79 results about "Orbit/Orbital" patented technology

The invention relates to an earth-moon space low-thrust spacecraft fuel optimal collision avoidance trajectory implementation method, which comprises the following steps of: constructing a kinetic model based on calculation parameters of a spacecraft collision avoidance problem, and establishing a fuel optimal performance index of the spacecraft collision avoidance problem through a homotopy method; the optimal control problem is converted into a two-point boundary value problem, targeting solving is carried out, and the state and co-state parameters of the optimal orbit are obtained; and according to the obtained state and co-state parameters of the optimal orbit, solving to obtain collision avoidance optimal control and collision probability of the spacecraft. According to the method, the fuel optimal control problem is solved by adopting the indirect method, the convergence difficulty of the low-thrust fuel optimal problem is reduced through the homotopy method, and when the thrust is calculated, the calculation efficiency is high, the precision is high, and the method is suitable for the fuel optimal collision avoidance scene of the low-thrust spacecraft for a single target in the earth-moon space.

The application provides an orbit transfer strategy optimization method and system for a large-ellipse frozen orbit satellite orbiting, comprising the following steps: acquiring a satellite-rocket separation orbit and a station parameter, calculating an elevation angle of the satellite and the station, and then determining a range of a true anomaly angle of pulse ignition meeting a measurement and control requirement; in combination with the range of the true anomaly angle of pulse ignition, generating a fuel most-probable pulse orbit change strategy of the satellite from an initial orbit to a target orbit, determining an orbit change speed increment, an ignition position and an ignition frequency; determining ignition-on and ignition-off conditions of each orbit change and a drift circle number between adjacent two orbit changes; according to the pulse orbit change strategy, respectively determining an initial value of an ignition time length, an ignition time and a thrust direction of each orbit change, and determining the final ignition time length, the ignition time and the thrust direction of each orbit change. The application adopts a non-apogee double-pulse orbit change strategy to design a fuel optimal orbit transfer strategy of a large-ellipse frozen orbit satellite orbiting, and solves the design method problem of the orbit transfer strategy.

The invention discloses a strict regression frozen orbit design method. The method comprises the following steps: calculating an initial estimated value of a target strict regression frozen orbit mean element number based on an analytic formula of a regression orbit and a frozen orbit under a third-order gravity field; under the high-order gravity field dynamic model, iteratively correcting the orbit semi-major axis and the orbit inclination angle based on the right ascension change rate of the ascending node and the deviation value of the longitude of the revisit ascending node; taking the geocentric vector difference of the revisit ascending node as an optimization target, taking the eccentricity rate as an optimization variable, and adopting an optimization algorithm to obtain a current optimal frozen eccentricity rate; iteratively correcting the semi-major axis of the orbit and the inclination angle of the orbit, and optimizing the eccentricity ratio to obtain the orbital number of the strictly frozen regression orbit.

The application discloses a Beidou satellite maneuvering orbit recovery method, which comprises the following steps: S1, giving initial satellite state information; S2, detecting whether satellite orbit maneuvering occurs for each ephemeris based on current satellite state information; if yes, entering step S4; if no, entering step S3; S3, updating state parameter values of all satellites and returning to step S2; S4, processing the maneuvering orbit by adopting combined short-arc orbit determination and grouping filtering method, and realizing maneuvering orbit recovery. The combined short-arc orbit determination and grouping filtering method realizes quick recovery of the Beidou satellite maneuvering orbit, so as to guarantee the navigation, positioning and timing service performance of the Beidou system in the case that the Beidou satellite occurs orbit maneuvering. The method can not only effectively improve the orbit determination precision in the case that the satellite occurs orbit maneuvering, but also greatly improve the reliability of the global service capability of the Beidou No. 3 system.

The application discloses a kind of space target ultra-short arc initial orbit intelligent determination method, belong to aerospace technology field.The application is based on artificial intelligence technology to realize space target ultra-short arc initial orbit determination, there is no strict requirement to complex nonlinear problem, and can develop more complex perturbation condition and initial orbit determination under the operation law of multiple body, coverage is wider;The application generates multiple groups of target orbits randomly and calculates the observation and state quantity of each visible arc segment, constructs the sample data set with ultra-large capacity, which is helpful for neural network to learn the nonlinear relationship between observation and state quantity using rich samples;The application builds and trains long short-term memory network model, determines the optimal network model weight parameter by designing and adjusting hyperparameter, can obtain the fast prediction solution of unknown space target ultra-short arc initial orbit according to observation;The application can meet more space target ultra-short arc initial orbit determination tasks by considering different inputs and outputs.

The application relates to a design method of a moon global positioning constellation based on a lunar space DRO reference orbit, wherein the initial orbit state and the orbit period of any DRO reference orbit in a DRO reference orbit to be deployed are determined based on the initial phase of the DRO reference orbit; the initial orbit state of any satellite to be deployed on the initial phase of the DRO reference orbit is determined based on the orbit state corresponding to the target phase; the orbit period of the any satellite is determined based on the orbit period of the any DRO reference orbit; the target orbit state corresponding to the any satellite is obtained based on the orbit period of the any satellite and the initial orbit state of the any satellite; and the moon global positioning constellation under a four-body model is constructed based on the target orbit state corresponding to each satellite. The moon global positioning constellation constructed by the application has higher stability and higher moon surface coverage capacity compared with the existing near-moon constellation.

This invention discloses a method and apparatus for planning a timed, fixed-point landing and timed takeoff for lunar sampling missions. The method includes: determining the orbital root number and orbital insertion time of the orbit in the first control phase of the lunar sampling mission, as well as the nominal parameters for timed, fixed-point landing and timed takeoff. Based on the nominal parameters, orbital root number, and orbital insertion time, the first control phase is jointly planned with subsequent control phases to obtain a first deviation, and the first deviation is iterated based on a first variable. When the first deviation is less than a convergence threshold, the target variable for the first control phase is determined, allowing the satellite to achieve timed, fixed-point landing and timed takeoff based on the target variables of each control phase. This achieves dynamic joint planning with timed, fixed-point landing and timed takeoff as the final goal for each control phase, effectively improving the mission's control reconfiguration capability, avoiding additional control operations, and reducing resource consumption.

This invention relates to the field of collision warning technology, specifically to a spacecraft collision warning method for a large number of space targets. The method includes: selecting a multi-data parallel screening warning approach; screening secondary target orbits within the TLE data set of the complete collision target orbits using set pre-screening thresholds for date expiration, apogee / perigee pre-screening safety margins, orbital intersection pre-screening thresholds, and time pre-screening thresholds; performing SGP4 orbit prediction on the primary target orbit and the initially screened secondary target orbits; obtaining the minimum spatial distance and collision probability between spacecraft within the primary and secondary target orbits; summarizing the primary and secondary target orbit numbering information, as well as the minimum spatial distance, collision probability, and the time corresponding to the minimum spatial distance, into a collision warning result; and summarizing all collision warning results to output a collision warning set information. This invention can filter out targets that are unlikely to intersect with the spacecraft to the greatest extent possible, reducing the computational load of subsequent orbit prediction.

The application discloses an initialization method of a conic surface under a projection linearization track surface, and belongs to the technical field of calculation, reckoning or counting. The method first obtains a preliminary estimation of a target track through three groups of observation data; determines a conic surface containing an initial track surface according to initial track information and geometric knowledge; estimates the missing height information in combination with the conic surface, and obtains more position information of the target in combination with the existing angle information; projects the position information of the target into a two-dimensional section, re-estimates the track parameters by using a least square algorithm, and more accurate track parameters can be obtained; then the iteration process is repeated by using the newly obtained track information until a certain initial orbit determination accuracy requirement is met. The algorithm firstly fits the height information of the target by using a conic surface coordinate system, and realizes linear fitting of the position vector by using a track surface projection method, so that the calculation amount is small and the accuracy is high.

The invention provides a nonlinear degree evaluation method suitable for a Lyapunov orbit near an earth-moon space stable translation point, and belongs to the field of celestial body measurement and celestial body mechanics. Comprising the following steps: generating a plurality of sample points near an initial value of a specified reference orbit according to Gaussian distribution of an initial error; calculating a propagation result of each sample under a floating-point number framework at any time length, gradually increasing an order k, integrating a reference orbit under a DA framework to obtain a k-order Taylor expansion of an integral end moment state quantity relative to an initial moment state quantity, calculating a k-order approximation value of a sample point by using the Taylor expansion, further calculating a local non-linear index eta, and researching the change of the local non-linear index eta; and when eta is smaller than the acceptable tolerance, determining the non-linear degree of the simulation data as the corresponding k at the moment, and ending the search order calculation. The high-order uncertainty propagation method based on the high-order uncertainty propagation method has the advantages that the high-order uncertainty propagation method ensures that the precision is close to the uncertainty expression methods such as Monte Carlo, and meanwhile, the efficiency is greatly improved.

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 application discloses a satellite orbit space environment disturbance coping method, equipment and computer program product. The method comprises the following steps: acquiring current position data and current space environment data of a target orbit, wherein the current space environment data comprises a first high-energy particle integrated flux at a target position of a high-energy particle flux and a first high-altitude atmospheric daily average atmospheric density; querying a second high-energy particle integrated flux at the target position of the high-energy particle flux and a second high-altitude atmospheric daily average atmospheric density when the target orbit environment is calm according to the current position data; calculating a high-energy particle radiation environment relative disturbance amount and a high-altitude atmospheric density maximum relative disturbance amount of the target orbit; determining a current high-energy particle radiation environment disturbance degree and a current atmospheric environment disturbance degree in combination with a disturbance degree discrimination range; and generating a risk coping suggestion. The application can accurately understand the disturbance situation of a spacecraft operating environment and improve the safety of the spacecraft in-orbit operation.

The invention discloses a navigation mode transition method, an electronic device and a medium, in a working process of a first navigation mode, geometric angles among a mother satellite, a service spacecraft and space debris are monitored, and when the geometric angles meet a mode switching requirement, the position of a target orbit under a J2000 coordinate system is obtained through double-satellite geometric estimation; and taking the position as an initial position calculated in the second navigation mode, and switching to the second navigation mode. Before the navigation mode is switched, the target orbit position is determined through double-satellite geometric estimation, the single-satellite navigation positioning precision can be rapidly improved, a good initial value is provided for double-satellite angle measurement filtering, and the convergence process is accelerated.

The application discloses an online trajectory planning method for the final stage of a carrier rocket aiming at an elliptical target orbit and belongs to the field of carrier rocket trajectory and guidance, which comprises the following steps: step one, constructing the orbit entry terminal constraint and planning performance index aiming at the elliptical target orbit; step two, establishing the system dynamics constraint of the orbit entry of the final stage of the carrier rocket; step three, convex approximation is conducted on the planning orbit entry terminal constraint and performance index function; step four, state affine approximation and pseudo-spectrum discretization are conducted on the system dynamics constraint of step two; and step five, the optimal trajectory of the final stage entering the elliptical orbit is generated online through sequential convex iteration. The method of the application is suitable for the joint reconstruction of the final stage trajectory and the adaptive guidance of the carrier rocket launching the elliptical target orbit mission, can realize the online generation of the optimal trajectory with high precision and fast convergence, and is favorable for improving the reliability of the carrier rocket and enhancing the launch service fulfillment capability.

The application discloses a method, system, device, medium and product for judging the instability degree of an electric power rope system, and relates to the technical field of spacecraft attitude control. The method comprises the following steps: acquiring the orbit inclination of the deployment orbit of a main star in the electric power rope system; determining the instantaneous equilibrium point distribution area of the electric power rope system based on the orbit inclination; positioning the optimal balance area of the electric power rope system in the instantaneous equilibrium point distribution area of the electric power rope system; acquiring the attitude angle of the electric power rope system at the current moment, and determining the instability degree of the electric power rope system at the current moment based on the attitude angle of the electric power rope system at the current moment, the instantaneous equilibrium point distribution area and the optimal balance area. The application improves the efficiency and reliability of the instability evaluation of the electric power rope system.

The application provides a satellite orbit prediction method, system and device based on revisiting nominal orbit, relates to the technical field of satellite orbit prediction, and the prediction method comprises the following steps: acquiring on-satellite GPS telemetry data of a satellite; obtaining a descending node time and a descending node longitude of an actual orbit of the satellite within a preset time period according to the on-satellite GPS telemetry data; matching the descending node longitude of the actual orbit with a plurality of descending node longitudes of a revisiting nominal orbit stored in advance, selecting a descending node state corresponding to the descending node longitude of the revisiting nominal orbit with the minimum deviation from the descending node longitude of the actual orbit as a matching nominal orbit state; and inputting the descending node time of the actual orbit and the matching nominal orbit state into an improved high-precision orbit dynamics model for orbit recursion to obtain predicted orbit data. The application can realize long-time and high-precision orbit prediction, and significantly improves the reliability and precision of satellite mission planning.

This invention proposes a method and system for remote sensing satellite orbit design based on multi-parameter fuzzy constraints. The method includes: receiving input orbit boundary constraints and design constraints; calculating a series of information about the corresponding orbit based on the altitude range, pixel resolution, revisit constraints, and number of mosaic operations for the sun-synchronous return circular orbit; calculating relevant information that can be directly filtered by subsequent design constraints based on the candidate orbit parameters; filtering the directly filtered relevant information using imaging angle constraints, swath width constraints, and integration time constraints; and filtering the orbit using latitude constraints, marking results that meet and do not meet the constraints. This invention integrates and coordinates multi-domain knowledge related to orbit design and the fragmented steps of orbit design, significantly improving the overall efficiency of satellite design.

The embodiment of the invention relates to a method and device for determining molecular system energy, equipment and a storage medium. The method comprises: for each fragment in a molecular system, determining a plurality of orbits associated with the fragment, the plurality of orbits including a fragment orbit of the fragment, a first environmental orbit strongly coupled to the fragment, and a second environmental orbit weakly coupled to the fragment; based on the corresponding occupation states of the plurality of tracks, the plurality of tracks are respectively corresponding to the first track space or the second track space, and for the tracks corresponding to the same track space, the occupation state of the second environment track is different from the occupation state of the segment track and the occupation state of the first environment track; selecting a set of target orbits from the plurality of orbits to construct a fragment orbital space of the fragment using reference information related to electron excitation determined based on the first orbital space and the second orbital space; and determining an estimate of system energy of the molecular system based on the fragment orbital spaces of the plurality of fragments.

The invention discloses a low-thrust spacecraft asymmetric halo orbit design method, and belongs to the field of spacecraft orbits. The problem that the periodicity and stability of the orbit of a low-thrust spacecraft cannot be guaranteed under the condition that the plane of the halo orbit is asymmetric is solved. The method comprises the following steps: calculating a halo orbit satisfying a resonance relation with a second gravitational body orbit period; the eccentricity rate of the second gravitational body orbit is continued from zero to the actual eccentricity rate through numerical continuation in the first step, and a zero-thrust resonance halo orbit is obtained; through numerical continuation in the second step, the thrust acceleration amplitude is continued from zero to a preset upper limit while the thrust direction of the spacecraft is kept in the sun-spacecraft connecting line direction; through numerical continuation in the third step, the thrust taper angle is continued from zero to the preset upper limit under the gravity plane coordinate system; and through numerical continuation in the fourth step, the thrust azimuth angle is continued from zero to a preset value under the gravity plane coordinate system, and finally the asymmetric halo orbit is obtained. The method is suitable for circular and elliptical restrictive three-body problems.

This invention discloses an orbital jump processing method, electronic device, and computer program product. The orbital jump processing method includes: determining a first mean square error in fitting the current orbit of the satellite to a first orbit, and a second mean square error in fitting the current orbit to a second orbit; the first orbit is the orbit preceding the current orbit of the satellite, and the second orbit is a predicted orbit obtained by batch processing based on the current orbit; determining the value of the orbital smoothing weight matrix in a preset first observation equation based on the first mean square error and the second mean square error; the orbital smoothing weight matrix characterizes the degree of dependence on the current orbit or the first orbit during orbit fitting; and determining the orbital value used for orbit updating based on the first observation equation.

The application discloses a low-fuel-consumption low-orbit satellite orbit recovery method, and belongs to the technical field of satellite orbit recovery.The application aims at the problem that the existing orbit maneuver method needs frequent orbit adjustment, which is not conducive to the long-term operation of the satellite in orbit with low fuel consumption.The application comprises the following steps: establishing an orbit dynamics equation considering the earth oblateness perturbation; using a semi-analytical method to solve the initial orbit elements of a low-fuel-consumption transition orbit according to the current orbit and the target orbit of the satellite; simultaneously, combining the orbit dynamics equation, calculating the drift time required for the satellite to drift on the transition orbit by using the earth oblateness perturbation without fuel consumption and the final orbit elements of the transition orbit at the end of the transition; and based on the Pontryagin maximum principle, calculating the optimal control rate sequence of the thruster fuel of the satellite from the current orbit to the transition orbit and from the transition orbit to the target orbit, and controlling the thruster to realize the orbit recovery of the satellite.The application is used for satellite orbit recovery.

The invention discloses a low-orbit satellite orbit maintenance method and system under GNSS denial, and the method comprises the following steps: S1, enabling a satellite to enter a nominal orbit, collecting the atmospheric density data of the satellite in real time through an atmospheric density sensor, and calculating the windward side flight resistance of the satellite and the side flight resistance of the satellite according to aerodynamics; s2, reasonably selecting the thruster ignition time of the satellite according to the service working period of the satellite; s3, the resistance impulse is calculated according to the flight resistance of the satellite, and the working duration and the working thrust required by the thruster are calculated according to the resistance impulse of the satellite; and S4, realizing stable flight of the satellite in a low orbit through closed-loop feedback control of the satellite. The problem that the satellite autonomously completes orbit attenuation compensation without depending on an external navigation means is achieved, and the survivability of the satellite in a low orbit is enhanced.

The invention relates to a common ground trajectory giant constellation fly-over orbit optimization method, device and equipment, and relates to the technical field of spaceflight. The method comprises the steps of obtaining task initial parameters; establishing an orbit optimization model; traversing different fly-over initial moments of a first target satellite in the whole task to obtain a maneuver-free fly-over nominal orbit and a corresponding target satellite fly-over number; deriving a plurality of maneuver-free flyover orbits for accessing the whole constellation; establishing a multi-pulse transfer optimization model, and solving the optimal speed increment of each transfer type; mapping each maneuver-free flyover track into an access node, constructing a node transfer cost matrix according to the optimal speed increment, and solving an optimal access sequence for accessing all nodes; and integrating the parameters of a plurality of maneuver-free flyover orbits, the optimal access sequence and the maneuver scheme of each transfer type to form a most economical propellant flyover scheme for accessing the whole common ground trajectory giant constellation. According to the method, multiple target satellites can fly over through one-time maneuvering, and efficiency and fuel economy are both considered.

The invention provides a closed-circuit guidance method and device, electronic equipment, a storage medium and a program product, and relates to the technical field of aerospace, and the method comprises the following steps: determining a target position vector of a target point according to the latitude and longitude and height information of the target point; determining a current position vector and a current speed vector of the aircraft according to the current state data of the aircraft at the current moment; on the basis of the current position vector and the target position vector, constructing an orbit constraint for describing an orbit to fly over and a time constraint for describing flight time; constructing an optimization problem which takes norm minimization of a difference between a required velocity vector and a current velocity vector as an optimization target, and solving the optimization problem under orbit constraint and time constraint; and performing outer layer iteration of the optimization problem by updating the target flight time estimation, and performing inner layer iteration of the optimization problem by updating the speed inclination angle through a Newton iteration method until a preset convergence condition is met, thereby obtaining a final target speed inclination angle and final target flight time.

A method of determining a precise orbit of a satellite through estimation of a parallactic refraction scale factor is proposed, the method includes inputting an initial estimate including initial orbit information of a satellite with respect to an observation epoch and the parallactic refraction scale factor; performing orbit propagation using a high-precision orbit propagator by applying a dynamics model; performing observer-centered satellite optical observation modeling including the parallactic refraction scale factor; calculating an observation residual between actual optical observation data and observation data calculated via the observation modeling reflecting the parallactic refraction; and precisely determining the orbit of the satellite by estimating the parallactic refraction scale factor and a satellite state vector using a batch least square estimation algorithm.

The application relates to a method, device and equipment for designing initial values of a moon landing orbit based on a velocity surface parameter. The method comprises the following steps: determining the value range of each variable according to engineering constraints and constants; calculating a first velocity surface parameter equation meeting an apogee height constraint; calculating a second velocity surface parameter equation meeting a perigee height constraint, and solving the equations simultaneously to construct a third velocity surface parameter equation of a general moon orbit; performing traversal assignment on each variable according to the value range of the variable; substituting each assignment into the third velocity surface parameter equation to calculate a tangential velocity; correcting the tangential velocity meeting a tolerance to obtain a corrected tangential velocity; obtaining a corresponding variable initial value set based on the corrected tangential velocity; and then screening the variable initial values in the variable initial value set through engineering constraints to obtain variable initial values meeting the conditions. The application can improve the orbit design precision and reliability.

The application discloses a satellite predicted orbit error compensation method, and utilizes fast Fourier transform to analyze satellite predicted orbit error, identifies and extracts the periodic characteristics of the satellite predicted orbit error, establishes a periodic term model of the predicted orbit error, obtains residual error of the predicted orbit error, and then utilizes a pre-trained long short-term memory network to output a fitting residual error of prediction, finally, the periodic term model and the neural network model are used to respectively predict and compensate the orbit error, so that the precision of the satellite predicted orbit is improved. The application integrates error processing, frequency domain analysis and neural network technologies, slows down the attenuation trend of the satellite predicted orbit, and improves the precision of the predicted orbit.

The application discloses a method for quickly changing orbit of an elliptical orbit out-of-plane target to the same orbit with the least fuel control on a satellite, and a method for independently solving the least fuel control on a satellite for quickly changing orbit from an out-of-plane orbit to the same orbit, which comprises the following steps: judging whether to perform normal orbit plane correction or orbit shape correction according to orbit elements of an initial satellite and a target; solving a least fuel normal orbit plane correction control strategy according to a normal orbit plane element difference between the satellite and the target; solving a least fuel orbit shape correction control strategy according to an in-orbit plane element difference between the satellite and the target; and iteratively correcting the least fuel control strategy by using a high-precision orbit extrapolation model. The method is suitable for orbit changing of any elliptical orbit satellite in near-earth space to any elliptical orbit target, and the control precision can reach ten meters; and the calculation amount is small, and the method is suitable for solving a task planning on a satellite.

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.

The invention relates to a method and device for determining the pose of a track, a controller and a product. The method includes determining a degree of damage to the track based on an image including the track. The method further comprises the step of determining a target model related to the damage degree from a plurality of preset models used for determining the pose based on the damage degree. The method further includes determining a pose of the orbit based on the target model. Through the mode, the most suitable pose determination mode can be flexibly selected according to different damage degrees of the orbit, so that the orbit pose determination robustness and accuracy are improved.