380 results about "Orbit" patented technology

The invention discloses an on-orbit spacecraft temperature prediction method and device based on a physical information neural network. The method comprises the following steps: S1, constructing an object model and an orbit heat source model according to a spacecraft to be predicted; s2, solving a space external heat flow of the orbit heat source model, and converting the space external heat flow into a volume heat flow of the object model; s3, constructing a physical information neural network; wherein a three-dimensional transient heat conduction equation is embedded in the physical information neural network; s4, designing a total loss function according to the volume heat flow and the physical information neural network; and S5, training the physical information neural network according to the total loss function to obtain a temperature prediction result of the spacecraft to be predicted. According to the method, the situation that model grids need to be divided in traditional finite element simulation is avoided, the calculation efficiency is improved, the calculation time is shortened, the temperature result can be predicted in real time, and the generalization ability and accuracy of the model are improved.

The invention relates to a conduction fusion enhancement method suitable for an earth-moon space high-orbit spacecraft, and the method comprises the steps: transmitting a communication and navigation fusion enhancement signal through a designed and planned HEO satellite constellation, and receiving and processing a Beidou navigation signal and an HEO enhancement signal in a compatible manner on the basis of not changing the hardware structure of an earth-moon space navigation receiver. The number of available navigational stars is increased through an observed quantity combination mode, the geometric distribution of constellations is improved, and the navigation solution accuracy and the positioning precision of earth-moon space are improved. And meanwhile, a communication function of forwarding information such as navigation ephemeris, receiver position and user customized data of uplink injection on the ground through an HEO constellation is supported, auxiliary information is provided for the earth-moon space navigation receiver, the receiving processing performance is improved, and communication and navigation fusion processing is realized. According to the method, the performance and the service capability of the high-orbit satellite navigation receiver can be improved, and the method has certain reference significance for spacecraft conduction fusion and integrated design.

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 provides a radar and optical observation segmental arc correlation method and system for space debris. A joint probability density function is obtained through a kernel density estimation algorithm. A radar observation arc section and an optical observation arc section are input. And converting the observation data into a geocentric celestial coordinate system. And acquiring system parameters of the observation equipment. Radar observation data and optical observation data are unified to the same dimension, and a distance measurement value weight and an angle measurement value weight are set. And calculating an initial orbit by the radar observation arc section. And calculating a joint objective function corresponding to the current orbit position speed. And if the convergence condition is met, outputting the minimum value of the joint objective function and the orbit position speed corresponding to the minimum value. Otherwise, updating the track position speed, and repeating until the convergence condition is met. And according to the minimum value of the joint objective function and the joint probability density function, judging whether association is carried out, and outputting confidence during association. The method is easy to implement and can be effectively applied to association of the space debris radar and the optical observation arc section.

The invention provides an optical navigation method and device for an earth-moon spacecraft, electronic equipment and a storage medium. The method comprises the steps that an optical image of the moon is shot through image acquisition equipment, the optical image is analyzed according to imaging parameters of the image acquisition equipment, the angle diameter of the moon is obtained, the target distance of the earth-moon spacecraft relative to the moon is calculated according to the angle diameter of the moon, and the optical image comprises the surface of the moon; measuring starlight angular distance information between the earth-moon spacecraft and the moon by adopting a star sensor, wherein the starlight angular distance information is used for representing the angular distance between the navigation star and the centroid of the moon detected by the star sensor; and calculating the current orbit position and the current attitude of the earth-moon spacecraft by adopting a navigation solution algorithm according to the target distance and the starlight angle distance information.

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 application discloses a fusion positioning algorithm based on low-orbit satellite Doppler measurement, DME and VOR, relates to the technical field of satellite navigation, and comprises the following steps: acquiring the Doppler frequency shift of a low-orbit satellite, the slant range information of a DME and the azimuth angle data of a VOR, constructing a combined observation equation, and obtaining a corresponding Jacobian matrix; based on a geometric dilution of precision, selecting stations and stars through a genetic algorithm, optimizing the selection of the DME, the VOR and the low-orbit satellite, and obtaining an optimal navigation source combination; constructing a motion model under different flight modes, and dynamically estimating the state of an aircraft by using an interactive multiple model Kalman filter, and updating the position information in real time. Therefore, the real-time positioning of the aircraft under complex conditions is realized, the anti-interference capability of the system is improved, and the overall positioning precision and stability of the system are improved.

The invention belongs to the technical field of aerospace, and relates to a TLE parameter prediction method and system based on TFT multi-source data fusion. The method comprises the following steps: generating a feature data set according to a historical TLE data set of an LEO uncontrolled space target, target static feature data and space weather data; adopting a # imgabs0 # model to obtain an angle parameter at the prediction moment according to the TLE data at the moment before the prediction moment; adopting a # imgabs1 # model to obtain core orbit parameters at a prediction moment according to the feature data set; acquiring a temporary TLE orbit parameter at the prediction moment according to the angle parameter at the prediction moment and the core orbit parameter at the prediction moment; updating a mean anomaly prediction value according to the current prediction step number and the temporary TLE orbit parameter at the previous moment; and replacing the mean anomaly preliminary prediction value in the temporary TLE orbit parameters at the prediction moment with the mean anomaly prediction value to obtain the TLE orbit parameters at the prediction moment. And the precision, real-time performance and efficiency of long-term orbit parameter prediction are improved.

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.

The invention discloses an air target optical image real-time simulation method and system suitable for approaching operation. The method comprises the following steps: receiving target and observation star motion orbit data generated by an orbit or attitude generator in real time; calculating a camera optical axis direction; screening out fixed stars which can be detected by the camera from the SAO star catalogue by combining the limit detection capability of the camera, and completing fixed star simulation; calculating the relative pose of the observation star and the target according to the data; performing imaging simulation on the target according to the characteristics of the space target; whether the moon appears in the visual field or not is judged, and the situation that the moon exists in the visual field is simulated; adding noise and stray light in the image; and outputting in a UDP transmission mode through an Ethernet port. According to the method, various space targets can be simulated, relative change of space illumination is considered, calculation is real-time and efficient, rapid image generation and transmission capacity is achieved, and vivid target image input can be provided for ground semi-physical performance verification of a space visual navigation system.

The invention provides an in-orbit star-observation calibration and correction method and system for geometric thermal stability of a space camera, and the method comprises the steps: calculating the positioning error of each star point in an along-orbit direction and a vertical-orbit direction through the star-observation imaging of a satellite under different initial temperature conditions, recording the satellite temperature and imaging time, and carrying out the calibration of the geometric thermal stability of the space camera. Fitting to obtain a temperature variation coefficient and a time variation coefficient of the exterior orientation element thermal stability of the space camera, and fitting to obtain a space camera interior orientation element zoom coefficient of each star observation task according to the positioning error of each star point after correction of the exterior orientation element thermal stability and the distribution condition of the star points in the camera field of view; the temperature change coefficient of the space camera interior orientation element thermal stability is obtained, and in-orbit calibration and correction of the space camera interior and exterior orientation element thermal stability are achieved. The method does not need a ground geometric field, has the advantages of high precision, high timeliness and low cost, and has good adaptability.

The invention relates to the technical field of satellite navigation, in particular to a cycle slip detection method, device and equipment for a low-orbit navigation enhanced carrier observation value, and the method comprises the steps: constructing a first target cycle slip detection quantity and a second target cycle slip detection quantity based on a low-orbit satellite carrier phase observation value, and determining a cycle slip detection threshold value according to a low-orbit satellite high angle, and performing cycle slip detection on the low-orbit satellite carrier phase observation value according to the first target cycle slip detection quantity, the second target cycle slip detection quantity and a cycle slip detection threshold value to obtain a cycle slip detection result, and performing cycle slip marking on the low-orbit satellite carrier phase observation value according to the cycle slip detection result to obtain a cycle slip detection condition. Therefore, the problems of a large amount of misjudgment on the cycle slip of the low-orbit carrier observation value caused by the change of the ionosphere error in the carrier observation value in a short time and the like are solved, two cycle slip detection quantities are given based on the characteristic of rapid change of the ionosphere error, and the cycle slip detection threshold is determined in combination with the satellite height angle. Therefore, the low-orbit carrier cycle slip phenomenon can be accurately detected.

The invention discloses a tethered satellite system dynamics modeling method considering perturbation factors. The tethered satellite system dynamics modeling method comprises the steps that a geocentric coordinate system, an orbital coordinate system of a Kepler orbit where a tethered satellite system is located and a body coordinate system of a unit satellite are established; according to the movement of the centroid of the tethered satellite system on the orbit, the generalized coordinates are the distance from the centroid of the system to the earth center and the true anomaly of the orbit, the Lagrange function of the centroid of the system is obtained; according to the movement of the unit star relative to the centroid, the generalized coordinates are the distance between the unit star and the centroid, the relative first rotation angle and the relative second rotation angle, the relative Lagrange function of the unit star is obtained; comprehensively considering perturbation factors to obtain non-spherical earth, third gravitational force and solar pressure perturbation potential energy; establishing a system kinetic equation considering comprehensive perturbation factors; and obtaining a complete Lagrange equation of the tethered satellite system. The error problem caused by the fact that the master satellite directly serves as the mass point of the system and the influence of comprehensive perturbation on system motion is not considered is solved, and the number of the slave satellites is not limited.

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.

A method and system for low Earth orbit (LEO) satellite multiple access and communication using grating lobes is disclosed. The method may include: acquiring ephemeris data for a plurality of LEO satellites; predicting an orbital trajectory for each of the plurality of LEO satellites based on the ephemeris data; selecting target satellites using the predicted orbital trajectories; adjusting an antenna array to form grating lobes in the directions of the selected target satellites; and transmitting and receiving data with the target satellites using the adjusted antenna array.

The invention discloses a spacecraft in-orbit operation health assessment method, and relates to the technical field of space flight and aviation. Environmental data such as gravity gradient, radiation intensity and dust particle density of an asteroid surface are collected in real time through a multi-mode sensor array, modeling and prediction are carried out by adopting a dynamic Bayesian network in combination with Kalman filtering, and the health assessment accuracy of the spacecraft in-orbit operation is improved. The method comprises the following steps: dynamically generating environment characteristic parameters including a risk level and an environment change trend, and in health state prediction, simulating stress distribution in a low-gravitation environment through digital twinborn body and finite element analysis, predicting a fatigue state and a failure risk of a contact device, and accurately evaluating the fatigue life of a high-stress area. The quantification capability of the failure risk is improved through the risk probability model; besides, according to a health prediction result, landing parameters including a damping mode and a landing speed are dynamically optimized, and a dynamic planning method is used for minimizing impact force and stress load, so that the detector can be in stable contact in a low-gravity environment.

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.

The invention provides a correlation method and system for space debris observation data and cataloguing orbit data. The method comprises the following steps: firstly, within a time range of currently processed batch observation data, recurring to an orbit matrix distributed at equal time intervals through an orbit of an orbit recursion cataloguing target, and generating a position interpolation function based on a time sequence and the orbit matrix; and for observation data in a processing time range, firstly unifying the observation data and an observation station position to a J2000.0 inertial coordinate system, and obtaining a multi-target position prediction value corresponding to observation time through an interpolation function. Calculating the difference between the predicted value and the measured value of the observation data; and calculating an associated object and confidence according to the difference. The method is easy to implement, high in efficiency during batch observation data processing and suitable for association of observation data of different observation devices and different types of orbit space debris and catalogued orbit data.

The invention provides a linear array camera exterior orientation element in-orbit calibration method and system based on fixed star observation. The method comprises the following steps: step S1, fixed star shooting and star point extraction; s2, calculating a visual coordinate of the shooting star point in the J2000 coordinate system; s3, carrying out rough matching on the shot star points and a real fixed star, and removing star points which are not successfully matched; s4, calculating real fixed star visual coordinates after aberration correction; s5, resolving a coarse compensation matrix of exterior orientation elements of the camera; and S6, finishing fine matching of the fixed star, and determining a fine compensation matrix of exterior orientation elements of the camera. According to the method, a two-stage matching mode of the shot star point and the real fixed star is adopted, namely rough matching is carried out firstly, then fine matching is carried out, the fixed star matching efficiency and matching precision are improved, and the method can be widely applied to on-orbit calibration of the exterior orientation elements of the linear array camera and has the advantages of high precision, high timeliness and low cost.

The invention relates to a lunar satellite formation orbit adaptive estimation method based on EM-EKF. Comprising the steps that an extended Kalman filtering module of a navigation algorithm predicts the orbit state of the lunar satellite formation at the current moment based on the orbit state of the lunar satellite formation at the previous moment, observation data at the current moment, a process noise covariance matrix and an observation noise covariance matrix; and updating a process noise covariance matrix and an observation noise covariance matrix by an expectation maximization module of a navigation algorithm based on a state sequence formed by a series of orbit states output by the extended Kalman filtering module within a period of time. According to the method provided by the invention, the process noise covariance matrix and the observation noise covariance matrix can be dynamically estimated, and online adaptive updating of the extended Kalman filtering parameters is realized.

The invention discloses a method and a system for predicting luminosity of a low-earth-orbit satellite during transit, and the method comprises the steps: obtaining historical observation data, and constructing a comprehensive parameter data set of luminosity, position and related observation information of the low-earth-orbit satellite at different observation moments; equivalently representing the optical scattering sectional area of the low-orbit satellite by using the sectional area of the sphere generating the same brightness; key factors influencing the equivalent sectional area are determined; integrating the obtained parameterized relationship of the equivalent sectional area into a ball model formula to obtain a luminosity prediction model; taking the obtained comprehensive parameter data set as the input of the luminosity prediction model, and optimizing the luminosity prediction model through a data driving method to obtain an optimized luminosity prediction model; and substituting to-be-predicted low-earth-orbit satellite data into the optimized luminosity prediction model to obtain a luminosity prediction value of the low-earth-orbit satellite in the transit period. According to the method, the correction ball model is fitted through observation data, the luminosity prediction model is constructed in combination with key parameters, and the brightness of the low-orbit satellite is accurately predicted.

The invention relates to the technical field of camera on-orbit calibration, in particular to a fisheye camera on-orbit calibration method based on a starlight vector, which comprises the following steps of: S1, acquiring a plurality of star maps observed by a fisheye camera on orbit, and extracting and denoising star points to form a star point set; s2, according to fisheye imaging distortion distribution, a star point located in an imaging center area is selected from a star point set, and a starlight vector is constructed based on image point coordinates of the star point in combination with star catalogue data subjected to space-time deviation and propagation path deviation correction; and S3, performing local geometrical relationship star point matching on the starlight vector in the step S2. According to the method, the multiple star maps observed by the fisheye camera in orbit are obtained and processed to form the star point set, so that the problems of poor in-orbit calibration applicability and low precision caused by the fact that most traditional fisheye camera calibration methods adopt a ground pre-calibration or space reference object arrangement mode are solved.

The invention discloses a multi-scene high-precision foundation optical imaging processing method based on focal length inverse solution and degradation processing, and belongs to the technical field of high-precision space target observation and astronomical research. The method comprises the following steps: collecting actual parameters under different orbit conditions to build a simulation space environment, adding targets according to actual requirements, simulating motion laws of space targets under different orbit conditions, and exporting a short-exposure image based on orbit dynamics and changes of an observation environment; performing superposition operation on the exported short-exposure images according to a time sequence to simulate a long-exposure image; constructing a degradation model, and performing degradation processing on the short-exposure image and the long-exposure image; and based on a peer-to-peer relationship between a sensor measurement view field and an actual observation view field and a camera imaging principle, reversely solving a focal length, and calculating angle measurement information of the space target in the degraded image. The method not only can obtain a high-quality ground-based image, but also can be used for generating a space-based image.

The invention discloses a zero-initial-value guess earth-moon Halo orbit low-thrust rendezvous trajectory optimization method, belongs to the field of low-thrust rendezvous trajectory optimization, and solves the problems that an optimal result cannot be ensured, the calculation time is long, the robustness is poor and the initial value sensitivity is poor. Comprising the steps of dispersing an original control trajectory and a state trajectory to obtain a target function meeting rendezvous task constraints; solving an intermediate variable of the target function, and calculating a discrete costate quantity through a costate mapping equation; in combination with the Pontryagin minimum principle, optimal control of a tracking spacecraft is obtained through Hamiltonian, and a two-point boundary value model is constructed; taking the discrete co-state quantity as an initial guess of the two-point boundary value model, and optimizing the discrete co-state quantity by using an SQP algorithm to obtain an output result; the kinetic equation and the co-state equation are positively integrated through the output result, and the optimal control rate and the small-thrust rendezvous trajectory of the tracking spacecraft are obtained; the method overcomes the problem of initial value sensitivity, shortens the calculation time, and ensures the optimal result and the robust calculation process.

The present disclosure provides a satellite system with orbital debris avoidance. In one embodiment, the satellite system includes debris sensor circuitry to scan orbital debris items in an orbital debris field; and sensor controller circuitry to determine a search window within the orbital debris field based on, at least, uncertainties associated with a velocity of the satellite (Vsatellite) and uncertainties associated with a velocity of at least one debris item (Vdebris), the sensor controller circuitry also to control the debris sensor circuitry to detect the at least one debris item within the search window.

The invention discloses a fly-around path autonomous planning method, which comprises the following steps of: 1, drawing up a mark track in a fly-around process as an elliptical track, and solving a minor semi-axis, a major semi-axis and a total fly-around angle of the elliptical track according to the current relative position and the preset tail end relative position of a tracking satellite and a target satellite; step 2, obtaining a motion mode of the aircraft on an elliptical orbit, and obtaining a formation vector design of an elliptical fly-around plane according to a minor semi-axis, a major semi-axis, a total fly-around angle, a total fly-around time and a current fly-around time of an ellipse, the formation vector comprising a position and a speed of a tracking star under a target star orbital coordinate system; and step 3, rotating the formation vector of the fly-around plane into an expected plane according to three parameters of the set fly-around elliptical arch line argument, the fly-around ascending intersection angle and the fly-around inclination angle. According to the method, the planned path is used as a nominal track, and flying-around approaching to the target or flying-around evacuating from the target can be realized by applying control.

The application provides a debris cloud disintegration timing inversion method based on space-time intersection feature statistics, belongs to the field of disintegration timing inversion, and solves the problem of how to improve the accuracy of disintegration timing determination; the method comprises the following steps: traversing the standardized debris cloud orbit data after orbit propagation, calculating the closest distance time of each two pieces of debris; using an Epanechnikov kernel function to perform non-parametric probability density estimation on the closest distance time; combining the inverse function of the cumulative distribution function with the significance level to construct a confidence interval, and removing the debris outside the confidence interval; traversing the debris in the confidence interval, calculating the closest distance time of each two pieces of debris; using an Epanechnikov kernel function to perform non-parametric probability density estimation on the closest distance time; and taking the closest distance time corresponding to the maximum probability density estimation value as the disintegration time of the space target; the application has strong universality and strong robustness.

The present invention relates to the technical field of data processing, and discloses a method and device for on-orbit calibration of the centroid offset of an inertial sensor. The method includes: obtaining the angular velocity and angular acceleration of the spacecraft after maneuvering through the inertial sensor; constructing a measurement observation model by using the angular velocity and angular acceleration of the spacecraft; processing the measurement observation model through a pre-constructed Kalman filter to obtain a predicted centroid offset; and calculating based on the predicted centroid offset and the acceleration output by the inertial sensor to obtain a calibration result of the centroid offset between the centroid of the inertial sensor and the center of gravity of the spacecraft. Through the above steps, the calibration of the centroid offset can be achieved within the short-period maneuver of the spacecraft. By means of one maneuver, the centroid offset and other parameters that need to be calibrated on orbit are calibrated simultaneously, reducing the maneuver cost.

The application discloses a small-sized separated penetrator for in-situ detection of celestial bodies, and relates to the technical field of penetrator equipment, comprising an aluminum honeycomb buffer module, which comprises an external protective shell and an aluminum honeycomb, and has the beneficial effects that the application is provided with in-situ detection of high-speed penetration of celestial bodies, and discloses a small-sized separated penetrator, which comprises a planet soil penetration module and a star table detection module, and realizes autonomous separation in the impact penetration process, the star table module can realize star table detection and communication with an orbiter, the penetration module is internally provided with an energy storage device and an impact hammer, and can continue to penetrate and dive relying on the energy storage device after impact; the star table module is connected with the penetration module through flexible wires, and can supply power and signal transmission to the latter, the technology enables the penetrator with the modular characteristic to have the in-situ detection capability of planet soil profile, aims to expand the application mode of the penetrator for celestial body detection, and provides a scientific basis for the research of star table structures.

The invention relates to a high and low orbit satellite tracking method based on geomagnetic correction. The method comprises the following steps: constructing a geomagnetic model, calculating an azimuth angle and carrying out magnetic field interference analysis; carrying out factory calibration by utilizing a three-dimensional geometric calibration mode, including horizontal correction and vertical correction; the current azimuth angle and pitch angle are continuously corrected based on the carrier coordinate system by using the signal intensity, and the deviation is obtained so as to continuously approach the real angle; utilizing the known coordinates of the three-dimensional space to inversely calculate the current position; and the current data is corrected by using the big data, so that the current data meets the use requirements. The current attitude angle and the satellite signal receiver are used for tracking the target satellite, and then the known azimuth angle is used for correcting the geomagnetic data, so that the azimuth angle precision is improved, the cost of a phased-array antenna or a mechanical antenna is reduced, the tracking precision is improved, and the tracking speed is improved.