28 results about "Orbital elements" 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.

Embodiments of the present application provide a space target orbit prediction method, a prediction device, an electronic device and a medium. The method comprises: obtaining two-line orbital elements of a space target at a current time, space environment data and perturbation coefficients, the perturbation coefficients being coefficients in a space target orbit prediction model; correcting the two-line orbital elements and / or the perturbation coefficients based on a pre-trained parameter correction model, the two-line orbital elements, the perturbation coefficients and the space environment data, to obtain corrected data, the corrected data comprising corrected two-line orbital elements and / or corrected perturbation coefficients; and predicting orbit information of the space target at a preset future time based on the space target orbit prediction model and the corrected data. The method makes the predicted orbit information more accurate.

The invention discloses a ground-based radar space object long-time energy accumulation method based on an orbital element model, and the method comprises the following implementation steps: 1, building an orbital element parameter space, and mapping the grid points of the parameter space into the distance between a target and a radar; and step 2, calculating an envelope position and a phase which need to be compensated according to the mapped distance history at each parameter space grid point. And step 3, compensating a corresponding phase for the echo signal after pulse compression along the calculated envelope position, and then performing coherent superposition, thereby realizing long-time coherent accumulation of echo energy. According to the method, the orbital elements are adopted to construct the parameter space, energy accumulation is carried out in the parameter domain, due to the fact that the orbital elements can better represent the actual motion of a space object, the energy accumulation effect can be improved, meanwhile, the orbital elements are small in parameter space dimension, and the algorithm calculation amount can be greatly reduced.

The application provides a method for analyzing a pointing angle of an inter-satellite link, comprising: obtaining orbital elements of a satellite and orbital parameters of a target satellite matched with an SGP4 model in a J2000.0 inertial coordinate system; determining a target recursion time t; obtaining positions of the satellite and the target satellite at the time t in the J2000.0 inertial coordinate system; determining a pointing vector from the satellite to the target satellite at the time t in the J2000.0 inertial coordinate system; determining the pointing vector in a satellite orbital coordinate system; determining the pointing vector in a satellite body coordinate system; determining the pointing vector in a satellite antenna coordinate system; determining a pointing angle of the satellite antenna coordinate system at the target recursion time; and determining whether a cycle number reaches a set cycle number, if the cycle number does not reach the set cycle number, continuously obtaining the pointing angle of the satellite antenna coordinate system at a next target recursion time until the pointing angle of the satellite antenna coordinate system at the set cycle number is obtained.

The invention provides an earth-moon transfer orbit iterative calculation method based on earth-moon relative motion, which comprises the following steps: firstly, calculating an ideal earth-moon transfer orbit number according to the deviation between the current flight time and the theoretical time, and taking the ideal earth-moon transfer orbit number as an iterative guidance target orbit; and then, by utilizing the characteristic that the shutdown time of the rocket in the flight section can be predicted in the iterative guidance calculation process, the influence of the flight time deviation on the earth-moon transfer orbit is compensated to the number of target orbits in real time, so that correction of a guidance program angle instruction is realized, and the rocket is controlled to fly towards a more ideal earth-moon transfer orbit all the time.

The invention relates to a multi-satellite collaborative observation probability estimation method for a window type target. The method comprises the following steps: inputting a statistical time range, an orbital element number and a load observation range of a guiding constellation A, and an orbital element number and a load observation range of a guided constellation B; and determining a to-be-observed area O, based on whether the constellation A is a geostationary orbit satellite, calculating the collaborative observation probability that the constellation A guides the low-orbit constellation B, and outputting a collaborative observation probability result. According to the method, the defect that the time dynamic characteristics of the window type target are not fully considered when the collaborative observation probability of the satellite to the land target is analyzed in the existing method is overcome, and the problem that the window type target collaborative observation success probability is inaccurately estimated in the existing method is solved.

The application aims to provide a remote sensing satellite descending node local time control method based on orbit height adjustment, which has simple control and high control precision. The method comprises the following steps: S1, acquiring orbit elements of a remote sensing satellite, wherein the orbit elements are instantaneous elements or two-line elements; S2, calculating a current descending node local time of the satellite based on the orbit elements; S3, calculating a drift deviation of the current descending node local time from a nominal descending node local time △T ; S4, calculating a current orbit plane rotation rate based on a J2 perturbation parameter of the earth; S5, calculating a difference between the current orbit plane rotation rate and an earth revolution rate; S6, designing an expected orbit plane rotation rate based on the difference calculated in the step S5; S7, calculating an expected orbit height according to the expected orbit plane rotation rate; and S8, adjusting the orbit height of the satellite to the expected orbit height. The application is applied to the technical field of remote sensing satellite measurement and control.

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 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 application relates to a satellite orbit and satellite attitude determination method and device after satellite-rocket separation. The method comprises the following steps: calculating a first position vector and a first velocity vector of a satellite in an epoch equatorial coordinate system according to orbit elements of a rocket in a second earth-centered equatorial coordinate system and a satellite-rocket separation time; determining a second position vector and a second velocity vector of the satellite in a J2000 earth-centered inertial coordinate system based on the first position vector and the first velocity vector; taking the second position vector and the second velocity vector as orbit initial values, and determining orbit parameters of the satellite before injection into an orbit by using an orbit recursion algorithm; determining an attitude direction cosine matrix of the satellite in the J2000 earth-centered inertial coordinate system at the satellite-rocket separation time based on a launch time; and taking the attitude direction cosine matrix as attitude initial values, and determining attitude parameters of the satellite before injection into the orbit by using a star gyroscope combination filtering method. The scheme can quickly determine the orbit and the attitude of the satellite, and improves the autonomous detection capability of the satellite.

The application discloses a scanning method and device of a non-cooperative target, electronic equipment and program product, and relates to the technical field of spaceflight measurement and control, wherein the scanning method comprises the following steps: acquiring initial orbit elements of the non-cooperative target, celestial body parameters of a target celestial body and antenna parameters of an antenna; determining the number of beam pointing positions required for the antenna to scan one orbit period of the non-cooperative target based on the initial orbit elements, the celestial body parameters and the antenna parameters; determining an acceleration multiple based on the orbit period of the non-cooperative target, the number of beam pointing positions and the antenna parameters; and generating guide data based on the initial orbit elements and the acceleration multiple, so as to control the antenna to scan the non-cooperative target through the guide data. The application solves the technical problem that the non-cooperative target cannot be efficiently scanned in the related art, thereby reducing the probability of capturing the non-cooperative target.

The invention discloses a calculation method and device for a satellite VDES load coverage area, a storage medium and calculation equipment, and the method comprises the steps: determining the position and attitude of a satellite according to the orbital element number of the satellite, a designated moment and a satellite working mode; determining the direction of a VDES antenna based on the position and attitude of the satellite and the load installation parameters; based on the VDES antenna pointing, constructing a polar plane of a satellite about the surface of an earth ellipsoid, and determining a parameter equation of a space ellipse where the polar plane intersects with the surface of the earth ellipsoid; according to the parameter equation, traversing angle parameters in the space ellipse to obtain a tangent point track and an intersection point track; and calculating a load coverage area of the VDES antenna based on the tangent point track and the intersection point track. The technical problem that the VDES load coverage area is difficult to calculate in the prior art is solved.

The invention discloses a satellite orbit thrust section inversion method based on pattern search, solves the problem of how to accurately invert parameters of a middle unknown thrust section by using two extremely short radar observation arc sections, and belongs to the field of satellite orbit thrust section inversion. Comprising the following steps: finding epoch moments of two observation arc segments; using a Laplace-LM algorithm to obtain an initial orbital element, and substituting the initial orbital element into POD to obtain a precise orbital element; constructing a satellite orbit dynamic model based on the initial dynamic parameters, and performing orbit propagation on a first precision orbit element of the first observation arc section from the first epoch moment to a second epoch moment of the second observation arc section to obtain a second propagation precision orbit element; constructing a target function according to the position difference between the second propagation precision orbit element and the second precision orbit determination result; introducing a pattern search algorithm to obtain a target dynamic parameter when the target function is minimum; according to the method, the thrust section can be accurately inverted, and the real propulsion strategy corresponding to the target dynamic parameter is summarized.

This invention discloses a method and system for interferometric imaging detection of satellite formations flying in the same direction in space. The method includes: setting initial parameters for a central satellite or virtual central satellite and N accompanying satellites in a formation; calculating the relative orbital elements of the N accompanying satellites; calculating the absolute orbital elements based on the relative orbital elements; predicting the position of each accompanying satellite based on the absolute orbital elements, and calculating the detection baseline formed by every two accompanying satellites; calculating the baseline distribution uniformity measurement function based on the detection baseline, using it as the objective function, and exploring optimization using an optimization algorithm to obtain the optimal solution; repeating the above steps based on the optimal solution to obtain a detection baseline with continuity and uniformity in both the radial and circumferential directions; and performing imaging inversion based on the detection baseline and the principle of interferometric imaging. This invention effectively improves the density of baseline coverage and reduces the deployment difficulty compared to deploying multiple satellites in a circle.

The present disclosure relates to a method and device for obtaining rigorous regression orbit parameters, and relates to the technical field of spacecraft, which can solve the problem of long time consumption of existing algorithms. The specific technical solution is: an orbit analytical method is adopted, and the J2 and J3 perturbation terms of the earth flattening are considered to obtain a set of rough regression orbit root numbers; a calculation function of the plane-instantaneous conversion is adopted to obtain the instantaneous orbit root numbers corresponding to the rough regression orbit root numbers; a numerical method is used to recursively correct the semi-major axis in the parameters of the instantaneous orbit root numbers, and the optimized initial value of the regression orbit root numbers is obtained; each orbit parameter and the regression period in the optimized initial value of the regression orbit root numbers are optimized and improved in a small range by using a nonlinear equation set, so that the rigorous regression orbit parameters with high regression accuracy are obtained. The present disclosure is simple and easy to understand in the calculation process, has high solving efficiency, short calculation time consumption, and can save a large amount of time for the design and analysis stage of the rigorous regression orbit.

This invention discloses a high-dynamic tracking method and system for low-Earth orbit satellites. The method is as follows: Step 1: Obtain the two-line orbital element number TLE of the satellite, and calculate the theoretical position and angular velocity of the satellite at the current moment using the orbital propagation model, as a priori guidance for antenna pointing; Step 2: Construct a particle swarm in a two-dimensional search space, where each particle represents a possible antenna pointing angle; Step 3: Collect the received signal strength (RSSI) corresponding to the positions of multiple particles in the particle swarm, and fit the current signal strength gradient vector using the least squares method based on the geometric differences in the spatial distribution of particles; Step 4: Calculate the position of each particle at the next moment according to the generalized velocity update equation, which includes three driving forces: swarm intelligence term, cooperative gradient term, and orbital feedforward term; Step 5: Convert the updated optimal swarm position into a servo control command to drive the antenna rotation.

The application discloses a kind of geostationary satellite multi-pulse high-precision autonomous position keeping control methods, comprising: according to the instantaneous orbital elements of satellite, the average orbital elements of satellite are determined;According to the average orbital elements of satellite, the in-plane and out-of-plane pulse control sequence of satellite orbit is determined;According to the in-plane and out-of-plane pulse control sequence of satellite orbit, the switch logic of thruster is determined;According to the switch logic of thruster, the switch of thruster is controlled, realizes geostationary satellite multi-pulse high-precision autonomous position keeping control.The method of the application realizes geostationary satellite multi-pulse high-precision autonomous position keeping control.

The invention relates to the technical field of satellite communication, and discloses a high-orbit and low-orbit satellite collaborative networking method and system, and the method comprises the steps: obtaining the orbit number of high-orbit and low-orbit satellites and the current phase of a low-orbit satellite, and predicting a future phase sequence based on an orbit dynamics model; a link quality prediction curve is generated in combination with a preset phase-link quality mapping relation, so that it is judged that a coverage blind area is about to enter before the link quality is lower than a service threshold value, and a compensation process is triggered; the compensation process comprises the following steps: predicting a departure moment, screening candidate high-orbit satellites and pre-allocating resources to a standby link before the departure moment, and then carrying out availability verification on the standby link; and after the verification is passed, migrating the service to the standby link before the separation moment, and releasing the original link resource. The system corresponds to the method. According to the invention, the active prediction and compensation of the coverage blind area are realized, and the service interruption caused by the high-speed movement of the low-orbit satellite is avoided.

The application provides a collision risk satellite combination processing sequencing optimization method, steps are as follows: according to TLE orbit root number data, orbit pre-division is carried out, satellites located on the same orbit plane are divided together; a perigee-apogee pre-screening method is used to carry out pre-screening of satellite collision combination, and satellite and debris combinations generating collision risks are obtained; TLE orbit root number data is combined with an orbit prediction model and a fast screening trajectory space intersection algorithm to determine the position of trajectory intersection and overlap in space; high collision risk satellite combinations and high collision risk regions are determined; a collision combination priority processing sequencing model is constructed in combination with satellite states and satellite characteristics; and a gradient descent method is used to optimize parameters of the multi-factor collision combination priority processing sequencing model. The application improves the screening speed of high collision risk combinations and improves the overall safety of a satellite constellation by using the geometric screening method of the intersection angle between two orbits and the gradient descent method to optimize multi-factor parameters.

The application discloses a physical decoupling adaptive envelope normalization maneuver detection method based on plane spring root numbers, and comprises the following steps: S1, acquiring a Cartesian state vector of a target satellite at a current time; S2, converting the Cartesian state vector into a non-singular plane spring orbit root number set; S3, performing adaptive normalization processing on a first-order difference value of each root number at the current time by using a first-order difference maximum noise envelope corresponding to each root number, and obtaining a normalized significance index of each root number; S4, based on the normalized significance index of each root number, three physically decoupled sub-detection factors are constructed; S5, the three physically decoupled sub-detection factors are compressed and fused to obtain a comprehensive maneuver detection statistic; and S6, the comprehensive maneuver detection statistic is compared with a preset detection threshold. The application can realize rapid and robust maneuver detection in a global orbit scene, especially in singular point and blind point working conditions in which classical root numbers fail.

The application discloses a satellite orbit thrust segment inversion method based on a mode search, solves the problem of how to accurately invert the parameters of an unknown thrust segment in the middle by using two extremely short radar observation arc segments, and belongs to the field of satellite orbit thrust segment inversion. The application comprises the following steps: finding the epoch time of the two observation arc segments; obtaining initial orbit elements by using a Laplace-LM algorithm and substituting the initial orbit elements into POD to obtain precise orbit elements; constructing a satellite orbit dynamics model based on initial dynamic parameters, performing orbit propagation on the first precise orbit elements of the first observation arc segment from the first epoch time to the second epoch time of the second observation arc segment to obtain second propagated precise orbit elements; constructing an objective function according to the position difference between the second propagated precise orbit elements and the second precise orbit determination result; and introducing a mode search algorithm to obtain target dynamic parameters corresponding to the minimum value of the objective function.

The invention provides a two-line orbit element acquisition method and a related device, and relates to the technical field of satellites. The two-row orbit element acquisition method comprises the following steps: acquiring position information and speed information of a target satellite at a target moment; converting the position information and the speed information into six instantaneous Kepler elements as initial values of the two rows of orbits; according to a preset correction strategy, correcting the initial value of the number of the two lines of orbits to obtain a corrected value of the number of the two lines of orbits; wherein the preset correction strategy comprises that at least one of an area integral constant, an orbit eccentricity rate and an orbit semi-major axis is used as a parameter correction convergence condition in the correction process. According to the technical scheme, the problem that in the prior art, the number of two lines of tracks is difficult to obtain can be solved.

The application provides a ground-visible space display system and relates to the technical field of satellite display. The system comprises annular display arrays, each annular display array comprises a plurality of satellite stars, the plurality of satellite stars perform relative annular flying motion around a virtual main star, a plurality of annular flying orbits are centered on the virtual main star, and the plurality of annular flying orbits are concentrically nested. The virtual main star is located on an elliptical orbit centered on the earth, the configuration parameters of the satellite stars satisfy a preset condition, and the orbit root number difference between the configuration parameters of the satellite stars and the orbit parameters of the virtual main star satisfies a correlation configuration model. By setting the configuration parameters of the satellite stars and the orbit parameters of the virtual main star, the orbit parameters of the satellite stars are obtained, the annular display arrays are formed, the brightness and color of the light emitting devices on the satellite stars are controlled, the expected display pattern is presented, the observer can see the delivered information without communication media, and global information delivery and transmission can be realized.

The application discloses a longitude-altitude coordinate system-based orbit extrapolation method, and solves the problem of low numerical propagation efficiency, and belongs to the field of orbit extrapolation; the method comprises the following steps: constructing a substar point longitude model based on orbit elements; obtaining a true anomaly model by solving; obtaining an optimized true anomaly according to a latitude amplitude iteration parameter; obtaining a first-order derivative and a second-order derivative of the true anomaly based on the optimized true anomaly, and combining the optimized true anomaly to form a derivative parameter; introducing the derivative parameter to obtain a first-order derivative and a second-order derivative of the substar point longitude; constructing an orbit altitude model from the difference between the spacecraft geocentric distance and the earth radius; introducing the derivative parameter to obtain a first-order derivative and a second-order derivative of the orbit altitude, and combining the first-order derivative and the second-order derivative of the substar point longitude to construct a state space equation; integrating the state space equation by using a numerical integrator, extrapolating a four-dimensional space state vector, and obtaining a longitude-altitude propagation model; and the method improves the numerical propagation efficiency.

This invention provides a method for creating a point-beam data transmission mission for a shipborne mobile receiving station to remote sensing satellites. The method includes: using the acquired approximate position and orbital element information of the shipborne receiving station, calculating a set of all approximate visible time windows for the entire day; then using this set to determine the approximate time window for the target data transmission mission; based on the approximate time window, sequentially calculating the ship's geocentric latitude, ship's travel distance, and spherical arc length; determining, based on the spherical arc length, whether the ship's course change exceeds a preset threshold; if so, performing segmented ship position prediction; otherwise, performing direct ship position prediction to obtain a precise predicted ship position; expanding the approximate time window based on the precise predicted ship position to obtain a precise time window; and then using the precise time window and the precise predicted ship position to construct the point-beam data transmission mission. This invention achieves precise planning of the satellite data transmission mission and ensures that the shipborne station antenna points towards the satellite to form a two-way closed link.

The application discloses a remote sensing shooting window planning method, device, equipment, medium and product, relates to the satellite shooting field, and the method comprises the following steps: acquiring the orbital elements of a satellite and the coordinates of a desired shooting point; calculating the instantaneous three-dimensional position and the satellite speed of the satellite according to the orbital elements; performing coordinate system conversion on the instantaneous three-dimensional position and the satellite speed and determining the satellite longitude, latitude and height coordinates; determining the real-time attitude angle parameters of the satellite according to the converted instantaneous three-dimensional position, the converted coordinates of the desired shooting point and the satellite longitude, latitude and height coordinates; determining the pan-tilt rotation angle according to the coordinates of the desired shooting point, the satellite longitude, latitude and height coordinates and the real-time attitude angle parameters; and the pan-tilt rotation angle is used for adjusting the remote sensing shooting window. The application can improve the autonomy of the shooting window, and improve the time matching accuracy and imaging positioning accuracy of the earth observation.

The invention relates to the technical field of spaceflight measurement and control and space situation awareness, and discloses a satellite orbit maneuver detection method based on multichannel residual imaging, which comprises the following steps: acquiring an orbit element actual observation value, acquiring an orbit element theoretical value, calculating an orbit element residual sequence based on the orbit element actual observation value and the orbit element theoretical value, and calculating the orbit element maneuver according to the orbit element residual sequence. The method comprises the following steps of: extracting multiple feature channels, respectively carrying out Gramer angle field transformation on a residual sequence of each feature channel to obtain a multi-channel GAF image, constructing a classification model based on a deep neural network, and carrying out feature extraction and orbit maneuver judgment on the multi-channel GAF image. According to the method, the Gramb angle field transformation is applied to satellite orbit data processing for the first time, unequally spaced TLE time sequences are converted into regular two-dimensional images, the mature computer vision technology can be directly applied to maneuvering detection tasks, the time sequence correlation of the time sequences is reserved through GAF transformation, the time sequences are coded into space structure features of the images, and the space structure features of the images are obtained. And deep learning model extraction and learning are facilitated.

The application discloses an in-plane small control quantity numerical iteration evaluation method based on orbit phase evolution, and through satellite pre-control and post-control orbit roots and high-precision orbit dynamics model, the time integral effect of in-plane small control quantity is utilized, the phase difference of post-control satellite orbit is taken as a control effect evaluation criterion, and in view of the limited calculation precision of an analytic model, numerical iteration is adopted in calibration, so that the effectiveness of the calibration method is ensured. The application can improve the control effect evaluation precision of a spacecraft when performing in-plane adjustment of a small control quantity orbit, thereby improving the orbit control precision of the satellite, effectively saving the fuel consumption of the satellite, and the method has high reliability, strong operability, and is easy to popularize and use, and has certain economic benefits for on-orbit operation of the spacecraft and important guiding significance for mission implementation.