CN112636446B - Power supply system of quick response small satellite - Google Patents
Power supply system of quick response small satellite Download PDFInfo
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- CN112636446B CN112636446B CN202011593426.4A CN202011593426A CN112636446B CN 112636446 B CN112636446 B CN 112636446B CN 202011593426 A CN202011593426 A CN 202011593426A CN 112636446 B CN112636446 B CN 112636446B
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- 230000004044 response Effects 0.000 title claims abstract description 20
- 238000003860 storage Methods 0.000 claims abstract description 25
- 238000009413 insulation Methods 0.000 claims description 14
- 238000009826 distribution Methods 0.000 claims description 11
- 108091092878 Microsatellite Proteins 0.000 claims description 9
- 229920000049 Carbon (fiber) Polymers 0.000 claims description 5
- 239000004917 carbon fiber Substances 0.000 claims description 5
- VNWKTOKETHGBQD-UHFFFAOYSA-N methane Chemical compound C VNWKTOKETHGBQD-UHFFFAOYSA-N 0.000 claims description 5
- XAGFODPZIPBFFR-UHFFFAOYSA-N aluminium Chemical group [Al] XAGFODPZIPBFFR-UHFFFAOYSA-N 0.000 claims description 3
- 229910052782 aluminium Inorganic materials 0.000 claims description 3
- 238000003825 pressing Methods 0.000 claims description 3
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- 238000013461 design Methods 0.000 abstract description 16
- 238000009434 installation Methods 0.000 abstract description 7
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- 230000006872 improvement Effects 0.000 description 9
- 238000012545 processing Methods 0.000 description 5
- 230000009471 action Effects 0.000 description 4
- 238000004519 manufacturing process Methods 0.000 description 4
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- 230000009286 beneficial effect Effects 0.000 description 2
- 238000010586 diagram Methods 0.000 description 2
- 238000005259 measurement Methods 0.000 description 2
- 238000012986 modification Methods 0.000 description 2
- 230000004048 modification Effects 0.000 description 2
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- 238000006243 chemical reaction Methods 0.000 description 1
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- 238000007599 discharging Methods 0.000 description 1
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- H—ELECTRICITY
- H02—GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
- H02J—CIRCUIT ARRANGEMENTS OR SYSTEMS FOR SUPPLYING OR DISTRIBUTING ELECTRIC POWER; SYSTEMS FOR STORING ELECTRIC ENERGY
- H02J7/00—Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries
- H02J7/34—Parallel operation in networks using both storage and other dc sources, e.g. providing buffering
- H02J7/35—Parallel operation in networks using both storage and other dc sources, e.g. providing buffering with light sensitive cells
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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
- B64G1/425—Power storage
-
- 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
- B64G1/428—Power distribution and management
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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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- H—ELECTRICITY
- H02—GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
- H02J—CIRCUIT ARRANGEMENTS OR SYSTEMS FOR SUPPLYING OR DISTRIBUTING ELECTRIC POWER; SYSTEMS FOR STORING ELECTRIC ENERGY
- H02J7/00—Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries
- H02J7/0042—Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries characterised by the mechanical construction
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- H—ELECTRICITY
- H02—GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
- H02J—CIRCUIT ARRANGEMENTS OR SYSTEMS FOR SUPPLYING OR DISTRIBUTING ELECTRIC POWER; SYSTEMS FOR STORING ELECTRIC ENERGY
- H02J7/00—Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries
- H02J7/0042—Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries characterised by the mechanical construction
- H02J7/0045—Circuit arrangements for charging or depolarising batteries or for supplying loads from batteries characterised by the mechanical construction concerning the insertion or the connection of the batteries
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- H—ELECTRICITY
- H02—GENERATION; CONVERSION OR DISTRIBUTION OF ELECTRIC POWER
- H02S—GENERATION OF ELECTRIC POWER BY CONVERSION OF INFRARED RADIATION, VISIBLE LIGHT OR ULTRAVIOLET LIGHT, e.g. USING PHOTOVOLTAIC [PV] MODULES
- H02S20/00—Supporting structures for PV modules
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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
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- Engineering & Computer Science (AREA)
- Power Engineering (AREA)
- Remote Sensing (AREA)
- Aviation & Aerospace Engineering (AREA)
- Life Sciences & Earth Sciences (AREA)
- Sustainable Development (AREA)
- Photovoltaic Devices (AREA)
- Charge And Discharge Circuits For Batteries Or The Like (AREA)
Abstract
The invention provides a power supply system of a quick response small satellite, which comprises: the solar cell array is used for converting solar energy into electric energy and then supplying power to the small satellite or charging the storage battery pack; the storage battery pack is connected with the solar cell array in parallel and is used for storing the electric energy converted by the solar cell array or outputting the electric energy to the small satellite; the power supply controller is used for energy management of the small satellites; the storage battery pack is electrically connected with the power supply controller, and the solar battery array is arranged on the outer side wall of the cabin board of the minisatellite through the fixed connecting device and has a gap with the outer side wall of the cabin board; the storage battery pack is arranged in a cabin plate of the small satellite. By the arrangement, on one hand, the solar cell array does not need to be unfolded after the small satellite enters the orbit, and the function that the small satellite enters the orbit to work is realized; on the other hand, the invention realizes the decoupling between the solar cell array structure installation and the whole satellite thermal design of the small satellite, thereby independently carrying out the thermal design on the small satellite body.
Description
Technical Field
The invention belongs to the technical field of power supplies, and particularly relates to a power supply system of a quick response small satellite.
Background
The conventional small satellite generally adopts an expanded solar cell array or an integrally-installed solar cell array as a power generation device in a power supply system of the small satellite. For the unfolding solar cell array mode, unlocking and unfolding actions are needed after the small satellite enters the orbit, a certain time is needed in the unfolding process, so that the small satellite cannot immediately execute tasks after entering the orbit, and the battery array structure is easy to vibrate in the unfolding process, so that the posture of the small satellite body is shaken, and the unfolding work is not facilitated after the small satellite enters the orbit;
for the integrated solar cell array mode, a power generation device on the solar cell array is usually used, and a solar cell sheet is directly attached to the surface of a cabin plate of a small satellite, and the mode does not need to perform related actions of on-orbit expansion, but the mode can ensure that the installation of the solar cell array and the implementation of a whole satellite thermal control means of the small satellite have certain coupling, namely, the thermal design of the whole satellite structure of the small satellite and the design of the solar cell array have coupling, so that the thermal design and the structure development process of the solar cell array are influenced mutually.
Therefore, a power supply system for a fast response small satellite is needed, which can not only realize the operation of the small satellite after the small satellite enters the orbit, but also decouple the installation of a solar cell array and the thermal design of the whole small satellite.
Disclosure of Invention
In order to solve the problems, the invention provides a power supply system for a quick response small satellite, which can not only realize the operation of the small satellite after the small satellite enters the orbit, but also decouple the installation of a solar cell array and the thermal design of the whole small satellite.
In order to achieve the above object, the present invention provides a power supply system for a rapid response small satellite, comprising:
the solar cell array is used for converting solar energy into electric energy and then supplying power to the small satellite or charging the storage battery pack;
the storage battery pack is connected with the solar cell array in parallel and is used for storing the electric energy converted by the solar cell array or outputting the electric energy to a small satellite;
the power supply controller is used for energy management of the small satellites;
the storage battery pack is electrically connected with the power supply controller, and the solar battery array is arranged on the outer side wall of the cabin board of the moonlet through a fixed connecting device and has a gap with the outer side wall of the cabin board; the storage battery pack is arranged in a cabin plate of the small satellite.
As a further improvement of the above scheme, the solar cell array comprises a main solar cell array and an auxiliary solar cell array which are separated from each other, and the main solar cell array and the auxiliary solar cell array are respectively installed on the outer sides of two opposite or adjacent side walls of the small satellite cabin plate through fixed connection devices and are respectively arranged in parallel with the corresponding side walls.
As a further improvement of the above scheme, the deck is a structural main body of a power supply system of a small satellite, and provides support for a solar cell array and a storage battery pack, and the deck is of a cubic structure.
As a further improvement of the above scheme, the fixed connection device includes a heat insulation pad and a connection member disposed on the heat insulation pad, the heat insulation pad is disposed between the deck and the solar cell array, so that a heat insulation gap is formed between the solar cell array and the deck, and the connection member is used for pressing the solar cell array and is fixedly connected with the deck.
As a further improvement of the above scheme, the solar cell array is a thermally decoupled solar cell array, and includes a base structure and a triple-junction gaas cell mounted on the base structure, where the base structure is fixedly mounted on a deck of a small satellite.
As a further improvement of the scheme, the base structure comprises two carbon fiber panels, and a honeycomb-shaped inner core is clamped between the two carbon fiber panels.
As a further improvement of the above scheme, the material of the honeycomb-shaped inner core is aluminum.
As a further improvement of the above scheme, the power controller includes a power management unit, a power control unit and two primary power distribution units, the two primary power distribution units are respectively connected with the solar cell main array and the solar cell auxiliary array, and the two primary power distribution units are respectively electrically connected with the power management unit and the power control unit.
Due to the adoption of the technical scheme, the invention has the beneficial effects that:
the invention provides a power supply system of a quick response minisatellite, which comprises a solar cell array, a storage battery pack and a power supply controller, wherein the storage battery pack is electrically connected with the power supply controller, and the solar cell array is arranged on the outer side wall of a cabin board of the minisatellite through a fixed connecting device and has a gap with the outer side wall of the cabin board; the storage battery pack is arranged in a cabin plate of the small satellite; on one hand, the solar cell array is fixedly arranged on the outer side wall of the cabin plate of the moonlet through the fixed connecting device, and the solar cell array does not need to be unfolded after the moonlet enters the orbit, so that the time consumed by unfolding the cell in the traditional implementation mode is avoided, the shaking of the posture of the moonlet caused by the unfolding action of the cell in the traditional implementation mode is also avoided, and the function of working when the moonlet enters the orbit is realized; on the other hand, the solar cell array is arranged on the outer side wall of the cabin board of the small satellite through the fixed connecting device, and a gap is formed between the solar cell array and the outer side wall of the cabin board, and compared with the traditional implementation mode (the solar cell array is directly attached to the surface of the small satellite body), the connection mode of the invention realizes decoupling between the installation of the solar cell array structure and the whole satellite thermal design of the small satellite, so that the thermal design can be independently carried out on the small satellite body, for example, the outer wall of the small satellite body is thermally coated. That is, when the thermal design is performed on the small satellite, the body of the small satellite can be directly and independently considered without considering an external solar cell array; on the other hand, the invention realizes the thermal decoupling between the solar cell array and the whole satellite thermal design of the small satellite, and also means the decoupling between the processing and manufacturing of the solar cell array and the processing and manufacturing of the small satellite body.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the drawings without creative efforts;
FIG. 1 is a schematic view of the three-dimensional structure of the power system of the quick response moonlet of the present invention mounted on the moonlet;
FIG. 2 is a schematic view of the power system of the quick response moonlet of the present invention mounted in a three-dimensional structure in the direction of the other side of the moonlet;
FIG. 3 is a schematic diagram of a power supply system for a fast response microsatellite according to the present invention;
FIG. 4 is a schematic diagram of an installation structure of a thermally decoupled solar array of a power system of a quick response microsatellite according to the present invention;
FIG. 5 is a schematic view of a portion of the area A in FIG. 4;
FIG. 6 is a partially enlarged structural view of the area B in FIG. 4;
the reference numbers are as follows:
1. a solar cell array; 11. A solar cell main array; 12. auxiliary array of solar cell; 2. a battery pack; 3. a power supply controller; 4. a fixed connection device; 41. a heat insulating pad; 42. a connecting member; 5. a deck board; s, clearance.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the drawings in the embodiments of the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
It should be noted that all the directional indicators such as the first, second, upper, lower, left, right, front and rear … … in the embodiment of the present invention are only used to explain the relative position relationship between the components, the movement situation, etc. in a specific posture as shown in the drawings, and if the specific posture is changed, the directional indicator is changed accordingly.
In addition, the technical solutions in the embodiments of the present invention may be combined with each other, but it must be based on the realization of those skilled in the art, and when the technical solutions are contradictory or cannot be realized, such a combination of technical solutions should not be considered to exist, and is not within the protection scope of the present invention.
The invention will be further described with reference to the following figures:
referring to fig. 1-6, the present invention provides a power supply system for a fast response minisatellite, comprising: the thermal decoupling solar cell array 11 and the storage battery pack 22 are respectively connected with the power controller 33, the thermal decoupling solar cell array 11 and the storage battery pack 22 are mutually independent and arranged in parallel with each other, fixedly installed on the cabin board 5 of the small satellite through a fixed connecting device, and arranged with a gap S between the cabin board 5 and the thermal decoupling solar cell array 11 and the cabin board 5 of the small satellite; the thermal decoupling solar cell array 11 converts solar energy into electric energy through solar photoelectric conversion, and generates current to supply power to the platform or charge the storage battery pack 2. Specifically, the thermally decoupled solar cell array 11 is connected to the power controller 33, and supplies power to the storage battery pack 22 or supplies power to a load through the charging shunting unit; the thermal decoupling type solar cell array 11 is used for converting solar energy into electric energy and supplying power to a small satellite or charging a storage battery pack 22; the storage battery pack 22 is arranged in the cabin board 5 of the small satellite and used for storing electric energy and supplying power to the small satellite when the small satellite enters a shadow period or the power supplied by the solar cell array 1 is insufficient; the power controller 32 is arranged inside the body of the small satellite, and is used for energy management of the small satellite, charge and discharge management of the storage battery pack 22, primary power distribution and remote control and remote measurement functions; the solar cell array 1 is fixedly arranged on the outer side wall of the cabin plate 5 of the moonlet through the fixed connecting device 4, and due to the arrangement, the solar cell array 1 does not need to be unfolded after the moonlet enters the orbit, so that the time consumed by unfolding the cell in the traditional implementation mode is avoided, the shaking of the posture of the moonlet caused by the unfolding action of the cell in the traditional implementation mode is also avoided, and the function of working when the moonlet enters the orbit is realized.
As a preferred embodiment, the thermal decoupling solar cell array 1 comprises solar cells separated from each otherThe solar cell auxiliary array 12 is arranged on the outer sides of two opposite side walls of the small satellite cabin plate 5 through the fixed connecting devices 4, and the solar cell main array 11 and the solar cell auxiliary array 12 are arranged in parallel with the corresponding side walls respectively. In the present embodiment, the area of the solar cell main array 11 is 1.54m2The output power is 330W; the area of the solar cell auxiliary array 12 is 0.82 m2The output power was 175W. At ordinary times, the solar cell main array 11 is in sun-facing mode, and the solar cell main array 11 is in sun-facing mode when the solar cell auxiliary array 12 is in fault. The solar cell auxiliary array 12 can be illuminated to generate electric energy at any time except in the normal attitude maneuver process, the solar cell auxiliary array 12 only captures the sun, and the attitude enters an uncontrolled state.
As a preferred embodiment, the deck 5 is a structural main body of a power supply system of a small satellite, and provides support for the solar cell array 1 and the storage battery pack 2; in the present embodiment, the deck boards 5 of the small satellites are in a cubic structure, and the deck boards 5 mounted on the solar cell main array 11 and the solar cell auxiliary array 12 are arranged in parallel, and are respectively mounted on the deck boards 5 on two sides of the satellite +/-Y.
As a preferred embodiment, in order to isolate the heat conduction of the solar cell array 1 to the satellite body, referring to fig. 4-6, the fixed connection device includes a heat insulation pad 41 and a connection member 42 disposed on the heat insulation pad 41, the heat insulation pad 41 is disposed between the deck 5 and the solar cell array 1, so that a heat insulation gap S is formed between the solar cell array 1 and the deck 5, and the connection member 42 is used for pressing the solar cell array 1 and is fixedly connected with the deck 5, so that the heat conduction to the satellite can be prevented, and the work of the satellite load can be prevented from being affected;
in a traditional implementation mode, a thermal coupling type solar battery array adopts a mounting type mode, and the structural installation of the solar battery array 1 and the whole satellite thermal design of a small satellite are coupled with each other in a correlation mode, so that the thermal design of the solar battery array 1 is considered while the thermal design is carried out on a small satellite body;
in the embodiment, the thermal decoupling solar cell array 1 is arranged on the outer side wall of the cabin plate 5 of the moonlet through the fixed connecting device 4, and a gap S is formed between the thermal decoupling solar cell array 1 and the outer side wall of the cabin plate 5. Taking the solar cell main array 11 as an example, 5 connection points are arranged on the cabin board 5 of the small satellite, and the solar cell main array 11 is respectively installed on the cabin board 5 through a fixed connection device at the 5 connection points. The heat insulation pad 41 of the fixed connecting device is arranged between the cabin plate 5 and the solar cell array 1, so that a heat insulation gap S is formed between the solar cell array 1 and the cabin plate 5, and heat of the solar cell main array 11 is prevented from being transferred to the moonlet body; so that the body of the microsatellite can be thermally designed separately, for example, the outer wall of the body of the microsatellite is thermally coated. That is, when the thermal design is performed on the small satellite, the body of the small satellite can be directly and independently considered without considering the external solar cell array 1; in addition, because the invention realizes the thermal decoupling between the solar cell array 1 and the whole satellite thermal design of the small satellite, which means the decoupling between the processing and manufacturing of the solar cell array 1 and the processing and manufacturing of the small satellite body, compared with the traditional integrated solar cell array 1 mode, the solar cell array 1 and the small satellite body can be independently and parallelly processed in the aspects of processing, development and final assembly test, and the test is carried out after the strict sequential assembly is not needed, thereby being beneficial to accelerating the development cycle of the small satellite.
As a preferred embodiment, the solar cell array 1 is a thermally decoupled solar cell array 1, which includes a base structure and a triple-junction gaas cell mounted on the base structure, wherein the base structure is fixedly mounted on a deck 5 of a small satellite, and the temperature range of the base structure is-100 to 145 ℃. In this embodiment, the base structure includes two carbon fiber panels, and a honeycomb core is sandwiched between the two carbon fiber panels, and the honeycomb core is made of aluminum.
As a preferred embodiment, the power controller 3 includes a power management unit, a power control unit and two primary power distribution units, the two primary power distribution units are respectively connected to the solar cell main array 11 and the solar cell auxiliary array 12, and the two primary power distribution units are respectively electrically connected to the power management unit and the power control unit. The charging and discharging management, primary power distribution and remote control and remote measurement functions are provided for the storage battery pack 2.
The foregoing is a detailed description of the invention, and specific examples are used herein to explain the principles and implementations of the invention, the above description being merely intended to facilitate an understanding of the principles and core concepts of the invention. It should be noted that, for those skilled in the art, it is possible to make various improvements and modifications to the present invention without departing from the principle of the present invention, and those improvements and modifications also fall within the scope of the claims of the present invention.
Claims (9)
1. A power system for a fast response microsatellite comprising:
the solar cell array is used for converting solar energy into electric energy and then supplying power to the small satellite or charging the storage battery pack;
the storage battery pack is connected with the solar cell array in parallel and is used for storing the electric energy converted by the solar cell array or outputting the electric energy to a small satellite;
the power supply controller is used for energy management of the small satellites;
the storage battery pack is electrically connected with the power supply controller, and the solar battery array is arranged on the outer side wall of the cabin board of the moonlet through a fixed connecting device and has a gap with the outer side wall of the cabin board; the fixed connection device comprises a heat insulation pad and a connecting piece arranged on the heat insulation pad, the heat insulation pad is arranged between the cabin board and the solar cell array, a heat insulation gap is formed between the solar cell array and the cabin board, and the connecting piece is used for pressing the solar cell array and is fixedly connected with the cabin board.
2. The power supply system of a rapid response microsatellite according to claim 1 wherein the solar cell array comprises a main solar cell array and an auxiliary solar cell array which are separated from each other, the main solar cell array and the auxiliary solar cell array are respectively installed on the outer sides of two opposite or adjacent side walls of the microsatellite deck through fixed connection devices and are respectively arranged in parallel with the corresponding side walls.
3. The power system of a quick response small satellite according to claim 1 or 2, wherein the solar cell array is a thermally decoupled solar cell array comprising a base structure and a cell sheet mounted on the base structure, and the base structure is fixedly mounted on a deck of the small satellite.
4. The power system of a rapid response microsatellite according to claim 3 wherein said cell is a triple junction GaAs cell.
5. The power system of a rapid response microsatellite according to claim 3 wherein the base structure comprises two carbon fiber face plates with a honeycomb core sandwiched therebetween.
6. The power system of claim 5, wherein the cellular core is aluminum.
7. The power system of claim 2, wherein the power controller comprises a power management unit, a power control unit and two primary power distribution units, the two primary power distribution units are respectively connected with the solar cell main array and the solar cell auxiliary array, and the two primary power distribution units are respectively electrically connected with the power management unit and the power control unit.
8. The power system of a quick response small satellite according to claim 1 or 2, wherein the deck is a structural body of the power system of the small satellite, and provides support for a solar cell array and a storage battery pack, and the deck is a cubic structure.
9. A rapid response moonlet power supply system according to claim 1 or 2, wherein the battery pack is provided in a deck of the moonlet.
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Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101383568A (en) * | 2008-10-24 | 2009-03-11 | 哈尔滨工业大学 | Synchronous expansion mechanism for foldable inflating expansion solar cell paddles |
| CN102897330A (en) * | 2012-10-22 | 2013-01-30 | 湖南航天机电设备与特种材料研究所 | Method for reliably connecting flexible thin-film solar cell and airship envelop |
| CN104136889A (en) * | 2012-01-04 | 2014-11-05 | 罗伯特·博世有限公司 | Sensor device for non-contact detection of a rotational property of a rotatable object |
| CN106410936A (en) * | 2016-08-31 | 2017-02-15 | 航天东方红卫星有限公司 | High-power high-efficiency satellite power supply system based on high voltage and low voltage double buses |
| CN108155869A (en) * | 2016-12-05 | 2018-06-12 | 波音公司 | The heat management system of the temperature of reflecting surface of the control with solar energy concentrator array |
| CN109335023A (en) * | 2018-08-31 | 2019-02-15 | 南京理工大学 | A kind of no cable high density cube star and its assembly method |
| CN111181238A (en) * | 2020-01-08 | 2020-05-19 | 航天行云科技有限公司 | DET power supply system of satellite |
Family Cites Families (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| US6127621A (en) * | 1999-04-02 | 2000-10-03 | The Aerospace Corporation | Power sphere |
| CN102648306A (en) * | 2009-10-22 | 2012-08-22 | 学校法人中央大学 | Large-scale ocean mobile solar power generation system |
| CN103023102A (en) * | 2012-11-26 | 2013-04-03 | 中国人民解放军国防科学技术大学 | Integrated power system for nano satellite |
| CN105375603B (en) * | 2015-08-05 | 2018-01-30 | 上海卫星工程研究所 | Satellite thermal vacuum test ground power supply system |
| CN106100096B (en) * | 2016-06-23 | 2018-08-21 | 航天东方红卫星有限公司 | A kind of micro-nano satellite low-voltage high-efficiency power-supply system |
| CN107994669A (en) * | 2017-11-30 | 2018-05-04 | 中国电子科技集团公司第四十八研究所 | A kind of missile-borne microsatellite energy control system |
| CN108321786A (en) * | 2018-03-16 | 2018-07-24 | 西北工业大学 | A kind of cube star integration power-supply system |
-
2020
- 2020-12-29 CN CN202011593426.4A patent/CN112636446B/en active Active
Patent Citations (7)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101383568A (en) * | 2008-10-24 | 2009-03-11 | 哈尔滨工业大学 | Synchronous expansion mechanism for foldable inflating expansion solar cell paddles |
| CN104136889A (en) * | 2012-01-04 | 2014-11-05 | 罗伯特·博世有限公司 | Sensor device for non-contact detection of a rotational property of a rotatable object |
| CN102897330A (en) * | 2012-10-22 | 2013-01-30 | 湖南航天机电设备与特种材料研究所 | Method for reliably connecting flexible thin-film solar cell and airship envelop |
| CN106410936A (en) * | 2016-08-31 | 2017-02-15 | 航天东方红卫星有限公司 | High-power high-efficiency satellite power supply system based on high voltage and low voltage double buses |
| CN108155869A (en) * | 2016-12-05 | 2018-06-12 | 波音公司 | The heat management system of the temperature of reflecting surface of the control with solar energy concentrator array |
| CN109335023A (en) * | 2018-08-31 | 2019-02-15 | 南京理工大学 | A kind of no cable high density cube star and its assembly method |
| CN111181238A (en) * | 2020-01-08 | 2020-05-19 | 航天行云科技有限公司 | DET power supply system of satellite |
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