115 results about "Orbit determination" patented technology

The invention discloses an unknown target short-arc high-precision initial orbit determination method based on filtering enhancement, and the method comprises the steps: firstly, carrying out the adaption of an initial orbit determination method according to an observation data type, after obtaining a plurality of groups of initial orbit candidate solutions, determining an initial orbit optimal solution according to a measurement residual error minimum principle, and converting the initial orbit optimal solution into a position vector and a velocity vector; configuring a dynamical model and parameters by combining a target track type and physical characteristics; establishing a target state equation based on the model, and introducing a process noise compensation model error; and constructing an observation equation by combining observation platform position information. In order to guarantee effective starting of the filter, parameters such as the orbit state, the state error covariance, the process noise and the measurement noise need to be calculated and initialized in advance. And finally, in combination with the state equation and the observation equation, the target orbit state is predicted and updated through an EKF technology until all observation epochs are traversed, and high-precision target initial orbit information is obtained.

The invention relates to the cross technical field of satellite navigation, satellite dynamics and the like, in particular to a dynamic orbit determination method based on integer ambiguity fixation, which can obviously improve orbit determination and extrapolation precision, and comprises the following steps: establishing a dynamic orbit determination model, establishing a GPS observation model, and then performing integer ambiguity fixation IAR, a single-difference IAR method adopting single receiver ambiguity correction specifically comprises the steps that firstly, GPS carrier phase observation is divided into continuous tracking segments for processing, for each tracking segment, an average Hatch-Melbourne-Wubbena (HMW) combination is constructed by means of double-frequency pseudo-range and carrier phase data, then cross-segment inconsistency is corrected, integers of wide lane ambiguity are achieved, and a single-receiver ambiguity correction result is obtained. And the corrected wide lane ambiguity is combined with an IF ambiguity floating point estimation value obtained by a least square method for use, and finally, the IF ambiguity difference is constrained as an integer value to realize the fixation of the single-difference ambiguity, and compared with the prior art, the orbit determination precision can be remarkably improved.

The invention provides a space target cataloguing library maintenance method and device fusing post-event multi-source observation data, and belongs to the field of space situation awareness. The method comprises the following steps: acquiring post observation data of an observed space target and preprocessing the post observation data; calculating an observation arc section by using the preprocessed observation data, taking an obtained initial orbit determination result as an orbit feature of the observation arc section, further matching the observation arc section with orbit data of a target in a cataloguing library, and taking the successfully matched observation arc section as an associated arc section of the target; and performing precise orbit determination by fusing a plurality of associated arc segments of the target, and updating and maintaining the cataloguing library by using a precise orbit determination result. According to the method, data preprocessing can be rapidly carried out according to the input multi-source observation data, association matching of the data and the space target is completed, finally, precise orbit determination is carried out by fusing the multi-source observation data, the calculation efficiency and the process completeness are guaranteed, and maintenance of a space target cataloguing library is achieved.

The invention discloses a simulation system and a simulation method for a deep space tracking measurement task, and relates to the technical field of deep space exploration tracking measurement. The system comprises a task simulation and visualization subsystem which is used for carrying out environment parameter configuration and task state control on a deep and far space simulation task and outputting operation information; the data set construction and verification subsystem is used for generating a navigation verification data set and performing functional performance simulation verification; and the algorithm performance evaluation subsystem is used for constructing a performance evaluation index system and evaluating the tracking measurement algorithm. According to the invention, integration of tracking measurement technologies such as space reference, one-way measurement, autonomous orbit determination, landing and patrol navigation and the like of deep and far space tasks and planet calendar, detection tasks and aircraft state simulation can be realized, test verification based on in-orbit actual measurement data is carried out, and a unified and standard evaluation index system is provided; the problem that an existing deep space task simulation system cannot provide accurate reference for a deep space ultra-large spatial-temporal scale simulation system is solved.

At a satellite, in each of a plurality of time periods, a method of communicating with first and second ground stations, to exchange a first downlink message and a first uplink message with the first ground station and a second downlink message and a second uplink message with the second ground station, wherein the first and second downlink messages and the first and second uplink messages are time-stamped messages, and determining one or more orbit parameters indicative of an orbit of the satellite based on known positions of respective first and second ground stations relating to one or more of the plurality of time periods and based on first timing information relating to times of reception and transmission of respective first uplink and downlink messages and second timing information relating to times of reception and transmission of respective second uplink and downlink messages in the one or more time periods.

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

An antenna and receiver design is described herein that allows a spacecraft to safely and accurately perform an autonomous flight safety system, an ascent, an entry, a re-entry, a landing, and / or an orbit determination. For example, a spacecraft can be equipped with 180-degree dual-opposed antennas and two receivers. A signal captured by a first antenna can be routed to a secondary radio frequency (RF) of the first receiver and to a combiner. A signal captured by a second antenna can be routed to a secondary RF of the second receiver and to the combiner. The mixer can perform the phase combination operation on the signals captured by the first receiver and the second receiver to produce a phase combined signal, and can route the phase combined signal to the primary RF of the first receiver and to the primary RF of the second receiver.

The application discloses a kind of only angle measurement lunar space Lagrange point orbit type intelligent identification method, belong to space technology field.The application establishes only angle measurement observation model, sets only angle measurement observation error covariance matrix;Establish the motion model of space Lagrange point orbit spacecraft in lunar space;According to the number of space Lagrange point orbit in lunar space, randomly generate space Lagrange point orbit in lunar space and recursion length, generate only angle measurement observation vector;The sample point required for training orbit type identification deep neural network is generated;Based on deep neural network, the fast and efficient identification of space Lagrange point orbit in lunar space is realized, and then the initial orbit determination and refined orbit estimation of space Lagrange point orbit in lunar space are realized.The application has the advantages of not depending on physical model, high identification accuracy, small amount of calculation, high efficiency and good robustness.The application is beneficial to improve the space situation awareness capability in lunar space, improve the determination accuracy and calculation speed of space Lagrange point orbit in lunar space.

This invention discloses a method for inverting pulse maneuvers in non-cooperative spacecraft based on minimizing orbit determination residuals. Addressing the difficulty in accurately inverting maneuver parameters for non-cooperative spacecraft under conditions of sparse observation data and unknown maneuver times, this invention constructs a sequence of historical orbits and the latest observed orbit states. Within a preset time window, candidate maneuver times are discretized. For each candidate time, orbit propagation is performed under the assumption of no maneuvering to obtain a predicted orbit. A residual vector is constructed based on the position difference between the observed and predicted orbits. A partial derivative matrix is ​​constructed through numerical differentiation. The least squares method is used to solve for the maneuver velocity increment. Candidate intervals and candidate times are selected based on the fitting error. Finally, the maneuver parameters are determined based on the criterion of optimal accuracy within the fit interval. This invention can achieve high-precision estimation of maneuver times and velocity increments under sparse observation conditions, reducing dependence on high-density observation data and prior maneuver information, and significantly improving the stability and reliability of maneuver inversion.

This invention discloses a method and system for monitoring and resetting the software heartbeat of a spaceborne receiver. The system includes a dual-channel GNSS positioning module, a dual-channel orbit determination module, and an antifuse FPGA chip. Both the GNSS positioning module and the orbit determination module have corresponding primary and backup FLASH memory. The antifuse FPGA acts as the heartbeat monitoring and control center, connecting to the GNSS positioning module and the orbit determination module via GPIO interfaces. It receives the software heartbeat signals from each module and sends reset commands to faulty modules via reset pins. This solution solves the problems of poor radiation resistance, low fault identification accuracy, and low recovery efficiency in existing spaceborne receiver software fault monitoring.

The application provides a GNSS / inter-satellite link distributed cooperative real-time orbit determination method and device for a mega constellation, first, each satellite in a cluster performs preliminary orbit determination based on local GNSS observation data and IMU data to obtain preliminary determination results of each satellite in the cluster; then, a cluster head satellite collects the preliminary determination results to determine relative orbit parameters in the cluster; then, a backbone node collects the relative orbit parameters to determine a regional absolute orbit reference; a global anchor satellite determines unified constellation orbit parameters; finally, according to the regional absolute orbit reference and the unified constellation orbit parameters, orbit determination results of each satellite are determined. Through a three-level hierarchical cooperative architecture, the application disperses data processing and determination tasks to different level nodes, avoids data processing bottlenecks of a central node and communication link congestion problems in centralized technology, reduces complex interaction between nodes and inter-satellite link resource occupation in a thousand-star level scale in traditional distributed technology, and shortens iteration convergence time.

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.

Provided are a URA calculation method and device for a low-orbit navigation satellite, a computer storage medium and a computer program product. The URA calculation method comprises: in precise orbit determination and clock error calculation for a low-orbit satellite, a covariance submatrix of a predicted orbit and clock error parameter in a predicted arc segment is given by using a transfer matrix, denoted as a first covariance submatrix; a projection matrix from a geocentric geodetic coordinate system to a satellite orbit coordinate system is calculated; the first covariance submatrix is converted into a second covariance submatrix of the satellite orbit coordinate system according to the projection matrix; the satellite orbit prediction accuracy and the clock error prediction accuracy are calculated according to the second covariance submatrix; and the URA is calculated according to the satellite orbit prediction accuracy and the clock error prediction accuracy. The present disclosure proposes a URA calculation method for a low-orbit satellite based on a covariance matrix, which guarantees the real-time performance and reliability of integrity monitoring.

The invention relates to a satellite orbit circle number forecasting method and device, computer equipment and a storage medium. Acquiring initial orbit data of the target satellite based on the first satellite identifier of the target satellite, determining corresponding initial ephemeris data according to the initial orbit data of the target satellite and a pre-configured first time parameter, and determining an initial circle number list corresponding to the initial ephemeris data, under the condition of receiving orbit measurement data for a target satellite, determining corresponding periodic ephemeris data according to the orbit measurement data, the initial orbit data and a pre-configured second time parameter, generating a periodic circle number list corresponding to the periodic ephemeris data according to the periodic ephemeris data and the initial circle number list, and sending the periodic circle number list to the target satellite; and generating a circle number forecasting list according to the initial circle number list and the periodic circle number list. And a final circle number forecasting list is generated based on initial short-term forecasting in combination with periodic forecasting of precise orbit determination, so that long-term high-precision forecasting of future circle numbers can be realized.

The invention discloses a method for enhancing the real-time orbit determination precision of a low-orbit satellite based on a satellite-ground link signal, and the method comprises the steps: obtaining pseudo-range and Doppler observation values through employing a satellite-ground link built during the visible period of the low-orbit satellite and a ground target while the low-orbit satellite receives GNSS ephemeris and observation data and constructs a GNSS observation equation, and obtaining the real-time orbit determination precision of the low-orbit satellite based on the pseudo-range and Doppler observation values; centimeter-level position, speed and clock error information provided by a ground target in real time is received; constructing a satellite-ground link observation equation by accurately calculating a ground state at a signal emission moment, and combining the satellite-ground link observation equation with a GNSS (Global Navigation Satellite System) observation equation; and an extended Kalman filtering algorithm is adopted to carry out joint solution, a residual error test mechanism is introduced, and finally high-precision orbit information is output. The invention aims to solve the defects of insufficient broadcast ephemeris precision and overlarge precision ephemeris time delay in the prior art, and can realize high-precision real-time orbit determination under the broadcast ephemeris condition only by a small number of ground targets without resolving GNSS ephemeris errors or depending on a global observation station network.

The application provides a GNSS autonomous orbit determination method for a spacecraft, which obtains observation data in real time after GNSS signal acquisition and tracking, combines an orbit dynamics model and an observation model, realizes real-time autonomous orbit determination of the spacecraft in orbit, and improves the positioning and orbit determination precision of the spacecraft. Meanwhile, the application proposes a GNSS observation data preprocessing method based on polynomial fitting and a convergence judgment condition of the autonomous orbit determination filter, so as to ensure the stability of the autonomous orbit determination filter of the spacecraft. The GNSS autonomous orbit determination method for the spacecraft realized by the application can be widely applied to high-precision navigation, positioning and autonomous orbit determination service scenes of high, medium and low orbit spacecrafts, and has a wide application prospect.

The application relates to a rotation error correction method of a distributed autonomous navigation constellation, and comprises the following steps: S1, each satellite calculates the orbit plane orientation error of the satellite in orbit and distributes the result to all visible satellites in the constellation in real time; S2, each satellite fits a local constellation rotation quantity according to the orbit plane orientation error of the satellite and the collected orbit plane orientation error; S3, each satellite determines a virtual satellite orbit observation quantity according to the local constellation rotation quantity of the satellite; S4, each satellite constructs a conditional constraint equation of virtual observation according to the virtual satellite orbit observation quantity and the second type of non-singularity orbit root number; and S5, each satellite solves the satellite orbit through filtering according to the conditional constraint equation of virtual observation, so that the overall rotation effect of the constellation is inhibited and corrected. The application can inhibit the rotation effect of the constellation, improve the distributed autonomous orbit determination precision of the satellite, does not increase the inter-satellite data interaction communication pressure, and can stably maintain the constellation configuration for a long time.

This application relates to a radar-optical hybrid dual-arc correlation orbit determination method, comprising: acquiring radar tracking arc segments and optical tracking arc segments; calculating the initial radar orbital elements based on the radar tracking arc segments using a Lambert boundary value problem; calculating the initial optical orbital elements based on the optical tracking arc segments using a distance search method to obtain candidate solutions; determining whether the radar tracking arc segments and optical tracking arc segments could originate from the same space target; if not, the process terminates; if possible, it continues; using the radar initial orbital elements to determine the estimated semi-major axis, the candidate solutions are screened, and a position observation sequence arc segment is generated; based on the position observation sequence arc segment and the radar tracking arc segment, a Lambert boundary value problem is constructed to perform dual-arc correlation orbit determination, and the solution is obtained using the Gooding method. This invention provides a radar-optical hybrid dual-arc correlation orbit determination method that constructs a Lambert boundary value problem to solve dual-arc orbit determination, achieving excellent radar-optical hybrid dual-arc correlation accuracy and dual-arc orbit determination precision.

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 discloses an initialization method of a conic surface under a projection linearization track surface, and belongs to the technical field of calculation, reckoning or counting. The method first obtains a preliminary estimation of a target track through three groups of observation data; determines a conic surface containing an initial track surface according to initial track information and geometric knowledge; estimates the missing height information in combination with the conic surface, and obtains more position information of the target in combination with the existing angle information; projects the position information of the target into a two-dimensional section, re-estimates the track parameters by using a least square algorithm, and more accurate track parameters can be obtained; then the iteration process is repeated by using the newly obtained track information until a certain initial orbit determination accuracy requirement is met. The algorithm firstly fits the height information of the target by using a conic surface coordinate system, and realizes linear fitting of the position vector by using a track surface projection method, so that the calculation amount is small and the accuracy is high.

An inter-satellite pointing target attitude high-precision real-time online prediction method, first, the ground uses the orbit determination program corrected by on-orbit data observation to generate a double-satellite orbit position and velocity value sequence in a predetermined time period, and samples according to the control period, then for each sampling time, the double-satellite orbit position and velocity is used to establish the local orbit coordinate system, and the component of the target star position vector in the local orbit coordinate system is calculated, the pitch and yaw target attitude angle values are calculated according to the target attitude angle definition, the pitch target angle sequence is linearly averaged and estimated, the average orbit period value and the long-term change rate parameter are calculated, and the modified sequence after deducting the long-term change rate is given, for the pitch target angle correction sequence and the yaw target angle sequence, data of a specified average orbit period number is taken from a specified initial time, the fitting coefficient is calculated, finally the generated fitting coefficient is uploaded to the satellite, and the satellite uses the high-precision fitting determination formula to calculate the inter-satellite pointing target attitude angle online every control period.

The present application relates to the satellite detection technical field of space target, specifically speaking, through the load characteristic, the orbit characteristic of the earth remote sensing satellite and the orbit characteristic of the high orbit target, the high orbit space target detection planning method based on the earth remote sensing satellite is proposed, using this method generates the imaging opportunity, calculates the relative angular velocity of the high orbit space target on the focal plane of the earth remote sensing satellite;And the attitude direction is calculated to carry out high orbit space target observation, both can enhance the brightness of the detection target and realize the detection of the star target, and long arc segment monitoring can be realized to meet the demand of space-based orbit determination to dense measurement, has the wide application scene.

According to the EDMA and McBSP-based inter-satellite-borne receiver DSP communication method provided by the invention, real-time data communication of an McBSP interface between DSPs is controlled by using the EDMA, and receiving and transmitting are not required to be allocated by a CPU, so that the DSP is ensured to have sufficient machine time to complete service processing functions such as navigation message processing, observed quantity calculation, positioning / orbit determination solution and inter-satellite measurement real-time processing; a multi-channel buffer serial port McBSP interface capable of being flexibly configured is adopted between the DSPs for communication, and the hardware interface design complexity is reduced on the premise that the real-time data communication rate between the DSPs is met. By means of reasonable embedded software design, full duplex communication of the McBSP is achieved only by inquiring an EDMA event completion mark in one interrupt (timer interrupt), and the software complexity is reduced while the real-time requirement of interface data communication is met; according to the method, the design of data communication hardware interfaces among the DSPs of the receiving boards in the current double-satellite measurement receiver is simplified, and the reliability and the real-time performance of data transmission among the DSPs of the multi-system and multifunctional double-satellite measurement receiver are greatly improved.

The application provides a satellite navigation signal antenna phase center calibration method and system, comprising: satellite inter-satellite link data is used for satellite precise clock difference estimation, and high-precision satellite clock difference of the inter-satellite link is obtained; the satellite inter-satellite link clock difference is fitted with a polynomial model, and the satellite inter-satellite link clock difference model coefficients, including clock difference deviation, clock speed and clock drift, are determined; the satellite inter-satellite link clock difference model is used to constrain the satellite clock difference parameters, the random walk estimation is carried out on the clock difference random term, the simplified dynamic orbit determination principle is used, the satellite orbit and the navigation signal antenna phase center deviation are jointly solved, and the on-orbit calibration of the satellite navigation signal (L band) antenna phase center deviation is realized. The method relieves the problem that the satellite antenna phase center deviation is related to the clock difference parameters, guarantees the accuracy of the satellite antenna phase center deviation correction, and supports the satellite precise orbit determination and the precise positioning service.

The invention belongs to the technical field of glass processing, relates to processing of an atomic clock microwave cavity, and relates to a tool and a method for high-precision processing of a large-size thin-wall glass cavity. The two ends of the glass crystal ingot are cut to be flat, and the parallelism of the two end faces after cutting is smaller than or equal to 0.05 mm; a first blank of a barrel-shaped structure is manufactured in the flatly-cut end face through drilling; milling and grinding the first blank to obtain a second blank; and the second blank is polished through the tool, and then the workpiece is obtained. The tool comprises a polishing main shaft, a top roller and a driving mechanism. According to the method provided by the invention, high-precision, high-efficiency and high-reliability processing can be carried out on the large-size thin-wall glass microwave cavity, so that the processing quality of the cavity and the performance of an atomic clock are improved, and the method plays an important role in the aspects of a satellite navigation system, deep space exploration and spacecraft orbit determination, information network synchronization, a national time frequency system and the like.

The application provides a phase difference data processing method and system based on dual-polarized GNSS occultation, comprising: obtaining a low-orbit satellite orbit through precise orbit determination of LEO observation data and positioning auxiliary data; calculating atmospheric additional phase data of H channel and V channel based on the low-orbit satellite orbit and time-registered phase data, and performing geometric co-registration to determine the space-time position of the occultation event; constructing polarized phase difference according to the geometric co-registered phase data to obtain phase difference data, and repairing phase difference cycle slip according to the phase difference data; and performing data processing on the cycle slip repaired differential phase sequence to output phase difference products. The application can suppress clock and geometric residual, eliminate phase breaks caused by half / integer cycle slip and open / closed loop conversion, and correct phase difference, thereby outputting stable and consistent differential phase observable.

The application discloses a satellite constellation GNSS receiver on-orbit fault diagnosis system, which is used for diagnosing a satellite constellation composed of three stars A, B and L, and is used for space-based orbit determination and autonomous navigation of a lunar spacecraft, carries multiple test loads, manages the loads on the stars, simultaneously performs networking communication between the stars and the ground, provides on-board computing resources for navigation tasks, and comprises a star health data processing unit, a star fault diagnosis unit, an interstellar health data processing unit and an interstellar fault diagnosis unit. The application establishes an interstellar transmission mechanism of the health state data of the satellite constellation on-orbit fault diagnosis system, simultaneously copes with the requirements of on-orbit fault diagnosis accuracy and real-time through fusing expert knowledge and machine learning methods, and improves the on-orbit fault diagnosis and disposal effect.

The invention provides an initial orbit determination method for primary pulse maneuver of a non-cooperative target under a space-based single-sight observation condition. The method comprises the following specific steps: carrying out maneuvering detection by adopting a sight line vector component based on a sliding window, and judging whether a target is maneuvered or not according to an absolute value of a slope difference value; a 3r iterative algorithm is put forward in the initial orbit determination stage, estimation of the pulse speed is added on the basis of a traditional double-r iterative method, and a prediction orbit with the minimum residual weighted quadratic sum is obtained through a grid search and optimization combined method and serves as an initial orbit determination result; and optimizing an initial orbit determination result by using a batch processing method, constructing a Jacobian matrix and a measurement error covariance matrix, and outputting an estimated covariance for use in a subsequent filtering algorithm. According to the invention, maneuvering detection and high-precision initial orbit determination can be completed only by using the observation sight line vector, and the method has the advantages of simple structure and high efficiency, is suitable for real-time monitoring of non-cooperative maneuvering targets, and can provide reliable initial value conditions for subsequent filtering.

The invention discloses an earth-moon space transfer spacecraft orbit determination method based on LEODRO inter-satellite measurement, and the method comprises the steps: S1, constructing an orbit dynamics model of a spacecraft, a pseudo-range measurement model of the spacecraft, and a clock correction observation model of the spacecraft, and carrying out the linear processing of the pseudo-range measurement model and the clock correction observation model; s2, the LEO satellite adopts a configured GNSS positioning system to carry out orbit determination; an inter-satellite link between the LEO satellite and the DRO satellite is constructed, and the LEO satellite performs orbit determination on the DRO satellite through the inter-satellite link; s3, according to the orbit dynamics model and the linearization processing result of the LEO satellite, adopting an extended Kalman filtering orbit determination algorithm to complete orbit determination of the spacecraft by the LEO satellite; s4, correcting the spacecraft orbit state and the covariance matrix obtained by orbit determination in the step S3 according to a linear processing result of the DRO satellite, and completing spacecraft earth-moon transfer orbit determination under the LEO-DRO satellite; and S5, returning to the step S3 when the spacecraft orbit determination is not terminated, and terminating the spacecraft orbit determination method when the spacecraft orbit determination is terminated.

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.