170results about "Instruments for comonautical navigation" patented technology

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 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 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 discloses a robust adaptive fusion filtering astronomical attitude determination method and system, and belongs to the technical field of radio navigation. The method comprises the following steps: constructing an integrated navigation model of a starlight and inertial integrated navigation system based on Lie group description; obtaining an observation residual error and an observation residual error covariance of a corresponding mode; robust and adaptive mode parallel sub-filtering is carried out, and posterior states and covariances in the two modes are updated; updating the mode probability; and fusing the updated posterior state with the covariance to obtain fusion output of parallel sub-filtering, taking the fusion output of the parallel sub-filtering as final estimation of the starlight and inertial integrated navigation system based on Lie group description, and outputting an astronomical attitude determination result containing an azimuth angle, a pitch angle and a roll angle. The high-precision astronomical attitude determination method has strong adaptability, strong robustness and accuracy when coping with complex noise and measuring outliers, and can realize high-precision astronomical attitude determination of starlight and inertial integrated navigation in a complex dynamic environment.

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 satellite orbit single-point state forecasting method based on deep learning. An orbit precision ephemeris is used as training data, a six-dimensional position speed state at a current moment is used as input, a state at a next moment is used as a target value, a prediction network in which LSTM and a one-dimensional CNN are connected in series is constructed, and a CBAM attention module is introduced to realize adaptive weighting of channel and time sequence features. And designing a weighted loss function, carrying out weighted fusion on a mean square error of the model to a real ephemeris, an HPOP equal-step prediction result and a residual mean square error output by the model, and improving prediction rationality through physical constraints. And training and evaluating through the training set, the verification set and the test set, and outputting a predicted ephemeris. According to the method, a historical sequence or TLE is not needed, high-precision and rapid forecasting can be achieved only through a single-point state, and unmodeled perturbative force is implicitly compensated.

The invention provides a spacecraft attitude orbit navigation initial value on-orbit establishment method and device, and belongs to the technical field of spacecraft navigation.The method comprises the steps that a navigation system single-machine product configuration scheme and a single-machine installation layout scheme are designed; judging whether the platform is successfully separated from the mother platform or not; the self angular rate is damped; performing slow rotation search around two axes of the body, and preliminarily establishing an inertial system attitude initial value; carrying out slow rotation search around two axes of the body, and preliminarily establishing an initial speed value of the earth fixed connection position; performing attitude maneuver based on the inertial system attitude initial value and the earth fixed connection system position speed initial value, and adjusting the attitude to a final stable direction capable of capturing the attitude and the position speed at the same time; and according to the attitude and position speed result, calculating the position speed and attitude result under the expected reference system as a final inertial navigation initial value. According to the invention, a software and hardware interface for information transmission does not need to be arranged between the spacecraft and the release mother platform, and the capability of on-orbit self-establishment of inertial navigation attitude and position speed initial values is realized.

The invention discloses a constellation communication-navigation coupling task scheduling method based on load balancing, and belongs to the technical field of aerospace mission planning. The method mainly improves the resource distribution efficiency under the device-level constraint, and achieves the purpose of meeting the geometric precision requirement of navigation service while guaranteeing the completion of a communication task. According to the method, satellite ephemeris data and task demand parameters are used as input, a multi-target mixed integer linear programming MILP model is constructed, an optimal binary allocation variable is solved through an MILP solver, service satellites and service time slots of tasks are extracted from the optimal binary allocation variable, and then a scheduling table is generated; according to the method, the problem of navigation task failure caused by the fact that a communication task occupies a multi-satellite convergence window in separated planning is avoided, simulation shows that the overall task completion rate is averagely increased by 12.3%, and the load variation coefficient is reduced by 66%-88%.

The invention provides a kinematics equation linearization method of normalized time, which converts a time variable under an absolute scale into a relative scale, thereby facilitating derivation and numerical calculation of an optimal control problem. Converting free time into time with boundary conditions by using normalized time, and converting time serving as a differential term in the motion equation into a multiplication term, so as to realize transformation of the motion equation; and according to the motion equation after time normalization, the time after normalization is optimized, and equivalent solution of an optimal control problem is realized. Equivalent solution of an original optimal control problem is realized, and derivation and numerical calculation of the optimal control problem are facilitated.

The invention relates to the technical field of spacecraft autonomous orbit determination, in particular to an HEO constellation enhanced earth-moon spacecraft satellite-borne GNSS autonomous orbit determination method and system, and the method comprises the steps: determining the weight of GNSS / HEO data; calculating a single-point positioning and speed measurement result by using GNSS / HEO data and HEO enhancement information, and initializing a state quantity and a covariance of a Kalman filter; forecasting the position and speed of the spacecraft by adopting a simplified kinetic model, and updating a state equation; calculating a GNSS / HEO innovation residual error sequence based on a forecast result, and detecting and marking gross errors according to a clock error consistency rule; sequentially carrying out Kalman measurement updating on the observation data without gross error according to the weight, and finishing the correction of the state quantity and the covariance; finally, the position and speed information of the spacecraft is output to a navigation guidance system, and autonomous orbit determination and time service are achieved. Therefore, the problems of low positioning precision, poor orbit determination and autonomous time maintenance capability and the like due to the fact that the earth-moon spacecraft cannot adapt to earth-moon space dynamic and observation conditions in the related technology are solved.

This invention discloses a method for improving the orbital accuracy of a lunar probe based on reflective prism constraints. The method first utilizes laser ranging data obtained from a known reflective prism on the lunar surface using an onboard laser altimeter, and constructs an error equation set by combining this data with a dynamic orbital model. Next, a high-resolution camera is used to perform multiple imaging operations on the lunar reflective prism from different orbits and perspectives. Using the bundle adjustment principle, the observed image point coordinates are combined with the dynamic model to construct the error equation set. Finally, laser ranging observations from the reflective prism and multi-temporal photogrammetry observations are used as two different types of geometric constraints to construct a joint adjustment model of multi-source observation data. The high-precision orbital parameters are then calculated using a least-squares algorithm. This invention uses laser ranging data and an optical image database to add new constraints to the estimated orbit, ultimately correcting and further improving the orbital accuracy of the lunar probe.

The present disclosure relates to an information processing device and an information processing method that allow provision of image capturing service in which images are captured using a camera mounted in an artificial satellite. The information processing device includes a display control unit for displaying a first screen including map display portion displaying, on a map of an astronomical body, artificial satellites circulating the astronomical body and the orbits of the artificial satellites, and a time information display portion displaying time information on the orbits of the artificial satellites. The display control unit displays, on the first screen, the positions of the artificial satellites on the map of the astronomical body provided by the map display portion and time information of the artificial satellites provided by the time information display portion such that the positions are synchronized with the time information. The technology of the present disclosure can be applied to, for example, an information processing device and the like that provide image capturing service in which images are captured by a camera mounted in an artificial satellite.

The application provides a spacecraft autonomous navigation method and electronic equipment. The spacecraft autonomous navigation method comprises the following steps: determining technical indexes of an optical sensor according to navigation requirements, and determining a configuration of the optical sensor according to orbit information of the spacecraft; collecting a star map by using the optical sensor with the determined technical indexes and configuration, processing the star map, identifying a refraction star and extracting a refraction angle; and using a Kalman filtering algorithm to perform information fusion on information output by an inertial navigation system and a refraction apparent height calculated based on the refraction angle, so as to correct navigation errors and obtain corrected navigation information. The application corrects drift errors of the inertial navigation by astronomical navigation, maintains continuous navigation information by using the inertial navigation when the astronomical system is unavailable, and achieves the purpose of improving navigation accuracy.

This disclosure relates to the field of relative navigation technology, and proposes a method, device, electronic device, and readable storage medium for locating a space target. The method includes: capturing an image of the space target using a space camera with imaging capabilities to obtain a spatial image and reflected light signal; determining the angular information of the space target relative to the orbital coordinate system of the probe platform based on the spatial image; determining the Fraunhofer line information of the reflected light from the space target based on the reflected light signal, and calculating the line-of-sight velocity of the space target based on the Fraunhofer line information and the solar Fraunhofer standard line; and calculating the target coordinates of the space target in the orbital coordinate system of the probe platform using a Kalman filter algorithm based on the angular information and the line-of-sight velocity. The technical solution provided by one or more embodiments of this disclosure is a passive detection method with strong concealment and anti-interference capabilities, enabling high-precision, low-power relative navigation of space targets.

The application discloses a kind of star group between heterogeneous measurement data association and high-precision fusion orbit determination method, belong to space relative navigation technical field.The method includes grouping to the member star in star group, constructs the measurement set of the angle measurement and ranging of all member stars in time t, then calculates the position assumption set of member star, calculates the distance assumption value between member stars under each association assumption, constructs association degree evaluation index by distance assumption value and actual distance measurement, constructs cost matrix according to association degree evaluation index, and then obtains optimal association result, obtains the ranging and angle measurement data corresponding to all member stars at current time relative orbit determination, including using optimal estimation method to carry out state estimation, and including determining system state, constructing system measurement equation and system equation;Finally, the above steps are repeatedly executed for each measurement time, until filter convergence, obtain the high-precision orbit determination result of member star.

The application belongs to the technical field of data processing, and discloses a Mars rover non-contact passability prediction method and system based on thermal inertia, which comprises the following steps: obtaining the ground temperature time sequence and environmental parameters of a target area; constructing a thermal inertia estimation model, inputting the ground temperature time sequence and environmental parameters into the thermal inertia estimation model, and obtaining the thermal inertia estimation value of the soil; constructing a thermal inertia and passability index relationship model, inputting the thermal inertia estimation value of the soil into the thermal inertia and passability index relationship model, and outputting a predicted passability index; determining the passability state level according to the predicted passability index, and generating a navigation control suggestion, so that the Mars rover passes the target area according to the navigation control suggestion of the Mars rover. The application can effectively improve the evaluation accuracy of the Mars rover passing the target area by extracting the thermal inertia estimation value of the soil of the target area to predict the passability index of the Mars rover, and ensures the driving safety of the Mars rover.

The invention provides a multi-star sensor attitude fusion calculation method and system, and the method comprises the steps: S1, calculating the projection of an optical axis of a star sensor j in a spacecraft body system according to an installation quaternion from a measurement coordinate system of the star sensor j to the spacecraft body system; s2, according to the quaternion from the inertial system of the star sensor j to the spacecraft body system, calculating the projection of the optical axis of the star sensor j under the inertial system; and S3, according to the vector sum, in combination with a Moore-Penrose generalized inverse theory of the matrix, calculating to obtain a spacecraft multi-star sensor fusion attitude conversion matrix. The invention aims to fully utilize the attitude information of all the star sensors carried by the spacecraft to carry out multi-star sensor attitude fusion calculation when the quantity of the star sensors carried by the spacecraft is more than 2, so as to realize high-precision spacecraft attitude measurement.

The invention discloses a cabin relative pose measurement method based on non-cooperative assembly characteristics, and the method employs a displacement sensor and a visual camera to form a double-sided measurement device, synchronously measures the assembly characteristics of a cabin, calculates and decouples the relative poses of six degrees of freedom, and provides a basis for pose adjustment assembly. According to the method, heterogeneous measurement data fusion and coordinate system unification of the displacement sensor and the visual camera are achieved, mutual decoupling and synchronous posture adjustment of six degrees of freedom can be achieved, the posture adjustment convergence rate is increased, and the assembly efficiency can be effectively improved.

The present application relates to a kind of small celestial body landing process using ToF camera to realize the method for aircraft positioning, belong to terrain relative navigation field.The method described in the present application can use the characteristics of ToF camera self-illumination and be aimed at the real-time rendering of the three-dimensional topography of terrain.After the generation of rendering image, the real image can be matched with rendering image using mask NCC method, and then the corresponding relationship of the center of field of view pointing to map coordinate is obtained.In addition, the depth information of ToF camera is also used to obtain the dimension of the above-mentioned pointing, and then the current aircraft position is calculated.The sensor type and light source type used in the present application are different from NFT algorithm, and only single matching is needed to complete the positioning function of aircraft.

The invention provides a state estimation and error identification method and device for an autonomous navigation system, and the method comprises the following steps: obtaining an observability evaluation index of the system by using an observability criterion; according to the observability evaluation index of the system, obtaining the accurate compensation degree of the measurement error of the optical sensor, and determining the achievable state estimation accuracy of the autonomous navigation system; constructing an observability optimization function; analyzing a coupling relationship between a measurement error and a system state based on an observability optimization function, and establishing an error response model; according to the error response model, constructing a system state parameter joint optimization model through the observability evaluation index of the navigation system; and optimizing the system state parameter joint optimization model and realizing on-orbit identification of errors and system states. The accuracy of observability judgment of the navigation system is improved, observation information and calculation resource allocation can be autonomously optimized, and the problem of state estimation under the weak observation condition is solved.

The invention relates to a lunar satellite formation relative navigation method based on factor graph optimization, which comprises the following steps: in a sliding window with a set length, constructing a factor graph by using relative orbit states of a series of satellite formations as state nodes and constraint factors connected with the state nodes; and optimizing the factor graph to obtain the optimal estimation of the relative orbit state of the satellite formation at the current moment. According to the relative navigation method, historical data are utilized, the historical orbit state is optimized and corrected again in the process of optimizing the factor graph, error accumulation is effectively restrained, and navigation precision and robustness are remarkably improved.

The invention provides an observability analysis method and device for deep space exploration optical autonomous navigation, and the method comprises the following steps: constructing an orbit dynamics model and an optical observation model according to the measurement information obtained by a deep space probe; constructing an optical autonomous navigation system model according to the orbit dynamics model and the optical observation model; observability parameters of the deep space exploration optical autonomous navigation system are given according to the optical autonomous navigation system model; constructing an observability characterization model of the autonomous navigation system; carrying out dimensionality reduction characterization on observability of the deep space exploration optical autonomous navigation system; therefore, an observability criterion is obtained; and realizing the observability autonomous determination of the deep space exploration optical autonomous navigation system according to the observability criterion. By analyzing key factors of complete representation of the observability of the system, judgment conditions of the observability of the system are deduced, and the observability quantitative expression method which is universal, simple and easy to autonomously realize on a satellite is provided.

The embodiment of the invention provides a space target maneuvering detection method and system based on an information energy momentum tensor, and relates to the technical field of spaceflight detection. Comprising the steps of obtaining orbital elements of an observation satellite and a space target; according to the number of the orbits, the number of the wide-area invariant opposite orbits is obtained; obtaining a space target relative position after dimension normalization under an observation satellite LVLH coordinate system according to the wide-area invariant relative orbital element; according to the relative position and the number of the relative orbits, the relative motion acting amount and the relative speed are obtained; carrying out filtering processing on the relative motion state of the observation satellite according to a mean filtering algorithm; acquiring a corresponding energy momentum tensor according to the filtered relative motion action quantity, relative position and relative speed; and obtaining a time-varying curve of the energy momentum tensor, and carrying out maneuvering detection on the space target. According to the method, the detection capability on low thrust and weak signal maneuver can be remarkably enhanced, and the limitation that a traditional method is high in dependence on model precision and insensitive to micro maneuver is overcome.

The invention relates to the technical field of space target correlation and positioning, and discloses a space target correlation matching and coordinate determination method, device and equipment and a medium, and the space target correlation matching and coordinate determination method comprises the steps: predicting the tracks of a plurality of observation stars and a plurality of space targets, and determining angle measurement observation data according to the target tracks; performing coordinate transformation on the angle measurement observation data; determining a projection distance cost matrix according to the unit relative position vector obtained by coordinate conversion; according to the projection distance cost matrix, matching space targets observed by the plurality of observation stars by adopting a Hungary allocation algorithm, and determining target position coordinates according to the position coordinates of the observation stars and the unit relative position vector in a matching result; according to the method, the unit vector is obtained by predicting the trajectory of the observation star and the space target and converting the coordinates, the cost matrix-based Hungary algorithm is constructed to match the target, and the space target coordinates of the position are fitted, so that the efficiency and accuracy of determining the space target coordinates are improved.

The invention discloses a grid star map identification method and system based on weighted voting and geometric verification, and the method specifically comprises the steps: firstly, constructing a feature mode database containing all-day navigation stars, and storing the rasterized geometric mode of an adjacent star sky region of each navigation star and the inverse document frequency (IDF) weight of each grid unit in the database; secondly, extracting a plurality of brightest observation stars as main stars from an observation star map obtained by a star map sensor, and generating an observation geometric mode for each main star; inquiring an offline database by utilizing an observation mode, generating a candidate navigation star list, dynamically filtering low-credibility candidate stars through an adaptive threshold strategy, and sorting the candidate stars by adopting a mixed scoring function; and finally, carrying out pose consistency check on the plurality of candidate stars with the highest scores through a geometric verification method, and identifying the most reliable matching result. On the premise of ensuring the real-time performance, the grid star map recognition robustness and accuracy under various noises and interferences are improved.

The present application relates to the field of space in-orbit close-range relative visual navigation technology, and particularly relates to a space target visual measurement method combined with acceleration drift estimation, a tracking spacecraft is provided with an acceleration sensor and a camera, S1: obtaining initial relative position information and relative velocity information through a measurement system, S2: obtaining acceleration values and space target images in real time; S3: establishing a relative motion equation, substituting the obtained acceleration values and the relative position and relative velocity information at the previous moment into the relative motion equation to calculate position information; S4: calculating an error amount, and estimating the errors of the relative position information, the relative velocity information and the acceleration values in the dynamic equation through a filtering algorithm according to the error amount to correct the calculation results obtained in the dynamic equation, so that accurate positioning of a target spacecraft can be realized.

The application provides a spacecraft attitude orbit navigation initial value on-orbit establishment method and device, and belongs to the technical field of spacecraft navigation, which comprises the following steps: designing a navigation system single machine product configuration scheme and a single machine installation layout scheme; judging whether the spacecraft has been successfully separated from a mother platform; damping the angular velocity of the spacecraft; performing slow rotation search around two axes of the spacecraft body to preliminarily establish an inertial system attitude initial value; performing slow rotation search around two axes of the spacecraft body to preliminarily establish an earth-fixed system position and velocity initial value; performing attitude maneuvering based on the inertial system attitude initial value and the earth-fixed system position and velocity initial value, and adjusting the attitude to a final stable pointing direction which can simultaneously capture the attitude and the position and velocity; and calculating the position and velocity and the attitude in an expected reference system according to the attitude and the position and velocity, and taking the position and velocity and the attitude as the final inertial navigation initial value. The application can make the spacecraft and the released mother platform not need to be equipped with information transmission hardware and software interfaces, and can realize the capability of establishing the inertial navigation attitude and position and velocity initial value on the spacecraft.

The invention discloses a navigation mode transition method, an electronic device and a medium, in a working process of a first navigation mode, geometric angles among a mother satellite, a service spacecraft and space debris are monitored, and when the geometric angles meet a mode switching requirement, the position of a target orbit under a J2000 coordinate system is obtained through double-satellite geometric estimation; and taking the position as an initial position calculated in the second navigation mode, and switching to the second navigation mode. Before the navigation mode is switched, the target orbit position is determined through double-satellite geometric estimation, the single-satellite navigation positioning precision can be rapidly improved, a good initial value is provided for double-satellite angle measurement filtering, and the convergence process is accelerated.

The application provides a method for autonomous rendezvous with small celestial bodies based on integrated design of line-of-sight measurement and trajectory maneuver, comprising: obtaining planned nominal maneuver trajectory and orbit transfer information on the ground; keeping the direction of the probe to the small celestial bodies after each orbit transfer, and measuring the central line-of-sight direction of the target small celestial bodies; according to the line-of-sight measurement results and the measurement results of the orbit transfer velocity increment, solving the initial relative position and initial velocity of the probe to the small celestial bodies, and the current relative position and velocity; according to the relative position at the orbit transfer time, the state transition matrix and the flight time, determining the required velocity increment to reach the next nominal orbit transfer position; controlling the probe to autonomously adjust the posture to the orbit control velocity increment direction, and after reaching the nominal orbit control time, controlling the probe to autonomously start the orbit control process; after completing the orbit transfer, autonomously judging whether the probe reaches the vicinity of the terminal position of the rendezvous section.

The application discloses a kind of GNSS optical angle-measuring autonomous orbit determination observation scheduling and consistency adaptive evaluation method for high-orbit spacecraft, belongs to spacecraft autonomous navigation technical field.The application includes: system initialization;GNSS target visibility pre-screening;Resource constraint modeling considering the dynamic performance of turntable;Optimal observation sequence generation under time budget constraint;Target imaging based on observation sequence and angle measurement extraction;Extended Kalman filtering with outlier suppression;Consistency evaluation based on innovation statistics and process noise closed-loop adaptive setting;Orbit determination result output and iteration.The application explicitly includes scheduling into the turntable maneuvering ability by resource constraint modeling, solves the observation efficiency problem under the limitation of spaceborne resource;Through optimal observation sequence generation, information acquisition maximization in limited time is realized;Through consistency closed-loop setting, the filter can be adapted to dynamic model error, and the statistical consistency of long arc segment operation is ensured.The application provides a kind of high-reliability, high-precision autonomous orbit determination technical approach for high-orbit spacecraft under GNSS signal blind area and weak ground support condition.