2031results about "Spacecraft guiding apparatus" patented technology

The invention discloses a satellite group orbit pursuit decision-making method based on multi-near-end reinforcement learning. The method comprises the following steps: S1, updating a satellite state by adopting an eight-order Runge-Kutta method; s2, all satellites in the satellite group collect state information of each other, a state vector is constructed, and a strategy and value network is formed according to the state vector; selecting and executing an action through a policy network, and updating a state vector after receiving an environment reward; calculating a time sequence difference error and a dominant function, and updating the network by using a near-end strategy optimization algorithm; meanwhile, fusing the strategy networks of the satellites by adopting a federal near-end algorithm; and S3, carrying out comprehensive performance evaluation on the satellite group, dynamically adjusting an updating strategy of each satellite strategy network parameter, and continuously optimizing the overall efficiency of the satellite group. According to the method, efficient cooperation of the satellite groups in a complex environment is ensured, the overall response speed and decision accuracy are improved through continuous learning and adjustment, and efficient and stable task execution is ensured.

The invention provides a lunar satellite formation autonomous navigation method based on inter-satellite ranging information. The lunar satellite formation autonomous navigation method comprises three parts of orbit dynamics model establishment, collaborative observation model establishment and navigation filtering algorithm data fusion. The invention provides an autonomous navigation method based on inter-satellite distance measurement and angle measurement information, which establishes a high-precision dynamic model containing 165-order moon non-spherical perturbation for moon satellite formation. The system can operate independently of a ground station, and high-precision autonomous navigation of the lunar satellite formation is realized only by using on-orbit observation data.

Intelligent optimization of mission planning and execution for spacecraft through intelligent analysis of mission objectives and payload operation is disclosed. Values for mission objectives, payload operation, and / or vehicle conditions may be defined during given time periods, including variable values for repeat operations or other factors. These values may then be used through a vehicle / mission simulation agent to optimize the mission plan through maximizing the mission score according to the values. The simulation agent may identify possible mission plans based on vehicle capability, and then score the mission plan based on the total objective score.

The invention provides a lunar satellite formation orbit determination method and system. The method comprises the following steps: S1, establishing a lunar satellite high-precision dynamic model; s2, establishing a lunar satellite laser ranging and inter-satellite ranging mathematical model; s3, measuring the distance, namely the absolute position, of the A-type satellite relative to the ground station through laser ranging, and measuring the relative position of the B-type satellite relative to the A-type satellite through inter-satellite ranging; s4, preprocessing the absolute position and relative position data of the satellite measured by laser ranging and inter-satellite ranging, and generating a format data file required by orbit calculation; and S5, resolving the position coordinates of each satellite in the monthly formation through the format data file. According to the method, laser ranging and inter-satellite ranging are combined for orbit determination, precise orbit determination of the lunar satellite formation is achieved, compared with a traditional orbit determination mode, the precision can be improved from the hundred-meter level to the meter level, the requirements for the arrangement number of ground stations and set distribution are reduced, and the measurement and control pressure of existing ground stations is relieved.

The invention discloses a service spacecraft multi-orbit maneuvering strategy approaching observation method based on J2 perturbation compensation, solves the problem of how to quickly improve orbit maneuvering accuracy, and belongs to the field of service spacecraft multi-orbit maneuvering strategy approaching observation. Comprising the following steps: constructing a motion equation under J2 item perturbation compensation under a target orbit coordinate system; constructing a service spacecraft multi-orbit maneuvering strategy model through the relative motion propagation relational expression and the motion equation; inputting the termination moment position conforming to the termination moment distance constraint into the model to output a single speed increment, and reserving the single speed increment conforming to the termination moment illumination angle constraint and the single speed increment constraint at the same time to obtain a speed increment solution set; constructing a circular region which takes the space target as a center and has a radius meeting a termination moment distance constraint; when the service spacecraft arrives at the circular area at a specified moment, marking the task successfully, and obtaining a service spacecraft multi-orbit maneuvering strategy for approaching observation; according to the invention, accurate and rapid track maneuvering is realized.

The invention relates to a propulsion system, in particular to an air-breathing radio frequency ion electric propulsion system with mixed working medium power and a multi-mode working method thereof, and aims to solve the problems that an existing air-breathing electric propulsion system only takes atmosphere as a single working medium, the atmosphere trapping efficiency is low, and the performance of a spacecraft is affected. The device comprises a propelling and gas trapping unit, a radio frequency neutralizer, a power and control unit and a xenon storage and supply unit, the propelling and gas trapping unit comprises a radio frequency ion thruster and an atmosphere trapping gas inlet channel; the power and control unit comprises a direct current module, a control module and a radio frequency module; the xenon storage and supply unit comprises a xenon bottle and a valve and sensor assembly; the control module and the valve are electrically connected with the sensor assembly. The radio frequency ion thruster is electrically connected with the radio frequency neutralizer through the radio frequency module. The first output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency neutralizer, and the second output end of the xenon storage and supply unit is connected with the gas path input end of the radio frequency ion thruster.

The invention discloses an ultra-low orbit guarantee method, system and equipment for a near-earth satellite. Comprising the following steps: determining a satellite orbit element at the current moment based on an atmospheric resistance perturbation item, a conventional perturbation item and a small thrust perturbation item, and inputting the satellite orbit element at the current moment into an orbit mechanical model; based on the orbit mechanical model and the satellite orbit elements, determining a first velocity position parameter of the satellite at the prediction moment under the ICRS coordinate system, and converting the first velocity position parameter into an ECEF coordinate system to obtain a second velocity position parameter at the prediction moment; inputting the time characteristic, the satellite state characteristic, the space environment characteristic and the second speed position parameter of the current moment into an orbit correction model to obtain a target error prediction value; adjusting the second speed position parameter based on the target error prediction value to obtain a target speed position parameter of the satellite at the prediction moment; and performing orbit attenuation risk early warning based on the target speed and position parameters of the satellite at the prediction moment.

The invention discloses an ultra-low orbit determination method and device for a near-earth satellite, a medium and a product. Comprising the following steps: acquiring a thrust vector of a satellite engine under a satellite body coordinate system under the condition that the thrust of an ion propeller of a satellite is constant; obtaining a first rotation matrix of the satellite attitude angle relative to the geocentric inertial coordinate system, and converting the thrust vector to the geocentric inertial coordinate system based on the first rotation matrix to obtain thrust components of the thrust vector on three coordinate axes of the geocentric inertial coordinate system; combining the satellite mass, the real-time on-off state of the engine and the thrust components on the three coordinate axes to obtain a small thrust perturbation item; and adding a small thrust perturbation item on the basis of the atmospheric resistance perturbation item and the conventional resistance perturbation item to obtain an input parameter, and inputting the input parameter to the orbit mechanical model, so that the orbit mechanical model determines the satellite orbit element based on the atmospheric resistance perturbation item, the conventional perturbation item and the small thrust perturbation item.

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.

A modular satellite system is provided comprising a main body member, one or more controllers, a communication system, a datastore, a power unit, and one or more orbital cameras. The controller(s) may be operable to perform a mission instruction. The communication system may be in communication with the controller(s) and with a client terminal. The datastore may store data accessible by the controller(s). The orbital camera(s) may be operable to capture one or more image(s) associated with the mission instruction, which may be defined as captured image(s). The controller(s) may be operable to detect and identify one or more predetermined object(s) in the one or more captured image(s). The controller(s) may generate a mission analytics packet, and the controller(s) may be operable to transmit the mission analytics packet to the client terminal.

Systems, apparatus, and method for momentum management for a spacecraft are disclosed. An example spacecraft includes a body; an appendage moveably coupled to the body; a thruster carried by the body; machine-readable instructions; and processor circuitry to execute the machine-readable instructions to detect a change in angular momentum at the spacecraft during firing of the thruster; and in response to the detection, cause the appendage to move relative to the body to adjust a center of mass of the spacecraft relative to a net thrust vector associated with the thruster.

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 invention belongs to the technical field of aerospace, and relates to an unscented batch processing orbit determination method and system for a continuous low-thrust maneuvering satellite. The method comprises the following steps: acquiring a radar observation vector according to radar measured data or satellite ephemeris data in combination with radar site coordinates; obtaining an initial orbit state at a to-be-estimated moment by adopting a Lambert algorithm according to the radar observation vector; generating an initial expansion state quantity according to the initial orbit state at the to-be-estimated moment; a simplified orbit kinetic equation containing earth non-spherical perturbation, atmospheric resistance perturbation and constant small thrust acceleration under an RTN coordinate system is constructed under the ECI coordinate system; based on the simplified orbit kinetic equation, obtaining an expansion state quantity precision value of the to-be-estimated moment through the initial expansion state quantity, the radar observation vector and an unscented batch processing filtering method; and substituting the precision value of the expansion state quantity of the to-be-estimated moment into the simplified orbit kinetic equation to carry out orbit forecasting at other moments. According to the method, the orbit determination precision is remarkably improved, and the influence of single-point observation errors is 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 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 method for constructing and screening orbital transfer speed increment required by on-orbit service of a pulse thrust spacecraft based on a CW equation, belongs to the field of on-orbit service of spacecrafts, and solves the problem of low precision of a maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft. Comprising the following steps: constructing a relative motion equation of a pulsed thrust spacecraft relative to space debris based on a CW equation; inputting the orbital transfer state transfer matrix and the speed transfer matrix into the relative motion equation at the same time to obtain an on-orbit service orbital transfer model; setting a termination state distance constraint, a termination state illumination angle constraint and an orbital transfer speed increment constraint according to task requirements; and inputting all the termination moment position vectors conforming to the termination state distance constraint into the on-orbit service orbital transfer model, and reserving orbital transfer speed increments conforming to the termination state illumination angle constraint and the orbital transfer speed increment constraint to obtain a screened orbital transfer speed increment solution set. According to the invention, the precision of the maneuvering orbital transfer strategy of the on-orbit service orbit of the pulse thrust spacecraft is improved.

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 simulation verification method for a small-thrust orbital transfer strategy of an all-electric propulsion satellite platform, and relates to the technical field of spacecraft orbit control and simulation. In order to solve the defect that a technical means for realizing high-precision simulation and verification of a small-thrust orbital transfer strategy of an all-electric propulsion satellite is lacked on the premise of comprehensively considering multi-perturbation influence in the prior art, the technical scheme provided by the invention comprises the following steps: acquiring a preset input model, inputting six orbital elements, thruster parameters and orbital transfer strategies of the satellites; based on a preset input model, an orbital transfer stage is judged, and an orbital transfer strategy is generated; acquiring a preset orbit dynamics model, calculating a thrust vector according to an orbit transfer strategy, and updating orbit parameters and mass consumption of the satellite; a preset attitude control model is collected, the attitude quaternion of the current satellite is solved, and whether the satellite attitude is stable or not is judged according to the solving result. The method is suitable for design and verification work of a small-thrust long-period orbital transfer strategy of an all-electric propulsion satellite platform.

The invention discloses an active orbital transfer method, device and product of an ultra-low orbit satellite, and relates to the technical field of satellite detection. Comprising the steps of inputting a first instantaneous orbit element and an attitude strategy to an orbit mechanical model under the condition that the first instantaneous orbit element of a satellite at a first moment and the attitude strategy of the satellite in an adjustment period are obtained, and obtaining a second instantaneous orbit element of the satellite at a second moment; converting the second instantaneous orbital element into an average orbital element, and making a difference between a semi-major axis parameter in the average orbital element and a target average orbital semi-major axis to obtain a predicted attenuation height of the satellite in the adjustment period; if the predicted attenuation height is smaller than zero, the next adjustment period is started at the third moment, and the interval duration between the first moment and the third moment is T / 2; if the predicted attenuation height is larger than zero and smaller than or equal to the first attenuation threshold value, the satellite is controlled to rise the predicted attenuation height between the first moment and the third moment, and the next adjustment period is started at the third moment.

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 relates to the technical field of spacecraft orbit control, and particularly discloses an orbit control plan generation method and system, electronic equipment and a storage medium, and the method comprises the steps: constructing an orbit control instruction set containing all preset orbit control instructions during the in-orbit flight of a target spacecraft, determining a target orbit control strategy of the target spacecraft in a target sub-time period after the preset time; determining each preset orbit control instruction screened out from the orbit control instruction set as a first orbit control instruction to be executed within a target time period after the preset moment according to the instruction issuing route, the spacecraft state and the orbit control mode of the target spacecraft at the preset moment; and based on the target orbit control strategy, the target order issuing time and the orbit control name, correcting the instruction basic parameter of each first orbit control instruction, and generating a target orbit control R sequence. According to the method, time consumption of orbit control preparation processes such as planning, countersigning and data generation can be reduced, the efficiency of orbit control planning is improved, and the method can be used as a backup means for collision avoidance in emergency situations.

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.

An azimuth ambiguity suppression method for a spaceborne SAR in spatially non-uniform sampling relates to the technical field of radars. The method includes the following steps: step 1: compensating for a phase difference between the channels according to the locations of transmitting and receiving stations for channel echoes; step 2: calculating an equivalent azimuth sampling location of each channel at the moment of each pulse transmission; step 3: sorting the equivalent azimuth sampling locations of all channels and their corresponding echoes after the phase difference between the channels is compensated in ascending order; and step 4: calculating reconstructed echoes from the sorted equivalent azimuth sampling locations and corresponding echoes after the phase difference between the channels is compensated, by using an adaptive weight-conjugate gradient-Toeplitz matrix (ACT) algorithm. According to the present disclosure, a relative distance between the channels is not required to remain unchanged, which greatly expands the scope of application of azimuth multi-channel SAR echo reconstruction.

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 discloses a quaternion-based rotation separation attitude adjustment method, which comprises the following steps of: aiming at attitude requirements and angular velocity requirements of satellite separation, dividing a program angle planning process into two stages: a first stage is a directional attitude adjustment stage, and forward backward pushing is carried out through a satellite separation time point to obtain an attitude quaternion corresponding to the angular velocity of 0; taking the result as a program quaternion at the attitude adjustment ending moment; the second section is a rotation starting, stabilizing and racemization section, and according to the angular speed requirement, the quaternion of a rotation starting, stabilizing rotation and racemization program is designed; after the attitude of the rocket body is adjusted to the attitude adjustment end moment program quaternion, the angular velocity is linearly accelerated from 0 to the angular velocity required by separation according to the rotation starting, stable rotation and racemization program quaternion, and the satellite separation attitude and the angular velocity both reach expected values at the separation moment. According to the method, a universal solution is provided for one-time safe separation of multiple satellites, the problem of attitude solution ambiguity in the attitude adjusting process is avoided, high universality is achieved, and the reliability of one-time safe separation of the multiple satellites is guaranteed.

The invention provides a flywheel unloading method based on an electric pushing mechanical arm in a position protection period. The flywheel unloading method comprises the steps that a joint corner combination is selected when the thrust direction is in a position protection allowable range; the control torque under the joint rotation angle combination of each mechanical arm is determined, and an initial table look-up matrix is established; determining an optimal torque and a torque error allowable range; according to the control torque, the optimal torque and the torque error allowable range, grade division is conducted on the joint corner combination; constructing a cost function according to the optimal torque and the control torque; on the basis of grade division, the smaller the cost function is, the higher the priority of the joint corner combination is; selecting a joint corner combination with the highest priority from the initial look-up table matrix, performing permutation and combination of joint corners around the joint corner combination in a fine step length manner, determining a corresponding control moment, forming a second-order look-up table matrix, and selecting a joint corner combination with the highest priority; and circularly selecting the joint corner combination with the highest priority to obtain a final joint corner combination.

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.