21 results about "Relative orbit" patented technology

The invention relates to the technical field of spacecraft orbit control and space safety, and discloses a space debris orbit prediction and avoidance method, which comprises the following steps that: a ground computing center screens candidate targets and generates enhanced data packets for uploading; the spaceborne computer combines real-time navigation data to carry out geometric feature screening, and an instant high-risk target list is generated; controlling the star sensor to execute optical observation, and resolving an optical measurement relative orbit state by utilizing displacement superposition and Kalman filtering; selecting a high confidence coefficient or a conservative probability threshold according to the optical measurement state acquisition condition, and generating an avoidance instruction when the collision probability exceeds the limit; and the propulsion system responds to the instruction to execute avoidance, verifies the effect by confirming the semi-major axis variation and executes track recovery by selecting an aircraft. According to the method, the ground-air collaborative screening and shift superposition enhancement technology is adopted, the problems that the low-precision ephemeris false alarm rate is high and the detection capability of a satellite-borne sensor on a dark and weak target is insufficient are solved, and the accuracy of evasion decision making and the execution reliability are improved.

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 specifically discloses a relative orbit dynamics multi-dimensional ground simulation method using a ducted fan, and relates to the technical field of spacecraft ground simulation. According to the method, a ducted fan compensation platform is designed, the ducted fan compensation platform comprises a ducted fan, a main structure frame and a fan driving unit, the ducted fan compensation platform is integrally fixed to the air floatation simulator, and the method comprises the specific steps that a target spacecraft reference coordinate system and a simulator body coordinate system are established, and a space-time scaling model is determined; a simulator pose is obtained in real time through a pose measurement system; starting a jet system to generate relative displacement, and inversely calculating an actual space relative distance and a nonlinear acceleration by combining a C-W equation and a space-time scaling model; the ducted fan is controlled to provide xoy plane nonlinear compensation force, and z-axis direction compensation force is calculated through mathematical simulation coupling. The problem that a traditional air floating platform cannot simulate the dynamic characteristics and multi-dimensional motion of a spacecraft orbit is solved, and the precision of a spacecraft ground simulation test is remarkably improved.

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

The application discloses a satellite orbit transfer strategy planning method and device, equipment and a storage medium. The method comprises the following steps: initializing an initial time of a satellite and performing orbit period extrapolation to update a t time to an extrapolated time; initializing an orbit transfer flag as not completed, calculating a first orbit parameter of the satellite at the t time and a second orbit parameter of a target virtual satellite at the t time, and determining a relative orbit parameter of the satellite to the target virtual satellite; determining each orbit parameter correction flag according to the relative orbit parameter and a preset orbit threshold value, and determining a corresponding orbit control sub-mode based on the each orbit parameter correction flag; when the orbit transfer flag is completed, determining a firing parameter of the satellite based on the orbit control sub-mode, and performing orbit transfer maneuver and orbit transfer phase adjustment processing. The scheme can quickly select the corresponding orbit control sub-mode according to the current orbit transfer requirement, quickly generate the calculation and simulation verification of the electric propulsion orbit transfer parameter, and make the satellite accurately enter the predetermined orbit position.

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

The invention provides an angle-measurement-only multi-pulse self-adaptive closed-loop rendezvous guidance method based on second-order cone programming, and belongs to the relative navigation and rendezvous guidance technology of spacecrafts. According to the method, firstly, an angle-only-measurement relative navigation dynamical model and a measurement model based on a CW equation are constructed under LVLH coordinates, system state observability analysis is carried out, the orbit maneuvering condition enabling the state to be observable is obtained, and the condition is used for constructing an optimal guidance index. Secondly, analyzing and modeling the rendezvous guidance error source, and deducing to obtain a relative orbit guidance error covariance model through error covariance analysis; thirdly, on the basis of a second-order cone programming method, guidance fuel consumption and system state observability conditions are fused, and a multi-target weighted cost function of optimal rendezvous guidance is jointly constructed; meanwhile, based on the established guidance error covariance model, a cost function weight index system with a self-adaptive adjustment function is designed, and real-time self-adaptive adjustment of a guidance strategy along with change of system performance is achieved. The method solves the problems of high fuel consumption, poor flexibility and the like caused by separate design of the relative navigation and the rendezvous guidance model in the traditional only-sight-line measurement rendezvous guidance method, and has the ability of adaptively adjusting and optimizing the rendezvous fuel consumption and the relative navigation system observability.

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 application discloses a spacecraft formation configuration decoupling control method and belongs to the technical field of satellite formation control. The application aims at solving the problems of serious approximation error accumulation and high fuel consumption of a spacecraft formation control method based on a CW equation. The method comprises the following steps: calculating relative orbit elements of a sub-satellite and a main satellite, obtaining a radial relative motion equation, a trace relative motion equation and a normal relative motion equation of the sub-satellite and the main satellite, and obtaining a modified equation; determining a velocity pulse according to a dynamic equation between the relative orbit elements and the velocity pulse; determining a required change amount of the relative orbit elements, and then determining a required change amount of a relative orbit inclination vector, a required change amount of a relative semi-major axis, a required change amount of a relative eccentricity vector and a required change amount of a relative angular distance; controlling the relative orbit inclination vector by adopting a single-pulse normal control strategy; controlling the relative semi-major axis and the relative eccentricity vector by adopting a double-pulse trace pulse strategy; and controlling the relative angular distance by adopting a multi-period trace control strategy. The application is used for formation control.

This invention discloses an autonomous approach guidance and control method, electronic equipment, and medium for spacecraft. It determines the approach control sequence and dwell time based on the spacecraft's relative orbit to the target, nominal approach position, attitude maneuver allowance, and nominal guidance transfer time. Then, it corrects the guidance transfer time and adjusts the approach control sequence accordingly. Next, it determines the dwell control sequence based on the spacecraft's relative orbit to the target, upper and lower limits of the nominal dwell time, and lateral acceleration. Finally, it determines the lateral control strategy, normal control strategy, and radial control strategy based on the spacecraft's relative orbit to the target and attitude maneuver allowance. This method combines fixed-time CW guidance with independent lateral, normal, and radial three-axis guidance, achieving high-precision approach control within a given time range.

The application relates to the technical field of satellite autonomous control, and discloses a satellite round game pursuit-escape game decision method based on Alpha-Zero, a terminal and a medium. The method sets physical constants and game rules, establishes a relative orbit motion model of satellites of both sides of pursuit and escape based on a C-W equation, and thus constructs a satellite round game pursuit-escape game environment; a strategy value neural network based on an Alpha-Zero framework is constructed, and when impulse orbit change decision needs to be made in each round, Monte Carlo tree search is performed based on the satellite round game pursuit-escape game environment and the strategy value neural network; the parameters of the strategy value neural network are iteratively optimized through self-game iteration, and an optimal strategy value neural network is obtained; after the real-time collected satellite orbit state is input into the optimal network after feature coding, an impulse orbit change strategy is generated and the satellite is controlled to execute. The application realizes efficient autonomous game of the satellite in a complex space environment.

The invention relates to a multi-low-orbit satellite cooperative communication method based on spectrum sensing assistance, which comprises the following steps: in a multi-low-orbit satellite cooperative system, a satellite with a relatively low relative orbit height senses a satellite with a relatively high relative orbit height by applying a spectrum sensing technology, and satellite-ground communication is carried out based on spectrum sensing assistance. The method specifically comprises the following steps: embedding signal cyclic delay characteristics at a sending end by adopting a cyclic delay diversity technology, and performing corresponding processing at a receiving end. According to the invention, through real-time efficient detection and utilization of the spectrum hole, the satellite-ground communication link quality and the transmission performance are effectively improved.

The invention discloses a spacecraft formation configuration decoupling control method, and belongs to the technical field of satellite formation control. The invention aims to solve the problems of serious approximate error accumulation and high fuel consumption of a spacecraft formation control method based on a CW equation. Comprising the following steps: calculating a relative orbital element of a sub-satellite and a main satellite, obtaining a radial, trace and normal relative motion equation of the sub-satellite and the main satellite, and obtaining a corrected equation; determining a speed pulse according to a kinetic equation between the relative orbital element and the speed pulse; determining the variable quantity required by the number of relative orbits, and further determining the variable quantity required by a relative orbit inclination angle vector, the variable quantity required by a relative semi-major axis, the variable quantity required by a relative eccentricity rate vector and the variable quantity required by a relative angular distance; adopting a monopulse normal control strategy to control the relative orbit inclination angle vector; adopting a double-pulse trace pulse strategy to control the relative semi-major axis and the relative eccentricity rate vector; and controlling the relative angular distance by adopting a multi-period trace direction control strategy. The method is used for formation control.

The invention discloses a spacecraft autonomous approaching guidance control method, electronic equipment and a medium, and the method comprises the steps: determining an approaching control sequence and a residence moment according to a relative orbit, an approaching nominal position, attitude maneuver reserved time and nominal guidance transfer time of a current spacecraft relative to a target, and then correcting guidance transfer time; correcting the approaching control sequence according to the corrected guidance transfer time, and then determining a residence control sequence according to the relative orbit of the current spacecraft relative to the target, the nominal residence time search upper and lower limits and the transverse acceleration; and finally, determining a transverse control strategy, a normal control strategy and a radial control strategy according to the relative orbit of the current spacecraft relative to the target and the attitude maneuver reserved time. According to the method, CW guidance at fixed time and transverse normal radial three-axis independent guidance are combined, and high-precision approaching control within a given time range is achieved.

The invention relates to the technical field of space target tracking, in particular to a space adaptive filtering positioning method and device based on a hybrid network model, and the method comprises the steps: determining an initial relative orbit state based on angle measurement observation data; iteratively executing a loop: inputting to a pre-trained CNN-LSTM hybrid network model, and outputting an adaptive state transition matrix and an adaptive model error factor at a next moment; state prediction is carried out, and prior estimation and a prior error covariance matrix of the relative orbit state at the next moment are obtained; updating is carried out, and the optimal estimation of the relative orbit state at the next moment and a posterior error covariance matrix are determined; outputting the optimal estimation of each time as a space target positioning result of the next moment; and subsequent Kalman filtering circulation is carried out. According to the method, the model is introduced, the adaptive parameters of Kalman filtering are dynamically predicted according to the current motion state, and the state estimation precision and the algorithm robustness in a maneuvering scene are remarkably improved.

This invention relates to the field of spacecraft orbit control and space safety technology, and discloses a method for space debris orbit prediction and avoidance. The method includes: a ground-based computing center screening candidate targets and generating enhanced data packets; an onboard computer performing geometric feature screening based on real-time navigation data to generate an immediate list of high-risk targets; controlling star sensors to perform optical observations, and using shift superposition and Kalman filtering to calculate the relative orbital state of the optical measurements; selecting a high-confidence or conservative probability threshold based on the acquired optical measurement state, and generating an avoidance command when the collision probability exceeds the limit; the propulsion system responding to the command to execute the avoidance, verifying the effect by confirming the change in the semi-major axis, and performing orbit restoration as appropriate. This invention employs ground-space collaborative screening and shift superposition enhancement technology, solving the problems of high false alarm rate due to low-precision ephemeris and insufficient detection capability of onboard sensors for faint targets, thus improving the accuracy of avoidance decisions and the reliability of execution.

The invention discloses a micro-nano satellite multi-pulse maneuvering method considering thrust constraint and J2 perturbation. The method aims at solving the problems that single speed increment implementation of the optimal maneuvering section of micro thrusters such as micro-nano satellite electric thrust / cold air and the like is limited, the maneuvering precision is reduced due to the fact that a long-term maneuvering orbit is influenced by J2 perturbation, and consequently the multi-satellite formation capture and reconstruction task of a near-circular orbit is difficult to complete. According to the method, the mapping relation between pulses and relative orbit parameters is established based on the relative orbit elements, a supplementary allocation method is provided for thrust constraint to achieve maneuvering pulse splitting, and a pulse sequence is corrected through iteration to compensate for J2 perturbation errors. According to the method, the multi-pulse maneuvering trajectory planning precision under the J2 perturbation influence can be effectively improved, the centimeter-level theoretical control precision can be achieved through 2-3 times of iterative computation, and the method has important practical value in the aerospace field.

The application discloses a kind of trace underactuated satellite incomplete information pursuit game control method, it belongs to satellite close-range orbit pursuit game technical field.The application solves the problem that three-dimensional pursuit game control of trace underactuated satellite cannot be realized using existing method.The application first derives the dynamics constraint imposed on tracker satellite and escape satellite by trace thrust loss through underactuated nonlinear and linear relative orbit dynamics, secondly derives trace underactuated three-dimensional pursuit game strategy under complete information using differential game theory, finally proposes an online calculation method based on differential Riccati equation control parameter, and derives trace underactuated pursuit game control strategy under incomplete information based on the same.The method of the application can be applied to trace underactuated satellite pursuit game control under incomplete information.

The invention relates to the technical field of satellite autonomous control, and discloses a satellite turn system pursuit game decision-making method based on Alpha-Zero, a terminal and a medium. According to the method, a physical constant and a game rule are set, and a relative orbital motion model of satellites of pursuit and escape parties is established based on a C-W equation, so that a satellite round system pursuit and escape game environment is constructed; constructing a strategy value neural network based on an Alpha-Zero framework, and executing Monte Carlo tree search based on a satellite round system pursuit game environment and the strategy value neural network when a pulse orbital transfer decision is needed in each round; parameters of the strategy value neural network are optimized through self-chess-playing iteration, and an optimal strategy value neural network is obtained; the satellite orbit state collected in real time is subjected to feature coding and then input into the optimal network, a pulse orbit transfer strategy is generated, and a satellite is controlled to execute. According to the invention, efficient autonomous gaming of satellites in a complex space environment is realized.

The application provides a non-cooperative spacecraft maneuver parameter quantitative identification method and system, and belongs to the technical field of aerospace. The non-cooperative spacecraft maneuver parameter quantitative identification method comprises the following steps: based on spacecraft relative orbit dynamics, an unknown impulse maneuver time parameter optimization model of a non-cooperative spacecraft is established; the motion state before and after one impulse maneuver of the non-cooperative spacecraft is acquired, and an optimization algorithm is used to solve the unknown impulse maneuver time parameter of the non-cooperative spacecraft; the motion state before and after the impulse maneuver of the maneuver point is calculated according to the maneuver time parameter and the optimization model, and finally the maneuver position and velocity pulse increment of the maneuver point are calculated. The application realizes the identification of the unknown impulse maneuver parameter of the non-cooperative spacecraft, is an important link for researching space situation information, has important theoretical research value, and can quantitatively analyze the maneuver parameter of the non-cooperative target.