High-compatibility remote sensing satellite commonality application support platform and arrangement method thereof

Abstract

The application provides a high-compatibility remote sensing satellite commonality application support platform and a layout method thereof, and belongs to the technical field of satellites. The problems that the satellite support platform in the prior art cannot be applied to different loads and different types of carrier rockets are solved. The camera mounting plate, the satellite-rocket separation mechanism support, the carried load mounting plate, two solar cell array assemblies and a flywheel support assembly are all mounted on an optical load light shield. Multiple carried loads are mounted on the carried load mounting plate. One end of a satellite-rocket separation mechanism is connected with the optical load light shield through two mass center adjusting structures, and the other end is connected with the satellite-rocket separation mechanism support through another two mass center adjusting structures. The mass center adjusting structure comprises a mass center adjusting moving block, a mass center adjusting support and an adjusting bolt. The mass center adjusting support is connected with the satellite-rocket separation mechanism, and the adjusting bolt connects the mass center adjusting moving block with the optical load light shield or the satellite-rocket separation mechanism support. The application has the characteristics of high compatibility of carried loads and carrier rockets.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of satellites, and particularly relates to a high-compatibility remote sensing satellite common application support platform and a layout method thereof. BACKGROUND

[0002] With the rapid development of the aerospace industry, a large number of new loads need to be verified in the real space environment, such as verifying the anti-radiation capability and long-term operation stability of the carried load, and with the shortening of the update cycle of aerospace technology, the rapid iteration of the load also needs high-frequency on-orbit verification support.

[0003] Since most new loads have limited research and development funds, it is difficult to develop a satellite dedicated to a load for on-orbit verification of the load, so multiple loads that need to be verified are carried on the same satellite for on-orbit verification in a "carpooling" manner, which is also an economical and effective verification method. Under the driving of diversified application scenarios, the volume, mass, installation interface, attitude stability, and micro-vibration environment requirements of different loads are different. Using the traditional development method, each satellite can only meet the needs of a few carried loads, and the change of the carried mass after the change of the carried load will cause the center of mass to deviate, affecting the satellite-rocket separation attitude and on-orbit attitude stability. At the same time, the mechanical interfaces of different carried loads are also different, so it is necessary to redesign the satellite layout and structure for new carried loads, which leads to a series of problems such as high development cost and long cycle.

[0004] Moreover, the microsatellite for verifying the carried load is generally launched in a carried manner rather than a primary satellite, and needs to be connected to different types of launch vehicles to seek launch opportunities. The mechanical interfaces and vibration environments of the satellite-rocket separation mechanisms of different types of launch vehicles are different, and the mechanical interfaces of the satellite-rocket separation mechanisms and the structure of the satellite need to be frequently changed to adapt to the vibration environment of the launch, further increasing the complexity of satellite development.

[0005] Therefore, there is an urgent need for a satellite support platform suitable for different loads and capable of carrying different types of launch vehicles. SUMMARY

[0006] Therefore, in order to solve the problem that the satellite support platform in the prior art cannot be applied to different loads and cannot carry different types of launch vehicles, the application provides a high-compatibility remote sensing satellite common application support platform and a layout method thereof.

[0007] To achieve the above-mentioned purpose, the application adopts the following technical solutions:

[0008] The application discloses a high-compatibility remote sensing satellite commonality application support platform, which comprises a satellite platform, a plurality of carried loads, a satellite-rocket separation mechanism and four mass center adjusting structures, wherein the satellite platform comprises an optical load sunshade, a platform substructure, two solar cell array assemblies, a flywheel support assembly, an optical camera, a plurality of platform single machines and a propulsion assembly; the plurality of platform single machines are installed on the optical load sunshade; the platform substructure comprises a camera mounting plate, a carried load mounting plate and a satellite-rocket separation mechanism support; the optical camera is installed on the camera mounting plate; the camera mounting plate is detachably installed on the X-axis negative direction side of the optical load sunshade; the carried load mounting plate is detachably installed on the Z-axis negative direction side of the optical load sunshade; the satellite-rocket separation mechanism support is detachably installed on the Z-axis positive direction side of the optical load sunshade; the two solar cell array assemblies are respectively installed on the Y-axis positive direction side and the Y-axis negative direction side of the optical load sunshade; the flywheel support assembly is installed on the Y-axis positive direction side of the optical load sunshade; and the propulsion assembly is installed on the Z-axis positive direction side of the optical load sunshade and embedded in the satellite-rocket separation mechanism.

[0009] One end of the satellite-rocket separation mechanism is connected with the optical load sunshade through the two mass center adjusting structures, and the other end is connected with the satellite-rocket separation mechanism support through the other two mass center adjusting structures.

[0010] The mass center adjusting structure comprises a mass center adjusting moving block, a mass center adjusting support and an adjusting bolt; the mass center adjusting support is connected with the satellite-rocket separation mechanism; the mass center adjusting support is provided with a regular polygonal pyramid-shaped hole; the mass center adjusting moving block is in a regular polygonal pyramid shape; the mass center adjusting moving block is arranged in the regular polygonal pyramid-shaped hole; the mass center adjusting moving block is provided with a waist-shaped through hole; the adjusting bolt is arranged in the waist-shaped through hole; the adjusting bolts in the two mass center adjusting structures connect the mass center adjusting moving block with the optical load sunshade; and the adjusting bolts in the other two mass center adjusting structures connect the mass center adjusting moving block with the satellite-rocket separation mechanism support.

[0011] As an optimal scheme of the high-compatibility remote sensing satellite commonality application support platform, the platform single machines with a height greater than a set height are installed on the Y-axis positive direction side of the optical load sunshade; a part of the platform single machines with a height less than the set height are installed on the Y-axis negative direction side of the optical load sunshade; and the other part of the platform single machines with a height less than the set height are installed on the camera mounting plate.

[0012] As an optimal scheme of the high-compatibility remote sensing satellite commonality application support platform, the mass center adjusting support is provided with a regular hexagonal pyramid-shaped hole, and the mass center adjusting moving block is in a regular hexagonal pyramid shape.

[0013] As a preferred scheme of the high-compatibility remote sensing satellite commonality application support platform, the satellite-rocket separation mechanism comprises a satellite-rocket separation mechanism frame, a satellite-rocket separation mechanism residual part and a satellite-rocket separation mechanism damper, the satellite-rocket separation mechanism frame is fixedly connected with the satellite-rocket separation mechanism residual part, the satellite-rocket separation mechanism residual part is connected with the mass center adjustment support, and the satellite-rocket separation mechanism frame is connected with the carrier rocket through the satellite-rocket separation mechanism damper.

[0014] As a preferred scheme of the high-compatibility remote sensing satellite commonality application support platform, the satellite-rocket separation mechanism damper is of a T-shaped structure and is made of butyl rubber.

[0015] As a preferred scheme of the high-compatibility remote sensing satellite commonality application support platform, the optical load sunshade comprises an optical load sunshade cylinder, a plurality of light blocking rings and a plurality of embedded parts, the plurality of light blocking rings are arranged in the interior of the optical load sunshade cylinder in a spaced manner along the extension direction of the optical load sunshade cylinder, the plurality of embedded parts are embedded in the interior of the optical load sunshade cylinder in advance, and the plurality of embedded parts are used for providing mounting interfaces for a camera mounting plate, a carried load mounting plate, a satellite-rocket separation mechanism support, a platform single machine and a mass center adjustment structure.

[0016] As a preferred scheme of the high-compatibility remote sensing satellite commonality application support platform, the optical load sunshade cylinder is made of M40J carbon fiber / epoxy resin prepreg.

[0017] As a preferred scheme of the high-compatibility remote sensing satellite commonality application support platform, the embedded part is made of an aluminum alloy material.

[0018] As a preferred scheme of the high-compatibility remote sensing satellite commonality application support platform, the flywheel support assembly comprises a flywheel support, a plurality of flywheels and a plurality of polyurethane dampers, the plurality of flywheels are mounted on the flywheel support, the flywheel support is connected with the optical load sunshade through the plurality of polyurethane dampers, and the polyurethane dampers can adjust the rigidity according to different task modes.

[0019] The application further provides a layout method of the high-compatibility remote sensing satellite commonality application support platform, and adopts the high-compatibility remote sensing satellite commonality application support platform.

[0020] When the carried load is arranged, the mass center position of the carried load and the inertia in the mass center coordinate system are calculated respectively, the outer envelope of the carried load is projected onto the carried load mounting surface and simplified as a convex polygon, and a single-target layout optimization method is adopted for layout, and the layout optimization design model can be expressed as:

[0021]

[0022] Wherein, X The layout scheme of the satellite is expressed.N Total number of carried payloads x i , y i Position coordinates of the carried payloads i α Rotation angle of the centroid adjustment moving block, which is a discrete variable d Moving distance of the adjustment screw along the waist-shaped through hole of the centroid adjustment moving block d max Length of the waist-shaped through hole of the centroid adjustment moving block f X Target function corresponding to the in-orbit disturbance force arm of the satellite x co , y co Centroid coordinates of the satellite in the in-orbit flight state x cd , y cd Windward face centroid coordinates of the satellite in the in-orbit state g 1 X Non-interference constraint between the carried payloads A ij Overlap area of the installation face projection envelope of the carried payloads i j g 2 X , g 3 X Centroid constraint of the satellite in the launch state x cl , y cl Centroid coordinates of the satellite in the launch state x s , y s Centroid coordinates of the satellite separation structure separation point δx s , δy s Satellite centroid and separation mechanism centroid deviation allowed by satellite and rocket separation g 4 X , g 5 X , and g 6 X In-orbit state rotational inertia constraint of the satellite I x , I y , I z ​​​​) represents the satellite inertia around three coordinate axes, δI x , δI y , δI z ) represents the maximum allowed satellite inertia around three coordinate axes.

[0023] Compared with the prior art, the high-compatibility remote sensing satellite common application support platform and the arrangement method thereof provided by the application have the beneficial effects that:

[0024] 1. The high-compatibility remote sensing satellite common application support platform and the arrangement method thereof provided by the application have the characteristics of high compatibility with the carried load and the launch vehicle. In the high-compatibility remote sensing satellite common application support platform, the satellite can meet the installation requirements of different carried loads by only changing the design of the carried load mounting plate, and can realize the optimal layout design of the carried load with high satellite-vehicle separation precision and high on-orbit attitude stability by designing the position of the carried load, the rotation angle of the center-of-mass adjusting block, and the position of the adjusting bolt in the waist-shaped through hole through the carried load layout optimization algorithm. The satellite-vehicle separation mechanism is rigidly connected with the launch vehicle through screws or connected with the launch vehicle through a rubber damper with a designed rigidity, so that the mechanical environment requirements of different launch vehicles can be adapted, and high launch vehicle compatibility is achieved.

[0025] 2. The high-compatibility remote sensing satellite common application support platform provided by the application has the characteristics of high satellite on-orbit working mode compatibility. In the high-compatibility remote sensing satellite common application support platform, the flywheel is installed on the flywheel support, and the flywheel support is connected with the optical load light shield through a polyurethane damper. The polyurethane damper can change the rigidity according to the requirements of the camera or the carried load on the micro-vibration environment and the requirements of the satellite platform maneuvering capability, so that the satellite on-orbit working mode compatibility is high. DETAILED DESCRIPTION

[0026] The accompanying drawings, which form a part of the present application, are used to provide further understanding of the present application, and the illustrative embodiments of the present application and their descriptions are used to explain the present application, and do not constitute improper limitations on the present application. In the drawings:

[0027] Figure 1 is a structural schematic view of the high-compatibility remote sensing satellite common application support platform provided by the specific embodiment of the present application in a solar cell array assembly retracted state along a first viewing angle;

[0028] Figure 2 is a structural schematic view of the high-compatibility remote sensing satellite common application support platform provided by the specific embodiment of the present application in a solar cell array assembly retracted state along a second viewing angle;

[0029] Figure 3is a structural schematic view of a high-compatibility remote sensing satellite commonality application support platform along a third visual angle in a solar cell array assembly unfolded state according to an embodiment of the present application;

[0030] Figure 4 is a structural schematic view of a high-compatibility remote sensing satellite commonality application support platform along a fourth visual angle in a solar cell array assembly unfolded state according to an embodiment of the present application;

[0031] Figure 5 is an assembly relationship schematic view of a high-compatibility remote sensing satellite commonality application support platform according to an embodiment of the present application;

[0032] Figure 6 is a structural schematic view of a mass center adjustment moving block of a high-compatibility remote sensing satellite commonality application support platform according to an embodiment of the present application;

[0033] Figure 7 is a structural schematic view of a mass center adjustment support of a high-compatibility remote sensing satellite commonality application support platform according to an embodiment of the present application;

[0034] Figure 8 is an assembly relationship schematic view of a mass center adjustment structure and a satellite platform of a high-compatibility remote sensing satellite commonality application support platform according to an embodiment of the present application;

[0035] Figure 9 is a cross-sectional view of a mass center adjustment structure and a satellite platform of a high-compatibility remote sensing satellite commonality application support platform after assembly according to an embodiment of the present application;

[0036] Figure 10 is a partial cross-sectional view of an optical load sunshade of a high-compatibility remote sensing satellite commonality application support platform according to an embodiment of the present application;

[0037] Figure 11 is a structural schematic view of a flywheel support assembly of a high-compatibility remote sensing satellite commonality application support platform according to an embodiment of the present application.

[0038] In the figure:

[0039] 1, satellite platform; 11, optical load sunshade; 111, optical load sunshade cylinder; 112, light blocking ring; 113, embedded part; 12, platform substructure; 121, camera mounting plate; 122, carried load mounting plate; 123, satellite-rocket separation mechanism support; 13, flywheel support assembly; 131, flywheel; 132, flywheel support; 133, polyurethane damper; 14, solar cell array assembly; 15, optical camera; 16, platform unit; 161, platform unit with a height greater than a set height; 162, platform unit with a height less than a set height; 17, propulsion assembly;

[0040] 2, center of mass adjusting structure; 21, center of mass adjusting block; 22, center of mass adjusting support;

[0041] 3, satellite-rocket separation mechanism; 31, satellite-rocket separation mechanism residual part; 32, satellite-rocket separation mechanism frame; 33, satellite-rocket separation mechanism damper;

[0042] 4, carried load. DETAILED DESCRIPTION

[0043] The technical solutions in the embodiments of the present application will be clearly and completely described below with reference to the drawings in the embodiments of the present application. It should be noted that the embodiments in the present application and the features in the embodiments can be combined with each other without conflict, and the described embodiments are only some of the embodiments of the present application, not all the embodiments.

[0044] In the description of the present application, unless otherwise explicitly specified and limited, the terms "connected", "connected", "fixed" should be understood broadly, for example, it can be fixedly connected, or it can be detachably connected, or it can be integrated; it can be mechanically connected, or it can be electrically connected; it can be directly connected, or it can be indirectly connected through an intermediate medium; it can be the internal communication of two elements or the interaction relationship between two elements. For ordinary skilled in the art, the specific meaning of the above terms in the present application can be understood according to the specific circumstances.

[0045] In the present application, unless otherwise explicitly specified and limited, the first feature "on" or "under" the second feature can include that the first and second features are in direct contact, or that the first and second features are not in direct contact but are in contact through another feature between them. Moreover, the first feature "on", "above" and "above" the second feature includes that the first feature is directly above and obliquely above the second feature, or only indicates that the horizontal height of the first feature is higher than that of the second feature. The first feature "below", "below" and "below" the second feature includes that the first feature is directly below and obliquely below the second feature, or only indicates that the horizontal height of the first feature is less than that of the second feature.

[0046] In the description of the present embodiment, the terms "up", "down", "right", and other orientation or position relationship are based on the orientation or position relationship shown in the drawings, and are only for the convenience of description and simplification of operation, and do not 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. In addition, the terms "first" and "second" are only used to distinguish in description and have no special meaning.

[0047] Reference Figures 1-11The application provides a high-compatibility remote sensing satellite common application support platform, which comprises a satellite platform 1, a plurality of carried loads 4, a satellite-rocket separation mechanism 3 and four mass center adjustment structures 2. The satellite platform 1 comprises an optical load light shield 11, a platform secondary structure 12, two solar cell array assemblies 14, a flywheel support assembly 13, an optical camera 15, a plurality of platform single machines 16 and a propulsion assembly 17. The plurality of platform single machines 16 are installed on the optical load light shield 11. The platform secondary structure 12 comprises a camera mounting plate 121, a carried load mounting plate 122 and a satellite-rocket separation mechanism support 123. The optical camera 15 is installed on the camera mounting plate 121. The camera mounting plate 121 is detachably installed on the X-axis negative direction side of the optical load light shield 11. The plurality of carried loads 4 are installed on the carried load mounting plate 122. The carried load mounting plate 122 is detachably installed on the Z-axis negative direction side of the optical load light shield 11. The satellite-rocket separation mechanism support 123 is detachably installed on the Z-axis positive direction side of the optical load light shield 11. The two solar cell array assemblies 14 are respectively installed on the Y-axis positive direction side and the Y-axis negative direction side of the optical load light shield 11. The flywheel support assembly 13 is installed on the Y-axis positive direction side of the optical load light shield 11. The propulsion assembly 17 is installed on the Z-axis positive direction side of the optical load light shield 11 and is embedded in the satellite-rocket separation mechanism 3. One end of the satellite-rocket separation mechanism 3 is connected with the optical load light shield 11 through the two mass center adjustment structures 2. The other end of the satellite-rocket separation mechanism 3 is connected with the satellite-rocket separation mechanism support 123 through the other two mass center adjustment structures 2. The mass center adjustment structure 2 comprises a mass center adjustment moving block 21, a mass center adjustment support 22 and an adjusting screw. The mass center adjustment support 22 is connected with the satellite-rocket separation mechanism 3. The mass center adjustment support 22 is provided with a positive polygonal pyramid hole. The mass center adjustment moving block 21 is in a positive polygonal pyramid shape. The mass center adjustment moving block 21 is arranged in the positive polygonal pyramid hole. The mass center adjustment moving block 21 is provided with a waist-shaped through hole. The adjusting screw is arranged in the waist-shaped through hole. The adjusting screw in the two mass center adjustment structures 2 connects the mass center adjustment moving block 21 with the optical load light shield 11. The adjusting screw in the other two mass center adjustment structures 2 connects the mass center adjustment moving block 21 with the satellite-rocket separation mechanism support 123.

[0048] The high-compatibility remote sensing satellite commonality application support platform is arranged around the optical load sunshade 11, and the optical load sunshade 11 also serves as the main bearing structure of the satellite platform 1; the load mounting plate 122, the camera mounting plate 121 and the satellite-rocket separation mechanism support 123 are connected to the optical load sunshade 11 through screws, the load mounting plate 122 is a metal structure plate, a plurality of loads 4 are connected to the load mounting plate 122 through screws, and the load mounting plate 122 can be redesigned according to different mechanical interfaces of the loads 4, that is, corresponding threaded holes are arranged according to the mounting hole positions of the loads 4, and when different loads 4 need to be carried, the load mounting plate 122 corresponding to the loads 4 can be replaced, without the need to redesign other structural members of the satellite platform 1.

[0049] Moreover, the center of mass adjustment structure 2 is arranged, and the position of the center of mass of the satellite in the separation plane (X-Y plane) can be adjusted through the center of mass adjustment structure 2, so that the position deviation of the center of mass of the satellite from the center of the separation point of the separation mechanism meets the design requirements. The center of mass adjustment support 22 is connected to the satellite-rocket separation mechanism 3 through screws, the center of mass adjustment support 22 is provided with a regular polygonal pyramid-shaped hole with the same number of edges and the same inclination angle as the center of mass adjustment moving block 21, the center of mass adjustment moving block 21 is arranged in the regular polygonal pyramid-shaped hole and can be rotated to adjust the position, the center of mass adjustment moving block 21 is in contact with and pressed against the inner wall of the regular polygonal pyramid-shaped hole, the center of mass adjustment moving block 21 is provided with a waist-shaped through hole, and an adjusting screw is arranged in the waist-shaped through hole and can slide in the waist-shaped through hole. After the position is determined, the adjusting screw is connected to the optical load sunshade 11 or the satellite-rocket separation mechanism support 123, the center of mass adjustment moving block 21 is connected to the optical load sunshade 11 or the satellite-rocket separation mechanism support 123 through the adjusting screw, so that the satellite-rocket separation mechanism 3 is connected to the optical load sunshade 11 and the satellite-rocket separation mechanism support 123, and the center of mass adjustment moving block 21 can be rotated and the adjusting screw can slide in the waist-shaped through hole, so that the installation position of the satellite-rocket separation mechanism 3 can be adjusted, thereby adjusting the center of mass, so that the position of the center of mass of the satellite in the satellite-rocket separation plane can be adjusted when the loads 4 change, to meet the satellite-rocket separation requirements.

[0050] The high-compatibility remote sensing satellite commonality application support platform has high compatibility with loads 4 and launch vehicles. In the high-compatibility remote sensing satellite commonality application support platform, the satellite can meet the installation requirements of different loads 4 by only redesigning the load mounting plate 122, and the position of the load 4 and the position of the adjusting screw in the waist-shaped through hole can be designed through a load layout optimization algorithm, so as to achieve the best layout design of the load 4 with high satellite-rocket separation precision and high on-orbit attitude stability. The satellite-rocket separation mechanism 3 is rigidly connected to the launch vehicle through screws or connected to the launch vehicle through a rubber damper with a designed rigidity, so as to adapt to the mechanical environment requirements of different launch vehicles and have high compatibility with launch vehicles.

[0051] Optionally, the flywheel support assembly 13 comprises a flywheel support 132, a plurality of flywheels 131 and a plurality of polyurethane dampers 133, the plurality of flywheels 131 are installed on the flywheel support 132, the flywheel support 132 is connected with the optical payload sunshade 11 through the plurality of polyurethane dampers 133, and the polyurethane dampers 133 can adjust the rigidity according to different mission modes.

[0052] In the embodiment, the number of flywheels 131 is 4, the number of polyurethane dampers 133 is 6, and the flywheel support 132 is connected with the optical payload sunshade 11 through the polyurethane dampers 133. The polyurethane dampers 133 can adjust the rigidity according to different mission modes, for example, when the satellite is imaging the earth or performing laser communication tasks, the rigidity of the polyurethane dampers 133 is adjusted to be low, so as to reduce the influence of flywheel 131 disturbance on the optical camera 15 or the carried load 4, and when the satellite performs a maneuvering task, the rigidity of the polyurethane dampers 133 is adjusted to be high, so as to reduce the influence of the rigidity of the flywheel 131 installation base on the stability of the satellite attitude system.

[0053] Optionally, the satellite-rocket separation mechanism 3 comprises a satellite-rocket separation mechanism frame 32, a satellite-rocket separation mechanism residual part 31 and a satellite-rocket separation mechanism damper 33, the satellite-rocket separation mechanism frame 32 is fixedly connected with the satellite-rocket separation mechanism residual part 31, the satellite-rocket separation mechanism residual part 31 is connected with the center-of-gravity adjusting support 22, and the satellite-rocket separation mechanism frame 32 is connected with the launch vehicle through the satellite-rocket separation mechanism damper 33.

[0054] The satellite-rocket separation mechanism frame 32 is connected with the launch vehicle through the satellite-rocket separation mechanism damper 33. The satellite-rocket separation mechanism damper 33 can design the connection rigidity between the satellite-rocket separation mechanism frame 32 and the launch vehicle, so that the satellite can meet the mechanical environment requirements of different launch vehicles on the satellite.

[0055] The top of the center-of-gravity adjusting support 22 is in contact with the satellite platform 1 and is pressed against the satellite platform 1 through the center-of-gravity adjusting moving block 21. The bottom of the center-of-gravity adjusting support 22 is connected with the satellite-rocket separation mechanism residual part 31 through four inner hexagonal screws, and the center-of-gravity adjusting support 22 and the satellite-rocket separation structure residual part are positioned through a 3mm diameter cylindrical pin.

[0056] Optionally, the platform single machine 161 with a height greater than a set height in the plurality of platform single machines 16 is installed on the Y-axis positive direction side of the optical payload sunshade 11, and a part of the platform single machine 162 with a height less than the set height is installed on the Y-axis negative direction side of the optical payload sunshade 11, and the other part is installed on the camera mounting plate 121.

[0057] The distance between the solar cell array assembly 14 on the positive direction side of the Y axis and the optical load sunshade 11 is large, the distance between the solar cell array assembly 14 on the negative direction side of the Y axis and the optical load sunshade 11 is small, the platform single machine 16 with a high height and the flywheel support assembly 13 are installed on the positive direction side of the Y axis of the optical load sunshade 11, part of the platform single machine 16 with a small height is installed on the negative direction side of the Y axis of the optical load sunshade 11, and the other part is installed on the camera mounting plate 121, and the propulsion assembly 17 is installed on the positive direction side of the Z axis of the optical load sunshade 11 and is embedded in the satellite-rocket separation mechanism frame 32. Such a layout can effectively shrink the satellite envelope and reduce the satellite moment of inertia.

[0058] The camera mounting plate 121 provides an installation interface for the optical camera 15 and part of the platform single machine 16. The camera mounting plate 121 is a honeycomb sandwich plate or a metal structure plate. The camera mounting plate 121 can be made of aluminum alloy. Threaded holes are arranged according to the installation position of the platform single machine 16, and a reinforcing rib is arranged through optimized design. The optical camera 15 is a commercial camera. The satellite can perform commercial remote sensing imaging while completing the verification of the carried load 4. At the same time, the optical camera 15 can provide image data for the carried load 4 that needs image data.

[0059] In the embodiment, the barycentric adjustment support 22 is provided with a regular hexagonal pyramid hole, and the barycentric adjustment moving block 21 is in the shape of a regular hexagonal pyramid. The more the number of edges, the more the angles that can be adjusted. The setting in the shape of a pyramid can make the axial force more uniform and the fixation more reliable.

[0060] Optionally, the optical load sunshade 11 includes an optical load sunshade cylinder 111, a plurality of light blocking rings 112, and a plurality of embedded parts 113. The plurality of light blocking rings 112 are arranged in the interior of the optical load sunshade cylinder 111 along the extension direction of the optical load sunshade cylinder 111. The plurality of embedded parts 113 are embedded in the interior of the optical load sunshade cylinder 111 in advance. The plurality of embedded parts 113 are used to provide installation interfaces for the camera mounting plate 121, the carried load mounting plate 122, the satellite-rocket separation mechanism support 123, the platform single machine 16, and the barycentric adjustment structure 2. The optical load sunshade 11 is made of carbon fiber structure and is provided with a plurality of circumferential light blocking rings 112 inside. The optical load sunshade 11 provides a non-glare environment for the optical camera 15. At the same time, through the optimized design of the light blocking rings 112, the strength and rigidity requirements of the optical load sunshade 11 as the main bearing structure are met. The optical load sunshade 11 provides installation interfaces for the camera mounting plate 121, the carried load mounting plate 122, the satellite-rocket separation mechanism support 123, the platform single machine 16, and the barycentric adjustment structure 2 through the embedded parts 113 made of metal materials. Optionally, the embedded parts 113 are made of aluminum alloy.

[0061] Optionally, the optical load baffle barrel 111 is made of M40J carbon fiber / epoxy resin prepreg. The wall thickness of the optical load baffle barrel 111 is 0.8mm, and the inside is provided with 9 levels of light blocking rings 112, the thickness of the third and fourth levels of light blocking rings 112 is 3mm, and the thickness of the remaining light blocking rings 112 is 2mm. Under the launch mechanical environment shown in Tables 1-2, the maximum strain of the carbon fiber is 1500με, the first-order fundamental frequency of the satellite is 22Hz, and the optical load baffle 11 as the main force bearing structure meets the strength and stiffness requirements of the structure.

[0062] Table 1: Overload coefficient (g) at the satellite centroid

[0063]

[0064] Table 2: Satellite low-frequency sinusoidal environment

[0065]

[0066] Optionally, the satellite-rocket separation mechanism damper 33 is a T-shaped structure, and the material is butyl rubber. The large-diameter part of the satellite-rocket separation mechanism damper 33 has a diameter of 35mm and a height of 15mm, and the small-diameter part has a diameter of 25mm and a height of 8mm. Under the input of Table 2, the sinusoidal response of each platform unit 16 of the satellite is less than 6g, and under the input of Table 3, the random response of each platform unit 16 of the satellite is less than 8grms, meeting the launch mechanical environment requirements. Table 3: Satellite random vibration environment

[0067]

[0068] The application also provides a layout method of a high-compatibility remote sensing satellite common application support platform, which adopts the high-compatibility remote sensing satellite common application support platform, and comprises:

[0069] When the carried load 4 is arranged, the inertia at the centroid position and in the centroid coordinate system of the carried load 4 is calculated respectively, the outer envelope of the carried load 4 is projected onto the installation surface of the carried load 4 and simplified as a convex polygon, and a single-target layout optimization method is adopted for layout. The layout optimization design model can be expressed as:

[0070]

[0071] Wherein, X The layout scheme of the satellite is expressed. N The total number of the carried load 4 is expressed. x i , y i The position coordinates of the carried load i are expressed. αThe rotation angle of the centroid adjustment moving block 21 is a discrete variable, which is an integer multiple of 60° in the embodiment. When the centroid adjustment moving block 21 is a circular truncated cone, the variable is a continuous variable; d The moving distance of the adjusting screw along the waist-shaped through hole of the centroid adjustment moving block 21 is adjusted; d max The length of the waist-shaped through hole of the centroid adjustment moving block 21 is 6 mm in the embodiment; f ( X ) represents the target function corresponding to the disturbance force arm of the satellite in orbit, ( x co , y co ) represents the centroid coordinates of the satellite in the in-orbit flight state; ( x cd , y cd ) represents the centroid coordinates of the satellite in the in-orbit windward surface; g 1( X ) represents the non-interference constraint between the carried loads 4; ΔA ij represents the carried load i and the carried load j overlap area of the projection envelope on the mounting surface; g 2( X ), g 3( X ) represents the centroid constraint of the satellite in the launch state; ( x cl , y cl ) is the centroid coordinates of the satellite in the launch state; ( x s , y s ) is the centroid coordinates of the satellite separation structure separation point; ( δx s , δy s ) is the allowed deviation of the satellite centroid and the separation mechanism centroid in the satellite-rocket separation, which is 3 mm in the embodiment; g 4( X ), g 5( X ) and g 6( X ) represents the rotational inertia constraint of the satellite in the in-orbit state; ( I x , I y , I z ) represents the moment of inertia of the satellite around three coordinate axes, ( δI x ,δI y , δI z ) represents the maximum value of the moment of inertia of the satellite around the three coordinate axes.

[0072] Based on the above design constraints and design objectives, the single-objective optimization algorithm is used to optimize and solve the layout model of the carried load, so as to obtain the optimal layout scheme of the carried load 4, the angle of the mass center adjusting moving block 21 rotation and the position of the adjusting bolt in the waist-shaped through hole.

[0073] Obviously, the above disclosed embodiments of the application are only used to help explain the application. The embodiments do not describe all the details and do not limit the application to the specific embodiments described. Many modifications and variations can be made in light of the content of the specification. The specification selects and specifically describes these embodiments in order to better explain the principles and practical applications of the application, so that those skilled in the art can well understand and utilize the application. It is not necessary and impossible to exhaust all the embodiments here.

Claims

1. A highly compatible remote sensing satellite common application support platform, characterized in that, include: The satellite platform (1), multiple payloads (4), a satellite-rocket separation mechanism (3), and four centroid adjustment structures (2) are included. The satellite platform (1) includes an optical payload shield (11), a platform substructure (12), two solar array components (14), a flywheel support assembly (13), an optical camera (15), multiple platform units (16), and a propulsion assembly (17). The multiple platform units (16) are mounted on the optical payload shield (11). The platform substructure (12) includes a camera mounting plate (121), a payload mounting plate (122), and a satellite-rocket separation mechanism support (123). The optical camera (15) is mounted on the camera mounting plate (121), which is detachably mounted on the satellite. On the negative X-axis side of the optical payload shield (11), multiple payloads (4) are mounted on the payload mounting plate (122). The payload mounting plate (122) is detachably mounted on the negative Z-axis side of the optical payload shield (11). The star-rocket separation mechanism bracket (123) is detachably mounted on the positive Z-axis side of the optical payload shield (11). Two solar cell array components (14) are respectively mounted on the positive Y-axis side and the negative Y-axis side of the optical payload shield (11). The flywheel bracket assembly (13) is mounted on the positive Y-axis side of the optical payload shield (11). The propulsion assembly (17) is mounted on the positive Z-axis side of the optical payload shield (11) and embedded in the star-rocket separation mechanism (3). One end of the star-rocket separation mechanism (3) is connected to the optical payload shield (11) through two centroid adjustment structures (2), and the other end is connected to the star-rocket separation mechanism support (123) through two other centroid adjustment structures (2). The center of mass adjustment structure (2) includes a center of mass adjustment block (21), a center of mass adjustment support (22), and an adjustment bolt. The center of mass adjustment support (22) is connected to the star-rocket separation mechanism (3). The center of mass adjustment support (22) has a regular polyhedral pyramidal hole. The center of mass adjustment block (21) is in the shape of a regular polyhedral pyramid. The center of mass adjustment block (21) passes through the regular polyhedral pyramidal hole. The center of mass adjustment block (21) has a waist-shaped through hole. The adjustment bolt passes through the waist-shaped through hole. The adjustment bolts in the two center of mass adjustment structures (2) connect the center of mass adjustment block (21) to the optical load shield (11). The adjustment bolts in the other two center of mass adjustment structures (2) connect the center of mass adjustment block (21) to the star-rocket separation mechanism support (123).

2. The high-compatibility remote sensing satellite common application support platform according to claim 1, characterized in that: Among the multiple platform units (16), the platform unit (161) with a height greater than the set height is installed on the positive Y-axis side of the optical load hood (11), and the platform unit (162) with a height less than the set height is partially installed on the negative Y-axis side of the optical load hood (11) and partially installed on the camera mounting plate (121).

3. The high-compatibility remote sensing satellite common application support platform according to claim 1, characterized in that: The center of mass adjustment support (22) is provided with a regular hexagonal pyramidal hole, and the center of mass adjustment moving block (21) is in the shape of a regular hexagonal pyramid.

4. The high-compatibility remote sensing satellite common application support platform according to claim 1, characterized in that: The star-rocket separation mechanism (3) includes a star-rocket separation mechanism frame (32), a star-rocket separation mechanism residual part (31), and a star-rocket separation mechanism vibration damper (33). The star-rocket separation mechanism frame (32) is fixedly connected to the star-rocket separation mechanism residual part (31), the star-rocket separation mechanism residual part (31) is connected to the center of mass adjustment support (22), and the star-rocket separation mechanism frame (32) is connected to the launch vehicle through the star-rocket separation mechanism vibration damper (33).

5. The high-compatibility remote sensing satellite common application support platform according to claim 4, characterized in that: The vibration damper (33) of the star-rocket separation mechanism is a T-shaped structure and is made of butyl rubber.

6. The high-compatibility remote sensing satellite common application support platform according to claim 1, characterized in that: The optical load shield (11) includes an optical load shield cylinder (111), multiple light-blocking rings (112) and multiple embedded parts (113). The multiple light-blocking rings (112) are spaced apart inside the optical load shield cylinder (111) along the extension direction of the optical load shield cylinder (111). The multiple embedded parts (113) are pre-embedded inside the optical load shield cylinder (111). The multiple embedded parts (113) are used to provide installation interfaces for the camera mounting plate (121), the load mounting plate (122), the star-rocket separation mechanism bracket (123), the platform unit (16), and the center of mass adjustment structure (2).

7. The high-compatibility remote sensing satellite common application support platform according to claim 6, characterized in that: The optical load shield cylinder (111) is made of M40J carbon fiber / epoxy resin prepreg.

8. The high-compatibility remote sensing satellite common application support platform according to claim 6, characterized in that: The embedded part (113) is made of aluminum alloy.

9. The high-compatibility remote sensing satellite common application support platform according to claim 1, characterized in that: The flywheel support assembly (13) includes a flywheel support (132), multiple flywheels (131) and multiple polyurethane dampers (133). The multiple flywheels (131) are mounted on the flywheel support (132). The flywheel support (132) is connected to the optical load shield (11) through the multiple polyurethane dampers (133). The polyurethane dampers (133) can adjust their stiffness according to different mission modes.

10. A method for deploying a highly compatible remote sensing satellite common application support platform, characterized in that: The high-compatibility remote sensing satellite common application support platform according to any one of claims 1-9 includes: When arranging the load (4), the inertia at the center of mass of the load (4) and in the coordinate system of the center of mass are calculated respectively. The outer envelope of the load (4) is projected onto the mounting surface of the load (4) and simplified into a convex polygon. The layout is carried out using a single-objective layout optimization method. The layout optimization design model can be expressed as: in, X Describe the satellite's layout scheme; N Describe the total number of payloads (4) carried; x i , y i ) indicates the payload being carried (4) i Position coordinates; α Let be the rotation angle of the center-of-mass adjusting block (21), which is a discrete variable; d To adjust the distance the bolt moves along the oblong through hole of the adjusting block at the center of gravity; d max The length of the waist-shaped through hole of the center-of-mass adjusting block (21); f ( X ) represents the objective function of the corresponding satellite's on-orbit disturbance arm, x co , y co () represents the coordinates of the satellite's centroid during its orbital flight; x cd , y cd This indicates the coordinates of the centroid of the satellite's windward side during orbit. g 1( X ) indicates that there is no interference constraint between the mounted loads (4); Δ A ij Indicates the payload carried (4) i With the payload (4) j The overlapping area of ​​the projected envelope on the mounting surface; g 2( X ), g 3( X ) represents the centroid constraint of the satellite launch state; x cl , y cl ( ) represents the centroid coordinates of the satellite at launch status; x s , y s ( ) represents the centroid coordinates of the separation point of the satellite separation structure; δx s , δy s This refers to the permissible deviation between the satellite's center of mass and the centroid of the separation mechanism during satellite-rocket separation. g 4( X ), g 5( X )and g 6( X ) indicates the satellite's rotational inertia constraint in its on-orbit state; I x , I y , I z ) represents the satellite's inertia around the three coordinate axes, ( δI x , δI y , δI z () indicates the maximum allowed moment of inertia of a satellite around the three coordinate axes.

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