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CN-121990186-A - High-temperature-resistant composite structure of spacecraft protective cover and manufacturing method

CN121990186ACN 121990186 ACN121990186 ACN 121990186ACN-121990186-A

Abstract

A high-temperature resistant composite structure of a spacecraft protective cover and a manufacturing method thereof belong to the technical field of spacecrafts. The high-temperature-resistant composite structure comprises a runner base body and an upper cover plate, wherein the runner base body and the upper cover plate are manufactured by adopting a laser powder bed melting additive manufacturing technology, an inner runner with a three-dimensional topological structure is arranged in the runner base body, a surface to be combined is arranged on the outer surface of the runner base body, at least part of the inner runner is communicated with the surface to be combined, the upper cover plate is connected with the surface to be combined in a sealing mode and encloses the inner runner to form a complete closed runner, micro-nano concave-convex structures for breaking the smooth shape of the outer surface to disturb and weaken a stable thermal boundary layer are densely distributed on the outer surface of the upper cover plate, the height of each micro-nano concave-convex structure is 0.5-3 mm, the distance between adjacent structures is 1-4 mm, and a gradient thermal barrier coating is prepared on the surface of each micro-nano concave-convex structure in a layering mode. The micro-nano concave-convex structure of the high temperature resistant composite structure can disturb and weaken a stable thermal boundary layer, improve the convection heat exchange capacity of the outer surface and reduce the temperature peak value of a hot end region.

Inventors

  • XU MING
  • LI HUAIXUE
  • MA JINGYI
  • HU QUANDONG

Assignees

  • 中国航空制造技术研究院

Dates

Publication Date
20260508
Application Date
20260409

Claims (10)

  1. 1. A high-temperature-resistant composite structure of a spacecraft shield is used for heat-resistant protection of a hot end of the spacecraft shield and is characterized by comprising a runner matrix and an upper cover plate, wherein the runner matrix and the upper cover plate are manufactured by adopting a laser powder bed melting additive manufacturing technology; the inner runner with the three-dimensional topological structure is arranged in the runner matrix, a surface to be combined is arranged on the outer surface of the runner matrix, and at least part of the inner runner is communicated with the surface to be combined; The upper cover plate is connected to the surface to be combined in a sealing mode and forms a complete closed flow channel with the inner flow channel, micro-nano concave-convex structures used for damaging the smooth form of the outer surface to disturb and weaken a stable thermal boundary layer are densely distributed on the outer surface of the upper cover plate, the height of each micro-nano concave-convex structure is 0.5-3 mm, the distance between adjacent structures is 1-4 mm, and a gradient thermal barrier coating is prepared on the surface of each micro-nano concave-convex structure in a layering mode.
  2. 2. The high temperature resistant composite structure of a spacecraft shield of claim 1, wherein said micro-nano relief structure is any one or a combination of the following: The structure I is a truncated cone array or a truncated pyramid array, wherein the unit height is 0.3-2.5 mm, and the array pitch is 1.0-4.0 mm; the second structure is V-shaped ridge rib or herringbone micro rib, the rib height is 0.2-1.5 mm, the rib width is 0.3-2.0 mm, and the rib spacing is 0.8-3.0 mm; a third structure is a sine wave or saw tooth wave structure, the wave height is 0.2-2.0 mm, and the wavelength is 1.0-6.0 mm; And the fourth structure is a grid crossed rib or a well-shaped reinforcing structure, the rib height is 0.2-1.2 mm, and the grid side length is 1.0-6.0 mm.
  3. 3. The high temperature resistant composite structure of a spacecraft shield according to claim 1, wherein the relation between the substrate thickness t of the upper cover plate along the normal direction of the plate surface and the maximum flight mach number M of the spacecraft in the atmosphere is: When M <7 is 5≤M, t is 1.5 to 2.5 mm; when M <10 is 7≤M, t is 2.5 to 4.0 mm; When M is not less than 10M, t is 4.0 to 6.0 mm.
  4. 4. The high temperature resistant composite structure of a spacecraft shield according to claim 1, wherein the relation between the substrate thickness t of the upper cover plate along the normal direction of the plate surface and the maximum flight Mach number M of the spacecraft in the atmosphere is that 。
  5. 5. The high-temperature-resistant composite structure of the spacecraft shield according to claim 1, wherein a uniform and compact functional coating is formed in the closed flow channel, the functional coating is a pure copper coating prepared by an electrochemical deposition process, or the functional coating is a silicon carbide or tungsten carbide-based cermet wear-resistant heat-conducting coating prepared by a supersonic flame spraying technology.
  6. 6. The high temperature resistant composite structure of a spacecraft shield of claim 1, wherein said gradient thermal barrier coating comprises a bond coat, a transition layer, and a face layer; the bonding layer is deposited by adopting supersonic flame spraying or vacuum plasma spraying, and the thickness of the bonding layer is 80-150 mu m; the transition layers adopt atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition, at least two transition layers are formed through double powder feeding or multiple powder feeding proportion control, and the total thickness of the transition layers is 150-400 mu m; the surface layer is a rare earth zirconate or 8YSZ temperature resistant ceramic surface layer deposited outside the transition layer, and the thickness is 200-450 mu m.
  7. 7. A method of manufacturing a high temperature resistant composite structure for a spacecraft shield as claimed in any of claims 1 to 6, comprising the steps of: The method comprises the steps of printing a runner matrix and an upper cover plate in a partition mode, namely, adopting high-precision laser powder bed melting additive manufacturing equipment to respectively finish three-dimensional printing forming of the runner matrix and the upper cover plate, and controlling process parameters to ensure that the density of a formed part is not lower than 99.5%, wherein micro-nano concave-convex structures on the outer surface of the upper cover plate are directly formed in a printing stage, the height of each micro-nano concave-convex structure is 0.5-3 mm, the distance between adjacent structures is 1-4 mm, and the structural dimensional tolerance is controlled within +/-0.05 mm; Welding and integrating, namely precisely assembling the runner base body and the upper cover plate, controlling the assembly clearance to be 0.02-0.05mm, and realizing the sealing connection of the runner base body and the upper cover plate by adopting a vacuum brazing or electron beam welding technology; the preparation method of the external gradient thermal barrier coating comprises the following steps: surface pretreatment, namely carrying out sand blasting roughening, cleaning and degreasing on the micro-nano concave-convex structure on the outer surface of the upper cover plate to ensure that the surface roughness Ra is 3-8 mu m; the bonding layer is prepared by adopting supersonic flame spraying or vacuum plasma spraying to deposit NiCoCrAlY or MCrAlY bonding layer with thickness of 80-150 μm; preparing transition layers, namely forming at least two transition layers by adopting atmospheric plasma spraying, suspension plasma spraying or electron beam physical vapor deposition and controlling the proportion of double powder feeding or multiple powder feeding, wherein the total thickness of the transition layers is 150-400 mu m; And (3) preparing a surface layer, namely depositing an 8YSZ or Gd 2 Zr 2 O 7 temperature-resistant ceramic surface layer outside the transition layer, wherein the thickness is 200-450 mu m.
  8. 8. The method of manufacturing a refractory composite structure for a spacecraft shield of claim 7, further comprising, prior to the welding integration step, the steps of: The inner surface functionalization treatment comprises the steps of carrying out surface cleaning, sand blasting roughening pretreatment on a printed runner substrate and an upper cover plate, carrying out functionalized coating preparation on the inner surface of the runner, wherein the manufacturing process is an electrochemical deposition copper plating process, and the specific process comprises the steps of taking a pretreated workpiece as a cathode, taking a pure copper plate as an anode, uniformly depositing a pure copper coating with the thickness of 50-100 mu m on the inner surface of the runner in a copper sulfate electrolyte system by regulating and controlling the current density, the deposition temperature and the deposition time, wherein the purity of the coating is not lower than 99.9%, and the coating has no pinholes and crack defects.
  9. 9. The method for manufacturing the high-temperature-resistant composite structure of the spacecraft shield according to claim 7, wherein the method is characterized by further comprising the following steps of carrying out surface functionalization treatment on the printed runner base body and the upper cover plate, carrying out surface cleaning and sand blasting roughening pretreatment on the printed runner base body and the upper cover plate, and then carrying out functional coating preparation on the inner surface of the runner, wherein the manufacturing process is a supersonic flame spraying heat conduction coating process, and the specific process is that a supersonic flame spraying device is adopted, silicon carbide or tungsten carbide cermet powder is adopted as a spraying material, and a wear-resistant heat conduction coating with the thickness of 80-150 mu m is prepared on the inner surface of the runner, wherein the bonding strength of the coating is not lower than 50MPa, and the microhardness HV0.3 is more than or equal to 800.
  10. 10. The method for manufacturing a high temperature resistant composite structure of a spacecraft shield according to claim 7, wherein when a vacuum brazing runner base body and an upper cover plate are adopted in the welding integration step, brazing filler metal matched with a base body material is selected, welding is performed in a vacuum furnace with a vacuum degree of not less than 1 x 10 -3 Pa according to a curve of heating-heat preservation-cooling, the heat preservation temperature is 950-1100 ℃, and the heat preservation time is 30-60min.

Description

High-temperature-resistant composite structure of spacecraft protective cover and manufacturing method Technical Field The invention relates to the technical field of spacecrafts, in particular to a high-temperature-resistant composite structure of a spacecraft protective cover and a manufacturing method thereof. Background With the development of reusable space vehicles, reentry vehicles and long-term in-orbit operation of the spacecraft, hypersonic aircraft protective cover components such as a local hot end cover plate, a heat protective cover, an antenna cover or a window outer cover protective cover of a fairing and the like bear remarkable pneumatic heating and radiation heat exchange effects under the working conditions of ascending section, reentry section or high heat flow, and have high local heat flow density, large temperature gradient and frequent heat circulation load. In order to ensure the thermal safety and mechanical reliability of the internal sensitive equipment, the structure cementing interface and the bearing framework, the outer surface of the protective cover is usually required to be provided with a thermal barrier coating to reduce the temperature rise of the matrix, and meanwhile, the heat in the structure is usually required to be conducted and discharged through a temperature-resistant runner and a thermal control loop to form an active and passive cooperative thermal protection system. Traditional thermal protection schemes mainly comprise three technical approaches of passive heat insulation, active cooling and thermal structure. The passive heat insulation technology relies on ceramic matrix composite materials or heat insulation tiles with low heat conductivity coefficients, and the ceramic matrix composite materials or heat insulation tiles can bear high temperature, but have poor toughness and insufficient thermal shock resistance, and are difficult to meet the requirement of aerodynamic shape precision. Active cooling techniques deliver coolant, such as cryogenic fuel or water, through internal microchannels, but the system is complex, heavy, and risks coolant leakage. The thermal structure technology adopts high-temperature alloy or intermetallic compound, but has limited temperature resistance limit and large weight. The existing heat-resistant scheme of the protective cover mainly has the following technical problems: (1) The outer surface is in a flat shape, a relatively stable thermal boundary layer is easy to form under the high-speed outflow condition, the surface convection heat exchange capacity is limited, the heat in the hot end area is accumulated, and the peak value of the surface temperature is higher. (2) Most of traditional thermal barrier coatings are a two-layer system of a bonding layer and a single ceramic surface layer, the thermal expansion coefficients of the coatings and a metal matrix are not matched, and under the effects of thermal circulation and vibration impact load, the interfacial shear stress is concentrated, so that cracking, layering and peeling are easy to occur. (3) The outer surface of the protective cover often comprises complex geometric characteristics such as curved surface transition, local reinforcing ribs, assembly boundaries and the like, the thickness and the tissue consistency of the coating in the corner, convex or concave areas are difficult to ensure, and local weak areas are easy to form and failure expansion is easy to induce. (4) When the cooling flow channel is integrated in the protective cover, the traditional split machining and welding assembly process is difficult to simultaneously meet the machining accessibility, connection sealing reliability and consistency requirements of outer surface coating preparation of the complex structure of the flow channel. In recent years, despite certain advances in thermal barrier coating material systems and cooling techniques, the fundamental technical difficulties described above have not been effectively resolved. In particular, no breakthrough technical scheme has been found in breaking the stable thermal boundary layer and improving the heat dissipation efficiency. The Chinese patent CN109264030A discloses a convection cooling active heat protection structure, belongs to the technical field of heat protection of spacecrafts, and solves the problems that the heat protection structure in the prior art excessively depends on the heat protection performance of materials, and has complex structure, high cost and low heat protection efficiency. The convection cooling active heat protection structure comprises a surface layer, an intermediate layer and an inner layer, wherein a cooling flow passage is arranged at the part, close to the surface layer, of the intermediate layer. The cooling flow channel is an S-shaped flow channel. This solution improves the insulation effect but does not break the stable thermal boundary layer. Chinese patent CN2792948Y discloses a heat-insulatin