A surface toughened gradient thermal protection structure and its preparation method and application

Abstract

The present invention relates to the technical field of preparation of thermal protection materials for spacecraft, and provides a surface-toughened gradient thermal protection structure and a preparation method thereof, wherein the thermal protection structure comprises: an outer layer mainly composed of a fiber-reinforced glass-based composite thermal protection panel; a ceramic fiber rigid insulation tile used in the thermal protection panel is denoted as a first rigid insulation tile; and an intermediate layer mainly composed of a second rigid insulation tile; and an inner layer mainly composed of a rigid insulation tile / silica aerogel composite material; the outer layer is used to reduce the surface temperature of the thermal protection structure from T1 to T2; the intermediate layer is used to reduce T2 to T3; and the inner layer does not shrink at the T3 temperature. The present invention adopts structural design and heat transfer design, which can take into account the dual effects of gradient cooling and maintaining the stability of the overall structure, so that the mechanical toughness, high temperature resistance and thermal insulation properties of the thermal protection structure can meet the requirements of hypersonic aircraft.

Description

Technical Field

[0001] The present invention relates to the technical field of preparation of spacecraft thermal protection materials, and in particular to a surface-toughened gradient thermal protection structure, a preparation method thereof, and an application thereof. Background Art

[0002] Rigid thermal insulation tiles are a classic reusable thermal protection material developed during the US Space Shuttle program. Since the Space Shuttle program, spacecraft using rigid thermal insulation tiles as the primary thermal protection have included: ① NASA's six space shuttles: Columbia, Challenger, Discovery, Atlantis, Endeavour, and Enterprise; ② the Soviet Buran space shuttle; ③ Boeing's X-series hypersonic technology demonstration vehicles, including the X-37B, X-43, and X-51; ④ SNC's small space cargo aircraft, Dreamchaser and Tenacity; ⑤ Space-X's Starship; and ⑥ the side thermal protection of the Orion deep space exploration spacecraft.

[0003] Typical rigid thermal insulation tile models include: LI-900 and LI-2200 (U.S. Patent 3,952,083) developed by Lockheed Martin, HTP (Thermophysical and Mechanical Properties of the HTP Family of Rigid Ceramic Insulation Materials, AIAA-85-1055; U.S. Patent 5,629,186), FRCI (U.S. Patent 4,148,962) and AETB (Options for Improving Rigidized Ceramic Heatshields, Ceramic Engineering and Science Proceedings, pp. 757-768, 1985) developed by NASA Ames Research Center, BRI (U.S. Patent US 6,716,782 B2) developed by Boeing, and Space-X used on starship spacecraft. Rigid insulating tiles.

[0004] The domestic rigid insulation tile preparation technology includes: CN105272322B disclosed by the Institute of Aerospace Special Materials and Process Technology, CN102199042A disclosed by the Institute of Aerospace Materials and Technology, CN101691138A disclosed by Shandong Industrial Ceramics Research and Design Institute, and CN106946579A disclosed by Harbin Institute of Technology.

[0005] For reusable spacecraft, such as the space shuttle and aerospace plane, the rigid thermal insulation tiles used in the large-area thermal protection systems on their exterior surfaces generally require a high-emissivity coating. For example, the high-emissivity coatings developed by the United States for use on rigid thermal insulation tiles during the space shuttle manned flight program primarily include RCG (Reaction Cured Glass and Glass Coatings, U.S. Patent No. 4,093,771) and TUFI (Toughened Unipiece Fibrous Insulation, U.S. Patent No. 5,079,082). High-emissivity coatings applied to rigid thermal insulation tiles only penetrate 50 to 500 microns into the fiber matrix, resulting in poor resistance to airflow erosion and impacts from micrometeoroids and orbital debris (MMOD) in the space environment.

[0006] For hypersonic spaceplanes, which are primarily used for military purposes and shuttle between near-space and the Earth's atmosphere, their flight speed is ≥5Ma, and high-speed flight will generate aerodynamic heat flow. Therefore, it places requirements on thermal protection structures in many aspects: First, a high-reliability thermal protection structure with stronger mechanical properties, high infrared emissivity, and high temperature resistance is required to cope with the impact of battlefield explosive micro-flow sheets; second, a thermal protection structure with high dimensionality is required during service to ensure the aerodynamic shape of the aircraft and thus ensure the controllability of the flight attitude; third, a thermal protection structure with higher thermal insulation performance is required to avoid the impact of the aerodynamic heat flow generated by the high-speed flight on the hypersonic spaceplane.

[0007] Rigid thermal insulation tiles are the material of choice for hypersonic aircraft thermal protection systems due to their excellent high-temperature mechanical properties and dimensional stability. However, their high-temperature resistance, resistance to hard impact, and thermal insulation remain major obstacles hindering their expanded application in hypersonic aircraft. In particular, in the field of high-temperature resistance, NASA's Ames Research Center developed a surface-toughened, heat-insulating, integrated wing leading edge component (TUFROC, U.S. Patent No. 7,314,648) during the development of the X-37B aircraft. This replaces the reinforced carbon / carbon composite (RCC) used in the space shuttle era and serves as a low-cost wing leading edge component for the X-37B space plane, with a maximum operating temperature of 1,600°C. For example, an improved TUFROC material is described in US Patent US2021 / 0179299A1, which is used on two space cargo aircraft Dreamchaser and Tenacity of Sierra Nevada Corporation in the United States. Its maximum operating temperature reaches 1700°C. In addition, high-quality continuous ceramic fibers such as carbon fiber and silicon carbide fiber, such as C f / C、C f / SiC、SiC f High-strength hot-end ceramic-based composites such as sintered silicon carbide (SiC) also offer certain advantages in temperature resistance. However, these carbide ceramic-based composites suffer from disadvantages such as high manufacturing cost, long production cycles, and poor oxidation resistance. For example, US Patent 5,284,806 discloses a glass-doped ceramic-based composite, and US Patent 5,552,215 discloses a fiber-reinforced glass-based composite for use as a gas turbine engine compressor inner cover. These overcome the shortcomings of these ceramic-based composites in terms of manufacturing cost, production cycle, and oxidation resistance, but have yet to be applied in the field of hypersonic aircraft. To improve thermal insulation performance, US Patents 20020061396A1 and 6770584B2, respectively, disclose rigid insulation tile / silica aerogel composites. These composites have lower equivalent thermal conductivity than rigid insulation tiles, but due to the large specific surface area of ​​the aerogel structure (600-1000 m2 / g), they exhibit strong sintering activity and experience severe sintering shrinkage at temperatures of 1100°C and above.

[0008] Therefore, how to give full play to the advantages of various thermal protection materials, avoid their respective shortcomings, and prepare a thermal protection structure that fully utilizes rigid insulation tiles and meets the requirements of hypersonic aircraft in all aspects is a technical problem that needs to be solved urgently. Summary of the Invention

[0009] To address the shortcomings of existing thermal protection structures in meeting the requirements of hypersonic vehicles for space-to-earth travel, such as mechanical toughness, high-temperature resistance, and thermal insulation, the present invention provides a surface-toughened gradient thermal protection structure and its preparation method. Through structural design and heat transfer calculations, a thermal protection structure is designed that includes a high-strength fiber-reinforced glass matrix composite (FRGMC) panel, medium-temperature rigid insulation tiles, and low-temperature insulation tile / aerogel composite materials.

[0010] Specifically, in the first aspect, the present invention provides a rigid thermal insulation tile that is resistant to high temperature and impact, which is an ultrafine ceramic fiber rigid thermal insulation tile mainly composed of chopped quartz fiber, chopped mullite fiber, chopped aluminum borosilicate fiber, silicon carbide infrared sunscreen powder and boron nitride powder.

[0011] The short-cut aluminum borosilicate fibers are mainly composed of short-cut aluminum borosilicate fibers with a diameter distribution of 480 to 520 nm and short-cut aluminum borosilicate fibers with a diameter distribution of 1 to 3 μm.

[0012] The ultrafine ceramic fiber rigid insulation tile has a porosity of more than 50%, an average pore size of less than 40 μm, and a density of 1.0 g / cm 3 the following.

[0013] By using short-cut aluminum borosilicate fibers with a bimodal distribution of 480-520 nm and 1-3 μm, and adding the same mass fraction, the number of sintering nodes of aluminum borosilicate fibers, quartz fibers and mullite fibers in the rigid insulation tiles obtained is increased, which significantly improves the mechanical properties of the rigid insulation tiles, especially the high-temperature mechanical properties. The Nextel-312 fiber with a diameter of 2 to 4 μm (produced by 3M Company in the United States, U.S. Patent No. 3795524) used in rigid insulation tiles is more excellent.

[0014] According to the high-temperature and impact-resistant rigid thermal insulation tile provided by the present invention, the mass ratio of the short-cut borosilicate aluminum fibers with a diameter distribution of 480 to 520 nm and the short-cut borosilicate aluminum fibers with a diameter distribution of 1 to 3 μm is 1:3 to 5, preferably 1:4.

[0015] According to the high-temperature and impact-resistant rigid thermal insulation tile provided by the present invention, the raw materials for preparing the ultrafine ceramic fiber rigid thermal insulation tile include, by weight:

[0016] 250-300 parts of chopped quartz fiber;

[0017] 85-90 parts of short-cut mullite fiber;

[0018] 40-50 parts of chopped aluminum borosilicate fibers;

[0019] 5-10 parts corn starch;

[0020] 5-15 parts of silicon carbide infrared sunscreen powder;

[0021] 3 to 6 parts of boron nitride powder.

[0022] In a second aspect, the present invention further provides a rigid thermal insulation tile material that is resistant to high temperature and impact and has high emissivity, comprising the rigid thermal insulation tile as described above and a coating on its surface;

[0023] The coating is obtained by applying glass glaze on the surface of the rigid insulation tile, sintering and cooling;

[0024] The glass glaze comprises quartz glass powder, core-shell structured high silica glass powder, silicon tetraboride powder, molybdenum disilicide powder, tantalum disilicide powder, tantalum diboride powder and silicon carbide whiskers;

[0025] Preferably, the particle size D50 of the powder raw material constituting the glass glaze frit is 2 to 3 μm.

[0026] In order to further improve the performance of the high temperature and impact resistant rigid insulation tile, the present invention specifically develops the glass glaze (denoted as UltraEmit-01 TM The sintered coating has an infrared hemispherical full-spectrum emissivity greater than or equal to 0.80, a temperature resistance greater than or equal to 1200°C, and a lower average linear expansion coefficient than the aforementioned high-temperature, impact-resistant rigid insulation tile. This extremely low linear expansion coefficient ensures that the high-temperature, impact-resistant, high-emissivity rigid insulation tile material possesses excellent thermal shock resistance, capable of withstanding the tremendous instantaneous heat flux generated by aerospace vehicles during atmospheric re-entry, gliding, and tactical maneuvers.

[0027] In a third aspect, the present invention also provides a thermal protection panel for a hypersonic aircraft, comprising: a fiber reinforced glass matrix composite (FRGMC) obtained by impregnating glass slurry into pores in a rigid insulation tile, followed by drying and sintering.

[0028] The rigid thermal insulation tile is an ultrafine ceramic fiber rigid thermal insulation tile; the porosity of the ultrafine ceramic fiber rigid thermal insulation tile is more than 50%, and the average pore size is less than 40 μm.

[0029] The glass paste includes quartz glass powder, high silica glass powder with a core-shell structure, silicon tetraboride powder, molybdenum disilicide powder, tantalum disilicide powder, tantalum diboride powder and silicon carbide whiskers; the particle size of the powder raw materials constituting the glass paste is all submicron; after sintering, the glass paste is a continuous phase in the rigid insulation tile.

[0030] As mentioned above, when rigid thermal insulation tiles are used for thermal protection of hypersonic aircraft, they still have deficiencies in high temperature resistance and mechanical toughness. However, the pore structure of rigid thermal insulation tiles can be used to prepare composite materials to improve the mechanical toughness of rigid thermal insulation tiles. Based on this, the present invention found in experiments that by optimizing the pore structure of rigid thermal insulation tiles and optimizing the raw material structure of glass paste, especially controlling the powder raw material at the submicron level, it is possible to construct a continuous phase glass matrix in rigid thermal insulation tiles, and the fiber structure of the rigid thermal insulation tiles themselves can be used to toughen the glass matrix. At the same time, the silicon carbide whiskers in the glass paste also toughen the glass matrix at a more microscopic level, thereby obtaining a toughened high-temperature, impact-resistant, and high-emissivity FRGMC panel. The high temperature resistance and mechanical toughness of the panel can meet the thermal protection needs of hypersonic aircraft.

[0031] According to the thermal protection panel for hypersonic aircraft provided by the present invention, the D50 of the powder raw materials constituting the glass slurry is less than or equal to 1 μm, and the D90 is less than 1.5 μm.

[0032] As the glass paste infiltrates the pores of the rigid insulation tiles, the movement of the powdered raw materials that comprise it is affected by the fibers within the tiles. While smaller particle sizes improve infiltration efficiency, nanoscale powders increase production costs. The present inventors have discovered that using the aforementioned powdered raw material structure, combined with the pore size of the rigid insulation tiles, can achieve the mechanical toughness and high-temperature resistance requirements of the thermal protection panels of hypersonic aircraft.

[0033] According to the thermal protection panel for hypersonic aircraft provided by the present invention, the aspect ratio L / D of the silicon carbide whiskers is in the range of 50≤L / D≤100.

[0034] When selecting silicon carbide whiskers of appropriate size, it is necessary to consider both that the silicon carbide whiskers must have a sufficient aspect ratio to play a toughening role and that if the aspect ratio is too large, they will be easily filtered by the fibers in the ceramic fiber rigid insulation tile and cannot penetrate.

[0035] According to the thermal protection panel for hypersonic aircraft provided by the present invention, the glass slurry further includes a ceramic dispersant.

[0036] Preferably, the ceramic dispersant is one or a combination of two or more of Darvan 821A ceramic dispersant, Tween-80, sodium polyacrylate, acrylamide, polyacrylic acid, polymethacrylic acid, methyl cellulose, polymethacrylate ammonium salt with a pH value of 7, ammonia water, polyvinyl alcohol, polyethylene glycol, and polyperfluorosulfonic acid.

[0037] More preferably, the ceramic dispersant is Darvan 821A ceramic dispersant.

[0038] For the thermal protection panel prepared by immersing glass slurry into the pores of rigid insulation tiles, the suspension stability of the glass slurry plays an important role in the structural uniformity and performance improvement of the thermal protection panel. Preferably, a ceramic dispersant is added to the glass slurry to stabilize the glass slurry so that it is in a suspended state during the immersion process.

[0039] The present invention adopts Darvan 821A ceramic dispersant. By adding the ceramic dispersant during the preparation of the glass slurry, the obtained glass slurry can maintain a good suspension state when it is left to stand for more than two weeks.

[0040] According to the thermal protection panel for hypersonic aircraft provided by the present invention, the rigid thermal insulation tile is the ultrafine ceramic fiber rigid thermal insulation tile as described above.

[0041] The ultrafine ceramic fiber rigid insulation tile has a low porosity and a pore size of micron level. The pores are interconnected, so that the glass slurry can fully penetrate the pores to form the continuous phase glass matrix. The overall structure of the ultrafine ceramic fiber rigid insulation tile is also retained. It works synergistically with the glass matrix to obtain a heat protection panel with high compressive strength and high surface hardness. Compared with the existing C f / C、C f / SiC、SiC f Thermal protection panels made of ceramic-based composite materials such as SiC have higher mechanical strength and reliability, as well as lower manufacturing costs and cycles.

[0042] According to the thermal protection panel for hypersonic aircraft provided by the present invention, the fiber-reinforced glass-based composite material has a hemispherical full-spectrum emissivity of infrared radiation greater than or equal to 0.9, a reusable temperature greater than or equal to 1600°C, and an impact strength greater than or equal to 10J.

[0043] In a fourth aspect, the present invention also provides a method for preparing the thermal protection panel for a hypersonic aircraft as described above, comprising: preparing the fiber-reinforced glass-based composite material by using the rigid insulation tile through one or more impregnation-drying-sintering processes.

[0044] The impregnation-drying-sintering process comprises: impregnating glass slurry in a stable suspension state into the pores of the rigid insulation tile under a vacuum environment to obtain a wet blank; and drying and sintering the wet blank.

[0045] Taking the preparation of the fiber-reinforced glass-based composite material by a two-step impregnation-drying-sintering process as an example, the preparation process of the fiber-reinforced glass-based composite material is as follows:

[0046] The first impregnation-drying-sintering process: the glass paste in a stable suspension state is impregnated into the pores of the rigid insulation tile under a vacuum environment to obtain a wet blank; the wet blank is dried and sintered to obtain the first composite material;

[0047] Second impregnation-drying-sintering process: the glass slurry in a stable suspension state is impregnated into the pores of the first composite material under a vacuum environment to obtain a wet blank; the wet blank is dried and sintered to obtain a fiber-reinforced glass-based composite material.

[0048] In the present invention, the glass slurry in a stable suspension state means that the glass slurry does not delaminate when stored for more than 7 days under normal temperature and pressure conditions.

[0049] In order to obtain a thermal protection panel with better mechanical toughness and temperature resistance, it is very important to maintain the overall structural stability of the rigid insulation tile and to make the glass slurry form a continuous phase. The present invention proposes to immerse the glass slurry into the pores of the rigid insulation tile under a vacuum environment. The operation is simple and can achieve the above-mentioned purpose.

[0050] Preferably, the pressure under the vacuum environment is less than or equal to 5 kPa.

[0051] In a fifth aspect, the present invention further provides a gradient thermal protection structure for a hypersonic aircraft, comprising: an outer layer, the outer layer mainly consisting of the thermal protection panel as described above; the rigid thermal insulation tile used in the thermal protection panel is denoted as a first rigid thermal insulation tile;

[0052] and, an intermediate layer, said intermediate layer being primarily composed of a second rigid insulation tile;

[0053] And, an inner layer, wherein the inner layer is mainly composed of a rigid thermal insulation tile / silica aerogel composite material obtained by supercritical drying the third rigid thermal insulation tile composite silica aerogel.

[0054] The outer layer is used to reduce the surface temperature of the heat protection structure from T1 to T2; the middle layer is used to reduce T2 to T3; and the inner layer does not shrink at the temperature T3.

[0055] In the thermal protection structures used for hypersonic aircraft, there has been a proposal to use a multi-layer combination of high-density rigid insulation tiles and low-density rigid insulation tiles to improve the insulation tiles' erosion resistance and temperature resistance. However, this simple multi-layer combination can only achieve limited improvements. More importantly, the multi-layer combination of rigid insulation tiles does not significantly improve thermal insulation performance. Although the use of rigid insulation tile / silica aerogel composite materials with very low equivalent thermal conductivity as part of a thermal protection structure has great potential, the rigid insulation tile / silica aerogel composite material undergoes severe sintering shrinkage at temperatures of 1100°C and above, making the entire thermal protection structure unstable and unsuitable for thermal protection of hypersonic aircraft.

[0056] In order to prepare a thermal protection structure that meets the requirements of hypersonic aircraft applications, the present invention designs a gradient thermal protection structure based on a layer structure, with the toughened fiber reinforced glass matrix composite (FRGMC) panel as the outer layer, the rigid insulation tile as the middle layer, and the rigid insulation tile / silica aerogel composite material as the inner layer. Among them, the main feature is that the thermal protection panel (outer layer) has an extremely high infrared emissivity, which can re-radiate most of the incident heat flux back to the background space. At the same time, the thermal protection panel and the middle layer are used in combination to achieve gradient cooling. When heat is transferred to the surface of the inner layer, the temperature is reduced, avoiding the shrinkage of the inner layer structure, thereby fully utilizing the excellent mechanical toughness and high temperature resistance of the outer layer and the low equivalent thermal conductivity of the inner layer in the present invention.

[0057] According to the gradient thermal protection structure for hypersonic aircraft provided by the present invention, the hemispherical full-spectrum emissivity of the outer layer to infrared radiation is greater than or equal to 0.9; the apparent thermal conductivity of the middle layer at 1200°C is less than 0.20 W / (m·K).

[0058] According to the gradient thermal protection structure for hypersonic aircraft provided by the present invention, the second rigid thermal insulation tile and / or the third rigid thermal insulation tile are the ultrafine ceramic fiber rigid thermal insulation tiles as described above.

[0059] According to the gradient thermal protection structure for a hypersonic aircraft provided by the present invention, the density of the first rigid thermal insulation tile is higher than the density of the second rigid thermal insulation tile; and / or the density of the second rigid thermal insulation tile is the same as the density of the third rigid thermal insulation tile.

[0060] Preferably, the rigid thermal insulation tiles of different densities in the present invention can be prepared by similar preparation methods, all of which refer to the method disclosed in invention patent CN105272322B.

[0061] According to the gradient thermal protection structure for a hypersonic aircraft provided by the present invention, the outer layer, the middle layer and the inner layer are mechanically interlocked with each other using dovetail grooves.

[0062] Dovetail groove mechanical interlocking offers numerous technical advantages. This invention utilizes its lack of freedom in the thickness direction, making installation easy. More importantly, experiments have shown that the three layers of the thermal protection structure have well-matched linear expansion coefficients, making it ideal for dovetail groove interlocking, eliminating the need for adhesives.

[0063] The present invention may also provide a ceramic-based high-temperature resistant adhesive between the dovetail groove interfaces to further stabilize the interface connection, such as GC-002 alumina-based high-temperature ceramic adhesive produced by Zircar Zirconia.

[0064] Preferably, the direction of the dovetail groove connecting the middle layer and the outer layer is perpendicular to the direction of the dovetail groove connecting the middle layer and the inner layer.

[0065] According to the gradient thermal protection structure for hypersonic aircraft provided by the present invention, the thickness of the thermal protection panel is 5 to 15 mm, the thickness of the second rigid thermal insulation tile is 20 to 30 mm, and the total thickness of the gradient thermal protection structure is 45 to 75 mm.

[0066] Preferably, the thickness of the heat protection panel is 10 mm, and the thickness of the second rigid insulation tile is 20 mm.

[0067] In a sixth aspect, the present invention also provides the use of the high-temperature and impact-resistant rigid thermal insulation tiles as described above, the high-temperature and impact-resistant high-emissivity rigid thermal insulation tile materials as described above, the thermal protection panels as described above, or the gradient thermal protection structures as described above in hypersonic aircraft.

[0068] The present invention provides a surface-reinforced gradient thermal protection structure, its preparation method and application, by designing a rigid thermal insulation tile with a special structure, and using the rigid thermal insulation tile as a substrate, designing a reinforced high-emissivity ceramic fiber reinforced glass-based composite material, and further using the reinforced high-emissivity ceramic fiber reinforced glass-based composite material as a thermal protection panel, and compounding it with a rigid thermal insulation tile with a specific structure and a rigid thermal insulation tile / silica aerogel composite material to form a multilayer structure, so that when the multilayer structure is used as a thermal protection structure for a hypersonic aircraft, its mechanical strength, high temperature resistance and thermal insulation performance all meet the requirements. BRIEF DESCRIPTION OF THE DRAWINGS

[0069] In order to more clearly illustrate the technical solutions in the present invention or the prior art, a brief introduction is given below to the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without paying any creative work.

[0070] Figure 1 It is a structural schematic diagram of the gradient thermal protection structure in Example 6 provided by the present invention.

[0071] Figure 2 This is a physical picture of the rigid thermal insulation tile in Example 2 provided by the present invention.

[0072] Figure 3 This is a microstructure diagram of the rigid thermal insulation tile in Example 2 provided by the present invention.

[0073] Figure 4 This is a physical picture of the high-temperature-resistant, impact-resistant, and high-emissivity rigid thermal insulation tile material prepared in Example 5 provided by the present invention.

[0074] Figure 5 This is a physical picture of the rigid thermal insulation tile material prepared in Comparative Example 1 provided by the present invention.

[0075] Figure 6 This is a SEM image of the rigid thermal insulation tile / silica aerogel composite material prepared in Example 4 provided by the present invention, with a magnification of 1000 times.

[0076] Figure 7 This is a partial enlarged view of the silica aerogel in the rigid thermal insulation tile / silica aerogel composite material prepared in Example 4 provided by the present invention, with a magnification of 50,000 times.

[0077] Figure 8 The present invention provides a blunt-headed surface-strengthened gradient thermal protection wing leading edge sample developed for a hypersonic space plane with a design speed of Ma=6.

[0078] Figure 9 It is the equivalent thermal conductivity of the rigid insulation tile in Example 2, the FRGMC panel prepared in Example 3, and the rigid insulation tile / silica aerogel composite material prepared in Example 4 provided by the present invention.

[0079] Figure 10 This is the temperature response at different depths of the hypersonic wind tunnel test of the gradient thermal protection structure in Example 6 provided by the present invention. DETAILED DESCRIPTION

[0080] To make the objectives, technical solutions, and advantages of the present invention more clear, the technical solutions of the present invention will be clearly and completely described below in conjunction with the accompanying drawings. Obviously, the embodiments described are only some of the embodiments of the present invention, not all of them. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative efforts shall fall within the scope of protection of the present invention.

[0081] The following combination Figures 1-10 The present invention describes a surface toughened gradient thermal protection structure and a preparation method thereof.

[0082] In an embodiment of the present invention, a high-temperature and impact-resistant rigid insulation tile is first provided, which is an ultrafine ceramic fiber rigid insulation tile mainly composed of chopped quartz fiber, chopped mullite fiber, chopped aluminum borosilicate fiber, silicon carbide infrared shading agent powder and boron nitride powder.

[0083] The short-cut aluminum borosilicate fibers are mainly composed of short-cut aluminum borosilicate fibers with a diameter distribution of 480 to 520 nm and short-cut aluminum borosilicate fibers with a diameter distribution of 1 to 3 μm.

[0084] The ultrafine ceramic fiber rigid insulation tile has a porosity of more than 50%, an average pore size of less than 40 μm, and a density of 1.0 g / cm 3 the following.

[0085] In some embodiments of the present invention, the mass ratio of the chopped aluminoborosilicate fibers with a diameter distribution of 480 to 520 nm to the chopped aluminoborosilicate fibers with a diameter distribution of 1 to 3 μm is 1:3 to 5, preferably 1:4.

[0086] In some embodiments of the present invention, the raw materials for preparing the ultrafine ceramic fiber rigid insulation tile include, by weight:

[0087] 250-300 parts of chopped quartz fiber;

[0088] 85-90 parts of short-cut mullite fiber;

[0089] 40-50 parts of chopped aluminum borosilicate fibers;

[0090] 5-10 parts corn starch;

[0091] 5-15 parts of silicon carbide infrared sunscreen powder;

[0092] 3 to 6 parts of boron nitride powder.

[0093] In an embodiment of the present invention, there is also provided a rigid thermal insulation tile material with high temperature resistance, impact resistance and high emissivity, comprising the rigid thermal insulation tile as described above and a coating on its surface;

[0094] The coating is obtained by applying glass glaze on the surface of the rigid insulation tile, sintering and cooling;

[0095] The glass glaze material comprises quartz glass powder, high silica glass powder with a core-shell structure, silicon tetraboride powder, molybdenum disilicide powder, tantalum disilicide powder, tantalum diboride powder and silicon carbide whiskers.

[0096] In some embodiments of the present invention, the particle size D50 of the powder raw material constituting the glass glaze frit is 2-3 μm.

[0097] In some embodiments of the present invention, the solid content of the glass frit is 20-40%.

[0098] In some embodiments of the present invention, the dispersion medium of the glass frit is water.

[0099] In some embodiments of the present invention, the sintering temperature of the coating is 1000-1200°C.

[0100] In some embodiments of the present invention, the glass glaze is applied to the surface of the rigid insulation tile by scraping or spraying.

[0101] In some embodiments of the present invention, the raw materials for preparing the glass enamel include, by weight:

[0102] 20-30 parts of quartz glass powder;

[0103] 20-30 parts of highly reactive core-shell structure high silica glass powder;

[0104] 2-3 parts of silicon tetraboride powder;

[0105] 0.5-1.5 parts of boron carbide powder;

[0106] 0.1-2 parts of alkaline silica sol;

[0107] 15-25 parts of molybdenum disilicide powder;

[0108] 20-25 parts of tantalum disilicide powder;

[0109] 2-3 parts of tantalum diboride powder;

[0110] 2-3 parts of silicon carbide whiskers;

[0111] 1 to 3 parts of ceramic dispersant.

[0112] In an embodiment of the present invention, a FRGMC thermal protection panel for a hypersonic aircraft is also provided, comprising: a fiber-reinforced glass-based composite material obtained by impregnating glass slurry into the pores of a ceramic fiber rigid insulation tile, followed by drying and sintering.

[0113] The rigid thermal insulation tile is an ultrafine ceramic fiber rigid thermal insulation tile; the porosity of the ultrafine ceramic fiber rigid thermal insulation tile is more than 50%, and the average pore size is less than 40 μm.

[0114] The glass paste includes quartz glass powder, high silica glass powder with a core-shell structure, silicon tetraboride powder, molybdenum disilicide powder, tantalum disilicide powder, tantalum diboride powder and silicon carbide whiskers; the particle size of the powder raw materials constituting the glass paste is all submicron; after sintering, the glass paste is a continuous phase in the rigid insulation tile.

[0115] In some embodiments of the present invention, the powder raw materials constituting the glass paste have a D50 less than or equal to 1 μm, and a D90 less than 1.5 μm.

[0116] In some embodiments of the present invention, the aspect ratio L / D of the silicon carbide whiskers is in the range of 50≤L / D≤100.

[0117] In some embodiments of the present invention, the glass slurry further includes a ceramic dispersant.

[0118] In some preferred embodiments of the present invention, the ceramic dispersant is one or a combination of two or more of Darvan821A, Tween-80, sodium polyacrylate, acrylamide, polyacrylic acid, polymethacrylic acid, methyl cellulose, polyacrylic acid ammonium salt with a pH value of 7, polymethacrylic acid ammonium salt with a pH value of 7, ammonia water, polyvinyl alcohol, polyethylene glycol, and polyperfluorosulfonic acid.

[0119] In some preferred embodiments of the present invention, the ceramic dispersant is a 20 wt % aqueous solution of polyacrylate ammonium salt with a pH of 7.

[0120] In some embodiments of the present invention, the preparation of the glass slurry of the present invention includes: charging various raw materials into a planetary ball mill and ball milling them.

[0121] For the fiber reinforcement phase of rigid thermal insulation tiles with different fiber compositions, it can be added to the glass paste Glass powder, Glass powder, silicon hexaboride powder, nano-alumina sol and other materials are used to adjust the sintering temperature of the FRGMC panel and the thermal expansion properties of the glass phase.

[0122] In some embodiments of the present invention, the solid content of the glass paste is 50-80 wt %.

[0123] In some embodiments of the present invention, the raw materials for preparing the glass paste include, by weight:

[0124] 20-30 parts of quartz glass powder;

[0125] 20-30 parts of highly reactive core-shell structure high silica glass powder;

[0126] 2-3 parts of silicon tetraboride powder;

[0127] 0.5-1.5 parts of boron carbide powder;

[0128] 0.1-2 parts of alkaline silica sol;

[0129] 15-25 parts of molybdenum disilicide powder;

[0130] 20-25 parts of tantalum disilicide powder;

[0131] 2-3 parts of tantalum diboride powder;

[0132] 2-3 parts of silicon carbide whiskers;

[0133] 1 to 3 parts of ceramic dispersant.

[0134] In some embodiments of the present invention, the rigid thermal insulation tile is the ultrafine ceramic fiber rigid thermal insulation tile described above.

[0135] In some embodiments of the present invention, the fiber-reinforced glass-based composite material has a hemispherical full-spectrum emissivity of infrared radiation greater than or equal to 0.9, a reusable temperature greater than or equal to 1600° C., and an impact strength greater than or equal to 10J.

[0136] An embodiment of the present invention also provides a method for preparing the thermal protection panel for a hypersonic aircraft as described above, comprising: preparing the fiber-reinforced glass-based composite material using a rigid insulation tile through one or more impregnation-drying-sintering processes; the impregnation-drying-sintering process comprises: impregnating a glass slurry in a stable suspension state into the pores of the rigid insulation tile under a vacuum environment to obtain a wet blank; and drying and sintering the wet blank.

[0137] In some embodiments of the present invention, the pressure in the vacuum environment is less than or equal to 5 kPa.

[0138] In some embodiments of the present invention, the method for preparing the heat protection panel comprises the following steps:

[0139] Place the ceramic fiber rigid insulation tile in a vacuum impregnation tank, turn on the vacuum pump, evacuate the tank until the pressure is ≤5kPa, close the vacuum pump outlet valve, and suck the glass slurry into the tank. The liquid level of the glass slurry submerges the upper surface of the rigid insulation tile. After soaking for a certain period of time, take it out to obtain a panel wet blank; the panel wet blank is dried and sintered at 110-130°C to obtain the thermal protection panel.

[0140] Preferably, the sintering is carried out in a muffle furnace, and the sintering procedure is:

[0141] The first stage: the muffle furnace room temperature is raised to 1100-1200℃, 2-5℃ / min.

[0142] The second stage: placing the panel blank obtained by drying at 1100-1200℃ into a muffle furnace.

[0143] The third stage: heating to 1250-1400℃ at 2-5℃ / min.

[0144] The fourth stage: keep warm at 1250-1400℃ for 20-60min.

[0145] The fifth stage: cooling down to 1100-1200℃ at 3-8℃ / min.

[0146] The sixth stage: 1100-1200℃, open the furnace door, take out and rapidly cool to room temperature in the air.

[0147] An embodiment of the present invention further provides a gradient thermal protection structure for a hypersonic aircraft, comprising: an outer layer, the outer layer mainly consisting of the thermal protection panel as described above; the rigid thermal insulation tile used in the thermal protection panel is denoted as a first rigid thermal insulation tile;

[0148] and, an intermediate layer, said intermediate layer being primarily composed of a second rigid insulation tile;

[0149] And, an inner layer, wherein the inner layer is mainly composed of a rigid thermal insulation tile / silica aerogel composite material obtained by supercritical drying the third rigid thermal insulation tile composite silica aerogel.

[0150] The outer layer is used to reduce the surface temperature of the heat protection structure from T1 to T2; the middle layer is used to reduce T2 to T3; and the inner layer does not shrink at the temperature T3.

[0151] In some embodiments of the present invention, the hemispherical full-spectrum emissivity of the outer layer to infrared radiation is greater than or equal to 0.9; and the apparent thermal conductivity of the middle layer at 1200° C. is less than 0.20 W / (m·K).

[0152] In some embodiments of the present invention, the second rigid thermal insulation tile and / or the third rigid thermal insulation tile is the ultrafine ceramic fiber rigid thermal insulation tile as described above.

[0153] In some embodiments of the present invention, the density of the first rigid insulation tile is higher than the density of the second rigid insulation tile; and / or the density of the second rigid insulation tile is the same as the density of the third rigid insulation tile.

[0154] In some embodiments of the present invention, the density of the first rigid thermal insulation tile is 0.3 to 0.5 g / cm higher than the density of the second rigid thermal insulation tile. 3 .

[0155] In some embodiments of the present invention, the apparent density of the second rigid thermal insulation tile is less than 0.3 g / cm3.

[0156] In some embodiments of the present invention, the method for preparing the rigid thermal insulation tile / silica aerogel composite material comprises the following steps:

[0157] (1) Preparation of silica sol.

[0158] (1.1) Mix ethyl orthosilicate, water, and anhydrous ethanol to obtain a mixed solution.

[0159] Preferably, the molar ratio of ethyl orthosilicate, water and anhydrous ethanol is 1:4-8:6-10.

[0160] (1.2) Adjust the pH value of the mixed solution to 2-3 with hydrochloric acid and stir at room temperature. During this process, the ethyl orthosilicate is fully hydrolyzed.

[0161] (1.3) Continue to add ammonia water to adjust the pH value of the mixed solution to neutral and stir evenly. During this process, the ammonia water should be added slowly and drop by drop. Adding a large amount will cause the local pH value to be too high and form flocculation.

[0162] (1.4) After step (1.3) is completed, continue stirring, and then add ammonium fluoride aqueous solution to obtain SiO2 sol for use.

[0163] In the present invention, the silica sol used to form the silica aerogel includes but is not limited to the silica sol prepared by the above-mentioned preparation process, and may also include: for example, the green silica aerogel preparation process using Si-40 polysiloxane, tetrabutylammonium fluoride and 1-ethanolamine as precursors published in accordance with Chinese invention patent authorization publication No. CN117303381B.

[0164] (2) Preparation of rigid thermal insulation tile / SiO2 wet gel composite materials.

[0165] (2.1) Place a piece of ultrafine ceramic fiber rigid insulation tile in a vacuum impregnation tank and pump the air pressure in the tank to an absolute pressure of ≤5kPa.

[0166] (2.2) The SiO2 sol prepared in step (1) is drawn into a vacuum impregnation tank. The sol level is at least 15% higher than the thickness of the insulation tile. Because the sol shrinks by approximately 5-10% during gelation, a larger amount of sol is required to ensure that the rigid insulation tile is fully impregnated in the thickness direction. The pipe connected to the vacuum pump is connected to a compressed air cylinder, and compressed air at 1-2 MPa is filled into the tank. The positive pressure on the gel surface during the gelation process is conducive to the formation of an aerogel with a more regular microstructure and better mechanical strength.

[0167] (2.3) After step (2.2) is completed, wait for the sol to autogel. The gelation time of this autogel is strongly correlated with the amount of ammonium fluoride catalyst added in step (1.4). When the concentration of the added ammonium fluoride aqueous solution is 1 mol / L, and the amounts of other materials in the experiment remain unchanged, the corresponding relationship between the amount of 1 mol / L ammonium fluoride aqueous solution and the gelation time is as follows:

[0168] Table 1

[0169] Mass ratio of 1 mol / L ammonium fluoride aqueous solution to water Gel time 0.5：9 20h, with an operating time window, acceptable 1：9 30s, too fast, no injection operation time window 0.2：9 34h, with an operating time window, acceptable <0.1∶9 No gel

[0170] (2.4) After the gel has aged at room temperature for 24 hours, open the lid of the vacuum impregnation tank, remove any residual liquid from the gel surface, and inject anhydrous ethanol until the gel surface is completely submerged. After a certain period of time, the ethanol in the tank is pumped out and fresh anhydrous ethanol is poured in to soak the composite material.

[0171] The solvent exchange process is designed to remove water from the pores of the wet gel through diffusion, driven by concentration differences. Water is not removed during supercritical drying, so it is essential to completely remove water from the gel pores before the supercritical drying process. Because water diffusion in nanopores is slow, aerogel products are typically thin. Extremely thick aerogel products require multiple solvent exchange cycles to remove water from the deep pores.

[0172] In addition, the aging temperature can be increased during aging, such as aging at 50°C, which can accelerate the strengthening process of the gel skeleton.

[0173] (3) Supercritical drying: Place the soaked composite material in a supercritical carbon dioxide reactor. Add liquid carbon dioxide into the reactor so that the liquid surface completely submerges the upper surface of the wet gel. Control the temperature in the reactor to 40-80°C and the pressure to 8-12 MPa. After 25-35 minutes, release the pressure in the reactor at a rate of 20-30 kPa / min.

[0174] (4) Hydrophobic treatment: The composite material obtained in step (3) is placed in a vacuum tank, the absolute pressure in the tank is evacuated to ≤1 kPa, methyltrimethoxysilane and glacial acetic acid are introduced into the tank, the vacuum tank is heated to 60-80°C, kept at a constant temperature for 2 hours, and then cooled to room temperature. The vacuum tank is opened to obtain a hydrophobic insulating tile / aerogel composite material.

[0175] (5) machining the hydrophobic composite material obtained in step (4) into the shape and size required by the design to obtain a rigid thermal insulation tile / silica aerogel composite material.

[0176] Furthermore, the second rigid thermal insulation tile, which serves as the intermediate layer, is also preferably treated with the aforementioned hydrophobic treatment to impart inherent hydrophobicity to the rigid thermal insulation tile. As a thermal protection material for the exterior surface of a hypersonic aircraft, hydrophobicity is essential to ensure that the thermal protection system does not absorb rainwater, which would increase its weight and reduce or even eliminate its thermal protection performance in rainy weather.

[0177] In some embodiments of the present invention, the outer layer, the middle layer and the inner layer are mechanically interlocked with each other using dovetail grooves.

[0178] In some embodiments of the present invention, the direction of the dovetail groove connecting the middle layer and the outer layer is perpendicular to the direction of the dovetail groove connecting the middle layer and the inner layer.

[0179] In some embodiments of the present invention, the thickness of the heat protection panel is 5-15 mm, the thickness of the second rigid insulation tile is 20-30 mm, and the total thickness of the gradient heat protection structure is 45-75 mm.

[0180] In some embodiments of the present invention, the thickness of the heat protection panel is 10 mm, and the thickness of the second rigid insulation tile is 20 mm.

[0181] The present invention also provides the application of the rigid thermal insulation tile resistant to high temperature and impact as described above, the thermal protection panel as described above, or the gradient thermal protection structure as described above in a hypersonic aircraft. Figure 8 As shown, this is a blunt-nosed surface-toughened gradient thermal protection wing leading edge sample developed by the present invention for a hypersonic space plane with a design speed of Ma=6.

[0182] Where specific techniques or conditions are not specified in the examples of the present invention, the techniques or conditions described in the literature in the field or in accordance with the product specifications were used. For devices, instruments, reagents, etc. used, where the manufacturer is not specified, all are conventional products that can be purchased through regular channels. The raw materials used in the present invention are all readily available in the domestic product market.

[0183] The chopped quartz fibers of the present invention are prepared according to the method disclosed in Chinese invention patent CN110424067A, and have a diameter of 1 to 3 microns and a chopped length of 4 to 6 mm.

[0184] The chopped mullite fiber of the present invention is prepared according to the method disclosed in Chinese invention patent CN110846741B, has a diameter of 2 to 4 microns, and a chopped length of 4 to 6 mm.

[0185] The corn starch of the present invention is purchased from Sinopharm Chemical Reagent Co., Ltd., with a purity of more than 99.5%.

[0186] The silicon carbide infrared sunscreen powder of the present invention is purchased from Beijing Micro-Nano Ultrafine Materials Co., Ltd., and has an α phase and a D50 of 3.5 μm.

[0187] The boron nitride powder of the present invention is purchased from Binhai Liaobin Fine Chemical Co., Ltd., has a D50 of 2 μm, and is a hexagonal phase.

[0188] The deionized water of the present invention is prepared by a laboratory pure water machine, and has a resistivity of ≥18MΩ.cm.

[0189] The concentrated ammonia water of the present invention is purchased from Bailingwei Chemical Reagent Co., Ltd. and contains 25% to 28% ammonia.

[0190] The Darvan 821A ceramic dispersant of the present invention is purchased from RT Vanderbilt Company, and its functional component is an aqueous solution of polyacrylate ammonium salt.

[0191] The quartz glass powder of the present invention is purchased from Pacific Quartz Co., Ltd., with a purity of ≥99.9%, D50=1 μm, and D90 less than 1.5 μm.

[0192] The highly reactive core-shell high silica glass powder of the present invention was purchased from Youyan Engineering Technology Research Institute Co., Ltd. The preparation method refers to Chinese invention patent CN113045206B, D50=1 μm, and D90 is less than 1.5 μm.

[0193] The silicon tetraboride powder of the present invention was purchased from Youyan Engineering Technology Research Institute Co., Ltd., and the preparation method was based on the invention patent CN114538457B, with D50=1 μm and D90 less than 1.5 μm.

[0194] The boron carbide powder of the present invention is purchased from Forsman (Beijing) Technology Co., Ltd., with D50=1 μm and D90 less than 1.5 μm.

[0195] The particle size of the molybdenum disilicide powder of the present invention is D50=1 μm, and D90 is less than 1.5 μm.

[0196] The tantalum disilicide powder of the present invention was purchased from Youyan Engineering Technology Research Institute Co., Ltd., and the preparation method was based on the invention patent CN114538450A, with D50=1 μm and D90 less than 1.5 μm.

[0197] The tantalum diboride powder of the present invention is purchased from Forsman (Beijing) Co., Ltd., with D50=1 μm and D90 less than 1.5 μm.

[0198] The silicon carbide whisker of the present invention has an α phase, D50=1 μm, and 50≤L / D≤100.

[0199] The pH value of the 30wt% high-purity alkaline silica sol of the present invention is 10.

[0200] The structure of the sintering kiln furniture of the present invention may adopt the structure disclosed in patent CN106403606B.

[0201] The testing method of the present invention comprises:

[0202] Field emission scanning electron microscope (FE-SEM): LEO-1530, Zeiss, Germany.

[0203] The infrared hemisphere full spectrum emissivity is tested by the National Infrared Products Quality Supervision and Inspection Center.

[0204] The equivalent thermal conductivity of different materials is tested according to the GJB 10252-2021 standard.

[0205] The impact resistance is tested in accordance with GB / T 39814-2021 "Ultra-thin glass impact strength test method - Drop ball impact method".

[0206] Example 1 Density 0.5g / cm 3 Thermoshield TM Ultrafine ceramic fiber rigid insulation tile

[0207] A method for preparing rigid thermal insulation tiles, the preparation process of which refers to the method disclosed in invention patent CN105272322B, and the specific steps are as follows:

[0208] (1) Raw material preparation

[0209] Table 2

[0210]

[0211] (2) Beating: Boron nitride powder as a sintering aid, silicon carbide infrared shading agent powder, corn starch, deionized water, concentrated ammonia water and Darvan 821A ceramic dispersant are mixed uniformly to obtain a suspension, and then short-cut quartz fiber, short-cut mullite fiber and short-cut aluminum borosilicate fiber are added to the suspension and beaten to obtain a ceramic fiber slurry with a beating degree of 20 to 50°SR.

[0212] (3) Wet blank forming: The ceramic fiber slurry is transferred to a filtering tool, and most of the water is quickly filtered out under vacuum assistance. The resulting filter cake is further filtered to a wet blank of designed thickness. The pressing process further orients the fibers in the blank, resulting in a higher degree of orientation.

[0213] (4) Drying of wet blanks: The wet blanks are transferred to a limited clamping fixture and placed in a blast drying oven for drying. The drying procedure is: first keep warm at 80°C for 2 hours, then keep warm at 120°C for ≥24 hours, preferably ≥48 hours, and more preferably ≥72 hours.

[0214] During the drying process, the starch undergoes gelatinization and cross-linking reactions, and is further dehydrated to form a continuous bonding phase, so that the dried blank remains in a complete block shape and is easy to transfer to the sintering kiln.

[0215] (5) Pressure sintering: Transfer the dry blank to the sintering kiln and sinter according to the following heating program:

[0216] The first stage: room temperature to 550℃, 3℃ / min.

[0217] The second stage: 550℃, keep warm for 30min.

[0218] The third section: 550℃ to 1280℃, 3℃ / min.

[0219] The fourth stage: keep warm at 1280℃ for 40min.

[0220] The fifth stage: the sintering process is terminated and the temperature is naturally cooled to room temperature.

[0221] During the sintering process, a silicon carbide plate is placed on the upper surface of the insulation tile body for pressure, and silicon carbide positioning pillars are placed on the side according to the designed thickness. Maintaining positive pressure on the body during sintering helps to sinter more thoroughly.

[0222] After sintering, the porosity was 79.2%, the average pore size was 35 μm, and the density was 0.5 g / cm 3 , rigid insulation tiles with length, width and height of 400mm×400mm×50mm.

[0223] Example 2 Density 0.24 g / cm 3 Thermoshield TM Ultrafine ceramic fiber rigid insulation tile

[0224] A method for preparing a rigid thermal insulation tile, wherein the preparation steps are basically the same as those in Example 1, except that the amount of various raw materials is reduced to 48% of that in (1), and the obtained density is 0.24 g / cm 3 , rigid insulation tile with a length, width and height of 400mm×400mm×50mm, recorded as ThermoShield TM -240, the actual picture is as follows Figure 2 As shown in the microstructure diagram Figure 3 As shown, the equivalent thermal conductivity is Figure 9 shown.

[0225] ThermoShield TM The main thermophysical performance indicators of -240 compared with LI-900 and HTP-12 are shown in the following table:

[0226] Table 3

[0227]

[0228] Example 3 FRGMC panel

[0229] A method for preparing a FRGMC panel, comprising the following steps:

[0230] (1) The materials shown in the following table were charged into a planetary ball mill according to their weight, milled at 500 rpm for 72 h, and filtered through a 600-mesh sieve to obtain a glass slurry.

[0231] Table 4

[0232] Substance name weight Quartz glass powder 25kg Highly reactive core-shell structure high silica glass powder 25kg Silicon tetraboride powder 2.5kg Boron carbide powder 1.0kg 30wt% high purity alkaline silica sol 2.0kg Molybdenum disilicide powder 20kg Tantalum disilicide powder 22.5kg Tantalum diboride powder 2.5kg Silicon carbide whiskers 2.5kg Deionized water 100kg Darvan821A ceramic dispersant 2kg

[0233] (2) Thermoshield with a density of 0.5 g / cm3 obtained in Example 1 was TM Place the ultrafine ceramic fiber rigid insulation tile (as the fiber base material for FRGMC panel) in a vacuum impregnation tank, turn on the vacuum pump, evacuate the tank to a pressure of ≤5kPa, close the vacuum pump outlet valve, and suck the glass slurry into the tank. The liquid level of the glass slurry submerges the upper surface of the rigid insulation tile. After soaking for 5 minutes, take out the wet FRGMC panel.

[0234] (3) Place the wet FRGMC panel blank in a blast drying oven and dry it at a temperature of 120°C for ≥24 hours until the moisture in the blank is completely dried.

[0235] (4) Place the dried FRGMC panel blank into a muffle furnace for sintering. The specific process is as follows:

[0236] The first stage: the room temperature is raised to 1150°C, 3°C / min.

[0237] The second stage: Place the FRGMC panel dry blank into the muffle furnace at 1150℃.

[0238] The third stage: heating to 1250℃ at 3℃ / min.

[0239] The fourth stage: keep warm at 1250℃ for 30min.

[0240] The fifth stage: cooling down to 1150℃ at 5℃ / min.

[0241] Stage 6: Open the furnace door at 1150℃, take out the sample and rapidly cool it to room temperature in air.

[0242] After cooling, it is precisely processed into the designed shape and size using CNC machine tools to obtain the FRGMC panel.

[0243] The prepared FRGMC panel was tested and the apparent density of the FRGMC panel was 1.2-1.4 g / cm3 , the surface infrared hemispherical full spectrum emissivity ≥ 0.80, the compressive strength ≥ 30MPa, the surface hardness reaches Mohs 7, the equivalent thermal conductivity is Figure 9 As shown in the figure, the actual Figure 4 Compared with the traditional large-area thermal insulation tile thermal protection solution, which laminates a 50-200μm thick high-emissivity glass coating with an infrared hemispherical full-spectrum emissivity of 0.8 on the surface of the thermal insulation tile, the FRGMC panel has higher mechanical strength and reliability.

[0244] Example 4 Rigid thermal insulation tile / silica aerogel composite material

[0245] A method for preparing a rigid thermal insulation tile / silica aerogel composite material, comprising the following steps:

[0246] (1) Preparation of silica sol.

[0247] (1.1) Mix ethyl orthosilicate, water, and anhydrous ethanol in a molar ratio of 1:5:8 and stir for 2 minutes to obtain a mixed solution.

[0248] (1.2) Adjust the pH value of the mixed solution to 2-3 with 1 mol / L hydrochloric acid and stir at room temperature for 2 h. During this process, the ethyl orthosilicate is fully hydrolyzed.

[0249] (1.3) Continue to add 0.5 mol / L ammonia water to adjust the pH value of the above mixed solution to approximately neutral, and stir evenly.

[0250] (1.4) After step (1.3) is completed, continue stirring for 5 minutes, then add 1 mol / L ammonium fluoride aqueous solution, the mass ratio of ammonium fluoride aqueous solution to water is 0.5:9, and stir for 2 minutes to obtain SiO2 sol for use.

[0251] (2) Preparation of rigid thermal insulation tile / SiO2 wet gel composite materials.

[0252] (2.1) The block of density 0.24 g / cm prepared in Example 2 3 Thermoshield TM Place the ultrafine ceramic fiber rigid insulation tile in the vacuum impregnation tank, cover the upper cover of the vacuum impregnation tank, tighten the surrounding fastening bolts, close the glue inlet pipe valve and the liquid discharge pipe valve, open the pipe valve connected to the vacuum pump, and pump the air pressure in the tank to an absolute pressure of ≤5kPa.

[0253] (2.2) Open the valve of the glue inlet pipe and draw the SiO2 sol prepared in step (1) into the vacuum impregnation tank. The liquid level of the sol should be 20% higher than the thickness of the insulation tile. Connect the pipe connected to the vacuum pump to the compressed air cylinder and fill the tank with 1-2 MPa compressed air.

[0254] (2.3) After step (2.2) is completed, wait for the sol to automatically gel.

[0255] (2.4) After aging the gel at room temperature and pressure for 24 hours, open the vacuum impregnation tank, remove any residual liquid from the gel surface, and inject anhydrous ethanol until the gel surface is completely submerged. After 6 hours, remove the ethanol from the tank and refill with fresh anhydrous ethanol to soak the composite material.

[0256] (3) Supercritical drying: Place the soaked composite material in a supercritical carbon dioxide reactor. Add liquid carbon dioxide into the reactor so that the liquid surface completely submerges the upper surface of the wet gel. Control the temperature in the reactor to 60°C and the pressure to 10 MPa. After 30 minutes, release the pressure in the reactor at a rate of 25 kPa / min.

[0257] (4) Hydrophobic treatment: The composite material obtained in step (3) is placed in a vacuum tank, the absolute pressure in the tank is evacuated to ≤1 kPa, methyltrimethoxysilane and glacial acetic acid are introduced into the tank, the vacuum tank is heated to 70°C, kept at this temperature for 2 hours, and then cooled to room temperature. The vacuum tank is opened to obtain a hydrophobic thermal insulation tile / aerogel composite material.

[0258] (5) machining the hydrophobic composite material obtained in step (4) into the shape and size required by the design to obtain a rigid thermal insulation tile / silica aerogel composite material.

[0259] The obtained rigid thermal insulation tile / silica aerogel composite material was tested, and its SEM image is shown in Figure 6 and Figure 7 , the thermal properties test results are as follows:

[0260] Table 5

[0261]

[0262] Example 5

[0263] A rigid thermal insulation tile material with high temperature resistance, impact resistance and high emissivity is prepared in Example 2 with a density of 0.24g / cm 3 Thermoshield TM The surface of the ultrafine ceramic fiber rigid insulation tile is coated with a layer of glass glaze by spraying, and then rapidly sintered at a temperature of 1150°C and rapidly cooled.

[0264] The chemical composition of the glass glaze is the same as that of the glass paste in Example 3, except that the particle size D50 of the powder raw material constituting the glass glaze is 2.5 μm, and the content of deionized water is adjusted to a solid content of 30%.

[0265] The obtained rigid thermal insulation tile material was tested and its infrared hemispherical full spectrum emissivity was greater than or equal to 0.80, its temperature resistance was greater than or equal to 1200℃, and the average linear expansion coefficient of the coating was 0.55×10 -6 mm / (mm·℃) (average value from room temperature to 1200℃), and impact strength is 0.1J.

[0266] Example 6 Gradient Thermal Protection Structure

[0267] A gradient thermal protection structure, comprising: a FRGMC panel obtained by machining Example 3, a 0.24 g / cm3 thick slab obtained by machining Example 2, and a 0.24 g / cm3 thick slab. 3 Thermoshield TM The ultrafine ceramic fiber rigid insulation tile and the rigid insulation tile / silica aerogel composite material prepared in Example 4 are rolled to the desired thickness.

[0268] like Figure 1 As shown, when the FRGMC panel is prepared in Example 3, a plurality of mutually parallel first dovetail grooves are processed on one side of the FRGMC panel during precision machining using a CNC machine tool.

[0269] Second dovetail grooves parallel to each other are processed on the first side of the rigid insulation tile obtained in Example 2, and third dovetail grooves parallel to each other are processed on the second side. The first side and the second side are opposite to each other, and the groove direction of the second dovetail groove is perpendicular to the groove direction of the third dovetail groove.

[0270] Example 4: When preparing the rigid thermal insulation tile / silica aerogel composite material, a plurality of mutually parallel fourth dovetail grooves are machined on one side of the composite material.

[0271] The above components are assembled to obtain a surface toughened gradient heat protection structure. The assembly method is: mechanically interlocking the first dovetail groove with the second dovetail groove, and mechanically interlocking the third dovetail groove with the fourth dovetail groove. The assembled heat protection structure is as follows: Figure 10 shown.

[0272] The gradient thermal protection structure obtained above was tested, and the results were:

[0273] Temperature resistance test results: long-term repeated use temperature is 1600℃, short-term (≤5 minutes) single use temperature is 2000℃.

[0274] Impact resistance test results: 10J no cracks, significantly better than the RCG coated rigid insulation tile (0.05J no cracks), the rigid insulation tile material prepared in Example 5 (0.1J no cracks), and the TUFI coated rigid insulation tile (0.2J no cracks).

[0275] Thermal insulation performance test results: Figure 10As shown, the surface of the thermal protection structure can withstand a high temperature of 1600°C. The maximum temperature at the junction of the middle layer and the outer layer is 1400°C, the maximum temperature at the junction of the inner layer and the middle layer is 1100°C, and the innermost temperature of the inner layer is always below 100°C.

[0276] Comparative Example 1

[0277] A rigid thermal insulation tile material is basically the same as Example 5, except that: a high emissivity glass coating is composited according to the invention patent CN201510632090.0, and the actual figure is as follows Figure 5 shown.

[0278] Experimental findings: Due to the difference between the coating and the Thermoshield TM The thermal expansion properties of ultrafine ceramic fiber rigid insulation tiles do not match, and the coating cracks after cooling.

[0279] The gradient thermal protection structure of the present invention achieves various index requirements such as mechanical strength, high temperature resistance and thermal insulation performance, and can meet the needs of hypersonic aircraft traveling back and forth between the earth and the sky.

[0280] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, rather than to limit it. Although the present invention has been described in detail with reference to the aforementioned embodiments, those skilled in the art should understand that they can still modify the technical solutions described in the aforementioned embodiments, or make equivalent replacements for some of the technical features therein. However, these modifications or replacements do not deviate the essence of the corresponding technical solutions from the spirit and scope of the technical solutions of the various embodiments of the present invention.

Claims

1. A gradient thermal protection structure for a hypersonic vehicle, characterized in that: include: The outer layer is mainly composed of a heat protection panel; the rigid heat insulation tile used in the heat protection panel is denoted as a first rigid heat insulation tile; and, an intermediate layer, said intermediate layer being primarily composed of a second rigid insulation tile; and, an inner layer, wherein the inner layer is mainly composed of a rigid thermal insulation tile / silica aerogel composite material obtained by supercritical drying the third rigid thermal insulation tile composite silica aerogel; The outer layer is used to reduce the surface temperature of the heat protection structure from T1 to T2; The intermediate layer is used to reduce T2 to T3; The inner layer does not shrink at the T3 temperature; The heat protection panel comprises: a fiber-reinforced glass-based composite material obtained by impregnating glass slurry into the pores of the first rigid insulation tile, followed by drying and sintering; The first rigid thermal insulation tile is an ultrafine ceramic fiber rigid thermal insulation tile; the ultrafine ceramic fiber rigid thermal insulation tile has a porosity of more than 50% and an average pore size of less than 40 μm; The glass paste includes quartz glass powder, high silica glass powder with a core-shell structure, silicon tetraboride powder, molybdenum disilicide powder, tantalum disilicide powder, tantalum diboride powder and silicon carbide whiskers; the particle size of the powder raw materials constituting the glass paste is all submicron; after sintering, the glass paste is a continuous phase in the rigid insulation tile.

2. The gradient thermal protection structure for a hypersonic aircraft according to claim 1, characterized in that: The hemispherical full-spectrum emissivity of the outer layer to infrared radiation is greater than or equal to 0.9; and the apparent thermal conductivity of the middle layer at 1200° C. is less than 0.20 W / (m·K).

3. The gradient thermal protection structure for a hypersonic aircraft according to claim 1 or 2, characterized in that: The second rigid thermal insulation tile and / or the third rigid thermal insulation tile are rigid thermal insulation tiles that are resistant to high temperatures and impacts; The high temperature and impact resistant rigid thermal insulation tile is an ultrafine ceramic fiber rigid thermal insulation tile mainly composed of chopped quartz fiber, chopped mullite fiber, chopped aluminum borosilicate fiber, silicon carbide infrared sunscreen powder and boron nitride powder; The short-cut aluminum borosilicate fibers are mainly composed of short-cut aluminum borosilicate fibers with a diameter distribution of 480-520 nm and short-cut aluminum borosilicate fibers with a diameter distribution of 1-3 μm; The porosity of the ultrafine ceramic fiber rigid insulation tile is more than 50%, the average pore size is less than 40 μm, and the density is 1.0 g / cm 3 the following.

4. The gradient thermal protection structure for a hypersonic aircraft according to claim 3, characterized in that: The mass ratio of the short-cut aluminum borosilicate fibers with a diameter distribution of 480-520 nm to the short-cut aluminum borosilicate fibers with a diameter distribution of 1-3 μm is 1:3-5.

5. The gradient thermal protection structure for a hypersonic aircraft according to claim 4, characterized in that: The mass ratio of the short-cut aluminum borosilicate fibers with a diameter distribution of 480-520 nm to the short-cut aluminum borosilicate fibers with a diameter distribution of 1-3 μm is 1:

4.

6. The gradient thermal protection structure for a hypersonic aircraft according to claim 3, characterized in that: The raw materials for preparing the ultrafine ceramic fiber rigid insulation tile include, by weight: 250~300 parts of chopped quartz fiber; 85-90 parts of short-cut mullite fiber; 40-50 parts of chopped aluminum borosilicate fiber; 5-10 parts corn starch; 5-15 parts of silicon carbide infrared sunscreen powder; 3-6 parts of boron nitride powder.

7. The gradient thermal protection structure for a hypersonic vehicle according to claim 4 or 5, characterized in that: The raw materials for preparing the ultrafine ceramic fiber rigid insulation tile include, by weight: 250~300 parts of chopped quartz fiber; 85-90 parts of short-cut mullite fiber; 40-50 parts of chopped aluminum borosilicate fiber; 5-10 parts corn starch; 5-15 parts of silicon carbide infrared sunscreen powder; 3-6 parts of boron nitride powder.

8. The gradient thermal protection structure for a hypersonic vehicle according to claim 1, characterized in that: The powder raw materials constituting the glass paste have a D50 of less than or equal to 1 μm and a D90 of less than 1.5 μm.

9. The gradient thermal protection structure for a hypersonic vehicle according to claim 1 or 8, characterized in that: The aspect ratio L / D of the silicon carbide whisker is in the range of 50≤L / D≤100.

10. The gradient thermal protection structure for a hypersonic vehicle according to claim 1 or 8, characterized in that: The glass paste further includes a ceramic dispersant.

11. The gradient thermal protection structure for a hypersonic vehicle according to claim 10, characterized in that: The ceramic dispersant is one or a combination of two or more of Darvan 821A, Tween-80, sodium polyacrylate, acrylamide, polyacrylic acid, polymethacrylic acid, methyl cellulose, polymethacrylate ammonium salt with a pH value of 7, ammonia water, polyvinyl alcohol, polyethylene glycol, and polyperfluorosulfonic acid.

12. The gradient thermal protection structure for a hypersonic aircraft according to claim 11, characterized in that: The ceramic dispersant is Darvan 821A ceramic dispersant.

13. The gradient thermal protection structure for a hypersonic vehicle according to claim 9, characterized in that: The glass paste further includes a ceramic dispersant.

14. The gradient thermal protection structure for a hypersonic aircraft according to claim 13, characterized in that: The ceramic dispersant is one or a combination of two or more of Darvan 821A, Tween-80, sodium polyacrylate, acrylamide, polyacrylic acid, polymethacrylic acid, methyl cellulose, polymethacrylate ammonium salt with a pH value of 7, ammonia water, polyvinyl alcohol, polyethylene glycol, and polyperfluorosulfonic acid.

15. The gradient thermal protection structure for a hypersonic vehicle according to claim 14, characterized in that: The ceramic dispersant is Darvan 821A ceramic dispersant.

16. The gradient thermal protection structure for a hypersonic vehicle according to claim 1, characterized in that: The method for preparing the heat protection panel comprises: preparing the fiber-reinforced glass-based composite material by subjecting the rigid insulation tile to one or more impregnation-drying-sintering processes; The impregnation-drying-sintering process comprises: impregnating glass slurry in a stable suspension state into the pores of the rigid insulation tile under a vacuum environment to obtain a wet blank; and drying and sintering the wet blank.

17. The gradient thermal protection structure for a hypersonic vehicle according to claim 1 or 2, characterized in that: The density of the first rigid thermal insulation tile is higher than the density of the second rigid thermal insulation tile; and / or the density of the second rigid thermal insulation tile is the same as the density of the third rigid thermal insulation tile.

18. The gradient thermal protection structure for a hypersonic vehicle according to claim 3, characterized in that: The density of the first rigid thermal insulation tile is higher than the density of the second rigid thermal insulation tile; and / or the density of the second rigid thermal insulation tile is the same as the density of the third rigid thermal insulation tile.

19. The gradient thermal protection structure for a hypersonic vehicle according to claim 1 or 2, characterized in that: The outer layer, the middle layer and the inner layer are mechanically interlocked with each other using dovetail grooves.

20. The gradient thermal protection structure for a hypersonic vehicle according to claim 19, characterized in that: The direction of the dovetail groove connecting the middle layer and the outer layer is perpendicular to the direction of the dovetail groove connecting the middle layer and the inner layer.

21. The gradient thermal protection structure for a hypersonic vehicle according to claim 3, characterized in that: The outer layer, the middle layer and the inner layer are mechanically interlocked with each other using dovetail grooves.

22. The gradient thermal protection structure for a hypersonic aircraft according to claim 21, characterized in that: The direction of the dovetail groove connecting the middle layer and the outer layer is perpendicular to the direction of the dovetail groove connecting the middle layer and the inner layer.

23. The gradient thermal protection structure for a hypersonic vehicle according to claim 17, characterized in that: The outer layer, the middle layer and the inner layer are mechanically interlocked with each other using dovetail grooves.

24. The gradient thermal protection structure for a hypersonic aircraft according to claim 23, characterized in that: The direction of the dovetail groove connecting the middle layer and the outer layer is perpendicular to the direction of the dovetail groove connecting the middle layer and the inner layer.

25. The gradient thermal protection structure for a hypersonic aircraft according to claim 18, characterized in that: The outer layer, the middle layer and the inner layer are mechanically interlocked with each other using dovetail grooves.

26. The gradient thermal protection structure for a hypersonic vehicle according to claim 25, characterized in that: The direction of the dovetail groove connecting the middle layer and the outer layer is perpendicular to the direction of the dovetail groove connecting the middle layer and the inner layer.

27. The gradient thermal protection structure for a hypersonic vehicle according to claim 1 or 2, characterized in that: The thickness of the heat protection panel is 5-15 mm, the thickness of the second rigid insulation tile is 20-30 mm, and the total thickness of the gradient heat protection structure is 45-75 mm.

28. The gradient thermal protection structure for a hypersonic vehicle according to claim 27, characterized in that: The thickness of the heat protection panel is 10 mm, and the thickness of the second rigid insulation tile is 20 mm.

29. The gradient thermal protection structure for a hypersonic vehicle according to claim 3, characterized in that: The thickness of the heat protection panel is 5-15 mm, the thickness of the second rigid insulation tile is 20-30 mm, and the total thickness of the gradient heat protection structure is 45-75 mm.

30. The gradient thermal protection structure for a hypersonic aircraft according to claim 29, characterized in that: The thickness of the heat protection panel is 10 mm, and the thickness of the second rigid insulation tile is 20 mm.

31. The gradient thermal protection structure for a hypersonic aircraft according to claim 17, characterized in that: The thickness of the heat protection panel is 5-15 mm, the thickness of the second rigid insulation tile is 20-30 mm, and the total thickness of the gradient heat protection structure is 45-75 mm.

32. The gradient thermal protection structure for a hypersonic aircraft according to claim 31, characterized in that: The thickness of the heat protection panel is 10 mm, and the thickness of the second rigid insulation tile is 20 mm.

33. The gradient thermal protection structure for a hypersonic aircraft according to claim 18, characterized in that: The thickness of the heat protection panel is 5-15 mm, the thickness of the second rigid insulation tile is 20-30 mm, and the total thickness of the gradient heat protection structure is 45-75 mm.

34. The gradient thermal protection structure for a hypersonic aircraft according to claim 33, characterized in that: The thickness of the heat protection panel is 10 mm, and the thickness of the second rigid insulation tile is 20 mm.

35. Use of the gradient thermal protection structure according to any one of claims 1 to 34 in a hypersonic aircraft.

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