CN112693637B - Thermal control method of hollow frame spacecraft - Google Patents
Thermal control method of hollow frame spacecraft Download PDFInfo
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- CN112693637B CN112693637B CN202011607171.2A CN202011607171A CN112693637B CN 112693637 B CN112693637 B CN 112693637B CN 202011607171 A CN202011607171 A CN 202011607171A CN 112693637 B CN112693637 B CN 112693637B
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Abstract
The invention relates to the technical field of aerospace, in particular to a thermal control method of a hollow frame spacecraft, which comprises the following steps: spraying black paint on the outer surface of the equipment to enable the surface infrared emissivity to be larger than 0.85; mounting a polyimide germanium-plated film on the outer side of the back sun surface of the cuboid hollow frame; and 15 units of heat insulation layers are coated on the outer side of the sunny side of the cuboid hollow frame. According to the method, the structural characteristics of the hollow-frame spacecraft and the direct infrared radiation heat dissipation property of the equipment are utilized, so that the cuboid hollow-frame and the equipment arranged in the frame can be always in a better temperature range, and the weight and the implementation difficulty of the spacecraft device are reduced.
Description
Technical Field
The invention relates to the technical field of aerospace, in particular to a thermal control method of a hollow frame spacecraft.
Background
The thermal control of the spacecraft controls internal heat transfer and external heat exchange, and ensures that the temperature of internal equipment/instruments of the spacecraft is kept within a specific working temperature range, thereby realizing normal working and long-life running of the equipment.
The conventional spacecraft is provided with a complete structural cabin plate as a bearing body of a radiating surface design, and internal equipment mainly adopts a heat pipe or a heat conducting structure to conduct heat to the radiating surface of the cabin plate and radiate the heat outwards through the characteristic of high infrared emissivity of the surface of the cabin plate. For the spacecraft with the frame hollow structure, the structure heat conduction capability is poor, and the surface of the incomplete structure is not used as a carrier for designing a heat dissipation surface. The conventional thermal control method of the spacecraft is to enhance the heat conduction between the equipment installation surface and the heat dissipation surface through a heat pipe or a metal structure, fill part of the hollow area of the hollow structure plate to be used as a design carrier of the heat dissipation surface, and spray a thermal control coating with low solar absorption ratio and high infrared emissivity on the filled structure plate. The conventional thermal design of the hollow-out structure spacecraft only emphasizes on considering the utilization structure as a radiating surface, and the radiation radiating capacity of the equipment and the change of the radiation angle coefficient of the hollow-out structure to a deep-air cooling background are not fully utilized. The conventional thermal control method brings the increase of the weight of the spacecraft, the rise of the development cost of thermal control and the complexity of the final assembly of the spacecraft. At the moment, a new thermal control method is needed, so that the requirement of the temperature index of the internal equipment of the spacecraft is met, and the weight, the cost and the final assembly implementation difficulty of the spacecraft are reduced.
Disclosure of Invention
In order to solve the technical problems, the invention provides a thermal control method of a hollow-frame spacecraft, which utilizes the structural characteristics of the hollow-frame spacecraft and the direct infrared radiation heat dissipation property of equipment to enable a cuboid hollow frame and the equipment arranged in the frame to be in a better temperature range all the time, thereby reducing the weight and the implementation difficulty of a spacecraft device.
In order to achieve the purpose, the invention adopts the following technical scheme:
a thermal control method of a hollow frame spacecraft is provided, a spacecraft device used by the thermal control method comprises a cuboid hollow frame and equipment arranged in the frame, and the thermal control method is characterized in that:
(1) spraying black paint on the outer surface of the equipment to enable the surface infrared emissivity to be larger than 0.85;
(2) mounting a polyimide germanium-plated film on the outer side of the back sun surface of the cuboid hollow frame;
(3) and 15 units of heat insulation layers are coated on the outer side of the sunny side of the cuboid hollow frame.
Further, the cuboid hollow-out frame adopts an aluminum alloy plate with the thickness of 2mm to 3 mm.
Furthermore, the heat insulation layer of each unit comprises a layer of double-sided aluminized film and a layer of polyester net.
Further, the heat dissipation capacity of the equipment is
Wherein Q is the heat dissipation capacity of the equipment (W/square meter); epsilon is the infrared emissivity of the equipment; σ is the radiation constant of the black body (stefan-boltzmann constant); lambda is the infrared transmittance of the polyimide germanium-plated film; s is the available radiation surface area of the equipment; t is the absolute temperature of the equipment.
Furthermore, the area of the polyimide germanium coating film attached to the back and the front side is reduced according to the heat dissipation capacity value Q, and the area of the reduced area of the polyimide germanium coating film attached to the back and the front side is replaced by an F46 film or 15 unit heat insulation layers.
The invention has the following beneficial effects:
1. spraying a high-infrared-emissivity coating on equipment in the spacecraft, and radiating heat to a deep-air cooling background by using the coating to radiate the infrared radiation to the outside;
2. heat is dissipated to the outer space through the polyimide germanium-plated film with certain infrared transmittance, and short-term direct sunlight and planet albedo are blocked by utilizing the characteristic of low solar absorption ratio of the polyimide germanium-plated film;
3. the heat insulation material is arranged on the outer side of the structural frame on the side of the spacecraft facing the sun to insulate heat, so that the requirement on the temperature index of equipment in the spacecraft is met, and the weight, the cost and the implementation difficulty of the spacecraft are reduced.
Drawings
FIG. 1 is an overall structural diagram of the hollow frame spacecraft of the present invention;
reference numerals: 1-equipment, 2-polyimide germanium film plating and 3-hollow frame.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments.
Referring to fig. 1, a thermal control method for a hollow frame spacecraft, the spacecraft device used in the thermal control method comprises a cuboid hollow frame 3 and equipment 1 installed inside the frame:
(1) spraying black paint on the outer surface of the equipment 1 to enable the surface infrared emissivity to be larger than 0.85;
(2) a polyimide germanium-plated film 2 is pasted on the outer side of the back sun surface of the cuboid hollow frame 3;
(3) and 15 units of heat insulation layers are coated on the outer side of the sunny side of the cuboid hollow frame 3.
In one embodiment of the invention, the outer side of one of six faces of a cuboid hollow frame 3 is coated with a heat insulation layer with 15 units, and the outer sides of the other five faces are pasted with polyimide germanium-plated films 2; the cuboid hollow frame 3 is made of an aluminum alloy plate with the thickness of 2mm to 3 mm; the heat insulation layer of each unit comprises a layer of double-sided aluminized film and a layer of terylene net.
The heat dissipation capacity of the equipment 1 arranged in the spacecraft hollow frame 3 mainly comprises two parts, on one hand, heat is conducted to the outer side of the spacecraft hollow frame 3 through the self heat conduction capacity of the spacecraft hollow frame 3 and is radiated to a deep air cooling background, and the part is a conventional heat dissipation mode; on one hand, the high radiation characteristic of the equipment 1 directly radiates heat to a deep-air cooling background through the polyimide germanium-coated film 2 with certain infrared transmittance, the part is a heat radiation mode specific to the structure of the hollow frame 3, and the heat radiation capability is
Wherein Q is the heat dissipation capacity (W/square meter) of the equipment 1; epsilon is the infrared emissivity of the device 1; σ is the radiation constant of the black body (stefan-boltzmann constant); lambda is the infrared transmittance of the polyimide germanium-plated film 2; s is the available radiation surface area of the device 1; t is the absolute temperature of the apparatus 1.
In one embodiment of the invention, the area for mounting the polyimide germanium-plating film 2 is reduced according to the heat dissipation force value Q, when the heat dissipation force value Q is too large, the area for mounting the polyimide germanium-plating film 2 on the back and the sun is reduced, and the area for mounting the polyimide germanium-plating film 2 is replaced by F46 films or 15 unit heat insulation layers.
In an embodiment of the invention, the region of the back and forth surface with the reduced area for mounting the polyimide germanium film 2 can be one of the back and forth surfaces of the cuboid hollow frame 3 or a part of one of the back and forth surfaces.
According to the embodiment of the invention, through calculation, when the temperature of the equipment is kept at 35 ℃, the heat dissipation capacity of the equipment to the space through the polyimide germanium film 2 can reach 88W/square meter, the coating quantity of the heat insulation layer to the sun is 15 units, and the polyimide germanium film 2 is pasted on the outer sides of the five back light surfaces of the cuboid hollow frame 3, so that the method is proved to have a great auxiliary heat dissipation effect under the condition that the heat conduction capacity of the body of the hollow frame spacecraft is insufficient.
The above-mentioned embodiments are intended to illustrate the objects, technical solutions and advantages of the present invention in further detail, and it should be understood that the above-mentioned embodiments are merely exemplary embodiments of the present invention, and are not intended to limit the scope of the present invention, and any modifications, equivalent substitutions, improvements and the like made within the spirit and principle of the present invention should be included in the scope of the present invention.
Claims (5)
1. A thermal control method of a hollow frame spacecraft is provided, a spacecraft device used by the thermal control method comprises a cuboid hollow frame and equipment arranged in the frame, and the thermal control method is characterized in that:
(1) spraying black paint on the outer surface of the equipment to enable the infrared emissivity of the surface of the equipment to be larger than 0.85;
(2) mounting a polyimide germanium-plated film on the outer side of the back sun surface of the cuboid hollow frame;
(3) and 15 units of heat insulation layers are coated on the outer side of the sunny side of the cuboid hollow frame.
2. The thermal control method of the hollow-frame spacecraft of claim 1, wherein the rectangular hollow frame is made of an aluminum alloy plate with the thickness of 2mm to 3 mm.
3. The thermal control method of the hollow-frame spacecraft of claim 1, wherein the thermal insulation layer of each unit comprises a double-sided aluminized film and a polyester net.
4. The thermal control method of the hollow-frame spacecraft of claim 1, wherein the heat dissipation capacity of the equipment is
Wherein Q is the heat dissipation capacity of the equipment, W/square meter; epsilon is the infrared emissivity of the equipment; σ is blackStefan-boltzmann constants of the body; lambda is the infrared transmittance of the polyimide germanium-plated film; s is the available radiation surface area of the equipment; t is the absolute temperature of the equipment.
5. The thermal control method of the hollow-frame spacecraft of claim 4, wherein the area of the polyimide germanium coating film attached to the back and the sun is reduced according to the heat dissipation force value Q, and the area of the reduced polyimide germanium coating film is replaced by an F46 film or 15 unit heat insulation layers.
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| Application Number | Priority Date | Filing Date | Title |
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| CN202011607171.2A CN112693637B (en) | 2020-12-30 | 2020-12-30 | Thermal control method of hollow frame spacecraft |
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| CN202011607171.2A CN112693637B (en) | 2020-12-30 | 2020-12-30 | Thermal control method of hollow frame spacecraft |
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| CN112693637B true CN112693637B (en) | 2022-04-22 |
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Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105659752B (en) * | 2010-11-19 | 2013-10-23 | 北京空间飞行器总体设计部 | Without the heat control method of attitude satellite |
| CN106481932A (en) * | 2016-09-08 | 2017-03-08 | 上海卫星工程研究所 | Wave transparent type multilayer insulation material structure and preparation method |
| CN109648971A (en) * | 2019-01-09 | 2019-04-19 | 上海卫星工程研究所 | A kind of space heat controlled thin film |
| JP2020001419A (en) * | 2018-06-25 | 2020-01-09 | 日本電気株式会社 | Heat control device |
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2020
- 2020-12-30 CN CN202011607171.2A patent/CN112693637B/en active Active
Patent Citations (4)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN105659752B (en) * | 2010-11-19 | 2013-10-23 | 北京空间飞行器总体设计部 | Without the heat control method of attitude satellite |
| CN106481932A (en) * | 2016-09-08 | 2017-03-08 | 上海卫星工程研究所 | Wave transparent type multilayer insulation material structure and preparation method |
| JP2020001419A (en) * | 2018-06-25 | 2020-01-09 | 日本電気株式会社 | Heat control device |
| CN109648971A (en) * | 2019-01-09 | 2019-04-19 | 上海卫星工程研究所 | A kind of space heat controlled thin film |
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| CN112693637A (en) | 2021-04-23 |
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