Carrier rocket satellite-rocket separation section variable target orbit control method and system thereof
By determining the trajectory change attitude, adjusting the rocket's attitude, and stabilizing the trajectory change, and by employing quaternion control and final-stage propulsion system adjustments, the problem of trajectory control during the launch vehicle's separation phase was solved, achieving efficient trajectory change and shortened cycle for the launch vehicle.
Patent Information
- Application Number
- CN202511960182.1
- Authority / Receiving Office
- CN · China
- Patent Type
- Applications(China)
- Current Assignee / Owner
- Filing Date
- 2025-12-23
- Publication Date
- 2026-02-06
AI Technical Summary
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 the launch vehicle, and reduce the launch capacity loss of the launch vehicle.
By determining the attitude required for the trajectory change, the rocket body attitude is adjusted and stabilized within a specified time. After stabilization, the trajectory change is performed until the target trajectory is reached. Quaternions are used for attitude control to avoid attitude singularities. The target trajectory adjustment is completed by using the final power system shutdown unit.
It enables real-time calculation of orbit changes based on target orbit parameters and current flight status, simplifying design and facilitating implementation. It also reduces the requirement for orbital consistency of satellites launched in the same batch, shortens the launch cycle of launch vehicles, and reduces capacity loss.
Smart Images
Figure CN121469892A_ABST
Abstract
Description
TECHNICAL FIELD
[0001] The present application relates to the field of aerospace, in particular, to a launch vehicle star-rocket separation section variable target orbit control method and system. BACKGROUND
[0002] In order to cope with the booming development of commercial satellite launch service market, while effectively utilizing the carrying capacity of national satellite launch tasks, the launch vehicle generally adopts the launch mode of multi-satellite splicing. The application of shared rocket launch mode further enhances the task adaptability of the launch vehicle, and can reduce the launch cost of a single satellite. At present, it has been popularized in international launch service projects, and has strong international competitiveness. In recent years, with the booming development of international commercial aerospace, major launch vehicles of countries around the world have launched high-performance "splicing" launch modes and related products. Among them, the United States Space Exploration Technology Corporation, Rocket Laboratory, European Ariane Company, and Russian State Space Group Corporation have completed multiple splicing launches relying on Falcon-9, Electron, Vega, and Union-2 launch vehicles, intensifying market competition. Chinese Long March-2C, Long March-6, Long March-8, and Force-1 launch vehicles have also launched multiple splicing launch configurations to meet commercial launch needs. At present, in order to reduce the cost of satellite launch, the launch vehicle launch will try to make the launch load close to the carrying capacity under the condition of sufficient carrying capacity. In the case that the satellite has no ability or limited ability to change the orbit, the satellite launched by the shared unified launch vehicle needs to meet the requirement that the satellite orbit is a unified orbit. Therefore, there is a demand for splicing of the launch vehicle satellite launch, and the same satellite orbit generally has constraints such as time and installation method, which reduces the frequency of rocket launch.
[0003] Therefore, how to provide a launch vehicle star-rocket separation section variable target orbit control method that reduces the requirement for consistency of the same satellite orbit in the same launch, further shortens the launch cycle of the launch vehicle, and reduces the loss of carrying capacity of the launch vehicle launch, has become a problem urgently to be solved in the field. SUMMARY
[0004] In order to solve the above problems, the present application provides a launch vehicle star-rocket separation section variable target orbit control method, which comprises the following steps: determining the required attitude of the variable orbit according to the variable orbit requirement; adjusting the attitude of the rocket body according to the required attitude; after the attitude adjustment of the rocket body is completed, stabilizing within a specified time, and after the stabilization is completed, performing variable orbit until reaching the target orbit; and closing the terminal correction power system after reaching the target orbit, and completing the target orbit adjustment.
[0005] The launch vehicle star-rocket separation section target orbit control method as claimed in any of the preceding claims, wherein the determination of the required attitude for the orbit change according to the orbit change requirement comprises: calculating a target orbit inclination to be increased speed according to the current flight parameters and the target orbit inclination; and calculating a target program angle for adjusting the attitude according to the target orbit inclination to be increased speed.
[0006] The launch vehicle star-rocket separation section target orbit control method as claimed in any of the preceding claims, wherein the target orbit inclination to be increased speed is expressed as: wherein T represents transposition, and AV x represents the component of the to-be-increased speed in the X axis, AV y represents the component of the to-be-increased speed in the Y axis, and AV z represents the component of the to-be-increased speed in the Z axis.
[0007] The launch vehicle star-rocket separation section target orbit control method as claimed in any of the preceding claims, wherein the target program angle is expressed as:
[0008]
[0009] The launch vehicle star-rocket separation section target orbit control method as claimed in any of the preceding claims, wherein after the completion of the adjustment of the attitude of the rocket body, the rocket body is stabilized within a specified time, and after the completion of the stabilization, the orbit is changed until the target orbit is reached, comprising the following sub-steps: tracking a closed-loop guidance program angle in real time; and performing an axis determination.
[0010] A launch vehicle star-rocket separation section target orbit control system, comprising: a required attitude for an orbit change determination unit, a rocket body attitude adjustment unit, an orbit change unit, and a terminal correction system shutdown unit; the required attitude for the orbit change determination unit is configured to determine a required attitude for an orbit change according to an orbit change requirement; the rocket body attitude adjustment unit is configured to adjust the attitude of the rocket body according to the required attitude; the orbit change unit is configured to stabilize the rocket body within a specified time after the completion of the adjustment of the attitude of the rocket body, and to change the orbit until the target orbit is reached after the completion of the stabilization; and the terminal correction system shutdown unit is configured to shut down a terminal correction power system after the target orbit is reached, and to complete the adjustment of the target orbit.
[0011] The launch vehicle star-rocket separation section target orbit control system as claimed in any of the preceding claims, wherein the required attitude for the orbit change determination unit determines the required attitude for the orbit change according to the orbit change requirement, comprising: calculating a target orbit inclination to be increased speed according to the current flight parameters and the target orbit inclination; and calculating a target program angle for adjusting the attitude according to the target orbit inclination to be increased speed.
[0012] The launch vehicle star-rocket separation section target orbit control system as claimed in any of the preceding claims, wherein the target orbit inclination to be increased speed in the required attitude for the orbit change determination unit is expressed as: wherein T represents transposition, and AV xrepresents the component of the to-be-increased speed in the X axis, AV y represents the component of the to-be-increased speed in the Y axis, AV z represents the component of the to-be-increased speed in the Z axis.
[0013] The launch vehicle star-rocket separation stage target orbit control system as described above, wherein the target program angle in the orbit change required attitude determination unit represents:
[0014]
[0015] The launch vehicle star-rocket separation stage target orbit control system as described above, wherein the orbit change unit is stable within a specified time, and after the stability is completed, the orbit change is performed until the target orbit is reached, including the following sub-steps: real-time tracking of the closed-loop guidance program angle; and performing axis determination.
[0016] The present application has the following beneficial effects:
[0017] (1) The present application can calculate the orbit change in real time according to the target orbit parameters and the current flight state parameters, and is simple in design and easy to implement, and has practicality.
[0018] (2) The present application uses quaternions for attitude control in the orbit change process, avoids singular phenomena of attitude in specific Euler angles, and ensures the reliability of the algorithm.
[0019] (3) The launch vehicle involved in the present application has the ability to change the orbit, which can reduce the requirement for the same satellite orbit consistency in the same launch, further shorten the launch period of the launch vehicle, and reduce the loss of launch capacity of the launch vehicle. BRIEF DESCRIPTION OF DRAWINGS
[0020] In order to more clearly illustrate the technical solutions in the embodiments of the present application or the prior art, the drawings needed to be used in the embodiments or prior art description will be briefly introduced below. Obviously, the drawings in the following description are only some embodiments described in the present application, and other drawings can also be obtained by those skilled in the art according to these drawings.
[0021] Figure 1 is a flowchart of a launch vehicle star-rocket separation stage target orbit control method provided according to an embodiment of the present application;
[0022] Figure 2 is a schematic diagram of the internal structure of a launch vehicle star-rocket separation stage target orbit control system according to an embodiment of the present application. DETAILED DESCRIPTION
[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 variable target orbit inclination angle and the velocity to be increased 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 Calculating the target angle of salary adjustment 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 orbit parameters accurately reach the target value, avoid over-adjustment or fuel waste, a double criterion of normal shutdown and time standby shutdown is set for the final correction propulsion system, and the final correction nozzle is closed when one of the two criteria is met.
[0051] The normal shutdown includes continuously monitoring the absolute value of the deviation of the orbit inclination from the target orbit inclination When the absolute value of the deviation is less than the second guidance constant ΔorbI'(typical design value is 0.001°, which can be referred to as about 20 control cycle running advance amount), the final correction nozzle is closed. The discrimination period is consistent with the fixed shaft discrimination period T1, ensuring that the deviation reaches the threshold value quickly.
[0052] The time standby shutdown includes real-time statistics of the opening time of the final correction propulsion system. If the opening time of the final correction propulsion system exceeds the preset final correction longest opening time in the transfer segment ), the final correction nozzle is closed. At this time, it is assumed that the final correction thrust has been sufficient to adjust the orbit to the target range within the longest working time, and forced shutdown is needed to avoid fuel depletion or attitude loss of control.
[0053] After the final correction nozzle is closed, the current program angle is maintained in the stable segment, and after the total transfer time is met, the transfer operation at this time is ended, and the next satellite-rocket separation segment is entered.
[0054] Further, it is not recommended to use incremental attitude adjustment when designing the satellite-rocket separation attitude trajectory after the target orbit inclination is changed. The final correction transfer stage has accurately adjusted the orbit parameters (such as inclination and semi-major axis) to the target value through iterative calculation of the program angle and supplement of the final correction transfer to-be-increased speed. At this time, if incremental attitude adjustment (through multiple small thrust adjustments) is used, it will directly introduce additional speed increment, causing the orbit state to deviate from the target.
[0055] The final correction transfer to-be-increased speed determined during the next satellite-rocket separation segment according to the transfer needs includes: determining the launch inertial system and orbit coordinate system conversion matrix, and determining the launch inertial system to-be-increased speed according to the launch inertial system and orbit coordinate system conversion matrix.
[0056] The launch inertial system and orbit coordinate system conversion matrix is expressed as:
[0057]
[0058] Wherein represents the unit vector of the launch inertial system position, represents the launch inertial system geocentric meridian loss, r d represents the modulus of the rocket geocentric meridian loss in the launch inertial system, 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; This 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 variable target orbit inclination angle and the velocity to be increased Represented as:
[0068]
[0069] Where T represents transpose, ΔV x ΔV represents the component of the acceleration on the X-axis. yΔVx represents the component of the velocity increment to be added in the X axis. z ΔVz represents the component of the velocity increment to be added in the Z axis.
[0070] According to the variable target orbit inclination velocity increment Calculate the target program angle of the adjustment Wherein The target pitch program angle is represented as: The target yaw program angle is represented as: the target roll program angle remains unchanged.
[0071] Wherein the target program angle is represented as:
[0072]
[0073] According to the first calculated closed-loop guidance program angle and the initial angle of the adjustment, the program angle is calculated according to the triangular wave adjustment
[0074] Through the above real-time calculation of the variable orbit according to the target orbit parameters and the current flight state parameters, the design is simple and easy to realize, and has practicality.
[0075] The missile body attitude adjustment unit 220 is used for adjusting the missile body attitude according to the required attitude.
[0076] When the missile body attitude is adjusted, the final correction power system is not started, and only the attitude control system is used to rotate the missile body to accurately align the direction of the final correction nozzle with the direction of the ΔV required for the variable orbit, for example, when the inclination is changed, the direction of the thrust needs to be perpendicular to the original orbit plane.
[0077] The variable orbit unit 230 is used for stabilizing the missile body attitude after the adjustment is completed within a specified time, and the final correction power system performs the variable orbit until the target orbit is reached after the stabilization is completed.
[0078] The variable orbit unit 230 performs the following sub-steps:
[0079] Step Q1: Real-time tracking of the closed-loop guidance program angle.
[0080] Wherein the stable segment before the final correction is started with the initial angle of the adjustment as the starting point, and the closed-loop guidance program angle is calculated iteratively in real time Maintain the direction of the thrust. When the stable time is met, the process is turned to the final correction working segment, so as to avoid the deviation of the thrust direction caused by the unstable attitude and affect the variable orbit precision.
[0081] It can be understood that, The parameters refer to the stable time before the final correction is started.
[0082] Step Q2: Perform the fixed axis discrimination.
[0083] After the final correction working section is started, the real-time tracking closed-loop guidance iterative calculation program angle is tracked, and the program angle is maintained after the fixed axis criterion is met.
[0084] Specifically, when the orbit parameters gradually approach the target values, the attitude of the missile body needs to be confirmed to be stably aligned with the target orbit direction through the fixed axis discrimination.
[0085] The fixed axis discrimination includes real-time calculation of the deviation absolute value of the orbit inclination and the target orbit inclination When the deviation absolute value is less than the first guidance constant ΔorbI (the typical design value is 0.01°), the fixed axis is determined. The deviation is repeatedly calculated according to the fixed discrimination period T1 to ensure the real-time nature of the criterion judgment.
[0086] After the fixed axis criterion is met, the attitude control system is switched from the tracking iterative program angle mode to the program angle maintenance mode, and the current attitude is locked.
[0087] The final correction system shutdown unit 240 is used to shut down the final correction propulsion system after reaching the target orbit, and complete the target orbit adjustment.
[0088] To ensure that the orbit parameters accurately reach the target values and avoid over-adjustment or fuel waste, the final correction propulsion system is provided with double criteria of normal shutdown and time backup shutdown, and the final correction nozzle is closed when one of the two criteria is met.
[0089] The normal shutdown includes continuous monitoring of the deviation absolute value of the orbit inclination and the target orbit inclination When the deviation absolute value is less than the second guidance constant ΔorbI' (the typical design value is 0.001°, which can be referred to as about 20 control period operation advance), the final correction nozzle is closed. The discrimination period is consistent with the fixed axis discrimination period T1 to ensure that the deviation reaches the threshold value and responds quickly.
[0090] The time backup shutdown includes real-time statistics of the opening time of the final correction propulsion system. If the opening time of the final correction propulsion system exceeds the preset final correction longest opening time of the orbit transfer section , the final correction nozzle is closed. At this time, it is assumed that the final correction thrust is sufficient to adjust the orbit to the target range within the longest working time, and forced shutdown is needed to avoid fuel depletion or attitude loss of control.
[0091] After the final correction nozzle is closed, the current program angle is maintained in the stable section, and after the total time of the orbit transfer is met, the current orbit transfer operation is ended and the next satellite-rocket separation section is entered.
[0092] Further, the incremental attitude adjustment mode is not recommended when designing the post-separation trajectory after changing the target orbit inclination. The orbit parameters (e.g. inclination, semi-major axis) are accurately adjusted to the target values by iteratively calculating the program angle and supplementing the final correction velocity during the final correction orbit phase. At this time, if the incremental attitude adjustment mode (adjusting the attitude through multiple small thrusts) is used, it will directly introduce additional velocity increments, causing the orbit state to deviate from the target.
[0093] The final correction velocity to be increased in the launch inertial system includes determining the launch inertial system to orbit coordinate system conversion matrix, and determining the launch inertial system to be increased velocity according to the launch inertial system to orbit coordinate system conversion matrix.
[0094] The launch inertial system to orbit coordinate system conversion matrix is expressed as:
[0095]
[0096] wherein represents the unit vector of the launch inertial system position, represents the geocentric meridian of the launch inertial system, r d represents the modulus of the geocentric meridian of the launch inertial system, represents the unit vector of the launch inertial system velocity, represents the launch inertial system velocity vector, V d represents the launch inertial system velocity modulus.
[0097] The launch inertial system to be increased velocity is expressed as:
[0098]
[0099]
[0100] wherein ΔV a represents the velocity to be increased in the orbit semi-major axis direction, Δa represents the orbit semi-major axis deviation amount, Δi represents the orbit inclination deviation amount, ΔV i represents the velocity to be increased in the orbit inclination deviation amount direction, Fm represents the earth's gravitational constant, represents the current target elements, and the semi-major axis at the end of the previous satellite-rocket separation stable segment before the orbit change is taken as the current target elements; a k is expressed as the orbit semi-major axis calculated in real time according to the current point information; represents the elements, i.e. the target orbit inclination during the orbit change; i k is expressed as the orbit inclination calculated in real time according to the current point information; e k is expressed as the eccentricity calculated in real time according to the current point information; u k is expressed as the latitude amplitude angle calculated in real time according to the current point information.
[0101] The application further provides a computer storage medium, which stores computer instructions, and the computer instructions are used to execute the launch vehicle target orbit changing control method in a launch vehicle separation phase when being invoked.
[0102] The disclosed embodiment provides a computer readable storage medium, which stores computer program instructions, and the computer program instructions make a computer execute the launch vehicle target orbit changing control method in a launch vehicle separation phase when being run on the computer.
[0103] The embodiment of the application provides a processor for processing the launch vehicle target orbit changing control method in a launch vehicle separation phase.
[0104] In the embodiment of the application, the processor can be an integrated circuit chip with signal processing capability. The processor can be a general processor, a digital signal processor (DSP), an application specific integrated circuit (ASIC), a field programmable gate array (FPGA) or other programmable logic device, a discrete gate or transistor logic device, a discrete hardware component.
[0105] The disclosed methods, steps and logic block diagrams in the embodiments of the application can be implemented or executed. The general processor can be a microprocessor or the processor can be any conventional processor. The steps of the method disclosed in combination with the embodiments of the application can be directly embodied as hardware code processor execution or executed by a combination of hardware and software modules in the code processor. The software module can be located in a random memory, a flash memory, a read-only memory, a programmable read-only memory or an electrically erasable programmable memory, a register or other mature storage medium in the field. The processor reads the information in the storage medium and combines the hardware to complete the steps of the above method.
[0106] The storage medium can be a memory, for example, can be a volatile memory or a non-volatile memory, or can include both volatile and non-volatile memories.
[0107] The non-volatile memory can be a Read-Only Memory (ROM), a Programmable ROM (PROM), an Erasable PROM (EPROM), an Electrically EPROM (EEPROM), or a flash memory. The volatile memory can be a Random Access Memory (RAM) used as an external cache. By way of example, and not limitation, many forms of RAM can be used, such as a Static RAM (SRAM), a Dynamic RAM (DRAM), a Synchronous DRAM (SDRAM), a Double Data Rate SDRAM (DDR SDRAM), an Enhanced SDRAM (ESDRAM), a Synchlink DRAM (SLDRAM), and a Direct Rambus RAM (DRRAM).
[0108] The present application has the following beneficial effects:
[0109] (1) The present application can calculate the orbit transfer in real time according to the target orbit parameters and the current flight state parameters, and is simple in design and easy to implement, and has practicality.
[0110] (2) The present application uses quaternions for attitude control in the process of orbit transfer, avoids singular phenomenon of attitude in the case of specific Euler angles, and ensures the reliability of the algorithm.
[0111] (3) The launch vehicle involved in the present application has the ability of orbit transfer, which can reduce the requirement for consistency of the same launch satellite orbit, further shorten the launch period of the launch vehicle, and reduce the loss of the launch capacity of the launch vehicle.
[0112] Although the examples referred to in the present application are described, they are only for the purpose of explanation and not limitation of the present application, and changes, additions and / or deletions can be made to the embodiments without departing from the scope of the present application.
[0113] The above merely provides the specific implementation of the present application, but the protection scope of the present application is not limited to this. Any person skilled in the art can easily think of the changes or replacements within the technical range disclosed by the present application, which should be covered in the protection scope of the present application. Therefore, the protection scope of the present application should be subject to the protection 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.
2. The method for controlling the variable target orbit 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, 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.
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.
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 target trajectory inclination angle and velocity to be increased in the attitude determination unit for trajectory change 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.
9. The launch vehicle satellite-rocket separation stage target trajectory 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.