20 results about "Transfer orbit" patented technology

The invention discloses a construction method and system for a low-thrust transfer orbit family between earth-moon translation point orbits, and belongs to the technical field of aerospace. The method comprises the following steps: firstly, solving a time optimal transfer orbit based on a continuous low-thrust spacecraft motion equation; then taking the orbit as a seed, performing forward and reverse phase continuation along a departure orbit and an arrival orbit by adopting a self-adaptive step length numerical continuation method, and generating a time optimal transfer orbit family covering different phases; secondly, mapping the time optimal orbit family into an energy optimal orbit family based on control law consistency, and expanding the transfer time range of the energy optimal orbit family through time continuation; and finally, gradually reducing homotopy parameters by using a homotopy method, detecting and processing thrust switching discontinuity in combination with a switching function, and converting the energy optimal orbit family into a fuel optimal transfer orbit family. According to the method, a continuous and complete orbit family is efficiently generated from a single solution, and systematicness and engineering applicability of orbit design are improved.

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 discloses a spacecraft approaching rendezvous guidance method based on limited thrust, and belongs to the technical field of spacecraft guidance. The invention aims to solve the problem that fuel consumption and hardware cost cannot be taken into consideration at the same time in an existing spacecraft short-distance guide section autonomous approaching method. Comprising the steps of calculating a theoretical speed increment of a current task stage; in the execution process of the task stage, the state of the propeller is judged according to a preset period; if the propeller is in the propelling process, the remaining speed increment is calculated by combining the theoretical speed increment and the single-cycle speed increment of the propeller; determining that the propeller is in a non-propelling state until the current residual speed increment is smaller than a set threshold value; the theoretical speed increment is calculated again and compared with a track safety belt threshold value to judge whether the tracking spacecraft exceeds the safety belt range or not, and if yes, the task state is updated, and the propeller is controlled to conduct propelling so as to correct the transfer orbit until the task stage is finished. According to the invention, high-precision guidance control is realized.

The application discloses a transfer orbit design method and system of a lunar orbit large ellipse frozen orbit based on eccentricity vector targeting. The transfer orbit lifting type is determined in advance in a transfer stage, the transfer orbit is selected as a lifting orbit, launching and arriving are lifted, the lunar orbit is ensured as a large ellipse frozen orbit with a far point above the south hemisphere of the moon, the eccentricity vector targeting is used to generate a transfer orbit meeting task constraints, the near point amplitude angle is adjusted near the far point after the moon is captured, the lunar orbit large ellipse frozen orbit is efficiently generated, and finally, a pulse is applied at the near point according to task requirements to transfer to a large ellipse frozen orbit with a required period.

The invention discloses an earth-moon L1 point track crossing behavior prediction method and system based on machine learning. According to the method, firstly, an initial state set is generated based on an earth-moon plane circular restrictive three-body model, and an orbit sample data set is constructed through numerical integration; secondly, a multi-class behavior classification system is established, and sample data is automatically labeled; on the basis, an XGBoost algorithm is adopted for model training, a category self-adaptive threshold optimization strategy based on a verification set is introduced for the problem of data extreme imbalance, and the overall prediction precision and category fairness of the model are remarkably improved by independently optimizing an optimal decision threshold for each category and applying the optimal decision threshold to normalized probabilistic decision. According to the method, the technical blank of spacecraft behavior prediction in the earth-moon translation point area is filled, and an effective means is provided for space situation awareness.

The application provides a spacecraft thruster fault response method, comprising: transferring a small geostationary orbit satellite from an earth transfer orbit to a geostationary orbit through three stages; the small geostationary orbit satellite comprises an orbit transfer chemical thruster, an attitude control chemical thruster and an electric thruster, wherein, in the second stage of the transfer of the small geostationary orbit satellite, a fault response scheme is started according to a fault response strategy of "first electric propulsion, then orbit transfer chemical thruster, and then attitude control chemical thruster", and the fault response scheme comprises: when the electric thruster is faulty, the orbit transfer chemical thruster is started; when both the electric thruster and the orbit transfer chemical thruster are faulty, the attitude control chemical thruster is started.

The invention particularly relates to a method and a system for estimating the minimum velocity increment of multi-circle double-pulse orbit transfer under J2 perturbation. The method comprises the following steps: uniformly representing origin-destination orbit states as six spring equinox elements taking flat longitude as a phase quantity; aiming at a transfer structure of a first pulse, middle drift and a second pulse, determining a limited transfer turn number search window by utilizing a flat longitude drift characteristic and a pulse capability; a closed equality constraint system with the total variable quantity = the first pulse variable quantity + the middle-section drift variable quantity + the second pulse variable quantity is constructed, and it is ensured that the track state is strictly closed; pulse is decomposed into tangential, normal and radial components for cooperative solution, and a phase closed loop is realized by updating a semi-diameter increment; and performing low-dimensional nonlinear optimization in a limited number of turns to quickly obtain minimum speed increment estimation. According to the method, the calculation efficiency is remarkably improved while the precision is guaranteed, and the method is suitable for large-scale sequence optimization in the near-earth orbit multi-target access task.

The invention discloses a real-time two-dimensional calculation method for a ground fire transfer orbit based on a variable step size method, and the method mainly comprises the following steps: 1, setting flight time, and obtaining gravitational constants of the earth, the Mars and the sun; 2, initializing the system; 3, judging a time stage; 4, screening time steps; 5, calculating the motion trail of the spacecraft based on the Newton's law of motion and the universal gravitation law; 6, solving the position and speed of the spacecraft through a numerical integration method; 7, updating the mass change of the spacecraft in real time; and 8, drawing a spacecraft orbit and outputting related data. The method can be used for simulating the orbit behavior of the spacecraft between the earth and the Mars, especially under the condition that a nuclear power propulsion system is considered. According to the method, the motion trail of the spacecraft in the complex gravitational field can be accurately calculated based on classical mechanics and numerical simulation technologies, and theoretical support is provided for design and optimization of aerospace tasks.

The invention discloses a lunar exploration mission rocket trajectory and spaceship orbit combined reconstruction method, which comprises the following steps: S1, after a thrust drop fault occurs in a three-level flight section of a carrier rocket, judging whether to continue to execute a lunar flight mission or not; s2, if the execution is continued, determining an initial equivalent departure moment and an initial transfer orbit parameter under the current fault condition based on an orbit perigee equivalent conversion method; s3, obtaining a new transfer orbit parameter satisfying the matching of the lunar orbit constraint and the earth-moon transfer time; s4, solving to obtain a reconstructed trajectory of the rocket and a corresponding actual orbit injection point state and an actual orbit injection moment; and S5, calculating the transfer time from the orbit perigee corresponding to the actual orbit injection point state to the actual orbit injection point. According to the method, after the thrust drop fault of the carrier rocket occurs, the rocket flight trajectory and the airship earth-moon transfer orbit can be quickly and synergistically reconstructed online, so that the airship can still enter the earth-moon orbit according to task requirements under the fault condition.

The present disclosure provides a spacecraft orbit transfer control method, device and storage medium suitable for deep space ferry mission, belonging to the field of spacecraft guidance and design. The method comprises: obtaining a dynamic model of a first spacecraft transferring from a first orbit to a second orbit; based on the dynamic model, an initial transfer orbit is obtained through a first homotopy parameter with energy optimization as the target; based on the initial transfer orbit, a target transfer orbit is obtained through a second homotopy parameter and a third homotopy parameter with fuel optimization as the target. The present disclosure provides an initial value guess for the fuel optimization problem by using the easily solvable energy optimization problem through the first homotopy parameter; through the synergistic effect of the second homotopy parameter and the third homotopy parameter, the target transfer orbit under the fuel optimization target is obtained, thereby reducing the sensitivity of the indirect method to the initial value and enhancing the convergence of the solution of the complex orbit transfer task.

Embodiments of the present application disclose a transfer orbit design method from GTO to the Sun-Earth collinear libration point, a device and a medium; the method comprises: calculating and obtaining an invariant manifold connected with the Halo orbit around the libration point according to the spacecraft dynamics equation and the potential function in the Sun-Earth barycenter coordinate system; taking the geosynchronous transfer orbit as the earth parking orbit to construct the constraint variable, taking the Halo orbit as the target orbit to construct the control variable, and correcting the control variable by using the least square differential correction method; based on the set constraint condition and performance index, the optimal HOI point corresponding to the optimal performance transfer orbit is selected from the transfer orbit meeting the constraint condition; according to the optimal HOI point, the relationship between the control variable and the parking orbit constraint variable is analyzed, and the differential correction initial value expression for determining the control variable initial value of the multi-constraint orbit design is constructed; the transfer orbit meeting multiple constraints is obtained by using the multi-stage differential correction method according to the differential correction initial value expression.

The invention discloses a transfer orbit design method in on-orbit service of a synchronous orbit satellite. The problem of how to effectively image a high-orbit target satellite with low fuel consumption is solved. Belongs to the technical field of space. Comprising the following steps: determining total transfer time according to different initial states of a service satellite and a target satellite; constructing a transfer orbit mathematical model based on the total transfer time to obtain the running time of the service satellite on the initial orbit, the transfer orbit and the approaching orbit in the transfer process and the speed increment for completing the Horman transfer; taking a relative position vector included angle and an illumination angle of two stars at the end of intersection as an objective function, taking total transfer time as a constraint condition, taking the operation time of a service star in an initial orbit and an approaching orbit as an optimization variable, optimizing the objective function by using a genetic algorithm based on an HPOP model, and correcting parameters of a transfer orbit mathematical model to obtain a transfer orbit mathematical model; obtaining a transfer orbit mathematical optimization model; according to the method, effective imaging of the high-orbit target with low burnup is realized through the mathematical optimization model of the transfer orbit.

The application discloses a transfer orbit electric propulsion satellite thrust drop detection method and system, comprising: establishing a single electric thruster real-time thrust-force variation relation model according to satellite data; calculating a satellite force tolerance range according to the single electric thruster real-time thrust-force variation relation model based on the single electric thruster real-time thrust-force variation relation model; if the electric thruster has a fault, determining the installation position of the faulty electric thruster first; according to the force size at different time within the satellite working time range and the installation position of the faulty electric thruster, determining the fault mode of the faulty electric thruster by using the fault mode function of the faulty electric thruster, and generating a faulty thruster thrust curve. The application can effectively track any form of thrust variation, lock the faulty thruster and determine the fault mode, so that timely correction measures are taken to ensure that the orbit and attitude of the satellite are kept in the predetermined target range.

The application discloses a spacecraft rendezvous guidance method based on limited thrust, and belongs to the technical field of spacecraft guidance. The application aims at the problem that the existing autonomous approach method of a spacecraft in a close distance guidance stage cannot simultaneously consider fuel consumption and hardware cost. The method comprises the following steps: calculating a theoretical speed increment of a current task stage; judging a state of a thruster according to a preset period during the execution of the task stage; if the thruster is in a propelling state, calculating a residual speed increment by combining the theoretical speed increment and a single-period speed increment of the thruster; until the current residual speed increment is less than a set threshold value, determining that the thruster is in a non-propelling state; then recalculating the theoretical speed increment, comparing the theoretical speed increment with a track safety belt threshold value, and judging whether a tracking spacecraft exceeds a safety belt range; if yes, updating a task state, and controlling the thruster to perform propelling, so as to correct a transfer orbit, until the task stage is ended. The application realizes high-precision guidance control.

The application provides a spacecraft on-orbit service large-scale task scheduling method and related equipment, comprising: establishing an orbit dynamics equation of a single service spacecraft and a single target satellite to determine the out-of-plane transfer phase and the coplanar transfer phase of the service spacecraft to the target satellite; taking the time and fuel consumption of the service spacecraft as performance indicators to establish a cost function; sampling multiple phase angles of the target satellite and performing nonlinear mapping on the orbit characteristics extracted based on the cost function to obtain a comprehensive cost, a time cost and an optimal transfer orbit; based on the Monte Carlo tree search method, the service spacecraft is assigned to each target satellite, and the optimal transfer orbit is selected through the established large-scale task scheduling model to schedule each service spacecraft to the corresponding target satellite; the exploration of the task state space in the process of the spacecraft on-orbit service large-scale task scheduling is avoided to fall into a local optimum, and the speed and accuracy of the large-scale on-orbit service task scheduling are improved.

The present invention relates to a propulsion system adapted for the transfer of an elliptic or circular transfer orbit to a GEO, MEO and / or HEO of a rideshare telecommunication satellite (1), the system comprising: at least two orbit raising thrusters (20), preferentially at least three 5 orbit raising thrusters, accommodated in a central portion (11) of a zenith face (10) of the satellite (1) opposed to the Earth when the satellite operates in transmission, at least three station keeping thrusters (30), preferentially at least four station keeping thrusters, each station keeping thruster being accommodated in a corner portion (12) of the zenith face (10).

The invention provides an orbital combination-based Lanbert transfer singularity avoidance method, which belongs to the technical field of spacecraft orbital design and orbital control, and comprises the following steps of: acquiring an initial moment position vector, a terminal moment position vector and flight time between two points; solving a Lanbert problem of an initial state to obtain a theoretical pulse speed increment instruction; calculating a normal unit vector of a transfer orbit plane at an initial moment; a control period is set, and track control calculation is carried out according to the following steps for each control moment from the initial moment; solving a normal unit vector of a transfer orbit plane at the current moment; calculating singularity measurement based on the center transfer angle; calculating an orbit combination weight coefficient according to the singular measurement; calculating a combined normal vector according to the normal unit vectors of the two transfer orbit planes and the combined weight coefficient; and solving the Lanbert problem by using the combined normal vector. According to the invention, whole-course smooth non-singular control is realized.

The invention discloses a finite thrust maneuvering planning method for coplanar multi-satellite formation and configuration adjustment. The method comprises the following steps: establishing a configuration adjustment control target decoupling description model for four-satellite formation; according to the four-satellite formation-oriented configuration adjustment control target decoupling description model, establishing a multi-circle transfer strategy oriented to coplanar multi-satellite formation configuration adjustment, and obtaining an orbit eccentricity rate adjustment amount, an orbit perigee angular distance total adjustment amount and a multi-circle transfer orbit maneuvering arc section control point; and according to the multi-circle transfer strategy for coplanar multi-satellite formation configuration adjustment, establishing a coplanar multi-satellite formation configuration adjustment finite thrust maneuver solving method. According to the invention, a flexible and variable coplanar satellite formation design is constructed, a rapid planning method is provided for a limited thrust trajectory maneuvering strategy of coplanar multi-satellite formation configuration adjustment, and the problems of flexible adjustment of the formation configuration and energy loss of on-satellite autonomous maintenance and control are solved; and a basis is provided for engineering implementation of a distributed formation system meeting multi-task requirements.

The application discloses a transfer orbit fast optimization method based on spatial conic curve splicing, and belongs to the field of deep space exploration. The application expands the plane conic curve splicing method to the optimization of spatial earth-moon transfer orbit on the basis of the plane conic curve splicing method used for orbit transfer, by using coordinate conversion and geometric analysis method to increase the azimuth angle parameters theta and psi used for describing the spatial orientation of the earth-centered launch orbit plane; the application can improve the orbit analysis and optimization efficiency by using the two-body orbit theory to deduce the two-body model and based on the double two-body model splicing with analytical characteristics; the application can realize the fast optimization and solution of the transfer orbit based on the spatial conic curve splicing by establishing the penalty function iterative optimization of four independent optimization quantities, i.e., the earth-moon transfer launch speed increment, the moon influence ball entry point phase, the azimuth angle parameters theta and psi, according to the orbit inclination and the perigee height constraints. The application has the advantages of good convergence, high optimization efficiency and wide application range. The application is suitable for various earth-moon transfer missions and moon flyby orbit optimization.

The application relates to a DRO orbit aircraft orbit design method for falling on the moon at a given angle and coordinates. The method comprises the following steps: DRO initial position and speed value integration, output of L2 coordinate system positions of the moon at different moments in a period, and mapping of the moon falling speed and transfer orbit and white plane intersection position. The mapping and DRO period position are combined to reversely coarsely calculate the moon falling moment, initial speed and transfer time length. The three are taken as iteration initial values and constraints are set to reversely accurately calculate the integral end moment, speed and orbit flight time length. The end moment is taken as the orbit transfer moment initial value, the end speed and the DRO speed difference value are taken as the orbit transfer speed increment initial value, the orbit flight time length is taken as the transfer time length initial value, the forward accurate calculation is carried out by setting the terminal constraint and the integral termination condition, and finally the transfer orbit transfer moment and speed increment projection are obtained. The method can be suitable for different task times and different DRO orbit sizes, the calculation result is high in precision, and the iteration is fast and reliable.