Satellite side plate body installs solar cell array heat insulation structure and satellite heat test heat leakage temporary protection method

Abstract

The present application relates to a kind of satellite side plate body installs solar cell array heat insulation structure and satellite heat test heat leakage temporary protection method, body installs solar cell array is set to the outer surface of satellite side plate towards cabin outside, so that the satellite side plate as battery array side plate, the heat insulation structure includes: multilayer heat insulation component, it is covered in the inner surface of battery array side plate towards cabin, is configured to block the heat radiation of battery array side plate to satellite cabin body structure plate transmission, and realizes the electrical isolation between multilayer heat insulation component and the battery array side plate;And multiple connecting pieces, it is configured to connect battery array side plate and satellite cabin body structure plate.The present application is reliable while realizing efficient heat insulation Electric insulation;Heat test verification, in ground test simulation screw hole in-orbit silicone rubber filling heat state, eliminate additional heat leakage interference, verify the effectiveness of heat insulation system design.

Description

Technical Field

[0001] This invention relates to the field of spacecraft thermal control system design and ground verification technology, and to a heat insulation structure for a satellite side plate-mounted solar cell array and a temporary protection method for heat leakage during satellite thermal tests. In particular, it relates to a heat insulation structure for a satellite side plate-mounted solar cell array and a temporary protection method for heat leakage during satellite thermal tests to reduce test deviations when using an infrared lamp array for thermal balance tests. Background Technology

[0002] To reduce weight, small satellites often employ a body-mounted solar panel configuration, where solar cells are directly attached to the outer surface of the satellite's main structural side panels. In this configuration, the body-mounted solar panel serves as both the satellite's main structure and the load-bearing structure for the solar array. Body-mounted solar panels experience long periods of on-orbit illumination and high external heat flux, resulting in high and fluctuating temperatures within the solar array itself. The transfer of heat from these panels to the cabin directly impacts the operating temperature of the equipment inside. Therefore, the effectiveness of the solar panel insulation design and the accurate verification of its insulation performance are two core tasks of the thermal control system.

[0003] The side panels of the solar array typically employ multilayer insulation (MLI) for radiative insulation, use thermal pads for conductive insulation, and install thermal bushings at screw connections to eliminate thermal bridges caused by direct metal-to-metal contact. However, this conventional insulation structure presents several problems: the solar panels typically use MLI for radiative insulation, and the reflector is a double-sided aluminized polyester film, which easily accumulates static charge in space and must be discharged via grounding. Traditional bulk-mounted substrates are generally thermally insulated and raised with the satellite side panels, with multiple layers laid on the outside of the satellite side panels, eliminating the need to consider insulation and conductivity between the multilayer layers and the substrate. However, when the satellite side panels and the bulk-mounted solar array share a common panel, the solar cells on the solar panel become charged during operation, requiring electrical insulation between the solar panel and the platform hull. This presents a contradiction: the multilayered metal reflector needs grounding, but the grounding path cannot disrupt the insulation between the solar panel and the hull. Furthermore, the multilayered modules require holes corresponding to the thermal pad locations; if the metal reflector cross-section at the edge of these holes is directly exposed, it may create unintended conductive paths.

[0004] Existing infrared lamps used for balancing tests have the following problems: During satellite flight, the screw holes connecting the solar panels and the cabin are filled with silicone rubber, and the screws are not directly exposed to the space thermal environment, resulting in no additional heat leakage. However, to meet the requirements of low cost and short cycle time in ground-based thermal balancing tests, the screw holes are usually not formally filled with silicone rubber (removing it after curing is difficult and easily damages the coating), leaving the screws exposed. Under infrared lamp illumination, the exposed screw heads have high absorptivity and low emissivity, absorbing heat and increasing temperature. Simultaneously, the semi-enclosed cavity formed by the screw head and the hole wall creates a black hole effect, exacerbating the temperature rise. Heat is conducted along the screw shaft to the cabin, forming an additional heat leakage channel that does not exist in orbit. Comparison shows that this factor can lead to inaccuracies in the on-orbit predictions of the thermal simulation model. Compared to previous satellites where the solar array substrate was raised and connected to the side panels with a fewer number of screws, the integrated solar array substrate and the supporting structure require consideration of both the rigidity and strength of the main structure. Therefore, there are more fasteners and contact surfaces, placing higher demands on the simulation accuracy in thermal vacuum tests. Summary of the Invention

[0005] This invention provides a heat insulation structure for a satellite side-panel solar array and a temporary protection method for heat leakage during satellite thermal testing. Specifically, it provides a heat insulation structure for a satellite side-panel solar array and a temporary protection method for heat leakage during satellite thermal testing to reduce test deviations when using an infrared lamp array for thermal balance testing. The heat insulation structure of this invention balances on-orbit heat insulation performance, insulation safety, and installation reliability. The temporary protection method for heat leakage during satellite thermal testing accurately simulates the on-orbit flight state of the solar array, and the screw protection design temporarily eliminates additional heat leakage from the screws during thermal balance testing.

[0006] The first aspect of this invention provides a heat insulation structure for a satellite side panel with a body-mounted solar array. The satellite side panel is the main load-bearing structure of the satellite, wherein the body-mounted solar array (used for on-orbit solar energy absorption and power generation) can be disposed on the outer surface of the satellite side panel facing outwards (towards space), so that the satellite side panel serves as the side panel of the solar array. The heat insulation structure includes: A multi-layer thermal insulation assembly, which covers the inner surface of the battery array side panel facing the interior of the cabin and is configured to block the heat radiation from the battery array side panel from being transmitted to the satellite cabin structural panel, and electrically isolates the multi-layer thermal insulation assembly from the battery array side panel; and Multiple connectors are configured to connect the battery array side panels to the satellite cabin structural panels.

[0007] Furthermore, from the side closest to the battery array side panel towards the interior of the cabin, the multi-layer heat insulation assembly sequentially includes: The first membrane layer is configured to provide insulation protection between the multilayer thermal insulation assembly and the battery array side panel; Thermal insulation units are configured to block radiative heat transfer from the battery array side panels to the satellite cabin structural panels (platform panels); and The second membrane layer, which has low emissivity, is configured to reduce heat exchange with the cabin's internal radiative heat, which is low-temperature radiation relative to the battery array sideplates.

[0008] Furthermore, the heat insulation unit includes: The reflective screen is configured to progressively attenuate the radiative heat flux in a vacuum environment, increase the radiative heat transfer resistance, and suppress radiative heat transfer between multi-layered insulation components; and The spacer layer is a low thermal conductivity isolation layer configured to isolate adjacent reflective screens, prevent direct contact between reflective screens, maintain the vacuum gap between layers, suppress solid heat conduction, and ensure the thermal insulation performance of the multilayer thermal insulation component; the spacer layer and reflective screen are stacked alternately in sequence from the first film layer to the second film layer.

[0009] Furthermore, the first film layer is a 50-60μm polyimide film; the reflective screen is a 5-6μm perforated double-sided aluminized polyester film; the spacer layer is a polyester mesh; and the second film layer is a perforated 16-20μm single-sided aluminized polyester film with the aluminized side facing inward.

[0010] Furthermore, the first membrane layer, the heat insulation unit, and the second membrane layer are stitched together with flame-retardant thread, and the layers are arranged in an overlapping manner.

[0011] Furthermore, the battery array side plate is provided with connector mounting holes, and the connector includes: The connector end is located on the side panel of the battery array near the interior of the compartment; and The connecting rod passes through the connecting mounting hole and extends into the interior of the satellite cabin structural plate.

[0012] Furthermore, the plurality of connectors are located at the docking and mounting surfaces of the battery array side plate and the satellite cabin structure plate; the satellite cabin structure plate constitutes a cabin enclosure structure, the plurality of connectors are arranged at the docking and mounting surfaces, the solar panel side plate is connected to the cabin structure plate, and the plurality of connectors are along the edge of the satellite cabin structure plate.

[0013] Furthermore, the connectors include titanium alloy connectors and stainless steel connectors.

[0014] Furthermore, it also includes: Multiple heat-insulating bushings are provided at one end of the connector rod near the connector end and are configured to block the solid heat conduction channel between the battery array side plate and the satellite cabin structural plate, while electrically insulating the connector rod from the battery array side plate.

[0015] Furthermore, the heat insulation bushing is a T-shaped heat insulation bushing with a T-shaped axial cross section, which is sleeved on the outer periphery of the connecting rod and accommodated in the mounting hole of the connecting member.

[0016] Furthermore, it also includes: Multiple heat insulation pads are provided, which surround the connector rod and abut one end against the inner surface of the battery array side plate facing the cabin, and are configured to block the solid heat conduction channel between the battery array side plate and the satellite cabin structural plate; holes are provided along the thickness direction of the multi-layer heat insulation assembly, which communicate with the mounting holes of the connector, so that when the connector is inserted through the holes, the multi-layer heat insulation assembly is arranged around the circumference of the heat insulation pads.

[0017] Furthermore, the heat insulation pad is an annular fiberglass heat insulation pad, which covers the connecting rod.

[0018] Furthermore, the axial length of the heat insulation pad is greater than the thickness of the multilayer heat insulation assembly, so that the heat insulation pad extends in a direction away from the end of the connector and passes through the second membrane layer of the multilayer heat insulation assembly.

[0019] Furthermore, the edges of the apertures are bound with polyimide single-sided adhesive to cover the reflective screen, isolating the metal section from electrical contact with the heat insulation pad; that is, the adhesive side of the polyimide single-sided adhesive faces the multilayer heat insulation assembly, and the non-adhesive side of the polyimide single-sided adhesive contacts the heat insulation pad; and / or The first film layer is attached to the side plate of the battery array with polyimide double-sided adhesive to form an insulating protection, which physically insulates the metal layer (reflective screen) of the multi-layer heat insulation component from the side plate of the battery array, thereby achieving electrical insulation and blocking; the first film layer extends 10~15mm beyond the heat insulation unit and the second film layer.

[0020] Furthermore, it also includes: A grounding wire is connected at one end to the heat insulation unit and the second membrane layer, and at the other end to the grounding terminal of the satellite cabin structural plate. Static electricity is discharged only through the grounding wire. The multi-layer grounding circuit is insulated from and independent of the battery panel side plate.

[0021] Furthermore, the connector is a screw, which includes a screw head and a screw shank.

[0022] Furthermore, the satellite cabin structure plate is provided with a first embedded part, which is adapted to the screw shank, so that the end of the screw shank away from the screw head is embedded in the first embedded part; the battery array side plate is provided with a screw mounting hole, and a second embedded part is provided in the screw mounting hole, so that after the screw passes through the screw mounting hole, the battery array side plate and the satellite cabin structure plate are fastened together.

[0023] Furthermore, a heat-insulating bushing is disposed at the end of the screw shank near the screw head and is configured to block the solid-state heat conduction channel between the battery array side plate and the satellite cabin structural plate, while electrically insulating the screw shank from the battery array side plate; a heat-insulating pad is disposed around the screw shank, with one end abutting against the inner surface of the battery array side plate facing the cabin, and is configured to block the solid-state heat conduction channel between the battery array side plate and the satellite cabin structural plate; the first embedded part is adapted to the screw shank, such that the end of the screw shank away from the screw head is embedded in the first embedded part; a screw mounting hole is provided on the battery array side plate, and a second pre-embedded part is provided in the screw mounting hole. An embedded part is provided so that after the screw passes through the screw mounting hole, it securely connects the battery array side plate and the satellite cabin structural plate. A hole is provided along the thickness direction of the multi-layer thermal insulation assembly, and the hole communicates with the screw mounting hole, so that when the screw passes through the hole, the multi-layer thermal insulation assembly surrounds the thermal insulation pad circumferentially. The axial length of the thermal insulation pad is greater than the thickness of the multi-layer thermal insulation assembly, so that the thermal insulation pad extends away from the screw head and passes through the second film layer of the multi-layer thermal insulation assembly. The thermal insulation bushing is fitted around the outer periphery of the screw shank and accommodated within the screw mounting hole. The thermal insulation pad covers the screw shank.

[0024] A second aspect of the present invention provides a temporary protection method for heat leakage during satellite thermal testing based on the thermal insulation structure described in the first aspect, comprising the following steps: Provide and cut flexible films; The flexible film is embedded into the connector mounting hole in an interference fit manner to seal the connector mounting hole (the flexible film covers the end of the connector). The edge of the flexible film is radially squeezed by the hole wall of the connector mounting hole to produce flexible deformation. It is then fixed inside the connector mounting hole by the elastic force of the flexible deformation. One-component room temperature vulcanizing silicone rubber is applied in dots at the gap between the edge of the flexible film and the wall of the mounting hole of the connector to form discrete auxiliary fixing adhesive dots, and left to stand at room temperature until the silicone rubber is surface dry. Using infrared lamp arrays or infrared heating cages to apply a thermal environment, spacecraft thermal balance tests are conducted, and operating temperature data is collected; and After the test, the flexible film was removed from the mounting hole of the connector, the residual silicone rubber in the hole was removed, and the satellite was restored to its original state.

[0025] Furthermore, the flexible film can be cut into circular sheets, polygonal sheets, or contoured sheets that conform to the outline of the mounting holes of the connectors.

[0026] Furthermore, the mounting holes for the connectors are screw mounting holes. Preferably, the flexible film is a flexible film disc adapted to fit the screw mounting holes for assembly. The flexible film disc covers the screw head, altering radiation characteristics and blocking the black hole effect, temporarily simulating the thermal state of silicone rubber filling in orbit during ground tests. Using this method, the additional heat leakage interference from the screws during ground thermal tests is effectively suppressed, and the thermal boundary conditions are closer to the actual in-orbit state.

[0027] Furthermore, the outer dimensions of the flexible film are 1mm to 3mm larger than the inner contour dimensions of the mounting hole of the connector.

[0028] Furthermore, a silver-plated secondary surface mirror or a white paint film made of perfluoroethylene propylene (F46) thin film with a thickness of 20~100µm can be used. The flexible film has a low solar absorptivity α. s ≤0.2, high infrared emissivity ε h Surface properties ≥0.75.

[0029] Furthermore, one side of the flexible film may be selectively laminated with an adhesive backing layer; when an adhesive backing layer is laminated, the thickness of the adhesive backing layer participates in the thickness filling of the interference fit and provides initial anti-slip friction during the initial embedding stage.

[0030] Furthermore, the middle part of the flexible film can make small-area contact with the end of the connector without generating adhesive force.

[0031] Furthermore, the number of discrete auxiliary fixing adhesive dots is 3 to 6.

[0032] The present invention has at least the following beneficial effects: 1) Integrated heat insulation and insulation: The present invention sets a polyimide film with an extended edge in the first film layer, which serves as both a multi-layer component structural layer for radiation heat insulation and an insulating layer to isolate the metal reflector and the platform panel. No additional insulation measures are required. With the help of the hole-position insulation edging, the internal conductive layer of the multi-layer heat insulation component is fully insulated from the external structure; 2) Independent grounding and insulation: The grounding end is set on the satellite body structural plate. The static electricity of the multi-layer heat insulation component is independently discharged through the grounding wire; the solar panel is kept insulated from the body through the heat insulation structure. The two circuits do not interfere with each other, taking into account both electrostatic protection and system insulation requirements; 3) Improved accuracy of on-orbit state simulation in thermal testing: The flexible film covers the end of the connector, changing the radiation characteristics and blocking the black hole effect. In the ground test, it temporarily simulates the thermal state of the silicone rubber filling in the on-orbit. After adopting this method, the additional heat leakage interference of the connector in the ground thermal test is effectively suppressed, and the thermal boundary conditions are closer to the real on-orbit state; 4) Completely reversible and low cost: The flexible film is mechanically fixed and can be removed after the test. There is no residual damage, no pollution, and it is easy to operate. Attached Figure Description

[0033] To further illustrate the above and other advantages and features of the various embodiments of the present invention, a more specific description of the embodiments of the invention will be presented with reference to the accompanying drawings. It is to be understood that these drawings depict only typical embodiments of the invention and are therefore not intended to limit its scope. In the drawings, identical or corresponding parts will be indicated by identical or similar reference numerals for clarity.

[0034] Figure 1 This diagram shows a cross-sectional view of the heat insulation structure of the satellite side plate solar cell array in some embodiments of the present invention. Figure 2 A schematic diagram showing the first film layer of the multilayer thermal insulation component extending beyond the epitaxial structure in some embodiments of the present invention is shown; Figure 3 This diagram illustrates the position and insulation edge binding status of the multilayer thermal insulation component and thermal insulation pad in some embodiments of the present invention. Figure 4 A schematic diagram of the grounding state of the multilayer thermal insulation assembly in some embodiments of the present invention is shown; Figure 5 A schematic diagram of a multilayer thermal insulation component covering a battery array side panel is shown in some embodiments of the present invention; Figure 6 This diagram shows a cross-sectional view of the heat insulation structure of the satellite side panel solar cell array in some embodiments of the present invention (F46 thin film disc protection state). Figure label: 1-Multi-layer thermal insulation component, 101-First film layer, 102-Insulation unit, 1021-Reflector, 1022-Spacer layer, 103-Second film layer, 2-Built-in solar cell array, 3-Cell array side plate, 4-Satellite cabin structure plate, 5-Screw, 501-Screw head, 502-Screw shank, 6-Insulation bushing, 7-Insulation pad, 8-First embedded part, 9-Second embedded part, 10-Back side of cell array side plate, 11-Polyimide double-sided adhesive, 12-Extended portion of first film layer, 13-Polyimide single-sided adhesive, 14-Grounding wire, 15-Grounding terminal, 16-Flexible film, 17-Single-component room temperature vulcanizing silicone rubber, 18-Outside cabin, 19-Inside cabin. Detailed Implementation

[0035] It should be noted that the components in the accompanying drawings may be shown exaggerated for illustrative purposes and may not be to scale.

[0036] In this invention, the various embodiments are merely intended to illustrate the solutions of the invention and should not be construed as limiting.

[0037] In this invention, unless otherwise specified, the quantifiers “a” and “one” do not exclude scenarios involving multiple elements.

[0038] It should also be noted that, in the embodiments of the present invention, only a portion of the parts or components may be shown for clarity and simplicity. However, those skilled in the art will understand that, under the teachings of the present invention, the required parts or components can be added as needed for specific scenarios.

[0039] It should also be noted that within the scope of this invention, the terms "same", "equal", and "equal to" do not mean that the two values ​​are absolutely equal, but allow for a certain reasonable error. In other words, the terms also cover "substantially the same", "substantially equal", and "substantially equal to".

[0040] It should also be noted that in the description of this invention, the terms "center," "longitudinal," "lateral," "upper," "lower," "front," "rear," "left," "right," "vertical," "horizontal," "top," "bottom," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing the invention and for simplifying the description, and do not explicitly or implicitly suggest that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on the invention. Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0041] Furthermore, the embodiments of the present invention describe the process steps in a specific order. However, this is only for the convenience of distinguishing each step, and is not a limitation on the order of each step. In different embodiments of the present invention, the order of each step can be adjusted according to the process.

[0042] The following embodiment provides a heat insulation structure for a satellite side panel-mounted solar cell array 2. Figure 1 This diagram shows a cross-sectional view of the thermal insulation structure of the satellite side panel's body-mounted solar array 2. The satellite side panel is the main load-bearing structure of the satellite. The body-mounted solar array 2 (used for on-orbit solar energy absorption and power generation) can be installed on the outer surface of the satellite side panel facing the outer side of the cabin 18 (facing the space side), so that the satellite side panel serves as the solar array side panel 3. The thermal insulation structure includes: The multi-layer heat insulation component 1 covers the inner surface of the battery array side panel 3 facing the cabin 19, with open sides for heat dissipation. It is configured to block the heat radiation of the battery array side panel 3 from being transmitted to the satellite cabin structure plate 4, and to electrically isolate the multi-layer heat insulation component 1 from the battery array side panel 3. This structure effectively isolates the impact of temperature changes of the battery array side panel 3 on the individual units in the cabin 19 from the multi-layer heat insulation component 1. The outer side panel 18 facing the cabin serves as a cooperating part of the heat insulation structure. The multi-layer heat insulation component 1 only covers the area of ​​the battery array side panel 10 in the cabin 19, while the outer side panel 18 is not covered by the multi-layer heat insulation component 1 and is directly exposed to the space environment. The outer side 18 facing the cabin serves as a radiative heat dissipation surface, dissipating some heat into the cold black space, thus preventing the solar panel temperature from becoming too high while providing heat insulation. Multiple screws 5 are configured to connect the battery array side plate 3 to the satellite cabin structure plate 4. Each screw 5 includes a screw head 501 and a screw shank 502. The screw head 501 is located on the side of the battery array side plate 3 near the cabin interior 19, and the screw shank 502 passes through the screw 5 mounting hole and extends into the interior of the satellite cabin structure plate 4. Multiple heat insulation bushings 6 are provided at one end of the screw shank 502 near the screw head 501 and are configured to block the solid heat conduction channel between the battery array side plate 3 and the satellite cabin structure plate 4, while electrically insulating the screw shank 502 from the battery array side plate 3; the heat insulation bushing 6 is a T-shaped heat insulation bushing with an axial cross section of T, which is sleeved on the outer periphery of the screw shank 502 and accommodated in the screw 5 mounting hole; Multiple heat insulation pads 7 surround the screw rod portion 502, with one end abutting against the inner surface of the battery array side plate 3 facing the cabin interior 19, and are configured to block the solid heat conduction channel between the battery array side plate 3 and the satellite cabin structural plate 4; holes are provided along the thickness direction of the multi-layer heat insulation assembly 1, and the holes communicate with the screw 5 mounting holes, so that when the screw 5 passes through the holes, the multi-layer heat insulation assembly 1 surrounds the heat insulation pads 7 circumferentially. Figure 3 Both the heat insulation pad 7 and the heat insulation bushing 6 are made of low thermal conductivity materials, together forming a heat-conducting insulation layer that blocks the heat transfer path from the side panel to the cabin; and The grounding wire 14 has one end electrically connected to each metal reflector 1021, and the other end connected to the grounding terminal 15 of the satellite cabin structure plate 4. Figure 5 Static electricity is discharged only through grounding wire 14. The multi-layer grounding circuit is insulated from and independent of the side plate of the solar panel. Figure 4 A schematic diagram of the grounding status of a multilayer thermal insulation assembly is shown.

[0043] From the side closest to the battery array side panel 3 towards the interior 19 (from the back of the battery array side panel 10), the multi-layer heat insulation assembly 1 sequentially includes: The first film layer 101 is configured to provide insulation protection between the multilayer thermal insulation component 1 and the battery array side panel 3; the first film layer 101 is a 50μm polyimide film. Thermal insulation unit 102 is configured to block radiative heat transfer from the battery array side panel 3 to the satellite cabin structural panel 4 (platform panel); and The second membrane layer 103, having low emissivity, is configured to reduce radiative heat exchange with the interior 19, where the heat radiation is low-temperature radiation relative to the battery array side plate 3. The second membrane layer 103 is a 16-20 μm single-sided aluminized polyester film, with the aluminized side facing the interior 19. The first membrane layer 101, the heat insulation unit 102, and the second membrane layer 103 are stitched together with flame-retardant thread, and the layers are arranged in an overlapping manner.

[0044] There are 20 insulation units 102, and each insulation unit 102 includes: The reflector 1021 is configured to progressively decrease the radiative heat flux under vacuum conditions, increase the radiative heat transfer resistance, and suppress radiative heat transfer between the multilayer insulation components 1; the reflector 1021 is a 6μm double-sided aluminized polyester film; and The spacer layer 1022 is a low thermal conductivity isolation layer and is configured to isolate adjacent reflective screens 1021, prevent direct contact between reflective screens 1021, maintain the vacuum gap between layers, suppress solid heat conduction, and ensure the heat insulation performance of the multilayer heat insulation component 1. The spacer layer 1022 is a polyester mesh, and the spacer layer 1022 and reflective screens 1021 are stacked alternately from the first film layer 101 to the second film layer 103.

[0045] The edges of the holes are bound with polyimide single-sided adhesive 13 so that the polyimide single-sided adhesive 13 covers the reflector screen 1021, isolating the metal section from the heat insulation pad 7; that is, the adhesive side of the polyimide single-sided adhesive 13 faces the multilayer heat insulation component 1, and the non-adhesive side of the polyimide single-sided adhesive 13 contacts the heat insulation pad 7; the first film layer 101 is attached to the battery array side plate 3 by polyimide double-sided adhesive 11, and the first film layer 101 forms an insulating protection, physically insulating the metal layer (reflector screen 1021) of the multilayer heat insulation component 1 from the battery array side plate 3, and achieving electrical insulation blocking; the first film layer 101 extends 10~15mm beyond the heat insulation unit 102 and the second film layer 103, and the extended part is defined as the first film layer extension 12 ( Figure 2 The first film layer 101 and the edges of the pores are edged with polyimide single-sided adhesive 13 and heat insulation bushing 6 to ensure electrical insulation of the battery panel side plate and platform compartment plate.

[0046] Screws 5 are located at the mating surfaces of the battery array side plate 3 and the satellite cabin structure plate 4. The satellite cabin structure plate 4 forms the cabin enclosure structure, and multiple screws 5 are arranged at the mating surfaces. The solar panel side plate is connected to the cabin structure plate, and multiple screws 5 are along the edge of the satellite cabin structure plate 4. A first embedded part 8 is provided in the satellite cabin structure plate 4. The first embedded part 8 is adapted to the screw shank 502, so that the end of the screw shank 502 away from the screw head 501 is embedded in the first embedded part 8. The battery array side plate 3 has screw 5 mounting holes, and a second embedded part 9 is provided in the screw 5 mounting holes, so that after the screw 5 passes through the screw 5 mounting holes, the battery array side plate 3 and the satellite cabin structure plate 4 are fastened together. The first embedded part 8 is adapted to the screw shank 502, so that the end of the screw shank 502 away from the screw head 501 is embedded in the first embedded part 8; the battery array side plate 3 is provided with a screw 5 mounting hole, and a second embedded part 9 is provided in the screw 5 mounting hole, so that after the screw 5 passes through the screw 5 mounting hole, the battery array side plate 3 and the satellite cabin structure plate 4 are fastened together; the axial length of the heat insulation pad 7 is greater than the thickness of the multi-layer heat insulation component 1, so that the heat insulation pad 7 extends in a direction away from the screw head 501 and passes through the second membrane layer 103 of the multi-layer heat insulation component 1; the heat insulation bushing 6 is sleeved on the outer periphery of the screw shank 502 and accommodated in the screw 5 mounting hole; the heat insulation pad 7 covers the screw shank 502.

[0047] The following embodiment provides a temporary heat leakage protection method for satellite thermal testing based on the above-mentioned thermal insulation structure, including the following steps: Provide and cut flexible film 16; Figure 6 A cross-sectional schematic diagram of the solar cell array heat insulation structure mounted on the satellite side panel is shown (F46 thin film disc protection state). The flexible thin film 16 can be cut into a circular disc, i.e., the flexible thin film 16 is a flexible thin film disc, adapted for use with the mounting holes of screw 5; the outer dimensions of the flexible thin film 16 are 1mm to 3mm larger than the inner contour dimensions of the mounting holes of screw 5; it can be made of polytetrafluoroethylene propylene (F46) thin film with silver-plated secondary surface mirror or white paint film, with a thickness of 20~100um. The flexible thin film 16 has a low solar absorptivity α. s ≤0.2, high infrared emissivity ε h Surface properties ≥0.75; The flexible film 16 is embedded in the mounting hole of the screw 5 in an interference fit manner to seal the mounting hole of the screw 5 (the flexible film 16 covers the screw head 501). The edge of the flexible film 16 is radially squeezed by the hole wall of the mounting hole of the screw 5 to produce flexible deformation, and is fixed inside the mounting hole of the screw 5 by the elastic force of the flexible deformation. One-component room temperature vulcanizing silicone rubber 17 is applied in dots at the gap between the edge of the flexible film 16 and the wall of the mounting hole of the screw 5 to form discrete auxiliary fixing adhesive dots. The number of discrete auxiliary fixing adhesive dots between a single flexible film and the hole wall is 3 to 6. The silicone rubber is allowed to dry at room temperature. Using infrared lamp arrays or infrared heating cages to apply a thermal environment, spacecraft thermal balance tests are conducted, and operating temperature data is collected; and After the test, the flexible film 16 was removed from the mounting hole of screw 5, the residual silicone rubber in the hole was removed, and the satellite was restored to its original state.

[0048] A flexible film 16 covers the screw head 501, altering its radiation characteristics and blocking the black hole effect, temporarily simulating the thermal state of silicone rubber filling in orbit during ground tests. Using this method, the additional heat leakage interference from screw 5 during ground thermal tests is effectively suppressed, and the thermal boundary conditions are closer to the actual in-orbit state.

[0049] Specifically, the thermal test verification method is implemented as follows: When in orbit, the mounting hole for screw 5 is filled with silicone rubber; during ground testing, screw 5 is exposed. Figure 1 After the heat insulation pad 7 and the multi-layer heat insulation component 1 block large-area heat transfer, the screw 5 becomes the remaining main thermal bridge. The exposed screw head 501 absorbs heat and heats up under infrared radiation, and the heat is conducted along the screw shank 502 to the satellite cabin structural plate 4, interfering with the evaluation of the heat insulation effect. Figure 6 Temporary protection was implemented using flexible film 16 (F46 disc): Flexible film 16 was cut to a thickness of 50 μm, α... s ≈0.14, ε h ≈0.82, outer diameter φ15mm (hole inner diameter φ12mm). Insert screw 5 into the mounting hole, securing it to the inside of the hole wall via interference fit, covering the screw head 501. The flexible film 16 disc alters the regional radiation characteristics and blocks the black hole effect, temporarily simulating the on-orbit silicone rubber filling state. Subsequently, a small amount of GD414C single-component room temperature vulcanizing silicone rubber 17 is dotted on the circumference where the edge of the flexible film 16 disc intersects with the hole wall, forming discrete auxiliary fixing dots. After the experiment, the disc is removed with tweezers and wiped clean with alcohol.

[0050] Performance verification: To verify the practical effect of the present invention, two isomorphic satellites, A and B, of a certain model were compared. The two satellites had identical panel structures, thermal control coating conditions, and screw specifications (5). During the thermal equilibrium test, the screw holes (5) on the surface of the panel of satellite A's body-mounted solar panels remained exposed without any temporary protective measures. During the thermal equilibrium test of satellite B, the screw holes (5) on the surface of its body-mounted solar panels were temporarily sealed using the F46 thin-film disc protective structure adopted in this embodiment. After the test, the discs were removed, and no irreversible damage was observed.

[0051] In this example, the first satellite did not implement temporary protection with screw 5 during the ground-based infrared thermal balance test, resulting in significant on-orbit deviations in the thermal simulation model corrected based on the test data. The second satellite employed the F46 thin-film disc temporary protection method of this invention during the ground-based infrared thermal balance test. Although the corrected thermal simulation model still showed some residual deviations from the test data after protection, the predicted on-orbit temperature (single-unit temperature on the battery side) of the corrected model showed good agreement with the satellite's on-orbit telemetry data. This indicates that using the disc protection scheme of this invention effectively improves the accuracy of on-orbit prediction.

[0052] (2) On-orbit verification of thermal insulation effect During on-orbit operation, the side panel 3 of the battery array is affected by solar radiation and Earth's infrared radiation, causing the temperature of the solar panel to fluctuate between 0 and 120°C. After the thermal insulation system of this invention is applied, the measured temperature fluctuation of the single unit 19 inside the cabin corresponding to the solar panel side is within 5°C in the initial stage of on-orbit operation, and the average temperature is not higher than 30°C. This indicates that the thermal insulation system effectively blocks the transmission of high temperature and drastic temperature changes from the solar panel to the cabin 19.

[0053] (3) On-orbit verification of insulation effect The insulation design between the solar panels and the platform ensures good insulation performance of the satellite during its orbit, and the on-orbit power supply and distribution system and individual units all function normally.

[0054] General scenario description: In this embodiment, the area of ​​the body-mounted solar panel is exposed to radiation for a long time in orbit, and the external heat flow changes drastically, amplifying the heat leakage effect of the screw holes 5. Therefore, the protection scheme of this invention is particularly effective. For areas with shorter on-orbit radiation time and relatively stable external heat flow (such as the dark side of the celestial body), although this scheme can still reduce heat leakage, the effect is less significant. In practical engineering applications, it is recommended to prioritize the implementation of this protection scheme for the screw holes in areas with long on-orbit radiation time, such as body-mounted solar panels and exposed cabin panels, to maximize cost-effectiveness.

[0055] In some embodiments, one side of the flexible film 16 may be selectively laminated with an adhesive backing layer; when the adhesive backing layer is laminated, the thickness of the adhesive backing layer participates in the thickness filling to form an interference fit and provides initial anti-slip friction during the initial embedding stage.

[0056] In some embodiments, the middle portion of the flexible film 16 may make small-area contact with the end of the screw 5 without generating adhesive force.

[0057] In some embodiments, as a further improvement to the above embodiments, a silicone rubber layer is coated on the side of the F46 film disc facing the screw head 501, so that the surface radiation characteristics of the disc in the screw hole 5 are closer to the silicone rubber filling state.

[0058] Implementation method: A layer of room temperature vulcanizing silicone rubber is evenly coated on the outward-facing side of the disc using a scraping method, and then allowed to cure at room temperature. The coated disc retains its flexibility and can be inserted into the screw hole 5 by interference fit and mechanically secured to the inside of the hole wall.

[0059] Beneficial effects: The solar absorptivity and infrared emissivity of the coated surface are closer to the state of silicone rubber filling in orbit, which makes the thermal radiation state of the screw head 501 area in the ground test closer to the real state in orbit, thereby more accurately simulating the thermal state of the screw 5 hole filled with silicone rubber in orbit and improving the accuracy of the simulation model's in-orbit prediction.

[0060] While some embodiments of the present invention have been described in this application, those skilled in the art will understand that these embodiments are merely illustrative. Numerous variations, alternatives, and improvements will arise in those skilled in the art under the teachings of this invention without departing from its scope. The appended claims are intended to define the scope of the invention and thereby cover methods and structures within the scope of the claims themselves and their equivalents.

Claims

1. A heat insulation structure for a satellite side panel with a bulk-mounted solar array, wherein the bulk-mounted solar array can be disposed on the outer surface of the satellite side panel facing outwards, so that the satellite side panel serves as the side panel of the solar array, characterized in that, The thermal insulation structure includes: A multi-layer thermal insulation assembly, which covers the inner surface of the battery array side panel facing the interior of the cabin and is configured to block the heat radiation from the battery array side panel from being transmitted to the satellite cabin structural panel, and electrically isolates the multi-layer thermal insulation assembly from the battery array side panel; and Multiple connectors are configured to connect the battery array side panels to the satellite cabin structural panels.

2. The satellite side plate-mounted solar cell array heat insulation structure according to claim 1, characterized in that, From the side closest to the battery array side panel towards the interior of the cabin, the multi-layer thermal insulation assembly sequentially includes: The first membrane layer is configured to provide insulation protection between the multilayer thermal insulation assembly and the battery array side panel; Thermal insulation units, configured to block radiative heat transfer from the battery array side panels to the satellite cabin structural panels; and The second membrane layer is configured to reduce heat exchange with the cabin's radiative interior.

3. The satellite side plate-mounted solar cell array heat insulation structure according to claim 2, characterized in that, The heat insulation unit includes: The reflective screen is configured to attenuate radiative heat flux in a vacuum environment, increase radiative heat transfer resistance, and suppress radiative heat transfer between multilayer insulation components; and The spacer layer is configured to isolate adjacent reflective screens, prevent direct contact between reflective screens, maintain the vacuum gap between layers, suppress solid heat conduction, and ensure the heat insulation performance of the multilayer heat insulation component; the spacer layer and reflective screen are stacked alternately in sequence from the first film layer to the second film layer.

4. The satellite side plate-mounted solar cell array heat insulation structure according to claim 3, characterized in that, The first film layer is a 50-60μm polyimide film; the reflective screen is a 5-6μm perforated double-sided aluminized polyester film; the spacer layer is a polyester mesh; the second film layer is a perforated 16-20μm single-sided aluminized polyester film, with the aluminized side facing inward.

5. The satellite side plate-mounted solar cell array heat insulation structure according to claim 3, characterized in that, The battery array side plate has mounting holes for connectors, and the connectors include: The connector end is located on the side panel of the battery array near the interior of the compartment; and The connecting rod passes through the connecting mounting hole and extends into the interior of the satellite cabin structural plate.

6. The satellite side plate-mounted solar cell array heat insulation structure according to claim 5, characterized in that, Also includes: Multiple heat-insulating bushings are provided at one end of the connector rod near the connector end and are configured to block the solid heat conduction channel between the battery array side plate and the satellite cabin structural plate, while electrically insulating the connector rod from the battery array side plate.

7. The satellite side plate-mounted solar cell array heat insulation structure according to claim 5, characterized in that, Also includes: Multiple heat insulation pads are provided around the connector rod and configured to block solid heat conduction channels between the battery array side plate and the satellite cabin structural plate. Holes are provided along the thickness direction of the multi-layer thermal insulation component, and the holes communicate with the mounting holes of the connectors, so that when the connectors are inserted through the holes, the multi-layer thermal insulation component surrounds the thermal insulation pad circumferentially.

8. The satellite side plate-mounted solar cell array heat insulation structure according to claim 7, characterized in that, The edges of the aperture are bound with polyimide single-sided adhesive so that the polyimide single-sided adhesive covers the reflective screen; and / or The first film layer is attached to the side plate of the battery array using polyimide double-sided adhesive; the first film layer extends 10-15 mm beyond the heat insulation unit and the second film layer.

9. The satellite side plate-mounted solar cell array heat insulation structure according to claim 3, characterized in that, Also includes: The grounding wire has one end electrically connected to the heat insulation unit and the second membrane layer, and the other end connected to the grounding terminal of the satellite cabin structure plate.

10. A temporary protection method for heat leakage during satellite thermal testing based on the thermal insulation structure described in any one of claims 1 to 9, characterized in that, Includes the following steps: Provide and cut flexible films; The flexible film is embedded into the connector mounting hole in an interference fit manner to seal the connector mounting hole. The edge of the flexible film is radially squeezed by the hole wall of the connector mounting hole to produce flexible deformation. It is then fixed inside the connector mounting hole by the elastic force of the flexible deformation. One-component room temperature vulcanizing silicone rubber is applied in dots at the gap between the edge of the flexible film and the wall of the mounting hole of the connector to form discrete auxiliary fixing adhesive dots, and left to stand at room temperature until the silicone rubber is surface dry. Using infrared lamp arrays or infrared heating cages to apply a thermal environment, spacecraft thermal balance tests are conducted, and operating temperature data is collected; and After the test, the flexible film was removed from the mounting hole of the connector, the residual silicone rubber in the hole was removed, and the satellite was restored to its original state.

Images