274 results about "Orbit" patented technology

A heat balance optimal satellite sun orientation control method comprises the following steps: firstly, determining sun orientation target parameters according to a sun vector direction calculated by an orbit, using preset sun orientation target parameters when regular preset parameters are set, and then calculating a sun orientation target attitude angle and angular velocity by using a formula provided by the method, and finally, the sun orientation of the satellite to the target attitude angle and angular velocity is realized according to the measurement of the terahertz gyroscope and the plane control of the jet phase.

The invention relates to a lunar satellite formation collaborative navigation method, which comprises the following steps of: inputting an orbit state and an observation residual error of a lunar satellite formation at a previous moment into an LSTM (Long Short Term Memory) prediction network of a navigation algorithm to predict a process noise covariance matrix and an observation noise covariance matrix, and multiplying the process noise covariance matrix and the observation noise covariance matrix by a scaling factor for scaling, obtaining a scaled prediction process noise covariance matrix and a scaled observation noise covariance matrix; and inputting the orbit state of the lunar satellite formation at the previous moment, the multi-source observation data at the current moment and the scaled process noise covariance matrix and observation noise covariance matrix into a self-adaptive extended Kalman filtering module in a navigation algorithm to predict the orbit state of the lunar satellite formation at the current moment.

The application discloses a fusion positioning algorithm based on low-orbit satellite Doppler measurement, DME and VOR, relates to the technical field of satellite navigation, and comprises the following steps: acquiring the Doppler frequency shift of a low-orbit satellite, the slant range information of a DME and the azimuth angle data of a VOR, constructing a combined observation equation, and obtaining a corresponding Jacobian matrix; based on a geometric dilution of precision, selecting stations and stars through a genetic algorithm, optimizing the selection of the DME, the VOR and the low-orbit satellite, and obtaining an optimal navigation source combination; constructing a motion model under different flight modes, and dynamically estimating the state of an aircraft by using an interactive multiple model Kalman filter, and updating the position information in real time. Therefore, the real-time positioning of the aircraft under complex conditions is realized, the anti-interference capability of the system is improved, and the overall positioning precision and stability of the system are improved.

This invention provides a method and system for satellite insertion into a highly elliptical frozen orbit, including multiple small-thrust maneuvers at apogee to raise the perigee altitude; naturally increasing the perigee angle using positive perturbation factors; evenly and rationally allocating the working time of each thruster; using inertial pointing to complete each ignition maneuver; positively offsetting the target orbit perigee altitude; performing small-thrust maneuvers near apogee to lower the perigee altitude to the target value, ultimately completing satellite insertion; the ignition altitude is selected above 28,000 kilometers. In this invention, the satellite can still enter its mission orbit under fault conditions, ensuring the satellite's service life.

The invention discloses a tethered satellite system dynamics modeling method considering perturbation factors. The tethered satellite system dynamics modeling method comprises the steps that a geocentric coordinate system, an orbital coordinate system of a Kepler orbit where a tethered satellite system is located and a body coordinate system of a unit satellite are established; according to the movement of the centroid of the tethered satellite system on the orbit, the generalized coordinates are the distance from the centroid of the system to the earth center and the true anomaly of the orbit, the Lagrange function of the centroid of the system is obtained; according to the movement of the unit star relative to the centroid, the generalized coordinates are the distance between the unit star and the centroid, the relative first rotation angle and the relative second rotation angle, the relative Lagrange function of the unit star is obtained; comprehensively considering perturbation factors to obtain non-spherical earth, third gravitational force and solar pressure perturbation potential energy; establishing a system kinetic equation considering comprehensive perturbation factors; and obtaining a complete Lagrange equation of the tethered satellite system. The error problem caused by the fact that the master satellite directly serves as the mass point of the system and the influence of comprehensive perturbation on system motion is not considered is solved, and the number of the slave satellites is not limited.

The invention provides a satellite-borne high-precision orbit extrapolation method, a satellite-borne high-precision orbit extrapolation device and electronic equipment. The satellite-borne high-precision orbit extrapolation method comprises the following steps: providing an orbit extrapolation input quantity of a satellite at an initial extrapolation moment; calculating an extrapolation duration based on the initial extrapolation moment and a preset threshold value, and determining a termination extrapolation moment of the satellite based on the extrapolation duration; calculating the non-spherical gravitational perturbation acceleration of the earth according to the non-spherical gravitational perturbation model of the earth; calculating a day-moon three-body gravitational acceleration according to a day-moon three-body gravitational model; determining a satellite thrust acceleration according to the current working state of the satellite; adding the earth non-spherical gravitational perturbation acceleration, the sun-moon three-body gravitational acceleration and the satellite thrust acceleration to obtain an extrapolation resultant acceleration; performing orbit extrapolation calculation on the basis of the extrapolation resultant acceleration by adopting a fourth-order Runge-Kutta method to obtain a target extrapolation position and a target extrapolation speed at an extrapolation termination moment; and according to the target extrapolation position and the target extrapolation speed, updating the orbit extrapolation input quantity at the next extrapolation moment.

A method and system for low Earth orbit (LEO) satellite multiple access and communication using grating lobes is disclosed. The method may include: acquiring ephemeris data for a plurality of LEO satellites; predicting an orbital trajectory for each of the plurality of LEO satellites based on the ephemeris data; selecting target satellites using the predicted orbital trajectories; adjusting an antenna array to form grating lobes in the directions of the selected target satellites; and transmitting and receiving data with the target satellites using the adjusted antenna array.

The invention discloses a spacecraft in-orbit operation health assessment method, and relates to the technical field of space flight and aviation. Environmental data such as gravity gradient, radiation intensity and dust particle density of an asteroid surface are collected in real time through a multi-mode sensor array, modeling and prediction are carried out by adopting a dynamic Bayesian network in combination with Kalman filtering, and the health assessment accuracy of the spacecraft in-orbit operation is improved. The method comprises the following steps: dynamically generating environment characteristic parameters including a risk level and an environment change trend, and in health state prediction, simulating stress distribution in a low-gravitation environment through digital twinborn body and finite element analysis, predicting a fatigue state and a failure risk of a contact device, and accurately evaluating the fatigue life of a high-stress area. The quantification capability of the failure risk is improved through the risk probability model; besides, according to a health prediction result, landing parameters including a damping mode and a landing speed are dynamically optimized, and a dynamic planning method is used for minimizing impact force and stress load, so that the detector can be in stable contact in a low-gravity environment.

The invention discloses an earth-moon translation point orbit spacecraft formation relative motion configuration design and control method, and belongs to the technical field of spacecraft orbit dynamics and control. The method comprises the following steps: firstly, constructing a double-layer dynamic model consisting of a circular restrictive three-body problem model and a high-precision ephemeris model, and defining an orbit amplitude difference and a phase difference as configuration parameters; then, according to different track amplitude difference and phase difference combinations, three types of natural relative motion configurations are set, a periodic track is rapidly generated under a circular restrictive three-body problem model, and multi-circle correction is conducted through a high-precision ephemeris model and two-stage multiple targeting; a hierarchical constraint model is established for a controlled task surrounding a fixed position, a plane and a space, and a control pulse is solved by adopting a two-stage optimization strategy. According to the method, efficient design and stable control of formation configuration in a complex dynamic environment are realized, and the method is suitable for various translation point orbit tasks.

The invention relates to a lunar satellite formation orbit adaptive estimation method based on EM-EKF. Comprising the steps that an extended Kalman filtering module of a navigation algorithm predicts the orbit state of the lunar satellite formation at the current moment based on the orbit state of the lunar satellite formation at the previous moment, observation data at the current moment, a process noise covariance matrix and an observation noise covariance matrix; and updating a process noise covariance matrix and an observation noise covariance matrix by an expectation maximization module of a navigation algorithm based on a state sequence formed by a series of orbit states output by the extended Kalman filtering module within a period of time. According to the method provided by the invention, the process noise covariance matrix and the observation noise covariance matrix can be dynamically estimated, and online adaptive updating of the extended Kalman filtering parameters is realized.

The invention relates to the technical field of camera on-orbit calibration, in particular to a fisheye camera on-orbit calibration method based on a starlight vector, which comprises the following steps of: S1, acquiring a plurality of star maps observed by a fisheye camera on orbit, and extracting and denoising star points to form a star point set; s2, according to fisheye imaging distortion distribution, a star point located in an imaging center area is selected from a star point set, and a starlight vector is constructed based on image point coordinates of the star point in combination with star catalogue data subjected to space-time deviation and propagation path deviation correction; and S3, performing local geometrical relationship star point matching on the starlight vector in the step S2. According to the method, the multiple star maps observed by the fisheye camera in orbit are obtained and processed to form the star point set, so that the problems of poor in-orbit calibration applicability and low precision caused by the fact that most traditional fisheye camera calibration methods adopt a ground pre-calibration or space reference object arrangement mode are solved.

The present disclosure provides a satellite system with orbital debris avoidance. In one embodiment, the satellite system includes debris sensor circuitry to scan orbital debris items in an orbital debris field; and sensor controller circuitry to determine a search window within the orbital debris field based on, at least, uncertainties associated with a velocity of the satellite (Vsatellite) and uncertainties associated with a velocity of at least one debris item (Vdebris), the sensor controller circuitry also to control the debris sensor circuitry to detect the at least one debris item within the search window.

The invention discloses a fly-around path autonomous planning method, which comprises the following steps of: 1, drawing up a mark track in a fly-around process as an elliptical track, and solving a minor semi-axis, a major semi-axis and a total fly-around angle of the elliptical track according to the current relative position and the preset tail end relative position of a tracking satellite and a target satellite; step 2, obtaining a motion mode of the aircraft on an elliptical orbit, and obtaining a formation vector design of an elliptical fly-around plane according to a minor semi-axis, a major semi-axis, a total fly-around angle, a total fly-around time and a current fly-around time of an ellipse, the formation vector comprising a position and a speed of a tracking star under a target star orbital coordinate system; and step 3, rotating the formation vector of the fly-around plane into an expected plane according to three parameters of the set fly-around elliptical arch line argument, the fly-around ascending intersection angle and the fly-around inclination angle. According to the method, the planned path is used as a nominal track, and flying-around approaching to the target or flying-around evacuating from the target can be realized by applying control.

The application provides a debris cloud disintegration timing inversion method based on space-time intersection feature statistics, belongs to the field of disintegration timing inversion, and solves the problem of how to improve the accuracy of disintegration timing determination; the method comprises the following steps: traversing the standardized debris cloud orbit data after orbit propagation, calculating the closest distance time of each two pieces of debris; using an Epanechnikov kernel function to perform non-parametric probability density estimation on the closest distance time; combining the inverse function of the cumulative distribution function with the significance level to construct a confidence interval, and removing the debris outside the confidence interval; traversing the debris in the confidence interval, calculating the closest distance time of each two pieces of debris; using an Epanechnikov kernel function to perform non-parametric probability density estimation on the closest distance time; and taking the closest distance time corresponding to the maximum probability density estimation value as the disintegration time of the space target; the application has strong universality and strong robustness.

The application discloses a small-sized separated penetrator for in-situ detection of celestial bodies, and relates to the technical field of penetrator equipment, comprising an aluminum honeycomb buffer module, which comprises an external protective shell and an aluminum honeycomb, and has the beneficial effects that the application is provided with in-situ detection of high-speed penetration of celestial bodies, and discloses a small-sized separated penetrator, which comprises a planet soil penetration module and a star table detection module, and realizes autonomous separation in the impact penetration process, the star table module can realize star table detection and communication with an orbiter, the penetration module is internally provided with an energy storage device and an impact hammer, and can continue to penetrate and dive relying on the energy storage device after impact; the star table module is connected with the penetration module through flexible wires, and can supply power and signal transmission to the latter, the technology enables the penetrator with the modular characteristic to have the in-situ detection capability of planet soil profile, aims to expand the application mode of the penetrator for celestial body detection, and provides a scientific basis for the research of star table structures.

The invention relates to a high and low orbit satellite tracking method based on geomagnetic correction. The method comprises the following steps: constructing a geomagnetic model, calculating an azimuth angle and carrying out magnetic field interference analysis; carrying out factory calibration by utilizing a three-dimensional geometric calibration mode, including horizontal correction and vertical correction; the current azimuth angle and pitch angle are continuously corrected based on the carrier coordinate system by using the signal intensity, and the deviation is obtained so as to continuously approach the real angle; utilizing the known coordinates of the three-dimensional space to inversely calculate the current position; and the current data is corrected by using the big data, so that the current data meets the use requirements. The current attitude angle and the satellite signal receiver are used for tracking the target satellite, and then the known azimuth angle is used for correcting the geomagnetic data, so that the azimuth angle precision is improved, the cost of a phased-array antenna or a mechanical antenna is reduced, the tracking precision is improved, and the tracking speed is improved.

The invention discloses a method for improving large-scale star orbit mask calculation efficiency. Relates to the technical field of space target observation, in particular to a method for improving large-scale star orbit mask calculation efficiency. The method is used for quickly removing dense fixed star orbits in space target observation image processing, so that the overall performance and the processing speed of target extraction are improved. The method comprises the following steps: constructing an original mask substrate according to the size of an observation image; clockwise rotating the original mask base plate around the original point according to the inclination direction of the star orbit to obtain a rotary mask base plate, and expanding around the rotary mask base plate to obtain a rotary canvas; solving a star orbit mask of each fixed star orbit in the rotating canvas in the original mask base plate; performing rotation inverse transformation on the rotating star orbit mask, and performing expansion around the rotating canvas after the inverse transformation to obtain an inverse transformation canvas; and cutting the expanded parts of the rotating canvas and the inverse transformation canvas to obtain a final star orbit mask.

The application relates to a method and device for automatically planning a spacecraft orbit calculation and an electronic device. The epoch time for which orbit calculation is to be performed is obtained, the last update time of Kepler six elements is obtained, the data delay duration is determined according to the epoch time and the last update time of the Kepler six elements, the target timeout threshold is determined based on at least one of the current orbit type, the spacecraft working mode and the spacecraft state, and the spacecraft orbit calculation is performed based on the Kepler six elements to determine the position and speed of the spacecraft when the data delay duration is less than or equal to the target timeout threshold. The comparison mechanism between the data delay duration and the dynamically configured target timeout threshold is established, intelligent decision of the orbit calculation mode is realized, the most suitable orbit calculation mode can be automatically selected according to the actual situation, the uncertainty caused by manual intervention is avoided, and the demand of different scenes can be met.

This invention discloses an on-orbit evaluation method for the single-frame attitude accuracy of a star sensor, comprising: S1, calculating the single-frame attitude covariance matrix of the star sensor by combining the synthesized value of the variance of noise from all star vector measurements; S2, evaluating the three-axis attitude error of the star sensor based on the attitude covariance matrix; S3, calculating the single-frame attitude quality factor of the star sensor based on the three-axis attitude error value; and S4, evaluating the attitude error of the star sensor based on the quality factor. This invention achieves on-orbit real-time calculation and evaluation of the single-frame attitude accuracy of a star sensor by calculating the covariance matrix of the star sensor and constructing a "quality factor".

To provide a method of routing data traffic in non-geostationary orbit (NGSO) satellite systems.SOLUTION: The method utilizes, at satellites 102-107, backbone route tables that define different snapshots of time-varying backbone topology of an NGSO satellite system 100. Because each satellite 102-107 may refer to a different portion of the backbone topology, the backbone route tables at the satellites 102-107 are different from each other. In addition, the backbone route tables at the satellites 102-107 are active at different times because the backbone topology changes over time. Different routing groups of backbone route tables are defined to realize different routing options for packets arriving at the satellites 102-107.SELECTED DRAWING: Figure 1

One or more embodiments of the present application provide a time calibration method, system, device, equipment and medium, a low-orbit satellite receives a first reference time, and calibrates a current first time using the first reference time; determines a clock difference with a GNSS according to at least one observation data sent by the GNSS, and calibrates the first time according to the clock difference to determine a second time; receives other clock differences sent by other low-orbit satellites, and calibrates the second time according to a clock difference average value of the other clock differences and the clock difference, and determines the calibrated second time as a calibrated current time. Since the low-orbit satellite determines a clock difference between the low-orbit satellite and the GNSS according to at least one observation data sent by the GNSS, and calibrates the time of the low-orbit satellite according to the clock difference and other clock differences sent by other low-orbit satellites, the time carried by the low-orbit satellite meets the timing accuracy required by GNSS navigation and positioning.

The invention provides a spacecraft attitude orbit navigation initial value on-orbit establishment method and device, and belongs to the technical field of spacecraft navigation.The method comprises the steps that a navigation system single-machine product configuration scheme and a single-machine installation layout scheme are designed; judging whether the platform is successfully separated from the mother platform or not; the self angular rate is damped; performing slow rotation search around two axes of the body, and preliminarily establishing an inertial system attitude initial value; carrying out slow rotation search around two axes of the body, and preliminarily establishing an initial speed value of the earth fixed connection position; performing attitude maneuver based on the inertial system attitude initial value and the earth fixed connection system position speed initial value, and adjusting the attitude to a final stable direction capable of capturing the attitude and the position speed at the same time; and according to the attitude and position speed result, calculating the position speed and attitude result under the expected reference system as a final inertial navigation initial value. According to the invention, a software and hardware interface for information transmission does not need to be arranged between the spacecraft and the release mother platform, and the capability of on-orbit self-establishment of inertial navigation attitude and position speed initial values is realized.

This application provides a method, device, and computer-readable storage medium for evaluating the on-orbit calibration error of a microwave radiometer calibration system used in geostationary orbit. The application evaluates the on-orbit calibration error of the microwave radiometer calibration system by determining the errors of the calibration cold source, calibration heat source, Earth observation radiation brightness temperature, and Earth target brightness temperature. By separately evaluating the errors of the calibration cold source, calibration heat source, Earth observation radiation brightness temperature, and Earth target brightness temperature in the microwave radiometer calibration system used in geostationary orbit, each error is quantified, thus quantifying the on-orbit calibration error of the microwave radiometer calibration system. Furthermore, the different radiation transmission paths are considered during the evaluation of each error, making the evaluation of the sources of on-orbit calibration error of the microwave radiometer calibration system used in geostationary orbit not only effective and accurate but also highly implementable.

This disclosure provides a method, apparatus, and storage medium for on-orbit optical axis calibration of a spectrometer, relating to the field of astrophysical detector calibration technology. The method includes: performing a helical scan of the celestial body under test; during the scan, continuously acquiring the imaging position of the centroid of the celestial body on the imaging target surface of the imaging device, and acquiring the thermal radiation energy of its pointing position at a first angle based on a preset acquisition frequency using the spectrometer; inverting the continuously acquired imaging positions onto the first target surface, and mapping the acquired thermal radiation energy to the corresponding position on the first target surface according to the acquisition time, generating first thermal radiation data; and using the position of the point with the highest thermal radiation energy in the first thermal radiation data on the first target surface as the optical axis pointing position of the spectrometer at the first angle. This disclosure achieves joint observation by synchronously controlling the imaging device and spectrometer of the astrophysical detector, and determines the true pointing position of the spectrometer's optical axis through the observation results, which is simple to operate and has high accuracy.

The application discloses a satellite attitude-orbit integrated control method for on-orbit capture, and belongs to the technical field of micro-satellite attitude-orbit integrated control. The application aims at the problem that the control accuracy is influenced by the cooperative delay of discrete actuators in the attitude-orbit integrated control. The application comprises the following steps: eight inclined thrusters are symmetrically arranged on two mounting surfaces and are symmetrically arranged on diagonal lines; based on the expression of the thrust and the force moment of the thrusters on the tracking star, a target function is constructed under the constraint condition, the thrust output meets the task requirement as a prerequisite, and the maximum moment output capacity is taken as a target to determine the mounting angle of the thrusters; a limiting amplitude correction factor is calculated by using a limiting amplitude method based on the thrust envelope, and the output capacity of the thrusters is maximally utilized as a target; a hybrid target function of the minimum fuel consumption and the minimum residual error of the moment distribution is constructed, the thrust value of each thruster is calculated, and the control of the thrusters is realized. The application is used for the satellite attitude-orbit integrated control.

This invention discloses a method for improving the orbital accuracy of a lunar probe based on reflective prism constraints. The method first utilizes laser ranging data obtained from a known reflective prism on the lunar surface using an onboard laser altimeter, and constructs an error equation set by combining this data with a dynamic orbital model. Next, a high-resolution camera is used to perform multiple imaging operations on the lunar reflective prism from different orbits and perspectives. Using the bundle adjustment principle, the observed image point coordinates are combined with the dynamic model to construct the error equation set. Finally, laser ranging observations from the reflective prism and multi-temporal photogrammetry observations are used as two different types of geometric constraints to construct a joint adjustment model of multi-source observation data. The high-precision orbital parameters are then calculated using a least-squares algorithm. This invention uses laser ranging data and an optical image database to add new constraints to the estimated orbit, ultimately correcting and further improving the orbital accuracy of the lunar probe.

The application provides a spacecraft autonomous navigation method and electronic equipment. The spacecraft autonomous navigation method comprises the following steps: determining technical indexes of an optical sensor according to navigation requirements, and determining a configuration of the optical sensor according to orbit information of the spacecraft; collecting a star map by using the optical sensor with the determined technical indexes and configuration, processing the star map, identifying a refraction star and extracting a refraction angle; and using a Kalman filtering algorithm to perform information fusion on information output by an inertial navigation system and a refraction apparent height calculated based on the refraction angle, so as to correct navigation errors and obtain corrected navigation information. The application corrects drift errors of the inertial navigation by astronomical navigation, maintains continuous navigation information by using the inertial navigation when the astronomical system is unavailable, and achieves the purpose of improving navigation accuracy.

A monitoring system (1) for an orbiting space object (5) comprising a ground-based radar system (2) (7) including an antenna (6) with a controllable pointing direction for moving a space observation window (11) within a field of view (12), the pointing direction having a predetermined value (Zref) associated with the space object (5), the monitoring system (1) comprising ground-based optical sensors (3) (7) and a control unit (4) configured to determine a first orbital restitution (O1) from first positions (P1) of the space object (5) measured by the optical sensors (3), and, if the first orbital restitution (O1) crosses the field of view (12) of the radar system (2) outside the space observation window (11) associated with the predetermined value (Zref) of the pointing direction, to determine a command for the pointing direction of the antenna (6). Abstract figure: Figure 1

The application provides a satellite on-orbit automatic shading heat control device and a satellite in the technical field of spacecraft heat control, which comprises a star body mechanism, an automatic unfolding and folding mechanism and a wave-transparent heat dissipation baffle, the star body mechanism is connected with the wave-transparent heat dissipation baffle through the automatic unfolding and folding mechanism, the wave-transparent heat dissipation baffle comprises a wave-transparent nano heat control film and a wave-transparent honeycomb, and the wave-transparent nano heat control film is arranged on two planes of the wave-transparent honeycomb. The application greatly reduces the reflection of the satellite to the ground, effectively shades light and increases the heat dissipation area, and is beneficial to the sun-shading and heat dissipation of the satellite.