285 results about "Orbit determination" patented technology

The invention discloses an enhanced information generation method and system fusing SBAS and PPP, and relates to a satellite navigation enhancement technology, and the method comprises the steps: receiving observation data of an SBAS and a global real-time monitoring station, and carrying out the file storage according to a standard format; carrying out precise orbit determination and forecasting by utilizing a stored standard format file and adopting a simplified dynamics method, and carrying out continuous calculation by taking one hour as a sliding window to generate a real-time precise orbit of the satellite; based on a satellite real-time precision orbit and real-time monitoring station data, a non-difference observation model and a sequential least square method are adopted for parameter solving, and a satellite real-time precision clock error is generated; solving an ephemeris and star clock differential correction number based on the broadcast ephemeris by using a satellite real-time precise orbit and a precise clock difference; and solving a covariance matrix and an integrity parameter by using the differential correction of the star calendar and the star clock. According to the method, two navigation enhancement technical means are fused, and the high-precision and high-integrity navigation positioning requirements are met.

The invention discloses an LEO on-orbit real-time precision orbit and clock error determination method and system based on PPP-B2b enhancement, and belongs to the technical field of low-orbit satellite high-precision autonomous orbit determination and clock error estimation. According to the method, enhanced information broadcasted by Beidou No.3 B2b signals is used for correcting a BDS-3 / GPS satellite orbit and clock correction, under the support of the regional BDS-3 / GPS enhanced information, when a new BDS-3 / GPS tracking satellite exists, before measurement updating, the inter-epoch clock correction variable quantity obtained through calculation by using a carrier phase time difference method is used as two orbit determination filtering clock correction parameters for constraint, so that the measurement accuracy is improved. In-orbit real-time separation of receiver clock error and non-difference floating point ambiguity parameters is carried out; and finally, substituting an LEO precision orbit output by orbit determination filtering and a non-difference floating point ambiguity parameter into a phase observation equation to calculate the precision clock error of the BDS-3 / GPS receiver. The method is suitable for the situation that only regional GNSS enhancement information support exists, and the influence of orbital maneuver is small.

The invention discloses an initial orbit determination method for passive angle measurement of a ring fire orbit for Mars sampling return, and the method comprises the steps: building an absolute dynamic model, obtaining an absolute motion state of a spacecraft, and deducing a relative motion state X; constructing a data set by using the measurement quantities H and X of the sensor; a network architecture is designed, input data I are position and speed coordinates of the tracker in a Mars equator inertial coordinate system and angle information converted from H, and output data are X; and finally, adjusting an observation interval dt and times k according to a ring fire observation condition, training a physical information neural network by using limited observation data, defining a total loss function F comprising a loss function FY and a physical law penalty term FD, and optimizing a weight coefficient and bias by minimizing the loss function through an F gradient descent method. According to the method, the neural network is trained offline, a network mapping model from a target line-of-sight angle measurement sequence to a target orbit is constructed, and the network mapping model is solidified on a returner for online use, so that sampler short-arc passive measurement rapid orbit determination is realized.

The invention discloses an astronomical positioning method for a space target elongated astronomical image of a large-field-of-view space-based monitoring platform, relates to the technical field of astronomical observation, and solves the problem that the elongated astronomical image is abandoned due to incapability of target association and precise orbit determination caused by inaccurate centering of the elongated astronomical image and low precision of a positioning method of an existing monitoring platform. According to the method, an observation image is divided into 16 sub-view fields, and each sub-view field is an independent view field and is calculated by an independent processor; removing fixed stars in the observation image by adopting an adjacent frame image difference method and extracting a space target star image; secondly, describing the length and angle of the lengthened astronomical image by adopting a minimum enclosing rectangle, and constructing a degradation model; performing motion blurred image restoration on the elongated star image; error points with large errors are eliminated by adopting a mutual comparison method, and track association of the space target is completed; and finally completing precise astronomical positioning of the elongated astronomical image and the like. According to the method, the processing efficiency is greatly improved, and a large number of lengthened astronomical images which are abandoned before are fully utilized again.

The invention provides a GNSS-based real-time positioning service stability improving method and system, and belongs to the field of satellite navigation and positioning services. According to the method, abnormal identification and elimination are carried out on monitoring receiver data with pseudo-range hopping and pseudo codes inconsistent with carriers, multiple receivers synchronous with a time reference are adopted to participate in clock correction calculation, the effectiveness of precise orbit determination and clock correction calculation is judged and identified, and meanwhile the consistency of calculation results is judged and repaired; estimating a satellite initial position vector and kinetic parameters, and forecasting a satellite orbit and a satellite clock error; and when the precise orbit determination solution and clock correction solution processing process is interrupted, the satellite orbit forecast result and the clock correction forecast result are output as a precise orbit and a clock correction, and the output robustness is guaranteed. According to the invention, the algorithm software of the system end is compensated, so that the calculation and storage burden of the user terminal is reduced, and the availability, continuity and stability of the real-time precise positioning service of the satellite navigation system are improved.

The invention discloses a low-orbit target orbit determination method and device. The method comprises the following steps: acquiring observation data at different moments; determining a target position vector according to the observation data and a pre-established target position relation model; the target position vector comprises a first target position vector corresponding to an initial observation moment, a second target position vector corresponding to an intermediate observation moment and a third target position vector corresponding to a termination observation moment; the target position relation model is determined based on a second target position vector corresponding to the intermediate observation moment and a position vector of the observation platform relative to the earth center; and determining a predicted target orbit according to the initial observation moment, the first target position vector, the termination observation moment and the third target position vector. According to the method, the target orbit parameters can be determined only based on the observation data at different moments under the conditions that an initial value is not needed and perturbation errors exist, and the accuracy and stability of low-orbit target orbit determination are improved.

The invention discloses a space-based directional observation method and device based on orbit precision evaluation. The method comprises the following steps: acquiring fixed star angular distance measurement data information, a Ka-band inter-satellite link observation value and an L-band satellite-ground link observation value of a Beidou satellite; using the information to construct a multi-dimensional observation matrix and carrying out resolving processing to obtain satellite position parameters corresponding to the three space-based directional observation modes; and performing comparison processing on the satellite position parameters by taking an MGEX precise ephemeris as a comparison baseline to obtain an optimal observation mode. Therefore, three space-based directional observation modes are provided, the orbit precision is used as an evaluation criterion, an observation mode system with high adaptability and high practicability is formed, and the orbit determination precision of the Beidou satellite is improved.

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 provides a satellite Ka wave band inter-satellite link antenna phase center change calibration method and system, and the method comprises the steps: collecting L wave band observation data and Ka wave band inter-satellite link observation data of a ground observation station, carrying out the gross error elimination, cycle slip detection, epoch normalization and combination, and generating a clock error-free and geometry-free observation quantity; analyzing the variation range of a nadir angle and an azimuth angle of the observation data, constructing a piecewise linear original PCV function model taking the nadir angle and the azimuth angle as independent variables, and synchronously resolving original PCV parameters of a satellite L / Ka wave band in a joint orbit determination process; decomposing the original PCV into a smooth PCV and a corresponding Z-direction PCO by adopting a least square separation technology, performing interpolation on a PCV value based on a discrete grid, and establishing a final PCV grid correction model; joint orbit determination error correction is introduced into the PCV correction model, a new round of joint orbit determination is carried out, the satellite L / Ka waveband PCO correction amount is calculated, and high-precision in-orbit calibration of the satellite Ka waveband inter-satellite link antenna phase center change is achieved.

A method for calibrating a PCO and a hardware delay of a downlink navigation antenna of an LEO satellite Ls includes: calculating, by using a precise orbit determination and timing result and ground calibrations, a ground calibrated phase center orbit initial value of the downlink navigation antenna and a ground calibrated satellite clock bias initial value; calculating, based on the ground calibrated phase center orbit initial value and the ground calibrated satellite clock bias initial value, a correction amount of the downlink navigation antenna by separating or combining GNSS signals and LEO satellite downlink navigation signals; and correcting, based on the correction amount of the downlink navigation antenna, the PCO of the downlink navigation antenna and the hardware delay of the downlink navigation antenna, so as to realize in-orbit calibration of the PCO of the downlink navigation antenna and the hardware delay of the downlink navigation antenna.

The present disclosure relates to a method of generating auxiliary data for controlling a satellite travelling in orbit around Earth, the method comprising: receiving tracking data for the satellite; applying an orbit determination algorithm including: estimating an orbit for the satellite based on the tracking data; and predicting, based on the estimated orbit, future ephemerides data of the satellite; generating auxiliary data comprising predicted future ephemerides data; and transmitting the auxiliary data to the satellite for use in the satellite's attitude and orbit control.

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 provides an earth rotation parameter determination method and system based on a GNSS data stream, and the method comprises the steps: achieving the purpose of increasing the estimation of earth rotation parameters in a process of achieving precise orbit determination through employing square root information filtering, regarding the acceleration values of polar shift and UT1 change as white noise in time updating, and forming a state equation; in the measurement updating process, a UT1 value in an ultra-fast ERP forecasting product provided by the IGS is added to serve as virtual observation to form a virtual observation equation, so that UT1 parameters and orbit parameters are decoupled and can be estimated, stable estimation of real-time earth rotation parameters is achieved, the precision is superior to that of earth rotation parameter forecasting products provided by the IGS and IERS, and the accuracy of the earth rotation parameter forecasting products provided by the IGS and the IERS is improved. And meanwhile, a precise track product with higher precision is obtained, and the real-time high-precision ERP application requirement can be better met.

The invention discloses a maneuvering spacecraft orbit determination method for adaptive state noise compensation, and belongs to the technical field of aerospace. The implementation method comprises the following steps: establishing a general dynamical model and an angle-only observation model for unknown mutation of the random nonlinear dynamical system; predicting a system state and a covariance matrix based on the dynamic model; calculating a system measurement residual error and a covariance matrix at the tk + 1 moment, and representing the matching degree of a prediction result of the system based on the kinetic model and the maneuver-free model through an index function lambda k + 1; comparing the index function at the tk + 1 moment with a mean value of the index functions from t1 to tk, when the difference between an evaluation index and the mean value is greater than a threshold value, judging that the state of the system is abruptly changed, and performing state noise compensation on a system prediction result; and calculating orbit determination algorithm gain based on a system state noise compensation result, weighting and outputting a tk + 1 system state and a covariance estimation value based on a system prediction result # imgabs0 # noise compensation result # imgabs1 # and a measurement result zk + 1, and iteratively solving a state estimation value at each moment so as to realize high-precision orbit determination of the maneuvering spacecraft.

The invention discloses an ionosphere time delay monitoring and orbit determination system and a related method, and belongs to the technical field of ionosphere time delay monitoring. The ionosphere time delay monitoring and orbit determination system comprises a satellite and a monitoring station which are in communication connection; the monitoring station broadcasts at least one uplink signal to the satellite, the satellite forwards at least two downlink signals to the monitoring station based on the at least one received uplink signal, and two forwarding ranging values are obtained through synchronous measurement; and determining the total electron number of the ionized layer on the ranging signal path between the satellite and the monitoring station and the satellite-ground space distance between the satellite and the monitoring station based on the two forwarding ranging values. The ionosphere time delay monitoring method provided by the invention has the advantages of high real-time performance, accurate monitoring result and the like.

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 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.

The invention provides a group optimization-based space-based angle-measurement-only short arc initial orbit determination method and system. The method comprises the following steps of: 1, setting fixed parameters; 2, initializing particles in the group; step 3, grouping the particles; 4, for all the particles, if the grouping identifier is 0, executing the step 5, otherwise, executing the step 6; 5, performing variation search and update on the particles in multiple directions; 6, updating the particles according to the tracking direction of the particles; 7, calculating a global optimal particle; and 8, repeating the steps 3-7 until a quit cycle condition is met, and outputting the orbit state quantity corresponding to the current global optimal particle. The method is easy to implement, has a certain anti-interference capability, and can stably and effectively solve the space-based angle-only-measurement short arc initial orbit determination problem.

The invention provides a time service and time keeping method based on satellite-borne GNSS high-precision real-time orbit determination, which has the function of resolving the clock error and clock drift of a receiver in real time, and can perform time model correction by using the clock error and clock drift, thereby finally realizing high-precision time service and time keeping functions. The method is characterized in that rough time information is extracted from a message of a navigation signal; performing high-precision orbit determination through a dual-frequency carrier phase, and resolving a clock error and clock drift in real time; calculating the hardware time delay of the whole signal link by using a high-precision hardware time delay estimation technology; real-time frequency modulation and phase modulation are carried out by using clock drift and clock difference information, and a high-precision time service technology is realized; after a navigation satellite signal is interrupted, a local time model is maintained and PPS (pulse per second) is output by utilizing orbit extrapolation, clock error and clock drift information and a time model, so that a high-precision time keeping function is realized; fusion and advantage complementation of the constant-temperature crystal oscillator and the GNSS receiver are realized, and the cost of high-precision time service is reduced.

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 discloses a point cloud intersection-based space target optical segmental arc rapid association method. The method comprises the following steps: 1, acquiring optical segmental arc data and constructing a constraint admissible domain; 2, generating a track point cloud; 3, forecasting parallel orbit point cloud; 4, fast point cloud intersection is carried out; and 5, determining an initial orbit based on point cloud intersection. Through the above process, the space target rapid segmental arc correlation method based on point cloud intersection is provided, an admissible domain is constructed by two optical observation segmental arcs and orbit point clouds are generated, and then the point clouds of the two segmental arcs are forecasted to the same moment through a parallel orbit point cloud forecasting technology. Quickly locking an overlapping region of the two point clouds by adopting a minimum bounding box method, and judging the relevance between the point clouds through a mahalanobis distance; and finally, based on the Bayesian estimation principle, carrying out initial orbit determination on the successfully associated arc segments.

The invention discloses a space target multi-satellite collaborative observation task planning method based on a greedy algorithm, which belongs to the technical field of spaceflight, and comprises the following steps: firstly, carrying out visibility analysis on a space target, and comprehensively considering factors such as earth shadow shielding, illumination conditions, detection distance and target characteristics; designing a method for calculating the visibility of the optical satellite to the space target; secondly, designing a hybrid heuristic rule, and preprocessing the visible window set; finally, a task planning method based on conflict resolution and priority dynamic adjustment is designed, and efficient arrangement of observation tasks is guided through a preset hybrid heuristic rule. According to the method, observation requirements of space target orbit determination are strictly considered, a multi-satellite multi-target multi-frequency observation plan can be rapidly formulated, observation of important targets is preferentially guaranteed on the premise of limited resources, meanwhile, other targets are considered, the requirements of space situation awareness tasks are met, and the method has practical engineering application value.

The embodiment of the invention discloses a low-orbit navigation satellite clock correction calculation method and device, electronic equipment and a storage medium, and the method comprises the steps: carrying out the ionosphere-free combined precise orbit determination of observation data received by an LEO through the satellite orbit parameters and clock correction products of a medium-high-orbit GNSS satellite, and obtaining the satellite orbit parameters of the LEO and the clock correction of a signal receiving end; based on the double-frequency mixed observation data, jointly resolving the signal transmitting end clock error of the LEO in the ground visible segmental arc; mathematical models are established for the signal receiving end clock error of the LEO and the signal transmitting end clock error of the LEO in the ground visible arc section respectively, and the difference of the clock error models is extracted; and according to the difference of the clock difference model, converting the clock difference of the signal receiving end of the LEO into the clock difference of the full-arc signal transmitting end of the LEO.

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 giant astronomical base partition centralized orbit determination and time synchronization method and system under the support of inter-satellite measurement, and relates to the field of giant astronomical bases, and the method comprises the steps: carrying out the space partitioning of all satellites in a giant astronomical base through employing a specific topological configuration, and obtaining a plurality of regional grids; node satellites in the current regional grid perform centralized Kalman filtering measurement updating by using all observation data to obtain a first updating result; sending the first updating result to each satellite in the current regional grid, and controlling each satellite to perform distributed Kalman filtering time updating to obtain a second updating result; the second updating result is sent to the node satellites in the current regional grid; and controlling the node satellites in the current regional grid to send all the received second update results to the node satellites in the adjacent regional grid, and repeating the steps to complete orbit determination and time synchronization of satellites of multiple batches. According to the invention, precise orbit determination and time synchronization of satellites in a giant satellite base are realized.

The invention provides a space-based optical observation high-orbit space debris short arc initial orbit determination method and system. The method comprises the steps that S1, angle measurement information of imaging of the same unidentified space debris by a space-based optical observation platform is acquired; s2, acquiring observation data according to the observation capability of the space-based optical observation platform, and setting the upper limit and the lower limit of the distance between the high-orbit space debris and the space-based optical observation platform; s3, calculating an actual observation distance upper limit and an actual observation distance lower limit corresponding to the current observation arc section according to the observation data; s4, randomly initializing a plurality of groups of particles within the upper limit and lower limit range of the actual observation distance; s5, calculating a loss function corresponding to each group of particles; s6, updating each group of particles, calculating a loss function corresponding to the updated particles, and recording each group of optimal particles and global optimal particles; and step S7, repeating the step S6 until the cycle quitting condition is met. The method is easy to implement, and the success rate of determining the high-orbit space debris short arc initial orbit can be increased.

The invention discloses a wavelet noise reduction multi-path extraction method for improving a threshold function, which can be used for effectively extracting and weakening multi-path errors of a global navigation satellite system (GNSS), solves the problem that signals of wavelet coefficients are discontinuous or distorted at a preset threshold, improves the traditional wavelet noise reduction effect and multi-path error modeling precision, and improves the accuracy of multi-path error modeling. Therefore, the influence of multi-path errors on GNSS positioning and orbit determination precision is effectively weakened, a feasible solution is provided for realizing high-precision positioning and orbit determination in a complex environment, and a solid foundation is laid for establishing 1mm space-time reference in the future.

The invention discloses an integrity monitoring method for real-time filtering timing and orbit determination of a low earth orbit satellite. The method comprises the following steps: acquiring a complete upper bound covariance matrix of a low earth orbit satellite system observation model of a current epoch; calculating an upper bound form covariance matrix of unknown parameters in the low earth orbit satellite system observation model according to the complete upper bound covariance matrix; the unknown parameters comprise clock correction and orbit parameters; obtaining a first upper bound form standard deviation of the clock error and a second upper bound form standard deviation of the orbital parameter according to corresponding elements of the clock error and the orbital parameter in the upper bound form covariance matrix; determining a first impact on the clock error and a second impact on the orbital parameter from the deviation transmissions of the plurality of epochs; obtaining a first protection level according to the first upper bound form standard deviation and the first influence, and obtaining a second protection level according to the second upper bound form standard deviation and the second influence; and integrity monitoring of low-orbit satellite real-time filtering timing and orbit determination is carried out based on the first protection level and the second protection level, so that the accuracy of integrity monitoring is improved.

The invention discloses a GEO binocular orbit determination method and system based on inter-satellite communication, and belongs to the technical field of aerospace, and the method comprises the steps: predicting the priori estimation of the state variable of a binocular orbit determination system through unscented transformation according to an orbit dynamics model of the binocular orbit determination system; simulating and generating space-based observation data based on an observation model of a binocular orbit determination system; according to the principle of unscented Kalman filtering, the priori estimation of the state variable and the space-based observation data are fused, and the posteriori estimation of the state variable of the binocular orbit determination system is obtained; repeatedly iterating until the posteriori estimation orbit determination precision of the state variable of the binocular orbit determination system meets the task requirement; according to the method, target observation data is obtained through binocular observation, observation data between observation satellites is obtained by considering inter-satellite communication, prediction data and the observation data are fused through an unscented Kalman filtering algorithm, and a high-precision orbit determination method is provided for a GEO spacecraft.

The embodiment of the invention discloses a multi-platform common-view orbit determination method for an earth-moon space target, and relates to the technical field of space target orbit determination, and the method comprises the steps: enabling a plurality of observation platforms to observe the same earth-moon space target in the same observation time period, and obtaining a motion data sequence of the observation platforms in the observation time period, obtaining an observation data sequence obtained by observing the earth-moon space target; determining a relative position relationship between the observation platform and the earth-moon space target, constructing a first observation model for the earth-moon space target, and calculating the position and speed of the earth-moon space target at each moment in a corresponding observation time period; and constructing a second observation model, and calculating to obtain an orbit arc section of the earth-moon space target in the corresponding observation time period. The method is small in type limitation of earth-moon space targets, and can adapt to multiple types of orbits; after the initial orbit is successfully determined, orbit refinement can be carried out again by combining original observation data, so that the initial orbit precision is improved.