CN117644996A - Optical satellite - Google Patents
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- CN117644996A CN117644996A CN202311667907.9A CN202311667907A CN117644996A CN 117644996 A CN117644996 A CN 117644996A CN 202311667907 A CN202311667907 A CN 202311667907A CN 117644996 A CN117644996 A CN 117644996A
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- 230000003287 optical effect Effects 0.000 title claims abstract description 83
- 238000003331 infrared imaging Methods 0.000 claims abstract description 54
- 238000009413 insulation Methods 0.000 claims abstract description 33
- 239000000758 substrate Substances 0.000 claims description 8
- 238000001816 cooling Methods 0.000 claims description 4
- 239000004642 Polyimide Substances 0.000 claims description 3
- 229920001721 polyimide Polymers 0.000 claims description 3
- 238000001228 spectrum Methods 0.000 abstract description 3
- 230000005540 biological transmission Effects 0.000 description 10
- 230000005855 radiation Effects 0.000 description 8
- 230000017525 heat dissipation Effects 0.000 description 6
- 238000013461 design Methods 0.000 description 4
- 238000010586 diagram Methods 0.000 description 4
- 238000012546 transfer Methods 0.000 description 3
- 238000003384 imaging method Methods 0.000 description 2
- 238000005259 measurement Methods 0.000 description 2
- 230000003595 spectral effect Effects 0.000 description 2
- 230000009286 beneficial effect Effects 0.000 description 1
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- 238000007906 compression Methods 0.000 description 1
- 238000001514 detection method Methods 0.000 description 1
- 230000000694 effects Effects 0.000 description 1
- 238000005265 energy consumption Methods 0.000 description 1
- 230000007613 environmental effect Effects 0.000 description 1
- 239000002360 explosive Substances 0.000 description 1
- 238000005286 illumination Methods 0.000 description 1
- 230000000977 initiatory effect Effects 0.000 description 1
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- 230000010354 integration Effects 0.000 description 1
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Abstract
The invention aims to provide an optical satellite which comprises a satellite body, a thermal infrared imaging load and a micro-optical multi-segment load, wherein the thermal infrared imaging load is arranged on one side of the satellite body along the positive direction x and is provided with a first radiating surface, and in an assembled state, the first radiating surface is positioned on one side of the optical satellite y along the positive direction y. The micro-light multiple section load is arranged on one side of the thermal infrared imaging load along the positive direction x and is provided with a second radiating surface, and in the assembled state, the second radiating surface is positioned on one side of the optical satellite along the positive direction y. And heat insulation pads are respectively arranged between the satellite body and the thermal infrared imaging load and between the thermal infrared imaging load and the glimmer-multiple-section load. The optical satellite can meet the working temperature requirements of thermal infrared imaging load and micro-light multi-spectrum load.
Description
Technical Field
The invention relates to the technical field of spacecrafts, in particular to an optical satellite.
Background
In the 'earth big data science engineering', an optical satellite is adopted to finely describe a 'human trace', and 'energy consumption, human residence pattern and offshore coastal environment' fine detection of a strong human activity area are developed, so that the acquired thermal infrared data products, town glimmer data products and offshore multispectral data products provide customized, autonomous, controllable, stable and reliable data sources for special items.
The optical satellite needs to be configured with thermal infrared imaging load and micro-light multi-spectrum load, and the two loads need to meet the requirement of coaxial same view field. For both loads, however, there is a large difference in the required operating temperature requirements for both. Thermal infrared imaging load imaging systems require a 100K (about-173 degrees celsius) low temperature environment, while the temperature requirement for low-light multispectral loads is around 20 ℃, which puts stringent demands on the configuration layout and thermal control of satellites.
It is desirable to provide an optical satellite that can meet different load operating temperatures.
Disclosure of Invention
The invention aims to provide an optical satellite which can meet the working temperature requirements of thermal infrared imaging load and micro-light multi-spectrum load.
In order to achieve the above object, an optical satellite has an x direction, a y direction, and a z direction, wherein a positive direction of the x direction is a forward direction of the optical satellite in an in-orbit flight state, a positive direction of the y direction is a negative normal direction of a track plane of the optical satellite in the in-orbit flight state, and a positive direction of the z direction is a ground observation direction of the optical satellite in the in-orbit flight state; the optical satellite includes:
a satellite body;
the thermal infrared imaging load is arranged on one side of the satellite body along the positive x direction and is provided with a first radiating surface, and in an assembled state, the first radiating surface is positioned on one side of the optical satellite in the positive y direction; and
the micro-light multiple section load is arranged on one side of the thermal infrared imaging load along the positive x direction and is provided with a second radiating surface, and in an assembled state, the second radiating surface is positioned on one side of the optical satellite in the positive y direction;
and heat insulation pads are respectively arranged between the satellite body and the thermal infrared imaging load and between the thermal infrared imaging load and the micro-light multiple section load.
In one or more embodiments, the heat insulation pad is fixed on the wall surface of the satellite body along the x-direction forward direction through screws, and a heat insulation bushing is arranged between the heat insulation pad and the wall surface of the satellite body along the x-direction forward direction; and
the thermal insulation pad is fixed on the wall surface of the thermal infrared imaging load along the positive x direction through screws, and a thermal insulation lining is arranged between the thermal insulation pad and the wall surface of the thermal infrared imaging load along the positive x direction.
In one or more embodiments, the insulation pad is made of polyimide.
In one or more embodiments, a star sensor unit is also included, the star sensor unit mounted to the micro-optic polypart load via a star sensor mount.
In one or more embodiments, the satellite further comprises a single-wing solar cell array assembly, wherein the single-wing solar cell array assembly is installed on one side of the satellite body along the negative positive y direction, and an included angle between the single-wing solar cell array assembly and the negative z direction after the single-wing solar cell array assembly is unfolded is 25-35 degrees.
In one or more embodiments, the single-wing solar cell array assembly is composed of a plurality of cell array substrates, the cell array substrates are connected through hinges, and the single-wing solar cell array assembly is connected with the satellite body through hinges.
In one or more embodiments, the thermal infrared imaging load further includes a baffle that covers the first cooling surface in a stowed state, and in a deployed state, the baffle has a normal direction that is co-directional with the z-direction and opens the first cooling surface toward the positive y-direction.
In one or more embodiments, the satellite further comprises a thruster assembly disposed on a sidewall of the satellite body along the x-direction negative direction.
In one or more embodiments, the satellite further includes an antenna assembly disposed on a sidewall of the satellite body along the positive z-direction.
In one or more embodiments, the antenna assembly includes a data antenna that is affixed to a side wall of the satellite body in the positive z-direction in a folded state and is locked by a locking mechanism.
The invention has the beneficial effects that:
in the satellite configuration, heat insulation pads are respectively arranged between the satellite body and the thermal infrared imaging load and between the thermal infrared imaging load and the micro-optical multiple section load. The thermal insulation pad can realize isolation on a heat transfer path between a thermal infrared imaging load and a micro-light multiple common section load and between a satellite body and the thermal infrared imaging load, and realize 'sectional' type independent temperature control. Simultaneously, set up the radiating surface of thermal infrared imaging load and little light many general section load towards one side of optical satellite y to the positive direction respectively for two radiating surfaces can have good radiating efficiency, thereby can satisfy the operating temperature demand of thermal infrared imaging load and little light many spectral band load.
The foregoing description is only an overview of the technical solutions of the present application, and may be implemented according to the content of the specification in order to make the technical means of the present application more clearly understood, and in order to make the above-mentioned and other objects, features and advantages of the present application more clearly understood, the following detailed description of the present application will be given.
Drawings
Various other advantages and benefits will become apparent to those of ordinary skill in the art upon reading the following detailed description of the preferred embodiments. The drawings are only for purposes of illustrating the preferred embodiments and are not to be construed as limiting the application. Also, like reference numerals are used to designate like parts throughout the accompanying drawings. In the drawings:
FIG. 1 illustrates a first perspective view of an optical satellite according to some embodiments of the present disclosure;
FIG. 2 illustrates a second perspective view of an optical satellite according to some embodiments of the present disclosure;
FIG. 3 illustrates an open state schematic diagram in accordance with some embodiments of the present optical satellite;
FIG. 4 illustrates an exploded view of some embodiments of the present optical satellite;
FIG. 5 illustrates a side schematic view of some embodiments of an optical satellite according to the present disclosure;
FIG. 6 illustrates a schematic diagram of a thermal pad installation in accordance with some embodiments of the present optical satellite;
FIG. 7 illustrates a perspective view of a single wing solar array module in an unfolded state according to some embodiments of the present optical satellite;
fig. 8 illustrates a side schematic view of a single wing solar array module in an unfolded state according to some embodiments of the present optical satellite.
Detailed Description
Embodiments of the technical solutions of the present application will be described in detail below with reference to the accompanying drawings. The following examples are only for more clearly illustrating the technical solutions of the present application, and thus are only examples, and are not intended to limit the scope of protection of the present application.
Unless defined otherwise, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application belongs; the terminology used herein is for the purpose of describing particular embodiments only and is not intended to be limiting of the application; the terms "comprising" and "having" and any variations thereof in the description and claims of the present application and in the description of the figures above are intended to cover non-exclusive inclusions.
In order to be able to meet the operating temperature requirements of thermal infrared imaging loads and micro-optical multi-band loads, according to some embodiments of the present application, an optical satellite is provided, fig. 1 shows a first perspective schematic view of an optical satellite according to some embodiments of the present application; FIG. 2 illustrates a second perspective view of an optical satellite according to some embodiments of the present disclosure; FIG. 3 illustrates an open state schematic diagram in accordance with some embodiments of the present optical satellite; FIG. 4 illustrates an exploded view of some embodiments of the present optical satellite; fig. 5 illustrates a side schematic view of some embodiments of the present optical satellite.
In order to explain the relative positional relationship between the components in the present optical satellite 100, the direction of the optical satellite 100 in the flying state is defined. The optical satellite 100 has an x-direction, a y-direction, and a z-direction, wherein the positive direction of the x-direction is defined by the +x arrow in the coordinate system shown in the figure, the positive direction of the y-direction is defined by the +y arrow in the coordinate system shown in the figure, and the positive direction of the z-direction is defined by the +z arrow in the coordinate system shown in the figure. The positive x-direction is the forward direction of the optical satellite 100 in the in-orbit flight state, the positive y-direction is the negative normal direction of the orbit plane of the optical satellite 100 in the in-orbit flight state, and the positive z-direction is the earth observation direction of the optical satellite 100 in the in-orbit flight state. It will be appreciated that, corresponding to the definition above, the negative x-direction is the backward direction of the optical satellite 100 in the in-orbit state, the negative y-direction is the positive normal direction of the orbit plane of the optical satellite 100 in the in-orbit state, and the positive z-direction is the opposite direction to the earth observation direction of the optical satellite 100 in the in-orbit state. After the x direction and the z direction are defined, the y direction is defined by a right-hand coordinate system.
The optical satellite 100 includes a satellite body 1, a thermal infrared imaging load 2 and a micro-optical multi-segment load 3, and referring to fig. 1 to 5, it can be seen that the thermal infrared imaging load 2 is disposed on one side of the satellite body 100 along the positive x-direction, the thermal infrared imaging load 2 has a first heat dissipation surface 20, and in an assembled state, the first heat dissipation surface 20 is located on one side of the optical satellite y along the positive x-direction. The micro-optical multi-segment load 3 is arranged on one side of the thermal infrared imaging load 2 along the positive direction x, and the micro-optical multi-segment load 3 is provided with a second radiating surface 30, and in the assembled state, the second radiating surface 30 is positioned on one side of the optical satellite y along the positive direction. The satellite body 1, the thermal infrared imaging load 2 and the micro-optical multi-segment load 3 assembled in the above manner are connected in series. In a specific embodiment, the thermal infrared imaging load 2 and the micro-micro multiple segment load 3 respectively have radiation plates, the first radiation surface 20 and the second radiation surface 30 are respectively surfaces of the radiation plates facing the positive y direction, and the ineffective heat source radiates heat to the deep space through the first radiation surface 20 and the second radiation surface 30.
And a heat insulation pad 4 is respectively arranged between the satellite body 1 and the thermal infrared imaging load 2 and between the thermal infrared imaging load 2 and the micro-optical multi-segment load 3. Through the heat insulation pad 4, the insulation on the heat transfer paths between the thermal infrared imaging load 2 and the micro-light multiple section load 3 and between the satellite body 1 and the thermal infrared imaging load 2 can be realized, and the sectional independent temperature control is realized. Simultaneously, set up the radiating surface of thermal infrared imaging load 2 and shimmer many section load 3 towards one side of optical satellite y to the positive direction respectively for two radiating surfaces can have good radiating efficiency, thereby can satisfy the operating temperature demand of thermal infrared imaging load and shimmer many spectral band load.
In the description of the embodiments of the present application, the technical terms "first", "second", such as "first heat dissipating surface" and "second heat dissipating surface", etc. are used only to distinguish between different objects, and are not to be construed as indicating or implying relative importance or implicitly indicating the number, specific order or primary and secondary relationships of the indicated technical features. In the description of the embodiments of the present application, the meaning of "plurality" is two or more unless explicitly defined otherwise.
Reference herein to "an embodiment" means that a particular feature, structure, or characteristic described in connection with the embodiment may be included in at least one embodiment of the present application. The appearances of such phrases in various places in the specification are not necessarily all referring to the same embodiment, nor are separate or alternative embodiments mutually exclusive of other embodiments. Those of skill in the art will explicitly and implicitly appreciate that the embodiments described herein may be combined with other embodiments.
Fig. 6 is a schematic diagram showing a mounting manner of a heat insulation pad according to some embodiments of the present optical satellite, in some embodiments of the present optical satellite, referring to fig. 4 and 6, the heat insulation pad 4 is fixed on a wall surface of the satellite body 1 along the x-direction by a screw 5, and a heat insulation bushing 6 is disposed between the heat insulation pad 4 and the wall surface of the satellite body along the x-direction. Similarly, the heat insulation pad 4 and the thermal infrared imaging load 2 have the same connection mode, namely, the heat insulation pad 4 is fixed on the wall surface of the thermal infrared imaging load 2 along the x-direction positive direction through the bolts 5, and the heat insulation lining 6 is arranged between the heat insulation pad 4 and the wall surface of the thermal infrared imaging load 2 along the x-direction positive direction. The heat transfer between the satellite body 1 and the thermal infrared imaging load 2 and between the thermal infrared imaging load 2 and the micro-optical multi-segment load 3 is further isolated by the heat insulation pad 4.
Further, in some embodiments of the present optical satellite, the insulation pad 4 is made of polyimide to accommodate the requirements for insulation in the space environment.
In some embodiments of the present optical satellite, the optical satellite 100 further includes a star sensor unit 71, where the star sensor unit 71 is a gesture determining component in the satellite, and the star sensor unit 71 is mounted on the micro-optical-multiple-segment load 3 through a star sensor bracket 72, so as to ensure that the star sensor unit 71 and the micro-optical-multiple-segment load 3 are in the same reference.
In some embodiments of the present optical satellite, the optical satellite 100 further includes a single-wing solar cell array assembly 73, where the single-wing solar cell array assembly 73 is mounted on one side of the satellite body 1 along the negative y direction, fig. 7 is a schematic perspective view illustrating the single-wing solar cell array assembly in an unfolded state according to some embodiments of the present optical satellite, and fig. 8 is a schematic side view illustrating the single-wing solar cell array assembly in an unfolded state according to some embodiments of the present optical satellite, and an angle between the single-wing solar cell array assembly 73 and the negative z direction is 25 ° to 35 ° after being unfolded. Preferably, the single wing solar array assembly 73 is unfolded at an angle of 30 ° to the negative z direction. By using a single wing solar array assembly 73 offset by an angle to accommodate the solar vector angle, optimal solar illumination is achieved.
Further, in some embodiments of the present optical satellite, the single-wing solar cell array assembly 73 is composed of a plurality of cell array substrates 730, the plurality of cell array substrates 730 are connected by a hinge 731, the single-wing solar cell array assembly 73 is connected with the satellite body 1 by the hinge 731, and the single-wing solar cell array assembly 73 is used as an energy input of the satellite. Specifically, a spring is provided at the hinge 731, by which the plurality of battery array substrates 730 can be unfolded to a designated position. The hinge 731 has a male hinge and a female hinge, and the positional relationship between the single wing solar array module 73 and the satellite body 1 and the positional relationship between the plurality of battery array substrates 730 are locked by the hinge 731 by locking pins and locking grooves on the male and female hinges.
Further, in some embodiments of the present optical satellite, the thermal infrared imaging load 2 further includes a baffle 21, where the baffle 21 covers the first heat dissipating surface 20 in the storage state, and in the unfolded state, a normal direction of the baffle 21 is aligned with the z direction, and the first heat dissipating surface 20 is opened toward the positive y direction. In one particular embodiment, the baffle 21 is an earth infrared radiation baffle. In the states shown in fig. 4, 5 and 7 to 8, the baffle 21 is in an unfolded state, and the baffle 21 does not cover the first heat dissipation surface 20 any more, so that the thermal infrared imaging load 2 can dissipate heat through the first heat dissipation surface 20. In the state shown in fig. 1, the baffle 21 is in the storage state, and at this time, the baffle 21 covers the first heat dissipation surface 20, so that the thermal infrared imaging load 2 cannot dissipate heat through the first heat dissipation surface 20. By controlling the position of the baffle 21, a proper heat external radiation channel is provided for the low temperature requirement of the thermal infrared imaging load 2, and the working temperature requirement of the thermal infrared imaging load 2 is met. In a particular embodiment, the shutter 21 is movable between the stowed and deployed conditions by means of, for example, a motor or a spring plus a locking member.
In some embodiments of the present optical satellite, the optical satellite 100 further includes a thruster assembly disposed on a sidewall of the satellite body 1 along the x-direction negative direction. In a specific embodiment, the thruster assembly realizes attitude orbit control sharing through a diagonal arrangement, and realizes satellite orbit maintenance, attitude pitching, attitude yawing and attitude rolling.
In some embodiments of the present optical satellite, the optical satellite 100 further includes an antenna assembly disposed on a side wall of the satellite body 1 along the positive z direction. In one embodiment, the antenna assembly includes a data transmission antenna, a measurement and control antenna, and a GNSS antenna.
Further, in some embodiments of the present optical satellite, the antenna assembly includes a data antenna, and the data antenna is attached to a side wall of the satellite body 1 along the positive z direction in a folded state, so as to correspond to the on-track flight to the ground, and is locked by a locking mechanism. The arrangement is that the data transmission antenna is folded on the wall surface of the satellite body 1 in the positive direction in the z direction when the satellite is transmitted, and the data transmission antenna can realize two-dimensional rotation after the optical satellite enters the orbit and is unlocked. In a specific embodiment, after the satellite enters the orbit, the compression point of the data transmission antenna is unlocked through the initiating explosive device, and after the unlocking, the data transmission antenna rotates in two dimensions through the azimuth and pitching motor. The unfolding locking mechanism is adaptively designed for the data transmission antenna, so that the folding state of the data transmission antenna is ensured to meet the requirement of the inner envelope of the carrying fairing, the data transmission antenna is unlocked in orbit, and the data transmission antenna rotates in two dimensions around the Y direction and the Z direction of the satellite body 1, so that the data transmission antenna can move in a large range, and is suitable for different working modes of satellites.
Further, in some embodiments of the present optical satellite, the GNSS antenna and the opposite-to-sky measurement and control antenna are mounted on the z-direction negative-direction wall surface of the satellite body 1, corresponding to the opposite-to-sky surface of the in-orbit flight.
Further, in some embodiments of the optical satellite, as shown in fig. 3, the satellite body 1 is in a cuboid shape, and is formed by connecting six side plates 10 with a frame, so as to form a frame panel type hexahedral structure, the load and the satellite body 1 are separately arranged, and the main bearing structure of the thermal infrared imaging load 2 is directly installed on the platform frame, so that the thermal boundary of the satellite body 1 is clear. The satellite body 1 adapts to the size envelope requirements of different loads and expands in three directions of X/Y/Z.
In the optical satellite, the requirements of a load observation field of view and the environmental requirements of emission and on-orbit operation, in particular to a low-temperature environment required by a thermal infrared imaging load imaging system are met through reasonable configuration layout design and whole satellite and load thermal control design. The configuration design of carrying the load with different temperature requirements and the same observation direction requirements on the ground remote sensing satellite is solved, the satellite carrying load can be conveniently expanded on the satellite platform, the satellite carrying load has good general assembly integration openness and operation convenience, the satellite carrying load is verified in practical engineering application, and references are provided for the general configuration design of the similar satellites.
In the description of the embodiments of the present application, unless explicitly specified and limited otherwise, the terms "mounted," "connected," "secured" and the like are to be construed broadly and may be, for example, fixedly connected, detachably connected, or integrally formed; or may be mechanically or electrically connected; can be directly connected or indirectly connected through an intermediate medium, and can be communicated with the inside of two elements or the interaction relationship of the two elements. The specific meaning of the above terms in the embodiments of the present application will be understood by those of ordinary skill in the art according to the specific circumstances.
It is to be understood that references herein to "along" a direction means that there is at least a component in that direction, preferably within 10 ° of that direction, more preferably within 5 °.
Finally, it should be noted that: the above embodiments are only for illustrating the technical solution of the present application, and not for limiting the same; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical scheme described in the foregoing embodiments can be modified or some or all of the technical features thereof can be replaced by equivalents; such modifications and substitutions do not depart from the spirit of the embodiments, and are intended to be included within the scope of the claims and description. In particular, the technical features mentioned in the respective embodiments may be combined in any manner as long as there is no structural conflict. The present application is not limited to the specific embodiments disclosed herein, but encompasses all technical solutions falling within the scope of the claims.
Claims (10)
1. An optical satellite is characterized by comprising an x direction, a y direction and a z direction, wherein the positive direction of the x direction is the advancing direction of the optical satellite in an in-orbit flight state, the positive direction of the y direction is the negative normal direction of a track surface of the optical satellite in the in-orbit flight state, and the positive direction of the z direction is the earth observation direction of the optical satellite in the in-orbit flight state; the optical satellite includes:
a satellite body;
the thermal infrared imaging load is arranged on one side of the satellite body along the positive x direction and is provided with a first radiating surface, and in an assembled state, the first radiating surface is positioned on one side of the optical satellite in the positive y direction; and
the micro-light multiple section load is arranged on one side of the thermal infrared imaging load along the positive x direction and is provided with a second radiating surface, and in an assembled state, the second radiating surface is positioned on one side of the optical satellite in the positive y direction;
and heat insulation pads are respectively arranged between the satellite body and the thermal infrared imaging load and between the thermal infrared imaging load and the micro-light multiple section load.
2. The optical satellite of claim 1 wherein the optical satellite comprises,
the heat insulation pad is fixed on the wall surface of the satellite body along the x-direction positive direction through a screw, and a heat insulation lining is arranged between the heat insulation pad and the wall surface of the satellite body along the x-direction positive direction; and
the thermal insulation pad is fixed on the wall surface of the thermal infrared imaging load along the positive x direction through screws, and a thermal insulation lining is arranged between the thermal insulation pad and the wall surface of the thermal infrared imaging load along the positive x direction.
3. The optical satellite of claim 1, wherein the thermal insulation pad is made of polyimide.
4. The optical satellite of claim 1, further comprising a star sensor unit mounted to the micro-optic dop-segment load by a star sensor mount.
5. The optical satellite of claim 1, further comprising a single-wing solar cell array assembly mounted to a side of the satellite body along the negative y-direction, the single-wing solar cell array assembly being unfolded at an angle of 25 ° to 35 ° to the negative z-direction.
6. The optical satellite of claim 5, wherein the single-wing solar array assembly is composed of a plurality of array substrates, the plurality of array substrates are connected by a hinge, and the single-wing solar array assembly is connected with the satellite body by a hinge.
7. The optical satellite of claim 1, wherein the thermal infrared imaging load further comprises a baffle that covers the first cooling surface in a stowed state, and wherein the baffle has a normal direction that is co-directional with the z-direction and opens the first cooling surface toward the positive y-direction in a deployed state.
8. The optical satellite of claim 1, further comprising a thruster assembly disposed on a sidewall of the satellite body along the negative x-direction.
9. The optical satellite of claim 1, further comprising an antenna assembly disposed on a sidewall of the satellite body along the positive z-direction.
10. The optical satellite of claim 9, wherein the antenna assembly comprises a data antenna that is attached to a side wall of the satellite body in the positive z-direction in a folded state and is locked by a locking mechanism.
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311667907.9A CN117644996A (en) | 2023-12-06 | 2023-12-06 | Optical satellite |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202311667907.9A CN117644996A (en) | 2023-12-06 | 2023-12-06 | Optical satellite |
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Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2010149659A (en) * | 2010-12-07 | 2012-06-20 | Федеральное государственное унитарное предприятие "Государственный космический научно-производственный центр имени М.В. Хруничева (ФГ | Microsatellite for remote sensing of the surface of the earth |
| US20140224939A1 (en) * | 2013-02-12 | 2014-08-14 | Lockheed Martin Corporation | Spacecraft east-west radiator assembly |
| CN104049643A (en) * | 2014-06-04 | 2014-09-17 | 北京空间飞行器总体设计部 | Adjustable vacuum thermal-insulation multilayer sunshading and cooling method |
| CN108674692A (en) * | 2018-04-04 | 2018-10-19 | 南京理工大学 | A kind of remote sensing microsatellite |
| US20190329912A1 (en) * | 2016-12-19 | 2019-10-31 | Mitsubishi Electric Corporation | Deployable radiator |
| KR102241874B1 (en) * | 2019-12-17 | 2021-04-19 | 한국항공우주연구원 | Star tracker and method for controlling the star tracker |
| CN113075788A (en) * | 2021-04-02 | 2021-07-06 | 中国科学院长春光学精密机械与物理研究所 | Multispectral multichannel common-caliber zooming imaging optical system |
| CN114721106A (en) * | 2022-03-18 | 2022-07-08 | 中国科学院西安光学精密机械研究所 | Many flexible coupling lightweight thermal-insulated bearing structure of cold science |
| CN114735252A (en) * | 2022-04-15 | 2022-07-12 | 中国科学院上海技术物理研究所 | Deep low-temperature heat dissipation system based on earth screen shielding |
-
2023
- 2023-12-06 CN CN202311667907.9A patent/CN117644996A/en active Pending
Patent Citations (9)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| RU2010149659A (en) * | 2010-12-07 | 2012-06-20 | Федеральное государственное унитарное предприятие "Государственный космический научно-производственный центр имени М.В. Хруничева (ФГ | Microsatellite for remote sensing of the surface of the earth |
| US20140224939A1 (en) * | 2013-02-12 | 2014-08-14 | Lockheed Martin Corporation | Spacecraft east-west radiator assembly |
| CN104049643A (en) * | 2014-06-04 | 2014-09-17 | 北京空间飞行器总体设计部 | Adjustable vacuum thermal-insulation multilayer sunshading and cooling method |
| US20190329912A1 (en) * | 2016-12-19 | 2019-10-31 | Mitsubishi Electric Corporation | Deployable radiator |
| CN108674692A (en) * | 2018-04-04 | 2018-10-19 | 南京理工大学 | A kind of remote sensing microsatellite |
| KR102241874B1 (en) * | 2019-12-17 | 2021-04-19 | 한국항공우주연구원 | Star tracker and method for controlling the star tracker |
| CN113075788A (en) * | 2021-04-02 | 2021-07-06 | 中国科学院长春光学精密机械与物理研究所 | Multispectral multichannel common-caliber zooming imaging optical system |
| CN114721106A (en) * | 2022-03-18 | 2022-07-08 | 中国科学院西安光学精密机械研究所 | Many flexible coupling lightweight thermal-insulated bearing structure of cold science |
| CN114735252A (en) * | 2022-04-15 | 2022-07-12 | 中国科学院上海技术物理研究所 | Deep low-temperature heat dissipation system based on earth screen shielding |
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