A thermal protection structure of a multi-layer hypersonic vehicle and a preparation method thereof

Abstract

A multi-layered thermal protection structure for hypersonic vehicles and its fabrication method are disclosed, relating to the field of hypersonic vehicle thermal protection technology. The structure comprises, from the outside in, a C / SiC composite material layer, a vacuum buffer layer, a white calcium carbonate layer, a solar cell module, and a thermoelectric generator. Through a five-layer synergistic design, multi-dimensional protection is achieved: the C / SiC composite material layer directly resists extreme high temperatures and ablation; the vacuum buffer layer blocks solid-state heat conduction, reducing heat transfer inwards; the white calcium carbonate layer reflects radiant heat, reducing heat input; and the solar cell module and thermoelectric generator convert residual heat into electrical energy, actively consuming heat. This integrated "blocking-reflection-conversion" design improves thermal protection efficiency by more than 40%, enabling the internal structure temperature to be controlled within a safe range of ≤100℃ even in extreme environments of 2000℃-3000℃. This invention achieves a comprehensive balance of high-temperature resistance, ablation resistance, low thermal conductivity, and high toughness.

Description

Technical Field

[0001] This invention relates to the field of thermal protection technology for hypersonic vehicles, and in particular to a multi-layered thermal protection structure for hypersonic vehicles and its preparation method. Background Technology

[0002] Hypersonic vehicles refer to aircraft with a cruising speed exceeding Mach 5, capable of reaching a battlefield 5,000 kilometers away within one hour. They possess advantages such as good stealth, strong penetration capability, high combat effectiveness, and wide operational space, enabling them to achieve strategic objectives such as precision strikes, rapid transport, and long-range reconnaissance, and represent a new strategic high ground in aerospace.

[0003] However, during flight, the intense friction between the surface of a hypersonic vehicle and the high-speed airflow generates aerodynamic heating, causing the surface temperature to rise rapidly. This can generate thermal stress on the internal structure of the vehicle, reducing its strength and compromising its structural integrity. Current technologies primarily address thermal protection by deploying suitable thermal protection structures in areas with harsh aerodynamic and thermal environments. Existing thermal protection structures suffer from the following problems: 1. They often rely on passive insulation using a single material, such as pure C / SiC coatings that only block heat through the material's high-temperature resistance; or they rely on a single method of heat dissipation, such as C / C composite materials that rely on high thermal conductivity to conduct heat, but have poor oxidation resistance. These structures suffer from severe heat accumulation, low energy utilization, and insufficient structural stability. 2. Existing composite materials, such as C / C and C / SiC... At high temperatures, thermal expansion can easily cause cracks, or vibration and impact can cause the single coating to peel off, leading to structural failure; 3. To ensure high-temperature resistance, existing thermal protection technologies often use thick-walled single materials, such as ultra-high temperature ceramic layers with a thickness of 5-10 mm, resulting in heavy structures that affect the range and maneuverability of the aircraft; 4. Existing ultra-high temperature composite materials also have obvious shortcomings, such as C / C composites having poor oxidation resistance and requiring additional coating protection; C / SiC composites having low thermal conductivity and easily accumulating heat; and UHTC ceramics having low toughness and being prone to brittle fracture; Therefore, in response to the thermal protection requirements of hypersonic vehicles in extreme high-temperature environments, proposing a multi-layered energy conversion type hypersonic vehicle thermal protection structure and its preparation method has become a basic demand of those skilled in the art. Summary of the Invention

[0004] The purpose of this invention is to overcome the shortcomings of the prior art and provide a multi-layered hypersonic vehicle thermal protection structure and its preparation method.

[0005] The technical solution adopted in this invention is:

[0006] A multi-layered thermal protection structure for a hypersonic vehicle includes, from the outside to the inside, a C / SiC composite material layer, a vacuum buffer layer, a white calcium carbonate layer, a solar cell module, and a thermoelectric generator. A square frame is fixed to the outside of the vacuum buffer layer. The vacuum buffer layer includes a support member and a buffer spring disposed between the C / SiC composite material layer and the white calcium carbonate layer. The support member has an "I"-shaped structure, including a cylindrical part and extension rings at the upper and lower ends of the cylindrical part. The extension rings of the support member are respectively bonded to the inner wall of the C / SiC composite material layer and the white calcium carbonate layer by high-temperature resistant ceramic adhesive. The outer wall of the calcium carbonate layer is bonded and fixed. Protrusions are fixed at the top and bottom of the cylindrical part. The buffer spring is sleeved on the outside of the cylindrical part, and the two ends of the buffer spring pass through the perforations on the protrusions and are interference-fitted with the perforations. The solar cell module includes multiple solar panels, which are connected in series or in parallel to form a square array structure. Each solar panel is fixedly connected to the white calcium carbonate layer through titanium alloy brackets at both ends. Thermally conductive silicone is coated at the contact surfaces between the solar panel and the white calcium carbonate layer, and at the contact surfaces between the solar panel and the thermoelectric generator to connect them.

[0007] The thermal protection structure of the multi-layered hypersonic vehicle uses an Al2O3-based high-temperature ceramic adhesive with a temperature resistance ≥1600℃, and an Inconel 718 high-temperature alloy. The solar panel is a GaAs-based high-temperature solar cell with a power density ≥50W / m², an operating temperature of -50℃ to 200℃, and a photoelectric conversion efficiency ≥25%. The thermoelectric generator is a Bi2Te3-based thermoelectric material with a power density ≥10W / m², a Seebeck coefficient ≥200μV / K, and a thickness of 3mm-5mm. The thermally conductive silicone has a thermal conductivity ≥1.5W / (m·K).

[0008] The thermal protection structure of the multi-layer hypersonic vehicle has an L-shaped titanium alloy bracket with a thickness of 1mm-2mm. Mo alloy bolts for fixing solar panels and white calcium carbonate layers are respectively inserted on the titanium alloy bracket. The Mo alloy bolts have a temperature resistance of ≥600℃.

[0009] The thermal protection structure of the multi-layer hypersonic vehicle has a pre-drilled joint on a square frame. A ceramic sealing joint and a molecular pump interface are fixed in the pre-drilled joint. The leads on the solar panel pass through the ceramic sealing joint from the square frame and are connected in parallel with the leads on the thermoelectric generator. The molecular pump interface is connected to the external molecular pump vacuum port.

[0010] A method for fabricating a multi-layered thermal protection structure for a hypersonic vehicle specifically includes the following steps:

[0011] S1. C / SiC composite material layers were prepared using a modified CVI method;

[0012] S2. Process the support and buffer spring of the vacuum buffer layer. After assembling the support and buffer spring, bond one end of the support to the inner wall of the C / SiC composite material layer with high temperature resistant ceramic adhesive.

[0013] S3, sintered white calcium carbonate layer, the other end of the support is bonded and fixed to the white calcium carbonate layer;

[0014] S4. Secure the solar panel to the white calcium carbonate layer using a titanium alloy bracket and Mo alloy bolts to complete the overall assembly;

[0015] S5. Perform vacuum treatment and overall sealing on the assembled thermal protection structure:

[0016] The interface of the vacuum buffer layer is connected to a molecular pump. First, a rough vacuum of 10 Pa is created, then the molecular pump is turned on to evacuate to 5 × 10 Pa. -4 Pa, maintain pressure for 30 minutes, pressure change ≤1Pa;

[0017] Close the vacuum valve, seal the interface with a copper gasket, and apply thermally conductive silicone to the contact surfaces of the solar panel and the white calcium carbonate layer, and the contact surfaces of the solar panel and the thermoelectric generator. Fill with thermally conductive silicone sealant, cover the surface with C / SiC tape, and secure with ceramic wire.

[0018] The method for preparing the thermal protection structure of the multi-layer hypersonic vehicle, specifically the steps for preparing the C / SiC composite material layer in step S1 are as follows:

[0019] S1.1. T700 grade carbon fiber is selected, and carbon fiber preforms are prepared using 3D four-way weaving technology, with a fiber volume fraction of 40%-45%;

[0020] S1.2 Place the carbon fiber preform into a horizontal CVI reactor, heat the furnace to 900°C, and purify it with inert argon gas (Ar) for 2 hours.

[0021] S1.3. Introduce reactive gases SiCl4 and CH4 at a flow rate ratio of 3:1. The flow rate of SiCl4 is 80-100 sccm and the flow rate of CH4 is 25-30 sccm. The pressure inside the furnace is controlled at 0.3 MPa to complete the vapor deposition.

[0022] S1.4 Perform pitch impregnation and pyrolysis densification operations on the carbon fiber preform that has completed vapor deposition in the previous step.

[0023] S1.5. Using a rotating tool at a speed of 5 r / min, the deposition is stopped after 4.5 cycles. The resulting C / SiC composite material layer has a thickness of 2 mm-3 mm, a SiC matrix grain size of 5 μm-10 μm, and is resistant to high temperatures of 1800℃-2200℃.

[0024] The specific processing steps of the vacuum buffer layer in step S2 of the fabrication method of the thermal protection structure of the multi-layer hypersonic vehicle are as follows:

[0025] S2.1. The support is prepared using the same CVI method as the previous step. The support is a C / SiC composite material with a diameter of 5mm-8mm and a height of 10mm-15mm. The surface is sandblasted with a roughness of Ra=5μm. Protrusions with perforations are machined at both ends of the cylindrical support.

[0026] S2.2. Select Inconel 718 alloy wire to process into a helical buffer spring with a free length of 8mm-12mm, 5 turns, elastic modulus of 12N / m, and compression ≤5mm.

[0027] S2.3 Mark the mounting points of the support on the inner wall of the C / SiC composite layer 1, coat one end of the support with high-temperature resistant ceramic adhesive, and bond it to the marked mounting point. Cur it in an oven at 250°C for 3 hours.

[0028] S2.4. The buffer spring is sleeved on the outside of the cylindrical support. The two ends of the buffer spring are passed through the holes on the upper and lower protrusions of the cylindrical support. The square frame is a C / SiC thin plate. Its upper and lower end faces are welded to the bottom of the C / SiC composite material layer and the top of the white calcium carbonate layer, respectively. A pre-drilled hole joint is opened on the square frame, and a ceramic sealing joint and a molecular pump interface are fixed in the pre-drilled hole joint, respectively.

[0029] The specific preparation steps of step S3 in the preparation method of the thermal protection structure of the multi-layer hypersonic vehicle are as follows:

[0030] S3.1 First, dry the calcium carbonate powder in a vacuum drying oven at 120℃ for 4 hours, and then pass it through a 200-mesh sieve;

[0031] S3.2 Fill the graphite mold, apply 5MPa pressure for pre-compression molding, place in a box furnace, sinter at 300℃-400℃ for 3 hours in air atmosphere, demold after natural cooling to room temperature, and grind the surface to Ra≤1.6μm, thickness 0.5mm-2mm, reflectivity≥85%;

[0032] S3.3. Coat the other end of the support with high-temperature resistant ceramic adhesive, bond it to the white calcium carbonate layer, and cure it in an oven at 250°C for 3 hours.

[0033] The method for preparing the thermal protection structure of the multi-layer hypersonic vehicle, wherein the installation steps of the solar panel, white calcium carbonate layer, and thermoelectric generator in step S4 are as follows:

[0034] S4.1 Connect the solar panels in series / parallel to form a square array, attach titanium alloy brackets to both sides, and fix them to the white calcium carbonate layer with Mo alloy bolts on the titanium alloy brackets;

[0035] S4.2 Attach the thermoelectric generator to the bottom of the solar panel, apply thermally conductive silicone to the hot side of the contact surface, apply 0.5MPa pressure and hold for 1 hour, attach a graphite thermally conductive sheet to the cold side, connect the lead wires on the thermoelectric generator to the circuit of the solar panel in parallel, and connect the total output terminal to the aircraft power supply system through a high-temperature resistant cable.

[0036] The preparation method of the thermal protection structure of the multi-layer hypersonic vehicle, the specific operation of the asphalt impregnation in step S1.4 is as follows: the carbon fiber preform that has completed vapor deposition is placed in a special high-pressure impregnation kettle. First, the special high-pressure impregnation kettle is evacuated to a vacuum degree ≤5Pa and kept for 30 minutes to remove residual air inside the preform. Then, molten mesophase asphalt with a softening point of 200℃-220℃ is injected into the special high-pressure impregnation kettle, the temperature is raised to 250℃ and a pressure of 0.8MPa-1.0MPa is applied, and the impregnation time is 2h-3h to ensure that the asphalt is fully impregnated into the micropores of the carbon fiber preform.

[0037] The steps of the pyrolysis densification operation are as follows: the carbon fiber preform impregnated with asphalt is transferred to an inert atmosphere sintering furnace, argon gas with a flow rate of 50 sccm-60 sccm is introduced as a protective atmosphere, the temperature is raised to 1000℃-1200℃ at a rate of 5℃ / min, and the temperature is held for 2 hours to complete the pyrolysis; the "impregnation-pyrolysis" process is repeated 1-2 times according to the porosity requirements.

[0038] Due to the adoption of the technical solution described above, the present invention has the following advantages:

[0039] 1. The thermal protection structure and its preparation method for the multi-layer hypersonic vehicle described in this invention achieve multi-dimensional protection through a five-layer synergistic design: the C / SiC composite material layer directly resists extreme high temperatures and ablation; the vacuum buffer layer blocks solid heat conduction, reducing heat transfer inward; the white calcium carbonate layer reflects radiant heat, reducing heat input; the solar panel and thermoelectric generator convert the remaining heat into electrical energy, actively consuming heat; this integrated "blocking-reflection-conversion" design improves thermal protection efficiency by more than 40%, and can control the internal structure temperature within a safe range of ≤100℃ in extreme environments of 2000℃-3000℃.

[0040] 2. The thermal protection structure and its preparation method of the multi-layer hypersonic vehicle described in this invention are based on the principle of energy conservation. The solar panel absorbs the radiant heat that penetrates the white calcium carbonate layer and converts it into electrical energy. The thermoelectric generator uses the temperature difference between the layers to generate electricity. Both types of energy can be converted into usable electrical energy by 10%-15% of the waste heat through an external recovery device to power the internal equipment of the vehicle, such as sensors and communication systems, thereby reducing the dependence on airborne energy.

[0041] 3. The thermal protection structure and its preparation method of the multi-layer hypersonic vehicle described in this invention ensure structural strength of the support components in the vacuum buffer layer, while having good compatibility with the outer layer material and reducing interface stress; the buffer spring is a buffer component that can buffer the deformation caused by thermal expansion of the support components (compression ≤ 5mm), alleviate vibration impact (adapted to vibration frequency of 10Hz-50Hz), prevent the support components from being damaged, and improve the integrity rate of the structure under repeated thermal cycling to more than 95%;

[0042] 4. The thermal protection structure of the multi-layer hypersonic vehicle and its preparation method described in this invention involve introducing reactive gases SiCl4 and CH4, performing preliminary vapor deposition of the SiC matrix according to predetermined parameters, and filling most of the pores of the carbon fiber preform; then, through a mesophase pitch impregnation process, ensuring that the pitch fully penetrates into the micropores of the carbon fiber preform; finally, during pyrolysis, the pitch carbonizes to form a carbon matrix, further filling the pores; repeating the "impregnation-pyrolysis" process 1-2 times, each process can reduce the porosity by 1%-2%, and combined with vapor deposition, the target porosity ≤5% can be achieved, shortening the cycle from the original 7-8 weeks to 4-5 weeks;

[0043] 5. The thermal protection structure and preparation method of the multi-layer hypersonic vehicle described in this invention have a C / SiC composite material layer thickness controlled at 2mm-3mm, balancing high temperature resistance and lightweight. The calcium carbonate layer thickness is only 0.5mm-2mm, with a reflection efficiency of over 85%, and the weight is reduced by 30% compared to traditional thermal insulation layers. The overall structural weight is reduced by 25%-30% compared to existing technologies, significantly improving the vehicle's penetration capability and endurance. Furthermore, the C / SiC composite material layer solves the problems of oxidation and ablation resistance, while the vacuum layer and calcium carbonate layer compensate for the low thermal conductivity of C / SiC. The elastic design of the support components and buffer springs alleviates the problem of low toughness and brittle fracture of UHTC ceramics used in existing technologies. This invention achieves a comprehensive balance of high temperature resistance, ablation resistance, low thermal conductivity, and high toughness. Attached Figure Description

[0044] Figure 1 This is a schematic diagram of the structure of the present invention.

[0045] Figure 2 This is a schematic diagram of the internal structure of the vacuum buffer layer of the present invention.

[0046] Figure 3 This is a schematic diagram of the structure of the support component of the present invention.

[0047] Figure 4 This is a schematic diagram of the pre-drilled hole connector on the square frame of the present invention.

[0048] Figure 5 This is a schematic diagram of the structure of the solar panel of the present invention, which is arranged in a square array.

[0049] In the diagram: 1. C / SiC composite material layer; 2. Buffer spring; 3. White calcium carbonate layer; 4. Solar panel; 5. Thermoelectric generator; 6. Graphite heat-conducting sheet; 7. Support component; 8. Protrusion; 9. Perforation; 10. Square frame; 11. Pre-drilled hole connector; 12. Titanium alloy bracket; 13. Mo alloy bolt; 14. Lead wire. Detailed Implementation

[0050] The present invention will be further explained and described below with reference to the accompanying drawings and embodiments. However, this should not be construed as limiting the scope of protection of the present invention. The purpose of disclosing the present invention is to protect all technical improvements within the scope of the present invention.

[0051] Combined with appendix Figure 1-5 The aforementioned multi-layered hypersonic vehicle thermal protection structure and its fabrication method include, from the outside to the inside, a C / SiC composite material layer 1, a vacuum buffer layer, a white calcium carbonate layer 3, a solar cell module, and a thermoelectric generator 5. A square frame 10 is fixed to the outside of the vacuum buffer layer. The vacuum buffer layer includes a support member 7 and a buffer spring 2 disposed between the C / SiC composite material layer 1 and the white calcium carbonate layer 3. The support member 7 has an "I"-shaped structure, including a cylindrical part and extension rings at the upper and lower ends of the cylindrical part. The extension rings of the support member 7 are bonded to the inner wall of the C / SiC composite material layer 1 and the outer wall of the white calcium carbonate layer 3 respectively using high-temperature resistant ceramic adhesive. The cylindrical component is fixed with protrusions 8 at both the top and bottom. A buffer spring 2 is sleeved on the outside of the cylindrical component, and both ends of the buffer spring 2 pass through the holes 9 on the protrusions 8 and are interference-fitted with the holes 9. The solar cell module includes multiple solar panels 4, which are connected in series or in parallel to form a square array structure. Each solar panel 4 is fixedly connected to the white calcium carbonate layer 3 through titanium alloy brackets 12 at both ends. Thermally conductive silicone is coated at the contact surfaces between the solar panel 4 and the white calcium carbonate layer 3, and at the contact surfaces between the solar panel 4 and the thermoelectric generator 5 to connect them. The thermal conductivity of the thermally conductive silicone is ≥1.5W / (m·k).

[0052] Specifically, the high-temperature resistant ceramic adhesive is an Al2O3-based adhesive with a temperature resistance of ≥1600℃, and the high-temperature alloy is Inconel 718;

[0053] Solar panel 4 is a GaAs-based high-temperature solar cell with a power density ≥50W / m², an operating temperature of -50℃ to 200℃, and a photoelectric conversion efficiency ≥25%.

[0054] The thermoelectric generator 5 is made of Bi2Te3-based thermoelectric material, with a power density ≥10W / m², Seebeck coefficient ≥200μV / K, and thickness of 3mm-5mm.

[0055] Specifically, the titanium alloy bracket 12 is set as an "L" shaped structure with a thickness of 1mm-2mm. Mo alloy bolts 13 for fixing the solar panel 4 and the white calcium carbonate layer 3 are respectively inserted on the titanium alloy bracket 12. The Mo alloy bolts 13 have a temperature resistance of ≥600℃.

[0056] Specifically, a pre-drilled connector 11 is provided on the square frame 10. A ceramic sealing connector and a molecular pump interface are fixed in the pre-drilled connector 11. The lead wires on the solar panel 4 pass through the ceramic sealing connector from the square frame 10 and are connected in parallel with the lead wires 14 on the thermoelectric generator 5. The molecular pump interface is connected to the external molecular pump vacuum port.

[0057] A method for fabricating a multi-layered thermal protection structure for a hypersonic vehicle specifically includes the following steps:

[0058] S1. The C / SiC composite layer 1 is prepared using the modified CVI method. The specific steps are as follows:

[0059] S1.1. T700 grade carbon fiber is selected and carbon fiber preforms are prepared using 3D four-way weaving process. The fiber volume fraction is 40%-45%. The shape is a pointed cone that matches the head cone. The bottom diameter is 300mm and the height is 500mm. The density of the preform is 0.8g / cm³.

[0060] S1.2 Place the carbon fiber preform into a horizontal CVI reactor, heat the furnace to 900°C, and purify it with inert argon gas (Ar) for 2 hours.

[0061] S1.3. Introduce reactive gases SiCl4 and CH4 at a flow rate ratio of 3:1. The flow rate of SiCl4 is 80-100 sccm and the flow rate of CH4 is 25-30 sccm. The pressure inside the furnace is controlled at 0.3 MPa to complete the vapor deposition.

[0062] S1.4 Perform pitch impregnation and pyrolysis densification operations on the carbon fiber preform that has completed vapor deposition in the previous step.

[0063] The specific operation of asphalt impregnation is as follows: The carbon fiber preform that has undergone vapor deposition is placed in a dedicated high-pressure impregnation vessel. First, the vessel is evacuated to a vacuum level ≤5 Pa and maintained for 30 minutes to remove residual air from the preform. Then, molten mesophase asphalt with a softening point of 200℃-220℃ is injected into the vessel. The temperature is raised to 250℃ and a pressure of 0.8MPa-1.0MPa is applied. The impregnation time is 2-3 hours to ensure that the asphalt is fully impregnated into the micropores of the carbon fiber preform.

[0064] The steps of the pyrolysis densification operation are as follows: the carbon fiber preform impregnated with pitch is transferred to an inert atmosphere sintering furnace, argon gas with a flow rate of 50 sccm-60 sccm is introduced as a protective atmosphere, the temperature is raised to 1000℃-1200℃ at a rate of 5℃ / min, and held for 2 hours to complete the pyrolysis; the "impregnation-pyrolysis" process is repeated 1-2 times according to the porosity requirements.

[0065] S1.5. By rotating the tool at a speed of 5 r / min, the deposition is stopped after 4.5 cycles. The resulting C / SiC composite material layer has a thickness of 2 mm-3 mm, a SiC matrix grain size of 5 μm-10 μm, and is resistant to high temperatures of 1800℃-2200℃. It also has both oxidation resistance and ablation resistance.

[0066] S2. Process the support component 7 and buffer spring 2 of the vacuum buffer layer. After assembling the support component 7 and buffer spring 2, bond one end of the support component 7 to the inner wall of the C / SiC composite material layer 1 with high-temperature resistant ceramic adhesive. The specific processing steps of the vacuum buffer layer are as follows:

[0067] S2.1. The support component 7 is prepared using the same CVI method as the previous step. The support component 7 is a C / SiC composite material. After SiCl4 / CH4 deposition, it is processed into an "I" shaped structure with a diameter of 6 mm and a height of 12 mm. The surface is sandblasted to a roughness of Ra=5μm and a compressive strength ≥250Mpa. The two ends of the cylindrical part of the support component 7 are processed with protrusions 8, and the protrusions 8 have a perforation 9 in the middle. The support component is then solution treated at 1000℃ for 2 hours.

[0068] S2.2. Inconel 718 alloy wire with a diameter of 0.8mm is selected and processed into a helical buffer spring 2 by a spring coiling machine. The free length is 10mm, the number of turns is 5, the elastic modulus is 12N / m, and the compression is ≤5mm. It is designed according to the vibration frequency of the aircraft 10Hz-50Hz to ensure a buffering efficiency of ≥80% during thermal expansion and deformation.

[0069] S2.3 Mark the mounting points of the support 7 on the inner wall of the C / SiC composite material layer 1, coat one end of the support 7 with Al2O3-based ceramic adhesive with a solid content of 60%, and bond it to the marked mounting points. Cur it in an oven at 250℃ for 3 hours with a heating rate of 5℃ / min.

[0070] S2.4. The buffer spring 2 is sleeved on the outside of the cylindrical support 7. The two ends of the buffer spring 2 are passed through the through holes 9 on the upper and lower protrusions 8 of the cylindrical support to ensure that the spring has no preload in its natural state. The square frame 10 is a 1mm thick C / SiC thin plate. Its upper and lower end faces are welded to the bottom of the C / SiC composite material layer 1 and the top of the white calcium carbonate layer 3, respectively. The weld width is 2mm. A pre-drilled hole joint 11 is opened on the square frame 10. A ceramic sealing joint with a temperature resistance of 300℃ and a molecular pump interface are welded and fixed in the pre-drilled hole joint 11. The interface for connecting the molecular pump is Φ8mm and the interface material is Mo alloy.

[0071] S3. Sintered white calcium carbonate layer 3, the other end of the support 7 is bonded and fixed to the white calcium carbonate layer 3, specifically including the following steps:

[0072] S3.1 First, dry the calcium carbonate powder with a particle size of 8μm in a vacuum drying oven at 120℃ for 4 hours to remove moisture, with a vacuum degree of -0.1MPa, and pass it through a 200-mesh sieve.

[0073] S3.2. The sieved calcium carbonate powder is filled into a graphite mold that matches the inner wall of the die head cone. The inner wall roughness of the graphite mold is Ra=0.8μm. A pressure of 5MPa is applied for pre-pressing and molding. The mold is then placed in a box furnace and sintered at 300℃-400℃ in air for 3 hours with a heating rate of 10℃ / min. After natural cooling to room temperature, the mold is demolded. The thickness of the sintered calcium carbonate layer is 1mm. The surface is ground to Ra=1.2μm with a thickness of 0.5mm-2mm. The measured reflectivity in the 2μm-10μm band is 88%.

[0074] S3.3. Coat the other end of the support 7 with the same Al2O3-based ceramic adhesive, bond it to the white calcium carbonate layer 3, and cure it in an oven at 250°C for 3 hours with a heating rate of 5°C / min to ensure a bonding strength ≥3MPa.

[0075] S4. Fix the solar panel 4 to the white calcium carbonate layer 3 using the titanium alloy bracket 12 and Mo alloy bolts 13 to complete the overall assembly. The size of a single solar panel 4 is 50mm×50mm, and the conversion efficiency is 26%. The installation steps of the solar panel 4, the white calcium carbonate layer 3, and the thermoelectric generator 5 are as follows:

[0076] S4.1 Connect the solar panels 4 in series / parallel to form a square array with a total power ≥30W. Attach titanium alloy brackets 12 to both sides. The thickness of the titanium alloy brackets 12 is 1.5mm. Fix them to the white calcium carbonate layer 3 with Mo alloy bolts 13 on the titanium alloy brackets 12. The Mo alloy bolts 13 are heat resistant up to 650℃ and have a preload torque of 8N·m.

[0077] S4.2 The thermoelectric generator 5 is a Bi2Te3-based thermoelectric generator with dimensions of 40mm×40mm×4mm and a Seebeck coefficient of 220μV / K. The thermoelectric generator 5 is attached to the underside of the solar panel 4. Thermally conductive silicone is applied to the hot side of the contact surface, and a pressure of 0.5MPa is applied and held for 1 hour to ensure that the interface contact thermal resistance is ≤0.08K·m² / W. A graphite thermally conductive sheet 6 with a thickness of 0.2mm is attached to the cold side. After the lead wire 14 on the thermoelectric generator 5 is connected in parallel with the line of the solar panel 4, the total output terminal is connected to the aircraft power supply system through a high-temperature resistant cable.

[0078] S5. Perform vacuum treatment and overall sealing on the assembled thermal protection structure:

[0079] The molecular pump interface of the vacuum buffer layer is connected to a molecular pump with a pumping speed of 500 L / s. First, a rough vacuum of 10 Pa is achieved, and then the molecular pump is turned on to pump to 5 × 10 Pa. -4 Pa, maintain pressure for 30 minutes, pressure change ≤1Pa;

[0080] The vacuum valve is closed, and the molecular pump interface is sealed with a copper gasket to ensure long-term vacuum stability with a leakage rate ≤5×10-9Pa·m³ / s. Thermally conductive silicone is applied to the contact surfaces between the solar panel 4 and the white calcium carbonate layer 3, and between the solar panel 4 and the thermoelectric generator 5. Thermally conductive silicone sealant is used to fill the gaps, and the surface is covered with a 20mm wide C / SiC tape, which is then tied and fixed with ceramic wire to complete the overall assembly and sealing. The assembled thermal protection structure is then installed on the missile nose simulator, and ground high-temperature wind tunnel tests can be conducted.

[0081] The parts of this invention not described in detail are prior art.

[0082] The embodiments selected herein for the purpose of disclosing the inventive objectives are currently considered suitable; however, it should be understood that the invention is intended to include all variations and modifications of the embodiments that fall within the scope of this concept and invention.

Claims

1. A thermal protection structure of a multi-layer hypersonic vehicle, comprising, from outside to inside, a C / SiC composite material layer, a vacuum buffer layer, a white calcium carbonate layer, a solar cell assembly and a thermoelectric power sheet; characterized in that: A square frame is fixed to the outside of the vacuum buffer layer. The vacuum buffer layer includes a support and a buffer spring disposed between the C / SiC composite material layer and the white calcium carbonate layer. The support has an "I" shaped structure, including a cylindrical part and extension rings at the upper and lower ends of the cylindrical part. The extension rings of the support are bonded and fixed to the inner wall of the C / SiC composite material layer and the outer wall of the white calcium carbonate layer respectively by high-temperature resistant ceramic adhesive. Protrusions are fixed at the upper and lower parts of the cylindrical part. The buffer spring is sleeved on the outside of the cylindrical part, and the two ends of the buffer spring pass through the perforations on the protrusions and are interference-fitted with the perforations. The solar cell module includes multiple solar panels, which are connected in series or in parallel to form a square array structure. Each solar panel is fixedly connected to the white calcium carbonate layer through titanium alloy brackets at both ends. Thermally conductive silicone is coated at the contact surfaces between the solar panel and the white calcium carbonate layer and between the solar panel and the thermoelectric generator to connect them.

2. The thermal protection structure of a multilayer hypersonic vehicle according to claim 1, characterized in that: The high-temperature resistant ceramic adhesive is an Al2O3-based adhesive with a temperature resistance of ≥1600℃. The high-temperature alloy is Inconel718. The solar panel is a GaAs-based high-temperature solar cell with a power density of ≥50W / m², an operating temperature of -50℃ to 200℃, and a photoelectric conversion efficiency of ≥25%. The thermoelectric generator is a Bi2Te3-based thermoelectric material with a power density of ≥10W / m², a Seebeck coefficient of ≥200μV / K, and a thickness of 3mm-5mm. The thermally conductive silicone has a thermal conductivity of ≥1.5W / (m·k).

3. The thermal protection structure of a multilayer hypersonic vehicle according to claim 1, characterized in that: The titanium alloy bracket is designed with an "L" shape and a thickness of 1mm-2mm. Mo alloy bolts for fixing the solar panel and the white calcium carbonate layer are respectively inserted into the titanium alloy bracket. The Mo alloy bolts have a temperature resistance of ≥600℃.

4. The thermal protection structure of a multilayer hypersonic vehicle according to claim 1, characterized in that: A pre-drilled connector is provided on the square frame. A ceramic sealing connector and a molecular pump interface are fixed in the pre-drilled connector. The lead wires on the solar panel pass through the ceramic sealing connector from the square frame and are connected in parallel with the lead wires on the thermoelectric cell. The molecular pump interface is connected to the external molecular pump vacuum port.

5. A method for preparing a thermal protection structure for a multi-layered hypersonic vehicle as described in any one of claims 1-4, characterized in that: Specifically, the following steps are included: S1. C / SiC composite material layers were prepared using a modified CVI method; S2. Process the support and buffer spring of the vacuum buffer layer. After assembling the support and buffer spring, bond one end of the support to the inner wall of the C / SiC composite material layer with high temperature resistant ceramic adhesive. S3, sintered white calcium carbonate layer, the other end of the support is bonded and fixed to the white calcium carbonate layer; S4. Secure the solar panel to the white calcium carbonate layer using a titanium alloy bracket and Mo alloy bolts to complete the overall assembly; S5. Perform vacuum treatment and overall sealing on the assembled thermal protection structure: The interface connection of the vacuum buffer layer is connected with a molecular pump. First, rough vacuum 10 Pa is extracted, and then the molecular pump is started to extract 5 x 10 -4 Pa, and pressure is kept for 30 minutes, and the pressure change is less than or equal to 1 Pa. Close the vacuum valve, seal the interface with a copper gasket, and apply thermally conductive silicone to the contact surfaces of the solar panel and the white calcium carbonate layer, and the contact surfaces of the solar panel and the thermoelectric generator. Fill with thermally conductive silicone sealant, cover the surface with C / SiC tape, and secure with ceramic wire.

6. The method of claim 5, wherein: The specific steps for preparing the C / SiC composite layer in step S1 are as follows: S1.

1. T700 grade carbon fiber is selected, and carbon fiber preforms are prepared using 3D four-way weaving technology, with a fiber volume fraction of 40%-45%; S1.2 Place the carbon fiber preform into a horizontal CVI reactor, heat the furnace to 900°C, and purify it with inert argon gas (Ar) for 2 hours. S1.

3. Introduce reactive gases SiCl4 and CH4 at a flow rate ratio of 3:

1. The flow rate of SiCl4 is 80-100 sccm and the flow rate of CH4 is 25-30 sccm. The pressure inside the furnace is controlled at 0.3 MPa to complete the vapor deposition. S1.4 Perform pitch impregnation and pyrolysis densification operations on the carbon fiber preform that has completed vapor deposition in the previous step. S1.

5. Using a rotating tool at a speed of 5 r / min, the deposition is stopped after 4.5 cycles. The resulting C / SiC composite material layer has a thickness of 2 mm-3 mm, a SiC matrix grain size of 5 μm-10 μm, and is resistant to high temperatures of 1800℃-2200℃.

7. The method of claim 5, wherein: The specific operation of bitumen impregnation in step S1.4 is as follows: the carbon fiber preform that has completed vapor deposition is placed into a special high-pressure impregnation vessel. First, the special high-pressure impregnation vessel is evacuated to a vacuum degree ≤5Pa and maintained for 30 minutes to remove residual air inside the preform. Then, molten mesophase bitumen with a softening point of 200℃-220℃ is injected into the special high-pressure impregnation vessel, the temperature is raised to 250℃ and a pressure of 0.8MPa-1.0MPa is applied, and the impregnation time is 2h-3h to ensure that the bitumen is fully impregnated into the micropores of the carbon fiber preform. The steps of the pyrolysis densification operation are as follows: the carbon fiber preform impregnated with pitch is transferred to an inert atmosphere sintering furnace, argon gas with a flow rate of 50 sccm-60 sccm is introduced as a protective atmosphere, the temperature is raised to 1000℃-1200℃ at a rate of 5℃ / min, and the temperature is held for 2 hours to complete the pyrolysis; the "impregnation-pyrolysis" process is repeated 1-2 times according to the porosity requirements.

8. The method of claim 5, wherein: The specific processing steps for the vacuum buffer layer in step S2 are as follows: S2.

1. The support is prepared using the same CVI method as the previous step. The support is a C / SiC composite material with a diameter of 5mm-8mm and a height of 10mm-15mm. The surface is sandblasted with a roughness of Ra=5μm. Protrusions with perforations are machined at both ends of the cylindrical support. S2.

2. Select Inconel 718 alloy wire to process into a helical buffer spring with a free length of 8mm-12mm, 5 turns, elastic modulus of 12N / m, and compression ≤5mm. S2.3 Mark the mounting points of the support on the inner wall of the C / SiC composite layer 1, coat one end of the support with high-temperature resistant ceramic adhesive, and bond it to the marked mounting point. Cur it in an oven at 250°C for 3 hours. S2.

4. The buffer spring is sleeved on the outside of the cylindrical support. The two ends of the buffer spring are passed through the holes on the upper and lower protrusions of the cylindrical support. The square frame is a C / SiC thin plate. Its upper and lower end faces are welded to the bottom of the C / SiC composite material layer and the top of the white calcium carbonate layer, respectively. A pre-drilled hole joint is opened on the square frame, and a ceramic sealing joint and a molecular pump interface are fixed in the pre-drilled hole joint, respectively.

9. The method of claim 5, wherein: The specific preparation steps for step S3 are as follows: S3.1 First, dry the calcium carbonate powder in a vacuum drying oven at 120℃ for 4 hours, and then pass it through a 200-mesh sieve; S3.2 Fill the graphite mold, apply 5MPa pressure for pre-compression molding, place in a box furnace, sinter at 300℃-400℃ for 3 hours in air atmosphere, demold after natural cooling to room temperature, and grind the surface to Ra≤1.6μm, thickness 0.5mm-2mm, reflectivity≥85%; S3.

3. Coat the other end of the support with high-temperature resistant ceramic adhesive, bond it to the white calcium carbonate layer, and cure it in an oven at 250°C for 3 hours.

10. The method of claim 5, wherein: The installation steps for the solar panel, white calcium carbonate layer, and thermoelectric generator described in step S4 are as follows: S4.1 Connect the solar panels in series / parallel to form a square array, attach titanium alloy brackets to both sides, and fix them to the white calcium carbonate layer with Mo alloy bolts on the titanium alloy brackets; S4.2 Attach the thermoelectric generator to the bottom of the solar panel, apply thermally conductive silicone to the hot side of the contact surface, apply 0.5MPa pressure and hold for 1 hour, attach a graphite thermally conductive sheet to the cold side, connect the lead wires on the thermoelectric generator to the circuit of the solar panel in parallel, and connect the total output terminal to the aircraft power supply system through a high-temperature resistant cable.

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