1184results about "Spacecraft guiding apparatus" patented technology

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 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 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 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.

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.

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 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.

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.

Methods and systems of creating accurate data representations of the positions and other characteristics of target space objects using data from a plurality of space-borne artificial satellites and ground-based sensors. Such methods and systems are deployable either on a single satellite, on a plurality or swarm of satellites, and / or in a ground-based system. Various artificial intelligence and machine learning processes described herein are employed to generate a latent data representation of passively-received broad spectrum multi-modal signal data systematically integrated with advanced multi-sensor data fusion algorithms.

Described herein is a method for wireless transmission of energy in space, wherein a constellation of transmitting satellites orbiting around a celestial body, such as the planet (P) Earth or the Moon (L), are equipped with a laser apparatus for sending a beam of coherent electromagnetic waves to an orbiting satellite or another receiving target, such as a mobile vehicle or a fixed base on said generic celestial body, in order to transmit energy thereto. The total power transmitted is divided among the orbiting satellites in the constellation, resulting in less contraindications than in the case of a single transmitting system with equivalent total power. The advantage of such division among a constellation of satellites mainly consists of shorter periods of solar eclipse, since sunlight is necessary for the operation of the transmission system. The receiving target or satellite converts the beam of electromagnetic waves into electric energy.

Provided is a semiconductor device which exhibits favorable electrical properties. This semiconductor device comprises an oxide semiconductor layer, a first insulating layer, a conductive layer having a portion overlapping the oxide semiconductor layer with the first insulating layer interposed therebetween, and a second insulating layer having a portion overlapping the first insulating layer with the oxide semiconductor layer interposed therebetween. The oxide semiconductor layer contains indium. The thermal expansion coefficient of the second insulating layer is less than the thermal expansion coefficient of the oxide semiconductor layer. The first insulating layer has a function of capturing oxygen and hydrogen, and the second insulating layer has a region containing excess oxygen.

Various embodiments comprise a multi-agent spacecraft system for rendezvous and proximity operations. The spacecraft system comprises a chief spacecraft and a deputy spacecraft. The chief spacecraft maneuvers to the orbit of a Resident Space Object (RSO) and deploys a deputy spacecraft to rendezvous with the RSO. The deputy spacecraft maneuvers to the RSO and senses the RSO during the rendezvous and proximity operations maneuver. The deputy spacecraft generates a Three-Dimensional (3D) surface reconstruction of the RSO based on the sensing. The deputy spacecraft determines a roll rate and a relative trajectory of the RSO based on the sensing. The deputy spacecraft returns to the chief spacecraft.

The invention provides a satellite-borne high-precision orbit extrapolation method, a satellite-borne high-precision orbit extrapolation device and electronic equipment. The satellite-borne high-precision orbit extrapolation method comprises the following steps: providing an orbit extrapolation input quantity of a satellite at an initial extrapolation moment; calculating an extrapolation duration based on the initial extrapolation moment and a preset threshold value, and determining a termination extrapolation moment of the satellite based on the extrapolation duration; calculating the non-spherical gravitational perturbation acceleration of the earth according to the non-spherical gravitational perturbation model of the earth; calculating a day-moon three-body gravitational acceleration according to a day-moon three-body gravitational model; determining a satellite thrust acceleration according to the current working state of the satellite; adding the earth non-spherical gravitational perturbation acceleration, the sun-moon three-body gravitational acceleration and the satellite thrust acceleration to obtain an extrapolation resultant acceleration; performing orbit extrapolation calculation on the basis of the extrapolation resultant acceleration by adopting a fourth-order Runge-Kutta method to obtain a target extrapolation position and a target extrapolation speed at an extrapolation termination moment; and according to the target extrapolation position and the target extrapolation speed, updating the orbit extrapolation input quantity at the next extrapolation moment.

The invention discloses a transient performance enhancement-oriented multi-spacecraft attitude cooperative tracking control method, which provides a fixed time controller through an inversion control strategy, realizes quick response of a multi-spacecraft attitude system, and ensures that a consistency error is converged to a sufficiently small domain within fixed time. A fuzzy weight updating law is designed, dependence on system model information is reduced, and adaptability and high robustness are achieved. According to the asymmetric fractional obstacle Lyapunov function, adherence of multiple spacecrafts to time-varying constraints is guaranteed, meanwhile, the problem that the control amplitude related to a large constraint boundary in a traditional logarithmic obstacle Lyapunov function is too large is solved, and therefore the proposed method can process constrained and unconstrained scenes. A saturation threshold event triggering strategy is designed, a multi-spacecraft system is allowed to select three different triggering strategies by adjusting parameters, the transmission frequency of control signals is flexibly changed, and control bus resources are reasonably distributed.

The invention relates to a spacecraft collaborative formation safety control method and system based on an adjustable self-adaptive control barrier function. The method comprises the following steps: establishing a six-degree-of-freedom model for describing relative motion of a spacecraft formation; multiple safety key constraints which need to be met by the spacecraft formation are established, and the constraints are formalized into a zero-hyperlevel set function of a state variable; a disturbance observer is designed, and unknown lumped disturbance existing in the six-degree-of-freedom model is estimated and compensated online; a safety controller is constructed, the safety controller comprises an adjustable self-adaptive control barrier function used for generating a safety constraint boundary changing along with time, and parameters of the adjustable self-adaptive control barrier function are adjusted online through the auxiliary dynamic system; and constructing a quadratic programming problem, and solving to obtain an optimal safety control instruction which minimizes the deviation between the actual control input and the nominal control input. Compared with the prior art, the method has the advantages of improving the task success rate and safety of the spacecraft formation in a complex, dynamic and uncertain environment and the like.

A method for avoiding strong light irradiation of a low-orbit spacecraft comprises the steps that 1, a low-orbit satellite starts ground strong light detection according to task requirements, and obtained detection telemetry information is sent to an on-satellite processing unit in real time; 2, the on-satellite processing unit judges and detects the effectiveness of the telemetry information; 3, the on-satellite processing unit interprets strong light energy and azimuth vector information, and obtains the current task state of the satellite, the solar wing attitude pointing information and the optical load pointing information at the same time; 4, evaluating the energy intensity of the strong light, judging the damage capability to the optical load and the solar wing, and carrying out corresponding operation; 5, after autonomous attitude avoidance and optical load pointing adjustment are completed, whether the strong light energy is eliminated or not continues to be judged; and after the strong light irradiation threat is relieved, the satellite-borne processing unit receives a ground remote control instruction and starts a new task. According to the method, the low-orbit satellite can avoid the threat of strong light, and the current situation that the current satellite has no prevention and resistance capability in orbit is changed.

The application discloses a kind of anti-saturation preset performance despun control methods of failed satellite and related devices, the method includes: obtaining the attitude of service spacecraft and failed satellite, establishes the dynamics model of service spacecraft and failed satellite;According to the dynamics model of service spacecraft and failed satellite, the flexible operating rod dynamics model of service spacecraft is established, the large deformation dynamics characteristics of despun brush are obtained, under the framework of the dynamics model of service spacecraft and the large deformation dynamics characteristics of despun brush, the rigid-flexible coupling dynamics model of service spacecraft is established by Lagrange multiplier method;Based on the rigid-flexible coupling dynamics model of service spacecraft, an anti-saturation preset performance controller is designed, and an anti-saturation full-state constraint control strategy is used for despun control.The application is aimed at service spacecraft with flexible operating rod, realizes efficient and accurate despun process, and greatly improves control performance.

Systems, methods, and instructions of computer-readable media may include obtaining, at a client machine, a user-selected configuration parameter for an orbit simulation; sending, from the client machine to a remote system, via a network connection, a first set of configuration parameters for the orbit simulation, wherein the first set of configuration parameters comprise the user-selected configuration parameter; receiving, at the client device from the remove device, via the network connection, a stream of orbital data comprising points along an orbit, wherein the points along the orbit are determined by the remote system based on the first set of configuration parameters; and presenting, at a display, a dynamic rendering of the orbit simulation, wherein the orbit simulation is based on the stream of orbital data.

The invention discloses an earth-moon translation point orbit spacecraft formation relative motion configuration design and control method, and belongs to the technical field of spacecraft orbit dynamics and control. The method comprises the following steps: firstly, constructing a double-layer dynamic model consisting of a circular restrictive three-body problem model and a high-precision ephemeris model, and defining an orbit amplitude difference and a phase difference as configuration parameters; then, according to different track amplitude difference and phase difference combinations, three types of natural relative motion configurations are set, a periodic track is rapidly generated under a circular restrictive three-body problem model, and multi-circle correction is conducted through a high-precision ephemeris model and two-stage multiple targeting; a hierarchical constraint model is established for a controlled task surrounding a fixed position, a plane and a space, and a control pulse is solved by adopting a two-stage optimization strategy. According to the method, efficient design and stable control of formation configuration in a complex dynamic environment are realized, and the method is suitable for various translation point orbit tasks.