CN114408221A - Satellite-used satellite-sensitive temperature control system - Google Patents
Satellite-used satellite-sensitive temperature control system Download PDFInfo
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- CN114408221A CN114408221A CN202210060731.XA CN202210060731A CN114408221A CN 114408221 A CN114408221 A CN 114408221A CN 202210060731 A CN202210060731 A CN 202210060731A CN 114408221 A CN114408221 A CN 114408221A
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- 238000009413 insulation Methods 0.000 claims abstract description 50
- 229910052684 Cerium Inorganic materials 0.000 claims abstract description 34
- GWXLDORMOJMVQZ-UHFFFAOYSA-N cerium Chemical compound [Ce] GWXLDORMOJMVQZ-UHFFFAOYSA-N 0.000 claims abstract description 34
- 239000011521 glass Substances 0.000 claims abstract description 34
- 230000017525 heat dissipation Effects 0.000 claims abstract description 24
- BASFCYQUMIYNBI-UHFFFAOYSA-N platinum Chemical compound [Pt] BASFCYQUMIYNBI-UHFFFAOYSA-N 0.000 claims description 23
- 229910052697 platinum Inorganic materials 0.000 claims description 11
- 238000009434 installation Methods 0.000 claims description 10
- 239000004519 grease Substances 0.000 claims description 9
- 229920001296 polysiloxane Polymers 0.000 claims description 8
- 239000007769 metal material Substances 0.000 claims description 6
- 238000001704 evaporation Methods 0.000 claims description 4
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- 229910000838 Al alloy Inorganic materials 0.000 claims description 3
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 3
- QVRYCILXIXLVCU-UHFFFAOYSA-N N.[AlH3] Chemical compound N.[AlH3] QVRYCILXIXLVCU-UHFFFAOYSA-N 0.000 claims description 3
- 239000004917 carbon fiber Substances 0.000 claims description 3
- 239000011152 fibreglass Substances 0.000 claims description 3
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 3
- 229910052755 nonmetal Inorganic materials 0.000 claims description 3
- 125000006850 spacer group Chemical group 0.000 claims description 3
- 238000000034 method Methods 0.000 description 19
- 238000013461 design Methods 0.000 description 10
- 238000005485 electric heating Methods 0.000 description 5
- 238000005259 measurement Methods 0.000 description 5
- 238000010438 heat treatment Methods 0.000 description 4
- 230000005855 radiation Effects 0.000 description 4
- 238000012360 testing method Methods 0.000 description 4
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- 238000002048 anodisation reaction Methods 0.000 description 1
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- 238000009529 body temperature measurement Methods 0.000 description 1
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- B—PERFORMING OPERATIONS; TRANSPORTING
- B64—AIRCRAFT; AVIATION; COSMONAUTICS
- B64G—COSMONAUTICS; VEHICLES OR EQUIPMENT THEREFOR
- B64G1/00—Cosmonautic vehicles
- B64G1/22—Parts of, or equipment specially adapted for fitting in or to, cosmonautic vehicles
- B64G1/24—Guiding or controlling apparatus, e.g. for attitude control
- B64G1/36—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors
- B64G1/361—Guiding or controlling apparatus, e.g. for attitude control using sensors, e.g. sun-sensors, horizon sensors using star sensors
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01C—MEASURING DISTANCES, LEVELS OR BEARINGS; SURVEYING; NAVIGATION; GYROSCOPIC INSTRUMENTS; PHOTOGRAMMETRY OR VIDEOGRAMMETRY
- G01C21/00—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00
- G01C21/02—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by astronomical means
- G01C21/025—Navigation; Navigational instruments not provided for in groups G01C1/00 - G01C19/00 by astronomical means with the use of startrackers
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- H—ELECTRICITY
- H05—ELECTRIC TECHNIQUES NOT OTHERWISE PROVIDED FOR
- H05K—PRINTED CIRCUITS; CASINGS OR CONSTRUCTIONAL DETAILS OF ELECTRIC APPARATUS; MANUFACTURE OF ASSEMBLAGES OF ELECTRICAL COMPONENTS
- H05K7/00—Constructional details common to different types of electric apparatus
- H05K7/20—Modifications to facilitate cooling, ventilating, or heating
-
- H—ELECTRICITY
- H05—ELECTRIC TECHNIQUES NOT OTHERWISE PROVIDED FOR
- H05K—PRINTED CIRCUITS; CASINGS OR CONSTRUCTIONAL DETAILS OF ELECTRIC APPARATUS; MANUFACTURE OF ASSEMBLAGES OF ELECTRICAL COMPONENTS
- H05K7/00—Constructional details common to different types of electric apparatus
- H05K7/20—Modifications to facilitate cooling, ventilating, or heating
- H05K7/2029—Modifications to facilitate cooling, ventilating, or heating using a liquid coolant with phase change in electronic enclosures
- H05K7/20336—Heat pipes, e.g. wicks or capillary pumps
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- Engineering & Computer Science (AREA)
- Physics & Mathematics (AREA)
- Remote Sensing (AREA)
- Radar, Positioning & Navigation (AREA)
- Microelectronics & Electronic Packaging (AREA)
- Thermal Sciences (AREA)
- Automation & Control Theory (AREA)
- Astronomy & Astrophysics (AREA)
- General Physics & Mathematics (AREA)
- Chemical & Material Sciences (AREA)
- Combustion & Propulsion (AREA)
- Aviation & Aerospace Engineering (AREA)
- Navigation (AREA)
Abstract
The invention provides a satellite-used satellite-sensitive temperature control system, which comprises a satellite sensor, a satellite-sensitive bracket, a light shield, a cerium glass secondary surface mirror, a heat pipe, a plurality of layers, a heat insulation pad, a satellite radiating surface and the like; the star sensor head is arranged on the star sensor support in a heat insulation manner, one end of the heat pipe is arranged on the star sensor head and is tightly attached to the star sensor head, and the other end of the heat pipe extends to the position of the heat dissipation surface; the inner part of the heat pipe cabin is coated with a plurality of layers, the part of the radiating surface of the cabin, which is opposite to the heat pipe, is coated with a plurality of layers and is suitable for a heat insulation pad to insulate heat from the heat pipe, and a cerium glass secondary surface lens is adhered to the outer surface of the heat pipe, wherein the absorptivity and emissivity of the cerium glass secondary surface lens are fixed, so that the heat pipe has a completely independent and stable radiating surface. The length of the heat pipe is adjusted according to the size of the star sensitive heat loss so as to achieve the purpose of changing the radiating surface. Therefore, the star sensor temperature can be controlled within a proper temperature range by fewer resources, and the precision reaches 1 ℃.
Description
Technical Field
The invention relates to the technical field of aerospace, in particular to a satellite-based temperature control system, and particularly relates to a satellite-based temperature control method with high applicability.
Background
With the development of aerospace technology, the breadth and depth of deep space exploration performed by human beings are continuously increased, various types of satellite sensors are more and more applied to spacecrafts, the satellite sensors are subjected to complicated and changeable severe thermal environment tests in the flight process, and high requirements are provided for the thermal design of the satellite sensors. Through the powerful and accurate heat transmission capability of the heat pipe, different satellite-sensitive heat consumptions are transmitted to the stable radiating surface to realize satellite-sensitive temperature control.
The patent document with the publication number of CN108759869A discloses a high-precision star sensor support thermal deformation test system, which comprises a temperature control system and a thermal deformation measurement system, wherein the temperature control system comprises a temperature loading system and a temperature measurement system and is used for accurately simulating an in-orbit temperature field of a star sensor support through feedback control on a loaded temperature load; the thermal deformation measuring system adopts an angle measuring system based on a photoelectric autocollimator and is used for measuring an optical axis vector of a star sensor reference prism under specified measuring conditions, the optical axis of the reference prism is used for representing the pointing angle of the star sensor, the pointing change condition of the star sensor is obtained through data analysis and processing, and the rationality of the thermal stability design of the star sensor support is verified. The inventor considers that the invention mainly describes a star sensor support thermal deformation test system.
The patent document with the publication number of CN103448925A discloses a satellite star sensor high-precision temperature control device, which comprises a star sensor mounting bracket, a star sensor, a thermistor, a heater, an externally-attached heat pipe, an independent heat dissipation surface, and a multi-layer heat insulation assembly, wherein the star sensor, the thermistor and the heater are mounted on a star sensor mounting surface of the star sensor mounting bracket, the externally-attached heat pipe is connected between the star sensor mounting bracket and the independent heat dissipation surface, and the multi-layer heat insulation assembly coats the star sensor mounting bracket, the star sensor, the thermistor and the heater. The inventors believe that the invention is primarily described for controlling the temperature of a star sensitive support by means of a heater and a conventional heat sink.
Patent document CN106292771B discloses a star sensor temperature field measurement and control device, which comprises: the plurality of independent temperature sensors are used for monitoring the temperature of the to-be-detected part of the star sensor; the heating power supply device comprises an electric heating sheet and a temperature control device; the electric heating sheet is used as an electric heater and arranged on the star sensor to-be-detected part; the temperature control equipment is connected with the electric heating piece and provides multiple paths of independent constant-voltage or constant-temperature power supply for the electric heating piece so as to control the temperature of the electric heating piece; and the processing control unit is connected with the heating power supply equipment and the plurality of independent temperature sensors and is used for controlling the test and outputting data. The inventor believes that the invention, which primarily describes the measurement and control of the star sensor temperature field, is a method for validating thermal design.
Patent document No. CN106940196A discloses a method for correcting the mounting thermal deformation of a star sensor, which comprises the following steps: s1, reserving a thermal deformation correction parameter uploading interface in the satellite remote control number-of-injection module; s2, accessing the first star sensor into the system, sorting the accessed multi-track telemetering download data, and extracting effective data of the satellite attitude angle, the measurement attitude angles of the first and second star sensors and the latitude argument of the satellite; s3, comparing the measurement attitude angle of the second star sensor with the satellite attitude angle to obtain the attitude deviation of the second star sensor; s4, expressing the relation between the measured attitude angle deviation of the second star sensor and the latitude argument of the satellite through a Fourier series fitting function, and solving the correlation coefficient of the Fourier series fitting function; and S5, injecting the fitted parameter values into the satellite remote control injection module through the remote control injection package.
Therefore, a technical solution is needed to improve the above technical problems.
Disclosure of Invention
Aiming at the defects in the prior art, the invention aims to provide a satellite-sensitive temperature control system.
The satellite-sensitive temperature control system for the satellite comprises a satellite-sensitive section, a heat insulation section, a heat dissipation section, a satellite-sensitive support, a light shield, a cerium glass secondary surface mirror, a heat pipe, a platinum resistor, a plurality of layers, a heat insulation pad, an installation and fixation accessory, a cabin plate and heat conduction silicone grease;
the heat pipe is fixedly combined with the satellite sensor through the mounting and fixing accessory and the heat insulation pad, heat-conducting silicone grease is coated on the combined surface, the platinum resistors are respectively adhered to the satellite sensor and the satellite sensor support, the satellite sensor support, the light shield and the heat pipe are subjected to multi-layer coating, the cerium glass secondary surface mirror is mounted on the heat pipe through the mounting and fixing accessory, and one unit of the multi-layer heat insulation assembly is composed of a reflecting layer and a spacer layer which are arranged at intervals.
Preferably, the star sensor comprises a star sensor support and a star sensor head, the star sensor head is installed on the star sensor support through a heat insulation pad and an installation fixing accessory, the star sensor head is a heating component and is in good contact with one end of a heat pipe, the star sensor head is subjected to black anodizing treatment in advance to enhance radiation heat exchange, and the whole star sensor head is coated with multiple layers.
Preferably, the heat pipe cabin inner section and the star sensitive head are coated with multiple layers; the heat pipe is also coated with a plurality of layers along the heat dissipation surface section; the outer end of the heat pipe cabin is pasted with a cerium glass secondary surface lens as a radiating surface.
Preferably, the star sensitive segment comprises a star sensitive bracket, a star sensitive head, multiple layers, a heat pipe evaporation end, a mounting accessory and a heat insulation pad; the star sensitive head is arranged on the star sensitive support through the heat insulation pad and the mounting accessory, the star sensitive support is coated with a plurality of layers, and the star sensitive head is firstly in close contact with the heat pipe and then coated with a plurality of layers.
Preferably, the heat dissipation section comprises a heat pipe heat dissipation section, multiple layers, an installation accessory, a heat insulation pad and a cerium glass secondary surface mirror; the heat pipe radiating section is fixed on the cabin plate through the heat insulation pad and the mounting accessory, wherein a cerium glass secondary surface lens is adhered to the surface opposite to the cold space to be used as a star sensitive head radiating surface, and the rest are coated with multiple layers to isolate the heat exchange between the environment and the heat pipe radiating section.
Preferably, the absorptivity and emissivity of the cerium glass secondary surface lens are fixed, the type of the heat pipe and the area of the heat dissipation section are adjusted according to different heat loss of the star sensitive head, and then the area of the cerium glass secondary surface lens is adjusted.
Preferably, the star sensor head is made of an aluminum alloy metal material;
the star sensor bracket and the light shield are made of carbon fiber non-metal materials.
Preferably, the heat pipe is an aluminum ammonia heat pipe for spaceflight;
the cerium glass secondary surface mirror is a cerium glass secondary surface mirror for spaceflight.
Preferably, one unit of the multilayer heat insulation assembly is formed by a reflecting layer and a spacing layer which are spaced;
the platinum resistor is of aerospace grade and is Pt100 in model number.
Preferably, the mounting and fixing accessories are screws for fixing the main mirror back plate on the structure body;
the heat insulation pad is made of glass fiber reinforced plastics.
Compared with the prior art, the invention has the following beneficial effects:
1. the invention can greatly reduce the influence of heat flux outside the space on the temperature of the main star sensor, realizes the temperature stability of the star sensor, and has the obvious advantages of high precision, high adaptability and high reliability;
2. the invention realizes the control of the temperature stability of the main star sensor in a radiation heat exchange mode through the design of heat pipe heat dissipation and multilayer heat insulation, has more uniformity and high reliability compared with other methods;
3. the method has wide application range, is suitable for star sensors with different heat consumption, and can meet the temperature requirement by adjusting the size of the heat dissipation surface of the cerium glass secondary surface mirror;
4. the thermal design method is reasonable, the material source is sufficient, the process is simple and reliable to realize, and the cost is low.
Drawings
Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:
FIG. 1 is an overview of the star sensor of the present invention;
FIG. 2 is a schematic view of a star sensor stent of the present invention;
FIG. 3 is a diagram of a heat dissipation device of the present invention;
FIG. 4 is a diagram of a satellite cooling panel of the present invention;
fig. 5 is an overall assembly view of the present invention.
Wherein:
star-sensitive 1 platinum resistor 6
Star sensitive support 2 multilayer 7
Mounting and fixing accessory 9 for cerium glass secondary surface mirror 4
Heat-conducting silicone grease 11
Detailed Description
The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.
Referring to fig. 1 and 2, the invention provides a satellite-based temperature control system, which comprises a satellite-based temperature control section, a heat insulation section and a heat dissipation section; the device comprises a star sensor 1, a light shield 3, a cerium glass secondary surface mirror 4, a heat pipe 5, a platinum resistor 6, a plurality of layers 7, a heat insulation pad 8, a satellite radiating surface, an installation fixing accessory 9 and the like; the star sensor 1 comprises a star sensor support 2 and a star sensor head, the star sensor head is arranged on the star sensor support 2 through a heat insulation pad 8 and a mounting and fixing accessory 9, the star sensor head is a heating part and is well contacted with one end of a heat pipe 5, and one end of the heat pipe 5 is an evaporation end in a cabin; the radiation heat exchange is enhanced by performing black anodizing treatment on the star sensitive head in advance, and the whole star sensitive head is coated with a plurality of layers 7 to realize the heat insulation effect; the inner section of the heat pipe cabin and the star sensitive head are coated with a plurality of layers 7 to realize the purpose of heat insulation with the external environment; the heat pipe 5 is also coated with a plurality of layers 7 along the heat dissipation surface section; the outer end of the heat pipe cabin is pasted with a cerium glass secondary surface lens as a radiating surface. The outer end of the heat pipe cabin is an outer condensation end. The method for coating the multilayer 7, pasting and fixing the cerium glass secondary surface mirror 4 and pasting the platinum resistor 6 is implemented according to the relevant specifications of space thermal control products.
The star sensor section comprises a star sensor bracket 2, a star sensor head, a plurality of layers 7, a heat pipe evaporation end, a mounting and fixing accessory 9 and a heat insulation pad 8; the star sensitive head is arranged on the star sensitive support 2 through a heat insulation pad 8 and an installation fixing accessory 9, the star sensitive support 2 is coated with a plurality of layers 7, and the star sensitive head is firstly in close contact with the heat pipe 5 and then coated with the plurality of layers 7; the purpose that the heat of the star sensitive head is transmitted to the radiating section only through the heat pipe 5 is achieved, and then the heat is transmitted through a single path.
The heat insulation section comprises a heat pipe insulation section, a plurality of layers 7, a mounting and fixing accessory 9 and a heat insulation pad 8; the heat pipe insulation section is fixed on the deck plate 10 through the heat insulation pad 8 and the mounting fixing accessory 9 and is wrapped by the multiple layers 7 to isolate the heat exchange between the environment and the heat pipe insulation section.
The heat dissipation section comprises a heat pipe heat dissipation section, a plurality of layers 7, an installation fixing accessory 9, a heat insulation pad 8 and a cerium glass secondary surface mirror 4, and the heat pipe heat dissipation section is a condensation end in the embodiment; the heat pipe radiating section is fixed on a cabin plate 10 through a heat insulation pad 8 and a mounting and fixing accessory 9, wherein a cerium glass secondary surface lens is adhered to the surface opposite to the cold space to be used as a star sensitive head radiating surface, and the rest is coated with a plurality of layers 7 to isolate the heat exchange between the environment and the heat pipe radiating section; by the method, the heat of the star sensitive head can be accurately transferred to the heat dissipation section of the heat pipe 5.
The absorptivity and emissivity of the cerium glass secondary surface lens are fixed, so that the heat pipe 5 has a completely independent and stable radiating surface, and the type of the heat pipe 5 and the area of a radiating section can be adjusted according to the heat consumption of different star sensitive heads, so that the area of the cerium glass secondary surface lens is adjusted; has high adaptability to different environmental conditions.
The star sensor head is made of metal materials such as aluminum alloy and the like; the star sensor bracket 2 and the light shield 3 are made of non-metal materials such as carbon fiber and the like; the heat pipe 5 is an aluminum ammonia heat pipe for spaceflight, and the type of the heat pipe 5 can be determined according to different heat consumptions of the star sensor 1 in the embodiment; the cerium glass secondary surface mirror 4 is a cerium glass secondary surface mirror for spaceflight; one unit of the multilayer 7 heat insulation component is formed by a reflecting layer and a spacing layer at intervals; the platinum resistor 6 is of aerospace grade, and the type is Pt 100; the mounting and fixing accessories 9 are screws with specific specifications and are used for fixing the main mirror back plate on the structure body; the heat insulation pad 8 is made of glass fiber reinforced plastics.
The star sensor 1 is fixed on a structural body of the star sensor support 2 through a mounting and fixing accessory 9 and a heat insulation pad 8, the star sensor 1 is subjected to black anodization, then the heat pipe 5 is fixedly combined with the star sensor 1 through the mounting accessory, and a combined surface is coated with heat-conducting silicone grease 11; platinum resistors 6 are respectively adhered to the star sensor 1 and the star sensor support 2; then, the star sensor 1, the star sensor bracket 2, the light shield 3 and the heat pipe 5 are wrapped by a plurality of layers 7, wherein the heat pipe 5 is a star sensor section and a heat insulation section; mounting a cerium glass secondary surface mirror 4 on a heat pipe 5 through a mounting and fixing accessory 9, wherein the heat pipe 5 is an extravehicular condensation section in the embodiment, and heat-conducting silicone grease 11 is coated on the joint surface part; one unit of the multilayer 7 heat insulation assembly is formed by a reflecting layer and a spacing layer at intervals; the heat conductive silicone grease 11 painting method and the black anodizing method are processed according to the spacecraft implementation method.
By adopting the thermal control design method and the thermal control design device, thermal analysis and calculation are carried out on the star sensor 1 component, necessary simplification and assumption are carried out on the model, and basic parameters of the satellite, such as the orbit, the attitude and the like, are set according to the overall technical requirements. The mounting accessories, the heat conducting grease and the multilayer 7 are aerospace standard materials.
The invention meets the requirement of star sensor 1 on temperature precision, simultaneously realizes the minimization of resource use, and has high safety and reliability. The star sensor 1 has wide application range for different heat consumption.
The invention can greatly reduce the influence of heat flux outside the space on the temperature of the main star sensor 1, realize the temperature stability of the star sensor 1, and the design has the obvious advantages of high precision, high adaptability and high reliability; the temperature stability of the main star sensor 1 is controlled in a radiation heat exchange mode through the design of heat dissipation of the heat pipe 5 and heat insulation of the multiple layers 7, and compared with other methods, the method has the advantages of higher uniformity and high reliability.
The method has wide application range, is suitable for the star sensors 1 with different heat consumption sizes, and can meet the temperature requirement by adjusting the size of the heat dissipation surface of the cerium glass secondary surface mirror 4; the thermal design method is reasonable, the material source is sufficient, the process is simple and reliable to realize, and the cost is low.
In the description of the present application, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience in describing the present application and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present application.
The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.
Claims (10)
1. A satellite-used satellite-sensitive temperature control system is characterized by comprising a satellite-sensitive section, a heat insulation section, a heat dissipation section, a satellite-sensitive (1), a satellite-sensitive support (2), a light shield (3), a cerium glass secondary surface mirror (4), a heat pipe (5), a platinum resistor (6), a plurality of layers (7), a heat insulation pad (8), a mounting and fixing accessory (9), a cabin plate (10) and heat-conducting silicone grease (11);
the star sensor is characterized in that the star sensor (1) is fixed on a structural body of a star sensor support (2) through an installation fixing accessory (9) and a heat insulation pad (8), the heat pipe (5) is fixedly combined with the star sensor (1) through the installation fixing accessory (9), a combined surface is coated with heat-conducting silicone grease (11), the star sensor (1) and the star sensor support (2) are respectively pasted with a platinum resistor (6), the star sensor (1), the star sensor support (2), the light shield (3) and the heat pipe (5) are processed in a multi-layer mode (7), the cerium glass secondary surface mirror (4) is installed on the heat pipe (5) through the installation fixing accessory (9), and one unit of the multi-layer (7) heat insulation assembly is composed of a reflecting layer and a spacer layer at intervals.
2. The satellite-used satellite-sensitive temperature control system according to claim 1, wherein the satellite-sensitive head (1) comprises a satellite-sensitive support (2) and a satellite-sensitive head, the satellite-sensitive head is mounted on the satellite-sensitive support (2) through a heat insulation pad (8) and a mounting and fixing accessory (9), the satellite-sensitive head is a heat generating component and is in good contact with one end of the heat pipe (5), the satellite-sensitive head is pre-blackened and anodized to enhance radiative heat exchange, and the whole satellite-sensitive head is coated with a plurality of layers (7).
3. The satellite-based satellite-sensitive temperature control system according to claim 1, wherein the heat pipe cabin inner section and the satellite-sensitive head are covered with multiple layers (7); the heat pipe (5) is also coated with a plurality of layers (7) along the heat dissipation surface section; the outer end of the heat pipe (5) is stuck with a cerium glass secondary surface mirror (4) sheet as a radiating surface.
4. The satellite-used satellite-sensitive temperature control system according to claim 1, wherein the satellite-sensitive segment comprises a satellite-sensitive support (2), a satellite-sensitive head, multiple layers (7), a heat pipe evaporation end, a mounting and fixing accessory (9), and a heat insulation pad (8); the star sensitive head is arranged on the star sensitive support (2) through a heat insulation pad (8) and an installation fixing accessory (9), the star sensitive support (2) is coated with a plurality of layers (7), and the star sensitive head is firstly in close contact with the heat pipe (5) and then coated with the plurality of layers (7).
5. The satellite-used satellite-sensitive temperature control system according to claim 1, wherein the heat dissipation section comprises a heat pipe heat dissipation section, a plurality of layers (7), a mounting and fixing accessory (9), a heat insulation pad (8) and a cerium glass secondary surface mirror (4); the heat pipe radiating section is fixed on a cabin plate (10) through a heat insulation pad (8) and a mounting and fixing accessory (9), wherein a cerium glass secondary surface mirror (4) sheet is pasted on the surface opposite to the cold space to be used as a star sensitive head radiating surface, and the rest are coated with a plurality of layers (7) to isolate the heat exchange between the environment and the heat pipe radiating section.
6. The satellite-used satellite-sensitive temperature control system according to claim 1, wherein the absorptivity and emissivity of the cerium glass secondary surface mirror (4) sheet are fixed, and the type of the heat pipe (5) and the area of the heat dissipation section are adjusted according to different heat consumption of the satellite-sensitive head, so as to adjust the area of the cerium glass secondary surface mirror (4) sheet.
7. The satellite-used star sensor temperature control system according to claim 1, wherein the star sensor (1) head is made of an aluminum alloy metal material;
the star sensor bracket (2) and the light shield (3) are made of carbon fiber non-metal materials.
8. The satellite-based satellite-sensitive temperature control system according to claim 1, wherein the heat pipe (5) is an aluminum ammonia heat pipe for aerospace;
the cerium glass secondary surface mirror (4) is a cerium glass secondary surface mirror for spaceflight.
9. The satellite-based satellite-temperature control system according to claim 1, wherein the multi-layer (7) thermal insulation assembly comprises a reflective layer and a spacer layer which are spaced apart from each other in one unit;
the platinum resistor (6) is aerospace grade and is Pt100 in model.
10. The satellite-used satellite-sensitive temperature control system according to claim 1, wherein the mounting and fixing accessories (9) are screws for fixing the back plate of the primary mirror on the structural body;
the heat insulation pad (8) is made of glass fiber reinforced plastics.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210060731.XA CN114408221B (en) | 2022-01-19 | 2022-01-19 | Satellite-sensitive temperature control system for satellite |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202210060731.XA CN114408221B (en) | 2022-01-19 | 2022-01-19 | Satellite-sensitive temperature control system for satellite |
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| CN114408221A true CN114408221A (en) | 2022-04-29 |
| CN114408221B CN114408221B (en) | 2024-05-07 |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN117699061A (en) * | 2023-12-05 | 2024-03-15 | 中国科学院国家空间科学中心 | Environment-impact-resistant thermal control device for small lunar-based equipment |
Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103448925A (en) * | 2013-08-08 | 2013-12-18 | 上海卫星工程研究所 | High-precision temperature control device for star sensors for satellites |
| CN103662088A (en) * | 2013-11-26 | 2014-03-26 | 中国空间技术研究院 | Thermal control distribution method for star sensors of GEO (geostationary earth orbit) satellite |
| CN203612228U (en) * | 2013-11-26 | 2014-05-28 | 中国空间技术研究院 | Novel star sensor thermal control small cabin |
| US20150069187A1 (en) * | 2013-09-09 | 2015-03-12 | Lockheed Martin Corporation | Hosted instrument radiator system |
| CN106336128A (en) * | 2016-08-19 | 2017-01-18 | 上海裕达实业有限公司 | Flexible OSR second surface mirror thermal control coating and preparation method and application thereof |
| CN110005672A (en) * | 2019-04-04 | 2019-07-12 | 北京卫星制造厂有限公司 | A kind of special-shaped structure piece outer surface stickup OSR blade technolgy design method |
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2022
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103448925A (en) * | 2013-08-08 | 2013-12-18 | 上海卫星工程研究所 | High-precision temperature control device for star sensors for satellites |
| US20150069187A1 (en) * | 2013-09-09 | 2015-03-12 | Lockheed Martin Corporation | Hosted instrument radiator system |
| CN103662088A (en) * | 2013-11-26 | 2014-03-26 | 中国空间技术研究院 | Thermal control distribution method for star sensors of GEO (geostationary earth orbit) satellite |
| CN203612228U (en) * | 2013-11-26 | 2014-05-28 | 中国空间技术研究院 | Novel star sensor thermal control small cabin |
| CN106336128A (en) * | 2016-08-19 | 2017-01-18 | 上海裕达实业有限公司 | Flexible OSR second surface mirror thermal control coating and preparation method and application thereof |
| CN110005672A (en) * | 2019-04-04 | 2019-07-12 | 北京卫星制造厂有限公司 | A kind of special-shaped structure piece outer surface stickup OSR blade technolgy design method |
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| Publication number | Priority date | Publication date | Assignee | Title |
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| CN117699061A (en) * | 2023-12-05 | 2024-03-15 | 中国科学院国家空间科学中心 | Environment-impact-resistant thermal control device for small lunar-based equipment |
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