503results about "Instruments for comonautical navigation" patented technology

The invention discloses a satellite collision avoidance system and method based on a satellite edge computing network. The system comprises a space monitoring network, a space traffic management subsystem and the satellite edge computing network. The space monitoring network is used for observing the running state of a space target to obtain multi-source observation data; the space traffic management subsystem is used for integrating the multi-source observation data to generate space target orbit data; the satellite edge computing network further comprises the steps that a GEO satellite broadcasts space target orbit data to an LEO satellite constellation and forwards collision risk screening information to a ground cloud computing center; the LEO satellite constellation calculates collision risk screening information according to the space target orbit data and formulates a collision avoidance control strategy; the LEO / GEO earth station is used for data transceiving and protocol conversion among the GEO satellite, the LEO satellite constellation and the cloud computing center; and the ground cloud computing center stores constellation orbit data through protocol conversion and satellite communication. According to the invention, the real-time performance of satellite collision avoidance is improved.

The invention relates to an autonomous navigation method for lunar satellite formation and a second-order filtering navigation architecture, and the autonomous navigation method comprises the steps: obtaining orbit prediction information based on a kinetic model of the lunar satellite formation, and obtaining collaborative theory observation information based on a collaborative observation model, carrying out fusion processing on orbit prediction information, collaborative theory observation information and multi-source observation data of in-orbit observation by using an extended Kalman filtering algorithm, and estimating an absolute orbit state of the lunar satellite formation; establishing a relative motion model of the slave satellite under the local coordinate system of the master satellite, and establishing a relative measurement model; and taking the absolute orbit state of the master satellite as a reference datum, obtaining a preliminary prediction relative orbit state based on the relative motion model of the slave satellite, and obtaining relative measurement data including distance measurement and angle measurement based on the relative measurement model, and fusing the relative measurement data and the preliminarily predicted relative orbit state through an extended Kalman filtering algorithm to estimate the relative orbit state of the slave satellite. Navigation may be independent of ground support.

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.

The invention relates to the technical field of unmanned aerial vehicle route planning, and discloses a low-altitude route risk map construction method and system based on artificial intelligence, and the system comprises a low-altitude environment perception constraint module, a low-altitude route feature extraction module, a dynamic risk constraint judgment module, a low-altitude route map construction module, and a low-altitude route risk analysis module. And the optimal path calibration output module is used for determining a low-altitude air route flight area by acquiring a first constraint condition of low-altitude air route environment perception, extracting key feature parameters from low-altitude air route data, performing second constraint condition analysis on the low-altitude air route, intelligently planning a low-altitude air route route according to an analysis result, and constructing a low-altitude air route map. According to the method, the low-altitude air route is acquired, the constant direction line risk analysis is performed on the low-altitude air route to calibrate the optimal path, the optimal path is output to the user terminal, the path three-dimensional display is performed in the low-altitude air route map, and the execution efficiency and the safety of the flight task are improved.

The invention discloses an on-orbit spacecraft attitude estimation method and system based on ISAR image feature selection, and belongs to the field of aerospace control systems. The method comprises the following steps: acquiring an ISAR image of a spacecraft, and acquiring a complex linear structure set and three-dimensional feature points of an on-orbit spacecraft; then, according to the obtained three-dimensional-two-dimensional projection model, the CRLB of each reference structure in the complex linear structure set is deduced to carry out attitude estimation error analysis; calculating the trace of the CRLB covariance matrix of each reference structure to select an optimal feature structure, correcting the scattering point trace of the optimal feature structure by using polynomial fitting, and switching the reference structures as a new optimal feature structure according to the scattering point loss rate and a preset sequence; and optimizing the spacecraft attitude angle solving function by using a particle swarm and LM hybrid algorithm to obtain attitude angle parameters. The target with high-precision target attitude real-time estimation can be completed aiming at the problems that high-order frequency change in a dynamic environment is difficult to capture and resolution and noise suppression are contradictory due to a fixed window time-frequency analysis method.

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 provides a lunar polar region sun synchronization # imgabs0 # path planning method, which can be applied to the technical field of celestial body image processing and path planning. The method comprises the following steps: independently training a plurality of enhanced U-Net models by using training data with time sequence slices and fusion types and a mixed loss function; dividing and extracting the reasoning data with a time sequence slice type and a time sequence fusion type by using a plurality of trained enhanced U-Net models to obtain a lunar polar region friendly illumination region, and preprocessing the reasoning data with the lunar polar region friendly illumination region and gradient data corresponding to the lunar polar region friendly illumination region; and based on the reasoning data after time alignment and the gradient data after space alignment, constructing a heuristic function by using mixed cost, and performing heuristic path planning by using an improved # imgabs1 # path planning algorithm with a time one-dimensional function and an in-situ waiting function to obtain a sun synchronization planning path of the lunar polar region.

The present disclosure relates to an information processing device and an information processing method for providing an imaging service for performing imaging with a camera mounted on an artificial satellite.An information processing device includes a display control unit that causes a first screen to be displayed, the first screen including a map display unit that displays an artificial satellite that orbits around a celestial object and an orbit of the artificial satellite on a map of the celestial object, and a time information display unit that displays time information on the orbit of the artificial satellite. The display control unit causes the first screen to synchronously display the position of the artificial satellite on the map of the celestial object on the map display unit and the time information about the artificial satellite on the time information display unit. The technology according to the present disclosure can be applied to an information processing device or the like that provides an imaging service for performing imaging with a camera mounted on an artificial satellite, for example.

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 discloses a lunar satellite orbit determination method and device based on radio and laser ranging, and relates to the technical field of lunar satellite orbit calculation, and the method comprises the steps: obtaining observation data of a radio tracking station and a laser station; establishing a high-precision kinetic model of the lunar satellite; establishing a lunar satellite radio and laser ranging mathematical model, and processing observation data according to an observation equation to obtain a kinetic parameter resolving value of the satellite at the initial orbit determination moment; and substituting the kinetic parameter resolving value into the satellite shot motion equation, and resolving to obtain an orbit determination result of the lunar satellite. According to the method, radio ranging and laser ranging are combined for orbit determination, precise orbit determination of the lunar satellite is achieved, compared with a traditional orbit determination mode, the precision and reliability are greatly improved, the requirements for the arrangement number and geometric distribution of ground stations are reduced, and the measurement and control pressure of existing ground stations is relieved.

The invention provides a two-line element generation method and orbit prediction method adaptive to an analytic propagation model, and the method comprises the steps: obtaining the historical osculating orbit elements at each epoch moment, and carrying out the orbit element optimization step of a plurality of iteration rounds on each historical osculating orbit element through employing the analytic propagation model, obtaining a target osculating orbital element number at each epoch moment, enabling the historical osculating orbital element number to be the same as the target osculating orbital element number, and performing inversion to obtain an average orbital element number; and extracting each average semi-major axis of each average orbital element, fitting each average semi-major axis by adopting a least square method to obtain a semi-major axis attenuation rate, inverting the semi-major axis attenuation rate to obtain an atmospheric resistance parameter, and equivalently converting the atmospheric resistance parameter into a standard resistance term in two rows of elements to obtain the two rows of elements. According to the method, the problems of inconsistent models, insensitive residual errors, unable physical explanation of parameters and the like in the two-line element generation process can be solved, and the ephemeris extrapolation orbit prediction precision is improved.

The invention provides an integrated navigation method combining starlight refraction positioning with inertial navigation in a near space, and belongs to the technical field of starlight refraction navigation positioning. The method comprises the following steps: firstly, determining a state model of starlight refraction positioning in a near space by taking an error equation of an inertial navigation system of a near space aircraft as a state equation; designing a near space starlight refraction positioning algorithm, identifying a refraction star by using a star sensor on the aircraft, acquiring a refraction angle as a measurement, and establishing a measurement model of starlight refraction positioning in a near space; and obtaining the optimal estimation value of the system state variable by using an extended Kalman filter (EKF) method. The method has the advantages of being suitable for near space, high in reliability, high in autonomy and the like, and can be finally used for pure astronomical positioning and astronomical / inertial integrated navigation.

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 provides a low-orbit constellation autonomous orbit determination method and system based on Beidou B2b and space between satellites, and belongs to the technical field of low-orbit navigation constellation orbit determination. Comprising the following steps: acquiring Beidou B2b correction data of a low-orbit satellite and correcting the Beidou B2b correction data to obtain an observation data set; obtaining original inter-satellite distance measurement data between two low-orbit satellites by using a bidirectional observation mode, and obtaining an inter-satellite distance solution; a Kalman filtering model is constructed based on the observation data set and the inter-satellite distance solution, and a dynamic orbit is predicted and corrected; and converting the Kalman filtering model into a square root information filtering model, and outputting a high-precision orbit solution through real-time joint filtering. According to the invention, on the basis of reducing calculation delay and burden, completely autonomous navigation and high-precision orbit determination based on inter-satellite links can be realized.

The invention discloses a star point coordinate extraction and star sensor attitude optimization method based on deep learning. The whole process of pseudo star map generation, star point coordinate extraction and attitude calculation is optimized. Multi-factor interference such as spacecraft jitter and optical distortion is simulated through dynamic star point modeling, and a high-fidelity pseudo star map is generated; a noise adaptive detection network is designed to fuse denoising and key point detection, and star point positioning precision in a low signal-to-noise ratio environment is improved; and optimizing attitude solution, constructing a loss function based on star point coordinates, and performing back propagation to optimize attitude parameters to realize an algorithm closed loop. According to the method, real data can be expanded to realize effective training of the model, and the positioning precision is improved; the star point coordinate positioning precision is improved to a sub-pixel level (RMSElt; 0.3 pixel); an end-to-end optimization framework reduces error accumulation in a traditional assembly line, and the robustness of the system is improved by 40%.

The invention discloses a landing positioning method and system based on fusion of vision and inertial navigation. The method comprises the following steps: collecting vision information of a current landing area; the visual information of the front landing area is mapped to an aligned grid of a map coordinate system through homography transformation, correlation matching is conducted on the visual information and a map through a frequency domain FFT algorithm, a matching result of rough matching is obtained, and rough matching is completed; the method comprises the following steps of: mapping an area subjected to rough matching and positioning to a grid aligned with a map coordinate system by utilizing homography transformation, performing spatial domain multi-window division on a larger range of images around the area subjected to rough matching and positioning, selecting a landmark area in each sub-window by applying an interest operator, performing spatial correlation matching on the landmark area and a map by utilizing a spatial correlation method, and performing spatial correlation matching on the landmark area and the map by utilizing a spatial correlation method. Finishing fine matching; inputting all matched observation results into an extended Kalman filter, and carrying out joint estimation on the matched observation results and the observation results of the inertial measurement unit; the method has the advantages that the robustness, the precision and the engineering applicability of an existing navigation positioning method are improved.

The invention provides a satellite earth observation path planning method and device, equipment and a storage medium, and the method comprises the steps: obtaining the earth-surrounding orbit and shooting breadth of a satellite to be subjected to a task, and a to-be-observed task, dividing a predicted observation ground region into a two-dimensional grid chart containing a plurality of uniform squares, carrying out the observation demand representation of the two-dimensional grid chart, and carrying out the observation demand representation of the two-dimensional grid chart; and according to the time sequence change of the around-the-ground orbit, stacking the two-dimensional grid charts represented by the observation demand according to a time sequence to obtain a three-dimensional space-time grid, searching path observation points based on the three-dimensional space-time grid to obtain a target path observation point set meeting the minimization of a pre-constructed optimization target function, and searching path observation points based on the target path observation point set. And performing path conversion on the target path observation point set to obtain an optimal earth observation path for the satellite to execute the to-be-observed task. According to the method, two-dimensional path planning is expanded into path planning on a three-dimensional space-time grid, the dynamic time-varying constraint processing process is effectively simplified, and the efficiency of satellite earth observation path planning is improved.

The present invention belongs to the field of signal processing technology and relates to a non-reconstruction direct positioning method based on mobile receiver data compression. The present invention aims to provide a non-reconstruction direct positioning method based on mobile receiver data compression. This method utilizes the characteristics of the Hadamard matrix to directly estimate the source position in the compressed measurement domain without the need for signal reconstruction and TDOA and FDOA parameter extraction. The present invention demonstrates that signals can still retain TDOA and FDOA information after Hadamard matrix compression, thereby reducing transmission throughput while maintaining positioning accuracy.

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 discloses a dual-path non-cooperative spacecraft pose estimation method based on uncertainty perception, and belongs to the technical field of spacecraft pose estimation. According to the method, a dual-path prediction framework based on a shared backbone network is constructed and comprises a geometric reasoning path and a global context sensing path. In the training stage, prediction results of the two paths are aligned through a geometric consistency loss function, and joint optimization is carried out by combining multi-task losses such as key point regression, classification, rotation and translation regression and the like. In the inference stage, the pose estimation result is adaptively fused according to the predicted variance output by each path, when the variance is low, the output of the global context sensing path is directly adopted, otherwise, the key point projections of the two paths are combined, and the final pose is recovered. The method has higher estimation precision, robustness and interpretability in complex space environments such as shielding, illumination variation and large-scale variation, and provides reliable visual navigation support for on-orbit autonomous tasks.

The invention relates to a satellite group cooperative area positioning method based on communication ranging, and belongs to the technical field of spacecraft navigation. Three reference satellites are selected in a satellite group space area, and a positioning reference is constructed; determining an initial relative motion state Xim0 of any member satellite m in the LVLH system based on a positioning reference; a triangular vector closed loop is formed by the relative positions of the satellite i, the satellite j and the satellite k; according to the triangular vector closed loop, establishing a relative orbit state consistency constraint model; introducing the consistency constraint model into a filtering algorithm, establishing a distributed Kalman filtering algorithm, and obtaining a measurement updating equation, namely star group distributed state estimation; according to the invention, cluster dynamic ad hoc network relative navigation is realized on the premise of limited satellite-borne measurement and control resources and complex space optical environment, the global positioning navigation problem under limited information interaction is solved, and the method has the capability of fast and accurate area positioning of an in-and-out satellite group system and global state sensing under limited resources.

The invention relates to the technical field of spacecraft navigation, and discloses a method for estimating the relative attitude of a failed spacecraft under the constraint of computing resources, and the method comprises the steps: carrying out the preprocessing of measurement data, obtaining the observation data of the relative attitude of a camera based on an ellipse recognition algorithm, and obtaining the measurement data of the relative attitude of a laser radar based on an iterative nearest point algorithm. Pseudo measurement factors are designed for the camera and the laser radar, and variational integral prediction factors are designed for attitude increments at adjacent moments of the failed spacecraft; factor graph optimization based on a moving window is carried out, and the relative attitude of the failed spacecraft is estimated; the momentum conservation characteristics of the failed spacecraft can be mined through a designed variational integral predictive factor, the constraint relation of attitude increments at adjacent moments is constructed, and the problem that the estimation precision of an existing scheme depends on high-frequency measurement input is solved; and low-frequency measurement data of the camera and the laser radar are fused through a built failure spacecraft full-state factor graph model, so that the relative pose estimation of the non-cooperative target under the low computing power load is realized.

The invention discloses a method for carrying out correlation and orbit determination on a space target based on a double-star optical observation result, and the method comprises the steps: obtaining the observation vectors of two observation stars for a certain space target in the respective visual fields, calculating the shortest distance between the observation vectors, and if the distance is always smaller than a threshold value in a time sequence, determining that the space target does not exist. And if not, determining that the two observation stars observe the same space target, calculating a midpoint of two observation vector common vertical lines of the same space target by the two observation stars as an orbit determination point, and performing orbit determination on the space target based on an orbit determination point sequence. According to the method, a low-cost space-based optical orbit determination mode is adopted, space target association is completed through an algorithm with low calculation cost, target consistency is judged based on observation vector distance constraints, a complex model or a large amount of calculation power is not needed, efficient processing can still be performed when the number of targets is large and the data size is large, and the cost and efficiency advantages are remarkable.

An almanac method for autonomous neural network navigation of a spacecraft includes performing steps via a computer onboard the spacecraft for determining a current navigation state of the spacecraft and determining a current target epoch for the spacecraft based on the current navigation state. The method also includes reading neural network model parameters from an almanac for the current target epoch. The almanac contains a plurality of neural network models, with each model being valid for an epoch corresponding to a portion of a trajectory. The method further includes executing the neural network model corresponding to the current target epoch to provide one or more thrust commands for the spacecraft.

The invention belongs to the technical field of data processing, and discloses a Mars rover non-contact trafficability estimation method and system based on thermal inertia, and the method comprises the steps: obtaining a ground surface temperature time sequence and environment parameters of a target region; constructing a thermal inertia estimation model, and inputting the earth surface temperature time sequence and the environmental parameters into the thermal inertia estimation model to obtain a thermal inertia estimation value of the fire soil; constructing a thermal inertia and trafficability index relation model, inputting the thermal inertia estimation value of the fire soil into the thermal inertia and trafficability index relation model, and outputting a predicted trafficability index; and determining a trafficability state grade according to the predicted trafficability index, and generating a navigation control suggestion, so that the Mars rover passes through the target area according to the Mars rover navigation control suggestion. According to the method, the trafficability index of the Mars rover is predicted by extracting the thermal inertia estimation value of the fire soil in the target area, the evaluation precision of the Mars rover passing through the target area can be effectively improved, and the driving safety of the Mars rover is ensured.

The invention relates to an ephemeris orbit calculation method suitable for orbit extrapolation on an elliptical orbit satellite. The ephemeris orbit calculation method comprises the following steps: calculating ephemeris orbit base point parameters on the ground; ephemeris orbit base point parameters which are injected from the ground are provided; and performing orbit extrapolation by using the ephemeris orbit base point parameters to calculate the position and the speed of the satellite in the inertial coordinate system. According to the method, orbit extrapolation is carried out on the basis of ephemeris orbit base point parameters injected on the ground, the calculation amount is low, the method is suitable for satellite-borne calculation, the calculation precision is high, the high-precision requirement of real-time orbit forecasting in the elliptic orbit satellite inter-satellite communication period can be met, the calculation capacity constraint of a satellite service system can be met, and the completion degree of an inter-satellite communication task is improved.

The invention provides a space debris angle-measurement-only short-arc rapid initial orbit determination method and a space debris angle-measurement-only short-arc rapid initial orbit determination system. The method comprises the following steps: S1, acquiring space debris angle-only segmental arc observation data; and S2, performing polynomial fitting on the observation data. And S3, calculating the pointing vector of the space debris relative to the observation station in the J2000 geocentric celestial coordinate system at the midpoint moment of the arc section. And S4, calculating the change rate of the pointing vector at the midpoint moment of the arc section. And S5, determining the distance at the midpoint moment of the arc section and the initial value of the distance change rate. And S6, calculating an objective function corresponding to the initial value. Step S7, if the target function does not reach the convergence condition, calculating a target function gradient corresponding to the current distance and the distance variation rate, updating the current distance and the distance variation rate, and executing steps S6 to S7; and if the target function reaches the convergence condition, calculating the corresponding position and speed. The implementation process is simple and easy to implement, and the method can be effectively applied to the space debris angle-only short arc initial orbit.

The invention provides a failure spacecraft relative pose estimation method based on dual algebra, and belongs to the field of spacecraft navigation, and the method comprises the following steps: building a failure spacecraft six-degree-of-freedom relative motion and dynamics model under dual algebra description; then, through a matrix vectorization operator, determining analysis forms of a state equation and an observation equation, and establishing a multiplicative extended Kalman filter; finally, unscented transformation is introduced, a process noise covariance and measurement noise covariance adaptive law is designed, the problems that parameter adjustment of a complex system is tedious and convergence is slow are solved, and accurate estimation of the near-operation real-time pose of the failed spacecraft is achieved. According to the invention, successful execution of the capture maintenance task can be ensured.

The invention discloses a moon leap trajectory planning method, system and device and a medium. The method comprises the steps of obtaining target landing point data; based on the acquired data, establishing a moon surface leap device kinetic model by defining a specific coordinate system and a state variable; based on the kinetic model, discretizing the kinetic equation through a numerical integration algorithm, constructing and solving a leap trajectory convex optimization problem with optimal burnup, and obtaining an obstacle avoidance-free optimal trajectory; based on the obstacle avoidance-free optimal trajectory, performing obstacle avoidance iteration through an iteration obstacle increase strategy to obtain a static obstacle avoidance trajectory; and on the basis of the static obstacle avoidance trajectory, the flight trajectory is adjusted again by calculating the trajectory error band width and the maximum extension distance, and a dynamic safety enhanced trajectory is obtained. According to the method, safe, robust and efficient moon surface leap trajectory planning in a complex and uncertain environment is realized.

The present invention discloses a robust optimization method for autonomous optical navigation landmarks for planetary landing, which belongs to the field of deep space exploration technology. The implementation method of the present invention is as follows: using the height information of the lander to determine the scale of the image plane projected onto the planetary surface, and deriving the quadrilateral vertex expression of the field of view based on the geometric relationship; using the relationship between the field of view vertex deviation and the lander posture error, converting the posture uncertainty to the field of view vertex uncertainty, and using the error ellipse to describe the field of view vertex uncertainty distribution; screening the quadrilateral area formed by the intersection of common tangents, which is the field of view intersection under the lander posture uncertainty; using the nominal trajectory and camera parameters to estimate the field of view sequence, combining the field of view intersection acquisition method to obtain the field of view intersection sequence, the field of view intersection sequence is obtained by offline estimation, and the amount of onboard calculation of the lander is reduced; using the observability of the navigation system to establish the landmark optimization index, optimize the selection of the landmarks in the field of view intersection sequence, and reduce the influence of posture uncertainty on the selected landmarks.