1601results about "Spacecraft guiding apparatus" patented technology

The present disclosure relates to a MEMS inertial sensor device in a semiconductor chip package that includes an integrated circuit configured to process inertial sensor data. Preferred implementations utilize inertial sensors having different sensitivity ranges to adjust operation of dynamic system control such as motion and or attitude control of autonomous vehicles.

The invention relates to the technical field of spacecraft orbit control and space safety, and discloses a space debris orbit prediction and avoidance method, which comprises the following steps that: a ground computing center screens candidate targets and generates enhanced data packets for uploading; the spaceborne computer combines real-time navigation data to carry out geometric feature screening, and an instant high-risk target list is generated; controlling the star sensor to execute optical observation, and resolving an optical measurement relative orbit state by utilizing displacement superposition and Kalman filtering; selecting a high confidence coefficient or a conservative probability threshold according to the optical measurement state acquisition condition, and generating an avoidance instruction when the collision probability exceeds the limit; and the propulsion system responds to the instruction to execute avoidance, verifies the effect by confirming the semi-major axis variation and executes track recovery by selecting an aircraft. According to the method, the ground-air collaborative screening and shift superposition enhancement technology is adopted, the problems that the low-precision ephemeris false alarm rate is high and the detection capability of a satellite-borne sensor on a dark and weak target is insufficient are solved, and the accuracy of evasion decision making and the execution reliability are improved.

The invention provides a satellite orbital transfer control method and device, electronic equipment and a storage medium, and relates to the technical field of satellite control. The satellite orbital transfer control method comprises the following steps: constructing a three-dimensional centralized monitoring area corresponding to a target satellite based on acquired first operation information of the target satellite; based on the three-dimensional centralized monitoring area, generating a monitoring instruction, and based on the monitoring instruction, controlling a monitoring terminal corresponding to the target satellite to perform centralized monitoring on the three-dimensional centralized monitoring area; when it is monitored that the target terminal exists in the three-dimensional centralized monitoring area, second operation information corresponding to the target terminal in a unit time period is obtained, and a control mode corresponding to the target satellite is determined based on the first operation information and the second operation information; and based on the control mode, controlling the target satellite to perform orbital transfer operation. According to the invention, the orbital transfer capability of the target satellite can be improved and the collision probability can be reduced.

The invention discloses a fuel-saving control on-satellite autonomous planning method for rapid accompanying flying, fly-around flying and sweeping flying of an elliptical orbit coplanar target, and relates to a fuel-saving control on-satellite autonomous solving method for rapid approaching and high-precision accompanying flying / fly-around flying / sweeping flying from a long-distance position on the same orbit surface. Comprising the following steps: providing a three-time planning and five-time maneuvering long-distance approaching accompanying flight / fly-around / grazing flight configuration forming method; the efficient iteration method is suitable for solving the initial value of a minimum fuel control analytical solution and the final value of an iterative precise solution of a high-precision orbit extrapolation model through a satellite autonomous solution two-body model. The method is suitable for a near-earth space random elliptical orbit satellite to quickly approach a long-distance target on an approximate same orbit to carry out accompanying flight / fly-around / grazing flight, and the control precision can reach a ten-meter level; and the calculation amount is small, and the method is suitable for on-satellite autonomous task planning and solving.

The present disclosure is a device and methods for satellite trajectory control and collision avoidance. Embodiments of the disclosure are comprised of a smart satellite device performing a process three steps. First, sensors collect data about the satellite's physical landing environment, passing information to satellite's database and processors. Second, the processors manipulate the information with a deep reinforcement learning program to produce instructions. Third, the instructions steer the satellite body by manipulating the satellite's panels for optimal trajectory and collision avoidance. The purpose for the present disclosure is to help solve the space debris problem.

The invention provides a low-orbit navigation constellation autonomous orbit determination method and system under a weak inter-satellite topological structure, and relates to the technical field of satellite orbit determination, and the method comprises the steps: obtaining various observation data of an inter-satellite link on a predetermined orbit, and carrying out the preprocessing of the various observation data; establishing a synchronous observation equation of orbit and clock correction parameters by adopting a segmented linear polynomial satellite clock correction modeling method, generating orbit, clock correction and related error parameters combining satellite-borne GNSS and inter-satellite link data, and forecasting; based on an error analysis result, multiple key features are extracted from the generated orbit, clock error and related error parameters to serve as input of a parameter prediction model, and a parameter time sequence of four types of prediction including relative weight, inter-satellite link hardware delay, low-orbit satellite clock error and unmodeled error is obtained through output; and uploading the predicted parameter time sequence and the broadcast ephemeris to a main satellite together, and transmitting to each sub-satellite by the main satellite through an inter-satellite link to realize distributed autonomous orbit determination of the sub-satellites.

The invention discloses a full-process closed-loop quick response satellite collision avoidance method, which comprises the following steps of: 1, observing and acquiring an orbit state of a space object, and then carrying out multi-source data fusion so as to construct a space object database; 2, screening out high-risk collision sources which have the possibility of collision with the own satellite; 3, obtaining a determined high-risk collision source according to the track state historical record of the possible high-risk collision source and the own track state; 4, generating an initial evasion maneuvering scheme, and optimizing the initial evasion maneuvering scheme through a multi-objective optimization algorithm; and obtaining a final evading maneuvering scheme for subsequent execution. And 5, according to the engine startup and shutdown time and the thrust vector contained in the final evasion maneuver scheme, the satellite engine is controlled to work with the specified thrust vector within the given startup and shutdown time, and collision evasion maneuver is executed. Compared with a conventional method, the method has the advantage that the response speed is higher.

The invention discloses a low-orbit target orbit determination method and device. The method comprises the following steps: acquiring observation data at different moments; determining a target position vector according to the observation data and a pre-established target position relation model; the target position vector comprises a first target position vector corresponding to an initial observation moment, a second target position vector corresponding to an intermediate observation moment and a third target position vector corresponding to a termination observation moment; the target position relation model is determined based on a second target position vector corresponding to the intermediate observation moment and a position vector of the observation platform relative to the earth center; and determining a predicted target orbit according to the initial observation moment, the first target position vector, the termination observation moment and the third target position vector. According to the method, the target orbit parameters can be determined only based on the observation data at different moments under the conditions that an initial value is not needed and perturbation errors exist, and the accuracy and stability of low-orbit target orbit determination are improved.

The invention provides a satellite weak and small target on-orbit observation system and method based on intelligent closed-loop feedback, and the system comprises a detection module which carries out the image enhancement and time sequence consistency enhancement of an infrared image, obtains an enhancement feature, and obtains a candidate target set based on the enhancement feature; the tracking module is used for carrying out target matching in the search area; when the matching succeeds, the candidate position is used as an observation value, and the observation value is input into a Kalman filter to obtain a target state vector; when the matching fails, taking a prediction state of the Kalman filter as a target state vector, expanding a search window by taking a prediction position as a center, and executing re-identification; the control module is used for mapping the target state vector into an attitude error and generating a control instruction by adopting a single-neuron self-adaptive PID (Proportion Integration Differentiation) controller; and the closed-loop scheduling module is used for dynamically adjusting operation parameters of at least one module according to the execution error and the detection confidence coefficient. According to the invention, the continuous tracking precision and attitude control stability of the weak and small target are improved.

A heat balance optimal satellite sun orientation control method comprises the following steps: firstly, determining sun orientation target parameters according to a sun vector direction calculated by an orbit, using preset sun orientation target parameters when regular preset parameters are set, and then calculating a sun orientation target attitude angle and angular velocity by using a formula provided by the method, and finally, the sun orientation of the satellite to the target attitude angle and angular velocity is realized according to the measurement of the terahertz gyroscope and the plane control of the jet phase.

The invention relates to the technical field of satellite attitude control, in particular to a satellite-borne laser semi-autonomous pointing control method, electronic equipment and a storage medium. The method is completed by cooperation of a ground system and an on-satellite system, the ground system calculates a top target passing moment T * in advance, generates a unified laser pointing control instruction and injects the instruction into a satellite; and the satellite-borne system receives and analyzes the instruction, obtains parameters such as T * and beam selection, and then autonomously controls a specified load for work preparation. And the satellite-borne system autonomously calculates the optimal moment T1 and the target attitude when passing through the calibration field through high-precision orbit extrapolation based on the latest navigation data, finally autonomously executes attitude maneuver and laser pointing tasks, and autonomously returns after the tasks are completed. Instructions in a uniform format are generated and injected from the ground, ground operation is greatly simplified, efficient cooperation of ground decision triggering and on-satellite autonomous execution is achieved, pointing errors caused by orbit extrapolation are greatly reduced, the pointing precision and the task success rate are improved, and universality is good.

The invention discloses a satellite attitude system anti-interference finite time control method under complex dynamic disturbance, and the method comprises the steps: building a rigid satellite attitude maneuver control model according to satellite attitude dynamics and kinematics equations; designing a self-adaptive fine disturbance observer to estimate the total disturbance of a satellite attitude system by combining the influence of the external disturbance variable on the rigid satellite attitude maneuver process, and outputting an estimated value of the external disturbance variable; a three-stage segmented sliding mode control method is designed, dynamic continuity of a three-stage sliding mode is ensured, and segmented sliding mode control output is obtained; and compensating the segmented sliding mode control output by using an estimated value of an external disturbance variable output by a self-adaptive fine disturbance observer to realize satellite attitude stable control. According to the method, the problem that the stability of the satellite is reduced in the attitude maneuvering performance process when external disturbance exists is solved, better observation performance and shorter attitude stabilization time are achieved, and the control precision of the satellite is effectively improved.

A satellite for software-defined satellite imagery includes an optical assembly. The satellite further includes an attitude control system. The attitude control system is configured to control an orientation of the satellite such that the optical assembly slews across a ground track of the satellite. The satellite further includes a set of image sensors configured to capture an image using light directed to the set of image sensors at least in part by the optical assembly.

A multi-satellite task scheduling hybrid control method driven by a large language model comprises the steps that a natural language task instruction input by a user is subjected to semantic analysis through an adaptive large language model, and structured scheduling parameters, heuristic scheduling algorithm control parameters and multi-target optimization weights are automatically generated; the method is used for guiding dynamic allocation and resource scheduling of satellite tasks. The method is combined with a heuristic algorithm to execute multi-objective optimization calculation, and a search strategy and a convergence behavior are adjusted according to task requirements and constraint conditions, so that the optimization effect of a scheduling scheme is ensured. Furthermore, the method supports a user to perform semantic-driven re-optimization based on natural language feedback of an initial scheduling result, realizes iterative update of a scheduling scheme, forms a closed-loop control process from natural language input to optimization result to feedback adjustment, and gives consideration to scheduling intelligence, controllability and multi-objective optimization performance. The method solves the problems that a traditional method is insufficient in environment dynamic adaptation and high in initial parameter and weight dependence, efficient multi-satellite task scheduling with controllable semantics and man-machine cooperation is achieved, and the method is suitable for scenes such as satellite group cooperation observation and real-time task planning.

An attitude control method for a spin detection satellite, which method belongs to the field of satellite attitude control. The method comprises: S10, establishing a satellite dynamic equation and converting same into a state-space equation; S20, designing a two-layer linear extended state observer, wherein the first-layer linear state observer is used for observing and acquiring a current angular velocity estimate and a current disturbance estimate of a satellite, and the second-layer state observer observes a tracking error of the first-layer linear state observer on the basis of the observation result of the first-layer linear state observer; and S30, then performing calculation on the basis of the current disturbance estimate and the tracking error thereof, so as to obtain the total internal and external disturbance torque estimate of the satellite, using attitude parameters representing the direction of a spin axis to design a PD control law, and compensating for the total internal and external disturbance torque estimate of the satellite, so as to calculate a control torque, which is used for satellite attitude control within a current cycle. The method has the advantages of fewer tuning parameters, ease of adjustment and stable performance, thereby greatly improving the observation precision.

The invention provides an air pressure regulation and control device, a cold air micro-propulsion system and a satellite, and relates to the technical field of cold air micro-propulsion. The air pressure regulation and control device comprises a high-pressure valve, a two-stage pressure reducing valve, a precise pressure regulating valve, a buffer gas cylinder and a pressure sensor; one end of the high-pressure valve is connected with the high-pressure gas cylinder, and the other end of the high-pressure valve is connected with the two-stage pressure reducing valve through a first pipeline; the precise pressure regulating valve is provided with a first connector, a second connector, a third connector and a fourth connector, the first connector is connected with the two-stage pressure reducing valve through a second pipeline, the second connector is connected with the pressure sensor through a third pipeline, the third connector is connected with the buffer gas cylinder through a fourth pipeline, and the fourth connector is connected with the thruster. Compared with the prior art, the air pressure regulation and control device has the advantages that the mode of combining two-stage pressure reduction, air capacity buffering and air pressure closed-loop precise regulation and control is adopted, stable and low-noise air can be supplied to the thruster, and therefore high-stability and low-noise thrust output is achieved.

The invention relates to a near-earth spacecraft strong transverse maneuvering reentry trajectory planning method. The method comprises the following steps: determining basic parameters of return reentry trajectory planning; calculating a reentry trajectory of the maneuver without transverse offset; distributing a transverse maneuver amount, and establishing a transverse maneuver planning equation; calculating a braking yaw angle; calculating an initial value of the heeling angle offset; establishing a funnel guidance law under the condition of transverse offset maneuvering; according to the determined return and reentry trajectory planning basic parameters, the total transverse offset maneuver amount, the braking yaw angle and the funnel guidance law under the transverse offset maneuver condition, the initial value of the heeling angle offset amount, the time of the braking point relative to the return circle ascending node and the braking starting duration serve as initial values, reentry trajectory parameters are iteratively solved, and reentry trajectory planning is completed. According to the method, the problem of rapid planning of the reentry trajectory when the ground proximity device returns to the designated position under the condition of large initial transverse deviation is solved, and meanwhile, the problem of rapid planning of the reentry trajectory when the landing point needs to avoid the designated area can be solved.

An artificial satellite and artificial satellite system has power transmission and reception efficiency and includes an expandable solar panel that generates power, expandable power transmitting antennas, and expandable beam members to connect the solar panel and the power transmitting antennas to each other. The beam members are expanded from being provided in the artificial satellite, on a plurality of rotational symmetry axes of an imaginary polyhedron formed with substantially a center of the artificial satellite as an origin. The beam members are arranged at substantially equal distances in directions of the rotational symmetry axes. A light receiving surface of the solar panel and installation surfaces of the power transmitting antennas, when expanded, are substantially perpendicular to a solar panel rotational symmetry axis passing through a center of the solar panel, and being arranged on opposite sides of a plane passing through the center of the artificial satellite.

The invention relates to an earth-moon space low-thrust spacecraft fuel optimal collision avoidance trajectory implementation method, which comprises the following steps of: constructing a kinetic model based on calculation parameters of a spacecraft collision avoidance problem, and establishing a fuel optimal performance index of the spacecraft collision avoidance problem through a homotopy method; the optimal control problem is converted into a two-point boundary value problem, targeting solving is carried out, and the state and co-state parameters of the optimal orbit are obtained; and according to the obtained state and co-state parameters of the optimal orbit, solving to obtain collision avoidance optimal control and collision probability of the spacecraft. According to the method, the fuel optimal control problem is solved by adopting the indirect method, the convergence difficulty of the low-thrust fuel optimal problem is reduced through the homotopy method, and when the thrust is calculated, the calculation efficiency is high, the precision is high, and the method is suitable for the fuel optimal collision avoidance scene of the low-thrust spacecraft for a single target in the earth-moon space.

The invention relates to the technical field of satellite attitude control, in particular to a satellite-borne laser full-autonomous pointing control method, electronic equipment and a storage medium. The method comprises the following steps: firstly, presetting the position of a ground fixed calibration field, a laser light emitting phase and pointing information on a satellite; the on-satellite system autonomously carries out long-term orbit extrapolation, judges whether a satellite is about to pass through a calibration field, calculates to obtain a rough target moment T * and a day and night state thereof, and autonomously notifies a ground user. And the satellite autonomously controls work preparation of laser and camera loads according to T *. And at the moment before T *, the satellite-borne system autonomously solves the optimal moment T1 and the target attitude of passing through the calibration field through iterative extrapolation calculation and preferential selection based on high-precision navigation data. And finally, the satellite autonomously maneuvers to a target attitude, laser is controlled to accurately point to the calibration field at the T1 moment, and the satellite autonomously returns after the task is completed. According to the method, on-satellite autonomy from task planning to execution of the whole process is realized, and the autonomy, the working efficiency and the pointing precision of the satellite are greatly improved.

In a space traffic management system (500), space traffic management devices (100) individually mounted in a plurality of mega-constellation business devices and in a debris removal business device (45) are connected to each other via a communication line (200). The debris removal device (45) performs Active Debris Removal (ADR) against debris formed by a satellite managed by a first mega-constellation business operator. The debris removal device (45) acquires real-time high-accuracy orbital information of a satellite group of a second mega-constellation business operator in a timeframe in which a debris removal satellite, during orbital descent, passes through an orbital altitude region where the satellite group of the second mega-constellation flies, the debris removal satellite passing through the satellite group while ensuring flight safety.

The invention relates to a lunar satellite formation collaborative navigation method, which comprises the following steps of: inputting an orbit state and an observation residual error of a lunar satellite formation at a previous moment into an LSTM (Long Short Term Memory) prediction network of a navigation algorithm to predict a process noise covariance matrix and an observation noise covariance matrix, and multiplying the process noise covariance matrix and the observation noise covariance matrix by a scaling factor for scaling, obtaining a scaled prediction process noise covariance matrix and a scaled observation noise covariance matrix; and inputting the orbit state of the lunar satellite formation at the previous moment, the multi-source observation data at the current moment and the scaled process noise covariance matrix and observation noise covariance matrix into a self-adaptive extended Kalman filtering module in a navigation algorithm to predict the orbit state of the lunar satellite formation at the current moment.

The invention discloses a fuel-optimal side-by-side four-star configuration periodical maintenance control method. The method comprises the following steps: establishing a complete kinetic model containing J2 perturbation and atmospheric perturbation effects in an LVLH coordinate system; based on the model, the change of the phase along with time under the influence of atmospheric resistance is obtained, and the monotonous change rule of the center deviation of different satellite pairs in the formation satellites along with time is obtained; the baseline length of the satellite pair and the center deviation distance between the first and fourth tracks of the satellite and the second and third tracks of the satellite are used as core constraint conditions. In the optimization framework, weighted fuel cost is used as a target function, and differentiated fuel weight distribution is carried out on different satellite pairs. A hybrid optimization strategy of global search and local refining is carried out by adopting a genetic algorithm, an optimal control variable is efficiently solved, and finally, the fuel consumption is minimized on the premise of meeting all constraints.

The invention discloses a time-consistent approaching multi-satellite cooperative control method for a space maneuvering target, and the method comprises the following steps: determining an approaching control expression of tracking satellites based on the control acceleration amplitude of each tracking satellite at a moment t, the remaining time of the tracking satellites approaching a maneuvering target, and the zero-control off-target vector of the tracking satellites; iteration is carried out based on the remaining time of each tracking satellite for separately approaching the escape target by adopting the maximum amplitude control acceleration, and the remaining time of the tracking satellites for approaching the escape target with consistent time is determined; determining the control acceleration amplitude of the tracking satellite based on the remaining time of approaching the escape target, the zero-control off-target distance and the maximum amplitude of the escape target control acceleration according to the time consistency; determining time consistent approaching control of the plurality of tracking satellites at the current moment according to the state vectors of each tracking satellite and the escape target in the local Oxyz coordinate system; and the time-consistent approaching of the maneuvering space target is realized.

The invention discloses a parameter self-adaptive control method for a low-orbit satellite with a solar panel driving device. Belongs to the technical field of aerospace, and particularly relates to the technical field of low-orbit satellite parameter control. The satellite attitude control method solves the technical problems of overcoming the influence of the rotational inertia of the satellite and compensating the influence of the SADA vibration disturbance torque and the atmospheric resistance torque on the satellite attitude control precision. The method comprises the following steps: constructing a satellite attitude kinematics and dynamics equation in a quaternion form, and constructing a satellite attitude kinematics and dynamics equation for designing an attitude controller on the basis of the satellite attitude kinematics and dynamics equation; defining an error attitude kinematics and dynamics equation of the satellite through the attitude control error and the angle control error; carrying out segmentation conversion on a to-be-estimated parameter and a known state variable, and expressing the negative influence of a rotational inertia error on the attitude by matrix operation; designing the control torque of the attitude controller by adopting a mode of combining nonlinear feedback control with self-adaptive control; and updating the self-adaptive estimated value in the control parameters of the attitude controller by adopting a parameter self-adaptive law.

The embodiment of the invention provides a collision avoidance method and device, equipment and a storage medium, which can be applied to the technical field of space debris, and the method comprises the steps: determining a plurality of time intervals for a spacecraft to execute collision avoidance on at least one dangerous object within a set time period; for any time interval, determining at least one avoidance strategy in the time interval based on the initial orbit of the spacecraft and the operation orbit of the at least one dangerous object; the evasion strategy comprises a target pulse moment and a target evasion speed pulse applied to the spacecraft at the target pulse moment; predicting whether a dangerous object with a new collision risk exists or not after the at least one avoidance strategy is executed; and the avoidance strategy without the new collision risk is determined as the target avoidance strategy, so that accurate avoidance of at least one dangerous object by executing the avoidance strategy once is realized, and the collision avoidance calculation automation and intelligence are improved.

The invention provides an in-orbit energy management method and system for a damaged satellite, network equipment, a computer readable storage medium and a computer program product. The method comprises the following steps: autonomously determining whether to execute the reasoning of the SADA rotation angle based on the current state data of a satellite, and selecting a reasoning model under the current energy limitation condition; simulating a regression relationship between the SADA rotation angle and the predicted current of the solar panel according to the selected reasoning model and the current state data of the satellite; and searching the corresponding SADA target rotation angle capable of enabling the predicted current of the solar panel to be maximum from the regression relationship based on a Bayesian optimization strategy, and converting the SADA target rotation angle into an SADA control instruction. According to the method, a damaged satellite energy management framework is constructed, real-time and accurate prediction of the optimal sun-pointing angle of the current peak value of the solar panel in the current damaged state is realized, energy fluctuation caused by satellite attitude and sun vector change is avoided, and an important basis is provided for autonomous maintenance of energy supply on a satellite.

The invention belongs to the technical field of satellite in-orbit management, and provides a low-orbit satellite inter-satellite cooperative planning execution method and system, and the method comprises the steps: planning an inter-satellite cooperative task, and transmitting an inter-satellite cooperative task package to a low-orbit satellite; the low-orbit satellites execute single-satellite inter-satellite cooperation tasks according to the inter-satellite cooperation task packages and the guiding information of the low-orbit satellites before the low-orbit satellites; and the low-orbit satellites send group guidance information to low-orbit rear satellites according to single-satellite inter-satellite cooperative task execution results. According to the method, a low-orbit multi-satellite inter-satellite collaborative architecture and an inter-satellite collaborative task planning method under information flow of the low-orbit multi-satellite inter-satellite collaborative architecture are established by disassembling a collaborative task execution division mechanism, a logic closed loop and a time closed loop which can meet the execution instruction chain time constraint and realize planning execution are considered, the preparation time for establishing a satellite working state can be effectively shortened, and the working efficiency of a satellite is improved. The efficiency of inter-satellite cooperation of the low-earth-orbit satellites is improved, and theoretical support and a systematic solution are provided for inter-satellite cooperation planning and execution of the low-earth-orbit satellites.

The invention discloses a high-reliability satellite-borne orbit control execution method and system, and the method comprises the steps: carrying out the orbit control strategy calculation at a certain time interval, carrying out the orbit control strategy consistency confirmation of N pieces of recorded orbit control strategy information after N times of calculation, and taking the orbit control strategy calculated at the last time as a selected orbit control strategy after N times of calculation; selecting an optimal thruster number and duty ratio combination meeting the orbit control time limit in the current orbit control strategy information; further confirming whether the execution time of the current orbit control strategy is sufficient or not; if yes, entering the next step; calculating fuel consumption of a thruster group and a duty ratio combination corresponding to the current orbit control strategy, if the fuel consumption is smaller than a task fuel expectation, considering that the strategy is safe and can be executed, and entering the next step; executing orbit control according to the orbit control strategy information, judging whether the orbit control is finished or not in the orbit control execution process, and performing shutdown protection when the orbit control exceeds the predicted orbit control time. The method has the characteristics of high reliability, high precision, high reusability and the like.

This invention provides a method and system for satellite insertion into a highly elliptical frozen orbit, including multiple small-thrust maneuvers at apogee to raise the perigee altitude; naturally increasing the perigee angle using positive perturbation factors; evenly and rationally allocating the working time of each thruster; using inertial pointing to complete each ignition maneuver; positively offsetting the target orbit perigee altitude; performing small-thrust maneuvers near apogee to lower the perigee altitude to the target value, ultimately completing satellite insertion; the ignition altitude is selected above 28,000 kilometers. In this invention, the satellite can still enter its mission orbit under fault conditions, ensuring the satellite's service life.