Satellite assembly structure and satellite launch control method

Abstract

The application relates to the field of aerospace technology and specifically discloses a satellite combination structure and a satellite launching control method. The satellite combination structure comprises a first satellite and at least two second satellites; each second satellite is arranged side by side on the same side of the first satellite, each second satellite is connected with the first satellite through a separation mechanism, and adjacent two second satellites are connected through a separation mechanism; the first satellite comprises a propulsion system and a first load-bearing structure, the propulsion system comprises a main propeller and a fuel tank, the fuel tank is located inside the first load-bearing structure, and the main propeller is connected to the outer side of the bottom of the first load-bearing structure. Therefore, without adding complex structures such as upper stages or propulsion power cabins, the carrying capacity is effectively saved under the condition of guaranteeing satellite functions and on-orbit deployment, new space garbage is avoided, satellite launching costs are saved, and system construction costs are reduced.

Description

Technical Field

[0001] This application relates to the field of aerospace technology, and in particular to a satellite assembly structure and a satellite launch control method. Background Technology

[0002] For medium and high orbit satellites, due to their powerful functions, large size, and heavy weight, they are mostly launched in a one-rocket-one-satellite, one-rocket-two-satellite, or one-rocket-three-satellite manner. At the same time, due to the limited carrying capacity of the rocket, the satellite is generally launched into a geostationary transfer orbit (GTO) or a medium Earth transfer orbit (MTO), and then the satellite or the upper stage carrying the satellite performs orbit transfer, so that the satellite finally enters a geostationary orbit (GEO) or a medium Earth orbit (MEO).

[0003] When launching one satellite with a single rocket, the launched satellite uses a large-capacity fuel tank to carry sufficient fuel for orbit transfer. However, when launching two or three satellites with a single rocket, due to structural limitations and space constraints within the fairing, it's impossible for each satellite to carry the large amount of fuel required for orbit transfer using a large-capacity fuel tank. Therefore, an upper stage or propulsion module is needed to carry the necessary fuel for orbit transfer, sending two or three satellites into their operational orbits before separation and deployment. The upper stage or propulsion module then separates from the satellites, is deactivated, and deorbited. This launch method has the following drawbacks: First, the upper stage or propulsion module remains in Earth orbit for an extended period, becoming useless debris. Second, the upper stage or propulsion module itself is ineffective "dead weight" during launch, reducing the rocket's payload capacity and significantly decreasing its effective carrying capacity. Third, the need for a dedicated upper stage or propulsion module increases satellite launch costs and the risks during launch and orbit insertion. Summary of the Invention

[0004] Therefore, it is necessary to provide a satellite assembly structure and satellite launch control method to address the above problems.

[0005] According to a first aspect of the embodiments of this application, a satellite combination structure is provided, the satellite combination structure including a first satellite and at least two second satellites; each second satellite is arranged side by side on the same side of the first satellite, each second satellite is connected to the first satellite via a separation mechanism, and adjacent second satellites are connected via a separation mechanism;

[0006] The first satellite includes a propulsion system and a first load-bearing structure. The propulsion system includes a main thruster and a fuel tank. The fuel tank is located inside the first load-bearing structure, and the main thruster is connected to the bottom outer side of the first load-bearing structure.

[0007] In one embodiment, the first load-bearing structure is a central load-bearing cylinder configuration, including a load-bearing cylinder, a plurality of first bulkheads, a plurality of first floor plates, and a plurality of first outer cabin plates;

[0008] Each of the first bulkheads is connected to the outer peripheral surface of the load-bearing cylinder at intervals along the circumferential direction of the load-bearing cylinder, each of the first floor plates is connected to the outer peripheral surface of the load-bearing cylinder at intervals along the axial direction of the load-bearing cylinder, and each of the first bulkheads and each of the first floor plates intersect to form a plurality of independent spaces located outside the load-bearing cylinder. The fuel tank is arranged in some of the independent spaces, and each of the first outer cabin plates is sequentially arranged outside the first load-bearing structure to enclose each of the independent spaces.

[0009] In one embodiment, the second satellite includes a second load-bearing structure. The second load-bearing structure is a hexahedron configuration, including a second partition board, a plurality of second floor plates, a plurality of second side plates, and a plurality of second outer cabin plates;

[0010] A plurality of the second outer cabin plates are sequentially connected and enclosed to form a hexahedron space. The second partition board is arranged in the hexahedron space. Each of the second side plates is perpendicular to the second partition board and is connected to one side surface of the second partition board at intervals. Each of the second floor plates is connected at intervals between two adjacent second side plates. The second floor plates are perpendicular to the second side plates and perpendicular to the second partition board.

[0011] In one embodiment, the main thruster includes an engine with a thrust of 750N.

[0012] In one embodiment, the separation mechanism includes a plurality of point-type connection and separation devices.

[0013] In one embodiment, a single second satellite and the first satellite are connected by four of the point-type connection and separation devices, and two adjacent second satellites are connected by four of the point-type connection and separation devices.

[0014] In one embodiment, the satellite combined structure includes one first satellite and two second satellites. The two second satellites are arranged side by side on the same side of the first satellite, forming a "pin" - shaped structure with the first satellite.

[0015] According to the second aspect of the embodiments of the present application, there is provided a satellite launch control method for performing launch control on the satellite combined structure as described above. The satellite launch control method includes:

[0016] Controlling the rocket carrying the satellite combined structure to be launched into a transfer orbit;

[0017] Upon receiving the satellite-rocket separation signal, the system controls the satellite assembly structure to separate from the rocket. Propelled by the first satellite, the satellite assembly structure enters its predetermined working orbit.

[0018] Upon receiving a separation command for the satellite assembly structure, the system controls the assembly of each of the second satellites to separate from the first satellite.

[0019] Once the first preset condition is met, control the separation of each of the second satellites.

[0020] In one embodiment, the first preset condition includes: the distance between the combination of each of the second satellites and the first satellite reaches a first preset distance.

[0021] In one embodiment, the step of controlling the satellite assembly structure to separate from the rocket after receiving the satellite-rocket separation signal, and the satellite assembly structure entering the predetermined working orbit under the propulsion of the first satellite, includes:

[0022] The first satellite enters rate-damped mode, and after meeting the corresponding conditions, the battery array deploys, enters cruise mode, and completes solar orientation. In apogee ignition mode, the orbital attitude of the satellite assembly structure is adjusted, and the satellite assembly structure is brought into the predetermined working orbit through multiple apogee orbital maneuvers.

[0023] After the step of controlling the separation of each of the second satellites upon meeting the first preset condition, the satellite launch control method further includes:

[0024] Once the second preset condition is met, the battery arrays of each of the second satellites deploy, enter cruise mode, and complete sun orientation.

[0025] The satellite assembly structure provided in this application embodiment includes a first satellite as the carrier satellite, which can carry at least two second satellites. Each second satellite is arranged side-by-side on the same side of the first satellite, and each second satellite is connected to the first satellite via a separation mechanism. Adjacent second satellites are also connected via separation mechanisms. In other words, the connection and separation of each satellite can be achieved through these separation mechanisms. In this application, the first satellite, as the carrier satellite, can inherit the design concept of traditional satellites. That is, the satellite assembly structure can be propelled from the transfer orbit to the target orbit by a main thruster connected to the outer bottom of the first load-bearing structure and a fuel tank containing the fuel required for orbit transfer, located inside the first load-bearing structure. This eliminates the need for complex structures such as upper stages or propulsion modules. While ensuring satellite functionality and on-orbit deployment conditions, this effectively saves carrying capacity, avoids generating new space debris, saves satellite launch costs, and reduces system construction costs. Attached Figure Description

[0026] Figure 1 This is a schematic diagram of a satellite assembly structure provided in an embodiment of this application;

[0027] Figure 2 This is a schematic diagram of a satellite assembly structure provided in an embodiment of this application;

[0028] Figure 3 This is a schematic diagram of the structure of the first satellite in a satellite assembly structure provided in an embodiment of this application;

[0029] Figure 4 This is a schematic diagram of the second load-bearing structure of the second satellite in a satellite assembly structure provided in an embodiment of this application;

[0030] Figure 5 A schematic diagram of the flight state of a satellite assembly structure provided in an embodiment of this application;

[0031] Figure 6 A schematic diagram of the on-orbit flight status of the first satellite in a satellite assembly structure provided in an embodiment of this application;

[0032] Figure 7 A schematic diagram of the on-orbit flight status of the second satellite in a satellite assembly structure provided in an embodiment of this application;

[0033] Figure 8 This is a flowchart of a satellite launch control method provided in an embodiment of this application.

[0034] Explanation of reference numerals in the attached figures:

[0035] 100. First satellite; 111. Main thruster; 112. Fuel tank; 121. Support cylinder; 122. First bulkhead; 123. First shelf; 124. First outer compartment panel; 200. Second satellite; 211. Second bulkhead; 212. Second shelf; 213. Second side panel; 214. Second outer compartment panel; 300. Separation mechanism; 310. Point-connected separation device; 400. Payload. Detailed Implementation

[0036] To facilitate understanding of this application, a more complete description will be provided below with reference to the accompanying drawings. Preferred embodiments of this application are shown in the drawings. However, this application can be implemented in many different forms and is not limited to the embodiments described herein. Rather, these embodiments are provided to provide a more thorough and complete understanding of the disclosure of this application.

[0037] In this application, unless otherwise expressly specified and limited, the terms "installation," "connection," "joining," and "fixing," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components, unless otherwise expressly limited. Those skilled in the art can understand the specific meaning of the above terms in this application according to the specific circumstances.

[0038] The terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Therefore, a feature defined as "first" or "second" may explicitly or implicitly include at least one of that feature. In the description of this application, "multiple" means at least two, such as two, three, etc., unless otherwise explicitly specified.

[0039] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by one of ordinary skill in the art to which this application belongs. The terminology used herein in the specification of this application is for the purpose of describing particular embodiments only and is not intended to be limiting of the application. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed items.

[0040] In the field of spacecraft launch technology, to improve launch efficiency, the application of multiple-satellite launches (MSLs) is becoming increasingly widespread, meaning launching more satellites into artificial Earth orbit at once. Given a fixed rocket carrying capacity and satellite orbit, the number of satellites launched in a single launch is related to factors such as the satellite's mass, structure, separation device, fairing space, satellite layout within the fairing, and separation method. Currently, multiple-satellite launches of low-Earth orbit (LEO) microsatellites typically use a central support structure or auxiliary support structure to secure the launched satellites, which then separate one by one after entering orbit. Alternatively, a stacked launch configuration can be used, but this is only suitable for LEO communication satellites of the same type that are easily designed as flat-panel structures. For medium- and high-Earth orbit (MEO) satellites with complex functions and special payloads, these methods are difficult to apply.

[0041] For medium and high orbit satellites, due to their powerful functions, large size, and heavy weight, they are mostly launched in a one-rocket-one-satellite, one-rocket-two-satellite, or one-rocket-three-satellite manner. At the same time, due to the limited carrying capacity of the rocket, the satellite is generally launched into a geostationary transfer orbit (GTO) or a medium Earth transfer orbit (MTO), and then the satellite or the upper stage carrying the satellite performs orbit transfer, so that the satellite finally enters a geostationary orbit (GEO) or a medium Earth orbit (MEO).

[0042] When launching one satellite with a single rocket, the launched satellite uses a large-capacity fuel tank to carry sufficient fuel for orbit transfer. However, when launching two or three satellites with a single rocket, due to structural limitations and space constraints within the fairing, it's impossible for each satellite to carry the large amount of fuel required for orbit transfer using a large-capacity fuel tank. Therefore, an upper stage or propulsion module is needed to carry the necessary fuel for orbit transfer, sending two or three satellites into their operational orbits before separation and deployment. The upper stage or propulsion module then separates from the satellites, is deactivated, and deorbited. This launch method has the following drawbacks: First, the upper stage or propulsion module remains in Earth orbit for an extended period, becoming useless debris. Second, the upper stage or propulsion module itself is ineffective "dead weight" during launch, reducing the rocket's payload capacity and significantly decreasing its effective carrying capacity. Third, the need for a dedicated upper stage or propulsion module increases satellite launch costs and the risks during launch and orbit insertion.

[0043] To address the aforementioned issues, this application provides a satellite assembly structure and a satellite launch control method.

[0044] In one embodiment, a satellite assembly structure is provided. Each satellite in this assembly structure can be a medium-to-high orbit satellite. This design avoids the drawbacks of traditional satellite structures and launch layouts, eliminating the need for an additional upper stage or propulsion module. While ensuring satellite functionality and on-orbit deployment conditions, it effectively saves launch capacity, avoids generating new space debris, and reduces satellite launch costs, thus lowering system construction costs.

[0045] Reference Figure 1 and Figure 2 The satellite assembly structure provided in this embodiment includes a first satellite 100 and at least two second satellites 200. Each second satellite 200 is arranged side by side on the same side of the first satellite 100, and each second satellite 200 is connected to the first satellite 100 via a separation mechanism 300. Adjacent second satellites 200 are also connected via a separation mechanism 300.

[0046] The first satellite 100 includes a propulsion system and a first load-bearing structure. The propulsion system includes a main thruster 111 and a fuel tank 112. The fuel tank 112 is located inside the first load-bearing structure, and the main thruster 111 is connected to the bottom outer side of the first load-bearing structure.

[0047] In the satellite assembly structure provided in this embodiment, the first satellite 100 serves as the carrier satellite, capable of carrying at least two second satellites 200. Each second satellite 200 is arranged side-by-side on the same side of the first satellite 100, and each second satellite 200 is connected to the first satellite 100 via a separation mechanism 300. Adjacent second satellites 200 are also connected via separation mechanisms. That is, the connection and separation of each satellite can be achieved through the separation mechanisms 300. In this application, the first satellite 100, as the carrier satellite, can inherit the design concept of traditional satellites. Specifically, the satellite assembly structure can be propelled from the transfer orbit to the target orbit via the main thruster 111 connected to the outer side of the bottom of the first load-bearing structure and the fuel tank 112 containing the fuel required for orbit transfer, located inside the first load-bearing structure. This eliminates the need for complex structures such as upper stages or propulsion modules. While ensuring satellite functionality and on-orbit deployment conditions, this effectively saves carrying capacity, avoids generating new space debris, saves satellite launch costs, and reduces system construction costs.

[0048] In the satellite assembly structure, the first satellite 100 serves as the carrier satellite, and each of the second satellites 200 serves as the carried satellite. Each of the second satellites 200 is carried by the first satellite 100. During launch and orbit transfer, the first satellite 100 and each of the second satellites 200 are connected as a single unit by separation mechanisms 300. The fuel tank 112 inside the first satellite 100 carries the fuel required for orbit transfer. The propulsion system of the first satellite 100 propels the entire satellite assembly structure from the transfer orbit to the target orbit. Then, the individual satellites are separated and deployed to complete their orbital insertion. In practical applications, the main thruster 111 of the first satellite 100 may include a 750N engine.

[0049] Reference Figure 3 In one embodiment, the first load-bearing structure is a central load-bearing cylinder 121 configuration, including a load-bearing cylinder 121, a plurality of first bulkheads 122, a plurality of first floorboards 123, and a plurality of first outer cabin panels 124.

[0050] In this configuration, each of the first partition frames 122 is circumferentially connected to the outer peripheral surface of the load-bearing cylinder 121, and each of the first layer plates 123 is axially connected to the outer peripheral surface of the load-bearing cylinder 121. Each of the first partition frames 122 and each of the first layer plates 123 intersect to form several independent spaces located around the load-bearing cylinder 121. The fuel tank 112 is disposed in some of the independent spaces. Each of the first outer panels 124 is sequentially disposed around the periphery of the first load-bearing structure to enclose each of the independent spaces.

[0051] The first load-bearing structure, which is formed by the combination of the load-bearing cylinder 121, several first bulkheads 122, several first layer plates 123 and several first outer cabin plates 124, can effectively increase the overall strength of the satellite assembly structure, with a moderate satellite altitude and good modal response.

[0052] The interface connecting the first satellite 100 to the launch vehicle can employ a relatively mature traditional encasing separation mechanism 300, which may include components such as explosive bolts, V-shaped clamps, limit springs, and tension springs. During launch, the first satellite 100 and the launch vehicle are fixedly connected via the encasing separation mechanism 300.

[0053] Reference Figure 4 In one embodiment, the second satellite 200 includes a second load-bearing structure, which is a hexahedral configuration and includes a second partition 211, a plurality of second layer plates 212, a plurality of second side plates 213 and a plurality of second outer cabin plates 214.

[0054] In this configuration, several second outer panels 214 are sequentially connected to form a hexahedral space. A second partition 211 is disposed within the hexahedral space. Each second side panel 213 is perpendicular to the second partition 211 and is spaced apart from one side surface of the second partition 211. Each second layer panel 212 is spaced apart between two adjacent second side panels 213. The second layer panel 212 is perpendicular to the second side panel 213 and perpendicular to the second partition 211.

[0055] The second load-bearing structure, consisting of a second bulkhead 211, several second layer plates 212, several second side plates 213, and several second outer cabin plates 214, forms the hexahedral configuration described above. This structure provides torsional stiffness and longitudinal shear stiffness, facilitating the layout of equipment within the second satellite 200 configuration and the installation of equipment inside and outside the cabin.

[0056] Reference Figure 6 and Figure 7 In one embodiment, the separation mechanism 300 includes a plurality of point-connected separation devices 310. By connecting two adjacent satellites with the plurality of point-connected separation devices 310, a stable connection between the satellites can be achieved, and the locking state of the connection points can be released at corresponding times, enabling the two adjacent satellites to separate quickly and reliably.

[0057] In one embodiment, a single second satellite 200 is connected to the first satellite 100 through four of the point - type connection and separation devices 310, and two adjacent second satellites 200 are connected through four of the point - type connection and separation devices 310. That is, four point - type connection and separation devices 310 can be used to connect between any two adjacent satellites. Among them, each point - type connection and separation device 310 can be arranged on the opposite surfaces of two adjacent satellites.

[0058] In practical applications, the satellite combination structure can include one first satellite 100 and two second satellites 200. The two second satellites 200 are arranged side by side on the same side of the first satellite 100, forming a "pin" - shaped structure with the first satellite 100. Among them, two adjacent satellites are connected through a separation mechanism 300 on their opposite surfaces. This facilitates the layout of the equipment outside the cabin, and the equipment does not interfere with each other.

[0059] Based on the same inventive concept, in another embodiment, a satellite launch control method is also provided. This satellite launch control method can be used to control the launch of the satellite combination structure as described above.

[0060] Refer to Figure 8 , the satellite launch control method provided in this embodiment includes the following steps:

[0061] Step S100: Control the rocket carrying the satellite combination structure to be launched into a transfer orbit.

[0062] Step S300: When receiving the satellite - rocket separation signal, control the satellite combination structure to separate from the rocket. Under the propulsion of the first satellite 100, the satellite combination structure enters the predetermined working orbit.

[0063] Specifically, when the launch vehicle carrying the satellite combination structure is launched into the transfer orbit, if the satellite platform of the first satellite 100 receives the satellite - rocket separation signal sent by the ground operation and control system, the first satellite 100 enters the rate damping mode. After meeting the corresponding conditions, the battery array is deployed, enters the cruise mode, and completes sun - pointing; in the apogee ignition mode, adjust the orbit - changing attitude of the satellite combination structure, and make the satellite combination structure enter the predetermined working orbit through multiple apogee orbit - changes. Figure 5 It is a schematic diagram of the flight state of the satellite combination structure.

[0064] In the process, after the first satellite 100 enters the rate-damped mode, a control torque is applied through the momentum wheel to consume the satellite's rotational kinetic energy, causing the satellite's rotational angular velocity to decrease to below 0.1° / s. At this point, the battery array begins to deploy, entering cruise mode, and using measurement information from star sensors, analog systems, and gyroscopes to achieve solar orientation. The satellite assembly structure in the transfer orbit, under the control of the ground control system, uses apogee ignition mode and jet control to adjust the satellite's orbital attitude, driving the engines to complete multiple apogee maneuvers, thereby enabling the satellite assembly structure to move from the transfer orbit to its predetermined working orbit.

[0065] Step S500: Upon receiving the separation command of the satellite assembly structure, control the assembly composed of each of the second satellites 200 to separate from the first satellite 100.

[0066] When the satellite assembly structure enters its predetermined working orbit, it receives a separation command from the ground control system instructing the satellite assembly structure to separate, and controls the assembly consisting of each second satellite 200 to separate from the first satellite 100. Figure 6 This is a schematic diagram of the first satellite's in-orbit flight status. Figure 6 The payload of the first satellite is shown as 400.

[0067] Step S700: When the first preset condition is met, control the separation between each of the second satellites 200.

[0068] The status of the combined structure consisting of the separated second satellites 200 and the first satellite 100 is monitored in real time. When a first preset condition is met, the individual second satellites 200 in the combined structure can be controlled to separate from each other. Subsequently, the status of each second satellite 200 is continuously monitored. If a second preset condition is met, the battery arrays of each second satellite 200 are deployed, entering cruise mode, and sun orientation is achieved using measurement information from star sensors, analog arrays, and gyroscopes. At this point, the separation of each satellite in the satellite combination structure is completed. Subsequently, according to the planning and control of the ground operation and control system, each satellite can be controlled to complete its on-orbit deployment as required. Figure 7 This is a schematic diagram illustrating the on-orbit flight status of the second satellite. Figure 7 The payload of the second satellite is shown as 400.

[0069] In one embodiment, the first preset condition may include: the distance between the combination of the second satellites 200 and the first satellite 100 reaches a first preset distance. The first preset distance may be set to 100m, 110m, 120m, or 150m, etc., and can be set according to actual needs.

[0070] In one embodiment, the second preset condition may include: the distance between each of the second satellites 200 reaches a second preset distance. The second preset distance may be set to 50m, 60m, 70m, or 80m, and can be set according to actual needs.

[0071] The satellite assembly structure and launch control method provided in this application can avoid the drawbacks of traditional satellite structures and launch layouts, while ensuring satellite functions and on-orbit deployment conditions, saving carrying capacity and not generating new space debris.

[0072] The technical features of the above embodiments can be combined in any way. For the sake of brevity, not all possible combinations of the technical features in the above embodiments are described. However, as long as there is no contradiction in the combination of these technical features, they should be considered to be within the scope of this specification.

[0073] The embodiments described above are merely illustrative of several implementation methods of this application, and while the descriptions are specific and detailed, they should not be construed as limiting the scope of the patent. It should be noted that those skilled in the art can make various modifications and improvements without departing from the concept of this application, and these all fall within the protection scope of this application. Therefore, the protection scope of this application should be determined by the appended claims.

Claims

1. A satellite assembly structure, characterized in that, The satellite combined structure includes a first satellite and at least two second satellites, and both the first satellite and the second satellites are medium and high orbit satellites; each of the second satellites is arranged side by side on the same side of the first satellite, and each of the second satellites is connected to the first satellite through a separation mechanism, and adjacent two of the second satellites are connected through a separation mechanism; The first satellite includes a propulsion system and a first load-bearing structure. The propulsion system includes a main thruster and a fuel tank. The fuel tank is located inside the first load-bearing structure, and the main thruster is connected to the outer side of the bottom of the first load-bearing structure; The first load-bearing structure is of a central load-bearing cylinder configuration, including a load-bearing cylinder, a plurality of first bulkheads, a plurality of first layer plates and a plurality of first outer cabin plates; each of the first bulkheads is connected to the outer peripheral surface of the load-bearing cylinder at intervals along the circumferential direction of the load-bearing cylinder, each of the first layer plates is connected to the outer peripheral surface of the load-bearing cylinder at intervals along the axial direction of the load-bearing cylinder, and each of the first bulkheads and each of the first layer plates cross to form a plurality of independent spaces located outside the load-bearing cylinder. The fuel tank is arranged in some of the independent spaces, and each of the first outer cabin plates is sequentially arranged outside the first load-bearing structure to enclose each of the independent spaces; The second satellite includes a second load-bearing structure. The second load-bearing structure is of a hexahedron configuration, including a second partition board, a plurality of second layer plates, a plurality of second side plates and a plurality of second outer cabin plates; a plurality of the second outer cabin plates are sequentially connected to enclose a hexahedron space, the second partition board is arranged in the hexahedron space, each of the second side plates is perpendicular to the second partition board and is connected to one side surface of the second partition board at intervals, and each of the second layer plates is connected at intervals between adjacent two of the second side plates. The second layer plates are perpendicular to the second side plates and perpendicular to the second partition board.

2. The satellite assembly structure according to claim 1, characterized in that, ​ 3. The satellite assembly structure according to claim 1, characterized in that, ​ 4. The satellite assembly structure according to claim 3, characterized in that, ​ 5. The satellite assembly structure according to claim 1, characterized in that, ​ 6. A satellite launch control method, characterized in that, ​ ​ Upon receiving the satellite-rocket separation signal, the system controls the separation of the satellite assembly from the rocket. Propelled by the first satellite, the satellite assembly enters its predetermined working orbit. This includes: the first satellite entering a rate-damped mode, applying a control torque via a momentum wheel to consume its rotational kinetic energy, causing its rotational angular velocity to decay to below 0.1° / s, after which its battery array deploys, entering cruise mode and completing solar orientation; in apogee ignition mode, the system adjusts the satellite assembly's orbital attitude and performs multiple apogee maneuvers to bring it into the predetermined working orbit. Upon receiving a separation command for the satellite assembly structure, the system controls the assembly of each of the second satellites to separate from the first satellite. Once the first preset condition is met, control the separation of each of the second satellites.

7. The satellite launch control method according to claim 6, characterized in that, The first preset condition includes: the distance between the combination of each of the second satellites and the first satellite reaches a first preset distance.

8. The satellite launch control method according to claim 6, characterized in that, After the step of controlling the separation of each of the second satellites upon meeting the first preset condition, the satellite launch control method further includes: Once the second preset condition is met, the battery arrays of each of the second satellites deploy, enter cruise mode, and complete sun orientation.

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