520 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 the technical field of low-orbit satellite measurement and control, and provides a low-orbit satellite measurement and control method, a satellite, electronic equipment and a medium, and the method comprises the steps: obtaining the space measurement and control signals of a reference satellite and a to-be-positioned satellite, and the orbit position data of the reference satellite; determining orbit information of the satellite to be positioned based on the space measurement and control signal and the orbit position data; calculating orbit information of other satellites to be positioned through a preset orbit positioning method and the orbit information of the positioned satellites through AI to obtain networking information of the whole satellite constellation; telemetry data sent and received through networking information among satellites are monitored in real time through AI sensing; and performing satellite measurement and control based on telemetry data. Therefore, measurement and control tasks are completed on the satellite, construction of ground measurement and control satellite stations can be reduced, manual uninterrupted measurement and control management can be reduced, economic cost can be reduced, pollution of space electromagnetic interference environment and harm to human bodies can be reduced, and intelligentization, modularization, fast response and low cost of a satellite measurement and control system can be realized.

The invention relates to the field of data processing, in particular to an on-orbit calibration method and device for a scale factor, and the method comprises the steps: obtaining attitude data of a spacecraft after maneuvering through a star sensor, converting the attitude data into an angular acceleration signal under a test mass coordinate system, constructing a Kalman filter in combination with an excitation voltage read by an inertial sensor, and carrying out the on-orbit calibration of the scale factor; on-orbit calibration of the angle scale factor is completed, and on-orbit calibration of the line scale factor is completed through the conversion relation between the angle scale factor and the line scale factor. By combining and considering related factors such as linear acceleration and angular acceleration, the calibration process can more comprehensively consider the motion state of the spacecraft, and more accurate calibration is provided. In the maneuvering process of calibrating the scale factor, other parameters such as centroid offset and rigidity of the inertial sensor of the spacecraft can be calibrated at the same time. Through the technical scheme of the invention, the additional mass increase can be reduced, the unnecessary control complexity is avoided, the reliability of the system is improved, and the cost is reduced.

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.

The invention belongs to the field of synthetic aperture radars, and discloses a high-orbit SAR sliding bunching working mode design method and system. The method comprises the steps that 1, resolution improvement factors of a high-orbit SAR in different orbit satellites are calculated; 2, according to resolution improvement factors of the high-orbit SAR in different orbit satellites, calculating a sliding speed of a sliding bunching mode antenna beam on the ground; 3, calculating the antenna beam angle scanning speed according to the sliding speed of the antenna beam in the sliding bunching mode on the ground; and 4, calculating the scanning angle of the antenna beam according to the observation breadth required by the azimuth. According to the method, the problem that a traditional sliding bunching working mode design and calculation method of the high-orbit SAR fails is solved, the influence of satellite speed time varying and earth rotation on the sliding bunching working mode design of the high-orbit SAR is considered, the precision of the working mode design of an SAR system can be greatly improved, and the method has important application prospects.

The invention belongs to the technical field of space photoelectric detection, and provides a high-precision tracking and pointing system and method capable of autonomously formulating an observation task in orbit. The system comprises an information processing unit, a power supply management unit, a task management unit, a thermal control unit, a GNSS unit, a star sensor unit, a gyroscope unit, a tracking and pointing calculation unit and a servo turntable unit. The system can receive a platform instruction, integrate information of sensors such as a GNSS, a star sensor and a gyroscope, calculate the accurate position of a tracked celestial body, and adjust the direction of a servo turntable in real time so as to accurately track a target. The working method comprises the steps of system starting and initializing, observation task making, tracking and aiming mode preparing, aiming direction correcting, observation target tracking and aiming, exception handling and the like. The system can realize high-precision observation of a celestial body target without depending on attitude adjustment of a spacecraft platform, improves the observation efficiency and reliability, and is suitable for observation of celestial bodies such as pulsar on a space station platform.

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 invention discloses a pseudo-spectral convex optimization-based method for optimizing a low-thrust trajectory near an asteroid, which comprises the following steps of: establishing a kinetic equation of an orbital motion of a spacecraft near the asteroid, discretizing the kinetic equation through a flipped Radau pseudo-spectral method, establishing a problem of no gravitational field and no anti-collision constraint, and solving by using a convex optimization method to obtain an initial nominal solution; constructing a gravitational field model in the kinetic equation based on a four-mass-point model, and establishing an asteroid nearby anti-collision model; establishing a convex optimization problem of an anti-collision small-thrust trajectory near the asteroid through convex methods such as nominal value replacement and Taylor expansion; establishing a successive solution model, and obtaining an optimal solution through sequence convex optimization; according to the method, a flipped Radau pseudo-spectrum method and convex optimization combined pseudo-spectrum convex optimization method is adopted, the motion low-thrust trajectory optimization of the spacecraft near the asteroid is realized, and the calculation efficiency is improved by using a four-mass-point gravitational field model.

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 provides a side-by-side multi-satellite formation method capable of realizing real-time high-precision relative measurement, and the method comprises the steps: building a formation multi-satellite relative motion model based on an orbital kinematics theory, obtaining a mathematical expression of a relation between a baseline vector and six orbital elements through derivation and simplification, and building a direct corresponding relation between a baseline requirement of a formation system and an orbital parameter; according to the method, a constraint relation of a baseline measurement system view field requirement to a baseline vector is established, and the constraint relation and a baseline expression are applied to formation satellite orbit parameter design and orbit maintenance strategy design to obtain an orbit real-time adjustment criterion, so that a formation system can realize inter-satellite measurement at any moment. The method can solve the problem of distributed multi-satellite formation system orbit and configuration design oriented to inter-satellite baseline requirements, solves the problem of baseline common-view shielding during operation of the formation system, can realize on-satellite autonomous task planning and execution in the whole process of in-orbit operation, is high in efficiency and speed, is beneficial to flexible task response, and is suitable for large-scale popularization and application. The method has practical engineering significance.

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 relates to a low-orbit star network-oriented random target tracking estimation method, which comprises the following steps of: establishing a dynamic filter AUKF of multiple targets in a geocentric inertial coordinate system by taking position, speed and thrust acceleration vectors of the targets j (j = 1, 2,..., Nt) in the geocentric inertial coordinate system as state vectors; establishing a maneuvering target space-based observation model based on the measurement equation of the target under the geocentric inertial coordinate system; based on the calculation result of the maneuvering target space-based observation model at the current moment, constructing a maneuvering starting and ending detector, when it is judged that the target starts maneuvering, using a maneuvering dynamics filter AUKF to track and estimate the position, speed and maneuvering acceleration information of the target, and when it is judged that the maneuvering of the target is ended, using the maneuvering dynamics filter AUKF to track and estimate the maneuvering speed of the target. And a standard Kalman filter UKF is switched to carry out state updating on the position and the speed of the target. According to the method, the relay observation and whole-course tracking performance estimation problems of the low-orbit star network on the uncertain maneuvering target are solved, and the method can be used for verification and evaluation of a large-scale star network system detection simulation test.

The invention provides a moon orbit formation navigation microsatellite inter-satellite measurement system and method, and relates to the technical field of deep space exploration, the system comprises a master satellite and a slave satellite, and a Ka inter-satellite link is established between the master satellite and the slave satellite; the master satellite and the slave satellite both send timing inter-satellite measurement signals, receive timing inter-satellite measurement signals from the opposite side, measure the time difference between a local timing signal and the received timing signal of the opposite side to obtain the pseudo-range of the satellite, and the two satellites exchange the pseudo-range through an inter-satellite link to solve the inter-satellite distance and the time difference. Accurate time synchronization of two satellites is not needed, a measurement result can be directly obtained on the satellites, ground calculation is not needed, and ground resources are saved; the satellite platform and the antenna design of the slave satellite are simplified, and the weight and the cost of the satellite are reduced.

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 an intelligent prediction guidance method for Mars pneumatic auxiliary orbit descending, and belongs to the field of deep space exploration. According to the method, a high-precision kinetic model of the Mars pneumatic-assisted orbit descent is constructed, a network training data set is generated by utilizing the high-precision model, a near-fire-point state intelligent forecasting model and a far-fire-point state intelligent forecasting model of the Mars pneumatic-assisted orbit descent are constructed, and meanwhile, near-fire-point heat flow constraints are constructed; designing a predictive guidance method by taking the orbit height deviation and the control input as optimization performance indexes; and optimizing and solving the control input to control the detector until the track height of the detector is reduced to be close to the preset track height, so that the detector is reduced to the preset track under the condition of meeting heat flow and control amplitude constraints, and meanwhile, the fuel consumption of the detector is reduced. According to the method, by solving the states of the near fire point and the far fire point under the two-body dynamic model, the network is constructed to predict the deviation value between the states of the near fire point and the far fire point under the high-precision dynamic model and the prediction result of the two-body model, and the training cost of the neural network is reduced.

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 an automatic operation lunar ladder modeling method and system based on a gravitational siphon effect. The method comprises the following steps: constructing a gravitational siphon lunar ladder model; performing bead point simplification processing on the gravitational siphon moon ladder model to obtain a multi-body dynamics simplified model of the gravitational siphon moon ladder; and performing coordinate system construction processing and dynamic analysis on the multi-body dynamic simplified model of the gravitational siphon moon ladder to obtain a gravitational siphon effect moon ladder dynamic model. By using the system, the transfer of substances from the surface of a rotating body to an orbit collection spacecraft can be realized by utilizing a gravitational siphon effect, so that the requirement of a traditional propulsion system is avoided, and the energy consumption required for improving the load is greatly reduced. The automatic operation lunar ladder modeling method and system based on the gravitational siphon effect can be widely applied to the technical field of aerospace modeling.

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.