Light agile remote sensing satellite platform configuration

Abstract

The invention provides a light agile remote sensing satellite platform configuration. The light agile remote sensing satellite platform configuration comprises a satellite body, a radar antenna, a solar cell array, a control moment gyroscope group, a hoisting structure, an electric propulsion system and a heat dissipation device, the star body comprises an integrated carbon fiber truss and a cellular board, the integrated carbon fiber truss serves as an installation foundation and is of a cuboid structure, the cellular board is arranged in the integrated carbon fiber truss, and the control moment gyroscope group is of a pentagonal pyramid structure and is fixedly installed on the cellular board; the solar cell array and the radar antenna are oppositely installed on the two sides of the integrated carbon fiber truss, the radar antenna comprises a plurality of movably connected sub-arrays, and the solar cell array comprises a plurality of movably connected sub-plates; a hoisting structure is arranged on the top of the integrated carbon fiber truss, the electric propulsion system is integrally installed on a star body in a modularized mode, and a heat dissipation device is arranged on the outer side face of the integrated carbon fiber truss. According to the invention, good effects of light weight, small envelope and rapid attitude maneuver are realized.

Description

Technical Field

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

[0002] Traditional domestic remote sensing satellites with planar phased array antennas are all large-platform configurations. Faced with the new demand for lightweight, miniaturized and agile remote sensing satellites with planar phased array antennas, there is an urgent need to solve new problems brought about by the satellite structure in the proportion of the whole satellite weight, the envelope limitations of large-size planar phased array radar antennas on the satellite, the envelope of large-area solar arrays on the satellite, rapid and agile attitude maneuvers of the satellite, rapid satellite lifting and modular electric propulsion configuration, etc.

[0003] The Remote Sensing Satellite No. 35 is a light and agile remote sensing satellite of my country. Its payload uses a planar phased array radar antenna. The weight constraint of the entire satellite is strict, the envelope restrictions are many, the satellite's on-orbit working mode is complex, and the rapid attitude maneuvers are frequent. Therefore, it is urgently necessary to rebuild a satellite platform configuration that is light in weight, small in envelope, adaptable to rapid attitude maneuvers on orbit, and highly modular, to meet the new needs of light and small remote sensing satellite platforms.

[0004] The existing Chinese patent with publication number CN111703592B discloses a large-scale 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 tank plate, a partition, a camera plate, an outer plate and a truss rod assembly. The tank plate and the partition are arranged on the docking plate. Multiple tank plates and partitions are respectively provided, and are arranged horizontally and vertically to form a tic-tac-toe structure. The camera plate is arranged above the tic-tac-toe structure and supported by a number of truss rod assemblies. A total of four outer plates are provided, which are arranged around the tic-tac-toe structure. Two adjacent 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. The single-machine equipment of each subsystem is installed on the tank plate and the partition; the propulsion subsystem assembly is installed on the docking plate, and the two tanks are installed symmetrically relative to the center of mass of the platform.

[0005] At present, the traditional remote sensing satellites with planar phased array antennas are all large platforms, and the commercial satellite platforms with light and small platforms are not adapted to the configuration requirements of planar phased array antennas. The weight of Remote Sensing Satellite No. 35 is less than 500 kilograms, and it can no longer directly inherit the existing mature technologies on traditional remote sensing satellites. It is necessary to innovate the configuration technology of new satellite platforms and break through the ideas based on its working characteristics of light weight, envelope limitation, frequent and rapid attitude maneuvers, and modular configuration. Therefore, the inventor believes that it is necessary to provide a lightweight and agile remote sensing satellite platform configuration to solve the above problems. Summary of the invention

[0006] In view of the defects in the prior art, an object of the present invention is to provide a lightweight agile remote sensing satellite platform configuration.

[0007] A lightweight agile remote sensing satellite platform configuration provided by the present invention includes: a satellite 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 includes an integrated carbon fiber truss and a honeycomb panel, wherein the integrated carbon fiber truss is a rectangular parallelepiped structure as a mounting base, the honeycomb panel is arranged inside the integrated carbon fiber truss, and the control moment gyro group is a pentagonal pyramid configuration and is fastened to the honeycomb panel;

[0009] The solar cell array and the radar antenna are installed on both sides of the integrated carbon fiber truss relative to each other, the radar antenna includes a plurality of movably connected sub-arrays, the radar antenna can switch between an expanded state and a folded state, and the solar cell array includes a plurality of movably connected sub-boards, the solar cell array can switch between an expanded state and a folded state;

[0010] The hoisting structure is arranged on the top of the integrated carbon fiber truss, the electric propulsion system is installed on the star body in a modularized manner, and the heat dissipation device is arranged on the outer side of the integrated carbon fiber truss.

[0011] Preferably, the honeycomb panel is used as a single-unit installation base on a satellite and adopts an I-shaped configuration. The honeycomb panel can be detachably installed inside the integrated carbon fiber truss, and the waist plate of the honeycomb panel is vertically arranged and parallel to the solar cell array.

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

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

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

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

[0016] When the solar cell array is in an unfolded state, a plurality of sub-panels form a complete solar cell array;

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

[0018] Preferably, the control moment gyro group includes five control moment gyros, which are installed at the center of the honeycomb panel through a bracket, and the rotation axes of the five control moment gyros intersect at a common point. A line connecting the center points of the five control moment gyros and the intersection points of the rotation axes forms a spatial pentagonal pyramid shape, and the cone apex angle is an acute angle.

[0019] Preferably, the lifting structure comprises a lifting 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 a carbon fiber plate, and the electric propulsion system is fastened to the star through the carbon fiber plate.

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

[0022] Preferably, an integrated star-sensing bracket is installed on the top of the integrated carbon fiber truss, and a plurality of mounting surfaces for mounting star sensors are provided on the integrated star-sensing bracket.

[0023] Preferably, two outer side surfaces of the integrated carbon fiber truss parallel to the radar antenna array are covered with thermal control multi-layers, and the other two outer side surfaces are provided with aluminum alloy heat dissipation plates.

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

[0025] 1. The present invention adopts integrated carbon fiber truss structure configuration technology, radar antenna splicing combination planar array configuration technology, body-mounted solar cell array configuration technology, control moment gyro group pentagonal pyramid configuration technology, structure top lifting technology, modular electric propulsion configuration technology and efficient heat dissipation configuration technology, which can solve the new demands brought about by the satellite's light weight, envelope limitation, and frequent and rapid attitude maneuvers, and achieves the good effects of light weight, small envelope, and rapid attitude maneuverability.

[0026] 2. The present invention adopts an integrated carbon fiber truss and honeycomb panel combination. The carbon fiber material is lighter in weight than the traditional honeycomb panel material, and has better bearing capacity and stiffness strength. An integrated bracket is designed on the top of the integrated carbon fiber truss to mount multiple star sensors, and an aluminum alloy metal joint is designed on the bottom to connect 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 and attached to the two outer sides of the integrated carbon fiber truss perpendicular to the radar antenna array, which can significantly provide heat dissipation capacity.

[0027] 3. The present invention uses the radar antenna splicing and combination plane configuration technology to divide the planar phased array radar antenna that exceeds the carrier envelope size into several small sub-arrays along one direction. The size of each sub-array cannot exceed the width of the integrated carbon fiber truss, and the sub-array is vertically folded and installed on the three outer sides of the integrated carbon fiber truss through a torsion spring hinge mechanism; the solar cell array that exceeds the carrier envelope size is divided into several sub-boards through the body-mounted solar cell array configuration technology, and the middle board is vertically fixed on an outer side of the integrated carbon fiber truss 11 structure, and the remaining sub-boards are folded and folded in succession and installed on the middle board through a torsion spring hinge mechanism; the satellite envelope restriction constraint is solved, and the satellite installation requirements of large-size planar phased array radar antennas and large-size solar cell arrays are met. The weight moment of inertia is small, and the driving mechanism and connecting rod of the traditional solar cell array are eliminated, thereby reducing the weight of the entire satellite.

[0028] 4. The present invention adopts the pentagonal pyramid configuration technology of the control moment gyro group to install five large torque control moment gyros on an I-shaped honeycomb plate through a bracket. The rotation axes of the five control moment gyros intersect at a common point. The center points of the five control moment gyros and the intersection points of the rotation axes are connected to form a spatial pentagonal pyramid shape with a fixed cone apex angle. The five control moment gyros are compactly installed, the on-board installation space envelope is small, and the requirements for rapid attitude maneuvering of the satellite in orbit can be met.

[0029] 5. The present invention designs the satellite lifting points at the four corner end rod joints at the top of the integrated carbon fiber truss structure, pre-embeds high-strength aluminum alloy embedded parts in the rod joints, and connects them into a whole through tooling and lifting tools, so as to provide for the rapid lifting of satellites. During launch, the tooling connected to the lifting points is removed, which can greatly reduce the weight of the satellite.

[0030] 6. The present invention realizes modular integration of the electric propulsion system by welding all the electric propulsion system products onto the carbon fiber plate, and connects the carbon fiber plate with the satellite structure plate as a whole, so as to realize modularization of the electric propulsion system as a whole, and quickly install and dismantle the satellite. BRIEF DESCRIPTION OF THE DRAWINGS

[0031] Other features, objects and advantages of the present invention will become more apparent from the detailed description of non-limiting embodiments made with reference to the following drawings:

[0032] Figure 1 The present invention mainly embodies the modular configuration combination schematic diagram of the lightweight agile remote sensing satellite platform configuration;

[0033] Figure 2 It is a schematic diagram of the structure of the star body mainly embodied in the present invention;

[0034] Figure 3This is a schematic diagram of the present invention mainly showing that an I-shaped honeycomb panel moves up and down within a truss structure;

[0035] Figure 4 This is a schematic diagram of the folded configuration of the radar antenna mainly embodied in the present invention;

[0036] Figure 5 This is a schematic diagram of the deployment configuration of a radar antenna mainly embodied in the present invention;

[0037] Figure 6 This is a schematic diagram of the folded configuration of the solar cell array mainly embodied in the present invention;

[0038] Figure 7 This is a schematic diagram of the unfolded configuration of a solar cell array mainly embodied in the present invention;

[0039] Figure 8 This is a schematic diagram of the configuration of a control moment gyro group mainly embodied in the present invention;

[0040] Fig. 9 This is a schematic diagram of the installation of an integrated star sensor bracket mainly embodied in the present invention;

[0041] Fig.10 This is a schematic diagram of the structure of the integrated star sensor support frame mainly embodied in the present invention;

[0042] Fig.11 This is a schematic diagram of the locations of the suspension points of the integrated carbon fiber truss mainly embodied in the present invention;

[0043] Fig.12 This is a schematic diagram of the configuration of an aluminum alloy joint that mainly embodies the integrated carbon fiber truss and the carrier connection of the present invention;

[0044] Fig.13 This is a schematic diagram of the modular configuration of electric propulsion that is mainly embodied in the present invention;

[0045] Fig.14 This is a schematic diagram of the structure of the aluminum alloy sheet mainly embodied in the present invention.

[0046] As shown in the figure:

[0047] Star 1 Radar antenna 2 Solar array 3

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

[0049] Honeycomb panel 12 DETAILED DESCRIPTION

[0050] The present invention is described in detail below in conjunction with specific embodiments. The following embodiments will help those skilled in the art to further understand the present invention, but are not intended to limit the present invention in any form. It should be noted that, for those of ordinary skill in the art, several changes and improvements can also be made without departing from the concept of the present invention. These all belong to the protection scope of the present invention.

[0051] like Figure 1-14 As shown, a light 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 gyro 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 is a rectangular parallelepiped structure as an installation base, the honeycomb panel 12 is arranged inside the integrated carbon fiber truss 11, and the control moment gyro group 4 is a pentagonal pyramid configuration and is fastened to the honeycomb panel 12; the solar cell array 3 is installed on both sides of the integrated carbon fiber truss 11 relative to the radar antenna 2, the radar antenna 2 includes a plurality of movably connected sub-arrays, the radar antenna 2 can switch between an expanded state and a retracted state, the solar cell array 3 includes a plurality of movably connected sub-panels, the solar cell array 3 can switch between an expanded state and a folded state; a hoisting structure 5 is arranged on the top of the integrated carbon fiber truss 11, the electric propulsion system is modularly installed on the satellite body 1, and a heat dissipation device is arranged on the outer side of the integrated carbon fiber truss 11.

[0052] The structure of the satellite body 1 is designed to be composed of an integrated carbon fiber truss 11 and a honeycomb panel 12, wherein the integrated carbon fiber truss 11 is in the form of a rectangular column, which is the main load-bearing structure, serving as the installation basis for the main load and solar array on the satellite, and mainly bearing the mechanical environment of the active phase of the launch vehicle. The honeycomb panel 12 is used as the installation basis for a single unit on the satellite, and adopts an I-shaped configuration design. The waist plate of the honeycomb panel 12 is vertically arranged and parallel to the solar array 3. The honeycomb panel 12 can be detachably installed inside the integrated carbon fiber truss 11, and the honeycomb panel 12 can be flexibly removed from the integrated carbon fiber truss 11 structure, generally moving from bottom to top in the integrated carbon fiber truss 11.

[0053] The radar antenna 2 is designed as a planar phased array configuration, by dividing the radar antenna 2 into a number of sub-arrays in the distance direction, and vertically folding them from three directions and installing them on the integrated carbon fiber truss 11. Before the radar antenna 2 works, all the sub-arrays are unfolded by a torsion spring hinge mechanism to form a complete planar phased array antenna. Preferably, the radar antenna 2 includes three vertically arranged sub-arrays, two adjacent sub-arrays are connected by a torsion spring hinge mechanism, and the sub-array in the middle is installed on the integrated carbon fiber truss 11; when the radar antenna 2 is in the unfolded state, the three sub-arrays are combined into a complete planar phased array antenna; when the radar antenna 2 is in the folded state, the three sub-arrays cover the three sides of the integrated carbon fiber truss 11.

[0054] The solar cell array 3 is a body-mounted configuration, and includes a plurality of sub-boards, two adjacent sub-boards are connected by a spring hinge mechanism, the middle board of the solar cell array 3 is fastened to the integrated carbon fiber truss 11, and the remaining sub-boards are folded and pressed against the middle board in sequence. When the satellite is launched into orbit, the remaining sub-boards of the solar cell array 3 are all unfolded in sequence through the spring hinge mechanism to form a complete solar cell array. Preferably, the solar cell array 3 includes three vertically arranged sub-boards, and when the solar cell array 3 is in an unfolded state, the three sub-boards form a complete solar cell array; when the solar cell array 3 is in a folded state, the sub-boards except the middle board are all folded and arranged on the middle board.

[0055] The control moment gyro group 4 is designed as a pentagonal pyramid configuration. The control moment gyro group 4 includes five control moment gyros. The control moment gyro is installed at the center of the honeycomb panel 12 through a bracket. The rotation axes of the five control moment gyros intersect at a common point. The connection between the center points of the five control moment gyros and the intersection points of the rotation axes is a spatial pentagonal pyramid shape, and the cone top angle is a fixed acute angle.

[0056] The hoisting structure 5 includes metal embedded parts of the lifting points arranged in the top rod joints of the integrated carbon fiber truss 11, that is, the four end joint areas located at the top of the rectangular column, which are designed as an integrated whole with the truss structure. High-strength metal embedded parts are pre-embedded in the rod joints in each end area. During the lifting process, hoists are installed at the four joints to lift the satellite vertically.

[0057] The propellant of the electric propulsion system is stored in gas cylinders. The electric propulsion system uses a 3mm carbon fiber plate to integrate the propulsion products together. As a modular single-machine configuration, it can independently complete the integrated welding, which is convenient for the parallel development and installation of the system. A high-pressure gas cylinder is set on the carbon fiber plate, and the high-pressure gas cylinder is connected to a pressure reducing valve and multiple valves. After completing the pipeline welding integration and propellant filling, the electric propulsion system is connected and fixed to the satellite 1 as a modular whole through the carbon fiber plate.

[0058] The four joints at the bottom of the integrated carbon fiber truss 11 are respectively provided with aluminum alloy joints for mechanical connection with the launch vehicle. The aluminum alloy metal structure in the joints at the four ends of the bottom of the integrated carbon fiber truss 11 is designed as an integrated whole with the truss structure to achieve mechanical connection with the launch vehicle.

[0059] An integrated star-sensing bracket is installed on the top side of the integrated carbon fiber truss 11, and its bottom surface is connected to the integrated carbon fiber truss 11 as a common installation surface. The integrated star-sensing bracket is provided with multiple installation surfaces for installing star sensors. The top of the integrated star-sensing bracket is provided with multiple installation surfaces, which are respectively connected to different star sensors. Preferably, three star sensors are installed on the top of the integrated star-sensing bracket. Multiple star sensors are installed on one integrated bracket, which is beneficial to both reducing the weight of the bracket and unifying the installation accuracy benchmarks of multiple star sensors.

[0060] The integrated carbon fiber truss 11 is a rectangular parallelepiped structure, and there is no honeycomb panel on the outside of the six directions. For the four outer sides of the integrated carbon fiber truss 11, the two outer sides of the integrated carbon fiber truss 11 parallel to the array of the radar antenna 2 are covered with thermal control multilayers, and the other two outer sides are provided with aluminum alloy heat sinks. In the direction of the outer side of the two integrated carbon fiber trusses 11 perpendicular to the array of the radar antenna 2, each is equipped with a complete 0.5 mm thick aluminum alloy metal sheet, which is directly screwed to the outer side of the integrated carbon fiber truss 11 to replace the traditional 25 mm thick honeycomb panel structure, which greatly reduces the weight. The thin plate is sprayed with paint to facilitate heat dissipation. On the outer side of the integrated carbon fiber truss 11 parallel to the array direction of the radar antenna 2, all are covered with thermal control multilayers, and buckles are attached to the rods of the integrated carbon fiber truss 11 for multi-layer pasting.

[0061] This application mainly adopts 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 pentagonal pyramid configuration technology, structure top lifting technology, modular electric propulsion configuration technology and efficient heat dissipation configuration technology, which can solve the new demands brought about by the satellite's light weight, envelope limitations, and frequent and rapid attitude maneuvers.

[0062] In response to the constraint of light weight of satellites, the present application is designed as a combination of an integrated carbon fiber truss 11 and a honeycomb panel 12. Carbon fiber materials are lighter than traditional honeycomb panel materials and have better bearing capacity and stiffness strength. The integrated carbon fiber truss 11 structural configuration technology is adopted, and the main structure of the entire satellite is designed from a traditional honeycomb panel structure to an integrated carbon fiber truss as the main body of the satellite structure's outer frame. All external side panels of the satellite structure are cancelled, and only an I-shaped honeycomb panel 12 combination is retained and built into the truss structure for installing a single unit on the satellite, and it can be flexibly moved out from the inside of the integrated carbon fiber truss 11 structure.

[0063] At the same time, an integrated bracket is designed on the top of the integrated carbon fiber truss 11, which can be equipped with three star sensors; an aluminum alloy metal joint is designed at the bottom 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 and attached to the two outer sides of the integrated carbon fiber truss 11 perpendicular to the array surface of the radar antenna 2, which can provide significant heat dissipation capabilities.

[0064] In view of the problem of strict satellite envelope restrictions, the radar antenna 2 splicing and combination plane configuration technology is used to divide the plane phased array radar antenna that exceeds the carrier envelope size into several small sub-arrays along one direction. The size of each sub-array cannot exceed the width of the integrated carbon fiber truss 11, and the sub-array is vertically folded and installed on the three outer sides of the integrated carbon fiber truss 11 through a torsion spring hinge mechanism. Through the body-mounted solar cell array 3 configuration technology, the solar cell array 3 that exceeds the carrier envelope size is divided into several sub-boards, and the middle board is vertically fixed on one outer side of the integrated carbon fiber truss 11 structure, and the remaining sub-boards are folded and folded successively and installed on the middle board through a torsion spring hinge mechanism. The above configuration technology solves the satellite envelope restriction constraints, meets the satellite installation requirements of large-sized plane phased array radar antennas and large-sized solar cell arrays, has a small weight moment of inertia, and eliminates the driving mechanism and connecting rods of the traditional solar cell array, reducing the weight of the entire satellite.

[0065] In order to meet the requirements of rapid and agile attitude maneuvering of remote sensing satellites in orbit, five high-torque control moment gyroscopes are installed on an I-shaped honeycomb plate 12 through brackets through the pentagonal pyramid configuration technology of the control moment gyroscope group 4. The rotation axes of the five control moment gyroscopes intersect at a common point. The lines connecting the center points of the five control moment gyroscopes and the intersection points of the rotation axes form a spatial pentagonal pyramid shape with a fixed cone apex angle. The five control moment gyroscopes are compactly installed, the on-board installation space envelope is small, and the requirements of rapid attitude maneuvering of satellites in orbit can be met.

[0066] In response to the need for rapid satellite lifting, the satellite lifting points are designed at the four corner end rod joints at the top of the integrated carbon fiber truss 11 structure. High-strength aluminum alloy embedded parts are pre-embedded in the rod joints and connected as a whole through tooling lifting fixtures for use in rapid satellite lifting.

[0067] In response to the configuration requirements of the electric propulsion system, a 3 mm thick carbon fiber plate was designed, and all the electric propulsion system products were welded and integrated on this carbon fiber plate, realizing the modular integration of the electric propulsion system, and the whole system is connected through the carbon fiber plate and the star structure plate.

[0068] This application aims to meet the urgent needs of new lightweight agile remote sensing satellites with light weight, small envelope and rapid in-orbit attitude maneuverability. Through the 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 pentagonal configuration technology and electric propulsion system modular configuration technology, it solves the problems that are difficult to solve for traditional large-platform remote sensing satellites and creates a new satellite platform configuration.

[0069] This application solves the difficult problem of configuring multiple types of lightweight and agile remote sensing satellite platforms, achieves the good effects of light weight, small envelope, and rapid attitude maneuverability, and fills the gap in the configuration design of planar phased array antenna remote sensing satellites on light and small satellite platforms.

[0070] This application is specifically described by the following data:

[0071] As a high-resolution remote sensing satellite with a weight of no more than 500 kilograms and equipped with a planar phased array antenna, the Remote Sensing Satellite No. 35 has clear constraints on the weight and envelope size of the entire satellite structure, radar antenna, solar array, and control system actuators. The Remote Sensing Satellite No. 35 requires that the structural weight accounts for no more than 10% of the entire satellite, the azimuth dimension of the radar antenna is 3390 mm, the range dimension is also 3390 mm, the solar array area is 7 square meters, and the on-orbit working mode switching requires that it can meet the attitude maneuvers in the front, back, left, and right directions, and the time cannot exceed 25 seconds. The carrier rocket that launches the satellite requires the satellite envelope to be no larger than 2200 mm.

[0072] In the present application, the main load-bearing structure of the entire satellite is designed to be an open integrated carbon fiber truss 11 structural configuration. The wall thickness of the truss rods is designed to be 1.2 mm, and all side panels outside the truss structure are cancelled. At the same time, in order to ensure the installation of a single machine, an I-shaped honeycomb panel 12 combination configuration is designed, which is built into the integrated carbon fiber truss 11 structure and can be flexibly moved out of the truss structure. Under the premise of ensuring the structural stiffness, bearing capacity and single-machine installation requirements of the entire satellite, the weight of the entire satellite structure accounts for only 9.3% of the entire satellite.

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

[0074] The body-mounted solar cell array 3 configuration technology of the present application divides the 7 square solar cell array 3 into three rectangular sub-boards, namely the left sub-board, the neutron board and the right sub-board, wherein the neutron board of the solar cell array 3 is directly fixed on the integrated carbon fiber truss 11 of the satellite body, the left sub-board is connected to the left side of the neutron board by a torsion spring hinge mechanism, the right sub-board is connected to the right side of the neutron board by a torsion spring hinge mechanism, the left sub-board and the right sub-board are folded and pressed on the neutron board in sequence, and the solar cell array configuration envelope under this configuration is obtained to be 2100 mm, which meets the launch envelope range requirements of the launch vehicle. This solar cell array configuration technology eliminates the traditional solar cell array drive mechanism and connecting rod, directly fixes the solar cell array sub-board to the satellite body, and greatly reduces the weight of the solar cell array.

[0075] The control moment gyro group 4 pentagonal pyramid configuration technology of the present application makes full use of the large torque performance of the control moment gyro, greatly reduces the use of flywheels, and installs five control moment gyros on the I-shaped honeycomb plate 12 through brackets, and the rotation axes of the five control moment gyros intersect at a common point. The center points of the five control moment gyros and the intersection points of the rotation axes are connected to form a spatial pentagonal pyramid configuration, and the cone vertex angle is designed to be 45°. The control moment gyro under this pentagonal pyramid configuration has a small demand for the space envelope of the star layout, which can not only solve the problem of tight space for small satellite layout, but also meet the star attitude maneuvers in different directions of the front, back, left, and right of the star attitude. It can be seen from the on-orbit measurement results that its attitude rapid maneuvering time does not exceed 25 seconds.

[0076] The top rapid lifting technology of the present application utilizes the high-strength characteristics of the integrated carbon fiber truss 11 structure, directly designs aluminum alloy embedded parts into the integrated carbon fiber truss 11 rod joints, and arranges them in the four corner end areas facing the sky at the top of the structure, and is connected as a whole through ground tooling. Normally, the tooling is kept connected to the lifting point, and the top can meet the needs of rapid lifting and placement of other deployment tooling. When launching, the tooling connected to the lifting point will be removed, which can greatly reduce the weight of the satellite.

[0077] The modular electric propulsion system technology of this application is designed with a 3 mm thick carbon fiber plate, which is hollowed out appropriately to reduce weight according to the installation interface on the satellite. The electric propulsion system uses gas cylinders to store propellant, which are fixedly connected to the carbon fiber plate. The remaining products are all layered and integrated around the gas cylinders and welded on the carbon fiber plate. Finally, the electric propulsion system can be modularized into a whole, which can be quickly installed and removed from the satellite.

[0078] This application innovatively solves the urgent needs of the weight, envelope, agile maneuverability and modular assembly of the Remote Sensing Satellite No. 35. It creates a new design method for the subsequent remote sensing satellite platform configuration technology with planar phased array antennas, especially for the lightweight and agile remote sensing satellite platform configuration, which will be widely used.

[0079] In the description of the present application, it should be understood that the terms "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inside", "outside", etc., indicating orientations or positional relationships, are based on the orientations or positional relationships shown in the accompanying drawings, and are only for the convenience of describing the present application and simplifying the description, and do not indicate or imply that the referred device or element must have a specific orientation, be constructed and operated in a specific orientation, and therefore should not be understood as a limitation on the present application.

[0080] The above describes the specific embodiments of the present invention. It should be understood that the present invention is not limited to the above specific embodiments, and those skilled in the art can make various changes or modifications within the scope of the claims, which does not affect the essence of the present invention. In the absence of conflict, the embodiments of the present application and the features in the embodiments can be combined with each other arbitrarily.

Claims

1. A lightweight and agile remote sensing satellite platform configuration, It is characterized in that include: A satellite body (1), a radar antenna (2), a solar array (3), a control moment gyro group (4), a lifting structure (5), an electric propulsion system, and a heat dissipation device; The star body (1) comprises an integrated carbon fiber truss (11) and a honeycomb panel (12); the integrated carbon fiber truss (11) is a rectangular parallelepiped structure as a mounting base; the honeycomb panel (12) is arranged inside the integrated carbon fiber truss (11); and the control moment gyro group (4) is a pentagonal pyramid configuration and is fixedly mounted on the honeycomb panel (12); The solar cell array (3) and the radar antenna (2) are mounted on both sides of the integrated carbon fiber truss (11) relative to each other, the radar antenna (2) comprises a plurality of movably connected sub-arrays, the radar antenna (2) can switch between an unfolded state and a folded state, the solar cell array (3) comprises a plurality of movably connected sub-boards, the solar cell array (3) can switch between an unfolded state and a folded state; The top of the integrated carbon fiber truss (11) is provided with the hoisting structure (5), the electric propulsion system is integrally mounted on the star body (1) in a modular manner, and the outer side surface of the integrated carbon fiber truss (11) is provided with the heat dissipation device.

2. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that The honeycomb panel (12) serves as a single-unit installation base on a satellite and adopts an I-shaped configuration. The honeycomb panel (12) is detachably installed inside the integrated carbon fiber truss (11). The waist plate of the honeycomb panel (12) is vertically arranged and parallel to the solar cell array (3).

3. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that The radar antenna (2) is of a planar phased array configuration, and comprises a plurality of vertically arranged sub-arrays, two adjacent sub-arrays are connected via a torsion spring hinge mechanism, and the sub-array in the middle is mounted on the integrated carbon fiber truss (11); When the radar antenna (2) is in an unfolded state, a plurality of the sub-arrays are combined into a complete planar phased array antenna; When the radar antenna (2) is in a folded state, the plurality of sub-arrays cover three side surfaces of the integrated carbon fiber truss (11).

4. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that The solar cell array (3) is of a body-mounted configuration, the solar cell array (3) comprises a plurality of sub-panels, two adjacent sub-panels are connected via a spring hinge mechanism, and the middle plate of the solar cell array (3) is fixedly mounted on the integrated carbon fiber truss (11); When the solar cell array (3) is in an unfolded state, a plurality of sub-panels form a complete solar cell array; When the solar cell array (3) is in a folded state, all sub-panels except the middle panel are folded and arranged on the middle panel.

5. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that The control moment gyro group (4) comprises five control moment gyros, which are mounted at the center of the honeycomb panel (12) via a bracket, and the rotation axes of the five control moment gyros intersect at a common point. A line connecting the center points of the five control moment gyros and the intersection points of the rotation axes forms a spatial pentagonal pyramid, and the cone vertex angle is an acute angle.

6. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that The lifting structure (5) comprises a lifting point metal embedded part arranged in a top rod joint of the integrated carbon fiber truss (11).

7. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that The propellant of the electric propulsion system is stored in a gas cylinder. The electric propulsion system is arranged on a carbon fiber plate. The electric propulsion system is firmly connected to the satellite (1) via the carbon fiber plate.

8. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that Aluminum alloy joints for mechanical connection with the launch vehicle are respectively arranged at four joints at the bottom of the integrated carbon fiber truss (11).

9. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that An integrated star-sensing bracket is installed 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-sensing bracket.

10. The lightweight agile remote sensing satellite platform configuration as claimed in claim 1, It is characterized in that Two outer side surfaces of the integrated carbon fiber truss (11) parallel to the array plane of the radar antenna (2) are covered with a thermal control multilayer, and the other two outer side surfaces are provided with aluminum alloy heat dissipation plates.

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