A spacecraft fixed solar wing heat dissipating device

Abstract

The application provides a heat dissipation device for a spacecraft fixed solar wing, comprising insulating material (1), a heat conducting film (2) and a secondary surface mirror (3), the insulating material (1) is attached to the back of a spacecraft fixed solar wing heat dissipation area, one end of the heat conducting film (2) is attached to the insulating material (1), the other end of the heat conducting film (2) is attached to the side of the spacecraft body which is not directly irradiated by the sun, and the secondary surface mirror (3) is attached to the heat conducting film (2). The heat dissipation device of the application can cool the fixed solar wing through the flexible graphene film without changing the body structure design, can meet the temperature index requirement after the number of fixed solar wing cloth sheets is reduced, and can make the solar wing output voltage meet specific requirements.

Description

Technical Field

[0001] This invention relates to the field of heat dissipation devices for spacecraft solar panels, and more particularly to a heat dissipation device for fixing spacecraft solar panels. Background Technology

[0002] ① Existing technical description:

[0003] Spacecraft fixed solar panels typically use heat-insulating materials to reduce their impact on the spacecraft's internal electronic equipment, and only radiate heat to space through the surface of the solar cells to ensure that the temperature of the fixed solar panels meets certain requirements.

[0004] The solar array radiates heat into space only through the solar cells on the front of the solar array, resulting in low heat dissipation efficiency. This causes the temperature of the fixed solar array to remain at a high level, typically reaching 120°C. Under such high temperatures, the number of solar cells needs to be increased to ensure a certain output voltage requirement, which also imposes greater constraints on the selection of materials for the solar array structure.

[0005] For fixed solar panels on spacecraft, their backs are shielded by the spacecraft body, blocking the channel for heat dissipation through radiation from the back of the fixed solar panels into the deep, cold space. The conventional heat dissipation method for such fixed solar panels is to radiate heat to space through the surface of the solar cells on the front of the solar panels, and to attach thermal control films with low solar absorptivity and high infrared emissivity between the solar cells.

[0006] ② Description of existing technical problems and defects:

[0007] Conventional thermal design for spacecraft fixed solar arrays focuses solely on utilizing the surface of the solar cells as a radiative heat dissipation surface, neglecting to fully leverage the radiative heat dissipation capabilities of other sides of the spacecraft and the ease of installation using flexible, highly thermally conductive materials. This conventional thermal design approach leads to an increase in the number of solar panels on the spacecraft's fixed solar array, higher requirements for the high-temperature resistance of the solar array substrate, and a larger area of ​​the fixed solar array. Therefore, a new thermal control design method is needed to both meet the electrical parameter output requirements of the spacecraft's fixed solar array and reduce its weight, size, and the special high-temperature processing requirements. Summary of the Invention

[0008] The present invention provides a heat dissipation device for fixing the solar array of a spacecraft, comprising an insulating material, a thermally conductive film, and a secondary surface mirror. The insulating material is attached to the back of the area of ​​the spacecraft's fixed solar array that needs to dissipate heat. One end of the thermally conductive film is attached to the insulating material, and the other end of the thermally conductive film is attached to other sides of the spacecraft body that are not directly exposed to sunlight. The secondary surface mirror is attached to the thermally conductive film.

[0009] As a further improvement of the invention, the heat dissipation device also includes a load-bearing member for providing load-bearing support to the heat-conducting film at the corner where the spacecraft's fixed solar array turns toward other sides of the spacecraft body.

[0010] As a further improvement of the present invention, the heat dissipation device also includes a multi-layer heat insulation component, which is installed between the heat-conducting film and the spacecraft body.

[0011] As a further improvement of the present invention, the insulating material is a polyimide film.

[0012] As a further improvement of the present invention, the polyimide film has a thickness of 20 μm.

[0013] As a further improvement of the present invention, the thermally conductive film is a graphene film; the secondary surface mirror is an F46 silver-plated secondary surface mirror.

[0014] As a further improvement of the present invention, the graphene film is a multilayer composite graphene film.

[0015] As a further improvement of the present invention, the multilayer composite graphene film is a 12-layer composite graphene film; the thermal conductivity of a single layer of the multilayer composite graphene film is greater than 1400 W / m² / ℃.

[0016] As a further improvement of the present invention, the load-bearing component is carbon nanotube copper foil.

[0017] As a further improvement of the present invention, the multilayer heat insulation component is composed of double-sided polyester film and polyester mesh overlapping each other.

[0018] The beneficial effects of the present invention are: the heat dissipation device of the present invention, without changing the main body structure design, cools the fixed solar panel by means of a flexible graphene film, which can meet the temperature index requirements after the number of fixed solar panel sheets is reduced, and allows the solar panel output voltage to meet specific requirements. Attached Figure Description

[0019] Figure 1 This is a structural diagram of the heat dissipation device of the present invention. Detailed Implementation

[0020] This invention discloses a heat dissipation device for fixing a solar array on a spacecraft, comprising an insulating material 1, a thermally conductive film 2, and a secondary surface mirror 3. The insulating material 1 is attached to the back of the area of ​​the spacecraft's fixed solar array 6 that needs heat dissipation. One end of the thermally conductive film 2 is attached to the insulating material 1, and the other end of the thermally conductive film 2 is attached to other sides of the spacecraft body that are not directly exposed to sunlight. The secondary surface mirror 3 is attached to the thermally conductive film 2.

[0021] The heat dissipation device also includes a load-bearing component 4, which provides load-bearing support to the heat-conducting film 2 at the corner where the spacecraft fixes the solar array 6 to other sides of the spacecraft body.

[0022] The load-bearing component is carbon nanotube copper foil.

[0023] The heat dissipation device also includes a multi-layer heat insulation component 5, which is installed between the heat-conducting film 2 and the spacecraft body. The multi-layer heat insulation component 5 is composed of double-sided polyester film and polyester mesh. Specifically, a single-layer heat insulation component is formed by one layer of double-sided polyester film and one layer of polyester mesh (spacer material), and 15 single-layer heat insulation components are stacked together to form the multi-layer heat insulation component 5.

[0024] The insulating material 1 is a polyimide film.

[0025] The polyimide film 1 has a thickness of 20 μm.

[0026] The thermally conductive film 2 is a graphene film.

[0027] The secondary surface mirror 3 is an F46 silver-plated secondary surface mirror.

[0028] The graphene film is a 12-layer composite graphene film, and the thermal conductivity of a single layer of the composite graphene film is greater than 1400 W / m² / ℃.

[0029] 1. Principle Introduction: A polyimide film is attached to the back of the heat-dissipating area of ​​the spacecraft's fixed solar array 6 to ensure insulation. A flexible, highly thermally conductive multilayer composite graphene film is then attached. Utilizing its high thermal conductivity, the graphene film conducts heat from the fixed solar array 6 to other sides of the spacecraft with lower external heat flow. The heat is then dissipated into outer space by F46 silver-plated secondary surface mirrors with low solar absorptivity and high infrared emissivity, which are attached to the graphene film. Simultaneously, 15 single-layer thermal insulation components composed of double-sided polyester film and polyester mesh are used to insulate the graphene film from the spacecraft structure, preventing the heat from the fixed solar array 6 from affecting the temperature of the spacecraft. Carbon nanotube copper foil is used for load-bearing at the corners where the graphene film turns from the fixed solar array 6 to other sides of the spacecraft. This method satisfies the electrical parameter output requirements of the fixed solar array 6 while reducing the weight, size, and high-temperature resistance requirements of the fixed solar array.

[0030] 2. Specific measures are as follows: A 20μm thick polyimide film is attached to the back of the area requiring heat dissipation of the spacecraft's fixed solar array 6, followed by a flexible, highly thermally conductive 12-layer composite graphene film. The thermal conductivity of a single-layer graphene film is greater than 1400W / ㎡ / ℃. Utilizing its high thermal conductivity, the heat from the fixed solar array 6 is conducted to other sides of the spacecraft that are not directly exposed to sunlight. The heat is then dissipated into outer space through F46 silver-plated secondary surface mirrors with low solar absorptivity and high infrared emissivity. Simultaneously, 15 single-layer heat insulation components composed of double-sided polyester film and polyester mesh are used to insulate the graphene film from the spacecraft's main structure, preventing temperature changes in the fixed solar array 6 from affecting the temperature of the main body. Carbon nanotube copper foil is used as a load-bearing element at the corners where the graphene film turns from the fixed solar array 6 to other sides of the main body.

[0031] 3. Implementation Principle: The heat dissipation capability of the spacecraft's fixed solar array 6 mainly consists of two parts. Firstly, heat is radiated to the deep-space cool background through the surface of the solar cells on the front of the solar array; this is a conventional heat dissipation method. Secondly, the heat is conducted to the multi-layered outer surface of other sides of the spacecraft body through the thermal conductivity of the flexible, highly thermally conductive graphene film, and then radiated to the deep-space cool background through the F46 silver-plated secondary surface mirrors. This part utilizes the unique enhanced heat dissipation method of the flexible, highly thermally conductive graphene film, and its radiative heat dissipation capability is:

[0032]

[0033] Where: q is the radiative heat dissipation capacity of this thermal control film combination (graphene film + carbon nanotube copper foil + F46 silver-plated secondary surface mirror), in W / m²;

[0034] ε is the infrared emissivity of the F46 silver-plated secondary surface mirror;

[0035] σ is the radiation constant of a blackbody (Stephen-Boltzmann constant);

[0036] S represents the radiation surface area of ​​the F46 silver-plated secondary surface mirror;

[0037] T represents the absolute temperature of the combined thermal control film, in Kelvin (K).

[0038] In the composite thermal control film, the main thermal conductive component is a flexible, highly thermally conductive graphene film, whose thermal conductivity is:

[0039]

[0040] Where: Q represents the thermal conductivity and heat dissipation capacity of the graphene film, in W;

[0041] K represents the thermal conductivity of the graphene film, in W / m² / ℃.

[0042] L represents the heat conduction path distance, in meters (m).

[0043] S represents the cross-sectional area of ​​the thermal conductivity path of the graphene film, in square meters (m²).

[0044] ΔT is the temperature difference between the hot and cold ends of the graphene film, in °C.

[0045] Analysis shows that the flexible, highly thermally conductive graphene film greatly assists in the heat dissipation of the spacecraft's fixed solar array 6, effectively reducing the solar array temperature to achieve the goal of reducing the number of fabric pieces and the area of ​​the solar array.

[0046] The beneficial effects of the present invention are: the heat dissipation device of the present invention, without changing the main body structure design, cools the fixed solar panel by means of a flexible graphene film, which can meet the temperature index requirements after the number of fixed solar panel sheets is reduced, and allows the solar panel output voltage to meet specific requirements.

[0047] The above description, in conjunction with specific preferred embodiments, provides a further detailed explanation of the present invention. It should not be construed that the specific implementation of the present invention is limited to these descriptions. For those skilled in the art, various simple deductions or substitutions can be made without departing from the concept of the present invention, and all such modifications and substitutions should be considered within the scope of protection of the present invention.

Claims

1. A heat dissipation device for fixing solar panels on a spacecraft, characterized in that: The device includes an insulating material (1), a thermally conductive film (2), and a secondary surface mirror (3). The insulating material (1) is attached to the back of the area of ​​the spacecraft's fixed solar array (6) that needs to dissipate heat. One end of the thermally conductive film (2) is attached to the insulating material (1), and the other end of the thermally conductive film (2) is attached to other sides of the spacecraft body that are not directly exposed to sunlight. The secondary surface mirror (3) is attached to the thermally conductive film (2). The heat dissipation device also includes a load-bearing component (4), which is used to provide load-bearing force to the heat-conducting film (2) at the corner where the spacecraft fixes the solar panel (6) turns toward other sides of the spacecraft body; the heat dissipation device also includes a multi-layer heat insulation component (5), which is installed between the heat-conducting film (2) and the spacecraft body.

2. The heat dissipation device according to claim 1, characterized in that: The insulating material (1) is a polyimide film.

3. The heat dissipation device according to claim 2, characterized in that: The polyimide film (1) has a thickness of 20 μm.

4. The heat dissipation device according to claim 1, characterized in that: The thermally conductive film (2) is a graphene film; the secondary surface mirror (3) is an F46 silver-plated secondary surface mirror.

5. The heat dissipation device according to claim 4, characterized in that: The graphene film is a multilayer composite graphene film.

6. The heat dissipation device according to claim 5, characterized in that: The multilayer composite graphene film is a 12-layer composite graphene film; the thermal conductivity of a single layer of the multilayer composite graphene film (2) is greater than 1400W / ㎡ / ℃.

7. The heat dissipation device according to claim 1, characterized in that: The load-bearing component is carbon nanotube copper foil.

8. The heat dissipation device according to claim 1, characterized in that: The multilayer thermal insulation component (5) is composed of double-sided polyester film and spacer material; the spacer material is polyester mesh.

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