High temperature resistant composite structure for spacecraft shields and method of manufacture

Abstract

A high-temperature resistant composite structure and manufacturing method for a spacecraft protective shield belongs to the field of spacecraft technology. The high-temperature resistant composite structure includes a flow channel substrate and a top cover plate manufactured using laser powder bed fusion additive manufacturing technology. The flow channel substrate has an internal flow channel with a three-dimensional topological structure. A bonding surface is provided on the outer surface of the flow channel substrate, and at least part of the internal flow channel connects to the bonding surface. The top cover plate is sealed to the bonding surface and forms a complete closed flow channel with the internal flow channel. The outer surface of the top cover plate is densely covered with micro-nano uneven structures to disrupt the smooth morphology of the outer surface, thereby disturbing and weakening the stable thermal boundary layer. The height of the micro-nano uneven structures is 0.5~3mm, and the spacing between adjacent structures is 1~4mm. A gradient thermal barrier coating is prepared on the surface of the micro-nano uneven structures in layers. The micro-nano uneven structure of the high-temperature resistant composite structure of this invention can disturb and weaken the stable thermal boundary layer, improve the convective heat transfer capacity of the outer surface, and reduce the peak temperature in the hot end region.

Description

Technical Field

[0001] This invention relates to the field of spacecraft technology, specifically to a high-temperature resistant composite structure for a spacecraft protective shield and its manufacturing method. Background Technology

[0002] With the development of reusable launch vehicles, reentry vehicles, and long-term on-orbit spacecraft, hypersonic vehicle protective components, such as fairing hot-end covers, thermal shields, radomes, or window covers, are subjected to significant aerodynamic heating and radiative heat transfer during the ascent, reentry, or high-heat-flux phases. These components experience high local heat flux density, large temperature gradients, and frequent thermal cycling loads. To ensure the thermal safety and mechanical reliability of sensitive internal equipment, structural bonding interfaces, and the load-bearing frame, thermal barrier coatings are typically applied to the outer surface of the protective shield to reduce substrate temperature rise. Simultaneously, the internal structure often requires heat dissipation through temperature-resistant channels and thermal control loops, forming a combined active and passive thermal protection system.

[0003] Traditional thermal protection solutions mainly include three technical approaches: passive insulation, active cooling, and thermal structures. Passive insulation relies on ceramic matrix composites or insulating tiles with low thermal conductivity. While these can withstand high temperatures, they suffer from poor toughness, insufficient thermal shock resistance, and difficulty meeting aerodynamic shape precision requirements. Active cooling technology delivers coolants such as cryogenic fuels or water through internal microchannels, but the systems are complex, heavy, and pose a risk of coolant leakage. Thermal structure technology uses high-temperature alloys or intermetallic compounds, but their temperature resistance limits are limited, and they are also quite heavy.

[0004] The existing heat-resistant protective cover solutions mainly have the following technical problems:

[0005] (1) The outer surface is usually flat and a relatively stable thermal boundary layer is easily formed under high-speed outflow conditions. The surface convection heat transfer capacity is limited, heat accumulates in the hot end area, and the surface temperature peak is relatively high.

[0006] (2) Traditional thermal barrier coatings are mostly two-layer systems of “adhesive layer + single ceramic surface layer”. The thermal expansion coefficients of the coating and the metal substrate are not matched. Under the action of thermal cycling and vibration impact load, the interface shear stress is concentrated, which easily leads to cracking, delamination and peeling.

[0007] (3) The outer surface of the protective cover often contains complex geometric features such as curved transitions, local reinforcing ribs and assembly boundaries. The thickness and consistency of the coating in corners, protrusions or pits are difficult to guarantee, which can easily form local weak areas and induce failure propagation.

[0008] (4) When the cooling channel is integrated inside the protective cover, it is difficult to simultaneously meet the requirements of the processing accessibility of the complex channel structure, the reliability of the connection and sealing, and the consistency of the outer surface coating by using traditional split processing and welding assembly processes.

[0009] In recent years, although some progress has been made in thermal barrier coating material systems and cooling technologies, the aforementioned fundamental technical challenges have not yet been effectively resolved. In particular, no breakthrough technical solutions have been found in breaking the stable thermal boundary layer and improving heat dissipation efficiency.

[0010] Chinese invention patent CN109264030A discloses a convection-cooled active thermal protection structure, belonging to the field of spacecraft thermal protection technology. It solves the problems of existing thermal protection structures that overly rely on the thermal protection performance of materials, resulting in complex structures, high costs, and low thermal protection efficiency. This convection-cooled active thermal protection structure includes a surface layer, a middle layer, and an inner layer. The middle layer, near the surface, has cooling channels. These cooling channels are S-shaped. This solution improves the thermal insulation effect but does not break the stable thermal boundary layer.

[0011] Chinese invention patent CN2792948Y discloses a spacecraft heat insulation device. The device includes a heat insulation pad on the outer surface of the cabin wall. Perpendicular to the cabin's central axis, there are annular grooves on the outer side of the cabin wall, arranged in a stepped pattern from front to back. The outer edge of the upper groove projects onto the central axis of the lower groove, which contains flocculent material. However, this device does not disclose how to break the stable thermal boundary layer. Furthermore, the heat insulation material within the grooves slows down airflow, reducing heat exchange efficiency. Moreover, the annular structure limits its application scenarios. Summary of the Invention

[0012] This invention provides a high-temperature resistant composite structure for spacecraft protective shields and a manufacturing method thereof, in order to at least partially solve the above-mentioned technical problems.

[0013] As a first aspect of the present invention, a high-temperature resistant composite structure for a spacecraft protective shield is provided for heat-resistant protection of the hot end of the spacecraft protective shield, comprising a flow channel matrix and an upper cover plate manufactured using laser powder bed fusion additive manufacturing technology.

[0014] The interior of the flow channel matrix has an internal flow channel with a three-dimensional topological structure, and a bonding surface is provided on the outer surface of the flow channel matrix. At least part of the internal flow channel is connected to the bonding surface.

[0015] The upper cover plate is sealed to the surface to be joined and forms a complete closed flow channel with the inner flow channel; the outer surface of the upper cover plate is densely covered with micro-nano uneven structures to disrupt the flat shape of the outer surface, thereby disturbing and weakening the stable thermal boundary layer. The height of the micro-nano uneven structures is 0.5~3mm, the spacing between adjacent structures is 1~4mm, and the surface of the micro-nano uneven structures is layered with a gradient thermal barrier coating.

[0016] The micro / nano uneven structure is any one of the following structures or a combination thereof:

[0017] Structure 1: truncated frustum array or truncated pyramid array, unit height 0.3-2.5 mm, array pitch 1.0-4.0 mm;

[0018] Structure 2: V-shaped ribs or herringbone microribs, rib height 0.2-1.5 mm, rib width 0.3-2.0 mm, rib spacing 0.8-3.0 mm;

[0019] Structure 3: Sine wave or sawtooth wave structure, wave height 0.2-2.0 mm, wavelength 1.0-6.0 mm;

[0020] Structure 4: Grid cross rib or grid-shaped reinforcement structure, rib height 0.2-1.2 mm, grid side length 1.0-6.0 mm.

[0021] The relationship between the base thickness t of the upper cover plate along the normal direction of the plate surface and the maximum Mach number M of the spacecraft in the atmosphere is as follows:

[0022] When 5 ≤ M < 7, t is 1.5 to 2.5 mm;

[0023] When 7 ≤ M < 10, t is 2.5 to 4.0 mm;

[0024] When M≥10M, t is 4.0 to 6.0 mm.

[0025] The relationship between the base thickness t of the upper cover plate along the normal direction of the plate surface and the maximum Mach number M of the spacecraft in the atmosphere is as follows:

[0026] .

[0027] The closed flow channel contains a uniform and dense functional coating; the functional coating is a pure copper coating prepared by electrochemical deposition; or the functional coating is a silicon carbide or tungsten carbide-based metal ceramic wear-resistant and thermally conductive coating prepared by supersonic flame spraying technology.

[0028] The gradient thermal barrier coating includes an adhesive layer, a transition layer, and a surface layer.

[0029] The adhesive layer is a NiCoCrAlY or MCrAlY adhesive layer deposited by supersonic flame spraying or vacuum plasma spraying, with a thickness of 80~150μm;

[0030] The transition layer is formed by atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition, and at least two transition layers are formed by controlling the ratio of dual powder feeding or multiple powder feeding. The total thickness of the transition layer is 150-400μm.

[0031] The surface layer is a rare earth zirconate or 8YSZ heat-resistant ceramic surface layer deposited outside the transition layer, with a thickness of 200-450μm.

[0032] As a second aspect of the present invention, a method for manufacturing a high-temperature resistant composite structure for a spacecraft protective shield is also provided, comprising the following steps:

[0033] Sectional printing of the flow channel substrate and the top cover plate: The 3D printing of the flow channel substrate and the top cover plate is completed separately using high-precision laser powder bed fusion additive manufacturing equipment. By controlling the process parameters, the density of the molded parts is ensured to be no less than 99.5%. The micro-nano uneven structure on the outer surface of the top cover plate is directly formed during the printing stage. The height of a single micro-nano uneven structure is 0.5~3mm, the spacing between adjacent structures is 1~4mm, and the structural dimensional tolerance is controlled within ±0.05mm.

[0034] Welding integration: The flow channel substrate and the top cover plate are precisely assembled with the assembly gap controlled at 0.02-0.05mm. Vacuum brazing or electron beam welding technology is used to achieve a sealed connection between the flow channel substrate and the top cover plate.

[0035] Preparation of external gradient thermal barrier coating includes the following steps:

[0036] Surface pretreatment: The micro-nano uneven structure on the outer surface of the top cover plate is roughened by sandblasting and cleaned and degreased to make the surface roughness Ra 3-8 μm;

[0037] Adhesive layer preparation: NiCoCrAlY or MCrAlY adhesive layers are deposited by supersonic flame spraying or vacuum plasma spraying, with a thickness of 80-150 μm;

[0038] Transition layer preparation: At least two transition layers are formed by atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition, with a total thickness of 150-400 μm, through dual or multiple powder feeding ratio control.

[0039] Surface layer preparation: Deposit a high-temperature resistant ceramic surface layer of 8YSZ or Gd2Zr2O7 outside the transition layer, with a thickness of 200-450 μm.

[0040] Prior to the welding and integration step, the following steps are also included:

[0041] Internal surface functionalization treatment: After the printed flow channel substrate and the top cover plate are pretreated by surface cleaning and sandblasting roughening, a functional coating is prepared on the inner surface of the flow channel. The manufacturing process is electrochemical deposition copper plating process. The specific process is as follows: using the pretreated workpiece as the cathode and the pure copper plate as the anode, in the copper sulfate electrolyte system, by controlling the current density, deposition temperature and deposition time, a pure copper plating layer with a thickness of 50-100μm is uniformly deposited on the inner surface of the flow channel. The purity of the plating layer is not less than 99.9%, and there are no pinholes or cracks.

[0042] Before the welding integration step, the following steps are included: internal surface functionalization treatment: After the printed flow channel substrate and the top cover plate are pretreated by surface cleaning and sandblasting roughening, a functional coating is prepared on the inner surface of the flow channel. The manufacturing process is a supersonic flame spraying thermal conductive coating process. Specifically, a wear-resistant and thermally conductive coating with a thickness of 80-150μm is prepared on the inner surface of the flow channel using a supersonic flame spraying equipment and silicon carbide or tungsten carbide metal ceramic powder as the spraying material. The coating bonding strength is not less than 50MPa and the microhardness HV0.3≥800.

[0043] In the welding integration step, when vacuum brazing the flow channel substrate and the upper cover plate, a brazing filler metal matching the substrate material is selected, and the vacuum degree is not less than 1×10⁻⁶. -3 Welding is carried out in a vacuum furnace with a pressure of Pa according to the curve of "heating-holding-cooling". The holding temperature is 950-1100℃ and the holding time is 30-60min.

[0044] Based on the above solution, it can be seen that the high-temperature resistant composite structure and manufacturing method of the spacecraft protective shield of the present invention have at least one of the following beneficial effects compared with the prior art:

[0045] 1) Due to the micro-nano uneven structure on the outer surface of the top cover, its uneven outer surface can disturb and weaken the stable thermal boundary layer, improve the convective heat transfer capacity of the outer surface, and reduce the temperature peak of the hot end region.

[0046] 2) In gradient thermal barrier coatings, the interfacial shear stress caused by thermal expansion mismatch is reduced by composition and thickness gradation, which significantly improves thermal cycle life and reduces the risk of peeling.

[0047] 3) The micro-nano textured structure enhances the mechanical interlocking and partitioning constraint of the coating, which can suppress crack propagation and large-area peeling;

[0048] 4) The built-in temperature-resistant flow channel and the external gradient thermal barrier coating work together to form a thermal management path of "internal conduction and external resistance", which takes into account both thermal protection and structural thermal control.

[0049] 5) By using a hybrid manufacturing method that combines laser powder bed fusion additive manufacturing with welding or brazing, it is possible to achieve integrated forming of complex flow channel structures and uneven outer surfaces, thereby improving sealing reliability and structural consistency. Attached Figure Description

[0050] To more clearly illustrate the technical solution of the present invention, the accompanying drawings used in the embodiments will be briefly described below.

[0051] Figure 1 This is a schematic diagram of the high-temperature resistant composite structure of the spacecraft protective shield of the present invention.

[0052] Figure 2 A schematic diagram of a V-shaped rib structure for micro / nano concave-convex structures;

[0053] Figure 3 A schematic diagram of a sinusoidal wave structure with micro / nano-concave and convex structures;

[0054] Figure 4 A schematic diagram of a grid-shaped reinforcement structure for micro / nano concave-convex structures;

[0055] Figure 5 A simulation diagram of the V-shaped rib structure;

[0056] Figure 6 This is a simulation diagram of a sinusoidal wave structure;

[0057] Figure 7 A diagram showing the relationship between the spacecraft's maximum speed within the atmosphere and the thickness of the substrate along the normal direction of the top cover.

[0058] Figure 8 A graph showing the relationship between substrate thickness and the maximum Mach number of a spacecraft in the atmosphere.

[0059] In the figure: 100, flow channel substrate; 110, inner flow channel; 120, surface to be bonded; 200, top cover plate; 210, micro-nano uneven structure; 220, gradient thermal barrier coating. Detailed Implementation

[0060] To better understand the technical solutions of the embodiments of the present invention, the embodiments of the present invention will be described in detail below with reference to the accompanying drawings.

[0061] It should be understood that the described embodiments are merely some, not all, of the embodiments of the present invention. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0062] like Figures 1 to 7As shown in the figure, this embodiment discloses a high-temperature resistant composite structure for a spacecraft protective shield, which is used to overcome the problems of significant thermal boundary layer effect, limited heat transfer on the outer surface, easy cracking and peeling of traditional thermal barrier coatings, and difficulty in integrated manufacturing and insufficient sealing reliability of complex flow channel structures under high heat flux and thermal cycling conditions.

[0063] See Figure 1 The present invention discloses a high-temperature resistant composite structure for a spacecraft protective shield, used for heat-resistant protection of the hot end of the spacecraft protective shield. This high-temperature resistant composite structure includes a flow channel substrate 100 and an upper cover plate 200 manufactured using laser powder bed fusion additive manufacturing technology.

[0064] The flow channel substrate 100 has an inner flow channel 110 with a three-dimensional topological structure inside. A bonding surface 120 is provided on the outer surface of the flow channel substrate 100, and at least a portion of the inner flow channel 110 is connected to the bonding surface 120.

[0065] The flow channel substrate 100, serving as the load-bearing foundation and core heat conduction carrier of the entire temperature-resistant structure, is integrally molded using laser powder bed fusion additive manufacturing technology. A complex internal flow channel network structure 110 with high thermal conductivity and low flow resistance is constructed through three-dimensional topology optimization design. The internal flow channel network 110 is arranged according to the heat flow distribution in the hot-end region of the protective shield, and can be in parallel, series-parallel, or multi-branch configurations to ensure controllable heat conduction and exhaust capacity and pressure drop under limited pumping power conditions. The substrate material is selected from titanium alloys, nickel-based high-temperature alloys, or other heat-resistant alloys that combine high strength, good high-temperature stability, and excellent thermal conductivity. The material type and internal flow channel 110 topology parameters can be flexibly adjusted according to the specific operating conditions of the spacecraft's hot-end components.

[0066] The upper cover plate 200 is sealed to the mating surface 120 and forms a complete closed flow channel with the inner flow channel 110. This allows the cooling medium in the inner flow channel 110 to contact the inner surface of the upper cover plate 200, achieving high heat dissipation efficiency for the upper cover plate 200. The upper cover plate 200 has a mating surface that is perfectly adapted to the mating surface 120 of the flow channel base 100.

[0067] The upper cover plate 200 is the outer structure that directly faces the high-temperature airflow. The outer surface of the upper cover plate 200 is densely covered with micro / nano-uneven structures 210 designed to disrupt the smoothness of the outer surface, thereby disturbing and weakening the stable thermal boundary layer. The height of each micro / nano-uneven structure 210 is 0.5–3 mm, and the spacing between adjacent structures is 1–4 mm. A gradient thermal barrier coating 220 is layered on the surface of each micro / nano-uneven structure 210.

[0068] The high-temperature resistant composite structure of the spacecraft protective cover in this embodiment, by setting a micro-nano convex-concave structure, disturbs the flat shape of the outer surface of the upper cover plate 200 and weakens the stable thermal boundary layer. When the spacecraft flies in the atmosphere, it can enhance the near-wall disturbance and mixing of the outer surface of the upper cover plate 200, thereby improving the convective heat transfer capacity of the outer surface and reducing the temperature peak of the hot end region.

[0069] In existing technologies, the outer surface of protective shields is mostly flat. According to boundary layer theory, when high-speed airflow passes over a flat surface, a stable laminar boundary layer is formed, and heat is transferred only through molecular diffusion, resulting in a low convective heat transfer coefficient. This application designs a micro-nano uneven structure with a height of 0.5~3mm and a spacing of 1~4mm. By disrupting the boundary layer stability, the laminar boundary layer is transformed into turbulent flow, increasing the airflow mixing intensity in the near-wall region by more than 30%, significantly enhancing convective heat transfer. At the same time, this size range precisely matches the typical 1~5mm boundary layer thickness of hypersonic airflow, effectively disturbing the boundary layer while avoiding a surge in aerodynamic drag due to excessive structure height.

[0070] By setting a gradient thermal barrier coating, the interfacial thermal stress can be reduced by more than 40%. At the same time, the micro-nano uneven structure increases the contact area between the coating and the top cover plate by 2 to 3 times, forming a mechanical interlocking effect. The crack propagation path is blocked by the uneven structure, the crack propagation rate is reduced by 50%, and the thermal cycle life is significantly improved.

[0071] By using laser powder bed fusion additive manufacturing technology, the three-dimensional topological internal flow channel and the top cover plate are integrated into one, ensuring the accuracy and density of the flow channel structure. The internal flow channel and the top cover plate are sealed together to form a synergistic thermal management system of "internal conduction and external resistance". The external gradient coating blocks the high-temperature heat flow, and the internal flow channel conducts and dissipates heat through the cooling medium, realizing the dual functions of thermal protection and thermal control. At the same time, it avoids the sealing leakage risk caused by traditional split manufacturing and solves the technical problems of the accessibility of complex flow channel manufacturing and the reliability of use.

[0072] Furthermore, the micro / nano bump structure is any one of the following structures or a combination thereof:

[0073] Structure 1: truncated frustum array or truncated pyramid array, with unit height of 0.3-2.5 mm and array pitch of 1.0-4.0 mm; the frustum is a cone, and the pyramid can be a triangular pyramid, a square pyramid, a pentagonal pyramid, etc. It has isotropic disturbance capability and is suitable for regions with uncertain airflow direction or three-dimensional flow around. The truncated design of the raised unit avoids sharp corner airflow separation and enhances the mechanical anchoring effect of the coating.

[0074] Structure 2: V-shaped ribs or herringbone microribs, rib height 0.2-1.5 mm, rib width 0.3-2.0 mm, rib spacing 0.8-3.0 mm, such as... Figure 2 As shown.

[0075] Structure 3: Sine wave or sawtooth wave structure, wave height 0.2-2.0 mm, wavelength 1.0-6.0 mm, such as... Figure 3 As shown;

[0076] Structure 4: A grid-like intersecting rib or lattice-shaped reinforcing structure, with rib heights of 0.2-1.2 mm and grid side lengths of 1.0-6.0 mm. Figure 4 As shown.

[0077] The reason for selecting the above four structures for the micro / nano uneven structure 210 is that these morphologies can simultaneously address boundary layer control, coating adhesion reliability, thermal stress release, and additive manufacturing feasibility under hypersonic outflow and thermal cycling coupling conditions. Specifically, the micro / nano uneven structure 210 can disrupt the stable thermal boundary layer formed on a flat surface without significantly sacrificing dimensional accuracy, enhancing near-wall disturbance and mixing, thereby improving the convective heat transfer capacity of the outer surface and reducing the peak temperature in the hot end region. At the same time, it enhances the adhesion strength of the gradient thermal barrier coating by increasing the effective bonding area and forming geometric interlocking, and plays a role in partitioning and blocking the propagation of coating cracks, thereby improving thermal cycling life and reducing the risk of peeling.

[0078] More preferably, the four types of external surface shapes correspond to different flow direction sensitivity, structural stiffness contribution, and coating preparation consistency requirements.

[0079] Firstly, truncated frustum arrays or truncated pyramid arrays are lattice-type three-dimensional raised units with approximately uniform perturbation capabilities in all directions. They are suitable for regions where the outflow direction is uncertain or where there is strong three-dimensional flow around the surface. The raised units can significantly increase the coating interface area and form a reliable mechanical anchor. Furthermore, the truncated and rounded corner treatments help reduce stress concentration at sharp corners and improve the conformal coverage consistency of the coating.

[0080] Secondly, V-shaped ribs or herringbone microribs are highly directional flow guiding and disturbance structures. They can be arranged along the mainstream direction to induce paired vortex structures and enhance near-wall shear and heat transfer. Simultaneously, the ribs have a significant reinforcing effect on the upper cover plate within a 200° plane, improving erosion resistance and reducing thermally induced deformation. They are suitable for areas with a clearly defined mainstream direction and high aerodynamic loads. See details... Figure 5 Software simulations show that when the tip of the V-shaped rib reaches nearly 800°C, its root can be maintained at around 400°C, demonstrating ideal heat exchange performance.

[0081] Third, sinusoidal or sawtooth corrugated structures are continuous undulating surfaces, characterized by smooth geometric transitions, low stress concentration, and good coating deposition continuity. Sinusoidal corrugations are better suited to balancing shape drag and coating thickness uniformity, while sawtooth corrugations focus more on enhancing disturbance strength and weakening boundary layer stability. They are suitable for areas requiring stable, manufacturable morphologies over large areas where improved coating consistency is desired. See details... Figure 6 Software simulations show that when the tip of the sinusoidal structure reaches nearly 800°C, its root can be maintained at around 450°C, demonstrating ideal heat exchange performance.

[0082] Fourth, the grid-like cross ribs or lattice-shaped reinforcement structure provides reinforcement and partitioning constraints in two orthogonal directions simultaneously. This can form a "rigidized" load-bearing and anti-stripping frame in curved transitions, assembly boundaries, or areas of concentrated local heat flow. On the one hand, it improves the overall and local bending stiffness of the top cover plate; on the other hand, it provides segmentation constraints and crack inhibition for the coating, thereby suppressing long-distance crack propagation and large-area spalling. This is suitable for areas with complex load directions or more significant thermal stress gradients. Furthermore, the geometric parameter ranges of the above four types of structures match the direct forming technology of laser powder bed fusion additive manufacturing and the subsequent conformal deposition of gradient thermal barrier coatings. This allows for meeting dimensional tolerances and continuous coating coverage requirements while avoiding forming defects and weak areas in the coating caused by excessively small features, thus ensuring the engineering feasibility of the structure and coating system. The grid-like cross ribs are set at 45°, while the lattice-shaped reinforcement structure is set at 90°. See details... Figure 7 According to software simulations, with the tip of the grid-shaped reinforced structure approaching 800°C, its root can be maintained at around 450°C, demonstrating ideal heat exchange performance.

[0083] Preferably, the base thickness t of the upper cover plate along the normal direction of the plate surface is selected in segments. The relationship between the base thickness t of the upper cover plate along the normal direction of the plate surface and the maximum Mach number M of the spacecraft in the atmosphere is as follows:

[0084] When 5 ≤ M < 7, t is 1.5 to 2.5 mm;

[0085] When 7 ≤ M < 10, t is 2.5 to 4.0 mm;

[0086] When M≥10M, t is 4.0 to 6.0 mm.

[0087] The aforementioned thickness range can be determined by adjusting the heat flux density distribution, allowable inner surface temperature, allowable deformation, and allowable high-temperature stress of the material in the hot-end region of the protective cover. Within the same speed range, if the local heat flux density peak in the hot-end region is higher or the height of the uneven structure on the outer surface is taken as the upper limit, a larger base thickness t should be taken to improve the thermal shock resistance of the valley region and reduce the thermal stress gradient at the weld or brazed joint.

[0088] In existing technologies, the design of the thickness of the upper cover plate of a spacecraft protective shield mostly relies on engineering experience or conservative estimates, failing to establish a systematic correlation between thickness and flight Mach number. This results in either excessive redundancy in thickness design, increasing structural weight, or insufficient adaptation, unable to cope with the strong aerodynamic heating and loads at high Mach numbers. This application creatively establishes a segmented quantitative relationship between the thickness of the upper cover plate and the maximum flight Mach number based on aerodynamic thermodynamics and the high-temperature mechanical properties of materials. The core innovation lies in: deeply coupling the three factors of "flight speed - aerodynamic heating intensity - structural load requirements," adapting to the differences in operating conditions in different Mach number ranges through segmented design. In the low Mach number range, aerodynamic heating and loads are mild, and a thinner thickness is used to balance lightweight and basic protection requirements. In the medium and high Mach number range, the aerodynamic heat flux density and aerodynamic load are significantly increased. By thickening the cover plate, the heat conduction path is extended, and the structural stiffness is improved, ensuring that the inner surface temperature is controlled within the allowable range of the material, while resisting aerodynamic impact and thermal stress deformation. This solves the core problems of poor versatility and imbalance between weight and performance in traditional thickness designs.

[0089] In the low Mach number range of 5 ≤ M < 7, aerodynamic heating is mainly based on convection heat transfer, with a relatively mild heat flux density. The high-temperature strength decay of the material is limited, and a relatively thin thickness can achieve thermal protection through its own heat conduction and internal flow channel cooling, while meeting the requirements for lightweight structure and avoiding the increase in launch costs caused by over-design.

[0090] In the Mach number range of 7 ≤ M < 10, aerodynamic heating enters the stage of convection and radiation coupling. The heat flux density and aerodynamic load increase simultaneously, and the high-temperature strength of the material decreases significantly. Thickening the cover plate can enhance thermal resistance to suppress the rise of the inner surface temperature, while improving the structure's resistance to aerodynamic erosion and thermal bending, and avoiding structural failure caused by local stress concentration.

[0091] In the high Mach number range of M≥10, aerodynamic heating reaches extreme levels, and heat flux density and temperature gradient reach peak values. Materials face multiple failure risks such as creep and oxidation. Further thickening the cover plate can form a dual protection of thickness insulation and strong cooling of the flow channel, while ensuring the integrity of the structure under extreme loads, filling the scientific basis gap in the design of the thickness of the protective cover for high Mach number spacecraft.

[0092] This segmented design balances four core constraints: thermal protection, structural strength, lightweighting, and manufacturing feasibility. The thickness gradient variation is highly compatible with the aerodynamic heating intensity gradient, the material's high-temperature performance degradation law, and the forming capability of laser powder bed fusion additive manufacturing. This avoids the weight redundancy caused by excessively thick designs at low Mach numbers and solves the performance deficiencies caused by excessively thin designs at high Mach numbers. At the same time, the thickness range of each segment matches the process requirements of subsequent welding integration and coating preparation, ensuring that the thickness design does not affect the manufacturing consistency and connection reliability of the overall structure. This achieves three-dimensional optimization that is adaptable to working conditions, meets performance standards, and is process-feasible, breaking through the limitations of traditional thickness design guided by a single constraint.

[0093] Further, see Figure 8 Based on the formula obtained by fitting the experimental results, the relationship between the substrate thickness t along the normal direction of the upper cover plate 200 and the maximum Mach number M of the spacecraft in the atmosphere is as follows:

[0094]

[0095] In other words, the thickness of the substrate, 't', increases with the maximum flight Mach number, 'M'. This thickness is used to simultaneously meet the requirements for internal surface temperature limits and structural strength limits under the combined effects of aerodynamic heating heat flux density and aerodynamic pressure. The substrate thickness 't' is defined as the normal distance from the bottom of the valley of the micro-nano uneven structure 210 on the outer surface of the upper cover 200 to the inner sealing surface of the upper cover 200. The substrate thickness 't' is selected based on the aerodynamic heating intensity and aerodynamic load level corresponding to the Mach number 'M'. Specifically, it ensures that the micro-nano uneven structure on the outer surface of the upper cover can be integrally formed and the coating continuously covers the surface, while ensuring that the highest temperature on the inner surface of the upper cover does not exceed the set temperature limit, and that the maximum equivalent stress of the upper cover under the combined effects of aerodynamic pressure and thermal gradient does not exceed the allowable high-temperature stress of the material.

[0096] This application creatively establishes a continuous quantitative correlation between thickness and Mach number through a formula. The formula is derived from extensive aerodynamic thermodynamic simulations and high-temperature material test data, enabling the calculation of the optimal thickness value based on a specific Mach number rather than just a range. This ensures a precise correspondence between thickness design and actual flight conditions, avoiding the empirical dependence on thickness selection within a range and significantly improving the scientific rigor and accuracy of the design. The relationship between thickness and Mach number in the formula incorporates the nonlinear law of aerodynamic heat flux density changing with Mach number, the correlation between material thermal conductivity and temperature, and the mechanical relationship between structural stress and thickness. It also integrates the forming accuracy constraints of laser powder bed fusion additive manufacturing and the preparation requirements of gradient thermal barrier coatings. The thickness calculated by the formula ensures that the heat conduction path meets the internal surface temperature limit and that the structure meets the strength requirements under the coupled effects of aerodynamic loads and thermal stress. Furthermore, it adapts to the forming capabilities of additive manufacturing and the uniformity of subsequent coating deposition, solving the problem of insufficient consideration of multi-physics coupling effects in traditional thickness design. This formula provides a unified scientific basis for the modular and serialized design of spacecraft protective shields. For different types of spacecraft with different flight parameters, it eliminates the need for repeated complex multiphysics simulations and experiments. The optimal shield thickness can be quickly calculated simply by inputting the maximum flight Mach number, significantly improving design efficiency. At the same time, the formula can be modified according to the high-temperature performance of different matrix materials and the heat transfer efficiency of different cooling media. It is compatible with various matrix materials such as titanium alloys and nickel-based high-temperature alloys and different thermal control circuit designs, possessing strong engineering scalability. It fills the technical gap in the rapid and accurate design of spacecraft protective shield thickness and provides an efficient solution for the mass production of hypersonic spacecraft.

[0097] Furthermore, a uniform and dense functional coating is formed within the closed flow channel enclosed by the upper cover plate 200 and the flow channel substrate 100; this functional coating is a pure copper coating prepared by electrochemical deposition; or the functional coating is a silicon carbide or tungsten carbide-based metal ceramic wear-resistant and thermally conductive coating prepared by supersonic flame spraying technology. The purpose of setting this functional coating is to further enhance the heat exchange efficiency and wear resistance of the inner wall of the flow channel.

[0098] When selecting a pure copper coating for functional coatings, the high thermal conductivity of copper can significantly improve the heat transfer capacity of the flow channel, accelerating heat dissipation. The electrochemical deposition process ensures a uniform and dense coating, free of pinholes and cracks, preventing cooling medium leakage, meeting sealing requirements, and making it suitable for applications with high thermal conductivity requirements.

[0099] When using supersonic flame spraying technology to prepare silicon carbide or tungsten carbide-based metal ceramic wear-resistant and thermally conductive coatings, the erosion and wear resistance of the inner wall of the flow channel can be significantly enhanced while ensuring a certain thermal conductivity. This makes it suitable for harsh working conditions where the cooling medium contains particles or is subjected to high-speed scouring. The coating is firmly bonded to the substrate and can withstand the interfacial stress caused by thermal cycling and fluid impact, preventing detachment and failure.

[0100] Electrochemical deposition is suitable for complex three-dimensional flow channels, enabling uniform coating deposition on inner surfaces and avoiding dead zones that are difficult to cover with traditional coating processes. Supersonic flame spraying can prepare high-density, high-bonding-strength metal-ceramic coatings, meeting the wear-resistant requirements of the inner surfaces of flow channels. Both processes are compatible with additive manufacturing flow channel structures, ensuring the feasibility and consistency of coating preparation and solving the technical challenge of "functional enhancement and process adaptation" for complex flow channel inner surfaces.

[0101] A gradient thermal barrier coating 220 is layered on the micro-nano uneven surface of the upper cover plate 200 to serve as a core protective barrier against high-temperature environments. By gradually changing the coating composition and microstructure, the mismatch in thermal expansion coefficients between the coating and the substrate is effectively mitigated, reducing the risk of thermal stress cracking, while simultaneously achieving efficient isolation of high-temperature heat flow.

[0102] Specifically, the gradient thermal barrier coating includes an adhesive layer, a transition layer, and a top layer;

[0103] The bonding layer is a NiCoCrAlY or MCrAlY bonding layer deposited by supersonic flame spraying or vacuum plasma spraying, with a thickness of 80~150μm;

[0104] The transition layer is formed by atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition, and at least two transition layers are formed by controlling the ratio of dual powder feeding or multiple powder feeding. The total thickness of the transition layer is 150-400μm.

[0105] The top layer is a rare-earth zirconate or 8YSZ high-temperature resistant ceramic layer deposited outside the transition layer, with a thickness of 200-450μm. The thickness of the adhesive layer ensures sufficient bonding strength and avoids internal stress accumulation caused by excessive thickness; the number and thickness of the transition layer match the gradual change in the coefficient of thermal expansion, achieving effective dispersion of thermal stress; the thickness of the top layer balances thermal insulation effect and structural integrity, avoiding the risk of thermal insulation failure due to excessive thinness or cracking due to excessive thickness. The coating characteristics adapted to different spraying processes ensure the density, bonding strength, and uniformity of each coating layer, solving the technical problems of poor bonding strength and insufficient thermal cycling stability of traditional coatings.

[0106] This invention also discloses a method for manufacturing a high-temperature resistant composite structure for a spacecraft protective shield, characterized by comprising the following steps:

[0107] The flow channel substrate 100 and the upper cover plate 200 are printed in sections: The flow channel substrate and the upper cover plate 200 are 3D printed separately using high-precision laser powder bed fusion additive manufacturing equipment. The density of the molded parts is ensured to be no less than 99.5% by controlling the process parameters. The micro-nano uneven structure 210 on the outer surface of the upper cover plate 200 is directly formed in the printing stage. The height of a single micro-nano uneven structure 210 is 0.5~3mm, the spacing between adjacent structures is 1~4mm, and the structural dimensional tolerance is controlled within ±0.05mm.

[0108] Welding and integration: The flow channel substrate 100 and the upper cover plate 200 are precisely assembled with the assembly gap controlled at 0.02-0.05mm. Vacuum brazing or electron beam welding technology is used to achieve a sealed connection between the flow channel substrate 100 and the upper cover plate 200, and helium mass spectrometry leak detection is performed to confirm that the sealing performance meets the usage requirements.

[0109] Preparation of external gradient thermal barrier coating includes the following steps:

[0110] Surface pretreatment: The micro-nano uneven structure 210 on the outer surface of the upper cover plate 200 is roughened by sandblasting and cleaned and degreased to make the surface roughness Ra 3-8 μm;

[0111] Adhesive layer preparation: NiCoCrAlY or MCrAlY adhesive layers are deposited by supersonic flame spraying or vacuum plasma spraying, with a thickness of 80-150 μm;

[0112] Transition layer preparation: At least two transition layers are formed by atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition, with a total thickness of 150-400 μm, through dual or multiple powder feeding ratio control.

[0113] Surface layer preparation: A high-temperature resistant ceramic surface layer of 8YSZ or Gd2Zr2O7 with a thickness of 200-450 μm is deposited outside the transition layer. Other rare earth zirconates can also be selected from Gd2Zr2O7.

[0114] This method employs a partitioned printing strategy, using laser powder bed fusion additive manufacturing equipment to separately form the flow channel substrate and the top cover plate. Precise control of process parameters ensures high density of the formed parts, avoiding defects such as porosity and cracks. The micro-nano textured structure on the outer surface of the top cover plate is directly formed during the printing stage, eliminating the need for subsequent processing and ensuring dimensional accuracy and surface consistency, thus avoiding structural damage caused by secondary processing. This process enables the direct manufacturing of complex three-dimensional topologies, solving the problem of complex structures that are difficult to achieve with traditional processing methods. High-precision assembly gap control, combined with vacuum brazing or electron beam welding technology, ensures a sealed connection between the flow channel substrate and the top cover plate, preventing cooling medium leakage. The welding process in a vacuum environment prevents oxidation of the substrate and weld, improving the corrosion resistance and mechanical properties of the weld joint and solving the reliability problem of complex structural connections. By improving the coating's bonding foundation through surface pretreatment, the adhesive layer, transition layer, and top layer are prepared sequentially. The process parameters of each step are matched in a coordinated manner. The adhesive layer process ensures a strong bond with the substrate, the transition layer process achieves a gradual change in performance, and the top layer process ensures the thermal insulation effect. At the same time, a conformal deposition strategy is adopted for micro-nano uneven structures to ensure uniform coating coverage on complex surfaces, thus solving the technical problem of poor deposition consistency of traditional coatings on complex surfaces.

[0115] More specifically, during the preparation of the external gradient thermal barrier coating, conformal deposition is performed on the micro / nano uneven structure 210. For ultra-high temperature conditions exceeding 1200℃, a double-ceramic layer structure of a high-toughness underlayer and a high-insulation toplayer is employed. The high fracture toughness of 8YSZ alleviates thermal mismatch stress, while the low thermal conductivity and high phase stability of Gd₂Zr₂O₇ block high-temperature heat flow, solving the problems of high-temperature sintering and phase transformation failure in traditional YSZ coatings. Each coating layer is continuously and permeably covered along the micro / nano uneven structure, with the surface normal thickness as the control variable. For pits or valleys, multi-angle rotational spraying or multi-axis trajectory control is used to ensure that the minimum thickness of the valley is not less than 70% of the design thickness, and to ensure that there are no through gaps or uncovered areas between adjacent uneven units, thereby avoiding the formation of weak zones with concentrated thermal stress.

[0116] Examples of possible implementation structures for gradient thermal barrier coatings include the following.

[0117] Implementation Method 1: This method employs a classic metal-based gradient. Specifically, the binder layer uses NiCoCrAlY with a thickness of 80-150 μm; transition layer 1 consists of NiCoCrAlY and 8YSZ at a mass ratio of 70:30 with a thickness of 60-120 μm; transition layer 2 consists of NiCoCrAlY and 8YSZ at a mass ratio of 40:60 with a thickness of 80-150 μm; and the surface layer is 8YSZ with a thickness of 200-450 μm. The binder layer is applied using supersonic flame spraying, while the ceramic layer is applied using atmospheric plasma spraying with dual powder feeding and segmented switching.

[0118] Implementation Method 2: This method achieves a higher temperature dual ceramic gradient. Specifically, the binder layer is NiCoCrAlY with a thickness of approximately 100 μm; transition layer 1 is NiCoCrAlY to 8YSZ in a 50:50 mass ratio with a thickness of approximately 120 μm; transition layer 2 is 8YSZ with a thickness of approximately 200 μm; transition layer 3 is 8YSZ to Gd₂Zr₂O₇ in a 50:50 mass ratio with a thickness of approximately 180 μm; and the surface layer is Gd₂Zr₂O₇ with a thickness of 200-350 μm. The ceramic layers are applied using atmospheric plasma spraying with multiple powder feeds or segmented powder feeds.

[0119] Method 3: This method prioritizes porosity gradient. Specifically, the bonding layer is MCrAlY with a thickness of 80-120 μm, the ceramic layer composition remains unchanged at 8YSZ, and the total thickness is 350-650 μm. By segmentally adjusting the atmospheric plasma spraying power, spraying distance, and powder feeding rate, the porosity on the side near the substrate is less than 5%, and the porosity on the outer side is 8-18%.

[0120] Implementation Method 4: The ceramic-based protective cover is optional in this method. When the upper cover is SiC / SiC or other ceramic-based materials, it is preferable to first prepare an environmental barrier coating (EBC) bonding layer on the outer surface of the substrate, and then prepare a rare earth silicate transition layer and a heat-resistant ceramic surface layer to form a composite gradient system of EBC and TBC.

[0121] Furthermore, prior to the welding integration step, the following steps are also included:

[0122] Internal surface functionalization: After surface cleaning and sandblasting roughening pretreatment, the printed flow channel substrate 100 and upper cover plate 200 undergo functional coating preparation on the inner surface of the flow channel. The manufacturing process is electrochemical deposition copper plating. Specifically, using the pretreated workpiece as the cathode and a pure copper plate as the anode, a pure copper plating layer with a thickness of 50-100 μm is uniformly deposited on the inner surface of the flow channel in a copper sulfate electrolyte system by controlling the current density, deposition temperature, and deposition time. The plating purity is not less than 99.9%, and there are no pinholes or cracks. By adopting electrochemical deposition copper plating, the process route is optimized for the three-dimensional topology of the flow channel: in the pretreatment stage, cleaning and sandblasting roughening remove surface oil and oxide film, increase surface active sites, and improve the adhesion between the coating and the substrate; during the deposition process, electrode configuration and process parameter control ensure that copper ions are uniformly reduced and deposited on the inner surface of the flow channel, forming a dense coating without pinholes or cracks; the high thermal conductivity of the pure copper coating significantly improves the heat exchange efficiency of the flow channel.

[0123] Alternatively, before the welding integration step, the following step is included: internal surface functionalization treatment. After the printed flow channel substrate and the top cover plate undergo surface cleaning and sandblasting roughening pretreatment, a functional coating is prepared on the inner surface of the flow channel. The manufacturing process is a supersonic flame spraying thermal conductive coating process. Specifically, a supersonic flame spraying equipment is used, with silicon carbide or tungsten carbide metal ceramic powder as the spraying material. Parameters such as spraying distance, flame temperature, and powder feeding rate are optimized to prepare a wear-resistant and thermally conductive coating with a thickness of 80-150μm on the inner surface of the flow channel. The coating bonding strength is not less than 50MPa, and the microhardness HV0.3≥800.

[0124] Furthermore, when vacuum brazing the flow channel substrate and the upper cover plate in the welding integration step, a brazing filler metal matching the substrate material is selected, and the vacuum degree is not less than 1×10⁻⁶. -3 Welding is carried out in a vacuum furnace with a pressure of Pa according to the curve of "heating-holding-cooling". The holding temperature is 950-1100℃ and the holding time is 30-60min.

[0125] In summary, the high-temperature resistant composite structure and manufacturing method of the spacecraft protective shield of the present invention can achieve the following beneficial effects.

[0126] 1) Due to the micro-nano uneven structure on the outer surface of the top cover, its uneven outer surface can disturb and weaken the stable thermal boundary layer, improve the convective heat transfer capacity of the outer surface, and reduce the temperature peak of the hot end region.

[0127] 2) In gradient thermal barrier coatings, the interfacial shear stress caused by thermal expansion mismatch is reduced by composition and thickness gradation, which significantly improves thermal cycle life and reduces the risk of peeling.

[0128] 3) The micro-nano textured structure enhances the mechanical interlocking and partitioning constraint of the coating, which can suppress crack propagation and large-area peeling;

[0129] 4) The built-in temperature-resistant flow channel and the external gradient thermal barrier coating work together to form a thermal management path of "internal conduction and external resistance", which takes into account both thermal protection and structural thermal control.

[0130] 5) By using a hybrid manufacturing method that combines laser powder bed fusion additive manufacturing with welding or brazing, it is possible to achieve integrated forming of complex flow channel structures and uneven outer surfaces, thereby improving sealing reliability and structural consistency.

[0131] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of the present invention should be included within the scope of protection of the present invention.

Claims

1. A high-temperature resistant composite structure for a spacecraft protective shield, used for heat-resistant protection of the hot end of a spacecraft protective shield, characterized in that, This includes the flow channel matrix and top cover plate manufactured using laser powder bed fusion additive manufacturing technology; The interior of the flow channel matrix has an internal flow channel with a three-dimensional topological structure, and a bonding surface is provided on the outer surface of the flow channel matrix. At least part of the internal flow channel is connected to the bonding surface. The upper cover plate is sealed to the surface to be joined and forms a complete closed flow channel with the inner flow channel. A uniform and dense functional coating is formed in the closed flow channel. The functional coating is a pure copper coating prepared by electrochemical deposition; or the functional coating is a silicon carbide-based or tungsten carbide-based metal-ceramic wear-resistant and thermally conductive coating prepared by supersonic flame spraying technology. The outer surface of the upper cover plate is densely covered with micro-nano uneven structures to disrupt the flat shape of the outer surface, thereby disturbing and weakening the stable thermal boundary layer and enhancing near-wall disturbance and mixing. The height of the micro-nano uneven structures is 0.5~3mm, and the spacing between adjacent structures is 1~4mm. The surface of the micro-nano uneven structures is layered with a gradient thermal barrier coating, which includes an adhesive layer, a transition layer and a surface layer. The relationship between the base thickness t of the upper cover plate along the normal direction of the plate surface and the maximum Mach number M of the spacecraft in the atmosphere is as follows: When 5 ≤ M < 7, t is 1.5 to 2.5 mm; When 7 ≤ M < 10, t is 2.5 to 4.0 mm; When M≥10, t is 4.0 to 6.0 mm.

2. The high-temperature resistant composite structure of the spacecraft protective shield according to claim 1, characterized in that, The micro / nano-concave-convex structure is any one of the following structures or a combination thereof: Structure 1: truncated frustum array or truncated pyramid array, unit height 0.3-2.5 mm, array pitch 1.0-4.0 mm; Structure 2: V-shaped ribs or herringbone microribs, rib height 0.2-1.5 mm, rib width 0.3-2.0 mm, rib spacing 0.8-3.0 mm; Structure 3: Sine wave or sawtooth wave structure, wave height 0.2-2.0 mm, wavelength 1.0-6.0 mm; Structure 4: Grid cross rib or grid-shaped reinforcement structure, rib height 0.2-1.2 mm, grid side length 1.0-6.0 mm.

3. The high-temperature resistant composite structure of the spacecraft protective shield according to claim 1, characterized in that, The thickness t of the substrate along the normal direction of the upper cover plate is further calculated using the following formula: 。 4. The high-temperature resistant composite structure of the spacecraft protective shield according to claim 1, characterized in that, The adhesive layer is a NiCoCrAlY or MCrAlY adhesive layer deposited by supersonic flame spraying or vacuum plasma spraying, with a thickness of 80~150μm; The transition layer is formed by atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition, and at least two transition layers are formed by controlling the ratio of dual powder feeding or multiple powder feeding. The total thickness of the transition layer is 150-400μm. The surface layer is a rare earth zirconate or 8YSZ heat-resistant ceramic surface layer deposited outside the transition layer, with a thickness of 200-450μm.

5. A method for manufacturing a high-temperature resistant composite structure for a spacecraft protective shield as described in any one of claims 1 to 4, characterized in that, Includes the following steps: Sectional printing of the flow channel substrate and the top cover plate: The 3D printing of the flow channel substrate and the top cover plate is completed separately using high-precision laser powder bed fusion additive manufacturing equipment. By controlling the process parameters, the density of the molded parts is ensured to be no less than 99.5%. The micro-nano uneven structure on the outer surface of the top cover plate is directly formed during the printing stage. The height of a single micro-nano uneven structure is 0.5~3mm, the spacing between adjacent structures is 1~4mm, and the structural dimensional tolerance is controlled within ±0.05mm. Welding integration: The flow channel substrate and the top cover plate are precisely assembled with the assembly gap controlled at 0.02-0.05mm. Vacuum brazing or electron beam welding technology is used to achieve a sealed connection between the flow channel substrate and the top cover plate. Preparation of external gradient thermal barrier coating includes the following steps: Surface pretreatment: The micro-nano uneven structure on the outer surface of the top cover plate is roughened by sandblasting and cleaned and degreased to make the surface roughness Ra 3-8 μm; Adhesive layer preparation: NiCoCrAlY or MCrAlY adhesive layers are deposited by supersonic flame spraying or vacuum plasma spraying, with a thickness of 80-150 μm; Transition layer preparation: At least two transition layers are formed by atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition, with a total thickness of 150-400 μm, through dual or multiple powder feeding ratio control. Surface layer preparation: Deposit a high-temperature resistant ceramic surface layer of 8YSZ or Gd2Zr2O7 outside the transition layer, with a thickness of 200-450 μm.

6. The method for manufacturing the high-temperature resistant composite structure of the spacecraft protective shield according to claim 5, characterized in that, Prior to the welding and integration step, the following steps are also included: Internal surface functionalization treatment: After the printed flow channel substrate and the top cover plate are pretreated by surface cleaning and sandblasting roughening, a functional coating is prepared on the inner surface of the flow channel. The manufacturing process is electrochemical deposition copper plating process. The specific process is as follows: using the pretreated workpiece as the cathode and the pure copper plate as the anode, in the copper sulfate electrolyte system, by controlling the current density, deposition temperature and deposition time, a pure copper plating layer with a thickness of 50-100μm is uniformly deposited on the inner surface of the flow channel. The purity of the plating layer is not less than 99.9%, and there are no pinholes or cracks.

7. The method for manufacturing the high-temperature resistant composite structure of the spacecraft protective shield according to claim 5, characterized in that, Before the welding integration step, the following steps are also included: internal surface functionalization treatment: after the printed flow channel substrate and the top cover plate are pretreated by surface cleaning and sandblasting roughening, a functional coating is prepared on the inner surface of the flow channel. The manufacturing process is supersonic flame spraying thermal conductive coating process. The specific process involves using supersonic flame spraying equipment and silicon carbide or tungsten carbide metal ceramic powder as the spraying material to prepare a wear-resistant and thermally conductive coating with a thickness of 80-150μm on the inner surface of the flow channel. The coating bonding strength is not less than 50MPa and the microhardness HV0.3≥800.

8. The method for manufacturing the high-temperature resistant composite structure of the spacecraft protective shield according to claim 5, characterized in that, When vacuum brazing the flow channel substrate and the top cover plate during the welding integration process, a brazing filler metal matching the substrate material should be selected, and the vacuum degree should be no less than 1×10⁻⁶. -3 Welding is carried out in a vacuum furnace with a pressure of Pa according to the curve of "heating-holding-cooling". The holding temperature is 950-1100℃ and the holding time is 30-60min.

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