Satellite combination structure and satellite launching control method

Abstract

The invention relates to the technical field of spaceflight, and particularly 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; the second satellites are arranged on the same side of the first satellite side by side, the second satellites are connected with the first satellite through separation mechanisms, and every two adjacent second satellites are connected through a separation mechanism; the first satellite comprises a propelling system and a first force bearing structure, the propelling system comprises a main propeller and a fuel tank, the fuel tank is located in the first force bearing structure, and the main propeller is connected to the outer side of the bottom of the first force bearing structure. Therefore, complex structures such as an upper stage or a propulsion power cabin do not need to be additionally arranged, under the condition that the satellite function and the on-orbit deployment condition are guaranteed, the carrying capacity is effectively saved, new space junk is avoided, meanwhile, the satellite launching cost is saved, and the system construction cost is reduced.

Description

Technical Field

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

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

[0003] When launching one satellite per rocket, the satellite uses a large-capacity fuel tank to carry sufficient fuel for orbital transfer. When launching two or three satellites per rocket, due to structural and space constraints within the fairing, each satellite cannot use a large-capacity fuel tank to carry the large amount of fuel required for orbital transfer. An upper stage or propulsion module is required to carry the required orbital transfer fuel. After the two or three satellites are placed into their operational orbit, they are separated and deployed. The upper stage or propulsion module is then separated from the satellites and deactivated and deorbited. This launch method has the following drawbacks: first, the upper stage or propulsion module will remain in Earth orbit for a long time, becoming useless waste. Second, the upper stage or propulsion module itself is ineffective "dead weight" during the launch process, squeezing the rocket's carrying capacity and significantly reducing its effective payload capacity. Third, the need for a dedicated upper stage or propulsion module increases the cost of satellite launches and the risks involved in the launch and orbital insertion process. Summary of the Invention

[0004] Based on this, it is necessary to provide a satellite combination structure and a satellite launch control method to address the above problems.

[0005] According to a first aspect of an embodiment of the present application, a satellite assembly structure is provided, comprising a first satellite and at least two second satellites; each of the second satellites is arranged side by side on the same side of the first satellite, each of the second satellites is connected to the first satellite via a separation mechanism, and two adjacent second satellites are connected via 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 outside of the bottom of the first load-bearing structure.

[0006] 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 layer plates, and a plurality of first outer cabin plates; Each of the first bulkheads is circumferentially spaced and connected to the outer peripheral surface of the load-bearing cylinder, and each of the first layer plates is axially spaced and connected to the outer peripheral surface of the load-bearing cylinder. Each of the first bulkheads and each of the first layer plates cross to form a number of independent spaces around the load-bearing cylinder, and the fuel tank is disposed in some of the independent spaces. Each of the first outer cabin plates is sequentially disposed around the first load-bearing structure to enclose each of the independent spaces.

[0007] In one embodiment, the second satellite includes a second load-bearing structure, and the second load-bearing structure is in a hexahedron configuration, including a second partition board, a number of second layer plates, a number of second side plates, and a number of second outer cabin plates; A number of the second outer cabin plates are sequentially connected and enclosed to form a hexahedron space. The second partition board is disposed in the hexahedron space. Each of the second side plates is perpendicular to the second partition board and is spaced and connected to one surface of the second partition board. Each of the second layer plates is spaced and connected between two adjacent second side plates. The second layer plate is perpendicular to the second side plate and perpendicular to the second partition board.

[0008] In one embodiment, the main thruster includes an engine with a thrust of 750N.

[0009] In one embodiment, the separation mechanism includes a number of point-type connection separation devices.

[0010] In one embodiment, a single second satellite and the first satellite are connected by four of the point-type connection separation devices, and two adjacent second satellites are connected by four of the point-type connection separation devices.

[0011] 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.

[0012] According to the second aspect of the embodiments of the present application, a satellite launch control method is provided for performing launch control on the satellite combined structure as described above. The satellite launch control method includes: Controlling a rocket carrying the satellite combined structure to be launched into a transfer orbit; When receiving a satellite-rocket separation signal, controlling the satellite combined structure to separate from the rocket. Under the propulsion of the first satellite, the satellite combined structure enters a predetermined working orbit; When receiving a separation instruction of the satellite combined structure, controlling the combined body formed by the second satellites to separate from the first satellite; When a first preset condition is satisfied, controlling the second satellites to separate from each other.

[0013] In one embodiment, the first preset condition includes: the distance between the combination of the second satellites and the first satellite reaches a first preset distance.

[0014] In one embodiment, upon receiving the satellite-rocket separation signal, controlling the satellite assembly structure to separate from the rocket, and the satellite assembly structure to enter a predetermined operating orbit under the propulsion of the first satellite, comprises: The first satellite enters the rate damping mode, and after meeting the corresponding conditions, the battery array is deployed, the satellite enters the cruise mode, and completes the solar orientation; in the apogee ignition mode, the orbit change attitude of the satellite assembly structure is adjusted, and the satellite assembly structure enters the predetermined working orbit through multiple apogee orbit changes.

[0015] After the step of controlling the separation of the second satellites when the first preset condition is met, the satellite launch control method further includes: When the second preset condition is met, the battery array of each of the second satellites is deployed, enters the cruise mode, and completes the solar orientation.

[0016] In the satellite combination structure provided in the embodiment of the present application, the first satellite serves as a 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. Two adjacent second satellites are also connected via a separation structure, that is, the connection and separation of each satellite can be achieved through each separation mechanism. In the present application, the first satellite serving as a carrier satellite can inherit the design concept of a traditional satellite, that is, the satellite combination structure can be propelled from the transfer orbit to the target orbit by a main thruster connected to the outside of the bottom of the first load-bearing structure and a fuel tank carrying the fuel required for orbit transfer arranged inside the first load-bearing structure. There is no need to add complex structures such as an upper stage or a propulsion power cabin. While ensuring the satellite function and on-orbit deployment conditions, the carrying capacity is effectively saved, the generation of new space debris is avoided, and the satellite launch cost is saved, thereby reducing the system construction cost. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 A schematic diagram of the structure of a satellite assembly provided in one embodiment of the present application; Figure 2 A schematic diagram of the structure of a satellite assembly provided in one embodiment of the present application; Figure 3 A schematic structural diagram of the first satellite in the satellite assembly structure provided in one embodiment of the present application; Figure 4A schematic structural diagram of a second load-bearing structure of a second satellite in a satellite assembly structure provided in one embodiment of the present application; Figure 5 A schematic diagram of the flight state of a satellite assembly structure provided in one embodiment of the present application; Figure 6 A schematic diagram of the in-orbit flight state of the first satellite in the satellite assembly structure provided in one embodiment of the present application; Figure 7 A schematic diagram of the in-orbit flight state of the second satellite in the satellite assembly structure provided in one embodiment of the present application; Figure 8 This is a flowchart of a satellite launch control method provided in one embodiment of the present application.

[0018] Description of reference numerals: 100. First satellite; 111. Main thruster; 112. Fuel tank; 121. Load-bearing cylinder; 122. First bulkhead; 123. First deck; 124. First outer cabin; 200. Second satellite; 211. Second bulkhead; 212. Second deck; 213. Second side panel; 214. Second outer cabin; 300. Separation mechanism; 310. Point-type connection and separation device; 400. Payload. DETAILED DESCRIPTION

[0019] To facilitate understanding of the present application, a more comprehensive description of the present application will be provided below with reference to the accompanying drawings. The accompanying drawings illustrate preferred embodiments of the present application. However, the present application may 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 comprehensive understanding of the disclosure of the present application.

[0020] In this application, unless otherwise specified or limited, the terms "installed," "connected," "connect," "fixed," etc. should be understood in a broad sense. For example, they can refer to fixed connection, detachable connection, or integration; mechanical connection or electrical connection; direct connection or indirect connection through an intermediate medium; internal communication between two elements or interaction between two elements, unless otherwise specified. Those skilled in the art will understand the specific meanings of the above terms in this application based on specific circumstances.

[0021] 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 the technical features being referred to. Thus, a feature specified as "first" or "second" may explicitly or implicitly include at least one of such features. Throughout the description of this application, "plurality" means at least two, for example, two, three, etc., unless otherwise specifically defined.

[0022] Unless otherwise defined, all technical and scientific terms used herein have the same meaning as commonly understood by those skilled in the art to which this application pertains. The terms used herein in the specification of this application are for the purpose of describing specific embodiments only and are not intended to limit this application. The term "and / or" as used herein includes any and all combinations of one or more of the associated listed items.

[0023] In the field of spacecraft launch technology, in order to improve launch efficiency, the use of multiple satellite launches with one rocket is becoming increasingly popular, that is, launching more satellites into artificial earth satellite orbits at one time. When the rocket's carrying capacity and satellite orbit are determined, the number of satellites launched at one time is related to the mass of the launched satellites, the satellite structure, the separation device, the space inside the fairing, the layout of the satellites inside the fairing, the separation method, etc. At present, the launch of multiple satellites with one rocket for small low-orbit satellites generally uses a central load-bearing bucket or auxiliary support structure to fix the launched satellites, and separate them one by one after entering orbit. In addition, a stacked launch method can also be used, but this method is only suitable for the launch of low-orbit communication satellites of the same model that can be easily designed into a flat-plate form. For medium and large satellites with complex functions and special payloads in medium and high orbits, the above methods are difficult to apply.

[0024] For medium and high-orbit satellites, due to their powerful functions, large size and heavy mass, they are mostly launched in a one-satellite-per-rocket, two-satellite-per-rocket or three-satellite-per-rocket manner. At the same time, due to the carrying capacity of the rocket, the satellite is generally launched into a synchronous transfer orbit (GTO) or a medium transfer orbit (MTO), and then the satellite or the upper stage carrying the satellite performs orbit transfer, so that the satellite finally enters the geostationary orbit (GEO) or medium earth orbit (MEO).

[0025] When launching one satellite per rocket, the satellite uses a large-capacity fuel tank to carry sufficient fuel for orbital transfer. When launching two or three satellites per rocket, due to structural and space constraints within the fairing, each satellite cannot use a large-capacity fuel tank to carry the large amount of fuel required for orbital transfer. An upper stage or propulsion module is required to carry the required orbital transfer fuel. After the two or three satellites are placed into their operational orbit, they are separated and deployed. The upper stage or propulsion module is then separated from the satellites and deactivated and deorbited. This launch method has the following drawbacks: first, the upper stage or propulsion module will remain in Earth orbit for a long time, becoming useless waste. Second, the upper stage or propulsion module itself is ineffective "dead weight" during the launch process, squeezing the rocket's carrying capacity and significantly reducing its effective payload capacity. Third, the need for a dedicated upper stage or propulsion module increases the cost of satellite launches and the risks involved in the launch and orbital insertion process.

[0026] In response to the above problems, the present application provides a satellite combination structure and a satellite launch control method.

[0027] In one embodiment, a satellite assembly structure is provided. Each satellite in the satellite assembly structure can be a medium-orbit satellite. The design of the satellite assembly structure 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 conserves carrying capacity, avoids the generation of new space debris, and simultaneously saves satellite launch costs and reduces system construction costs.

[0028] 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. Two adjacent second satellites 200 are also connected via a separation mechanism 300. 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 outside of the bottom of the first load-bearing structure.

[0029] In the satellite assembly structure provided in this embodiment, a first satellite 100 serves as a carrier satellite and can carry 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. Two adjacent second satellites 200 are also connected via a separation structure. That is, each satellite can be connected and separated by each separation mechanism 300. In this application, the first satellite 100, serving as a 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 111 connected to the outer bottom of the first bearing structure and a fuel tank 112 containing the fuel required for orbit transfer, which is arranged inside the first bearing structure. There is no need to add complex structures such as an upper stage or a propulsion power cabin. While ensuring satellite functions and on-orbit deployment conditions, it effectively saves carrying capacity, avoids the generation of new space debris, and saves satellite launch costs, reducing system construction costs.

[0030] In the satellite assembly structure, the first satellite 100 serves as the carrier satellite, and each second satellite 200 serves as the hosted satellite. Each second satellite 200 is carried by the first satellite 100. During launch and orbit transfer, the first satellite 100 and each second satellite 200 are connected to form a single structure using a separation mechanism 300. The fuel tank 112 within the first satellite 100 carries the fuel required for orbit transfer. The propulsion system of the first satellite 100 propels the entire satellite assembly from the transfer orbit into the target orbit. The satellites are then separated and deployed, completing their orbital insertion. In practical applications, the main thruster 111 of the first satellite 100 may include a 750N engine.

[0031] Reference Figure 3 In one embodiment, the first load-bearing structure is a central load-bearing cylinder 121 configuration, including the load-bearing cylinder 121, a plurality of first bulkheads 122, a plurality of first layer plates 123 and a plurality of first outer cabin plates 124.

[0032] Among them, each first bulkhead 122 is connected to the outer peripheral surface of the bearing cylinder 121 at intervals along the circumference of the bearing cylinder 121, and each first layer plate 123 is connected to the outer peripheral surface of the bearing cylinder 121 at intervals along the axial direction of the bearing cylinder 121. Each first bulkhead 122 and each first layer plate 123 intersect to form a plurality of independent spaces located outside the bearing cylinder 121. The fuel tank 112 is arranged in part of the independent spaces, and each first outer cabin plate 124 is sequentially arranged on the periphery of the first bearing structure to enclose each independent space.

[0033] The first load-bearing structure of the central load-bearing cylinder 121 configuration is formed by a combination of a load-bearing cylinder 121, a plurality of first bulkheads 122, a plurality of first layer plates 123 and a plurality of first outer cabin plates 124, which can effectively increase the overall strength of the satellite assembly structure, and the satellite height is moderate and the modal response is good.

[0034] The interface connecting the first satellite 100 to the launch vehicle can utilize a conventional, relatively mature, strap-on 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 is securely connected to the launch vehicle via the strap-on separation mechanism 300.

[0035] 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 panels 212, a plurality of second side panels 213 and a plurality of second outer cabin panels 214.

[0036] Among them, a plurality of the second outer cabin plates 214 are sequentially connected and enclosed to form a hexahedron space. The second partition plate 211 is disposed within the hexahedron space. Each of the second side plates 213 is perpendicular to the second partition plate 211 and is spaced and connected to one side surface of the second partition plate 211. Each of the second layer plates 212 is spaced and connected between two adjacent second side plates 213. The second layer plate 212 is perpendicular to the second side plate 213 and perpendicular to the second partition plate 211.

[0037] The second force-bearing structure of the above-mentioned hexahedron configuration is constituted by the second partition plate 211, a plurality of second layer plates 212, a plurality of second side plates 213, and a plurality of second outer cabin plates 214, which can provide torsional stiffness and longitudinal shear stiffness, facilitating the layout of equipment within the configuration of the second satellite 200 and the installation of equipment inside and outside the cabin.

[0038] Refer to Figure 6 and Figure 7 In one embodiment, the separation mechanism 300 includes a plurality of point-type connection and separation devices 310. By connecting adjacent satellites through a plurality of point-type connection and separation devices 310, stable connection between satellites can be achieved, and at the corresponding moment, the locking state of the connection points can be released, enabling adjacent satellites to be separated quickly and reliably.

[0039] 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 adjacent satellites. Among them, each point-type connection and separation device 310 can be disposed on the opposite surfaces of adjacent satellites.

[0040] In practical applications, the satellite combination structure may 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, adjacent satellites are connected through the separation mechanism 300 on their opposite surfaces. This facilitates the layout of equipment outside the cabin, and the equipment does not interfere with each other.

[0041] Based on the same inventive concept, in another embodiment, a satellite launch control method is further provided. This satellite launch control method can be used for launching and controlling the satellite combination structure as described above.

[0042] <000012> Figure 8 Referring to Figure 8 the satellite launch control method provided in this embodiment includes the following steps: Step S100: Control the launch of a rocket carrying the satellite assembly structure into a transfer orbit.

[0043] Step S300: After receiving the satellite-rocket separation signal, the satellite assembly structure is controlled to separate from the rocket. Under the propulsion of the first satellite 100, the satellite assembly structure enters a predetermined working orbit.

[0044] Specifically, after the carrier rocket carries the satellite assembly structure 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, and after the corresponding conditions are met, the battery array is deployed, the system enters the cruise mode, and completes the solar orientation; in the apogee ignition mode, the orbit change attitude of the satellite assembly structure is adjusted, and the satellite assembly structure is put into the predetermined working orbit through multiple apogee orbit changes. Figure 5 Schematic diagram of the flight status of the satellite assembly structure.

[0045] After the first satellite 100 enters rate damping mode, a control torque is applied via the momentum wheels, dissipating the satellite's rotational kinetic energy and reducing its angular velocity to below 0.1° / s. At this point, the battery array deploys, and the system enters cruise mode, using measurements from the star sensor, module, and gyroscope to achieve solar orientation. The satellite assembly in transfer orbit, under the control of the ground-based operation and control system, operates in apogee ignition mode, adjusting the satellite's orbital attitude through jet control and driving the engine to complete multiple apogee orbit changes, thereby moving the satellite assembly from the transfer orbit into the intended operating orbit.

[0046] Step S500: After receiving the separation instruction of the satellite assembly structure, the assembly formed by the second satellites 200 is controlled to separate from the first satellite 100.

[0047] When the satellite assembly structure enters the predetermined working orbit and receives a separation instruction from the ground operation and control system to instruct the satellite assembly structure to separate, the assembly consisting of the second satellites 200 is controlled 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 A payload 400 of a first satellite is shown.

[0048] Step S700: When a first preset condition is met, the second satellites 200 are controlled to separate from each other.

[0049] The status of the assembly of separated second satellites 200 and first satellite 100 is monitored in real time. When a first pre-set condition is met, the second satellites 200 in the assembly are controlled to separate from each other. Subsequently, the status of each second satellite 200 is continuously monitored. If a second pre-set condition is met, the battery array of each second satellite 200 is controlled to deploy, entering cruise mode, and using star sensor, module, and gyroscopic measurements to achieve solar orientation. At this point, the separation of the satellites in the satellite assembly is complete. Subsequently, the satellites can be individually controlled to complete their on-orbit deployment as required, based on the planning and control of the ground operation and control system. Figure 7 This is a schematic diagram of the second satellite's in-orbit flight status. Figure 7 A payload 400 of a second satellite is shown.

[0050] 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 may be set according to actual needs.

[0051] In one embodiment, the second preset condition may include: the distance between the second satellites 200 reaches a second preset distance. The second preset distance may be set to 50m, 60m, 70m, or 80m, which may be set according to actual needs.

[0052] The satellite combination structure and launch control method thereof provided in the embodiments of the present application can avoid the disadvantages brought about by traditional satellite structures and launch layouts, save carrying capacity while ensuring satellite functions and on-orbit deployment conditions, and avoid generating new space debris.

[0053] The technical features of the above-mentioned embodiments can be combined arbitrarily. In order to make the description concise, not all possible combinations of the technical features in the above-mentioned 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.

[0054] The above-described embodiments merely represent several implementation methods of the present application. While the descriptions are relatively specific and detailed, they should not be construed as limiting the scope of the patent. It should be noted that a person of ordinary skill in the art may make various modifications and improvements without departing from the spirit of the present application, and such modifications and improvements are all within the scope of protection of the present application. Therefore, the scope of protection of the present application shall 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; the second satellites are arranged side by side on the same side of the first satellite, and each second satellite is connected to the first satellite through a separation mechanism, and adjacent second satellites are connected to each other 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.

2. The satellite assembly structure according to claim 1, characterized in that: 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 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.

3. The satellite assembly structure according to claim 1, characterized in that: 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. Each of the second layer plates is connected at intervals between adjacent second side plates. The second layer plates are perpendicular to the second side plates and perpendicular to the second partition board.

4. The satellite assembly structure according to claim 1, characterized in that: The main thruster includes an engine of 750N.

5. The satellite assembly structure according to claim 1, characterized in that: The separation mechanism includes a plurality of point-type connection and separation devices.

6. The satellite assembly structure according to claim 5, characterized in that: A single second satellite is connected to the first satellite through four of the point-type connection and separation devices, and adjacent second satellites are connected to each other through four of the point-type connection and separation devices.

7. The satellite assembly structure according to claim 1, characterized in that: 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" character structure with the first satellite.

8. A satellite launch control method, characterized in that: For launching and controlling the satellite combined structure according to any one of claims 1-7, the satellite launch control method includes: Controlling a rocket carrying the satellite combined structure to be launched into a transfer orbit; When receiving a satellite-rocket separation signal, controlling the satellite combined structure to separate from the rocket. Under the propulsion of the first satellite, the satellite combined structure enters a predetermined working orbit; When receiving a separation instruction of the satellite combined structure, controlling the combined body formed by the second satellites to separate from the first satellite; When a first preset condition is satisfied, controlling the second satellites to separate from each other.

9. The satellite launch control method according to claim 8, characterized in that: The first preset condition includes: the distance between the combined body formed by the second satellites and the first satellite reaches a first preset distance.

10. The satellite launch control method according to claim 8, characterized in that: The step of controlling the satellite assembly structure to separate from the rocket upon receiving the satellite-rocket separation signal, and the satellite assembly structure to enter a predetermined operating orbit under the propulsion of the first satellite, comprises: The first satellite is controlled to enter the rate damping mode. When the corresponding conditions are met, the battery array is deployed, the satellite enters the cruise mode, and the solar orientation is completed. In the apogee ignition mode, the orbit change attitude of the satellite assembly structure is adjusted, and the satellite assembly structure is put into the predetermined working orbit through multiple apogee orbit changes. After the step of controlling the separation of the second satellites when the first preset condition is met, the satellite launch control method further includes: When the second preset condition is met, the battery array of each of the second satellites is deployed, enters the cruise mode, and completes the solar orientation.

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