Satellite fixed structure and installation system based on one rocket two satellites parallel launch method

Abstract

The present application provides a satellite fixing structure and its installation system based on the one-rocket-two-satellite parallel launch method, which is used to promote the realization of the satellite fixing structure for satellites in the one-rocket-two-satellite parallel launch method with a streamlined and simple solar wing design structure, thereby significantly reducing the required production costs.

Description

Technical Field

[0001] The present application relates to the field of satellites, and specifically to a satellite fixing structure and an installation system thereof based on a one-rocket two-satellite parallel launch method. Background Art

[0002] In recent years, with the rapid development of commercial aerospace at home and abroad, the market demand for 100-kilogram small communication satellites for low-orbit networking has increased sharply. At the same time, many types of small rockets have been developed to adapt to the market, and their fairing diameters are mostly between 1m and 1.4m.

[0003] In order to effectively improve the competitiveness of satellites in the market, small communication satellites need to carry antennas as large as possible within a limited space. Large antennas often mean that more energy needs to be supplied. At the same time, in order to provide a better experience for more users on the ground, these antennas mounted on satellites often need to be powered on for a long time, which further consumes the energy of the entire satellite. Combining the above two characteristics, it can be seen that small satellites need a larger solar wing area. In a limited space, there are generally two ways to increase the solar wing area. One is to increase the number of folds of each wing, such as changing a single fold to a double fold, or a double fold to a triple fold. The second is to add another wing, such as changing a double wing to a triple wing or a quadruple wing. This requires designers to make a trade-off between cost and efficiency.

[0004] During the research process of existing related technologies, the inventors found that in the design scheme of increasing the area of ​​the solar wing, such as the double-wing multi-fold structure scheme, trusses and special flip vehicles are required when the product is installed, which puts high demands on the site, equipment and personnel. The corresponding product installation cycle will also increase. In other words, it has a higher production cost. Summary of the invention

[0005] The present application provides a satellite fixing structure and its installation system based on the one-rocket-two-satellite parallel launch method, which is used to promote the realization of the satellite fixing structure for satellites in the one-rocket-two-satellite parallel launch method with a streamlined and simple solar wing design structure, thereby significantly reducing the required production costs.

[0006] In the first aspect, the present application provides a satellite fixing structure based on a one-rocket two-satellite parallel launch method, in which two identical satellites are placed back to back in a fairing 01, and each satellite is equipped with a solar wing with a three-wing single-fold structure;

[0007] The two satellites are spatially symmetrically arranged, and the two satellites are rectangular parallelepiped structures with six faces. The two satellites include a set of parallel +Z faces and -Z faces, a set of parallel +X faces and -X faces, and a set of parallel +Y faces and -Y faces. The -Z faces of the two satellites are parallel, and the -Z faces of the two satellites are combined with the fairing 01 to form a flat area between the two satellites. For any of the satellites, the -Z wing 11 with the largest area in the solar wing is folded on the -Z face, the +X wing 12 in the solar wing is folded on the +X face, and the -X wing 13 in the solar wing is folded on the -X face.

[0008] When the fairing 01 is removed and the satellite is in the deployed state, the -Z wing 11, +X wing 12 and -X wing 13 adjust the wing surfaces to be parallel to the -Y plane through the hinge structure, and the -Y plane is the sun-facing plane.

[0009] In a first possible implementation of the first aspect of the present application, an intersatellite laser communication device 22 is installed on the +Y plane;

[0010] When the fairing 01 is removed and the satellite is in the deployed state, the +Y surface is the shaded surface.

[0011] In a second possible implementation of the first aspect of the present application, a parabolic antenna 21 for ground communication and a ground measurement and control antenna 41 are installed on the +Z plane, and the parabolic antenna 21 and the ground measurement and control antenna 41 are located in the arched area formed by the +Z plane and the fairing 01.

[0012] In a third possible implementation of the first aspect of the present application, two brackets 51, one bracket 52 and one bracket 53 are installed on the -Y surface, the bracket 51 is installed with one hinge 14, the bracket 52 is installed with two hinges 14, and the bracket 53 is installed with two hinges 14;

[0013] A hinge 14 mounted on a bracket 51 and a hinge 14 mounted on a bracket 53 are both connected to the +X wing 12;

[0014] Another hinge 14 mounted on the bracket 53 and one hinge 14 mounted on the bracket 52 are both connected to the -Z wing 11;

[0015] Another hinge 14 mounted on the bracket 51 and another hinge 14 mounted on the bracket 52 are both connected to the X-wing 13 .

[0016] In combination with the fourth possible implementation of the first aspect of the present application, in the fifth possible implementation of the first aspect of the present application, a solar wing locking device 15 is also installed on the -Z surface, the +X surface and the -X surface respectively.

[0017] In a fifth possible implementation of the first aspect of the present application, a docking ring 34 is installed on the -Y surface for connecting to a rocket adapter;

[0018] The -Y surface is also equipped with a propulsion device 43 with a gas cylinder. The propulsion device 43 is installed on the -Y surface from the outside of the satellite to the inside of the satellite, passing through the middle area of ​​the docking ring 34.

[0019] In the sixth possible implementation of the first aspect of the present application, the satellite provides installation positions for various components through the main frame 32. The main frame 32 is in an open state on the +Z plane and -Z plane sides. A +Z structural plate 31 for installing the parabolic antenna 21 is also provided on the main frame 32. Four independent operating covers 33 are also provided on the main frame 32 on the -Z plane side. The four operating covers 33 correspond to the four open openings reserved by the main frame 32 respectively. The main frame 32 is also provided with a GNSS antenna 49 on the -Z plane side.

[0020] In combination with the sixth possible implementation of the first aspect of the present application, in a seventh possible implementation of the first aspect of the present application, the main frame 32 is integrally formed of aluminum alloy.

[0021] In combination with the sixth possible implementation of the first aspect of the present application, in an eighth possible implementation of the first aspect of the present application, inside the main frame 32, the whole satellite integrated electronic unit 61 and the parabolic antenna integrated processor 23 are installed on the heat dissipation surfaces at both sides;

[0022] A battery pack 62, two fiber optic gyroscopes 45, two magnetometers 48 and three magnetic rods 47 are installed in the middle position. The battery pack 62 and the two fiber optic gyroscopes 45 are installed in an upside-down manner under the middle plate of the main frame 32. The two fiber optic gyroscopes 45 are symmetrical left and right along the battery pack 62, and the two magnetometers 48 are symmetrical front and back. Four obliquely installed reaction flywheels 46 are installed on the -Y surface inside the satellite through a bracket with a certain slope and are evenly distributed around.

[0023] In a second aspect, the present application provides an installation system for a satellite fixing structure based on a one-rocket-two-satellite parallel launch method, and the system is used to install a satellite fixing structure based on a one-rocket-two-satellite parallel launch method as provided in the first aspect of the present application or any possible implementation method of the first aspect.

[0024] From the above content, it can be concluded that the present application has the following beneficial effects:

[0025] Compared with the traditional solar wing structure design scheme that has the problem of high production cost, the present application is a three-wing single-fold solar wing design structure configured for satellites with a one-rocket two-satellite parallel launch method. The installation operation is simple, no truss is required, and only a general flip car is needed. It has low requirements for site, equipment and personnel, can effectively shorten the installation cycle, further reduce the production cost of satellite products, simplify the satellite assembly and disassembly process, and improve the production efficiency of satellites, which is in line with the future development trend of small commercial satellites. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] In order to more clearly illustrate the technical solutions in the embodiments of the present application, the drawings required for use in the description of the embodiments will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present application. For those skilled in the art, other drawings can be obtained based on these drawings without creative work.

[0027] Figure 1 This is an oblique schematic diagram of a layout of the binary satellite in the fairing of this application;

[0028] Figure 2 This is a bird's-eye view schematic diagram of a layout of the binary satellite in the fairing of this application;

[0029] Figure 3 This is a top-down oblique schematic diagram of the satellite retracting state of the present application;

[0030] Figure 4 This is a bottom-up oblique schematic diagram of the satellite in the deployed state of the present application;

[0031] Figure 5 This is a top-down oblique schematic diagram of the satellite in the deployed state of the present application;

[0032] Figure 6 This is a bottom-up oblique schematic diagram of the satellite retracted state of the present application;

[0033] Figure 7 A schematic diagram of the main frame of the satellite of this application;

[0034] Figure 8 A front view schematic diagram of the interior of the satellite of this application;

[0035] Fig. 9 A schematic diagram of the back of the interior of the satellite of this application;

[0036] Fig.10 This is a schematic diagram of the installation of the upper layer equipment of the satellite of this application;

[0037] Fig.11 This is a schematic diagram of the installation of the satellite mid-layer equipment of this application;

[0038] Fig.12 This is a schematic diagram of the installation of the lower-layer equipment of the satellite in this application. DETAILED DESCRIPTION

[0039] The following will be combined with the drawings in the embodiments of the present application to clearly and completely describe the technical solutions in the embodiments of the present application. Obviously, the described embodiments are only part of the embodiments of the present application, not all of the embodiments. Based on the embodiments in the present application, all other embodiments obtained by those skilled in the art without creative work are within the scope of protection of this application.

[0040] The terms "first", "second", etc. in the specification and claims of this application and the above-mentioned drawings are used to distinguish similar objects, and are not necessarily used to describe a specific order or sequence. It should be understood that the data used in this way can be interchangeable where appropriate, so that the embodiments described here can be implemented in an order other than that illustrated or described here. In addition, the terms "including" and "having" and any of their variations are intended to cover non-exclusive inclusions, for example, a process, method, system, product or device that includes a series of steps or modules is not necessarily limited to those steps or modules clearly listed, but may include other steps or modules that are not clearly listed or inherent to these processes, methods, products or devices. The naming or numbering of steps in this application does not mean that the steps in the method flow must be executed in the time / logical sequence indicated by the naming or numbering. The process steps that have been named or numbered can change the execution order according to the technical purpose to be achieved, as long as the same or similar technical effects can be achieved.

[0041] The division of modules in this application is a logical division. There may be other division methods when it is implemented in actual applications. For example, multiple modules can be combined or integrated into another system, or some features can be ignored or not executed. In addition, the mutual coupling or direct coupling or communication connection shown or discussed can be through some interfaces, and the indirect coupling or communication connection between modules can be electrical or other similar forms, which are not limited in this application. In addition, the modules or submodules described as separate components may or may not be physically separated, may or may not be physical modules, or may be distributed in multiple circuit modules, and some or all of the modules may be selected according to actual needs to achieve the purpose of the present application.

[0042] First, see Figure 1 The schematic diagram of the layout of the dual satellites in the fairing of the present application is shown in FIG. Figure 2 An overhead schematic diagram of the layout of the two satellites of the present application in the fairing is shown. The present application provides a satellite fixing structure based on the parallel launch of two satellites with one rocket. In its satellite fixing structure, two identical satellites are placed back to back in the fairing 01, and each satellite is equipped with a solar wing with a three-wing single-fold structure (including a -Z wing 11, a +X wing 12 and a -X wing 13).

[0043] from Figure 1 , 2 It can also be seen that in the satellite fixing structure involved in the present application, the two satellites are spatially symmetrically arranged, and the two satellites are six-sided rectangular parallelepiped structures.

[0044] Continue reading Figure 3 A schematic diagram of the satellite folding state of the present application is shown from top to bottom. Both satellites include a set of parallel +Z planes and -Z planes, a set of parallel +X planes and -X planes, and a set of parallel +Y planes and -Y planes (defined according to the coordinate system of a single satellite).

[0045] Combination Figure 1 , Figure 2 It can also be seen that the spatial relationship between the two satellites is: 1. The -Z planes of the two satellites are parallel, and 2. The -Z planes of the two satellites combined with fairing 01 form a flat area between the two satellites (that is, there is a gap between the two satellites, which is flat and has a large area).

[0046] Furthermore, for any of the satellites, the -Z wing 11 with the largest area among the solar wing is folded on the -Z plane (it can be seen from the layout that the shape of the flat area between the two satellites is very consistent with the flat shape of the solar wing, so the wing with the largest area among the three wings is folded on this plane of the satellite), the +X wing 12 among the solar wing is folded on the +X plane, and the -X wing 13 among the solar wing is folded on the -X plane.

[0047] The -Z wing 11, +X wing 12 and -X wing 13 are specifically connected to the satellite through a hinge structure, so that the attitude adjustment can be performed after the satellite is launched and in the space environment. Specifically, when the fairing 01 is removed and the satellite is in the unfolded state, Figure 4 A schematic diagram of the satellite of the present application in an unfolded state is shown obliquely from bottom to top. The -Z wing 11, +X wing 12 and -X wing 13 are adjusted to be parallel to the -Y plane through a hinge structure. The -Y plane is the sun-facing plane (in the sun-facing direction, with the side where the battery is installed pointing to the sun).

[0048] Under the above three-wing single-fold solar wing design structure, two identical satellites are arranged back to back in the rocket fairing 01, so that the wing with a larger area can make full use of the space at the diameter of the fairing. At the same time, compared with traditional double-wing multi-fold solar wing and other solar wing structures, the three-wing single-fold can reduce the use of high-value products such as hinges, so that satellite products can effectively and fully utilize the space in the launch vehicle fairing while reducing costs, further reducing the volume of satellite products and making them more compact.

[0049] It can be seen from the above embodiments that compared with the traditional solar wing structure design scheme, there is a problem of high production cost. The present application is a three-wing single-fold structure solar wing design structure configured for a satellite with a one-rocket two-satellite parallel launch method. The installation operation is simple, no truss is required, and only a general flip car is required. It has low requirements for site, equipment and personnel, can effectively shorten the installation cycle, further reduce the production cost of satellite products, simplify the satellite assembly and disassembly process, and improve the production efficiency of satellites, which is in line with the future development trend of small commercial satellites.

[0050] In addition, the solar wing design structure based on the three-wing single-fold structure can also be combined with the fixed structure design of other components to bring other advantages and thus achieve better application value.

[0051] For example, from Figure 3 It can be seen that, as a specific implementation method, an intersatellite laser communication device 22 is installed on the +Y plane of the satellite; when the fairing 01 is removed and the satellite is in the unfolded state, the +Y plane is the shaded side.

[0052] It can be understood that after the three-wing single-fold solar wing of the present application is unfolded in orbit, the solar wing is closer to the main body of the satellite, which can form a shielding effect on the equipment on the shady side of the satellite. Generally speaking, networked communication satellites carry laser payloads for inter-satellite communications, which are easily affected by solar radiation. At this time, the solar wing closer to the satellite can provide better shielding protection for the laser payload.

[0053] In this case, the -Z wing 11, +X wing 12 and -X wing 13 are tightly wrapped around the satellite body, and the intersatellite laser communication equipment 22 is installed on the opposite surface away from the plane where the solar wing is unfolded, which can help the intersatellite laser communication equipment 22 to shield the sunlight to a certain extent. At the same time, considering that the +Y surface of the satellite is on the shady side with a lower temperature, and the intersatellite laser communication equipment 22 that is turned on for a long time has a higher power consumption, installing it on the +Y surface can just achieve a good heat dissipation effect, thereby ensuring that the intersatellite laser communication equipment 22 is in a more suitable ambient temperature, and then can stably and reliably complete its communication work.

[0054] In addition, taking the satellite deployment state as an example, see Figure 5 A schematic diagram of the satellite of the present application in an unfolded state is shown diagonally from top to bottom. On both sides of the intersatellite laser communication device 22 on the +Y plane, a star sensor 42 is connected to each of them through a bracket with an oblique angle, which is used to measure the angular position of a brighter star relative to the satellite and compare it with the angular position parameters of the star in the ephemeris, so as to determine the satellite's spatial attitude.

[0055] In addition, it can be understood that ground communication equipment is installed on the side of the satellite away from the larger solar wing. The general forms of this communication equipment include parabolic antenna 21, foldable and unfoldable antenna and multi-element array antenna, which can just make use of the arched area formed between the satellite plane and the fairing 01. At the same time, the heavier communication equipment and the solar wing on the opposite side of the satellite can have a balancing effect, so that the center of mass position of the satellite in the longitudinal direction can be as centered as possible, which is conducive to the operation of other systems of the satellite.

[0056] As an example, Figure 1 and Figure 2 As shown, the +Z plane of the satellite is equipped with a parabolic antenna 21 for ground communication. The parabolic antenna 21 is placed on the surface opposite to the larger solar wing installation surface. The parabolic antenna 21 is in the arched area formed by the +Z plane and the fairing 01.

[0057] In addition, if Figure 3 As shown, a ground tracking antenna 41 can also be installed on the +Z plane of the satellite. The installation position of the tracking antenna 41 is subject to not being affected by the parabolic antenna 21. If necessary, a bracket can be designed to increase the height. In addition, the remaining space on the +Z plane can be used to carry some equipment that needs to be pointed to the ground.

[0058] In addition, from Figure 4 , Figure 5 It can also be seen that a solar wing locking device 15 can be installed on the -Z surface, +X surface and -X surface respectively, so as to further reinforce the corresponding solar wing.

[0059] In addition, returning to the hinge structure itself involved in the solar wing, see Figure 6 A schematic diagram of the present invention showing the folded state of the satellite is shown from bottom to top. For the hinge structure involved in the solar wing, there may be the following specific implementation schemes:

[0060] -The Y surface is installed with two brackets 51, one bracket 52 and one bracket 53. Bracket 51 is installed with one hinge 14, bracket 52 is installed with two hinges 14, and bracket 53 is installed with two hinges 14. There are six hinges 14 in total, which are divided into three groups and are respectively connected to the corresponding solar wings.

[0061] Among them, a hinge 14 installed on a bracket 51 and a hinge 14 installed on a bracket 53 are both connected to the +X wing 12, and these two hinge structures are close to the +X plane, and the +X wing 12 is folded on the +X plane through the hinge structure;

[0062] Another hinge 14 installed on the bracket 53 and one hinge 14 installed on the bracket 52 are both connected to the -Z wing 11. The two hinge structures are close to the -Z plane, and the -Z wing 11 is folded on the -Z plane through the hinge structure;

[0063] Another bracket 51 and another hinge 14 installed on bracket 52 are both connected to the -X wing 13. These two hinge structures are close to the -X plane, and through these hinge structures, the -X wing 13 folds on the -X plane.

[0064] Furthermore, Figure 6 It can also be seen that digital sun sensors 44 are respectively installed on bracket 52 and bracket 53 for detecting the position of sunlight, thereby helping to determine the orientation of the satellite relative to the solar coordinate system and determine the spatial attitude of the satellite.

[0065] Continuing to refer to Figure 6 , corresponding to the assembly part between the satellite and the rocket, a docking ring 34 can also be installed on the -Y plane for connecting with the rocket adapter;

[0066] A propulsion device 43 with a gas cylinder can also be installed on the -Y plane. The propulsion device 43 is installed on the -Y plane from the outside of the satellite to the inside of the satellite through the middle area of the docking ring 34. In this way, while the satellite is assembled to the rocket, it is also convenient to independently remove the propulsion device 43 after the overall satellite assembly is completed.

[0067] Furthermore, from Figure 3 , 4 , 5, and 6, it can also be seen that the relevant components involved in the above content are specifically installed at the corresponding positions of the main frame 32, or rather, the satellite provides the installation positions of each component through the main frame 32.

[0068] Regarding this main frame 32, refer to Figure 7 a schematic diagram of the satellite main frame of the present application shown in

[0069] The main frame 32 is in an open state on the +Z plane and -Z plane sides, so as to facilitate the installation of internal equipment and the operation of personnel.

[0070] In addition, a horizontal and vertical cross-plane can be designed inside the main frame 32 to enhance the stiffness of the entire satellite and provide more equipment installation surfaces.

[0070] As Figure 4 shown, a +Z structural plate 31 (an independent structural plate) for installing the parabolic antenna 21 can also be provided on the main frame 32. After the installation of the satellite internal equipment is completed, the +Z structural plate 31 with the antenna is then sealed, and at the same time, the +Z structural plate 31 can also play an effect of further enhancing the stiffness of the entire satellite.

[0071] In addition, four independent operation covers 33 are provided on the -Z plane side of the main frame 32, and the four operation covers 33 respectively correspond to four open openings (in a cross shape) left by the main frame 32.

[0072] Since there are fewer devices arranged on the -Z surface, the sealing plates on this surface are installed last. In order to improve the maintainability of the entire satellite, considering that the cross frame (corresponding to four openings) divides the internal space of the satellite into four areas, the sealing plates on the -Z surface are designed as four independent pieces. Once a faulty device occurs, it is only necessary to open the corresponding operating cover 33 on the -Z surface instead of all the operating covers 33, which reduces the difficulty of operation and improves the maintainability of the satellite.

[0073] In addition, on the side of the -Z plane away from the -Z solar wing, the main frame 32 is also provided with a GNSS antenna 49 on the -Z plane side, wherein the antenna is directly mounted on the main frame 32 of the satellite. At the same time, the locking device 15 of the -Z solar wing is also directly mounted on the main frame 32 of the satellite. It can be seen that when the four independent operating covers 33 are removed, there is no need to repeatedly disassemble the above two devices, which reduces the repetition of work and the difficulty of operation.

[0074] Among them, the main frame 32 can be specifically an integrated molding of aluminum alloy (obtained by 3D printing, casting or stir friction welding, etc.). The integrated molding solution can, on the one hand, ensure the rapid installation of equipment with installation accuracy requirements through early mechanical processing, reduce the precision measurement time in the satellite assembly stage, and even eliminate the precision measurement link; on the other hand, the integrated molding of aluminum alloy can ensure that the entire satellite has high thermal conductivity, reduce the complexity of the thermal control design of the entire satellite, reduce or even eliminate the use of high-value products such as heat pipes, and also effectively reduce the difficulty of assembly, which can further achieve the purpose of reducing costs and increasing efficiency.

[0075] from Figure 7 It can also be seen that there are two openings on the +Y side and -Y side of the satellite main frame 32, which can be used for the installation of the inter-satellite laser communication equipment 22 and the propulsion device 43 from the outside to the inside respectively.

[0076] The present application will continue to describe the contents of the structural optimization configuration inside the satellite.

[0077] See also Figure 8 A front view schematic diagram of the interior of the satellite of the present application is shown. Fig. 9 A schematic diagram of the back of the satellite of the present application is shown. The interior of the satellite, that is, the interior of the main frame 32, can be roughly divided into two areas (divided into upper and lower areas along the Y axis). The lower area is mainly composed of platform equipment, and the upper area is mainly composed of payload equipment. Among them, equipment with high power consumption, such as the whole satellite integrated electronic unit 61 and the parabolic antenna integrated processor 23, are directly installed on the heat dissipation surfaces on both sides.

[0078] See also Fig.10An installation diagram of the upper layer equipment of the satellite of the present application is shown. Inside the main frame 32, the upper layer is mainly installed with the whole satellite integrated electronic unit 61 and the parabolic antenna integrated processor 23. Both of them have high power consumption, so they are directly installed on both sides of the satellite respectively. The upper layer of the satellite is the side away from the sun and has a lower temperature. The two are placed in the upper half of the satellite, which can achieve a good heat dissipation effect.

[0079] Taking into account that a large number of cables will be drawn from the satellite integrated electronic unit 61, the parabolic antenna integrated processor 23 and the inter-satellite laser communication equipment 22, a reinforcement frame 55 can be designed between the three to fix the cables and enhance the local rigidity of the satellite.

[0080] Look again Figure 8 , Fig. 9 Inside the main frame 32, the equipment that needs to be kept warm is located in the middle, such as the battery pack 62 and the two fiber optic gyroscopes 45.

[0081] Combination Fig.11 A schematic diagram of the installation of the middle-level equipment of the present application is shown, in which a battery pack 62, two fiber optic gyroscopes 45, two magnetometers 48 and three magnetic rods 47 are installed in the middle position (the middle position divided along the upper and lower Y axes). The battery pack 62 and the two fiber optic gyroscopes 45 are installed in an upside-down manner under the middle plate of the main frame 32, which can effectively utilize the lower space inside the satellite.

[0082] The two fiber optic gyroscopes 45 are symmetrical left and right along the battery pack 62 (divided into left and right along the X axis), and the two magnetometers 48 are symmetrical front and back (divided into front and back along the Z axis, both located on the -X side) and reasonably distributed. The above measures are taken to effectively control the center of mass of the entire satellite in the horizontal direction (X and Z directions).

[0083] In addition, a transverse support member 54 is provided between the battery pack 62 and the propulsion device 43 to enhance the local rigidity of the satellite and also to provide a path for the cable network.

[0084] Look again Fig.12 An installation diagram of the lower-level equipment of the satellite of the present application is shown, in which four obliquely installed reaction flywheels 46 are installed on the -Y surface inside the satellite through brackets with a certain slope and are evenly distributed around (such as around the propulsion device 43). Such an installation method is also conducive to controlling the center of mass of the entire satellite.

[0085] From the above content, it can be seen that the satellite fixing structure based on the one-rocket two-satellite parallel launch method provided by the present application not only adjusts the external fixing structure involved in the satellite from the outside, but also adjusts the fixing structure of the relevant components involved in the satellite from the inside, from the initial streamlined and simple solar wing design structure to other aspects of the optimized design to promote the good realization of the satellite fixing structure, thereby significantly reducing the required production cost.

[0086] On the basis of the satellite fixing structure based on the one-rocket-two-satellite parallel launch mode provided in the above present application, the present application may further be configured with an installation system based on its structural characteristics in actual applications, that is, an installation system for the satellite fixing structure based on the one-rocket-two-satellite parallel launch mode. This system is matched with the satellite fixing structure based on the one-rocket-two-satellite parallel launch mode provided in the present application, and is used to install the satellite fixing structure based on the one-rocket-two-satellite parallel launch mode provided in the present application. Specifically, it may be an installation system of a rocket, or an installation system built into a rocket, or a system independent of the installation system of a rocket. It may involve different components for installation, which are not specifically limited here.

[0087] The above is a detailed introduction to the satellite fixing structure and its installation system based on the one-rocket two-satellite parallel launch method provided by the present application. Specific examples are used in this article to illustrate the principles and implementation methods of the present application. The description of the above embodiments is only used to help understand the content of the present application and its core idea; at the same time, for technical personnel in this field, according to the idea of ​​the present application, there will be changes in the specific implementation method and application scope. In summary, the content of this specification should not be understood as a limitation on the present application.

Claims

1. A satellite fixing structure based on a two-satellite-in-one-rocket parallel launch method, characterized in that: In the satellite fixing structure, two identical satellites are placed back to back in the fairing, and each of the satellites is equipped with a solar wing with a three-wing single-fold structure; The two satellites are spatially symmetrically arranged, the two satellites are six-sided rectangular parallelepiped structures, the two satellites include a group of parallel +Z and -Z surfaces, a group of parallel +X and -X surfaces, and a group of parallel +Y and -Y surfaces, the -Z surfaces of the two satellites are parallel, and the -Z surfaces of the two satellites form a flat area between the two satellites in combination with the fairing, and for any of the satellites, the -Z wing with the largest area among the solar wings is folded on the -Z surface, the +X wing among the solar wings is folded on the +X surface, and the -X wing among the solar wings is folded on the -X surface; When the fairing is removed and the satellite is in the deployed state, the -Z wing, the +X wing and the -X wing adjust the wing surfaces to be parallel to the -Y plane through the hinge structure, the -Y plane is the sun-facing plane, and the +Y plane is the shady side; The satellite provides installation positions for various components through a main frame, the main frame is in an open state on the +Z plane and the -Z plane side, and a +Z structural plate for installing a parabolic antenna is also provided on the main frame, and four independent operation covers are also provided on the main frame on the -Z plane side, and the four operation covers correspond to the four openings left by the main frame respectively, and the main frame is also provided with a GNSS antenna on the -Z plane side; The +Y surface is equipped with intersatellite laser communication equipment; The interior of the main frame is divided into two layers along the +Y plane to the -Y plane. The integrated electronic unit of the whole satellite and the integrated processor of the parabolic antenna located in the upper layer are installed on the heat dissipation surfaces at both sides. A reinforcement frame is also designed between the integrated electronic unit of the whole satellite, the integrated processor of the parabolic antenna and the intersatellite laser communication equipment. A battery pack, two fiber optic gyroscopes, two magnetometers and three magnetic rods are installed in the middle position. The battery pack and the two fiber optic gyroscopes are installed in an upside-down manner under the middle plate of the main frame. The two fiber optic gyroscopes are symmetrical left and right along the battery pack, and the two magnetometers are symmetrical front and back. A horizontal support is arranged between the battery pack and the propulsion device. Four obliquely installed reaction flywheels are installed on the -Y plane inside the satellite through brackets with a certain slope and are evenly distributed around.

2. The satellite fixing structure according to claim 1, characterized in that: The +Z surface is installed with a parabolic antenna for ground communication and a ground measurement and control antenna, and the parabolic antenna and the ground measurement and control antenna are located in an arched area formed by the +Z surface and the fairing.

3. The satellite fixing structure according to claim 1, characterized in that: The -Y surface is installed with two first brackets, a second bracket and a third bracket, the first bracket is installed with a hinge, the second bracket is installed with two hinges, and the third bracket is installed with two hinges; A hinge mounted on the first bracket and a hinge mounted on the third bracket are both connected to the +X wing; Another hinge mounted on the third bracket and one hinge mounted on the second bracket are both connected to the -Z wing; Another hinge mounted on said first bracket and another hinge mounted on said second bracket are both connected to said X-wing.

4. The satellite fixing structure according to claim 3, characterized in that: The -Z surface, the +X surface and the -X surface are also respectively equipped with a solar wing locking device.

5. The satellite fixing structure according to claim 1, characterized in that: The -Y surface is provided with a docking ring for connecting to a rocket adapter; The -Y surface is also equipped with a propulsion device with a gas cylinder, and the propulsion device is installed on the -Y surface from the outside of the satellite to the inside of the satellite, passing through the middle area of ​​the docking ring.

6. The satellite fixing structure according to claim 1, characterized in that: The main frame is integrally formed from aluminum alloy.

7. A satellite fixed structure installation system based on a two-satellite-in-one-rocket parallel launch method, characterized in that: The system is used to install a satellite fixing structure based on a one-rocket two-satellite parallel launch method as described in any one of claims 1 to 6.

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