hypersonic missile shells, aero-engine nacelles, and high-speed train front fairings, etc., which have stringent requirements for heat resistance and lightweighting. It has good versatility and engineering application prospects. Detailed Implementation

[0029] To make the technical solutions of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below. It should be noted that the following embodiments are only used to better understand the technical solutions of the present invention and should not be construed as limiting the present invention.

[0030] Example 1. This embodiment provides a high-temperature resistant lightweight composite material, which is a three-layer gradient integrated co-curing structure. Its raw material selection, layup design, and preparation process are as follows: 1. Raw material selection: Outer heat-insulating layer: The matrix resin is obtained by copolymerizing bisphenol A type epoxy resin (E-51) and cyanate ester resin (CE-01) at a mass ratio of 40:60. Simultaneously, 10 wt% of carboxyl-terminated liquid nitrile rubber (CTBN) is added as a toughening agent. The curing agent is diaminodiphenylmethane (DDM), used at 12% of the total resin mass. The reinforcing filler consists of 5 wt% fumed silica and 1 wt% multi-walled carbon nanotubes (10-20 nm in diameter) at the total resin mass. The reinforcement is a plain weave fabric made of high-silica glass fiber and T300 carbon fiber woven at a mass ratio of 60:40, with an areal density of 300 g / m². 2 .

[0031] Intermediate transition layer: Uses Rohacell 71 HT high-temperature resistant polymethacrylamide (PMI) foam with a density of 75 kg / m³. 3 The thickness is 12mm.

[0032] Inner Support Layer: The matrix resin is obtained by blending cyanate ester resin (CE-01) and phenolic epoxy resin (F-51) at a mass ratio of 70:30, with 15 wt% polyethersulfone (PES) added as a thermoplastic toughening agent. Dicyandiamide is used as the latent curing agent, at a dosage of 5 wt% of the total resin mass. The reinforcement is made of T700SC-12K carbon fiber plain weave fabric with an areal density of 320 g / m². 2 .

[0033] 2. Layer design: The composite material layup in this embodiment, from the outside to the inside, is as follows: The outer heat-insulating layer is made of four layers of the above-mentioned mixed woven fabric composite modified epoxy / cyanate copolymer resin, with a total thickness of 1.0 mm; The intermediate transition layer consists of one layer of the aforementioned PMI foam, with a thickness of 12mm; The inner load-bearing layer is made of 6 layers of the above-mentioned carbon fiber cloth composite epoxy modified cyanate resin, with a total thickness of 1.5mm.

[0034] 3. The preparation process is as follows: This embodiment uses a vacuum-assisted resin transfer molding (VARTM) combined with an oven co-curing process. The specific steps are as follows: First, prepare the mold and lay up the layers. Use a steel female mold and evenly coat the mold surface with a high-temperature release agent. Lay out the outer heat-insulating layer of dry fiber preform, the pre-formed PMI foam core material, and the inner load-bearing layer of dry fiber preform in the design order. Then, lay out the flow guide net and release cloth, seal the vacuum bag and build a complete vacuum system, control the vacuum degree to ≤100Pa, and perform pressure holding and leak detection.

[0035] Simultaneous preparation of two-component resin adhesives: Preparation method of the outer heat-insulating layer resin adhesive: Add epoxy resin and cyanate ester resin to the reactor in proportion, heat to 90℃ and stir for 30 minutes, add fumed silica and disperse at high speed of 2500 rpm for 30 minutes until uniformly dispersed, add carbon nanotubes that have been pre-dispersed in a small amount of epoxy resin using a three-roll mill, continue to disperse for 30 minutes, cool to below 60℃ and add curing agent and toughening agent, stir evenly and then vacuum degas for later use; Preparation method of the inner bearing layer resin adhesive: Mix cyanate ester resin and epoxy resin in proportion, heat to 110℃ and stir until the system is transparent, add thermoplastic toughening agent polyethersulfone, continue stirring for 2.5 hours until completely dissolved, cool to 80℃ and add latent curing agent, stir evenly and then vacuum degas for later use.

[0036] During the resin injection stage, the resin liquid of the outer heat-insulating layer, preheated to 80°C, is injected from the outside of the mold, while the resin liquid of the inner bearing layer, preheated to 80°C, is injected from the inside of the mold. The injection speed is controlled to avoid dry spots. The total injection time is about 2 hours, which completes the fiber impregnation.

[0037] After the adhesive is injected, the injection port is closed, and a vacuum state is maintained. The mixture is then transferred to an oven with a pressurization device, where an external pressure of 0.3 MPa is applied. A stepped temperature increase regime is used for co-curing: first, the temperature is increased to 120°C and held for 2 hours, then increased to 150°C and held for 2 hours, then increased to 180°C and held for 4 hours, and finally increased to 200°C and held for 2 hours to complete the curing process. After the curing process is completed, the mixture is allowed to cool naturally to below 60°C before demolding. The demolded components are then trimmed and subjected to non-destructive testing to obtain the finished composite material.

[0038] Example 2. This embodiment provides an optimized high-temperature resistant lightweight composite material. Its core structure and preparation process are the same as those in Example 1, with the only differences being the optimization and adjustment of raw material selection, layup design, and core material pretreatment process, as detailed below: The outer heat-insulating layer is reinforced with a plain weave fabric made of a 50:50 mass ratio of high-silica glass fiber and T700 carbon fiber, with an areal density of 350 g / m². 2 The amount of reinforcing filler added to the matrix resin was adjusted to 6 wt% fumed silica and 1.5 wt% multi-walled carbon nanotubes by the total resin mass.

[0039] The intermediate transition layer uses high-temperature resistant PMI foam with a density gradient structure, with a total thickness of 16mm and a density of 110kg / m³ from the outside to the inside. 3 Outer layer (thickness 3mm), density 75kg / m³ 3 The core layer (10mm thick) has a density of 110kg / m³. 3 The inner layer (3mm thick); and before the PMI foam is laid, the foam surface is modified for 5 minutes using a 300W plasma device.

[0040] The inner load-bearing layer is reinforced with T800SC-12K carbon fiber plain weave fabric with a surface density of 280g / m² and the number of layers is adjusted to 8.

[0041] The curing process in this embodiment is consistent with that in Example 1, and the optimized high-temperature resistant lightweight composite material is finally obtained.

[0042] Example 3. This embodiment is an engineering verification of a full-size rocket fairing section. The fabricated fairing section has a diameter of 5.2m and a height of 3.5m, with a wall panel thickness of approximately 18mm. Its core material system and fabrication process principle are consistent with those of Example 1. The specific process parameters and layup design are as follows: The mold uses a composite material female mold adapted to the curvature of the fairing, and the mold surface is sprayed with a high-temperature resistant release agent. The layer design from the outside to the inside is as follows: the outer heat insulation layer is made of 6 layers of high silica fiber / carbon fiber hybrid fabric, and the middle transition layer is made of 75kg / m³ density fabric. 3 A 15mm thick high-temperature resistant PMI foam was used, with an inner load-bearing layer consisting of 8 layers of T700SC-12K carbon fiber cloth. During preparation, a two-component injection system was used to inject resin, with a total injection volume of approximately 150kg and an injection time of 4 hours. The injection temperature was controlled at 60-80℃. The curing process was completed in a large autoclave at a curing pressure of 0.4MPa, using a stepped temperature curing regime identical to that of Example 1. After curing, the material was cooled and demolded, then trimmed, machined, and subjected to full-size non-destructive testing to obtain a full-size fairing component.

[0043] Comparative Example 1. Traditional carbon fiber / epoxy + honeycomb sandwich structure: This comparative example is a conventional composite material structure for rocket fairings in the prior art, which adopts a separate design of an external heat-insulating layer + carbon fiber / epoxy load-bearing layer + aramid paper honeycomb core material, as detailed below: The external heat-insulating layer uses a low-density ablation material and is bonded to the surface of the outer load-bearing panel with epoxy structural adhesive; both the outer and inner load-bearing panels are made of T700 carbon fiber / bisphenol A type epoxy resin composite material, with a panel thickness of 1.5mm; the middle layer uses conventional aramid paper honeycomb with a density of 80kg / m³. 3 The thickness is 12mm. The manufacturing process adopts autoclave molding combined with secondary bonding process. First, carbon fiber / epoxy panels and outer heat protection layer are prepared separately. Then, structural adhesive is used to bond and cure the heat protection layer, panel and honeycomb core material in a secondary process to finally obtain sandwich structure composite material.

[0044] Comparative Example 2. Traditional aluminum honeycomb sandwich structure: This comparative example shows the existing technology of conventional aluminum alloy honeycomb sandwich structure for rocket fairings, as detailed below: The panel is made of 2024 aluminum alloy sheet, 1.0mm thick; the middle layer is 5052 aluminum alloy honeycomb core material, with a density of 150kg / m³. 3 The thickness is 12mm; the panel and the honeycomb core are connected by brazing, and an additional cork heat-insulating layer is glued to the outer surface to finally obtain an aluminum honeycomb sandwich structure component.

[0045] Example of results. I. Engineering verification was performed on the full-size component obtained in Example 3, and the results are as follows: 1. Non-destructive testing results: Ultrasonic C-scan showed that the internal quality of the component was good, with no large-area defects and an overall porosity of ≤1.5%, which meets the quality requirements of aerospace structural components; 2. Ground thermal test results: When the quartz lamp is heated to 250℃ on the outer surface and held for 15 minutes, the highest temperature on the inner surface is ≤85℃. The components show no deformation or delamination, and the structural integrity is good. 3. Static test results: Under aerodynamic load simulation, the structural failure load is 1.52 times the design load, which meets the strength design requirements of the rocket fairing; 4. Modal test results: The measured values ​​of the natural frequencies of the components deviate from the design values ​​by ≤5%, and the dynamic performance meets the design expectations.

[0046] 2. The properties of the materials prepared in each embodiment were tested, and the test results are shown in Table 1.

[0047] Table 1

[0048] The test results show that the glass transition temperature of the outer heat-resistant layer of the composite material described in this invention is not lower than 208℃, and the glass transition temperature of the inner structural layer is not lower than 186℃. After heat aging at 200℃ for 100 hours, the tensile strength retention rates of the inner and outer layers reach over 86% and 82% respectively, which are far higher than those of traditional carbon fiber / epoxy composite materials. This indicates that it has excellent high-temperature resistance and thermal stability, and can meet the long-term heat resistance requirements under high Mach number flight and reusable conditions. The overall density of the composite material of this invention is 0.78~0.85 g / cm³. 3 The material exhibits significant lightweight properties, while maintaining a tensile strength of at least 862 MPa for the inner load-bearing layer and an interlaminar shear strength of at least 26.5 MPa, achieving a balance between low density and high mechanical load-bearing capacity. Its post-impact compressive strength is at least 208 MPa, significantly superior to traditional structural materials, indicating that the composite material possesses higher impact toughness and damage tolerance, effectively resisting aerodynamic loads and acoustic vibration impacts. The thermal conductivity in the thickness direction is only 0.26~0.29 W / (m). The material exhibits excellent thermal insulation properties; the interface strength between the panel and the core material is high, showing no delamination or cracking after 100 cycles of thermal cycling from -100℃ to 200℃. It demonstrates strong structural integrity and environmental adaptability, solving the problems of easy debonding, warping, and peeling in traditional split structures. The high-temperature resistant PMI foam core material used in this invention meets the structural support requirements in terms of compressive strength, and its temperature resistance level matches that of the panel resin, preventing softening and instability of the core material at high temperatures. The overall performance of the sandwich layer is significantly superior to conventional foam and honeycomb core materials.

[0049] The full-size fairing component with a diameter of 5.2m prepared using the process of this invention showed good internal quality after ultrasonic C-scan testing, with a porosity of no more than 1.5%. Ground thermal tests, static tests, and modal tests all met the design specifications, indicating that the process of this invention is suitable for the integral molding of large-size, high-curvature components, with stable and reliable molding quality, and has good prospects for engineering applications.

[0050] Obviously, the above embodiments are merely illustrative examples for clear explanation and are not intended to limit the implementation. Those skilled in the art will recognize that other variations or modifications can be made based on the above description. It is neither necessary nor possible to exhaustively list all possible implementations here. However, obvious variations or modifications derived therefrom are still within the scope of protection of this invention.

Claims

1. A high-temperature resistant lightweight composite material, characterized in that, It includes a three-layer structure consisting of an outer heat-insulating layer, an intermediate transition layer, and an inner load-bearing layer, and the three-layer structure is an integrated co-curing molding structure; The outer heat-insulating layer uses modified epoxy / cyanate copolymer resin as the matrix resin and high-silica glass fiber and carbon fiber blended fabric as the reinforcement. The matrix resin is also compounded with fumed silica and carbon nanotube synergistic reinforcing fillers. The intermediate transition layer is polymethacrylimide foam; The inner bearing layer uses epoxy-modified cyanate ester resin as the matrix resin and T700 grade and / or T800 grade carbon fiber biaxial plain weave fabric as the reinforcement. The outer heat-insulating layer has a total thickness of 0.5~1.5mm and has 2~6 layers; the intermediate transition layer has a thickness of 5~20mm; and the inner load-bearing layer has a thickness of 0.8~2.0mm and has 4~8 layers. The modified epoxy / cyanate copolymer resin is obtained by copolymerizing epoxy resin and cyanate resin at a mass ratio of 30:70 to 50:50; the modified epoxy / cyanate copolymer resin contains 5 to 15 wt% toughening agent, which is carboxyl-terminated nitrile rubber or core-shell rubber. The mass ratio of high silica glass fiber to carbon fiber in the high silica glass fiber and carbon fiber blended fabric is 50:50~70:30; The outer heat-insulating layer contains 2-8 wt% fumed silica and 0.5-2 wt% carbon nanotubes. The epoxy-modified cyanate ester resin is obtained by blending cyanate ester resin and epoxy resin at a mass ratio of 60:40 to 80:20; 10 to 20 wt% of a thermoplastic toughening agent is added to the epoxy-modified cyanate ester resin, and the thermoplastic toughening agent is polyethersulfone or polyetherimide.

2. The high-temperature resistant lightweight composite material according to claim 1, characterized in that, The polymethacrylimide foam adopts a density gradient distribution with a denser outer layer and a looser inner layer.

3. A method for preparing a high-temperature resistant lightweight composite material as described in claim 1 or 2, characterized in that, Includes the following steps: S1. Coat the mold surface with a release agent, and lay out the dry fibers of the outer heat-insulating layer, the high-temperature resistant polymethacrylimide foam core material, and the dry fibers of the inner load-bearing layer in the design order, lay out the flow guiding system and seal the vacuum system. S2. After vacuuming and pressure testing the vacuum system, a two-component injection system is used to inject modified epoxy / cyanate ester copolymer resin into the dry fibers of the outer heat-insulating layer and epoxy modified cyanate ester resin into the dry fibers of the inner load-bearing layer to complete fiber impregnation. S3. After the glue injection is completed, close the glue injection port and maintain vacuum. Apply a pressure of 0.2~0.4MPa and use a stepped temperature increase regime for co-curing. After curing, cool down and demold. S4. Trim, machine and non-destructive test the demolded components to obtain the finished product.

4. The method for preparing the high-temperature resistant lightweight composite material according to claim 3, characterized in that, The modified epoxy / cyanate ester copolymer resin solution is prepared as follows: epoxy resin and cyanate ester resin are added to a reaction vessel in proportion, heated to 80~100℃ and stirred for 30min; fumed silica is added and dispersed at high speed of 2000~3000rpm for 30min; pre-dispersed carbon nanotubes are added and dispersed for another 30min; the temperature is lowered to below 60℃, curing agent and toughening agent are added, stirred evenly and then vacuum degassed.

5. The method for preparing the high-temperature resistant lightweight composite material according to claim 3, characterized in that, The preparation method of the epoxy modified cyanate ester resin solution is as follows: mix cyanate ester resin and epoxy resin in proportion, heat to 100~120℃ and stir until transparent; add thermoplastic toughening agent and stir to dissolve for 2~3 hours; cool down to 80℃, add curing agent, stir evenly and then vacuum degas.

6. The method for preparing the high-temperature resistant lightweight composite material according to claim 3, characterized in that, The dispensing temperature in step S2 is 60~80℃; The stepped temperature curing process in step S3 is as follows: first, heat to 120℃ and hold for 2 hours, then heat to 150℃ and hold for 2 hours, then heat to 180℃ and hold for 4 hours, and finally heat to 200℃ and hold for 2 hours for post-curing; after curing, cool down to below 60℃ to demold.

7. The application of a high-temperature resistant lightweight composite material as described in claim 1 or 2, characterized in that, It is used in rocket fairings, hypersonic vehicle skins, spacecraft reentry capsule thermal protection structures, hypersonic missile shells, aircraft engine nacelles, or high-speed train front fairings.

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