38results about How to "Improve the accuracy of orbit determination" patented technology

The invention discloses a space-based phased-array radar space multi-target orbit determination method, which mainly solves the problems that the space weak target cannot be evaluated effectively and the target orbit determination precision is low in the prior art. The method comprises the following steps of: processing target echo data by a zero-setting conformal algorithm, acquiring distance prior information by utilizing a distance pulse compression principle, and segmentally processing echo signals according to the distance prior information; segmenting the echo signals according to the number of the targets, wherein each segment of data is a target signal and an adjacent unit signal; performing multi-target two-dimensional angle evaluation on each segment of data by utilizing a sum-difference multi-beam angle-measuring principle; performing coordinate conversion based on a space target tracking result and detection satellite orbit information; and performing orbit determination on different space targets by a Laplace type iterative algorithm, and improving the orbit determination precision by a least square algorithm. According to the method, the influence of strong signals on weak targets can be reduced and the parameters of the weak targets can be evaluated accurately. The method can be applied in the actual application fields of space situation awareness, orbit resource management and the like.

The invention discloses an autonomous orbit determination method of a Lagrange navigation satellite. The method comprises the following steps: acquiring three groups of inter-satellite ranging information by virtue of a navigation constellation containing at least four satellites; updating a neural network weight by utilizing the inter-satellite ranging information; estimating a nonlinear perturbing term according to the neural network weight; and structuring a neural network state observer by utilizing the obtained nonlinear perturbing term, and estimating orbit information of the Lagrange satellite. Based on elliptic restrictive three-body problems, the perturbing force suffered by the Lagrange navigation satellite is precisely estimated by virtue of the neural network, so that the model precision of the determined orbit is improved; and the state observer is used for precisely estimating the state of the Lagrange navigation satellite, so that no limitation is made for the system noise and the observation noise, and the universality is better.

The embodiments of the invention disclose an orbit determination method, device and system of a low-orbit satellite. Firstly, broadcast ephemeris sent by a navigation satellite is received, and the orbit and clock difference of the navigation satellite are calculated and obtained according to the broadcast ephemeris. Then, the correction information of the positioning information sent by a groundsystem is received to correct the positioning information, and the corrected positioning information is obtained. Then the revised positioning information and observation data are used for determiningthe orbit of the low-orbit satellite. That is, the embodiments of the invention introduce the correction information of the positioning information, and the accurate corrected positioning informationis used for determining the orbit the low-orbit satellite, so that the orbit determination accuracy of the low-orbit satellite is improved.

The invention relates to a satellite passive orbit measurement method and system, and relates to satellite technology. The satellite passive orbit measurement method of the present invention is to passively receive the radio signal broadcast by the satellite through the ground station, so as to realize the position measurement and orbit measurement of space objects such as satellites. The specific method is to set up 2 to 4 antennas in each measurement station, and use carrier signal waveform correlation to obtain the range difference D between the two antennas when receiving satellite signals. The rendezvous measurement model obtains the position and orbit of the satellite in space. While completing the above-mentioned rough orbit determination, 10-100 satellite position accuracy monitoring reference stations shall be established within the coverage area of ​​satellite signals. The vector correction number of the satellite position is obtained, so as to achieve the second high-precision orbit determination of the satellite.

The invention discloses an image-based remote sensing satellite multi-mode autonomous orbit determination method. The remote sensing satellite multi-mode autonomous orbit determination method comprises the following steps of: step 1, preprocessing a remote sensing image; step 2, matching the preprocessed remote sensing image with a ground feature point library and a remote sensing image at a previous moment, and extracting orbit determination feature points; step 3, establishing a multi-mode orbit determination measurement equation based on monocular vision and multi-view geometric principles;step 4, establishing an orbit determination state equation based on an orbit dynamics model; step 5, and performing high-precision real-time optimal estimation on an orbit determination state quantity through adopting a filtering algorithm based on the orbit determination measurement equation and the orbit determination state equation. According to the remote sensing satellite multi-mode autonomous orbit determination method, the problems of poor orbital determination autonomy and low orbital determination precision of the remote sensing satellite are solved, and the multi-mode autonomous orbital determination observation model based on the image is constructed according to the monocular vision and multi-view geometric principle, so that the improvement of the precision and reliability ofautonomous orbital determination of the remote sensing satellite and the enhancement of the autonomy are realized.

The invention discloses an atmospheric resistance perturbation modeling and calculating method for a low earth orbit satellite. The method is specifically implemented according to the following steps:1, establishing a Box-Wing model of a satellite; 2, establishing a piecewise linear atmospheric resistance perturbation model; 3, according to the Box-Wing model established in the step 1 and the piecewise linear atmospheric resistance perturbation model established in the step 2, calculating the folded atmospheric resistance perturbation acceleration at the moment t. By utilizing the method to calculate the atmospheric resistance perturbation of the satellite, atmospheric resistance perturbation calculation errors caused by the fact that a traditional single atmospheric resistance factor isdifficult to accurately describe atmospheric changes can be overcome, the orbit determination precision is effectively improved, and technical support is provided for precise orbit determination of alow-orbit satellite, especially precise orbit determination under special conditions such as magnetic storm.

The invention belongs to the technical field of space target data processing, and discloses a geometric method for determining a space target initial orbit through optical telescope common-view observation and a system, which directly solve target position information through employing the geometric method in a triangle formed by a space target and two observation stations based on binocular telescope common-view observation data, and directly solve speed information of the target by using a difference method according to the continuity of an observed quantity. The geometric method specifically comprises the following steps of: synchronizing the time of two pieces of observation data through adopting an interpolation method; calculating the distance of the target in a triangle ABS by usinga geometric approach; and acquiring a position vector of the target through utilizing the azimuth and distance information. Compared with a vector method in the prior art, the geometric method has the advantages that two unknown numbers are solved by two equations in the geometric method, two unknown numbers are solved by three equations in the vector method, the geometric method is accurate, andthe solution is unique. The invention further provides a position difference method for solving the speed, and the whole solving process completely gets rid of a mechanical equation; and the positiondifference method has obvious advantages for an orbital transfer target.

The invention relates to a navigation satellite orbit determination and improvement non-singular point technology under nonconservative force effects. The technology is characterized by comprising obtaining a theoretical expression of a nonconservative force perturbation through a Gauss equation, forming a theoretical expression of a disturbed orbit without singular points according to a Gauss-Xu equation, and performing theoretical fitting according to observation to as to determine a navigational satellite orbit or improve the orbit. Since the orbits of beidou navigation satellites are all near-circular or near equator, and the singular points exist in all classic orbit theories, orbit determination or orbit improvement fails. The technology provided by the invention can be applied to the orbit determination or improvement of the beidou navigation satellite orbits, and no singular points exist. Besides, the technology is simple, computational complexity is reduced, and the technology is suitable for satellite automatic orbit determination. Since a satellite perturbative force comprises two kinds, i.e., a potential force and a nonconservative force, the invention, together with the invention with an invention right registry number 201310545077.2, i.e., a navigation satellite orbit determination and improvement non-singular point technology under the potential force, forms a complete non-singular point orbit determination and improvement technology through combination application.

The invention discloses a system and method suitable for combined orbit determination of networking low-orbit satellites. The system comprises an LE0 constellation ground centralized orbit determination module and an LEO constellation distributed orbit determination module. The LE0 constellation ground centralized orbit determination module is used for acquiring observation data and measurement data, and acquiring predicted orbit information of each LE0 constellation satellite according to the observation data and the measurement data; and the LEO constellation distributed orbit determination module is used for acquiring the orbit information of the LEO constellation satellite according to the measurement data and the predicted orbit information. The invention aims to provide a system and a method suitable for combined orbit determination of networking low-orbit satellites, and solves the problem that autonomous, real-time and high-precision orbit determination of LEO constellation satellites cannot be met under the condition that only a part of areas have LEO satellite ground tracking stations.