CN109703784B - Microsatellite with integrated electronic integrated case as main body - Google Patents
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- CN109703784B CN109703784B CN201910023711.3A CN201910023711A CN109703784B CN 109703784 B CN109703784 B CN 109703784B CN 201910023711 A CN201910023711 A CN 201910023711A CN 109703784 B CN109703784 B CN 109703784B
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Abstract
The invention provides a microsatellite taking an integrated electronic comprehensive case as a main body, which comprises a central cabin, a propelling device, a solar cell array and a truss; the origin of the satellite coordinate is the center of the carrying butt joint surface, + X points to the carrying direction, the reverse side of the fixed solar cell array patch direction is the + Z direction, + Y and the two directions form a right-handed system; mounting the propulsion device at the top of the center nacelle; the solar cell array is connected with the central cabin body through a truss; the central cabin is a plug board type integrated electronic case and comprises functional board cards and a bottom board, wherein the bottom board is parallel to the + Z surface and is arranged at the bottom of the central cabin, and each functional board card is parallel to the bottom board and is inserted in the case.
Description
Technical Field
The invention relates to the technical field of satellite configuration, in particular to a microsatellite structure with an integrated electronic cabin as a satellite main body.
Background
The microsatellite is taken as an important direction of the current satellite development and receives more and more attention, and because the microsatellite has small volume, the microsatellite needs to carry satellite load and satellite equipment has the characteristics of small volume and light weight, and the microsatellite meets the technical development direction.
The main body of a general satellite is a central cabin or a central cabin body, the central cabin body uses an aluminum honeycomb panel as a main structure, each functional module is a single machine in various forms, the single machine is installed on a structural plate through screws, and the single machines are connected through external connectors.
The common integrated comprehensive electronic cabin is used as a piece of equipment of the satellite and is arranged on the structure of the satellite, and the structure of the comprehensive electronic cabin is mainly used for the strength of single-machine equipment and does not bear the structural function of the whole satellite of the satellite. The satellite integrated electronic case generally means that a plurality of functional devices are concentrated in one case, so that higher space utilization rate is realized, and the number of devices is reduced.
In the prior art, a structural plate is mostly used as a core bearing part, and the common plate frame type and plate cylinder type configurations are adopted, wherein the structural plate is a honeycomb plate, and a central cylinder is a carbon fiber cylinder. The disadvantages of the prior art are mainly shown in that:
(1) compared with a microsatellite, the structure has heavy weight and the bearing capacity of the satellite is small
If the existing micro satellite structure adopts the configuration and the structural mode of the traditional large satellite, when the weight of the satellite is less than 100kg, the quantity of equipment on the satellite occupies most space and most structural weight of the satellite, the space for installing loads of the satellite is extremely limited, the satellite is not beneficial to bearing different loads, and the platform efficiency is reduced.
(2) The space is limited, and the integrated electronic case is favorable for reducing the volume of the satellite body
The existing satellite bearing structural plate generally uses a honeycomb plate with the thickness of more than 20mm, the space utilization rate of a satellite with the side length of only 200-300 mm is low, the shapes of single-machine equipment are inconsistent, the space topological structure is very difficult under the condition of not using comprehensive electrons, and meanwhile, a large amount of satellite space is wasted.
(3) External connector
In the existing satellite configuration, all the function units are respectively installed on the wall of the satellite cabin and connected with each other through connectors and wires, so the configuration space must be reserved for the connectors to connect and insert. Besides increasing the weight and the occupied volume of the socket, the risk of inserting the connector incorrectly can be increased, and the reliability of the connector can be reduced by frequent inserting.
(4) Is not beneficial to realizing modularization and serialization
The non-uniformity of the hardware interfaces makes modularization and serialization difficult, so that each single machine needs to be independently developed, and the development period and the development cost are increased.
Disclosure of Invention
In order to solve the problems in the prior art, the invention provides the microsatellite taking an integrated electronic integrated case as a main body, combines the advantages of the integrated electronic case with the actual requirements of the microsatellite, and utilizes various advantages of the integrated electronic case as a central body to improve the functional capacity of the microsatellite. The invention is realized by the following technical scheme:
the microsatellite taking the integrated electronic comprehensive case as a main body comprises a central cabin, a propulsion device, a solar cell array and a truss; the origin of the satellite coordinates is the center of a carrying butt joint surface, + X points to the carrying direction, the reverse side of the fixed solar cell array patch direction is the + Z direction, + Y and the two directions form a right-handed system; mounting the propulsion device at the top of the center nacelle; the solar cell array is connected with the central cabin body through a truss; the central cabin is a plug board type integrated electronic case and comprises functional board cards and a bottom board, wherein the bottom board is parallel to the + Z surface and is arranged at the bottom of the central cabin, and each functional board card is parallel to the bottom board and is inserted in the case.
As a further improvement of the invention, the microsatellite is provided with three solar cell arrays, wherein one solar cell array is fixed on the-Z top surface of the satellite, and the other two solar cell arrays are symmetrically arranged on the +/-Y side surface of the satellite and form a certain angle with the XOZ surface.
As a further improvement of the invention, the satellite and the carrier are connected through four-point explosive bolts, and four explosive bolt mounting corner boxes are symmetrically mounted at four corners of the + X surface of the central cabin.
As a further improvement of the invention, the propulsion device is an integrated cold air propulsion device.
As a further improvement of the invention, the central cabin is made of hard aluminum, and the solar cell array is made of carbon fiber skin aluminum honeycomb composite material.
As a further improvement of the invention, a DCS radio frequency processing board, a first DCS signal processing board, a second DCS signal processing board, a CPU board, an interface board, a measurement and control board, a secondary power supply board and a power supply controller are arranged from the bottom board to the top in sequence.
As a further improvement of the present invention, the interface conversion function is implemented on the backplane.
As a further improvement of the invention, a cover plate is additionally arranged on the + Z surface of the integrated electronic cabin and is used for installing an antenna on the ground.
As a further improvement of the invention, a momentum wheel, a magnetic torquer, a lithium ion battery, an MEMS gyroscope and a strong magnetometer are arranged on a laminate of the integrated electronic cabin.
As a further improvement of the invention, a sun sensor, a GPS antenna and an antenna for measuring and controlling the sky are arranged on the solar cell array.
The invention has the beneficial effects that: the seriation and the platformization of the microsatellite can effectively reduce the development and the production cost of the satellite, particularly, the microsatellite has the characteristics of small volume and large quantity, and the seriation has important significance.
(1) Reducing the structural weight of a type of microsatellite
The integrated electronic case of the invention is equivalent to the case that a plurality of devices of the traditional satellite are integrated in one integrated electronic case, so that the structural weight of the case of each single device can be effectively reduced. Meanwhile, the integrated electronic case structure is also used as a main structure of the satellite, so that the weight of the satellite structure can be reduced, and the weight of the redundant structure of the satellite can be effectively reduced.
(2) Improve space utilization rate
Various functional boards are installed inside the integrated electronic case, various other loads are installed outside the case, the solar cell array is supported through the truss, the whole satellite is compact in structure, and compared with a structure that multiple devices of a traditional satellite are installed inside the structure, the space utilization rate can be obviously reduced; meanwhile, the directions of the outlets of the connectors of the function boards are consistent, so that the plugging space and the plugging direction are consistent, and the space waste can be effectively avoided.
(3) Improved reliability
The various board cards installed inside the comprehensive electronic case are directly connected through the connectors without wires or cables, so that the plugging link of the connectors is reduced, the possibility of wrong plugging is avoided, the risk of the cables is reduced, and the risk factors are reduced. The reliability of the satellite is greatly improved in both the technical level and the operational level.
(4) Improved modularity of satellites
The satellite can make design specifications and standards of relevant sizes, develop series board cards and is very beneficial to rapid design, rapid assembly and rapid launching of the satellite; aiming at different load requirements, the board cards with different configurations can be conveniently selected, and satellites with different functions and different purposes are realized. Because of the same specification and standard, the interchange between the satellites of the type can be conveniently realized, and the on-orbit maintenance can be realized in the future.
Drawings
FIG. 1 is a schematic diagram of satellite coordinate definition;
FIG. 2 is a schematic diagram of the launching state of a microsatellite according to the present invention;
FIG. 3 is a schematic diagram of the operating state of the microsatellite of the present invention;
FIG. 4 is a diagram of the overall dimensions of the microsatellite in the launch state of the invention;
FIG. 5 is a schematic diagram of the relationship between the microsatellite and the carrier interface of the present invention;
FIG. 6 is a schematic diagram of a preliminary scheme of the microsatellite structure of the present invention;
FIG. 7 is a schematic view of the microsatellite configuration of the present invention;
FIG. 8 is a schematic diagram of a device layout outside a solar array;
fig. 9 is a schematic view of a center cabin layout.
Detailed Description
The invention is further described in the following description and embodiments with reference to the drawings.
The invention mainly provides a truss type microsatellite configuration taking a plugboard type integrated electronic as a central cabin, which has the advantages of small volume, high functional density, easy realization of serialization and the like.
As shown in fig. 1, the origin of the satellite coordinates is the center of the carrying interface, the + X direction points to the carrying direction, the reverse side of the fixed solar cell array patch direction is the + Z direction, and the + Y direction and the two directions form a right-handed system.
The appearance of the launching state of the microsatellite is similar to a hexahedron with a trapezoidal cross section and consists of a central cabin body, an antenna, a solar wing and the like. In order to reduce the weight and the overall dimension, the satellite body adopts the structure and the electronic cabinet integrated design.
The central cabin adopts an integrated comprehensive electronic cabin as a central cabin body, a box section type electronic cabin module is superposed, and an integrated cold air propelling device is installed at the top of the central cabin. The satellite has three solar cell arrays in total, one of which is fixed on the-Z top surface of the satellite, and the other two are symmetrically arranged on the +/-Y side surface of the satellite and form a certain angle (preferably 42 degrees) with the XOZ surface. The transmission state and flight state of the satellite are shown in fig. 2 and 3, respectively.
The three-dimensional size of the satellite in the transmitting state is 600mm multiplied by 1060mm multiplied by 423mm, wherein the size of the central cabin is 243mm multiplied by 240mm multiplied by 460mm, the area of the-Z surface in the three solar cell arrays is 300mm multiplied by 600mm, the other two areas are 600mm multiplied by 500mm, and the total area is 0.78m2The size of the satellite is shown in fig. 4.
Considering the excessive weight of the docking ring, the microsatellite uses a four-point connection with the carrying connection points distributed on a circumference of phi 350mm centered on the origin of the satellite coordinates, as shown in figure 5. The satellite and the carrying adopt a connection mode of a separation nut.
As shown in fig. 6, the satellite structure mainly includes an integrated electronic cabin 1, a propulsion device 2, a solar cell array 3, and a truss 4. The satellite is provided with three fixed solar cell arrays of-Z and +/-Y, the fixed solar cell arrays are connected with a central cabin body of the satellite through a truss 4, and the satellite is connected with a carrier through four-point explosion bolts. The integrated electronic cabin 1 is made of hard aluminum, and the solar cell array is made of carbon fiber skin aluminum honeycomb composite material. Four explosive bolt mounting corner boxes 5 are symmetrically mounted at four corners of the + X surface of the integrated electronic cabin, and a cover plate is additionally mounted at the + Z surface of the integrated electronic cabin and used for mounting an antenna on the ground. The structure of the satellite, the assembly bracket and the cabin body are designed integrally.
As shown in fig. 7, besides the standardized integrated electronic cabin equipment, other onboard equipment such as a momentum wheel 11, a magnetic torquer 12, a lithium ion battery 13, a MEMS gyroscope 14, a strong magnetometer 15, etc. are installed on a side plate of the integrated electronic cabin 1; an L-frequency-band transmitting antenna 9, an L-frequency-band receiving antenna 10, a ground measurement and control antenna 7, a data transmission antenna 8 and the like are installed on the integrated electronic cabin cover plate 6.
As shown in fig. 8, three sun sensors 26, a GPS antenna 25 and a sky survey antenna 24 are mounted on the solar cell array.
The plate card requirement of the satellite center cabin body is shown in detail in the detailed demonstration analysis of the satellite platform scheme, according to the use requirement, the center cabin has 8 functional plate cards and a bottom plate, the bottom plate is parallel to the + Z plane and is installed at the bottom of the center cabin, and other plate cards are arranged as shown in fig. 9 and sequentially arranged from the bottom plate up: a DCS radio frequency processing board 23, a first DCS signal processing board 22, a second DCS signal processing board 21, a CPU board 20, an interface board 19, a measurement and control board 18, a secondary power supply board 17 and a power supply controller 16.
The board cards of the cube star are interconnected through the upper and lower layers of contact pins, and a bottom plate is not provided, so that only interconnection among a small number of board cards can be realized, and along with the increase of the board cards, the number of the board contact pins sharply rises, occupies a large area of the board cards, and finally becomes unusable. The cubic star board card can not be inserted and pulled out, the assembly sequence and the like have uniqueness, the testing of the microsatellite is not facilitated, and the flexible assembly and replacement are also not facilitated. The invention is more suitable for the use of the microsatellite. The weight range of the microsatellite is 10-500 kg, the usable load range of the cubic satellite is very limited, and the microsatellite has wider application prospect due to the capability of bearing larger load.
The above description is only a preferred embodiment of the present invention, and is not intended to limit the scope of the present invention, and all equivalent structures or equivalent processes performed by the present invention or directly or indirectly applied to other related technical fields will be covered by the scope of the present invention.
For those skilled in the art, it should be understood that various changes, substitutions and alterations can be made herein without departing from the spirit and scope of the invention as defined by the appended claims.
Claims (6)
1. The utility model provides a microsatellite of quick-witted case is synthesized to integration electron as main part which characterized in that: the microsatellite comprises a central cabin, a propulsion device, a solar cell array and a truss; the origin of the satellite coordinate is the center of the carrying butt joint surface, + X points to the carrying direction, the reverse side of the fixed solar cell array patch direction is the + Z direction, + Y and the two directions form a right-handed system; mounting the propulsion device at the top of the center nacelle; the microsatellite is provided with three solar cell arrays, wherein one solar cell array is fixed on the-Z top surface of the satellite, the other two solar cell arrays are symmetrically arranged on the +/-Y side surface of the satellite and form a certain angle with the XOZ surface, and the solar cell arrays are connected with the central cabin body through a truss; the central cabin is a plug board type integrated electronic case and comprises a functional board card and a bottom board, an interface conversion function is realized on the bottom board, and a DCS radio frequency processing board, a first DCS signal processing board, a second DCS signal processing board, a CPU board, an interface board, a measurement and control board, a secondary power supply board and a power supply controller are sequentially arranged from the bottom board to the top; the base plate is parallel to the + Z surface and is arranged at the bottom of the central cabin, and each functional board card is parallel to the base plate and is inserted in the case; and a momentum wheel, a magnetic torquer, a lithium ion battery, an MEMS gyroscope and a strong magnetometer are arranged on a side plate of the comprehensive electronic cabin.
2. A microsatellite according to claim 1 wherein: the satellite is connected with the carrier through four-point explosion bolts, and four explosion bolt mounting corner boxes are symmetrically mounted at four corners of the + X surface of the central cabin.
3. A microsatellite according to claim 1 wherein: the propulsion device is an integrated cold air propulsion device.
4. A microsatellite according to claim 1 wherein: the central cabin is made of hard aluminum, and the solar cell array is made of carbon fiber skin aluminum honeycomb composite material.
5. A microsatellite according to claim 1 wherein: and a cover plate is additionally arranged on the + Z surface of the integrated electronic cabin and is used for installing an antenna on the ground.
6. A microsatellite according to claim 1 wherein: a sun sensor, a GPS antenna and an opposite-sky measurement and control antenna are arranged on the solar cell array.
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| CN113064194A (en) * | 2020-08-12 | 2021-07-02 | 中国科学院微小卫星创新研究院 | Internet constellation-based pico-nano satellite communication, navigation and attitude determination method |
| CN116534291B (en) * | 2023-07-05 | 2023-09-12 | 北京未来宇航空间科技研究院有限公司 | Reusable optical calibration hatch cover mechanism and satellite remote sensing camera calibration system |
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Address after: 518000 whole building of satellite building, 61 Gaoxin South Jiudao, Yuehai street, Nanshan District, Shenzhen City, Guangdong Province Patentee after: Shenzhen Aerospace Dongfanghong Satellite Co.,Ltd. Address before: 518000 whole building of satellite building, 61 Gaoxin South Jiudao, Yuehai street, Nanshan District, Shenzhen City, Guangdong Province Patentee before: AEROSPACE DONGFANGHONG DEVELOPMENT Ltd. |
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