A launch vehicle and satellite separation section variable target orbit control method and system thereof

Abstract

The application discloses a launch vehicle and satellite separation section variable target orbit control method and a system thereof, wherein the launch vehicle and satellite separation section variable target orbit control method comprises the following steps: determining a required attitude for changing an orbit according to a variable orbit requirement; adjusting an attitude of a rocket body according to the required attitude; after the attitude adjustment of the rocket body is completed, stabilizing for a specified time, and after the stabilization is completed, changing the orbit until the target orbit is reached; and after the target orbit is reached, closing a terminal correction power system, and completing target orbit adjustment. The application can calculate the variable orbit in real time according to target orbit parameters and current flight state parameters, is simple in design, and is convenient to realize and has practicality.

Description

Technical Field

[0001] This application relates to the aerospace field, specifically to a method and system for controlling the variable target orbit during the separation phase of a launch vehicle. Background Technology

[0002] To address the booming commercial satellite launch service market and effectively utilize the payload capacity of national satellite launch missions, launch vehicles generally adopt a multi-satellite launch configuration. The application of shared rocket launch models further enhances the mission adaptability of launch vehicles and reduces the launch cost per satellite. This model has been adopted in international launch service projects and possesses strong international competitiveness. In recent years, with the rapid development of international commercial spaceflight, major launch vehicles worldwide have launched cost-effective multi-satellite launch modes and related products. SpaceX, Rocket Lab, Arianespace, and Roscosmos have completed numerous multi-satellite launches using Falcon 9, Electro, Vega, and Soyuz-2 rockets, respectively, intensifying market competition. China's Long March 2C, Long March 6, Long March 8, and Lijian-1 launch vehicles have also launched various multi-satellite launch configurations to meet commercial launch demands. Currently, to reduce satellite launch costs, launch vehicles, when capacity is sufficient, aim to launch payloads as close to their carrying capacity as possible. When satellites lack the ability to alter their orbits or have limited ability to do so, satellites launched by the same launch vehicle must share the same orbit. Therefore, there is a need for rocket-to-satellite launches, and when satellites in the same orbit are launched together, there are generally constraints such as time and installation methods, which reduces the frequency of rocket launches.

[0003] Therefore, how to provide a method for controlling the variable target orbit during the satellite-rocket separation stage of a launch vehicle to reduce the requirements for orbital consistency of satellites in the same launch, further shorten the launch cycle of launch vehicles, and reduce the launch capacity loss of launch vehicles has become an urgent problem to be solved in this field. Summary of the Invention

[0004] To address the aforementioned issues, this application proposes a target trajectory control method for the separation stage of a launch vehicle, comprising the following steps: determining the required attitude for trajectory change based on the trajectory change requirements; adjusting the rocket body attitude according to the required attitude; stabilizing the rocket body attitude within a specified time after the attitude adjustment is completed; performing trajectory change after stabilization until reaching the target trajectory; and shutting down the final propulsion system after reaching the target trajectory to complete the target trajectory adjustment.

[0005] The target trajectory control method for the separation stage of a launch vehicle described above, wherein determining the required attitude for trajectory change according to the trajectory change requirements includes calculating the target trajectory inclination increment based on the current flight parameters and the target trajectory inclination, and calculating the attitude adjustment target program angle based on the target trajectory inclination increment.

[0006] The above-described method for controlling the variable target orbit during the launch vehicle's satellite-rocket separation phase includes a variable target orbit inclination angle and a velocity increment. Represented as: Where T represents transpose, ΔV x ΔV represents the component of the acceleration on the X-axis. y ΔV represents the component of the acceleration on the Y-axis. z This represents the component of the acceleration to be increased along the Z-axis.

[0007] The above-described method for controlling the variable target orbit during the separation phase of a launch vehicle and its satellite includes a target program angle. Represented as:

[0008]

[0009] The target trajectory control method for the separation stage of a launch vehicle, as described above, includes the following sub-steps: after the rocket body attitude adjustment is completed, it stabilizes within a specified time. After stabilization, the trajectory is changed until the target trajectory is reached: real-time tracking of the closed-circuit guidance program angle; and axis determination.

[0010] A target trajectory change control system for the separation stage of a launch vehicle includes: an attitude determination unit for trajectory change, a rocket body attitude adjustment unit, a trajectory change unit, and a final repair system shutdown unit. The attitude determination unit is used to determine the attitude required for trajectory change based on the trajectory change requirements. The rocket body attitude adjustment unit is used to adjust the rocket body attitude according to the required attitude. The trajectory change unit is used to stabilize the rocket body attitude within a specified time after the attitude adjustment is completed, and then perform trajectory change until the target trajectory is reached. The final repair system shutdown unit is used to shut down the final repair propulsion system after reaching the target trajectory, thus completing the target trajectory adjustment.

[0011] As described above, in the launch vehicle satellite-rocket separation stage target trajectory control system, the attitude determination unit for trajectory change determines the attitude required for trajectory change according to the trajectory change requirements, including calculating the target trajectory inclination increment based on the current flight parameters and the target trajectory inclination, and calculating the attitude adjustment target program angle based on the target trajectory inclination increment.

[0012] As described above, in the launch vehicle satellite-rocket separation stage variable target orbit control system, the variable target orbit inclination angle and velocity increment are determined in the attitude determination unit required for orbit change. Represented as: Where T represents transpose, ΔV xΔV represents the component of the acceleration on the X-axis. y ΔV represents the component of the acceleration on the Y-axis. z This represents the component of the acceleration to be increased along the Z-axis.

[0013] As described above, in the launch vehicle satellite-rocket separation stage variable target orbit control system, the target program angle in the attitude determination unit required for orbit change... Represented as:

[0014]

[0015] The target trajectory control system for the launch vehicle satellite-rocket separation stage, as described above, includes the following sub-steps: the trajectory change unit stabilizes within a specified time, and after stabilization, the trajectory change is performed until the target trajectory is reached: real-time tracking of the closed-loop guidance program angle; and axis determination.

[0016] This application has the following beneficial effects:

[0017] (1) This application can calculate the orbit change in real time according to the target orbit parameters and the current flight status parameters. It is simple in design, easy to implement and practical.

[0018] (2) In this application, quaternions are used for attitude control during the orbit change process to avoid the occurrence of singularity in attitude under specific Euler angles and to ensure the reliability of the algorithm.

[0019] (3) The launch vehicle involved in this application has the ability to change orbits, which can reduce the requirement for the consistency of satellite orbits in the same launch, further shorten the launch cycle of the launch vehicle, and reduce the launch capacity loss of the launch vehicle. Attached Figure Description

[0020] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in this application. For those skilled in the art, other drawings can be obtained based on these drawings.

[0021] Figure 1 This is a flowchart illustrating the target trajectory control method for the launch vehicle satellite-rocket separation stage according to an embodiment of this application.

[0022] Figure 2 This is a schematic diagram of the internal structure of the target trajectory control system for the separation stage of a launch vehicle provided in an embodiment of this application. Detailed Implementation

[0023] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application.

[0024] Example 1

[0025] like Figure 1 As shown, this embodiment provides a method for controlling the changing target orbit during the satellite-rocket separation phase of a launch vehicle. Specifically, it is executed during the satellite-rocket separation phase and includes the following steps:

[0026] Step S1: Determine the attitude required for orbit change based on the orbit change requirements.

[0027] The attitude adjustment segment calculates the target orbit inclination to be increased velocity based on the current flight parameters and the target orbit inclination, and then calculates the attitude adjustment target program angle.

[0028] Among them, the target trajectory inclination angle needs to be increased speed Represented as:

[0029]

[0030] Where T represents transpose, ΔV x ΔV represents the component of the acceleration on the X-axis. y ΔV represents the component of the acceleration on the Y-axis. z This represents the component of the acceleration to be increased along the Z-axis.

[0031] Based on the variable target trajectory inclination angle, the required acceleration Calculate the target angle of salary adjustment procedure in Indicates the target pitch program angle. This indicates the target yaw program angle. The target roll program angle remains unchanged.

[0032] The target program angle is represented as:

[0033]

[0034] Based on the initially calculated closed-circuit guidance program angle and attitude adjustment initial angle (i.e., the program angle at the start of attitude adjustment), the program angle is calculated using triangular wave attitude adjustment.

[0035] Understandably, triangular wave attitude adjustment is an existing attitude adjustment method, and the specific principles will not be elaborated here.

[0036] The above-described method of calculating orbit changes in real time based on target orbit parameters and current flight status parameters is simple in design, easy to implement, and practical.

[0037] Step S2: Adjust the arrow body posture according to the required posture.

[0038] When adjusting the rocket's attitude, the final repair power system is not activated. The rocket body is rotated only by the attitude control system so that the thrust direction of the final repair nozzle is precisely aligned with the ΔV direction required for the trajectory change. For example, when changing the inclination angle, the thrust direction must be perpendicular to the original track plane.

[0039] Step S3: After the rocket body attitude is adjusted, it is stabilized within a specified time. After stabilization, the trajectory is changed until the target trajectory is reached.

[0040] Step S3 includes the following sub-steps:

[0041] Step S31: Real-time tracking of closed-circuit guidance procedure angle.

[0042] The stabilization phase before the final repair begins starts with the initial attitude adjustment angle, and the closed-circuit guidance iterative calculation program angle is tracked in real time. Maintain thrust direction. Within the settling time... Then it is moved to the final repair section to avoid attitude instability causing the thrust direction to deviate and affecting the trajectory change accuracy.

[0043] Understandable. This refers to the period of stability before the final stage of cultivation begins.

[0044] Step S32: Perform axis determination.

[0045] After the final repair work section is started, the closed-circuit guidance is iteratively calculated in real time to maintain the program angle. The program angle is maintained after the axis determination criterion is met.

[0046] Specifically, as the trajectory parameters gradually approach the target value, it is necessary to confirm that the rocket's attitude has been stably aligned with the target trajectory direction through axis determination.

[0047] Fixed-axis discrimination includes real-time calculation of the parameters of the orbital inclination and the target orbital inclination. If the absolute value of the deviation is less than the first guidance constant ΔorbI (typically designed to be 0.01°), then axis fixing is performed. The deviation is repeatedly calculated at a fixed discrimination period T1 to ensure the real-time nature of the criterion judgment.

[0048] Once the axis-fixing criterion is met, the attitude control system switches from the tracking iteration program angle mode to the program angle holding mode, locking the current attitude.

[0049] Step S4: After reaching the target track, shut down the final repair power system to complete the target track adjustment.

[0050] To ensure that the track parameters accurately reach the target value and avoid over-adjustment or fuel waste, a dual criterion is set for the final maintenance power system: normal shutdown and time-based backup shutdown. If either criterion is met, the final maintenance nozzle will be shut down.

[0051] Normal shutdown includes continuing to monitor the orbital inclination and target orbital inclination parameters. The absolute value of the deviation is determined. When the absolute value of the deviation is less than the second guidance constant ΔorbI' (typical design value is 0.001°, which corresponds to approximately 20 control cycle advances), the final-stage nozzle is shut off. The discrimination period is consistent with the fixed-axis discrimination period T1 to ensure a rapid response when the deviation reaches the threshold.

[0052] The time-based backup shutdown includes real-time statistics of the operating time of the final maintenance power system. If the operating time of the final maintenance power system exceeds the preset maximum operating time for the final maintenance of the track-changing section ( When the engine is in operation, the final repair nozzle is shut down. This assumes that the final repair thrust is sufficient to adjust the trajectory to the target range within the maximum operating time. To avoid fuel exhaustion or attitude loss, a forced shutdown is required.

[0053] After the final nozzle is closed, the system enters the stabilization phase, maintaining the current program angle, while meeting the total trajectory change duration. After that, the orbital change operation ends, and the next satellite-rocket separation phase begins.

[0054] Furthermore, incremental attitude adjustment is not recommended when designing the satellite-launch separation attitude trajectory after the target orbit inclination has been adjusted. During the final orbital change phase, the orbital parameters (such as inclination and semi-major axis) have been precisely adjusted to the target values ​​by iteratively calculating the program angle and supplementing the velocity to be increased during the final orbital change. If incremental attitude adjustment (adjusting the attitude through multiple small thrusts) is used at this point, it will directly introduce additional velocity increments, causing the orbital state to deviate from the target.

[0055] During the next satellite-rocket separation phase, the determined final orbit change acceleration will be determined as follows: the transformation matrix between the launch inertial system and the orbital coordinate system will be determined, and the acceleration to be increased in the launch inertial system will be determined based on the transformation matrix.

[0056] The transformation matrix between the launch inertial frame and the orbital coordinate system. Represented as:

[0057]

[0058] in A unit vector representing the position of the launching inertial frame. Indicates the geocentric deviation of the launch inertial frame, r d This indicates the magnitude of the rocket's displacement from the Earth's center in its launch inertial frame. The unit vector representing the velocity of the launching inertial frame. V represents the velocity vector of the launching inertial frame. d This represents the velocity magnitude of the launching inertial frame.

[0059] Launch inertial frame to be increased velocity Represented as:

[0060]

[0061]

[0062] Where ΔV a Δa represents the velocity to be increased along the semi-major axis of the track, Δi represents the deviation of the semi-major axis of the track, and ΔV represents the deviation of the track inclination angle. i This represents the velocity to be increased in the direction of the orbital inclination deviation, and Fm represents the Earth's gravitational constant. The current target parameters are represented by the semi-major axis at the end of the stable phase of the satellite-launch separation before the orbit change; a k This is represented as the semi-major axis of the track, calculated in real time based on the current point information; Indicates the parameters, namely the inclination angle of the target orbit for the orbit change; This indicates the orbital inclination angle calculated in real time based on the current point information; e k This represents the eccentricity calculated in real time based on the current point information; u k This indicates the latitude argument calculated in real time based on the current point information.

[0063] Example 2

[0064] like Figure 2 As shown in the embodiment of this application, a target trajectory change control system for the separation stage of a launch vehicle is provided, which specifically includes: an attitude determination unit 210 for trajectory change, a rocket body attitude adjustment unit 220, a trajectory change unit 230, and a final repair system shutdown unit 240.

[0065] The attitude determination unit 210 for orbit change is used to determine the attitude required for orbit change based on the orbit change requirements.

[0066] The attitude adjustment segment calculates the target orbit inclination to be increased velocity based on the current flight parameters and the target orbit inclination, and then calculates the attitude adjustment target program angle.

[0067] Among them, the target trajectory inclination angle needs to be increased speed Represented as:

[0068]

[0069] Where T represents transpose, ΔV x ΔV represents the component of the acceleration on the X-axis. yΔV represents the component of the acceleration on the Y-axis. z This represents the component of the acceleration to be increased along the Z-axis.

[0070] Based on the variable target trajectory inclination angle, the required acceleration Calculate the target angle of salary adjustment procedure in Indicates the target pitch program angle. This indicates the target yaw program angle. The target roll program angle remains unchanged.

[0071] The target program angle is represented as:

[0072]

[0073] Based on the initially calculated closed-circuit guidance program angle and attitude adjustment initial angle, the program angle is calculated using triangular wave attitude adjustment.

[0074] The above-described method of calculating orbit changes in real time based on target orbit parameters and current flight status parameters is simple in design, easy to implement, and practical.

[0075] Arrow body attitude adjustment unit 220 is used to adjust the arrow body attitude according to the required attitude.

[0076] When adjusting the rocket's attitude, the final repair power system is not activated. The rocket body is rotated only by the attitude control system so that the thrust direction of the final repair nozzle is precisely aligned with the ΔV direction required for the trajectory change. For example, when changing the inclination angle, the thrust direction must be perpendicular to the original track plane.

[0077] The trajectory change unit 230 is used to stabilize the rocket body after the attitude adjustment is completed within a specified time. After stabilization, the final power system performs trajectory change until the target trajectory is reached.

[0078] The trajectory-changing unit 230 performs the following sub-steps:

[0079] Step Q1: Real-time tracking of closed-circuit guidance procedure angle.

[0080] The stabilization phase before the final repair begins starts with the initial attitude adjustment angle, and the closed-circuit guidance iterative calculation program angle is tracked in real time. Maintain thrust direction. Within the settling time... Then it is moved to the final repair section to avoid attitude instability causing the thrust direction to deviate and affecting the trajectory change accuracy.

[0081] Understandable. This refers to the period of stability before the final stage of cultivation begins.

[0082] Step Q2: Perform axis determination.

[0083] After the final repair work section is started, the closed-circuit guidance is iteratively calculated in real time to maintain the program angle. The program angle is maintained after the axis determination criterion is met.

[0084] Specifically, as the trajectory parameters gradually approach the target value, it is necessary to confirm that the rocket's attitude has been stably aligned with the target trajectory direction through axis determination.

[0085] Fixed-axis discrimination includes real-time calculation of the parameters of the orbital inclination and the target orbital inclination. If the absolute value of the deviation is less than the first guidance constant ΔorbI (typically designed to be 0.01°), then axis fixing is performed. The deviation is repeatedly calculated at a fixed discrimination period T1 to ensure the real-time nature of the criterion judgment.

[0086] Once the axis-fixing criterion is met, the attitude control system switches from the tracking iteration program angle mode to the program angle holding mode, locking the current attitude.

[0087] The final repair system shutdown unit 240 is used to shut down the final repair power system after reaching the target track, thus completing the target track adjustment.

[0088] To ensure that the track parameters accurately reach the target value and avoid over-adjustment or fuel waste, a dual criterion is set for the final maintenance power system: normal shutdown and time-based backup shutdown. If either criterion is met, the final maintenance nozzle will be shut down.

[0089] Normal shutdown includes continuing to monitor the orbital inclination and target orbital inclination parameters. The absolute value of the deviation is determined. When the absolute value of the deviation is less than the second guidance constant ΔorbI' (typical design value is 0.001°, which corresponds to approximately 20 control cycle advances), the final-stage nozzle is shut off. The discrimination period is consistent with the fixed-axis discrimination period T1 to ensure a rapid response when the deviation reaches the threshold.

[0090] The time-based backup shutdown includes real-time statistics of the operating time of the final maintenance power system. If the operating time of the final maintenance power system exceeds the preset maximum operating time for the final maintenance of the track-changing section ( When the engine is in operation, the final repair nozzle is shut down. This assumes that the final repair thrust is sufficient to adjust the trajectory to the target range within the maximum operating time. To avoid fuel exhaustion or attitude loss, a forced shutdown is required.

[0091] After the final nozzle is closed, the system enters the stabilization phase, maintaining the current program angle, while meeting the total trajectory change duration. After that, the orbital change operation ends, and the next satellite-rocket separation phase begins.

[0092] Furthermore, incremental attitude adjustment is not recommended when designing the satellite-launch separation attitude trajectory after the target orbit inclination has been adjusted. During the final orbital change phase, the orbital parameters (such as inclination and semi-major axis) have been precisely adjusted to the target values ​​by iteratively calculating the program angle and supplementing the velocity to be increased during the final orbital change. If incremental attitude adjustment (adjusting the attitude through multiple small thrusts) is used at this point, it will directly introduce additional velocity increments, causing the orbital state to deviate from the target.

[0093] During the next satellite-rocket separation phase, the determined final orbit change acceleration will be determined as follows: the transformation matrix between the launch inertial system and the orbital coordinate system will be determined, and the acceleration to be increased in the launch inertial system will be determined based on the transformation matrix.

[0094] The transformation matrix between the launch inertial frame and the orbital coordinate system. Represented as:

[0095]

[0096] in A unit vector representing the position of the launching inertial frame. Indicates the geocentric deviation of the launch inertial frame, r d This indicates the magnitude of the rocket's displacement from the Earth's center in its launch inertial frame. The unit vector representing the velocity of the launching inertial frame. V represents the velocity vector of the launching inertial frame. d This represents the velocity magnitude of the launching inertial frame.

[0097] The velocity to be increased in the launch inertial frame is expressed as:

[0098]

[0099]

[0100] Where ΔV a Δa represents the velocity to be increased along the semi-major axis of the track, Δi represents the deviation of the semi-major axis of the track, and ΔV represents the deviation of the track inclination angle. i This represents the velocity to be increased in the direction of the orbital inclination deviation, and Fm represents the Earth's gravitational constant. The current target parameters are represented by the semi-major axis at the end of the stable phase of the satellite-launch separation before the orbit change; a k This is represented as the semi-major axis of the track, calculated in real time based on the current point information; Indicates the parameters, namely the inclination angle of the target orbit for the orbit change; i k This indicates the orbital inclination angle calculated in real time based on the current point information; e k This represents the eccentricity calculated in real time based on the current point information; u k This indicates the latitude argument calculated in real time based on the current point information.

[0101] This application also provides a computer storage medium storing computer instructions, which, when invoked, are used to execute the launch vehicle satellite-rocket separation stage target orbit control method.

[0102] The embodiments disclosed in this invention provide a computer-readable storage medium storing computer program instructions. When the computer program instructions are executed on a computer, the computer causes the computer to perform the above-described method for controlling the variable target orbit of a launch vehicle's satellite-rocket separation stage.

[0103] This invention provides a processor for processing the above-described method for controlling the orbit of a variable target during the separation phase of a launch vehicle.

[0104] In this embodiment of the invention, the processor can be an integrated circuit chip with signal processing capabilities. The processor can be a general-purpose processor, a digital signal processor (DSP), an application-specific integrated circuit (ASIC), a field-programmable gate array (FPGA), or other programmable logic devices, discrete gate or transistor logic devices, or discrete hardware components.

[0105] The various methods, steps, and logic diagrams disclosed in the embodiments of this invention can be implemented or executed. The general-purpose processor can be a microprocessor or any conventional processor. The steps of the methods disclosed in the embodiments of this invention can be directly implemented by a hardware decoding processor, or implemented by a combination of hardware and software modules in the decoding processor. The software modules can reside in random access memory, flash memory, read-only memory, programmable read-only memory, electrically erasable programmable memory, registers, or other mature storage media in the art. The processor reads information from the storage medium and, in conjunction with its hardware, completes the steps of the above methods.

[0106] The storage medium can be memory, such as volatile memory or non-volatile memory, or may include both volatile and non-volatile memory.

[0107] Non-volatile memory can be read-only memory (ROM), programmable read-only memory (PROM), erasable programmable read-only memory (EPROM), electrically erasable programmable read-only memory (EEPROM), or flash memory. Volatile memory can be random access memory (RAM), which is used as an external cache. By way of example, but not limitation, many forms of RAM are available, such as Static Random Access Memory (SRAM), Dynamic Random Access Memory (DRAM), Synchronous DRAM (SDRAM), Double Data Rate Synchronous DRAM (DDRSDRAM), Enhanced Synchronous DRAM (ESDRAM), Synchlink DRAM (SLDRAM), and Direct Rambus RAM (DRRAM).

[0108] This application has the following beneficial effects:

[0109] (1) This application can calculate the orbit change in real time according to the target orbit parameters and the current flight status parameters. It is simple in design, easy to implement and practical.

[0110] (2) In this application, quaternions are used for attitude control during the orbit change process to avoid the occurrence of singularity in attitude under specific Euler angles and to ensure the reliability of the algorithm.

[0111] (3) The launch vehicle involved in this application has the ability to change orbits, which can reduce the requirement for the consistency of satellite orbits in the same launch, further shorten the launch cycle of the launch vehicle, and reduce the launch capacity loss of the launch vehicle.

[0112] Although the examples referred to in the present application are described for illustrative purposes only and not for limiting the scope of the application, changes, additions and / or deletions to the implementation may be made without departing from the scope of the application.

[0113] The above description is merely a specific embodiment of this application, but the scope of protection of this application is not limited thereto. Any variations or substitutions that can be easily conceived by those skilled in the art within the scope of the technology disclosed in this application should be included within the scope of protection of this application. Therefore, the scope of protection of this application should be determined by the scope of the claims.

Claims

1. A method for controlling the variable target trajectory during the separation phase of a launch vehicle, characterized in that, Includes the following steps: Determine the required attitude for the orbit change based on the orbit change requirements; Adjust the arrow's posture according to the required posture; After the rocket body attitude is adjusted, it stabilizes within a specified time. After stabilization, it changes course until it reaches the target trajectory. Upon reaching the target track, shut down the final repair power system to complete the target track adjustment; After the final power system is shut down, the system enters the stabilization phase and maintains the current program angle. After the total orbit change time is met, the orbit change operation ends and the system moves into the next satellite-rocket separation phase. During the next satellite-rocket separation phase, when orbit changes are required, the determined final orbit change acceleration includes: determining the transformation matrix between the launch inertial frame and the orbital coordinate system, and determining the launch inertial frame acceleration based on the transformation matrix. The transformation matrix between the launch inertial frame and the orbital coordinate system. Represented as: in , ; A unit vector representing the position of the launching inertial frame. This indicates that the launch inertial frame has lost its geocentric length. This indicates the magnitude of the rocket's displacement from the Earth's center in its launch inertial frame. The unit vector representing the velocity of the launching inertial frame. This represents the velocity vector of the launching inertial frame. This represents the velocity magnitude of the launching inertial frame. Launch inertial frame to be increased velocity Represented as: in , ; in This indicates the velocity to be increased along the semi-major axis of the track. This indicates the deviation of the semi-major axis of the track. This indicates the deviation of the track inclination angle. This represents the velocity to be increased in the direction of the orbital inclination deviation. Represents the Earth's gravitational constant. The current target parameters are represented by the semi-major axis at the end of the stable phase of the satellite-rocket separation before the orbit change. This is represented as the semi-major axis of the track, calculated in real time based on the current point information; Indicates the parameters, namely the inclination angle of the target orbit for the orbit change; This indicates the orbital inclination angle calculated in real time based on the current point information; This indicates the eccentricity calculated in real time based on the current point information; This indicates the latitude argument calculated in real time based on the current point information.

2. The method for controlling the variable target trajectory during the separation phase of a launch vehicle as described in claim 1, characterized in that, Determining the required attitude for orbit change based on the orbit change requirements includes calculating the required velocity increment for the target orbit inclination angle based on the current flight parameters and the target orbit inclination angle, and calculating the attitude adjustment target program angle based on the required velocity increment for the target orbit inclination angle.

3. The method for controlling the variable target orbit during the separation phase of a launch vehicle as described in claim 2, characterized in that, Variable target trajectory inclination to increase speed Represented as: Where T represents transpose. This represents the component of the acceleration on the X-axis. This represents the component of the acceleration on the Y-axis. This represents the component of the acceleration to be increased along the Z-axis.

4. The method for controlling the variable target orbit during the separation phase of a launch vehicle as described in claim 3, characterized in that, Target program angle , Represented as: 。 5. The method for controlling the variable target trajectory during the separation phase of a launch vehicle as described in claim 1, characterized in that, After the rocket's attitude is adjusted, it stabilizes within a specified time. Once stabilized, it changes trajectory until it reaches the target trajectory, including the following sub-steps: Real-time tracking of closed-circuit guidance procedure angles; Perform axis determination.

6. A variable target orbit control system for the satellite-rocket separation stage of a launch vehicle, characterized in that, include: The trajectory change attitude determination unit, the rocket body attitude adjustment unit, the trajectory change unit, and the final repair system shutdown unit; The attitude determination unit for orbit change is used to determine the attitude required for orbit change based on the orbit change requirements. Arrow body attitude adjustment unit, used to adjust the arrow body attitude according to the required attitude; The trajectory-changing unit is used to stabilize the rocket body after the attitude adjustment is completed within a specified time, and then perform trajectory changing until the target trajectory is reached. The final repair system shutdown unit is used to shut down the final repair power system after reaching the target track, thus completing the target track adjustment. After the final power system is shut down, the system enters the stabilization phase and maintains the current program angle. After the total orbit change time is met, the orbit change operation ends and the system moves into the next satellite-rocket separation phase. During the next satellite-rocket separation phase, when orbit changes are required, the determined final orbit change acceleration includes: determining the transformation matrix between the launch inertial frame and the orbital coordinate system, and determining the launch inertial frame acceleration based on the transformation matrix. The transformation matrix between the launch inertial frame and the orbital coordinate system. Represented as: in , ; A unit vector representing the position of the launching inertial frame. This indicates that the launch inertial frame has lost its geocentric length. This indicates the magnitude of the rocket's displacement from the Earth's center in its launch inertial frame. The unit vector representing the velocity of the launching inertial frame. This represents the velocity vector of the launching inertial frame. This represents the velocity magnitude of the launching inertial frame. Launch inertial frame to be increased velocity Represented as: in , ; in This indicates the velocity to be increased along the semi-major axis of the track. This indicates the deviation of the semi-major axis of the track. This indicates the deviation of the track inclination angle. This represents the velocity to be increased in the direction of the orbital inclination deviation. Represents the Earth's gravitational constant. The current target parameters are represented by the semi-major axis at the end of the stable phase of the satellite-rocket separation before the orbit change. This is represented as the semi-major axis of the track, calculated in real time based on the current point information; Indicates the parameters, namely the inclination angle of the target orbit for the orbit change; This indicates the orbital inclination angle calculated in real time based on the current point information; This indicates the eccentricity calculated in real time based on the current point information; This indicates the latitude argument calculated in real time based on the current point information.

7. The launch vehicle satellite-rocket separation stage variable target orbit control system as described in claim 6, characterized in that, The attitude determination unit for orbit change determines the attitude required for orbit change based on the orbit change requirements, including calculating the target orbit inclination increment velocity based on the current flight parameters and the target orbit inclination, and calculating the attitude adjustment target program angle based on the target orbit inclination increment velocity.

8. The launch vehicle satellite-rocket separation stage target trajectory control system as described in claim 7, characterized in that, The attitude determination unit for changing the trajectory requires adjusting the target trajectory inclination angle and the velocity to be increased. Represented as: Where T represents transpose. This represents the component of the acceleration on the X-axis. This represents the component of the acceleration on the Y-axis. This represents the component of the acceleration to be increased along the Z-axis.

9. The launch vehicle satellite-rocket separation stage variable target orbit control system as described in claim 8, characterized in that, Target program angle in attitude determination unit required for trajectory change , Represented as: 。 10. The launch vehicle satellite-rocket separation stage target trajectory control system as described in claim 6, characterized in that, The orbit-changing unit stabilizes within a specified time. After stabilization, the orbit-changing process continues until the target orbit is reached, including the following sub-steps: Real-time tracking of closed-circuit guidance procedure angles; Perform axis determination.

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