A spacecraft multilayer insulation assembly thermal bridge suppression bolted fixture structure

Abstract

The application relates to the technical field of spacecraft thermal control, and particularly provides a spacecraft multilayer heat insulation assembly thermal bridge suppression type bolt fixing structure which comprises a fixing assembly, a heat insulation bushing, a first heat insulation gasket, a second heat insulation gasket, a multilayer heat insulation assembly and a structural member; a stud passes through the second heat insulation gasket, the structural member and the multilayer heat insulation assembly in sequence; the outer surface of the stud is covered with the heat insulation bushing; the heat insulation bushing is a gradient functional composite structure; the heat insulation bushing comprises, from inside to outside, a heat insulation damping layer, a heat reflection layer and a structural enhancement radiation resistance layer; the contact surface between the second heat insulation gasket and the structural member is provided with a micro-cone array structure; the contact surface between the first heat insulation gasket and the top of the heat insulation bushing is provided with a micro-cone array structure; and the micro-cone array structure is used for filling assembly gaps and reducing solid heat conduction. The application has the advantages of low heat leakage, strong space adaptability and high thermal resistance stability.

Description

Technical Field

[0001] This invention belongs to the field of spacecraft thermal control technology, and particularly relates to a thermal bridge suppression bolt fixing structure for a multi-layer thermal insulation component of a spacecraft. Background Technology

[0002] In the field of spacecraft thermal control, multilayer thermal insulation components are widely used to isolate external heat flow and maintain internal temperature stability; theoretically, their equivalent thermal conductivity can reach 10. -5 With a temperature range in the order of w / m•℃, the material is very lightweight, produces no dust, and will not pollute the surrounding environment. The fixing method of multilayer insulation components (MLI) significantly affects their insulation performance, mainly by introducing thermal bridges, increasing solid contact between layers, disrupting the vacuum layer, and affecting the interlayer arrangement, thereby increasing solid thermal conductivity and reducing insulation efficiency.

[0003] Common installation methods for cryogenic multilayer thermal insulation components widely used on spacecraft include nylon mesh binding, nylon hook-and-loop fastening, and pin-and-clamp assembly fixing. Existing technologies often face a trade-off between "fixing reliability" and "thermal insulation performance." Specifically, to prevent loosening due to launch vibration, a large preload or contact area must be applied, which inevitably leads to a sharp increase in thermal conductivity.

[0004] Nylon mesh bags are generally used to fix multi-layer thermal insulation components with complex shapes (cylinders, cones). Since the objects being covered are mostly irregular in structure, it is difficult to control the tightness of the nylon mesh bags, which leads to a decrease in the thermal insulation performance of the multi-layer thermal insulation components.

[0005] Nylon hook and loop fasteners are easy to assemble and disassemble repeatedly, and rely on the "hooks" and "fuzz" of the hook and loop for adhesion, which has limited firmness. Moreover, the hook and loop fasteners in the fixed position are sewn together with multiple layers, and the multiple layers are in a compressed state, which reduces the heat insulation performance. Heat leakage is also likely to occur at the nylon hook and loop fastener joint.

[0006] In traditional pin-pressor assemblies, the pin head is typically exposed to the cold, dark space, while the tail is fixed to the structural component. Heat conduction through the pin can lead to heat leakage from the structural component. Furthermore, the multiple layers at the pressor fixing point are under compression, which also reduces the thermal insulation capacity of these layers. In conventional pin-pressor assemblies, the base and pin head of the pin are machined from the same material. The base is usually flat and fixed with silicone rubber. This type of pin head structure is prone to wear and deformation from repeated installation and removal of the pressor. The silicone rubber fixing makes disassembly of the pin very difficult after curing, and the disassembly process can easily damage the spacecraft surface. Moreover, the traditional planar contact significantly reduces contact thermal resistance under high preload, failing to meet the requirements of high-precision thermal control. Summary of the Invention

[0007] In view of this, the present invention aims to provide a thermal bridge suppression bolt fixing structure for a multi-layer thermal insulation component of a spacecraft, including a hollow titanium alloy stud, a thermal insulation bushing with a gradient functional composite structure covering the outer surface of the stud, and a thermal insulation gasket with a microcone array structure. The inner cavity of the stud forms a vacuum insulation layer, and the thermal insulation bushing is composed of three composite layers: a thermal insulation damping layer, a heat reflection layer, and a structurally reinforced radiation-resistant layer. The microcone array structure undergoes elastic or plastic deformation under preload to form a microscopic vacuum-sealed thermal resistance layer, thereby reducing solid heat conduction.

[0008] To achieve the above objectives, the technical solution created by this invention is implemented as follows: This invention provides a thermal bridge suppression bolt fixing structure for a spacecraft multi-layer thermal insulation assembly, comprising: a fixing component, a thermal insulation bushing, a first thermal insulation gasket, a second thermal insulation gasket, a multi-layer thermal insulation assembly, and structural components; The fixing assembly includes a nut, a stud, and a bolt, and the fixing assembly has a hollow structure; the stud passes through the second thermal insulation gasket, the structural component, and the multi-layer thermal insulation assembly in sequence; The outer surface of the stud is covered with a heat insulation bushing. The heat insulation bushing has a gradient functional composite structure. From the inside to the outside, the heat insulation bushing consists of a heat insulation damping layer, a heat reflection layer, and a structurally reinforced radiation-resistant layer. The upper surface of the second heat insulation pad contacts the lower surface of the structural component, the bottom of the heat insulation bushing contacts the upper surface of the structural component, and the top of the heat insulation bushing contacts the first heat insulation pad. The contact surfaces of the second heat insulation pad and the structural component are provided with a micro-cone array structure, and the contact surfaces of the top of the first heat insulation pad and the heat insulation bushing are provided with a micro-cone array structure. The micro-cone array structure is used to fill the assembly gap and reduce solid heat conduction. The center hole of the first heat insulation pad passes through the stud and contacts the heat insulation bushing and the multi-layer heat insulation assembly; the nut is located above the first heat insulation pad and is threadedly connected to the stud to fix the multi-layer heat insulation assembly.

[0009] Preferably, the fixing component is made of titanium alloy, the stud is a hollow cylindrical structure, the radial thermal conductivity of the stud is ≤1.5W / (m·K), and the stud wall thickness is greater than or equal to 0.2mm and less than or equal to 0.3mm.

[0010] Preferably, the surfaces of the nuts and bolts are treated with black anodizing, and the infrared emissivity ε≥0.85.

[0011] Preferably, the layers of the thermal insulation bushing are integrated through a hot-pressing process to prevent delamination under spatial vibration.

[0012] Preferably, the material of the thermal insulation damping layer is polyimide impregnated with nano-aerogel, with a thickness of ≥0.1mm and ≤0.3mm, a thermal conductivity ≤0.03W / (m·K), and a porosity of ≥95% for the nano-aerogel.

[0013] Preferably, the material of the heat reflective layer is an aluminized polyester film, the aluminum layer thickness is greater than or equal to 50 nm and less than or equal to 100 nm, and the reflectivity is ≥95%.

[0014] Preferably, the material of the structurally reinforced radiation-resistant layer is carbon fiber reinforced polyetheretherketone material with a thickness of ≥0.1 mm and ≤0.2 mm, an elastic modulus ≥5 GPa, and a fiber volume fraction of ≥40% and ≤50%.

[0015] Preferably, the material of the microcone array structure is silicone rubber, and the microcone array structure includes multiple cones with a height of greater than or equal to 50 μm and less than or equal to 100 μm and a cone angle of 60°.

[0016] Preferably, the cone angle of the microcone array structure is greater than or equal to 50° and less than or equal to 70°.

[0017] Preferably, the first and second heat insulation pads are made of carbon fiber substrate with a thickness of ≥0.2mm and ≤0.5mm.

[0018] Compared with the prior art, the present invention can achieve the following beneficial effects: This invention effectively reduces heat leakage at bolt connection points through the synergistic design of hollow titanium alloy fixing components, heat insulation bushings, and microcone array structures.

[0019] The outer layer of the thermal insulation bushing is made of carbon fiber reinforced polyetheretherketone material, which has a radiation resistance of over 200kGy, effectively solving the problem of polymer embrittlement in traditional thermal insulation bushings; the microcone array structure utilizes the resilience generated by its deformation to effectively prevent vacuum cold welding while maintaining high thermal resistance.

[0020] The hollow bolt design reduces weight and can withstand horizontal tensile forces of 0N to 100N without falling off; the inner layer of the thermal insulation bushing uses a nano-aerogel impregnated polyimide layer to ensure the fluffy state of the multi-layer thermal insulation components. Attached Figure Description

[0021] The accompanying drawings, which form part of this invention, are used to provide a further understanding of the invention. The illustrative embodiments and descriptions of the invention are used to explain the invention and do not constitute an undue limitation of the invention. In the drawings: Figure 1This is a cross-sectional view of a thermal bridge suppression bolt fixing structure for a spacecraft multilayer thermal insulation assembly provided in an embodiment of the present invention.

[0022] The reference numerals in the figures include: 2. Insulating bushing; 3. First insulating gasket; 4. Second insulating gasket; 5. Multi-layer insulating assembly; 6. Structural component; 7. Insulating damping layer; 8. Heat reflective layer; 9. Structurally reinforced radiation-resistant layer; 10. Microcone array structure; 11. Nut; 12. Stud; 13. Nut; 14. Thermal resistance layer. Detailed Implementation

[0023] To make the objectives, technical solutions, and advantages of this invention clearer, the invention will be further described in detail below with reference to the accompanying drawings and specific embodiments. It should be understood that the specific embodiments described herein are only for explaining the invention and do not constitute a limitation thereof. Similar elements in different embodiments are referred to by associated similar element reference numerals. In the following embodiments, many details are described to facilitate a better understanding of the invention. However, those skilled in the art will readily recognize that some features may be omitted in different situations, or may be replaced by other elements, materials, or methods. In some cases, some operations related to the invention are not shown or described in the specification. This is to avoid obscuring the core parts of the invention with excessive description. For those skilled in the art, detailed description of these related operations is not necessary; the relevant operations can be fully understood based on the description in the specification and general technical knowledge in the art.

[0024] It should be noted that, unless otherwise specified, the embodiments and features described in this invention can be combined to form various implementations. Furthermore, the order of the steps or actions in the method description can be changed or adjusted in a manner readily apparent to those skilled in the art. Therefore, the various orders in the specification and drawings are merely for the clear description of a particular embodiment and do not imply a mandatory order, unless otherwise stated that a particular order must be followed.

[0025] In the description of this invention, it should be understood that the terms "center," "longitudinal," "lateral," "length," "width," "thickness," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," "outer," "clockwise," and "counterclockwise," etc., indicating orientations or positional relationships based on the orientations or positional relationships shown in the accompanying drawings, are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation, and therefore should not be construed as a limitation on this invention. Furthermore, the terms "first," "second," etc., are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, features defined with "first," "second," etc., may explicitly or implicitly include one or more of that feature. In the description of this invention, unless otherwise stated, "a plurality of" means two or more.

[0026] In the description of this invention, it should be noted that, unless otherwise explicitly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection of two components. Those skilled in the art will understand the specific meaning of the above terms in this invention based on the specific circumstances.

[0027] The invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0028] Please see Figure 1 In one embodiment of the present invention, a thermal bridge suppression bolt fixing structure for a spacecraft multi-layer thermal insulation assembly is provided, comprising: a fixing assembly, a thermal insulation bushing 2, a first thermal insulation gasket 3, a second thermal insulation gasket 4, a multi-layer thermal insulation assembly 5, and a structural component 6; The fixing assembly includes a nut 11, a stud 12, and a nut 13. The fixing assembly has a hollow structure. The stud 12 passes through the second heat insulation gasket 4, the structural component 6, and the multi-layer heat insulation assembly 5 in sequence. The outer surface of the stud 12 is covered with a heat insulation bushing 2. The heat insulation bushing 2 is a gradient functional composite structure. From the inside to the outside, the heat insulation bushing 2 consists of a heat insulation damping layer 7, a heat reflection layer 8, and a structurally reinforced radiation-resistant layer 9. The upper surface of the second heat insulation pad 4 is in contact with the lower surface of the structural component 6, the bottom of the heat insulation bushing 2 is in contact with the upper surface of the structural component 6, and the top of the heat insulation bushing 2 is in contact with the first heat insulation pad 3. The contact surfaces of the second heat insulation pad 4 and the structural component 6 are provided with micro-cone array structures 10, and the contact surfaces of the top of the first heat insulation pad 3 and the heat insulation bushing 2 are provided with micro-cone array structures 10. The micro-cone array structures 10 are used to fill the assembly gaps and reduce solid heat conduction. The center hole of the first heat insulation pad 3 passes through the stud 12 and contacts the heat insulation bushing 2 and the multi-layer heat insulation assembly 5; the nut 13 is located above the first heat insulation pad 3 and is threadedly connected to the stud 12 to fix the multi-layer heat insulation assembly 5.

[0029] As an optional embodiment, the fixing assembly includes a nut 11, a stud 12, and a nut 13. The fixing assembly is a hollow structure made of titanium alloy. The stud 12 is a hollow cylindrical structure with a wall thickness greater than or equal to 0.2 mm and less than or equal to 0.3 mm, and a length greater than or equal to 12 mm and less than or equal to 15 mm. The inner cavity of the stud 12 forms a vacuum insulation gap layer with a radial thermal conductivity ≤1.5 W / (m·K). The nut 11 and the stud 12 are an integral structure. The stud 12 passes through the second heat insulation gasket 4, the structural component 6, and the multi-layer heat insulation assembly 5 in sequence. The upper surface of the second heat insulation gasket 4 contacts the lower surface of the structural component 6. The contact surface between the second heat insulation gasket 4 and the structural component 6 is provided with a micro-cone array structure 10.

[0030] The stud 12 is covered with a heat-insulating bushing 2, which is a gradient functional composite structure consisting of, from the inside out, a heat-insulating damping layer 7, a heat-reflecting layer 8, and a structurally reinforced radiation-resistant layer 9. The heat-insulating damping layer 7 is made of nano-aerogel-impregnated polyimide material with a thickness greater than or equal to 0.1 mm and less than or equal to 0.3 mm, a thermal conductivity ≤0.03 W / (m·K), and an aerogel porosity ≥95%. The heat-insulating damping layer 7 is used to reduce interfacial solid thermal conductivity. The heat-reflecting layer 8 is made of aluminized polyester film with a reflectivity ≥95% and an aluminum layer thickness greater than or equal to 50 nm and less than or equal to 100 nm. The heat-reflecting layer 8 reduces radiative heat transfer through high reflectivity. The structurally reinforced radiation-resistant layer 9 is made of carbon fiber-reinforced polyetheretherketone material with a thickness greater than or equal to 0.1 mm and less than or equal to 0.2 mm, an elastic modulus ≥5 GPa, and a fiber volume fraction greater than or equal to 40% and less than or equal to 50%. The structurally reinforced radiation-resistant layer 9 is used to provide mechanical support and improve radiation resistance and environmental adaptability. The layers are integrated through a hot-pressing process to prevent delamination under spatial vibration.

[0031] The bottom of the thermal insulation bushing 2 contacts the upper surface of the structural component 6, and the top of the thermal insulation bushing 2 contacts the first thermal insulation pad 3, forming a conformal interface and creating a load transfer path. This allows the bolt preload to be transferred through the thermal insulation bushing 2, thereby protecting the multi-layer thermal insulation component 5 filled on the outside of the thermal insulation bushing 2 from compaction and maintaining its theoretically high thermal insulation efficiency. A micro-cone array structure 10 is arranged on the contact surface between the first thermal insulation pad 3 and the top of the thermal insulation bushing 2. Both the first thermal insulation pad 3 and the second thermal insulation pad 4 are made of carbon fiber substrate with a thickness greater than or equal to 0.2 mm and less than or equal to 0.5 mm. The micro-cone array structure 10 includes multiple cones with a height greater than or equal to 50 μm and less than or equal to 100 μm, a cone angle of 60°, and is made of GD414 silicone rubber. The micro-cone array structure 10 is formed by molding GD414 silicone rubber onto the carbon fiber substrate. Without preload, there are interconnected macroscopic assembly gaps between the interfaces, and no closed microcavity structure is formed. Under a preload of ≥100N, the microcone array structure 10 undergoes elastic or plastic deformation, filling the interface gaps and forming multiple isolated microcavities enclosed by the deformed cone sidewalls, creating a broken gas cavity. This broken gas cavity suppresses residual gas heat conduction by dividing the continuous gas transmission path into multiple isolated microcavities. In aerospace vacuum or thermal vacuum environments, residual gas in the broken gas cavity is extracted, forming multiple closed microscopic vacuum cavities, thereby constructing a vacuum-sealed thermal resistance layer 14 (thermal resistance ≥0.8m²·K / W) at the bolt connection interface, improving contact thermal resistance and suppressing interfacial heat transfer. The microcone array structure 10 utilizes the rebound force generated by its deformation to effectively prevent vacuum cold welding while maintaining high thermal resistance. If the cone height of the microcone array structure 10 is less than 50 μm, it cannot effectively form a vacuum cavity; if the cone height is greater than 100 μm, it is prone to buckling instability under preload.

[0032] The central hole of the first heat insulation pad 3 passes through the stud 12, making it contact the heat insulation bushing 2 and the multi-layer heat insulation assembly 5. The nut 13 is located above the first heat insulation pad 3, and the lower surface of the nut 13 contacts the upper surface of the first heat insulation pad 3. The multi-layer heat insulation assembly 5 is pressed and fixed by the threaded connection between the nut 13 and the stud 12. The contact surface between the nut 13 and the first heat insulation pad 3 is provided with a micro-cone array structure 10 to prevent the nut 13 from loosening.

[0033] The surfaces of the nut 11 and the bolt 13 are black anodized, with an infrared emissivity ε≥0.85. The horizontal tensile force that the bolt can withstand at the nut 13 is greater than or equal to 0N and less than or equal to 100N. Within the horizontal tensile force range, the bolt will not fall off, and the increase in heat leakage is ≤5%.

[0034] When heat is conducted to the second heat insulation pad 4 through the structural component 6, the microcone array structure 10 blocks about 50% of the heat flow. When the residual heat is conducted along the stud 12, the hollow structure and the heat insulation bushing 2 block 84% of the heat flow, and the radiant heat flow is reflected by the heat reflection layer 8 by 96%.

[0035] As an optional embodiment, the cone angle of the microcone array structure 10 is greater than or equal to 50° and less than or equal to 70°, and the thickness of the heat insulation bushing 2 can be finely adjusted as needed to accommodate different preload forces.

[0036] This embodiment can be extended to other spacecraft thermal control systems, such as satellite heat pipe fixing or space station thermal insulation component installation.

[0037] In summary, the above description is merely a preferred embodiment of this specification and is not intended to limit the scope of protection of this specification. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principles of this specification should be included within the scope of protection of this specification.

[0038] The systems, apparatuses, modules, or units described in one or more of the above embodiments may be implemented by a computer chip or entity, or by a product having a certain function. A typical implementation device is a computer. Specifically, a computer may be, for example, a personal computer, a laptop computer, a cellular phone, a camera phone, a smartphone, a personal digital assistant, a media player, a navigation device, an email device, a game console, a tablet computer, a wearable device, or any combination of these devices.

[0039] It should also be noted that the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitation, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0040] The various embodiments in this specification are described in a progressive manner. Similar or identical parts between embodiments can be referred to interchangeably. Each embodiment focuses on describing the differences from other embodiments. In particular, the system embodiments are basically similar to the method embodiments, so the description is relatively simple; relevant parts can be referred to the descriptions in the method embodiments.

Claims

1. A thermal bridge suppression bolt fixing structure for a spacecraft multi-layer thermal insulation assembly, characterized in that, include: Fixed components, thermal insulation bushings, first thermal insulation gaskets, second thermal insulation gaskets, multi-layer thermal insulation components, and structural components; The fixing component includes a nut, a stud, and a bolt, and the fixing component has a hollow structure; the stud passes sequentially through the second heat insulation gasket, the structural component, and the multi-layer heat insulation component; The outer surface of the stud is covered with a heat insulation bushing, which is a gradient functional composite structure. From the inside to the outside, the heat insulation bushing consists of a heat insulation damping layer, a heat reflection layer, and a structurally reinforced radiation-resistant layer. The upper surface of the second heat insulation pad is in contact with the lower surface of the structural member, the bottom of the heat insulation bushing is in contact with the upper surface of the structural member, and the top of the heat insulation bushing is in contact with the first heat insulation pad. The contact surfaces of the second heat insulation pad and the structural component are provided with a micro-cone array structure, and the contact surfaces of the top of the first heat insulation pad and the heat insulation bushing are provided with a micro-cone array structure. The micro-cone array structure is used to fill the assembly gap and reduce solid heat conduction. The center hole of the first heat insulation pad passes through the stud and contacts the heat insulation bushing and the multi-layer heat insulation assembly; the nut is located above the first heat insulation pad and is threadedly connected to the stud to fix the multi-layer heat insulation assembly.

2. The spacecraft multi-layer thermal insulation assembly thermal bridge suppression bolt fixing structure according to claim 1, characterized in that, The fixing component is made of titanium alloy, the stud is a hollow cylindrical structure, the radial thermal conductivity of the stud is ≤1.5W / (m·K), and the wall thickness of the stud is greater than or equal to 0.2mm and less than or equal to 0.3mm.

3. The thermal bridge suppression bolt fixing structure for spacecraft multi-layer thermal insulation components according to claim 1, characterized in that, The surfaces of the nut and the bolt are subjected to black anodizing treatment, and the infrared emissivity ε≥0.

85.

4. The thermal bridge suppression bolt fixing structure for spacecraft multi-layer thermal insulation components according to claim 1, characterized in that, The layers of the thermal insulation bushing are integrated through a hot-pressing process to prevent delamination under spatial vibration.

5. The thermal bridge suppression bolt fixing structure for spacecraft multi-layer thermal insulation components according to claim 1, characterized in that, The thermal insulation damping layer is made of nano-aerogel impregnated polyimide with a thickness of ≥0.1mm and ≤0.3mm, a thermal conductivity ≤0.03W / (m·K), and a porosity of ≥95% for the nano-aerogel.

6. The thermal bridge suppression bolt fixing structure for spacecraft multi-layer thermal insulation components according to claim 1, characterized in that, The heat reflective layer is made of aluminum-coated polyester film, with an aluminum layer thickness of ≥50nm and ≤100nm and a reflectivity of ≥95%.

7. The thermal bridge suppression bolt fixing structure for spacecraft multi-layer thermal insulation components according to claim 1, characterized in that, The material of the structurally reinforced radiation-resistant layer is carbon fiber reinforced polyetheretherketone (PEEK), with a thickness of ≥0.1 mm and ≤0.2 mm, an elastic modulus ≥5 GPa, and a fiber volume fraction of ≥40% and ≤50%.

8. The thermal bridge suppression bolt fixing structure for spacecraft multi-layer thermal insulation components according to claim 1, characterized in that, The material of the microcone array structure is silicone rubber, and the microcone array structure includes multiple cones with a height of greater than or equal to 50 μm and less than or equal to 100 μm and a cone angle of 60°.

9. The thermal bridge suppression bolt fixing structure for a spacecraft multi-layer thermal insulation assembly according to claim 1, characterized in that, The cone angle of the microcone array structure is greater than or equal to 50° and less than or equal to 70°.

10. The thermal bridge suppression bolt fixing structure for a spacecraft multi-layer thermal insulation assembly according to claim 1, characterized in that, The first and second heat insulation pads are made of carbon fiber substrate with a thickness of ≥0.2mm and ≤0.5mm.

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