93 results about "Orbit prediction" patented technology

The invention relates to a method and system for automatically tracking a medium earth orbit navigational satellite through a servo antenna, and belongs to the technical field of satellite navigation signal quality monitoring. The system comprises an information prediction unit, a servo control unit, a rotary table unit, the feed-forward large-diameter parabolic antenna and a spectrum analyzer. The information prediction unit is used for the operation of orbit prediction and angle calculation, the servo control unit is used for adjusting antenna postures in a closed loop feedback mode, the rotary table unit is used for driving an antenna rotary table to rotate in the azimuth axial direction and in the pitch axial direction, the feed-forward large-diameter parabolic antenna is used for receiving target satellite signals, and the spectrum analyzer is used for observing quality of received signals. The method and system for automatically tracking the medium earth orbit navigational satellite can achieve an ideal tracking effect in the state of being divorced from manual monitoring.

The invention relates to an embedded satellite laser ranging control system, and solves problems that the satellite laser ranging control technology implemented through a main control computer is complex in ranging method, is low response speed and causes the data loss. A telescope tracking controller reads a current observation moment, calculates the azimuth and height position at the current moment through orbit prediction interpolating of an observation station, and outputs the azimuth and height position to a servo system of a motor. A distance gate controller outputs an igniting signal of a laser according to a distance measuring frequency, locks a main wave signal, and calculates the moment when a distance gate signal corresponding to a main wave is generated. A gate control signal is outputted through the distance gate controller. A laser beam direction controller recognizes a light beam image, obtains a light peak position, calculates the deviation between the light peak point and the center of a receiving visual field, and corrects the direction deviation of a laser beam through a control algorithm. The system can meet the demands of higher-repetition-frequency laser ranging through employing an FPGA+DSP heterogeneous processor configuration, and facilitates the upgrading of the whole ranging system.

The invention discloses a system and method for dynamics simulation of a space robot. The method for dynamics simulation of the space robot comprises the steps that a space robot dynamics module outputs a value representing the motion state of the space robot in the remote guiding stage according to the control force and the control moment which are received by a space robot platform and the control moment of a mechanical arm system joint and after the returned control amount is received, a value representing the motion state of the space robot at the next moment is output; after a space robot control module receives the value which is input by the space robot dynamics module and represents the motion state of the space robot in the remote guiding stage, the control amount is returned to the space robot dynamics module by the space robot control module; an STK long-term orbit prediction module outputs the motion state of the space robot and the motion state of a target in the remote guiding stage according to a value representing the motion initial state of the target, the value representing the motion state of the space robot and the orbital transfer sequence of the space robot. According to the system and method for dynamics simulation of the space robot, complete simulation of the whole task stage of the space robot can be provided.

The invention discloses a mid-term orbit prediction method and device of a low-orbit spacecraft and a storage medium. The method includes the steps that target parameters are determined, wherein the target parameters at least comprise the atmospheric density parameter of the position where the target spacecraft is located, and the equivalent windward area and the atmospheric drag coefficient of the target spacecraft; and the mid-term orbit of the target spacecraft is predicted through the target parameters. The effect that the mid-term orbit prediction accuracy of the low-orbit spacecraft is improved is achieved through the method and device.

The embodiment of the invention provides an application-oriented heterogeneous constellation space-ground integrated task planning method and device, wherein the method comprises the steps of synthesizing an observation task demand; collecting the basic parameters and real-time state information of each satellite in the heterogeneous constellation; configuring the measurement and control and datatransmission resources of executable tasks; giving a time function of each satellite orbit position by utilizing accurate orbit prediction; sorting all the sub-tasks, processing conflicts among the sub-tasks, and judging whether a conflict relationship between the current to-be-added observation sub-task and each observation sub-task in the current observation sub-task set meets constraints or not; carrying out downlink scheduling; scheduling an observation data route by adopting a greedy algorithm; and generating a task planning scheme, estimating the duration of each observation and data transmission activity, the conversion time of satellite attitudes and the change of satellite related state parameters, and generating a task arrangement scheme containing the observation activities andthe data transmission activities.

The invention discloses a high-precision satellite-borne orbit prediction method based on numerical fitting. The method comprises the following steps: firstly, obtaining atmospheric environment parameters including the latest solar radiation flow and a geomagnetic index, satellite star parameters and the latest precise orbital elements of the satellite, combining the star and space environment parameters, carrying out orbit prediction by utilizing a ground high-precision numerical method so as to obtain a set of satellite instantaneous orbital elements in a prediction arc section, then, converting into an average element to perform numerical fitting so as to obtain a change item coefficient of each orbital element, arranging multiple sets of primary orbits to carry out segmented extrapolation according to an satellite-borne predication requirement, carrying out orbit prediction according to the arranged initial orbital element and the change item coefficient to obtain the average element of the satellite at the prediction moment, if the prediction precision requirement is high, arranging a plurality of sets of initial values for staged prediction, converting the average element ofthe satellite at the prediction moment into the instantaneous element, so that the instantaneous position of the satellite is obtained. According to the method in the invention, the problem that the precision of satellite-borne orbit prediction is poor in the prior art is solved.

The invention provides a method and a device for correcting the orientation of a three-axis antenna, and the method comprises the steps: obtaining an approximate value of an orientation angle set by an antenna under an equipment horizon system and an antenna guide angle under an equipment measurement system according to a target position, provided by orbit prediction, under the equipment horizon system through coordinate transformation; obtaining a corresponding pointing angle of the approximate value of the antenna guide angle under the equipment horizon system through correction calculation;and obtaining an accurate value of the antenna guiding angle set by the antenna under the equipment horizon system under the equipment measurement system through recursive calculation and by utilizing the corresponding pointing angle of the approximate value of the antenna guiding angle under the equipment horizon system. By utilizing the method, digital tracking or program tracking can be completed in real time and with high precision under the working condition of the high frequency band of the three-axis antenna.

The invention relates to a satellite accurate orbit prediction method. Deep neural network automatic learning training is carried out according to comparison results of a prediction ephemeris calculated through a satellite ephemeris and kinetic models and a real ephemeris to obtain a deep neural network compensator, and then the prediction of satellite accurate obits is obtained by adding the error compensation obtained through the deep neural network compensator and prediction orbit results obtained through the kinetic models. According to the satellite accurate orbit prediction method, the kinetic prediction models and deep neural networks are two independent parts, the structure is simple, the method is easy to achieve, compared with pure kinetic models or other neural network algorithms, prediction accuracy and control over prediction errors in long-term prediction are improved.

A multiple-constraint multi-spacecraft flight interval indicating and collision avoidance method includes the following steps: (1) confirming magnitude of attitude control average force according to telemeasuring data value of propellant consumed by attitude control of launched spacecrafts on an orbit under the cruising attitude, (2) conducting high-accuracy orbit prediction according to initial ephemeris information of each spacecraft and the attitude control average force confirmed in the step one, calculating ephemeris of each spacecraft in an inertial frame at any time and relative distance between each spacecraft at any time, and confirming the minimum relative distance between the spacecrafts, (3) changing the direction of the attitude control average force, repeating the step one and the step two, calculating the minimum value of the minimum relative distance between each spacecraft, namely, the direction of the attitude control average force corresponding to the minimum value is the worst situation, (4) comparing the minimum value confirmed in the step three with the minimum safe distance, when the minimum value is greater than the minimum safe distance, the spacecrafts has no risk of collision, otherwise, within a first circle observation and controlling track segmental arc, one spacecraft which is selected conducts orbital maneuver, and thereby the distance between the spacecrafts is widened and the collision risk of the spacecrafts is avoided.

The invention relates to a low-earth-orbit satellite long-term orbit prediction method based on a two line element. The method comprises the following steps: performing related processing on a two line element in an orbital coordiante system at a moment t, thus obtaining a satellite orbital element initial value applicable to an orbit prediction dynamical model in a J2000.0 geocentric celestial coordinate system; and solving the orbit prediction dynamical model by adopting a numerical method, thus obtaining a satellite orbital element prediction result, wherein the related processing comprisesreduction processing performed on angular velocity of peace movement in the two line element, conversion of a mean element into an instant element in the orbital coordinate system at the moment t andconversion of the instant element in the orbital coordinate system at the moment t into an instant element in the J2000.0 geocentric celestial coordinate system. The low-earth-orbit satellite long-term orbit prediction method based on the two line element is high in prediction precision.

The present invention provides a resistance coefficient adaptive modulation method based on precise ephemeris. Measurement data of a measurement station and a high-precision dynamical model are combined to optimize and correct a resistance modulation coefficient of TLE data, correlation of equal-height satellite correction is analyzed in a space environmental change condition, correlation of the TLE data of the equal-height satellite in the same day and the precise ephemeris is employed to design an orbit prediction algorithm for correction of a TLE+SGP4 model when the satellite is difficult to obtain enough outer measurement data and improve the TLE short-term prediction precision in a space environmental dramatic change period. The resistance coefficient adaptive modulation method basedon precise ephemeris can achieve adaptive modulation of abnormal satellite resistance coefficients, is high in adaptability and improves the prediction precision.

The invention provides a modeling method of a space debris environment long-term evolution model. The current space debris group is used as initial input; the current space debris group is subjected to orbit prediction by a simplified semi-analysis method; and further space object emission, on-orbit object collision disassembly analysis, orbit discarding analysis and active debris clearing are considered, so that the space debris environment long-term evolution result is obtained. The evolution model can describe the space debris environment evolution condition at a random moment in next two hundred years at different emission levels by using different debris retarding and clearing measures, so that strategy analysis and model support can be provided for China for active debris clearing in further.

The invention discloses a mars probe onboard autonomous orbit calculation method, which is a calculation method based on radial basis function neural network curve fitting. The method includes the following steps: building a three-layer radial basis function network model; getting orbit prediction data of a mars probe within a period of time according to ground orbit determination, and training the radial basis function network model by taking the data as a training sample; and finally, uploading the trained radial basis function neural network model as an onboard orbit prediction model to the mars probe. There is neither need for establishment of a complex dynamic model on a satellite nor need for ephemeris calculation. The prediction precision is almost equal to the ground orbit prediction precision. The method is also suitable for engineering implementation. The method not only can satisfy the engineering precision constraints of the recursion result, but also can satisfy the constraints of limited resources of onboard computers.

The invention discloses a formation satellite relative orbit determination method facilitating satellite-borne on-orbit real-time processing and belongs to the field of satellite formation relative navigation. According to the method, a GPS/BDS broadcast ephemeris and a simplified relative kinetic model are used, the in-orbit real-time processing is carried out on GPS/BDS double-frequency observation data of formation satellites, in consideration of limited computing resources of a satellite-borne processor, an operation mode of dynamic orbit prediction matched with orbit interpolation is designed, a core complex data processing process is realized in multiple steps in a time-sharing manner to reduce the computing time consumption of a single epoch, and high-precision absolute and relativeorbit results of the formation satellites with adjustable frequencies can be output. The method is suitable for inter-satellite formation with a large inter-satellite distance range, a smooth navigation result meeting certain precision requirements still can be continuously output under the condition that GPS/BDS signals are interrupted, and the method has the advantages of high orbit determination precision, good stability, low requirement for the performance of the satellite-borne processor, easy engineering achievement and the like.

The invention discloses an on-satellite autonomous satellite orbit prediction method. Firstly, a user comprehensively considers prediction precision of an orbit and the acceptable running time of an actual project and presets an orbit prediction period and segment time intervals, and a ground control system performs precise prediction on orbit prediction requirements proposed by the user to generate all segment initial orbital parameters in the orbit prediction period; secondly, the ground control system inputs the generated segment initial orbital parameters into an on-satellite control system sequentially in accordance with the preset segment time intervals as the input of the on-satellite control system autonomous satellite orbit prediction; and finally, the on-satellite control systemperforms on-satellite autonomous orbit prediction of based on the segment initial orbital parameters. Based on satellite-ground cooperation, the high computing power of the ground is effectively exerted. The operation time of an on-satellite autonomous satellite orbit prediction algorithm is reduced. Engineering practice precision of orbit prediction is balanced.

The present invention discloses an inter-satellite link antenna scanning acquisition device and method. The device comprises: an orbit prediction module configured to predicate antenna parameters; a scanning parameter setting module configured to obtain scanning parameters; a maximum capture time calculation module configured to obtain the required time of scanning the whole determinate region according to the antenna parameters when two satellite antennas are rotated at opposite directions; a timer module configured to limit the scanning time of the two satellite antennas being rotated; a first scanning strategy module configured to perform scanning of the rotation of the two satellite antennas at the opposite directions and determine whether the acquisition is successful or not; and a second scanning strategy module configured to allow the two satellite antennas to return back to the direction of the opposite antenna through orbit prediction at a constant linear velocity and perform scanning in a mode of only one satellite antenna rotation when the two satellite antennas cannot be subjected to successful acquisition in a maximum acquisition time, and determine whether the acquisition can be successful or not. The device can effectively reduce the acquisition time and improve the probability of successful acquisition in a short time.

A method and apparatus for processing and transmitting precise orbit predictions of satellites in a Global Navigation Satellite System such as Navstar-GPS which employs curve fitting techniques and a polynomial and sinusoidal model to encode ephemerides, and particularly ephemerides of durations 6 hours or longer, in order to minimize bandwidth requirements over-the-air and NVRAM storage requirements. The methods also apply to GNSS constellations such as Galileo or GLONASS. Also, methods are disclosed for recovering the ephemeris on a GNSS receiver device in the constellation's native format.

The invention discloses a star orbit prediction method and device. The method comprises the following steps: acquiring geographic position information of a current geographic position; taking a starry sky picture at the current geographic position; and generating a star orbit image of the current geographic position according to the starry sky picture and the geographic position information. Through adoption of the star orbit prediction method and device, the starry sky picture under a current shooting condition can be predicted through a mobile terminal, so that a user can pre-judge the starry orbit picture formed under the current shooting condition conveniently.

The invention discloses a two-line element generation method based on model error compensation. According to the method, first, observed position and speed vectors are converted to a true equator mean equinox coordinate system through a geocentric equatorial inertial coordinate system, then corresponding instant orbital elements are computed and are used as the initial estimated values of two-line element iterative computation; the position and speed vectors at the corresponding sampling instant are respectively predicted according to the initial estimated values, the difference values between the position and speed vectors and the position and speed vectors at the practical observing sampling instant are computed, whether a termination condition is met is judged; through a numerical differentiation method, a partial derivative matrix about the two-line elements of a simplified pervasive perturbation orbit prediction function at the sampling instant is computed; by solving a compensation least square problem, the correction of the two-line elements is obtained; new two-line elements are computed, and repeated iterative operation is carried out until the accuracy requirement is met. According to the method, the situations of single-point fitting and sampling fitting can be processed, and average errors are small when the two-line elements obtained by fitting are used for orbit prediction.

The invention discloses a satellite measurement and control station antenna automatic control method based on cloud service, and the method comprises the steps that a cloud server generates an orbit prediction and time window set for related satellites according to a measurement and control task specified by a user through an SGP4 algorithm; and then the server performs task planning on the antenna resources of each measurement and control station by using a genetic algorithm based on the generated time window set to generate an optimal scheduling scheme meeting requirements. The server allocates tasks to the measurement and control stations in various places according to the scheme, the measurement and control stations in various places read the tasks and then automatically control the antennas of the stations to track, capture and release related satellites, and the task execution conditions are fed back to the server. The invention has the advantages that automatic control over themeasurement and control antenna is achieved, meanwhile, the utilization rate of measurement and control resources is increased through algorithm planning and optimization, and the overall measurementand control efficiency of the system is also improved.

The invention discloses a time prediction method for earth-moon picture shooting by an ultraviolet sensor at cislunar fight, which comprises steps of (1) selecting an orbit arc meeting the light condition for earth-moon shooting, (2) selecting an orbit arc, in which earth-moon appears, within viewing field area of the ultraviolet sensor, (3) taking the orbit position simultaneously meeting the step (1) and the step (2) as a selectable orbit for shooting, inputting an orbit prediction position corresponding to the selectable orbit, attitude matrix at a certain time before the selectable orbit, sun-moon-earth ephemeredes and a satellite attitude motion model after launching the satellite and before the coming of the selectable orbit, calculating the attitude matrix and a selenocentric/geocentric vector within the prediction time period and making a curve, changing along the time, of an included angle of the selenocentric/geocentric vector and a viewing field central vector, then obtaining continuous visible time range(t1-t2) of earth/moon by the curve and the size of the viewing field of the ultraviolet sensor and calculating the best time according to the obtained continuous visible time range.