CN116620569B - Integrated thermal control device and method for solar cell wing and electronic equipment - Google Patents
Integrated thermal control device and method for solar cell wing and electronic equipment Download PDFInfo
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- CN116620569B CN116620569B CN202310496394.3A CN202310496394A CN116620569B CN 116620569 B CN116620569 B CN 116620569B CN 202310496394 A CN202310496394 A CN 202310496394A CN 116620569 B CN116620569 B CN 116620569B
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Classifications
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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
- B64G1/42—Arrangements or adaptations of power supply systems
- B64G1/44—Arrangements or adaptations of power supply systems using radiation, e.g. deployable solar arrays
- B64G1/443—Photovoltaic cell arrays
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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
- B64G1/46—Arrangements or adaptations of devices for control of environment or living conditions
- B64G1/50—Arrangements or adaptations of devices for control of environment or living conditions for temperature control
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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
- B64G1/46—Arrangements or adaptations of devices for control of environment or living conditions
- B64G1/50—Arrangements or adaptations of devices for control of environment or living conditions for temperature control
- B64G1/506—Heat pipes
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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
- B64G1/46—Arrangements or adaptations of devices for control of environment or living conditions
- B64G1/50—Arrangements or adaptations of devices for control of environment or living conditions for temperature control
- B64G1/503—Radiator panels
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- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02E—REDUCTION OF GREENHOUSE GAS [GHG] EMISSIONS, RELATED TO ENERGY GENERATION, TRANSMISSION OR DISTRIBUTION
- Y02E10/00—Energy generation through renewable energy sources
- Y02E10/50—Photovoltaic [PV] energy
Abstract
The invention provides an integrated thermal control device and method for solar cell wings and electronic equipment, which relate to the technical field of satellite thermal control and comprise the following steps: satellite body, electronic equipment, solar cell wings, radiator plate and heat pipe; the satellite body comprises a plurality of cabin boards, and the electronic equipment is arranged on the inner wall of any cabin board; the solar cell wings and the radiator plates are respectively arranged on the outer surfaces of different cabin plates, and the cabin plates on the solar cell wings are adjacent to the cabin plates on the radiator plates; the heat pipe comprises an embedded heat pipe and an externally attached heat pipe, one end of the embedded heat pipe is embedded in the inner wall of the cabin board on the wing side of the solar cell, the other end of the embedded heat pipe is embedded in the radiator board, one end of the externally attached heat pipe is arranged on the outer surface of the cabin board on the side of the electronic equipment, and the other end of the externally attached heat pipe is arranged on the surface of the radiator board. The invention can relieve the problem of overhigh average temperature of the body-mounted solar cell wing, and simultaneously utilizes the thermal control device to construct a heat transmission and heat dissipation channel for electronic equipment on a satellite so as to achieve the aim of integrated design.
Description
Technical Field
The invention relates to the technical field of satellite thermal control, in particular to an integrated thermal control device and method for solar cell wings and electronic equipment.
Background
The solar cell wing is used as an important component of a spacecraft power supply subsystem, is a common power supply component of a spacecraft, is widely applied to satellites at present, and is a main guarantee for successfully completing a spacecraft task due to stable and reliable work. It is counted that the output power is reduced by 0.4% to 0.5% for every 1 deg.c increase in the temperature of the battery. Therefore, to ensure long-term, efficient and stable operation of the solar cell wing, the primary condition is to reduce the operating temperature of the cell as much as possible.
Considering satellite configuration and layout, whole satellite power resource requirement and the like, some existing remote sensing optical satellites are provided with solar cell wings on the-Z surface of a satellite body, namely the solar cell wings are arranged on the body, in order to obtain better illumination conditions, the satellites usually adopt a-Z pair sun gesture during non-task operation, partial on-board equipment is in a standby state, and the solar cell wings are irradiated for a long time and have higher average temperature; when the satellite performs tasks, the posture is adjusted to +Z to the ground, the standby equipment starts up to work, and the average temperature of the body-mounted solar cell wings is reduced due to posture adjustment. In order to reduce the influence of the body-mounted solar cell wings on the temperature of the in-satellite equipment, the back surfaces of the solar cell wings are coated with a plurality of layers of heat insulation assemblies and are installed in a heat insulation mode with other cabin plates. According to the radiating device for the fixed solar wing of the spacecraft, provided by the related technology, the fixed solar wing is cooled through the flexible graphene film, so that the temperature index requirement of the reduced number of solar wing cloth pieces can be met.
Simulation analysis shows that if the back of the body-mounted solar cell wing is coated with the multi-layer heat insulation component, when the satellite attitude is in the-Z pair day for a long time, the highest temperature of the solar cell wing can reach 130 ℃, and the solar cell wing is seriously beyond the normal working temperature range (generally-90 ℃ to +90 ℃), so that the reliability of long-term working of the solar cell wing is not facilitated. A heat dissipating device for fixing solar wings of a spacecraft, such as disclosed in the related art, can reduce the temperature of the solar wings, but is not applicable to bulk solar battery wings with larger area because of being limited by the heat conducting capability of graphene on one hand and the size of graphene on the other hand. In addition, the device directly dissipates the heat of the solar wing to a cold space, so that resource waste is caused. If the heat pipe or the pump driving fluid loop for heat transfer is directly installed on the body-mounted solar cell wing, the heat pipe or the pump driving fluid loop is limited by the requirement of normal working temperature (-60 ℃ to +60 ℃) of the heat pipe or the pump driving fluid loop, so that the heat pipe or the pump driving fluid loop may not work normally, and the reliability of the system is reduced.
Disclosure of Invention
In view of the above, the present invention aims to provide an integrated thermal control device and method for solar cell wings and electronic devices, which can alleviate the problem of overhigh average temperature of the body-mounted solar cell wings, and simultaneously utilize the thermal control device to construct heat transmission and heat dissipation channels for electronic devices on satellites, so as to achieve the integrated design goal.
In a first aspect, an embodiment of the present invention provides an integrated thermal control device for a solar cell wing and an electronic device, including: satellite body, electronic equipment, solar cell wings, radiator plate and heat pipe; wherein,
the satellite body comprises a plurality of cabin boards, and the electronic equipment is arranged on the inner wall of any cabin board;
the solar cell wings and the radiator plates are respectively arranged on the outer surfaces of different cabin plates, and the cabin plates on the solar cell wing sides are adjacent to the cabin plates on the radiator plate sides;
the heat pipe comprises an embedded heat pipe and an externally attached heat pipe, one end of the embedded heat pipe is embedded in the inner wall of the cabin plate on the wing side of the solar cell, the other end of the embedded heat pipe is embedded in the radiator plate, one end of the externally attached heat pipe is installed on the outer surface of the cabin plate on the side of the electronic equipment, and the other end of the externally attached heat pipe is installed on the surface of the radiator plate.
In one embodiment, the solar cell wing comprises a first unfolding mechanism and a first compression release mechanism, wherein the solar cell wing comprises a first sub-wing, a second sub-wing and a third sub-wing, and the first sub-wing is installed on the outer surface of the first cabin board in a heat insulation way; wherein,
the first unfolding mechanism is used for connecting the second sub-wing and the third sub-wing with the first sub-wing respectively;
the first compression release mechanism is arranged on the upper surface of the first cabin board and is used for locking the second sub-wing and the third sub-wing to the upper surface of the first sub-wing so as to enable the second sub-wing and the third sub-wing to be in a folded state; and the first unfolding mechanism is used for releasing the second sub-wing and the third sub-wing to unfold the second sub-wing and the third sub-wing to the set positions in a rotating way.
In one embodiment, the device further comprises a second deployment mechanism and a second compression release mechanism; wherein,
the second unfolding mechanism is used for connecting the radiator plate with a second cabin plate;
the second hold-down release mechanism is for locking the radiator plate to an outer surface of the second deck, and for releasing the radiator plate to rotatably deploy the radiator plate to a set position by the second deployment mechanism.
In one embodiment, the radiator plate is an aluminum honeycomb panel with an aluminum skin, and comprises an inner skin, an aluminum honeycomb and an outer skin, wherein the inner skin and the outer skin are respectively arranged on the upper side and the lower side of the aluminum honeycomb; and the surfaces of the inner skin and the outer skin are provided with thermal control coatings.
In one embodiment, the embedded heat pipe comprises a first embedded rigid section, an embedded flexible section and a second embedded rigid section; wherein,
the embedded flexible section is used for connecting the first embedded rigid section and the second embedded rigid section;
the first embedded rigid section is embedded in the inner wall of the cabin plate at the wing side of the solar cell;
the second embedded rigid section is embedded in the radiator plate, and the second embedded rigid section is connected with the inner skin of the radiator plate through an adhesive film.
In one embodiment, the external heat pipe comprises a first external rigid section, an external flexible section, and a second external rigid section; wherein,
the external-paste flexible section is used for connecting the first external-paste rigid section and the second external-paste rigid section;
the first externally-attached rigid section is arranged on the outer surface of the cabin board at the electronic equipment side, and heat conduction silicone grease is filled between the first externally-attached rigid section and the outer surface of the cabin board at the electronic equipment side;
the second externally-mounted rigid section is mounted on the surface of the radiator plate, and heat conduction silicone grease is filled between the second externally-mounted rigid section and the surface of the radiator plate.
In one embodiment, the embedded heat pipe and the externally attached heat pipe are respectively internally provided with a heat pipe working medium, wherein the heat pipe working medium is used for absorbing heat at a hot end, changing from a liquid state to a gas state, flowing from the hot end to a cold end under the action of internal pressure, and releasing heat at the cold end, changing from the gas state to the liquid state.
In one embodiment, each of the modules is an aluminum skin aluminum honeycomb panel, and the outer surface of the module on the solar cell wing side is provided with a thermal control coating.
In a second aspect, an embodiment of the present invention further provides an integrated thermal control method for a solar cell wing and an electronic device, which is applied to the integrated thermal control apparatus for a solar cell wing and an electronic device provided in any one of the first aspect, where the method includes:
absorbing first heat by the solar cell wings in a state that the satellite is in a non-mission stage and the solar cell wings are opposite to each other, and transmitting the first heat from the solar cell wings to the cabin plate on the solar cell wing side in a radiation form;
transferring the first heat from the cabin plate on the wing side of the solar cell to the radiator plate through the pre-buried heat pipe, and discharging the first heat through the radiator plate; and transmitting the first heat from the cabin plate on the wing side of the solar cell to the electronic equipment through the externally attached heat pipe so as to electrically heat and compensate the electronic equipment.
In one embodiment, the method further comprises:
transferring second heat generated by the electronic equipment to the radiator plate through the external heat pipe when the satellite is in a task stage, and discharging the second heat through the radiator plate;
and transmitting the second heat to the solar cell wing side cabin plate through the embedded heat pipe, and transmitting the second heat from the solar cell wing side cabin plate to the solar cell wing in a radiation mode so as to reduce temperature fluctuation of the solar cell wing.
The embodiment of the invention provides an integrated thermal control device and method for solar cell wings and electronic equipment, comprising the following steps: satellite body, electronic equipment, solar cell wings, radiator plate and heat pipe; the satellite body comprises a plurality of cabin boards, and the electronic equipment is arranged on the inner wall of any cabin board; the solar cell wings and the radiator plates are respectively arranged on the outer surfaces of different cabin plates, and the cabin plates on the solar cell wings are adjacent to the cabin plates on the radiator plates; the heat pipe comprises an embedded heat pipe and an externally attached heat pipe, one end of the embedded heat pipe is embedded in the inner wall of the cabin board on the wing side of the solar cell, the other end of the embedded heat pipe is embedded in the radiator board, one end of the externally attached heat pipe is arranged on the outer surface of the cabin board on the side of the electronic equipment, and the other end of the externally attached heat pipe is arranged on the surface of the radiator board. According to the device, through radiation heat exchange between the solar cell wings and the cabin plate, conduction heat exchange of the heat pipes embedded in the cabin plate and the radiator plate is achieved, heat transfer between the solar cell wings and the radiator plate is greatly reduced, the highest temperature of the body-mounted solar cell wings is greatly reduced, long-term working reliability of the solar cell wings is improved, in addition, heat transfer is conducted between the radiator plate and electronic equipment through the externally attached heat pipes, and therefore heat transfer and heat dissipation channels among the electronic equipment, the body-mounted solar cell wings and the radiator plate are constructed, and therefore the problem that the average temperature of the body-mounted solar cell wings is too high can be relieved.
Additional features and advantages of the invention will be set forth in the description which follows, and in part will be obvious from the description, or may be learned by practice of the invention. The objectives and other advantages of the invention will be realized and attained by the structure particularly pointed out in the written description and claims hereof as well as the appended drawings.
In order to make the above objects, features and advantages of the present invention more comprehensible, preferred embodiments accompanied with figures are described in detail below.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings that are needed in the description of the embodiments or the prior art will be briefly described, and it is obvious that the drawings in the description below are some embodiments of the present invention, and other drawings can be obtained according to the drawings without inventive effort for a person skilled in the art.
Fig. 1 is a schematic structural diagram of an integrated thermal control device for solar cell wings and an electronic device according to an embodiment of the present invention;
fig. 2 is a schematic drawing illustrating a folded state of an integrated thermal control device for solar cell wings and an electronic device according to an embodiment of the present invention;
fig. 3 is a top view illustrating an expanded state of an integrated thermal control device for solar cell wings and an electronic device according to an embodiment of the present invention;
fig. 4 is a bottom view of an integrated thermal control device for solar cell wings and an electronic device according to an embodiment of the present invention;
FIG. 5 is a cross-sectional view of a radiator plate according to an embodiment of the present invention;
fig. 6 is a structural exploded view of an integrated thermal control device for solar cell wings and electronic devices according to an embodiment of the present invention;
fig. 7 is a schematic heat dissipation diagram of an integrated thermal control device for solar cell wings and an electronic device according to an embodiment of the present invention;
FIG. 8 is a schematic structural diagram of a pre-buried heat pipe at a 90-degree viewing angle according to an embodiment of the present invention;
fig. 9 is a schematic structural diagram of a pre-buried heat pipe at a view angle of 180 ° according to an embodiment of the present invention;
fig. 10 is a schematic structural diagram of an externally attached heat pipe at a viewing angle of 0 ° according to an embodiment of the present invention;
fig. 11 is a schematic structural diagram of an externally attached heat pipe at a 90 ° viewing angle according to an embodiment of the present invention;
fig. 12 is a schematic flow chart of an integrated thermal control method for solar cell wings and electronic devices according to an embodiment of the present invention;
fig. 13 is a schematic flow chart of another integrated thermal control method for solar cell wings and electronic devices according to an embodiment of the present invention.
Icon: 1-a satellite body; 11- +X deck; 12- -X deck; 13- +Y deck; 14- -Y deck; 15- +Z deck; 16- -Z deck; 2-solar cell wings; 21-a first sub-wing; 22-a second sub-wing; 23-third sub-wing; 3-an electronic device; 31-a first electronic device; 32-a second electronic device; 4-radiator plates; 41-an inner skin; 42-an outer skin; 43-aluminum honeycomb; 5-a heat pipe; 511-a first pre-buried rigid segment; 512-a second pre-buried rigid section; 513-embedding flexible segments; 514-glue film; 521-a first external rigid section; 522-a second external rigid section; 523-externally attaching a flexible section; 524-thermally conductive silicone grease; 61-a first deployment mechanism; 62-a second deployment mechanism; 71-a first compression release mechanism; 72-a second compression release mechanism.
Detailed Description
For the purpose of making the objects, technical solutions and advantages of the embodiments of the present invention more apparent, the technical solutions of the present invention will be clearly and completely described in conjunction with the embodiments, and it is apparent that the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments, which can be made by those skilled in the art based on the embodiments of the invention without making any inventive effort, are intended to be within the scope of the invention.
At present, a heat dissipation device for fixing solar wings of a spacecraft of a related art cannot be well suitable for body-mounted solar battery wings with larger area, and resource waste is caused by directly dissipating heat of the solar battery wings to a cold space.
For the sake of understanding the present embodiment, first, a detailed description will be given of an integrated thermal control device for solar cell wings and electronic devices according to an embodiment of the present invention, referring to a schematic structural diagram of an integrated thermal control device for solar cell wings and electronic devices shown in fig. 1, the integrated thermal control device for solar cell wings and electronic devices includes: satellite body 1, solar cell wings 2, electronics 3, radiator plate 4 and heat pipe 5.
In one embodiment, the satellite body 1 includes a plurality of panels, and the electronic device 3 is mounted on an inner wall of any of the panels. In one example, the satellite body 1 is in a cuboid structure and is composed of 6 cabin plates, including a +/-X cabin plate, a +/-Y cabin plate and a +/-Z cabin plate, wherein the +/-X cabin plate is arranged opposite to the-X cabin plate, the +/-Y cabin plate is arranged opposite to the-Y cabin plate, and the +/-Z cabin plate is arranged opposite to the-Z cabin plate. The electronic device 3 may be mounted on the inner wall of any of the decks for controlling the satellite to perform tasks during the mission phase.
In one embodiment, the solar cell wings 2 and the radiator plate 4 are mounted on the outer surfaces of different panels, respectively, the panel on the solar cell wing 2 side being adjacent to the panel on the radiator plate 4 side. In one example, solar cell fins 2 are mounted to the outer surface of the-Z module and radiator plate 4 is mounted to the outer surface of the +y module.
In one embodiment, the heat pipe 5 includes an embedded heat pipe and an externally attached heat pipe, one end of the embedded heat pipe is embedded in the inner wall of the cabin board on the solar cell wing 2 side, the other end of the embedded heat pipe is embedded in the radiator board 4, one end of the externally attached heat pipe is installed on the outer surface of the cabin board on the electronic device 3 side, and the other end of the externally attached heat pipe is installed on the surface of the radiator board 4.
In a specific implementation, when the satellite is in a non-mission stage and the attitude is in a-Z cabin board pair day, the heat consumption of the electronic equipment 3 is reduced, the average temperature of the solar cell wings 2 is increased when the solar cell wings are irradiated for a long time, the absorbed heat is transmitted to the-Z cabin board through radiation, then the heat absorbed by the-Z cabin board is transmitted to the radiator board 4 by utilizing the pre-buried heat pipes inside the-Z cabin board and the radiator board 4, and a part of the heat is discharged to a cold space through a thermal control coating on the surface of the radiator board 4; meanwhile, the other part of heat is transferred to the electronic equipment 3 through the external heat pipe, so that the temperature level of the electronic equipment 3 is improved, and the electric heating compensation when the electronic equipment 3 does not work is reduced. When the satellite performs a task, the satellite attitude is changed into +Z cabin board to the ground, the heat consumption of the electronic equipment 3 is increased, the external heat flow received by the solar cell wing 2 is reduced, the average temperature is reduced, the external heat pipe is utilized to transmit the heat generated when the electronic equipment 3 works to the radiator board 4, one part of the heat is discharged to a cold space in a radiation mode through a thermal control coating on the surface of the radiator board 4, and the other part of the heat is transmitted to the-Z cabin board through the pre-buried heat pipe and is transmitted to the solar cell wing 2 through radiation heat exchange between the-Z cabin board and the solar cell wing 2. On one hand, the embodiment of the invention can improve the temperature of the solar cell wing 2 and reduce the temperature fluctuation of the solar cell wing 2; on the other hand, the radiator plate 4 is used for radiating heat to a cold space when the electronic device 3 works, so that the total heat radiating capacity of the whole star is improved.
According to the integrated heat control device for the solar cell wing and the electronic equipment, provided by the embodiment of the invention, through radiation heat exchange between the solar cell wing and the cabin plate and conduction heat exchange between the cabin plate and the radiator plate by the embedded heat pipes, heat transfer between the solar cell wing and the radiator plate is completed, the highest temperature of the body-mounted solar cell wing is greatly reduced, the long-term working reliability of the solar cell wing is improved, and in addition, heat transfer is carried out between the radiator plate and the electronic equipment through the externally-attached heat pipes, so that a heat transfer and heat dissipation channel among the electronic equipment, the body-mounted solar cell wing and the radiator plate is constructed.
In one embodiment, satellite body 1 includes +x deck 11, -X deck 12, +y deck 13, -Y deck 14, +z deck 15, and-Z deck 16, each of which is an aluminum skin aluminum honeycomb panel. Wherein, the outer surface of the cabin board on the solar cell wing 2 side is provided with a thermal control coating, and the outer surface is one side facing the body-mounted solar cell wing 2. Alternatively, the thermal control coating is selected from KS-ZA white paint or OSR (Optical Solar Reflecor) commonly used for satellite thermal control products.
In one embodiment, the solar cell wing 2 includes a first sub-wing 21, a second sub-wing 22, and a third sub-wing 23. In one example, the first sub-wing 21 is mounted on the outer surface of a first panel, which may be the-Z panel 16, and the panel on the solar cell wing 2 side is the first panel. Specifically, the solar cell wing 2 is divided into 3 sub-wings, the first sub-wing 21 positioned in the middle is in a body-mounted type, the second sub-wing 22 arranged on the +X side and the third sub-wing 23 arranged on the-X side of the satellite body in a heat insulation manner are in an extensible design. Wherein the back side of the first sub-wing 21 (facing the-Z deck 16 side) is left untreated and the distance to the-Z deck 16 is as large as possible without affecting the overall envelope of the satellite.
In one embodiment, the integrated thermal control device of the solar cell wing and the electronic device further comprises a first deployment mechanism 61 and a first compression release mechanism 71. Specific:
in one example, the first deployment mechanism 61 is configured to connect the second sub-wing 22 and the third sub-wing 23 to the first sub-wing 21, respectively.
In one example, a first compression release mechanism 71 is provided on the upper surface of the first deck, the first compression release mechanism 71 being used to lock the second and third sub-wings 22 and 23 to the upper surface of the first sub-wing 21 so that the second and third sub-wings 22 and 23 are in a collapsed state. Specifically, the first compression release mechanism 71 is used to fix the second sub-wing 22 and the third sub-wing 23 in the collapsed state. Optionally, the second sub-wing 22 and the third sub-wing 23 are in a folded state before satellite takeoff, such as a folded state schematic diagram of an integrated thermal control device for solar cell wings and electronic devices shown in fig. 2.
In one example, the first compression release mechanism 71 is also used to release the second and third sub-wings 22 and 23 to rotatably deploy the second and third sub-wings 22 and 23 to the set positions by the first deployment mechanism 61. Specifically, after the satellite is in orbit, the first compression release mechanism 71 is unlocked, and the second sub-wing 22 and the third sub-wing 23 are rotatably deployed to the set positions by the first deployment mechanism 61 and locked, such as a top view of the deployed state of the integrated thermal control device of the solar cell wing and the electronic apparatus shown in fig. 3, and a bottom view of the deployed state of the integrated thermal control device of the solar cell wing and the electronic apparatus shown in fig. 4.
In one embodiment, the radiator plate 4 is an aluminum skin aluminum honeycomb plate, and the radiator plate 4 includes an inner skin 41, an aluminum honeycomb 43, and an outer skin 42, referring to a cross-sectional view of the radiator plate shown in fig. 5, wherein the inner skin 41 and the outer skin 42 are disposed on upper and lower sides of the aluminum honeycomb 43, respectively; the surfaces of the inner skin 41 and the outer skin 42 are each provided with a thermal control coating. In specific implementation, the thermal control coating can be selected from KS-ZA white paint or OSR commonly used for satellite thermal control products, and the area of the thermal control coating is adjusted according to the equipment heat consumption and the space environment.
In one embodiment, the integrated thermal control device for solar cell wings and electronic devices provided by the embodiments of the present invention further includes a second deployment mechanism 62 and a second compression release mechanism 72. Specific:
in one example, a second deployment mechanism 62 is used to connect the radiator panel 4 to the second deck. Wherein the second deck may be a +y deck 13. In practice, to reduce the effect on the carrying envelope, the radiator panel 4 is connected to the +y deck 13 by a second deployment mechanism 62 and is secured to the +y deck 13 by a second hold-down release mechanism 72, the radiator panel 4 being in a collapsed condition prior to take-off.
In one example, the second compression release mechanism 72 is used to lock the radiator plate 4 to the outer surface of the second deck and also to release the radiator plate 4 to rotationally deploy the radiator plate 4 to the set position by the second deployment mechanism 62. In practice, after the satellite is in orbit, the second compression release mechanism 72 is unlocked, and the radiator plate 4 is rotatably unfolded to the set position by the second unfolding mechanism 62 and locked.
In practical applications, the first compression release mechanism 71 and the second compression release mechanism 72 may be compression release devices commonly used in satellites.
In one embodiment, the electronic device 3 may include a first electronic device 31 and a second electronic device 32, where the first electronic device 31 and the second electronic device 32 are devices commonly used for satellite platforms, and when the satellite is in a mission phase, the working heat consumption of the satellite is increased, and when the satellite is in a non-mission phase, the heat consumption of the satellite is reduced, and the first electronic device 31 and the second electronic device 32 are both installed on the inner wall of the +y-board 13, such as a structural explosion diagram of an integrated thermal control device of a solar cell wing and an electronic device shown in fig. 6, and fig. 6 only illustrates a relative position relationship diagram of each component, which does not represent an actual structure.
In one embodiment, the heat pipe 5 includes two types: the heat pipe is pre-buried and the heat pipe is externally attached. The two heat pipes are flexible heat pipes, namely, two ends of the heat pipe are rigid sections, the middle part of the heat pipe is a flexible section, the rigid section provides a heat conduction installation contact plane between the heat pipe and the cabin board and is used for fixing the heat pipe, and the flexible section is used for connecting the rigid sections at the two ends of the heat pipe so as to adapt to track change in the unfolding process of the radiator board.
Furthermore, heat pipe working media are arranged in the embedded heat pipe and the externally attached heat pipe, and are used for absorbing heat at the hot end to be changed from liquid to gaseous, flowing from the hot end to the cold end under the action of internal pressure, and releasing heat at the cold end to be changed from gaseous to liquid. In practical application, the working medium in the heat pipe absorbs heat at the hot end, the working medium changes from liquid state to gas state, the gas working medium flows to the cold end under the action of internal pressure and releases heat, the condensed gas working medium changes into liquid state, and flows back to the hot end under the action of the capillary core to form a cycle.
For the sake of easy understanding, the embodiment of the present invention further provides a heat dissipation schematic diagram of an integrated thermal control device for solar cell wings and electronic devices as shown in fig. 7, and for the sake of easy understanding of fig. 7, the embodiment of the present invention further explains the pre-buried heat pipe and the externally attached heat pipe respectively, specifically:
the first embedded heat pipe comprises a first embedded rigid section 511, an embedded flexible section 513 and a second embedded rigid section 512. Referring to fig. 8, a schematic structural diagram of a heat pipe embedded at a 90 ° view angle, and fig. 9, a schematic structural diagram of a heat pipe embedded at a 180 ° view angle, in which an embedded flexible section 513 is used to connect a first embedded rigid section 511 and a second embedded rigid section 512; the first embedded rigid section 511 is embedded in the inner wall of the cabin board at the solar cell wing side, namely the first embedded rigid section 511 is embedded in the-Z cabin board 16; the second pre-buried rigid section 512 is pre-buried inside the radiator plate 4. Further, the heat pipe is connected to the structural board by an adhesive film 514, such as the second rigid section 512 is connected to the inner skin 41 of the radiator board 4 by the adhesive film 514.
The second external heat pipe includes a first external rigid section 521, an external flexible section 523, and a second external rigid section 522. Referring to fig. 10, a schematic structural diagram of an external heat pipe at a 0 ° viewing angle is shown, and fig. 11, a schematic structural diagram of an external heat pipe at a 90 ° viewing angle is shown, wherein an external flexible section 523 is used to connect a first external rigid section 521 and a second external rigid section 522; the first external rigid section 521 is mounted on the outer surface of the electronic-device-side deck, that is, the first external rigid section 521 is mounted +y on the outer surface of the deck 13; a second external rigid section 522 is mounted to the surface of the radiator plate 4. Further, a thermal silicone grease 524 is filled between the heat pipe and the contact surface, such as between the first rigid section 521 and the outer surface of the cabin board of the electronic device 3, and between the second rigid section 522 and the surface of the radiator board 4.
In summary, the integrated thermal control device for solar cell wings and electronic devices provided by the embodiment of the invention has at least the following characteristics:
(a) The back of the body-mounted solar cell wing is not treated, -a thermal control coating is arranged on the outer surface of the Z cabin plate, and heat exchange is carried out between the Z cabin plate and the thermal control coating through radiation;
(b) The method comprises the steps of arranging an expandable radiator plate, wherein heat exchange is carried out between a Z cabin plate and the radiator plate through pre-buried heat pipes, and heat exchange is carried out between electronic equipment and the radiator plate through externally attached heat pipes, so that the Z cabin plate and the radiator plate are integrally designed;
(c) The heat pipe is a flexible heat pipe and comprises a rigid section and a flexible section, wherein the flexible section has 90-180 DEG bending capability.
In addition, the integrated thermal control device for the solar cell wing and the electronic equipment provided by the embodiment of the invention can achieve the following technical effects:
(1) Compared with the prior art, the heat transfer between the body-mounted solar cell wing and the radiator plate is completed through the radiation heat exchange between the body-mounted solar cell wing and the Z cabin plate and the conduction heat exchange of the heat pipe embedded in the radiator plate, so that the highest temperature of the body-mounted solar cell wing is greatly reduced, and the long-term working reliability of the solar cell wing is improved;
(2) The radiator plate and the electronic equipment are subjected to heat transfer through the externally attached heat pipe, so that a heat transmission and heat dissipation channel among the electronic equipment, the body-mounted solar cell wing and the radiator plate is constructed, and an integrated design is realized;
(3) The heat control device is utilized to guide redundant heat of the body-mounted solar cell wing to the electronic equipment end, so that the electric heating compensation power of the equipment is reduced, and the whole star power resource is saved;
(4) The expandable radiator plate is arranged, so that the total heat dissipation capacity of the whole star is improved; in addition, the area of the expandable radiator plate is adjusted according to the actual internal thermal environment and the space thermal environment, so that the adaptability is high;
(5) The expandable radiator plate is arranged, and before taking off, the radiator plate is in a furled state, so that the envelope size of the whole star can be effectively reduced, and the influence on the carrying envelope is reduced;
(6) The heat pipe comprises a rigid section and a flexible section, wherein the flexible section has 0-180 DEG bending capability and can adapt to the track change in the unfolding process of the radiator plate;
(7) The heat pipe is not directly arranged on the body-mounted solar cell wing, the use environment of the heat pipe is optimized, the working reliability of the heat pipe is improved, and the robustness is good.
For the integrated thermal control device for a solar cell wing and an electronic device provided in the foregoing embodiment, the embodiment of the present invention further provides an integrated thermal control method for a solar cell wing and an electronic device, where the method is applied to the integrated thermal control device for a solar cell wing and an electronic device provided in the foregoing embodiment, referring to a schematic flow chart of an integrated thermal control method for a solar cell wing and an electronic device shown in fig. 12, the method mainly includes steps S1202 to S1204:
step S1202, absorbing the first heat by the solar cell wings and transmitting the first heat from the solar cell wings to the cabin board on the solar cell wing side in a radiation mode when the satellite is in a non-mission stage and the solar cell wings are opposite to each other;
step S1204, transferring the first heat from the cabin plate on the solar cell wing side to the radiator plate through the pre-buried heat pipe, and discharging the first heat through the radiator plate; and transferring the first heat from the cabin plate on the wing side of the solar cell to the electronic equipment through the externally attached heat pipe so as to electrically heat and compensate the electronic equipment.
In particular implementations, embodiments of the present invention may utilize the radiative heat transfer between the-Z module 16 and the first sub-wing 21 to indirectly control the temperature of the solar cell wing 2, and in turn, the heat transfer function of the heat pipe 5 to control the temperature of the radiator plate 4 and the electronic device 3. Namely, when the satellite is in a non-mission stage and the attitude is in the-Z cabin plate 16 pair day, the heat consumption of the first electronic equipment 31 and the second electronic equipment 32 is reduced, the average temperature of the body-mounted solar cell wing 2 is increased when the time under irradiation is prolonged, part of the absorbed heat is transmitted to the-Z cabin plate 16 through radiation, then the heat absorbed by the-Z cabin plate 16 is transmitted to the radiator plate 4 by utilizing the pre-buried heat pipes inside the-Z cabin plate 16 and the radiator plate 4, and part of the heat is discharged to a cold space through a thermal control coating on the surface of the radiator plate 4, so that compared with the multilayer heat insulation component coated on the back of the body-mounted solar cell wing 2, the temperature of the body-mounted solar cell wing 2 is greatly reduced; meanwhile, another part of heat is transferred to the first electronic device 31 and the second electronic device 32 through the external heat pipe, so that the temperature level of the first electronic device 31 and the second electronic device 32 is improved, and the electric heating compensation when the first electronic device 31 and the second electronic device 32 do not work is reduced.
The embodiment of the invention also provides another integrated thermal control method for solar cell wings and electronic equipment, referring to a flow chart of the integrated thermal control method for solar cell wings and electronic equipment shown in fig. 13, the method mainly comprises the following steps S1302 to S1304:
step S1302, when the satellite is in a task stage, transferring the second heat generated by the electronic device to the radiator plate through the external heat pipe, and discharging the second heat through the radiator plate;
in step S1304, the second heat is transferred to the solar cell wing side panel through the pre-buried heat pipe, and the second heat is transferred from the solar cell wing side panel to the solar cell wing in a radiation form, so as to reduce the temperature fluctuation of the solar cell wing.
In a specific implementation, when the satellite performs a task, the satellite attitude is converted into +z cabin board 15 to the ground, the heat consumption of the electronic equipment is increased, and the external heat flow received by the solar cell wing 2 is reduced, the average temperature is reduced, the heat generated when the first electronic equipment 31 and the second electronic equipment 32 work is transmitted to the radiator board 4 by using the external heat pipe, one part of the heat is discharged to a cold space in a radiation mode through a thermal control coating on the surface of the radiator board 4, the other part of the heat is transmitted to the-Z cabin board 16 through the embedded heat pipe and is transmitted to the first sub-wing 21 through radiation heat exchange between the-Z cabin board 16 and the first sub-wing 21, so that on one hand, the temperature of the body-mounted solar cell wing 2 can be increased, the temperature fluctuation of the solar cell wing 2 is reduced, and on the other hand, the heat generated when the first electronic equipment 31 and the second electronic equipment 32 work is discharged to the cold space by using the radiator board 4, and the total heat dissipation capacity of the whole satellite is improved.
According to the integrated heat control method for the solar cell wing and the electronic equipment, through radiation heat exchange between the solar cell wing and the cabin plate and conduction heat exchange of the heat pipes embedded in the cabin plate and the radiator plate, heat transfer between the solar cell wing and the radiator plate is completed, the highest temperature of the body-mounted solar cell wing is greatly reduced, the long-term working reliability of the solar cell wing is improved, and in addition, heat transfer is carried out between the radiator plate and the electronic equipment through the externally-attached heat pipes, so that a heat transfer and heat dissipation channel among the electronic equipment, the body-mounted solar cell wing and the radiator plate is constructed.
It will be clearly understood by those skilled in the art that, for convenience and brevity of description, the specific working process of the integrated thermal control method for solar cell wings and electronic devices described above may refer to the corresponding process in the foregoing embodiment, and will not be repeated herein.
In the description of embodiments of the present invention, unless explicitly specified and limited otherwise, the terms "mounted," "connected," and "connected" are to be construed broadly, and may be, for example, fixedly connected, detachably connected, or integrally connected; can be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communication between two elements. The specific meaning of the above terms in the present invention will be understood by those skilled in the art in specific cases.
In the description of the present invention, it should be noted that the directions or positional relationships indicated by the terms "center", "upper", "lower", "left", "right", "vertical", "horizontal", "inner", "outer", etc. are based on the directions or positional relationships shown in the drawings, are merely for convenience of describing the present invention and simplifying the description, and do not indicate or imply that the devices or elements referred to must have a specific orientation, be configured and operated in a specific orientation, and thus should not be construed as limiting the present invention. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
Finally, it should be noted that: the above examples are only specific embodiments of the present invention, and are not intended to limit the scope of the present invention, but it should be understood by those skilled in the art that the present invention is not limited thereto, and that the present invention is described in detail with reference to the foregoing examples: any person skilled in the art may modify or easily conceive of the technical solution described in the foregoing embodiments, or perform equivalent substitution of some of the technical features, while remaining within the technical scope of the present disclosure; such modifications, changes or substitutions do not depart from the spirit and scope of the technical solutions of the embodiments of the present invention, and are intended to be included in the scope of the present invention. Therefore, the protection scope of the present invention shall be subject to the protection scope of the claims.
Claims (8)
1. An integrated thermal control device for solar cell wings and electronic equipment, comprising: satellite body, electronic equipment, solar cell wings, radiator plate and heat pipe; wherein,
the satellite body comprises a plurality of cabin boards, and the electronic equipment is arranged on the inner wall of any cabin board;
the solar cell wings and the radiator plates are respectively arranged on the outer surfaces of different cabin plates, and the cabin plates on the solar cell wing sides are adjacent to the cabin plates on the radiator plate sides;
the heat pipe comprises an embedded heat pipe and an externally attached heat pipe, one end of the embedded heat pipe is embedded in the inner wall of the cabin plate on the wing side of the solar cell, the other end of the embedded heat pipe is embedded in the radiator plate, one end of the externally attached heat pipe is arranged on the outer surface of the cabin plate on the electronic equipment side, and the other end of the externally attached heat pipe is arranged on the surface of the radiator plate;
the solar cell wing comprises a first sub-wing, a second sub-wing and a third sub-wing, and the first sub-wing is installed on the outer surface of the first cabin board in a heat insulation way; wherein,
the first unfolding mechanism is used for connecting the second sub-wing and the third sub-wing with the first sub-wing respectively;
the first compression release mechanism is arranged on the upper surface of the first cabin board and is used for locking the second sub-wing and the third sub-wing to the upper surface of the first sub-wing so as to enable the second sub-wing and the third sub-wing to be in a folded state; the first unfolding mechanism is used for unfolding the second sub-wing and the third sub-wing to a set position in a rotating way;
the device also comprises a second unfolding mechanism and a second compression release mechanism; wherein,
the second unfolding mechanism is used for connecting the radiator plate with a second cabin plate;
the second hold-down release mechanism is for locking the radiator plate to an outer surface of the second deck, and for releasing the radiator plate to rotatably deploy the radiator plate to a set position by the second deployment mechanism.
2. The integrated thermal control device of solar cell wings and electronic equipment according to claim 1, wherein the radiator plate is an aluminum skin aluminum honeycomb plate, the radiator plate comprises an inner skin, an aluminum honeycomb and an outer skin, and the inner skin and the outer skin are respectively arranged on the upper side and the lower side of the aluminum honeycomb; and the surfaces of the inner skin and the outer skin are provided with thermal control coatings.
3. The integrated thermal control device of solar cell wings and electronic equipment of claim 1, wherein the pre-buried heat pipe comprises a first pre-buried rigid segment, a pre-buried flexible segment, and a second pre-buried rigid segment; wherein,
the embedded flexible section is used for connecting the first embedded rigid section and the second embedded rigid section;
the first embedded rigid section is embedded in the inner wall of the cabin plate at the wing side of the solar cell;
the second embedded rigid section is embedded in the radiator plate, and the second embedded rigid section is connected with the inner skin of the radiator plate through an adhesive film.
4. The integrated thermal control device of solar cell wings and electronic equipment of claim 1, wherein the external heat pipe comprises a first external rigid section, an external flexible section, and a second external rigid section; wherein,
the external-paste flexible section is used for connecting the first external-paste rigid section and the second external-paste rigid section;
the first externally-attached rigid section is arranged on the outer surface of the cabin board at the electronic equipment side, and heat conduction silicone grease is filled between the first externally-attached rigid section and the outer surface of the cabin board at the electronic equipment side;
the second externally-mounted rigid section is mounted on the surface of the radiator plate, and heat conduction silicone grease is filled between the second externally-mounted rigid section and the surface of the radiator plate.
5. The integrated thermal control device of solar cell wings and electronic equipment according to claim 3 or 4, wherein heat pipe working media are placed in the embedded heat pipe and the externally attached heat pipe, and the heat pipe working media are used for absorbing heat at a hot end to be changed into a gas state, flowing from the hot end to a cold end under the action of internal pressure, and releasing heat at the cold end to be changed into a liquid state from a gas state.
6. The integrated thermal control device of solar cell wings and electronic equipment of claim 1, wherein each of the panels is an aluminum skin aluminum honeycomb panel, and the outer surfaces of the panels on the solar cell wing sides are provided with a thermal control coating.
7. A method of integrated thermal control of solar cell wings and electronic devices, characterized in that it is applied to an integrated thermal control apparatus of solar cell wings and electronic devices according to any one of claims 1 to 6, said method comprising:
absorbing first heat by the solar cell wings in a state that the satellite is in a non-mission stage and the solar cell wings are opposite to each other, and transmitting the first heat from the solar cell wings to the cabin plate on the solar cell wing side in a radiation form;
transferring the first heat from the cabin plate on the wing side of the solar cell to the radiator plate through the pre-buried heat pipe, and discharging the first heat through the radiator plate; and transmitting the first heat from the cabin plate on the wing side of the solar cell to the electronic equipment through the externally attached heat pipe so as to electrically heat and compensate the electronic equipment.
8. The integrated thermal control method of a solar cell wing and electronic device of claim 7, further comprising:
transferring second heat generated by the electronic equipment to the radiator plate through the external heat pipe when the satellite is in a task stage, and discharging the second heat through the radiator plate;
and transmitting the second heat to the solar cell wing side cabin plate through the embedded heat pipe, and transmitting the second heat from the solar cell wing side cabin plate to the solar cell wing in a radiation mode so as to reduce temperature fluctuation of the solar cell wing.
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