Lightweight agile remote sensing satellite platform configuration

Abstract

The application provides a light agile remote sensing satellite platform configuration, which comprises a star body, a radar antenna, a solar cell array, a control moment gyro group, a hoisting structure, an electric propulsion system and a heat dissipation device; the star body comprises an integrated carbon fiber truss and a honeycomb plate, the integrated carbon fiber truss is in a cuboid structure as a mounting base, the honeycomb plate is arranged in the integrated carbon fiber truss, the control moment gyro group is in a five-prism pyramid configuration and is tightly mounted on the honeycomb plate; the solar cell array and the radar antenna are oppositely mounted on two sides of the integrated carbon fiber truss, the radar antenna comprises a plurality of sub-arrays which are movably connected, and the solar cell array comprises a plurality of sub-panels which are movably connected; the hoisting structure is arranged on the top of the integrated carbon fiber truss, the electric propulsion system is modularly integrally mounted on the star body, and the heat dissipation device is arranged on the outer side of the integrated carbon fiber truss. The application achieves the good effects of light weight, small envelope and fast attitude maneuvering.

Description

TECHNICAL FIELD

[0001] The present application relates to the field of satellite overall configuration layout design, in particular to a light agile remote sensing satellite platform configuration. BACKGROUND

[0002] The remote sensing satellites of the traditional planar phased array antenna in China are all large platform configurations. In the face of the new demand for light and small agile remote sensing satellites of planar phased array antennas, it is urgent to solve the new problems brought by the satellite structure in the aspects of the proportion of the overall weight of the satellite, the large-size planar phased array radar antenna installation envelope limitation, the large-area solar cell array installation envelope, the rapid agile attitude maneuver of the satellite body, the rapid hoisting of the satellite, and the modularized electric propulsion configuration.

[0003] Remote Sensing Satellite No. 35 is a light agile remote sensing satellite of China. Its payload adopts a planar phased array radar antenna. The overall weight is strictly constrained, the envelope is limited, the satellite on-orbit working mode is complex, and the attitude rapid maneuver is frequent. Therefore, it is urgently needed to create a satellite platform configuration with light weight, small envelope, and high modularity to adapt to the rapid on-orbit attitude maneuver, so as to meet the new demand for light and small remote sensing satellite platforms.

[0004] The existing Chinese patent with the publication number CN111703592B discloses a large commercial remote sensing satellite platform configuration and assembly method. The platform configuration includes a satellite platform main structure, a docking ring, and a solar cell array. The satellite platform main structure includes a docking plate, a storage box plate, a partition plate, a camera plate, an outer plate, and a truss rod assembly. The storage box plate and the partition plate are arranged on the docking plate. The storage box plate and the partition plate are each provided with multiple plates and are arranged in a vertical and horizontal manner to form a cross-shaped structure. The camera plate is arranged above the cross-shaped structure and is supported by the truss rod assembly. The outer plate is arranged around the cross-shaped structure. The adjacent two outer plates are connected by an outer skin. The solar cell array is arranged on the outer surface of the outer plate. The docking ring is fixed below the docking plate. Each subsystem single machine device is installed on the storage box plate and the partition plate. The propulsion subsystem assembly is installed on the docking plate. Two storage boxes are symmetrically installed relative to the platform center of mass.

[0005] At present, the remote sensing satellites of the traditional planar phased array antenna are all large platforms, and the commercial satellite platforms are light and small. There is no configuration demand suitable for the planar phased array antenna. Remote Sensing Satellite No. 35 has a weight of less than 500 kg. It cannot directly inherit the mature technology on the traditional remote sensing satellite. It is necessary to innovate the new satellite platform configuration technology and break through the ideas according to the working characteristics of light weight, envelope limitation, frequent and rapid attitude maneuver, and modularized configuration. Therefore, the inventor believes that it is necessary to provide a light agile remote sensing satellite platform configuration to solve the above problems. SUMMARY

[0006] Aiming at the defects in the prior art, the present application aims to provide a light agile remote sensing satellite platform configuration.

[0007] The light agile remote sensing satellite platform configuration provided by the present application comprises a star body, a radar antenna, a solar cell array, a control moment gyro group, a hoisting structure, an electric propulsion system and a heat dissipation device.

[0008] The star body comprises an integrated carbon fiber truss and a honeycomb panel, the integrated carbon fiber truss is in a cuboid structure as a mounting base, the honeycomb panel is arranged inside the integrated carbon fiber truss, and the control moment gyro group is in a five-prism pyramid configuration and is fastened and mounted on the honeycomb panel.

[0009] The solar cell array and the radar antenna are oppositely mounted on both sides of the integrated carbon fiber truss, the radar antenna comprises a plurality of sub-arrays connected movably, the radar antenna can switch between an unfolded state and a folded state, the solar cell array comprises a plurality of sub-panels connected movably, and the solar cell array can switch between an unfolded state and a folded state.

[0010] The top of the integrated carbon fiber truss is provided with the hoisting structure, the electric propulsion system is modularly integrally mounted on the star body, and the outer side of the integrated carbon fiber truss is provided with the heat dissipation device.

[0011] Preferably, the honeycomb panel is used as a single-machine mounting base on the star and adopts an I-shaped configuration, the honeycomb panel is detachably mounted inside the integrated carbon fiber truss, and the web of the honeycomb panel is vertically arranged and is arranged in parallel with the solar cell array.

[0012] Preferably, the radar antenna is in a planar phased array configuration, the radar antenna comprises a plurality of vertically arranged sub-arrays, two adjacent sub-arrays are connected through a torsional spring hinge mechanism, and a middle sub-array is mounted on the integrated carbon fiber truss.

[0013] When the radar antenna is in the unfolded state, the plurality of sub-arrays are combined into a complete planar phased array antenna.

[0014] When the radar antenna is in the folded state, the plurality of sub-arrays cover three sides of the integrated carbon fiber truss.

[0015] Preferably, the solar cell array is in a body-mounted configuration, the solar cell array comprises a plurality of sub-panels, two adjacent sub-panels are connected through a spring hinge mechanism, and a middle panel of the solar cell array is fastened and mounted on the integrated carbon fiber truss.

[0016] When the solar cell array is in the unfolded state, the plurality of sub-panels are combined into a complete solar cell array.

[0017] When the solar cell array is in the folded state, the sub-panels other than the middle panel are folded and arranged on the middle panel.

[0018] Preferably, the control moment gyro group comprises five control moment gyroscopes which are mounted at the center of the honeycomb panel through a support, the rotation axes of the five control moment gyroscopes intersect at a common point, and the line connecting the center point of the five control moment gyroscopes and the intersection point of the rotation axes is in the shape of a space five-prism, and the top angle of the prism is an acute angle.

[0019] Preferably, the hoisting structure comprises a hoisting point metal embedded part arranged in the top rod joint of the integrated carbon fiber truss.

[0020] Preferably, the propellant of the electric propulsion system is stored in a gas cylinder, the electric propulsion system is arranged on the carbon fiber panel, and the electric propulsion system is fastened to the star body through the carbon fiber panel.

[0021] Preferably, aluminum alloy joints mechanically connected to the launch vehicle are arranged at the four bottom joints of the integrated carbon fiber truss.

[0022] Preferably, an integrated star sensor support is arranged at the top of the integrated carbon fiber truss, and a plurality of mounting surfaces for mounting star sensors are arranged on the integrated star sensor support.

[0023] Preferably, the two outer sides of the integrated carbon fiber truss parallel to the radar antenna array surface are covered with a heat control multilayer cover, and the other two outer sides are provided with aluminum alloy heat sinks.

[0024] Compared with the prior art, the present application has the following beneficial effects:

[0025] 1. By adopting the integrated carbon fiber truss structure configuration technology, the radar antenna splicing and combining planar array configuration technology, the body-mounted solar cell array configuration technology, the control moment gyro group five-prism configuration technology, the structure top hoisting technology, the modular electric propulsion configuration technology and the high-efficiency heat dissipation configuration technology, the present application can solve the new requirements brought by light weight, envelope limitation and frequent and rapid attitude maneuvering of the satellite, and achieve good effects of light weight, small envelope and rapid attitude maneuvering.

[0026] 2. By adopting the integrated carbon fiber truss and honeycomb panel combination form, the carbon fiber material is lighter in weight, and has better carrying capacity and rigidity strength than the traditional honeycomb panel material; an integrated support is designed at the top of the integrated carbon fiber truss, a plurality of star sensors can be assembled, aluminum alloy metal joints are designed at the bottom for connection with the launch vehicle; and two thin aluminum alloy metal plates are designed to adhere to the two outer sides of the integrated carbon fiber truss perpendicular to the radar antenna array surface, so as to significantly provide heat dissipation capacity.

[0027] 3. The application solves the satellite envelope restriction constraint, meets the large-size planar phased array radar antenna and large-size solar cell array satellite loading requirements, has small weight and moment of inertia, and cancels the driving mechanism and connecting rod of the traditional solar cell array, thereby reducing the weight of the whole satellite.

[0028] 4. The application solves the satellite envelope restriction constraint, meets the large-size planar phased array radar antenna and large-size solar cell array satellite loading requirements, has small weight and moment of inertia, and cancels the driving mechanism and connecting rod of the traditional solar cell array, thereby reducing the weight of the whole satellite.

[0029] 5. The application solves the satellite envelope restriction constraint, meets the large-size planar phased array radar antenna and large-size solar cell array satellite loading requirements, has small weight and moment of inertia, and cancels the driving mechanism and connecting rod of the traditional solar cell array, thereby reducing the weight of the whole satellite.

[0030] 6. The application solves the satellite envelope restriction constraint, meets the large-size planar phased array radar antenna and large-size solar cell array satellite loading requirements, has small weight and moment of inertia, and cancels the driving mechanism and connecting rod of the traditional solar cell array, thereby reducing the weight of the whole satellite. BRIEF DESCRIPTION OF DRAWINGS

[0031] Other features, objects, and advantages of the application will become more apparent from the following detailed description of non-limiting embodiments, read in conjunction with the accompanying drawings:

[0032] Figure 1 The application mainly embodies the modular configuration combination schematic diagram of a light agile remote sensing satellite platform configuration;

[0033] Figure 2 The application mainly embodies the structural schematic diagram of a satellite body;

[0034] Figure 3Fig. 1 is a schematic diagram showing the I-beam honeycomb panel moving up and down from the truss structure, which is the main embodiment of the present application.

[0035] Figure 4 Fig. 2 is a schematic diagram showing the radar antenna retracted configuration, which is the main embodiment of the present application.

[0036] Figure 5 Fig. 3 is a schematic diagram showing the radar antenna deployed configuration, which is the main embodiment of the present application.

[0037] Figure 6 Fig. 4 is a schematic diagram showing the solar array retracted configuration, which is the main embodiment of the present application.

[0038] Figure 7 Fig. 5 is a schematic diagram showing the solar array deployed configuration, which is the main embodiment of the present application.

[0039] Figure 8 Fig. 6 is a schematic diagram showing the control moment gyro cluster configuration, which is the main embodiment of the present application.

[0040] Figure 9 Fig. 7 is a schematic diagram showing the integrated star sensor support installation, which is the main embodiment of the present application.

[0041] Figure 10 Fig. 8 is a schematic diagram showing the integrated star sensor support configuration, which is the main embodiment of the present application.

[0042] Figure 11 Fig. 9 is a schematic diagram showing the integrated carbon fiber truss hoisting point position, which is the main embodiment of the present application.

[0043] Figure 12 Fig. 10 is a schematic diagram showing the integrated carbon fiber truss and the aluminum alloy joint connected to the carrier, which is the main embodiment of the present application.

[0044] Figure 13 Fig. 11 is a schematic diagram showing the electric propulsion modular configuration, which is the main embodiment of the present application.

[0045] Figure 14 Fig. 12 is a schematic diagram showing the aluminum alloy sheet structure, which is the main embodiment of the present application.

[0046] In the drawings:

[0047] Star body 1 Radar antenna 2 Solar array 3

[0048] Control moment gyro cluster 4 Hoisting structure 5 Integrated carbon fiber truss 11

[0049] Honeycomb panel 12 DETAILED DESCRIPTION

[0050] The present invention will now be described in detail with reference to specific embodiments. These embodiments will help those skilled in the art to further understand the present invention, but do not limit the invention in any way. It should be noted that those skilled in the art can make several changes and improvements without departing from the concept of the present invention. These all fall within the protection scope of the present invention.

[0051] like Figures 1-14 As shown, a lightweight and agile remote sensing satellite platform configuration provided by the present invention includes: a satellite body 1, a radar antenna 2, a solar cell array 3, a control moment gyroscope group 4, a hoisting structure 5, an electric propulsion system, and a heat dissipation device; the satellite body 1 includes an integrated carbon fiber truss 11 and a honeycomb panel 12. The integrated carbon fiber truss 11 serves as the mounting base and has a cuboid structure. The honeycomb panel 12 is disposed inside the integrated carbon fiber truss 11. The control moment gyroscope group 4 has a pentagonal pyramidal structure and is securely mounted on the honeycomb panel 12; the solar cell array 3 is mounted opposite to the radar antenna 2 on both sides of the integrated carbon fiber truss 11. The radar antenna 2 includes multiple movable subarrays and can switch between an extended state and a retracted state. The solar cell array 3 includes multiple movable subplates and can switch between an extended state and a folded state; a hoisting structure 5 is provided on the top of the integrated carbon fiber truss 11; the electric propulsion system is modularly and integrally mounted on the satellite body 1; and a heat dissipation device is provided on the outer surface of the integrated carbon fiber truss 11.

[0052] The structure of Satellite 1 consists of an integrated carbon fiber truss 11 and honeycomb panels 12. The integrated carbon fiber truss 11 is a rectangular column and serves as the main load-bearing structure, providing the mounting foundation for the main payloads and solar arrays on the satellite, and primarily bearing the mechanical environment of the launch vehicle's active phase. The honeycomb panels 12 serve as the mounting foundation for individual units on the satellite, employing an I-beam configuration. The waist plates of the honeycomb panels 12 are vertically positioned and parallel to the solar array 3. The honeycomb panels 12 are detachably installed inside the integrated carbon fiber truss 11 and can be flexibly removed from the integrated carbon fiber truss 11 structure, generally moving upwards within the integrated carbon fiber truss 11.

[0053] The radar antenna 2 is designed as a planar phased array. It is constructed by dividing the radar antenna 2 into several subarrays along the range direction, which are then vertically retracted and mounted on an integrated carbon fiber truss 11 from three directions. Before operation, the radar antenna 2 unfolds all subarrays via a torsion spring hinge mechanism, assembling them into a complete planar phased array antenna. Preferably, the radar antenna 2 includes three vertically arranged subarrays, with adjacent subarrays connected by a torsion spring hinge mechanism. The middle subarray is mounted on the integrated carbon fiber truss 11. When the radar antenna 2 is in the unfolded state, the three subarrays combine to form a complete planar phased array antenna. When the radar antenna 2 is in the retracted state, the three subarrays cover the three sides of the integrated carbon fiber truss 11.

[0054] The solar cell array 3 is in a body-mounted configuration, and includes a plurality of sub-panels. Two adjacent sub-panels are connected by a spring hinge mechanism. The middle panel of the solar cell array 3 is fixedly mounted on the integrated carbon fiber truss 11. The remaining sub-panels are sequentially folded and compressed on the middle panel. When the satellite is launched into orbit, the remaining sub-panels of the solar cell array 3 are sequentially unfolded by the spring hinge mechanism to form a complete solar cell array. Preferably, the solar cell array 3 includes three vertically arranged sub-panels. When the solar cell array 3 is in an unfolded state, the three sub-panels form a complete solar cell array. When the solar cell array 3 is in a folded state, the sub-panels other than the middle panel are folded and arranged on the middle panel.

[0055] The control moment gyro group 4 is designed in a pentagonal pyramid configuration. The control moment gyro group 4 includes five control moment gyroscopes. The control moment gyroscopes are mounted at the center of the honeycomb panel 12 by a support. The rotation axes of the five control moment gyroscopes intersect at a common point. The line connecting the center point of the five control moment gyroscopes and the intersection point of the rotation axes is in the shape of a space pentagonal pyramid. The apex angle of the pyramid is a fixed acute angle.

[0056] The hoisting structure 5 includes hoisting point metal inserts arranged in the top bar joint of the integrated carbon fiber truss 11, i.e., four end joint areas at the top of the rectangular column, which are designed in an integrated manner with the truss structure. High-strength metal inserts are pre-embedded in the bar joints in each end area. During hoisting, lifting lugs are installed at the four joints for vertical hoisting of the satellite.

[0057] The propellant of the electric propulsion system is stored in gas cylinders. The electric propulsion system integrates the propulsion products together using a 3mm carbon fiber plate, as a modular single structure, which can independently complete integrated welding and is convenient for parallel development and installation of the system. The carbon fiber plate is provided with high-pressure gas cylinders. The high-pressure gas cylinders are connected with a pressure reducing valve and a plurality of valves. After completing pipeline welding integration and propellant filling, the electric propulsion system is connected and fixed to the satellite body 1 as a modular whole through the carbon fiber plate.

[0058] The bottom four joints of the integrated carbon fiber truss 11 are respectively provided with aluminum alloy joints mechanically connected with the launch vehicle. The aluminum alloy metal structure in the joints of the bottom four ends of the integrated carbon fiber truss 11 is designed in an integrated manner with the truss structure, achieving mechanical connection with the launch vehicle.

[0059] The top side of the integrated carbon fiber truss 11 is provided with an integrated star sensor support, the bottom side of which is connected to the integrated carbon fiber truss 11 as a common mounting surface, and a plurality of mounting surfaces for mounting star sensors are arranged on the integrated star sensor support. The top of the integrated star sensor support is provided with a plurality of mounting surfaces, each of which is connected to a different star sensor. Preferably, three star sensors are mounted on the top of the integrated star sensor support. Mounting a plurality of star sensors on one integrated support not only helps to reduce the weight of the support, but also helps to unify the installation precision reference of the plurality of star sensors.

[0060] The integrated carbon fiber truss 11 has a cuboid structure and does not have honeycomb panels on the six outer sides. For the four outer sides of the integrated carbon fiber truss 11, two outer sides parallel to the array surface of the radar antenna 2 adopt thermal control multilayer covering, and the other two outer sides are provided with aluminum alloy heat dissipation plates. In the direction of the two outer sides of the integrated carbon fiber truss 11 perpendicular to the array surface of the radar antenna 2, a complete 0.5mm thick aluminum alloy metal sheet is used for mounting, which is directly screwed to the outer side of the integrated carbon fiber truss 11 to replace the traditional 25mm thick honeycomb panel structure, greatly reducing the weight, and the sheet is painted for easy heat dissipation. On the outer sides of the integrated carbon fiber truss 11 parallel to the array surface of the radar antenna 2, thermal control multilayer covering is used, and buckles are attached to the members of the integrated carbon fiber truss 11 for multilayer pasting.

[0061] The present application mainly adopts integrated carbon fiber truss structure configuration technology, radar antenna splicing combined planar array configuration technology, body-mounted solar cell array configuration technology, control moment gyro group five-prism configuration technology, structure top hoisting technology, and modular electric propulsion configuration technology and high-efficiency heat dissipation configuration technology, which can solve the new requirements brought by light weight, envelope limitation and frequent and rapid attitude maneuvering of satellites.

[0062] In view of the constraint problem of light weight of satellites, the present application is designed in the form of integrated carbon fiber truss 11 and honeycomb panel 12 combination. Carbon fiber material is lighter in weight than traditional honeycomb panel material, and has better carrying capacity and stiffness strength. The integrated carbon fiber truss 11 structure configuration technology is adopted to design the whole satellite main structure from the traditional honeycomb panel structure to the integrated carbon fiber truss, which is used as the outer frame main body of the satellite structure, cancels all the external side plates of the satellite structure, and only retains a I-shaped configuration honeycomb panel 12 combined and embedded in the truss structure for mounting single machines on the satellite, and it can be flexibly removed from the inside of the integrated carbon fiber truss 11 structure.

[0063] At the same time, the top of the integrated carbon fiber truss 11 is designed with an integrated support, which can assemble three star sensors; the bottom is designed with an aluminum alloy metal joint, which is used for connection with the launch vehicle; at the same time, in order to ensure the temperature environment of the single machine in the cabin, two thin aluminum alloy metal plates are designed, which are attached to the two outer sides of the integrated carbon fiber truss 11 perpendicular to the array plane of the radar antenna 2, which can significantly provide heat dissipation capacity.

[0064] In view of the problem of strict satellite envelope limitation, through the radar antenna 2 splicing combination plane configuration technology, the planar phased array radar antenna exceeding the envelope size of the carrier is divided into several small sub-arrays along one direction, the division size of each sub-array cannot exceed the width of the integrated carbon fiber truss 11, and the sub-arrays are vertically folded and installed on the three outer sides of the integrated carbon fiber truss 11 through the torsion spring hinge mechanism. Through the body-mounted solar cell array 3 configuration technology, the solar cell array 3 exceeding the envelope size of the carrier is divided into several sub-panels, and the middle panel is vertically fixed to one outer side of the integrated carbon fiber truss 11 structure, and the remaining sub-panels are successively folded and installed on the middle panel through the torsion spring hinge mechanism. The above configuration technology solves the satellite envelope limitation constraint, meets the demand of large-size planar phased array radar antenna and large-size solar cell array, has small weight and moment of inertia, and cancels the driving mechanism and connecting rod of the traditional solar cell array, thereby reducing the weight of the whole satellite.

[0065] In view of the demand of on-orbit rapid agile attitude maneuvering of remote sensing satellites, through the five-pyramid configuration technology of the control moment gyro group 4, five large-torque control moment gyroscopes are installed on the I-shaped configuration honeycomb panel 12 through a support, the rotation axes of the five control moment gyroscopes intersect at a common point, the connecting line between the center points of the five control moment gyroscopes and the intersection points of the rotation axes forms a spatial five-pyramid shape, and the top angle of the pyramid is a fixed angle. The five control moment gyroscopes are installed compactly, the on-orbit installation space envelope of the satellite is small, and the on-orbit rapid attitude maneuvering requirement of the satellite can be met.

[0066] In view of the demand of satellite rapid lifting, the satellite lifting point is designed at the four corner end member joint at the top of the integrated carbon fiber truss 11 structure, a high-strength aluminum alloy embedded part is pre-embedded in the member joint, and the lifting point is connected as a whole through a tooling lifting device, thereby providing a satellite rapid lifting use.

[0067] In view of the configuration demand of the electric propulsion system, a 3mm-thick carbon fiber plate is designed, and the electric propulsion system products are all welded and integrated on the carbon fiber plate, so as to realize the modular integration of the electric propulsion system, and the whole passes through the connection between the carbon fiber plate and the structure plate of the satellite body.

[0068] The application aims at the urgent demand of light weight, small envelope and rapid attitude maneuvering in orbit of new light agile remote sensing satellite, solves the problem that traditional large platform remote sensing satellite cannot solve by integrated carbon fiber truss structure configuration technology, radar antenna splicing combination planar array configuration technology, body-mounted solar cell array configuration technology, control moment gyroscope group five-prism configuration technology and electric propulsion system modular configuration technology, and creates a new satellite platform configuration.

[0069] The application solves the problem of multi-type light agile remote sensing satellite platform configuration, realizes good effect of light weight, small envelope and rapid attitude maneuvering, and fills the configuration design blank of planar phased array antenna remote sensing satellite on light small satellite platform.

[0070] The application is specifically illustrated by the following data:

[0071] As a not more than 500 kg level high-resolution remote sensing satellite configured with planar phased array antenna, remote sensing satellite No. 35 has clear constraints on the weight and envelope size of the whole satellite structure, radar antenna, solar cell array and control system actuator products. The structure weight of remote sensing satellite No. 35 accounts for not more than 10% of the whole satellite, the azimuth size of the radar antenna is 3390 mm, the distance size is also 3390 mm, the solar cell array area is 7 square meters, the on-orbit working mode switching requires that the attitude maneuvering in front, back, left and right directions can be met, and the time cannot exceed 25 seconds. The launch vehicle requires that the satellite envelope is not more than 2200 mm.

[0072] The main load-bearing structure of the whole satellite is designed as an open integrated carbon fiber truss 11 structure configuration, the wall thickness of the truss bar is designed to be 1.2 mm, all the side plates outside the truss structure are cancelled, and at the same time, a I-shaped honeycomb plate 12 combination configuration is designed to ensure single machine installation, which is built-in the integrated carbon fiber truss 11 structure and can be flexibly moved out from the truss structure. On the premise of ensuring the structural stiffness, bearing capacity and single machine installation demand of the whole satellite, the weight of the whole satellite structure accounts for only 9.3% of the whole satellite.

[0073] The radar antenna splicing combination planar array configuration technology of the application divides the radar antenna 3 along the distance direction into three rectangular sub-arrays, the distance size of which is 1130 mm, and the azimuth size remains 3390 mm, which are respectively vertically installed on the integrated carbon fiber truss 11 structure of the satellite body to form a "π" configuration. The horizontal envelope size under this configuration is 2070 mm, which meets the launch envelope requirement of the launch vehicle. After the satellite is launched into orbit, the three rectangular sub-arrays are spliced into a complete planar phased array antenna through the action of the torsion spring hinge mechanism, and the size of the antenna is azimuth 3390 mm and distance 3390 mm.

[0074] The body-mounted solar cell array 3 configuration technology of the application divides a 7 square solar cell array 3 into three rectangular sub-panels, which are left sub-panel, middle sub-panel and right sub-panel. The middle sub-panel of the solar cell array 3 is directly fixed on the integrated carbon fiber truss 11 of the satellite body. The left sub-panel and the left side of the middle sub-panel are connected by a torsion spring hinge mechanism. The right sub-panel and the right side of the middle sub-panel are connected by a torsion spring hinge mechanism. The left sub-panel and the right sub-panel are folded and pressed on the middle sub-panel in turn, and the solar cell array configuration envelope under this configuration is 2100mm, which meets the launch envelope range requirements of the launch vehicle. This solar cell array configuration technology cancels the traditional solar cell array driving mechanism and connecting rod, directly connects the solar cell array sub-panel and the satellite body, and greatly reduces the weight of the solar cell array.

[0075] The control moment gyro group 4 five pyramid configuration technology of the application fully utilizes the large torque performance of the control moment gyro, greatly reduces the use of the flywheel, and installs the five control moment gyro through the support on the I-shaped honeycomb panel 12. The rotation axes of the five control moment gyro intersect at a common point. The connecting line between the center points of the five control moment gyro and the intersection points of the rotation axes forms a spatial five pyramid configuration, and the top angle of the pyramid is designed as 45°. The control moment gyro under this five pyramid configuration has small space envelope requirements for the satellite layout, which can solve the situation of tight satellite layout space and meet the satellite attitude maneuver in different directions of the satellite attitude. Through the on-orbit measurement results, it is known that the attitude rapid maneuvering time is not more than 25 seconds.

[0076] The top rapid lifting technology of the application utilizes the high strength characteristics of the integrated carbon fiber truss 11 structure, directly designs the aluminum alloy embedded part in the rod joint of the integrated carbon fiber truss 11, and arranges it in the four corner end area of the top structure facing the sky. Through the ground tool connection, it forms a whole. The tool and the lifting point are connected at ordinary times. The top can meet the requirements of rapid lifting and placing other deployment tools. When launching, the tool connected with the lifting point is removed, which can greatly reduce the weight of the satellite.

[0077] The modular electric propulsion system technology of the application designs a 3mm thick carbon fiber plate, which is appropriately hollowed out according to the installation interface on the satellite to reduce weight. The propellant of the electric propulsion system is stored in a gas cylinder. The gas cylinder is fixed to the carbon fiber plate. The remaining products are all layered and integrated on the carbon fiber plate. Finally, the electric propulsion system can be modularized as a whole, which can be quickly installed on the satellite and removed.

[0078] The application innovatively solves the urgent needs of the remote sensing No. 35 satellite in weight, envelope, agile maneuvering and modular assembly. The application creates a new design method for the remote sensing satellite platform configuration technology with subsequent planar phased array antennas, which is particularly widely used for light agile remote sensing satellite platform configuration.

[0079] In the description of the present application, it should be understood that the terms "upper", "lower", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer" and the like refer to the orientation or positional relationship shown in the drawings, and are only intended to facilitate the description of the present application and simplify the description, and are not intended to indicate or imply that the device or element referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore cannot be understood as a limitation on the present application.

[0080] The specific embodiments of the present application are described above. It should be understood that the present application is not limited to the above specific embodiments, and various changes or modifications can be made by those skilled in the art within the scope of the claims, which do not affect the essential content of the present application. The embodiments of the present application and the features in the embodiments can be arbitrarily combined with each other without conflict, provided that they do not conflict.

Claims

1. A light-weight agile remote sensing satellite platform configuration, characterized by, The application relates to a satellite (1), a radar antenna (2), a solar cell array (3), a control moment gyro group (4), a hoisting structure (5), an electric propulsion system and a heat dissipation device. The satellite (1) comprises an integrated carbon fiber truss (11) and a honeycomb plate (12), the integrated carbon fiber truss (11) is in a cuboid structure as a mounting base, the honeycomb plate (12) is arranged inside the integrated carbon fiber truss (11), the control moment gyro group (4) is in a five-prism pyramid configuration and is tightly mounted on the honeycomb plate (12); The solar cell array (3) is oppositely mounted on both sides of the integrated carbon fiber truss (11) with the radar antenna (2), the radar antenna (2) comprises a plurality of sub-arrays which are movably connected, the radar antenna (2) can be switched between an unfolded state and a folded state, the solar cell array (3) comprises a plurality of sub-panels which are movably connected, and the solar cell array (3) can be switched between an unfolded state and a folded state; The hoisting structure (5) is arranged on the top of the integrated carbon fiber truss (11), the electric propulsion system is modularly integrally mounted on the satellite (1), and the outer side of the integrated carbon fiber truss (11) is provided with the heat dissipation device; The honeycomb plate (12) is used as a single-machine mounting base on the satellite and adopts an I-shaped configuration, the honeycomb plate (12) can be detachably mounted inside the integrated carbon fiber truss (11), and the waist plate of the honeycomb plate (12) is vertically arranged and parallel to the solar cell array (3); The radar antenna (2) is in a planar phased array configuration, the radar antenna (2) comprises a plurality of vertically arranged sub-arrays, two adjacent sub-arrays are connected through a torsional spring hinge mechanism, and a middle sub-array is mounted on the integrated carbon fiber truss (11); When the radar antenna (2) is in the unfolded state, a plurality of the sub-arrays are combined into a complete planar phased array antenna; When the radar antenna (2) is in the folded state, a plurality of the sub-arrays cover three sides of the integrated carbon fiber truss (11); The solar cell array (3) is in a body-mounted configuration, the solar cell array (3) comprises a plurality of sub-panels, two adjacent sub-panels are connected through a spring hinge mechanism, and a middle plate of the solar cell array (3) is tightly mounted on the integrated carbon fiber truss (11); When the solar cell array (3) is in the unfolded state, a plurality of the sub-panels are combined into a complete solar cell array; When the solar cell array (3) is in the folded state, all the sub-panels except the middle plate are folded and arranged on the middle plate; The control moment gyro group (4) comprises five control moment gyroscopes, the control moment gyroscopes are mounted on the center of the honeycomb plate (12) through supports, the rotation axes of the five control moment gyroscopes intersect at a common point, a line connecting the center point of the five control moment gyroscopes and the intersection point of the rotation axes is in a space five-prism pyramid shape, and the top angle of the pyramid is an acute angle; Two outer sides of the integrated carbon fiber truss (11) which are parallel to the array surface of the radar antenna (2) adopt a heat control multilayer covering, and the other two outer sides are provided with aluminum alloy heat dissipation plates. ​ 2. The light-weight agile remote sensing satellite platform configuration of claim 1, wherein, The hoisting structure (5) comprises a hoisting point metal embedded part arranged in a top rod joint of the integrated carbon fiber truss (11).

3. The light-weight agile remote sensing satellite platform configuration of claim 1, wherein, The propellant of the electric propulsion system is stored in a gas cylinder, the electric propulsion system is arranged on a carbon fiber plate, and the electric propulsion system is fastened and connected with the star body (1) through the carbon fiber plate.

4. The light-weight agile remote sensing satellite platform configuration of claim 1, wherein, Four bottom joints of the integrated carbon fiber truss (11) are respectively provided with aluminum alloy joints mechanically connected with the carrier rocket.

5. The light-weight agile remote sensing satellite platform configuration of claim 1, wherein, An integrated star sensor support is arranged on the top of the integrated carbon fiber truss (11), and a plurality of mounting surfaces for mounting star sensors are arranged on the integrated star sensor support.

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