CN212667705U - Low-orbit communication satellite structure - Google Patents
Low-orbit communication satellite structure Download PDFInfo
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- CN212667705U CN212667705U CN202021113034.9U CN202021113034U CN212667705U CN 212667705 U CN212667705 U CN 212667705U CN 202021113034 U CN202021113034 U CN 202021113034U CN 212667705 U CN212667705 U CN 212667705U
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Abstract
A low-orbit communication satellite structure relates to a satellite structure and solves the problem that the existing thermally induced vibration phenomenon can damage the work of a load antenna of a low-orbit communication satellite, and the low-orbit communication satellite structure comprises a satellite body and the load antenna, wherein the load antenna is positioned at the top of an outer shell of the satellite body, a supporting plate is arranged at the back of an antenna surface of the load antenna, the supporting plate is formed by splicing a plurality of polygonal plates, and each polygonal plate comprises an upper constraint layer, a first vibration reduction layer and a lower constraint layer; the first vibration damping layer is arranged between the upper constraint layer and the lower constraint layer and is in an inclined shape; the first vibration reduction layer is made of a viscoelastic damping material, and the upper constraint layer and the lower constraint layer are made of composite material pressing plates. The vibration reduction layer can provide damping force for vibration in the horizontal direction and the vertical direction, vibration is reduced, and stability of the satellite load antenna is guaranteed.
Description
Technical Field
The utility model relates to a satellite structure, in particular to low earth orbit communication satellite structure.
Background
The low earth orbit communication satellite has low orbit height, so that the transmission delay is short. The path loss is small, the constellation formed by a plurality of satellites can realize real global coverage, and the frequency reuse is more effective; on the other hand, cellular communication, multiple access, spot beam, frequency reuse and other technologies also provide technical support for low-orbit satellite mobile communication. Therefore, the LEO system is considered to be the latest and most promising satellite mobile communication system. Due to the commercial use of communication satellites, low earth orbit communication satellites are transmitting more and more. The space industry is seeking greater commercial incentives as more and more companies rely on the data provided by the constellation of low earth orbit communications satellites.
The low-orbit satellite mobile communication system consists of a satellite constellation, a gateway earth station, a system control center, a network control center, a user unit and the like. A plurality of satellites are disposed in a plurality of orbital planes, and the satellites in the plurality of orbital planes are coupled by a communication link. The whole constellation is like a large platform which is structurally connected into a whole, a cellular service cell is formed on the surface of the earth, users in the service cell are covered by at least one satellite, and the users can access the system at any time.
The satellite, especially for an exposed load antenna, has a very severe working environment in space, and each period in operation needs to experience severe temperature changes, which can cause the satellite load antenna to generate a thermally induced vibration phenomenon, affect the service life of the satellite load antenna, and in severe cases, can cause the damage of key parts of the satellite, and the thermally induced vibration phenomenon can damage the stability of the satellite load antenna, and affect the operation of the satellite load antenna. In view of the thermally induced vibration phenomenon, the prior art has conducted on-orbit thermal actuation analysis of satellites and the dynamics of solar panels deployed in space environments, and has also attempted to reduce the vibration by changing the installation and structural collocation of the spacecraft, but none of the results are ideal.
Disclosure of Invention
To the problem that current thermally-induced vibration phenomenon can destroy low earth orbit communication satellite load antenna work, the utility model provides a reduce low earth orbit communication satellite structure of load antenna vibration.
The utility model discloses a low earth orbit communication satellite, including satellite body and load antenna 1, load antenna 1 is located the shell body top of satellite body, and the back of the antenna face of load antenna 1 is equipped with backup pad 110, backup pad 110 is formed by the concatenation of a plurality of polygonal plates, and each polygonal plate includes upper restraint layer 111, first damping layer 112 and lower restraint layer 113;
the first vibration damping layer 112 is arranged between the upper constraint layer 111 and the lower constraint layer 113 and is inclined;
the first vibration attenuation layer 112 is made of viscoelastic damping materials, and the upper constraint layer 111 and the lower constraint layer 113 are made of composite material pressing plates.
Preferably, the satellite body comprises an outer shell and three solar panels; the three solar panels are respectively a body installation plate 121 and two expansion plates 122, the body installation plate 121 is fixed on the outer shell, the two expansion plates 122 are distributed on two sides of the body installation plate 121 and are respectively hinged with two sides of the body installation plate 121, and the back plates and the outer shell of the three solar panels respectively comprise an inner shell plate 1212, a honeycomb layer 1213, a first constraint layer 1214, a second vibration reduction layer 1215 and a second constraint layer 1216;
the inner shell plate 1212, the honeycomb layer 1213, the first constraint layer 1214, the second vibration attenuation layer 1215 and the second constraint layer 1216 are sequentially arranged from outside to inside, wherein the second vibration attenuation layer 1215 is inclined, the first vibration attenuation layer 112 adopts viscoelastic damping material, and the inner shell plate 1212, the honeycomb layer 1213, the first constraint layer 1214 and the second constraint layer 1216 adopt composite material pressing plates.
Preferably, the satellite body further comprises a sun sensor 6, a measurement and control transponder 7, a power controller 8, a reaction flywheel 9, a measurement and control antenna 10, a star computer 5, a star sensor 4, a fiber-optic gyroscope 3, a storage battery 2 and a docking ring 11;
the outer shell comprises six panels which form a closed space, and the load antenna 1 and the sun sensor 6 are arranged at the top of the upper panel;
the star sensor 4 is fixed on the outer side of the first side panel;
the inner side of a second side panel adjacent to the first side panel is provided with a measurement and control transponder 7, a power controller 8, a reaction flywheel 9, a housekeeping computer 5, a fiber-optic gyroscope 3 and a storage battery 2, the outer shell further comprises a partition board used for forming a closed space by six panels into each space, wherein the fiber-optic gyroscope 3 and the housekeeping computer 5 are arranged in the same space, the fiber-optic gyroscope 3 is arranged above the housekeeping computer 5, the measurement and control transponder 7 and the reaction flywheel 9 are positioned in the same space, the measurement and control transponder 7 is positioned above the reaction flywheel 9, and the storage battery 2 and the power controller 8 are respectively positioned in an independent space;
the measurement and control antenna 10 is arranged on the bottom surface of the lower panel and distributed on the corners, and the measurement and control antenna 10 faces the outer side of the satellite;
the docking ring 11 is located on the bottom surface of the lower panel.
The beneficial effects of the utility model reside in that: the satellite structure of the utility model arranges a support plate at the back of the antenna surface of the load antenna 1, the support plate 110 is formed by splicing a plurality of polygonal plates, and a structure which is more laminated with the curved surface of the antenna surface can be spliced by splicing a plurality of polygons, so that the antenna surface can be effectively supported; the utility model discloses a backup pad 110 is a vibration damper, and this vibration damper adopts the damping layer of filling damping material in the clearance between two restraint layers and is in the inclined position, and when the load antenna takes place the thermal vibration, the damping layer can provide damping force for the vibration of level and two vertical directions, reduces the vibration, guarantees satellite load antenna's steady. The utility model discloses still provide the workspace, position layout, the antenna of each inside functional system of satellite directional etc. when practicing thrift the space, realize reasonable layout.
Drawings
Fig. 1 is a schematic diagram of a low earth orbit communication satellite structure according to the present invention;
FIG. 2 is a schematic view of another angle of FIG. 1;
FIG. 3 is a schematic diagram of the three solar panels of FIG. 2 shown in an unfolded state;
FIG. 4 is a schematic view of a loaded antenna;
FIG. 5 is a schematic structural view of a support plate for a loaded antenna;
FIG. 6 is a schematic view of a polygonal structure of the support plate;
FIG. 7 is a schematic view of a back sheet of a solar panel;
fig. 8 is an internal distribution diagram of a low earth orbit communication satellite structure according to the present invention.
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 the embodiments and features of the embodiments may be combined with each other without conflict.
The present invention will be further described with reference to the accompanying drawings and specific embodiments, but the present invention is not limited thereto.
The low-orbit communication satellite of the embodiment comprises a satellite body and a load antenna 1, wherein the load antenna 1 is positioned at the top of an outer shell of the satellite body, as shown in fig. 4, a support plate 110 is arranged at the back of an antenna surface of the load antenna 1 of the embodiment, as shown in fig. 5, the support plate 110 of the embodiment is formed by splicing a plurality of polygonal plates, a structure which is more attached to a curved surface of the antenna surface is spliced, and the antenna surface can be effectively supported, as shown in fig. 6, each polygonal plate comprises an upper constraint layer 111, a first vibration reduction layer 112 and a lower constraint layer 113;
the first vibration damping layer 112 is arranged between the upper constraint layer 111 and the lower constraint layer 113 and is inclined;
the first vibration attenuation layer 112 is made of viscoelastic damping materials, and the upper constraint layer 111 and the lower constraint layer 113 are made of composite material pressing plates.
The constraint layer of the embodiment is added on the basis of a free damping structure, and due to the constraint of the constraint layer, the vibration reduction layer bears larger shearing deformation and a part of tension-compression deformation when the structure vibrates, so that vibration energy is lost through the damage of the molecular chain of long-chain molecules of the viscoelastic damping material under the action of synthesized alternating stress, the damping effect of the whole structure is increased, and the amplitude of a vibration response resonance peak value can be effectively reduced under the condition that the rigidity and the mass of the shell are not obviously changed.
When the antenna surface of the load antenna vibrates in the vertical direction, the first vibration reduction layer 112 can dissipate a large amount of vibration energy due to the shearing deformation and the compression deformation of the upper constraint layer 111 and the lower constraint layer 113, so that the effect of damping the vibration in the vertical direction is achieved, and the high-frequency vibration of the antenna surface of the load antenna in the vertical direction during working can be effectively inhibited by the embodiment.
The satellite body of the embodiment comprises three solar panels; the three solar cell panels are respectively a body installation plate 121 and two expansion plates 122, the body installation plate 121 is fixed on the outer shell, the two expansion plates 122 are distributed on two sides of the body installation plate 121 and are respectively hinged with two sides of the body installation plate 121, the satellite structure of the solar cell panel in a folding state is shown in figure 2, and the satellite structure of the solar cell panel in an expansion state is shown in figure 3; as shown in fig. 7, the solar cell panel includes a solar receiving panel 1211 and a back sheet, the solar receiving panel 1211 and the back sheet are attached together, and the back sheet includes an inner skin 1212, a honeycomb layer 1213, a first constraining layer 1214, a second vibration damping layer 1215, and a second constraining layer 1216;
the inner shell plate 1212, the honeycomb layer 1213, the first constraint layer 1214, the second vibration attenuation layer 1215 and the second constraint layer 1216 are sequentially arranged from outside to inside, wherein the second vibration attenuation layer 1215 is inclined, the first vibration attenuation layer 112 is made of a viscoelastic damping material, and the inner shell plate 1212, the honeycomb layer 1213, the first constraint layer 1214 and the second constraint layer 1216 are made of a composite material laminated plate.
When the panel surface of the solar panel vibrates in the vertical direction, the vibration damping layer 131 can dissipate a large amount of vibration energy due to the shear deformation and the compression deformation of the first constraint layer 1214 and the second constraint layer 1216, so that the vibration damping effect on the vertical direction vibration is achieved, and the vibration of the solar panel is reduced.
The outer shell of the satellite body is the same as the back plate of the solar cell panel, and the vibration reduction function is achieved.
The composite material pressing plate adopts the carbon fiber resin layer/epoxy resin one-way plate, is firstly manually laid and pasted, and then is subjected to hot pressing and curing to form the composite material pressing plate; the viscoelastic damping material of the embodiment selects ZN-1 type butyl damping rubber, a glue film with the thickness of 0.1mm to 1mm, then the glue film and the carbon fiber veneer are cut together, manually laid, and finally high-temperature co-curing molding is carried out; the satellite of the embodiment adopts a honeycomb panel structure form to realize the structural connection of all functional subsystems of the satellite and the installation and support of components;
in addition, when the satellite structure is designed, the requirements of working space, position layout, antenna direction and the like of each functional system are also considered. The structure of each functional system in the satellite is shown in fig. 8. Comprises a measurement and control system, a house affair computer 5, an attitude and orbit control system, a power supply and distribution system and a butt joint ring;
the measurement and control system comprises a measurement and control transponder 7, a measurement and control antenna 10 and a load antenna 1, and has the main functions of receiving and demodulating an uplink remote control signal transmitted by a ground station, generating and executing an instruction, or transmitting the instruction to the integrated electronic system module for execution, modulating and then downloading engineering remote measurement parameters transmitted by the integrated electronic system module, acquiring the engineering parameters of the measurement and control module, transmitting the engineering parameters to the integrated electronic system module, and simultaneously finishing downlink data transmission of load data.
The satellite affair computer 5 has the main functions of performing satellite affair management on the satellite, managing each single machine component on the satellite, controlling the attitude and the orbit of the satellite, receiving and executing remote control instructions and injection data, collecting, packaging and sending remote measurement information, and has the functions of whole satellite thermal control, power supply and propulsion management.
The attitude and orbit control system comprises a reaction flywheel 9, a fiber optic gyroscope 3, a sun sensor 6 and a star sensor 4, and has the main functions of determining and controlling the attitude and the orbit and the like: the attitude control function of the satellite is used for acquiring the attitude of the satellite and applying a control moment to stabilize and maintain the current attitude of the satellite or control the attitude of the satellite to transit from the current attitude to another attitude, and the orbit control function is used for adjusting the orbit according to the task requirements of the satellite to ensure that the orbit meets the constraint of the task.
The power supply and distribution system comprises a solar cell panel, a power supply controller 8 and a storage battery 2, and mainly has the functions of providing power for the satellite during the in-orbit operation period of the satellite and meeting the power requirements of various instruments and equipment; the power supply subsystem generates power by utilizing the solar cell array in the illumination period, supplies power to all equipment of the satellite and charges the storage battery, and the storage battery supplies power to all equipment of the satellite in the earth shadow period.
The butt-joint ring 11 is used as a satellite and arrow butt-joint device to ensure reliable connection and separation between satellites and arrows.
The outer shell of the satellite body comprises six panels which form a closed space, and the load antenna 1 and the sun sensor 6 are arranged at the top of the upper panel;
the star sensor 4 is fixed on the outer side of the first side panel;
the inner side of a second side panel adjacent to the first side panel is provided with a measurement and control transponder 7, a power controller 8, a reaction flywheel 9, a housekeeping computer 5, a fiber-optic gyroscope 3 and a storage battery 2, the outer shell further comprises a partition board used for forming a closed space by six panels into each space, wherein the fiber-optic gyroscope 3 and the housekeeping computer 5 are arranged in the same space, the fiber-optic gyroscope 3 is arranged above the housekeeping computer 5, the measurement and control transponder 7 and the reaction flywheel 9 are positioned in the same space, the measurement and control transponder 7 is positioned above the reaction flywheel 9, and the storage battery 2 and the power controller 8 are respectively positioned in an independent space;
the measurement and control antenna 10 is arranged on the bottom surface of the lower panel and distributed on the corners, and the measurement and control antenna 10 faces the outer side of the satellite;
the docking ring 11 is located on the bottom surface of the lower panel.
Although the invention herein has been described with reference to particular embodiments, it is to be understood that these embodiments are merely illustrative of the principles and applications of the present invention. It is therefore to be understood that numerous modifications may be made to the illustrative embodiments and that other arrangements may be devised without departing from the spirit and scope of the present invention as defined by the appended claims. It should be understood that features described in different dependent claims and herein may be combined in ways different from those described in the original claims. It is also to be understood that features described in connection with individual embodiments may be used in other described embodiments.
Claims (3)
1. A low-earth-orbit communication satellite comprises a satellite body and a load antenna (1), and is characterized in that the load antenna (1) is positioned at the top of an outer shell of the satellite body, a supporting plate (110) is arranged at the back of an antenna surface of the load antenna (1), the supporting plate (110) is formed by splicing a plurality of polygonal plates, and each polygonal plate comprises an upper constraint layer (111), a first vibration damping layer (112) and a lower constraint layer (113);
the first damping layer (112) is arranged between the upper restraint layer (111) and the lower restraint layer (113) and is in an inclined shape;
the first vibration attenuation layer (112) is made of viscoelastic damping materials, and the upper constraint layer (111) and the lower constraint layer (113) are made of composite material pressing plates.
2. The low earth orbit communication satellite of claim 1, wherein the satellite body comprises an outer housing and three solar panels; the three solar panels are respectively a body mounting plate (121) and two expansion plates (122), the body mounting plate (121) is fixed on the outer shell, the two expansion plates (122) are distributed on two sides of the body mounting plate (121) and are respectively hinged with two sides of the body mounting plate (121), and the back plates and the outer shell of the three solar panels respectively comprise an inner shell plate (1212), a honeycomb layer (1213), a first constraint layer (1214), a second vibration reduction layer (1215) and a second constraint layer (1216);
the inner shell plate (1212), the honeycomb layer (1213), the first constraint layer (1214), the second damping layer (1215) and the second constraint layer (1216) are sequentially arranged from outside to inside, wherein the second damping layer (1215) is inclined, and the inner shell plate (1212), the honeycomb layer (1213), the first constraint layer (1214) and the second constraint layer (1216) are made of composite material pressing plates.
3. The low earth orbit communication satellite of claim 2, wherein the satellite body further comprises a sun sensor (6), a measurement and control transponder (7), a power controller (8), a reaction flywheel (9), a measurement and control antenna (10), a star computer (5), a star sensor (4), a fiber-optic gyroscope (3), a storage battery (2) and a docking ring (11);
the outer shell comprises six panels which form a closed space, and the load antenna (1) and the sun sensor (6) are arranged at the top of the upper panel;
the star sensor (4) is fixed on the outer side of the first side panel;
the inner side of a second side panel adjacent to the first side panel is provided with a measurement and control transponder (7), a power controller (8), a reaction flywheel (9), a star computer (5), a fiber-optic gyroscope (3) and a storage battery (2), the outer shell further comprises a partition board for forming a closed space with six panels into each space, wherein the fiber-optic gyroscope (3) and the star computer (5) are arranged in the same space, the fiber-optic gyroscope (3) is arranged above the star computer (5), the measurement and control transponder (7) and the reaction flywheel (9) are located in the same space, the measurement and control transponder (7) is located above the reaction flywheel (9), and the storage battery (2) and the power controller (8) are located in an independent space;
the measurement and control antenna (10) is arranged on the bottom surface of the lower panel and distributed on the corners, and the measurement and control antenna (10) faces the outer side of the satellite;
the butt-joint ring (11) is positioned on the bottom surface of the lower panel.
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| CN202021113034.9U CN212667705U (en) | 2020-06-16 | 2020-06-16 | Low-orbit communication satellite structure |
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| CN202021113034.9U CN212667705U (en) | 2020-06-16 | 2020-06-16 | Low-orbit communication satellite structure |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113879561A (en) * | 2021-11-16 | 2022-01-04 | 北京微纳星空科技有限公司 | Cube star platform and cube star |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN113879561A (en) * | 2021-11-16 | 2022-01-04 | 北京微纳星空科技有限公司 | Cube star platform and cube star |
| CN113879561B (en) * | 2021-11-16 | 2022-09-16 | 北京微纳星空科技有限公司 | Cube star platform and cube star |
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