Long-distance retrograde orbit satellite load layout structure and satellite assembly

Abstract

The invention relates to a long-distance retrograde orbit satellite load layout structure and a satellite assembly. The long-distance retrograde orbit satellite load layout structure comprises a satellite body, a satellite-borne atomic clock, a satellite-ground or inter-satellite link measurement and communication load, a space observation camera, a laser communication test load, a laser ranging time difference measurement load and a space scientific detection load, wherein the satellite-borne atomic clock is mounted on the inner side surface of the satellite body in a heat conduction manner; the satellite-ground or inter-satellite link measurement and communication load comprises an in-satellite transmitting assembly and an out-satellite antenna, the in-satellite transmitting assembly is installed on the inner side face of the satellite body in a heat conduction mode, and the out-satellite antenna is installed on the outer side face of the satellite body in a heat insulation mode. The space observation camera is installed on the outer side of the satellite body. The laser communication test loads are installed on the inner side and the outer side of the satellite body. The laser ranging time difference measurement load is thermally mounted on the inner side surface of the satellite body; space scientific detection loads are installed on the inner side face and the outer side face of the satellite body. And through reasonable load layout, low-energy-consumption orbit injection of the DRO satellite can be realized.

Description

Technical Field

[0001] This invention relates to the field of satellite-related technologies, specifically to a long-distance retrograde orbit satellite payload layout structure and satellite components. Background Technology

[0002] The "DRO Exploration and Research in Lunar Space" is one of the strategic leading science and technology projects. It aims to achieve low-energy orbit insertion and long-term stable parking of DROs, verify the autonomous cruise capability of wide-area flight in Lunar space based on DROs, and carry out gamma-ray detection research based on DROs. It can play a strategic leading role in the development and utilization of Lunar space.

[0003] The DRO project includes the development of three small satellites (DRO satellite, cruise satellite, and LEO satellite) and more than 10 payloads. The total mass of satellite A (DRO satellite) and satellite B (cruise satellite) will not exceed 600 kg, and the final mass must match the launch vehicle's carrying capacity. Satellite L (LEO satellite) will not exceed 120 kg. Based on the mission objectives, mission indicators, overall mission plan, and taking full account of the rocket's carrying capacity constraints, satellite A (~320 kg, DRO) and satellite B (~275 kg, Earth-Moon resonance orbit) will be launched as a dual-satellite combination on a single rocket, carrying an application payload with a mass not exceeding 104.6 kg. Satellite L (~120 kg, LEO) will carry an application payload with a mass of approximately 35 kg. The payloads on the three small satellites are designed to support the construction of inter-satellite links in the Earth-Moon constellation, inter-satellite communication measurements, space-based autonomous orbit determination, wide-area autonomous cruise, and other functional requirements, as well as to support space science exploration using the Earth-Moon space. Therefore, the satellite payload layout needs to consider heat dissipation, weight requirements, and link construction requirements. Summary of the Invention

[0004] In order to solve one or more technical problems existing in the prior art, the present invention provides a long-distance retrograde orbit satellite payload layout structure and satellite components.

[0005] The technical solution of this invention to solve the above-mentioned technical problems is as follows: A long-distance retrograde orbit satellite payload layout structure includes a satellite body, an onboard atomic clock, a satellite-to-ground or inter-satellite link measurement and communication payload, a space observation camera, a laser communication experimental payload, a laser ranging time difference measurement payload, and a space science exploration payload. The onboard atomic clock is thermally mounted on the inner side of the satellite body. The satellite-to-ground or inter-satellite link measurement and communication payload includes an internal transmitting component and an external antenna. The internal transmitting component is thermally mounted on the inner side of the satellite body, and the external antenna is thermally insulated and mounted on the outer side of the satellite body. The space observation camera is thermally insulated and mounted on the outer side of the satellite body. Part of the laser communication experimental payload is thermally mounted on the inner side of the satellite body, and another part of the laser communication experimental payload is thermally insulated and mounted on the outer side of the satellite body. The laser ranging time difference measurement payload is thermally mounted on the inner side of the satellite body. Part of the space science exploration payload is thermally mounted on the inner side of the satellite body, and another part of the space science exploration payload is thermally insulated and mounted on the outer side of the satellite body.

[0006] The beneficial effects of this invention are as follows: The payload layout structure for a long-distance retrograde orbit (DRO) satellite, through a reasonable payload arrangement, can meet the weight requirements of the DRO satellite (i.e., the long-distance retrograde orbit satellite of this invention), enabling low-energy orbit insertion and significantly reducing the cost of accessing the Earth-Moon space, thus opening a new path for large-scale Earth-Moon space development and utilization. This invention also facilitates the construction of large-scale satellite constellations in the Earth-Moon space, enabling the exploration of the unique properties and application value of long-distance retrograde orbits. The payload layout structure for a long-distance retrograde orbit satellite, i.e., a long-distance retrograde orbit satellite, requires that the payload layout meet heat dissipation requirements. Payloads of the same type are placed close together, and the cable paths connecting them are shortened, reducing the satellite's weight.

[0007] Based on the above technical solution, the present invention can be further improved as follows.

[0008] Furthermore, the spaceborne atomic clock includes a hydrogen clock, a first rubidium clock, and a second rubidium clock, all of which are thermally mounted on the inner side of the -Y plate of the satellite body via screws.

[0009] Furthermore, the onboard transmission assembly includes a transmitter, a processor, a power amplifier, and a microwave network. The transmitter, processor, and power amplifier are thermally mounted on the inner side of the -Y plate of the satellite body using screws, and the microwave network is thermally mounted on the inner side of the +Z plate of the satellite body using screws. The offboard antenna includes a K-load 0.6m aperture antenna, a K-load 0.22m aperture antenna, and a K-load quasi-omnidirectional antenna. The K-load 0.6m aperture antenna and the K-load 0.22m aperture antenna are thermally insulated and mounted on the outer side of the +Z plate of the satellite body, and the K-load quasi-omnidirectional antenna is thermally insulated and mounted on the outer side of the -X plate of the satellite body.

[0010] The beneficial effects of adopting the above-mentioned further scheme are as follows: considering the satellite's requirements for long-distance high-precision measurement, beam coverage of the Earth, and short-distance wide beam coverage, a high-gain large-aperture antenna (K-load 0.6m aperture antenna), a medium-gain medium-aperture antenna (K-load 0.22m aperture antenna), and a low-gain small-aperture horn antenna (K-load quasi-omnidirectional antenna) are configured simultaneously.

[0011] Furthermore, the laser communication test payload includes an optomechanical body, an optomechanical control box, and a signal processor. The optomechanical body is thermally insulated and installed on the outer side of the +Z plate of the satellite body. The optomechanical control box and the signal processor are both thermally conductively installed on the inner side of the -Y plate of the satellite body using screws.

[0012] Furthermore, the laser ranging time difference measurement payload includes a laser time difference measuring instrument, a passive laser corner reflector, and a laser corner anti-shading cover. The laser time difference measuring instrument is thermally mounted on the inner side of the +Z plate of the satellite body via screws. The passive laser corner reflector is thermally insulated and mounted on the outer side of the +Z plate of the satellite body. The laser corner anti-shading cover is thermally insulated and mounted on the outer side of the +Z plate of the satellite body and blocks the outside of the passive laser corner reflector.

[0013] Furthermore, the space science exploration payload includes a GTM electrical control box and multiple GTP probes. The GTM electrical control box is thermally mounted on the inner side of the +Y plate of the satellite body via screws, and the multiple GTP probes are thermally insulated and mounted on the outer side of the satellite body.

[0014] Furthermore, at least one GTP probe is thermally insulated and installed on the outer surfaces of the -Z plate, -X plate, -Y plate, and +Y plate of the satellite body.

[0015] Furthermore, the thermal insulation installation includes using fiberglass thermal insulation pads for thermal insulation installation; the inner surface of the satellite body is provided with heat pipes for heat dissipation.

[0016] The present invention also provides a satellite assembly, including the long-distance retrograde orbit satellite payload layout structure as described above, and also including a cruise satellite; During launch: The +X plate of the cruise satellite is connected to the outer side of the +X plate of the satellite body through the satellite-rocket separation assembly; Orbit insertion phase: The cruise satellite separates from the satellite body, and the long-distance retrograde orbit satellite and the cruise satellite enter the DRO orbit. The long-distance retrograde orbit satellite remains in the DRO orbit for a long time, and the cruise satellite will subsequently enter the RO orbit; the outer side of the +Z plate of the satellite body faces the ground, and the outer side of the +Z plate of the cruise satellite faces the ground.

[0017] The beneficial effects of this invention are: This invention sends long-distance retrograde orbit satellites to DRO orbits. DRO is a type of orbit that is about 310,000 to 450,000 kilometers away from Earth and about 70,000 to 100,000 kilometers away from the Moon. It can be regarded as a transportation hub connecting Earth, the Moon and deep space. It has the advantages of satellites being able to enter with low energy consumption, park stably, and reach the entire region with low energy consumption.

[0018] Furthermore, it also includes a LEO satellite, which enters a sun-synchronous orbit during the orbit insertion phase, with its +Z plate pointing towards the satellite body. Attached Figure Description

[0019] Figure 1 This is a schematic diagram of the payload layout structure of the long-distance retrograde orbit satellite of the present invention. Figure 1 ; Figure 2 This is a schematic diagram of the payload layout structure of the long-distance retrograde orbit satellite of the present invention. Figure 2 ; Figure 3 This is a schematic diagram of the payload layout structure of the long-distance retrograde orbit satellite of the present invention. Figure 3 ; Figure 4 This is a schematic diagram of the payload layout structure of the long-distance retrograde orbit satellite of the present invention. Figure 4 ; Figure 5 This is a schematic diagram of the payload layout structure of the long-distance retrograde orbit satellite of the present invention. Figure 5 ; Figure 6 This is a schematic diagram of the payload layout structure of the long-distance retrograde orbit satellite of the present invention. Figure 6 .

[0020] The attached diagram lists the components represented by each number as follows: 100. +X plate; 101. -X plate; 102. +Y plate; 103. -Y plate; 104. +Z plate; 105. -Z plate; 106. Heat pipe; 107. Solar panel; 108. Star sensor; 109. Payload management unit; 200. Hydrogen clock; 201. First rubidium clock; 202. Second rubidium clock; 300. Transmitter; 301. Processor; 302. Power Amplifier; 303. Microwave Network; 400, K-load 0.6m aperture antenna; 401, K-load 0.22m aperture antenna; 402, K-load quasi-omnidirectional antenna; 500. Optomechanical main body; 501. Optomechanical electrical control box; 502. Signal processor; 600. Laser time difference measuring instrument; 601. Passive laser corner reflector; 602. Laser corner reflector; 700, GTM electrical control box; 701, GTP probe; 800. Space observation camera. Detailed Implementation

[0021] The principles and features of the present invention are described below. The examples given are only for explaining the present invention and are not intended to limit the scope of the present invention.

[0022] Example 1 like Figures 1-6 As shown, this embodiment of a long-distance retrograde orbit satellite payload layout structure includes a satellite body, an onboard atomic clock, a satellite-to-ground or inter-satellite link measurement and communication payload, a space observation camera 800, a laser communication experimental payload, a laser ranging time difference measurement payload, and a space science exploration payload. The onboard atomic clock is thermally mounted on the inner side of the satellite body. The satellite-to-ground or inter-satellite link measurement and communication payload includes an internal transmitting component and an external antenna. The internal transmitting component is thermally mounted on the inner side of the satellite body, and the external antenna is thermally insulated and mounted on the outer side of the satellite body. The space observation camera 800 is thermally insulated and mounted on the outer side of the satellite body. Part of the laser communication experimental payload is thermally mounted on the inner side of the satellite body, and another part of the laser communication experimental payload is thermally insulated and mounted on the outer side of the satellite body. The laser ranging time difference measurement payload is thermally mounted on the inner side of the satellite body. Part of the space science exploration payload is thermally mounted on the inner side of the satellite body, and another part of the space science exploration payload is thermally insulated and mounted on the outer side of the satellite body.

[0023] like Figure 1 and Figure 4 As shown, the space observation camera 800 in this embodiment is a narrow field-of-view camera, specifically arranged on the outer surface of the +Z plate 104 of the satellite body and adjacent to the -X plate 101 and -Y plate 103. It detects and images targets at relatively long distances (including spacecraft targets hundreds of kilometers away, the Earth, the Moon, etc.), providing technical accumulation for situational awareness in the Earth-Moon space. Due to platform resource limitations, a lightweight space camera solution is adopted.

[0024] like Figure 5 and Figure 6As shown, in one specific embodiment, the onboard atomic clock includes a hydrogen clock 200, a first rubidium clock 201, and a second rubidium clock 202. All three clocks are thermally mounted on the inner side of the -Y plate 103 of the satellite body using screws. In this embodiment, the hydrogen clock 200, the first rubidium clock 201, and the second rubidium clock 202 are used for onboard timing. Onboard timing is a prerequisite and foundation for the satellite to achieve autonomous flight and space-based autonomous orbit determination services. Using the hydrogen clock 200, the first rubidium clock 201, and the second rubidium clock 202 enables precise autonomous timekeeping and time synchronization between the satellite and ground stations, as well as between satellites. Furthermore, the hydrogen clock 200, the first rubidium clock 201, and the second rubidium clock 202 are arranged sequentially and close together along the Z-axis on the inner side of the -Y plate 103 of the satellite body, near the +X plate, facilitating heat dissipation and reducing the cable paths between them, thus decreasing the satellite's weight. The hydrogen clock 200, the first rubidium clock 201, and the second rubidium clock 202 can provide high-precision frequency signals.

[0025] like Figure 5 and Figure 6As shown, specifically, the onboard transmission assembly includes a transmitter 300, a processor 301, a power amplifier 302, and a microwave network 303. The transmitter 300, processor 301, and power amplifier 302 are thermally mounted on the inner side of the -Y plate 103 of the satellite body using screws. The microwave network 303 is thermally mounted on the inner side of the +Z plate 104 of the satellite body using screws. The offboard antenna includes a K-load 0.6m aperture antenna 400, a K-load 0.22m aperture antenna 401, and a K-load quasi-omnidirectional antenna 402. The K-load 0.6m aperture antenna 400 and the K-load 0.22m aperture antenna 401 are thermally insulated and mounted on the outer side of the +Z plate 104 of the satellite body. The K-load quasi-omnidirectional antenna 402 is thermally insulated and mounted on the outer side of the -X plate 101 of the satellite body. The onboard transmitting component also includes waveguides, both inside and outside the satellite, primarily connecting the power amplifiers, microwave network, and antennas of the microwave link, and also connecting the transmitter and antennas. By configuring a 0.6m aperture K-band antenna 400, a 0.22m aperture K-band antenna 401, and a quasi-omnidirectional K-band antenna 402, the three satellites in the subsequent satellite assembly mainly utilize K-band microwave links to achieve inter-satellite link communication networking and tracking measurements. K-band microwaves are characterized by high frequency, narrow beam, and concentrated energy, making them suitable for constructing long-distance inter-satellite links in the Earth-Moon space, enabling inter-satellite ranging, time-frequency transfer, and inter-satellite communication. The power consumption and quality requirements of the K-band microwave link payload are suitable for microsatellite platforms and have been applied in inter-satellite communication and onboard relay communication of the BeiDou satellite system. The key technologies of the hardware system are relatively mature, while the software system varies depending on the mission. All three satellites in the subsequent satellite assembly will be equipped with K-band microwave transmitting / receiving systems, with specific configuration schemes varying according to the inter-satellite link communication requirements, using a nominal design scheme that can support at least a 500,000-kilometer link construction.

[0026] like Figure 5 and Figure 6 As shown, specifically, the laser communication test payload includes an optomechanical main body 500, an optomechanical control box 501, and a signal processor 502. The optomechanical main body 500 is thermally insulated and mounted on the outer surface of the +Z plate 104 of the satellite body. The optomechanical control box 501 and the signal processor 502 are both thermally conductively mounted on the inner surface of the -Y plate 103 of the satellite body using screws. The laser communication test payload is used to conduct experiments on Earth-Moon distance communication technology, and also accommodates modulated continuous wave ranging codes.

[0027] Specifically, such as Figure 5 and Figure 6As shown, in this embodiment, the transmitter 300, processor 301, power amplifier 302, optomechanical control box 501, and signal processor 502 are all installed on the inner side of the -Y plate 103 of the satellite body. The heat pipe 106 can pass through these components for heat dissipation. That is, the transmitter 300, processor 301, power amplifier 302, optomechanical control box 501, and signal processor 502 are pressed onto the heat pipe 106. These components are the electronic control and processing equipment of the satellite body. They are arranged close to each other on the -Y plate 103 of the satellite body for easy heat dissipation. Moreover, the cable paths between them are short, which reduces the weight of the satellite.

[0028] like Figure 5 and Figure 6 As shown, specifically, the laser ranging time difference measurement payload includes a laser time difference measuring instrument 600, a passive laser corner reflector 601, and a laser corner reflector shield 602. The laser time difference measuring instrument 600 is thermally mounted on the inner side of the +Z plate 104 of the satellite body via screws. The passive laser corner reflector 601 is thermally insulated and mounted on the outer side of the +Z plate 104 of the satellite body. The laser corner reflector shield 602 is thermally insulated and mounted on the outer side of the +Z plate 104 of the satellite body, and blocks the outside of the passive laser corner reflector 601. The passive laser corner reflector 601 and the laser time difference measuring instrument 600 provide a cost-effective, reliable, and stable independent verification method for satellite-to-ground (or inter-satellite) ranging and time synchronization (periodically observed). The laser communication test payload primarily aims to achieve high-speed satellite-to-ground (or inter-satellite) communication while also considering ranging and time synchronization. The passive laser corner reflector 601 is used for satellite-to-ground laser reflection ranging, and periodic ranging measurements are used to verify the precise orbit determination results. The Laser Time Difference Meter 600 is used to record the arrival time of ground laser pulses, supports satellite-to-ground time comparison, and performs periodic measurements to verify the accuracy of satellite-to-ground clock difference estimation.

[0029] Compared to K-band microwave links, laser links offer advantages in measurement accuracy and communication capacity. They also have lower payload weight and power consumption, making them easier to lightweight. However, they present greater technical challenges in link construction and aiming / tracking. Nevertheless, with advancements in pointing control technology, their tracking performance is continuously improving. Long-distance retrograde orbit satellites also employ passive laser corner reflectors to verify the accuracy of satellite-to-ground distance measurements and orbit determination, and are equipped with laser time difference measurement instruments to verify the accuracy of satellite-to-ground clock reconciliation.

[0030] like Figure 5 and Figure 6As shown, specifically, the space science payload includes a GTM control box 700 and multiple GTP probes 701. The GTM control box 700 is thermally mounted on the inner side of the +Y plate 102 of the satellite body via screws, and the multiple GTP probes 701 are thermally insulated and mounted on the outer side of the satellite body. The space science payload is also a gamma-ray transient source monitor, used to detect gamma rays and generate transient source early warning information.

[0031] like Figures 1-6 As shown, in this embodiment, at least one GTP probe 701 is thermally insulated and installed on the outer surfaces of the -Z plate 105, -X plate 101, -Y plate 103 and +Y plate 102 of the satellite body.

[0032] The long-distance retrograde orbit satellite payload layout structure of this embodiment also includes a payload management unit 109. The payload management unit 109 is thermally insulated and mounted on the outer side of the +Y plate 102. The payload management unit 109 provides unified management and scheduling for each application payload, and is in an independent chassis structure. The payload management unit interacts with the platform management unit via a CAN bus and with each payload through various interfaces. As the information transmission hub between the payload system and the satellite platform, the payload management unit provides information communication services to the payload, provides uplink and downlink channels for payload data, and realizes payload function control, data management, application data acquisition, processing, storage, and data transmission. Through efficient inter-satellite / satellite-to-ground communication protocols and routing planning, it constructs a lunar-Earth space-based orbit determination and autonomous navigation information network, achieving efficient utilization of inter-satellite / satellite-to-ground links.

[0033] Furthermore, the heat insulation installation includes using a fiberglass heat insulation pad for heat insulation installation; the inner side of the satellite body is provided with a heat pipe 106 for heat dissipation, which can be installed by passing the hydrogen clock 200, the first rubidium clock 201 and the second rubidium clock 202, that is, the hydrogen clock 200, the first rubidium clock 201 and the second rubidium clock 202 are pressed onto the heat pipe 106, and heat dissipation is carried out by the heat pipe 106.

[0034] In this embodiment, solar panels 107 are mounted on the outer surfaces of both the +Y plate 102 and the -Y plate 103 of the long-distance retrograde orbit satellite payload layout structure. The structure and installation method of the solar panels 107 can adopt the common satellite solar panel structure and installation method. A star sensor 108 is mounted on the outer surface of the -Z plate 105. The star sensor 108 is a common satellite component, and its installation method and structure are also common satellite structures.

[0035] In this embodiment, the commonly used XYZ coordinates of satellites are used to indicate the structural orientation of the satellite body. -X, -Y, and -Z are the opposite directions to +X, +Y, and +Z. In this embodiment of a long-distance retrograde orbit satellite payload layout structure, the main focus is on setting the layout of the entire satellite payload to meet the requirements of heat dissipation, weight, and subsequent space link. The payloads on the satellite body can all use commonly used satellite equipment or structural principles.

[0036] This embodiment presents a payload layout structure for a long-distance retrograde orbit (DRO) satellite. Through a reasonable payload layout, it can meet the weight requirements of the DRO satellite (i.e., the long-distance retrograde orbit satellite of this invention), enabling the DRO satellite to enter orbit with low energy consumption. This significantly reduces the cost of accessing the Earth-Moon space and opens up a new path for large-scale development and utilization of the Earth-Moon space. This invention also facilitates the construction of a large-scale satellite constellation in the Earth-Moon space, enabling the exploration of the unique properties and application value of long-distance retrograde orbits.

[0037] Example 2 Based on Embodiment 1, this embodiment provides a satellite component, including a long-distance retrograde orbit satellite payload layout structure as described above, and also includes a cruise satellite; During launch: The +X plate of the cruise satellite is connected to the outer side of the +X plate 100 of the satellite body via the satellite-rocket separation assembly; Orbit insertion phase: The cruise satellite separates from the satellite body, and the long-distance retrograde orbit satellite and the cruise satellite enter the DRO orbit. The long-distance retrograde orbit satellite remains in the DRO orbit for a long time, and the cruise satellite will subsequently enter the RO orbit; the outer surface of the +Z plate 104 of the satellite body faces the ground, and the outer surface of the +Z plate of the cruise satellite faces the ground.

[0038] Furthermore, the satellite components in this embodiment also include a LEO satellite, which enters a sun-synchronous orbit during the orbit insertion phase, with its +Z plate pointing towards the satellite body.

[0039] This embodiment sends a long-distance retrograde orbit satellite to a DRO orbit. DRO is a type of orbit that is about 310,000 to 450,000 kilometers from Earth and about 70,000 to 100,000 kilometers from the Moon. It can be regarded as a transportation hub connecting Earth, the Moon and deep space, and has the advantages of satellites being able to enter with low energy consumption, park stably, and reach the entire region with low energy consumption.

[0040] Three satellites—a long-range retrograde orbit satellite (DRO satellite, satellite A), a cruise satellite (satellite B), and a LEO satellite (satellite L)—are each equipped with a K-payload and deployed in different orbits. Satellite B, cruising in Earth-Moon orbit, determines its own orbit and clock bias through two-way communication measurements with the LEO and DRO satellites. Satellites A and B simultaneously orbit the Moon, and can also fly alongside each other (satellite B orbiting satellite A). The primary mission of the space camera is to search for and explore satellite A during satellite B's approach; and to image satellites B and A together through their respective space cameras during their flybys and orbits. Additionally, the space camera on satellite A can also image the Earth and the Moon.

[0041] In this embodiment, the satellite assembly involves three satellites and one ground station during missions. Each satellite has different resource conditions and applicable distances for inter-satellite measurements, therefore, the payload configuration and technical requirements for each satellite also differ. Considering its suitability for long-distance high-precision measurements, beam coverage of the Earth, and short-range wide-beam coverage, the long-range retrograde orbit satellite with the above-described layout is equipped with a high-gain large-aperture antenna, a medium-gain medium-aperture antenna, and a low-gain small-aperture horn antenna.

[0042] In the description of this invention, it should be understood that the terms “X”, “Y”, “Z”, etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limiting this invention.

[0043] Furthermore, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this invention, "a plurality of" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0044] In this invention, unless otherwise explicitly specified and limited, the terms "installation," "connection," "linking," and "fixing," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components, unless otherwise explicitly limited. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0045] In this invention, unless otherwise explicitly specified and limited, "above" or "below" the second feature can mean that the first feature is in direct contact with the second feature, or that the first feature is in indirect contact with the second feature through an intermediate medium. Furthermore, "above," "over," and "on top" of the second feature can mean that the first feature is directly above or diagonally above the second feature, or simply that the first feature is at a higher horizontal level than the second feature. "Below," "below," and "under" the second feature can mean that the first feature is directly below or diagonally below the second feature, or simply that the first feature is at a lower horizontal level than the second feature.

[0046] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., indicate that a specific feature, structure, material, or characteristic described in connection with that embodiment or example is included in at least one embodiment or example of the present invention. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. Moreover, without contradiction, those skilled in the art can combine and integrate the different embodiments or examples described in this specification, as well as the features of different embodiments or examples.

[0047] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention.

Claims

1. A long-distance retrograde orbit satellite payload layout structure, characterized in that, The system includes a satellite body, an onboard atomic clock, a satellite-to-ground or inter-satellite link measurement and communication payload, a space observation camera, a laser communication experimental payload, a laser ranging time difference measurement payload, and a space science exploration payload. The onboard atomic clock is thermally mounted on the inner surface of the satellite body. The satellite-to-ground or inter-satellite link measurement and communication payload includes an internal transmitting component and an external antenna. The internal transmitting component is thermally mounted on the inner surface of the satellite body, and the external antenna is thermally insulated and mounted on the outer surface of the satellite body. The space observation camera is thermally insulated and mounted on the outer surface of the satellite body. Part of the laser communication experimental payload is thermally mounted on the inner surface of the satellite body, while another part is thermally insulated and mounted on the outer surface. The laser ranging time difference measurement payload is thermally mounted on the inner surface of the satellite body. A portion of the space science exploration payload is thermally mounted on the inner surface of the satellite body, while another part is thermally insulated and mounted on the outer surface.

2. The long-distance retrograde orbit satellite payload layout structure according to claim 1, characterized in that, The spaceborne atomic clock includes a hydrogen clock, a first rubidium clock, and a second rubidium clock. The hydrogen clock, the first rubidium clock, and the second rubidium clock are all thermally mounted on the inner side of the -Y plate of the satellite body using screws.

3. The long-distance retrograde orbit satellite payload layout structure according to claim 1, characterized in that, The onboard transmission assembly includes a transmitter, a processor, a power amplifier, and a microwave network. The transmitter, processor, and power amplifier are thermally mounted on the inner side of the -Y plate of the satellite body using screws. The microwave network is thermally mounted on the inner side of the +Z plate of the satellite body using screws. The offboard antennas include a K-load 0.6m aperture antenna, a K-load 0.22m aperture antenna, and a K-load quasi-omnidirectional antenna. The K-load 0.6m aperture antenna and the K-load 0.22m aperture antenna are thermally insulated and mounted on the outer side of the +Z plate of the satellite body. The K-load quasi-omnidirectional antenna is thermally insulated and mounted on the outer side of the -X plate of the satellite body.

4. The long-distance retrograde orbit satellite payload layout structure according to claim 1, characterized in that, The laser communication test payload includes an optomechanical main body, an optomechanical control box, and a signal processor. The optomechanical main body is thermally insulated and installed on the outer side of the +Z plate of the satellite body. The optomechanical control box and the signal processor are both thermally conductively installed on the inner side of the -Y plate of the satellite body using screws.

5. The long-distance retrograde orbit satellite payload layout structure according to claim 1, characterized in that, The laser ranging time difference measurement payload includes a laser time difference meter, a passive laser corner reflector, and a laser corner anti-shading cover. The laser time difference meter is thermally mounted on the inner side of the +Z plate of the satellite body via screws. The passive laser corner reflector is thermally insulated and mounted on the outer side of the +Z plate of the satellite body. The laser corner anti-shading cover is thermally insulated and mounted on the outer side of the +Z plate of the satellite body and blocks the outside of the passive laser corner reflector.

6. The long-distance retrograde orbit satellite payload layout structure according to claim 1, characterized in that, The space science exploration payload includes a GTM electrical control box and multiple GTP probes. The GTM electrical control box is thermally mounted on the inner side of the +Y plate of the satellite body via screws, and the multiple GTP probes are thermally insulated and mounted on the outer side of the satellite body.

7. The long-distance retrograde orbit satellite payload layout structure according to claim 6, characterized in that, At least one GTP probe is thermally insulated and installed on the outer surfaces of the -Z plate, -X plate, -Y plate, and +Y plate of the satellite body.

8. The long-distance retrograde orbit satellite payload layout structure according to claim 1, characterized in that, The thermal insulation installation includes using fiberglass thermal insulation pads for thermal insulation installation; the inner side of the satellite body is provided with heat pipes for heat dissipation.

9. A satellite component, characterized in that, Including a long-distance retrograde orbit satellite payload layout structure as described in any one of claims 1 to 8, it also includes cruise satellites; During launch: The +X plate of the cruise satellite is connected to the outer side of the +X plate of the satellite body through the satellite-rocket separation assembly; Orbit insertion phase: The cruise satellite separates from the satellite body, the long-distance retrograde orbit satellite enters the DRO orbit, and the cruise satellite enters the RO orbit; The outer side of the +Z plate of the satellite body faces the ground, and the outer side of the +Z plate of the cruise satellite faces the ground.

10. A satellite component according to claim 9, characterized in that, It also includes LEO satellites, which enter a sun-synchronous orbit during the orbit insertion phase, with the +Z plate of the LEO satellite pointing towards the satellite body.

Images