26 results about "Orbital inclination" patented technology

The invention relates to an interferometric synthetic aperture radar (InSAR)-based measurement method and system, and the method comprises the steps: determining the prior positions of a plurality of targets, and providing the prior positions to a ground processing unit; shooting an area where a plurality of targets are located by using an SAR satellite, and sending the obtained data to a ground processing unit; determining a phase difference between the targets for a single shot of the target; according to the determined phase difference and the SAR load wavelength, the slope distance between the targets is calculated through a ground processing unit; according to the slope distance and the satellite orbit information, the position relation between the targets is determined through a ground processing unit. The measuring system at least comprises a ground processing unit and an acquisition unit. According to the invention, shooting at the same track inclination angle and different track inclination angles can be realized, the limitation of heavy track shooting conditions in the prior art is broken through, and high-frequency and high-precision surface micro deformation measurement can be realized for different application scenes.

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 north-south electric propulsion inclination angle keeping control method and device. The application separates the long-term term and the half-year period term of the GEO satellite orbit inclination angle, respectively constructs motion equations, and constructs keeping constraint conditions by making full use of the relationship between the half-year period term perturbation law and the north-south keeping range, so that the keeping strategy calculated based on the keeping constraint conditions not only takes into account the advantage of high control stability of the electric propulsion method, but also improves the accuracy of the control time and the ignition time by respectively calculating the control strategy of the long-term term and the half-year period, can effectively reduce the electric propulsion ignition time length, realizes the premise that the orbit inclination angle meets the keeping requirements, reduces the influence of the electric propulsion method on the satellite longitude, reduces fuel consumption, and the ignition point longitude fluctuation in the control strategy cycle is small.

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 cooperative detection and tracking constellation design method for high-maneuvering cross-domain targets, comprising the following steps: determining a high-orbit constellation configuration, initializing and determining the number of high-orbit satellites; calculating the relationship curve between the double coverage percentage of the high-orbit satellites and the orbit inclination curve, and selecting the orbit inclination corresponding to the optimal coverage performance; randomly generating a plurality of high-maneuvering cross-domain targets, and calculating the maximum position error and velocity error of the high-orbit constellation tracking targets; judging whether the handover of the low-orbit satellite detector is successful according to the field of view size of the low-orbit satellite detector; determining a high-orbit constellation configuration, initializing and determining the number of low-orbit satellites; setting the initial tracking error according to the high-maneuvering cross-domain targets, and calculating the maximum position error and velocity error of the low-orbit constellation tracking targets; judging whether the maximum position error and velocity error of the low-orbit constellation tracking targets meet the index requirements, and if not, repeating the method until the index requirements are met.

The invention discloses an integrated antenna mechanism driver, and relates to an antenna mechanism driver. The integrated antenna mechanism driver comprises a communication assembly, a controller assembly, a driver assembly and a posture collection assembly, wherein the communication assembly, the controller assembly, the driver assembly and the posture collection assembly are integrated together; the communication assembly, the controller assembly, the driver assembly and the posture collection assembly are connected through connectors. The integrated antenna mechanism driver is used for controlling and driving an external antenna, the single-satellite cost needs to be reduced through modular design, and the transmission reliability needs to be achieved through a space-air-ground integrated network. The driving mechanism is used for the low-orbit communication satellite, the satellite orbit inclination angle change can be obtained, the azimuth angle / pitch angle can be calculated, and the service life and the communication stability of the satellite are guaranteed.

The application discloses a kind of latitude zone full coverage nonlinear time-varying load constellation configuration analytical design methods, comprising: according to satellite orbit altitude and load detection geometry constraint, the analytical model of fan-shaped load ground detection coverage is established, the latitude and longitude extreme value of fan-shaped coverage area and nonlinear time-varying relationship changing with orbit position are obtained;According to the latitude and longitude extreme value of fan-shaped coverage area and nonlinear time-varying relationship changing with orbit position, and target latitude zone range, by solving orbit inclination, satellite number in orbit plane, orbit plane number and orbit plane satellite phase difference gradually through analytical constraint model, the constellation configuration parameters satisfying target latitude zone full coverage requirement are obtained;Satellite constellation system is designed and deployed according to constellation configuration parameters.The method of the application can be directly used for satellite launch into orbit after orbit deployment and on-orbit operation control, without additional numerical iteration optimization.

The invention discloses an automatic division method and system for a satellite constellation orbital plane, and relates to the technical field of spaceflight measurement and control. The method comprises the following steps: firstly, calculating an orbital plane unit normal vector based on a satellite orbit inclination angle and an ascending node right ascension, and performing first-level spatial pointing clustering by adopting a density clustering algorithm to form an orbital plane group; secondly, in the same orbital plane group, by taking a semi-major axis and an eccentricity rate as characteristics, automatically determining an optimal shell layer number by utilizing a Gaussian mixture model and a Bayesian information criterion, and realizing second-stage orbital energy shell layer division; and finally, in the same shell layer, sorting and interval statistical analysis are performed on the right ascension of the ascending node, recursive subdivision is performed through dynamic threshold segmentation in combination with a range-based dual-condition verification mechanism, and fine division of a third-level orbital plane is completed. The method realizes full-process automation from orbit data to division results, has the advantages of high division precision, strong adaptability, good engineering practicability and the like, and is suitable for operation and maintenance management of large-scale heterogeneous satellite constellations.

The application discloses a rocket launch parameter and a detector orbit injection orbit parameter determination method and device, belongs to the technical field of spacecraft orbit design, and is used for jointly and rapidly determining the rocket launch parameter and the detector orbit injection orbit parameter. In the method, the rocket inclination value range and the corresponding rocket sliding time value range are determined according to the detector orbit injection orbit inclination, the expected rocket sliding time is determined based on the detector orbit injection orbit perigee amplitude, and the expected rocket orbit injection orbit ascending node longitude is determined based on the expected rocket sliding time. The expected detector orbit injection time is determined according to the ascending node longitude difference between the rocket and the detector. The rocket launch time and the launch azimuth angle are determined according to the detector orbit injection orbit inclination, the rocket inclination value range, the expected rocket sliding time, the rocket sliding time value range and the expected detector orbit injection time, and the rocket launch parameter consistent with the detector orbit injection orbit parameter is obtained.

The invention discloses a method, system and device for judging the instability degree of an electric power rope system, a medium and a product, and relates to the technical field of spacecraft attitude control, and the method comprises the steps: obtaining an orbit inclination angle of a deployment orbit of a main satellite in the electric power rope system; determining an instantaneous balance point distribution area of the electric power rope system based on the track inclination angle; the optimal balance area of the electric power rope system is positioned in the instantaneous balance point distribution area of the electric power rope system; and acquiring an attitude angle of the electrodynamic force rope system at the current moment, and determining the instability degree of the electrodynamic force rope system at the current moment based on the attitude angle of the electrodynamic force rope system at the current moment, the instantaneous balance point distribution area and the optimal balance area. According to the invention, the instability evaluation efficiency and reliability of the electric power rope system are improved.

The present application belongs to the technical field of satellite remote sensing, and particularly relates to a kind of orbit inclination design method.The technical scheme is a kind of optimal orbit inclination calculation method for realizing high revisit rate of specified latitude range, and the features are as follows: when the satellite is at the intersection point of ascending orbit and descending orbit, the edge of scanning strip close to the earth's poles coincides with the highest coverage latitude of the remote sensing area; and the edge of scanning strip close to the earth's equator coincides with the most coverage times latitude of the remote sensing area, at this time the orbit inclination is optimal, and the calculation formula of optimal orbit inclination i is: ∠AOa is the most coverage times latitude, which is a known value; and φ is the corresponding earth center angle of satellite scanning strip width.The present application studies the law that the orbit inclination has optimal solution with the orbit inclination of the most coverage times latitude ∠AOa, provides a kind of optimal orbit inclination calculation method for realizing high revisit rate of specified latitude, so that the optimal orbit inclination is obtained by calculation, and the satellite satisfies the most single-day coverage times at the most coverage times latitude.

A satellite constellation includes a plurality of artificial satellites flying on an inclined circular orbit for each of six or more orbital planes having a common orbital inclination and arranged so that south and north axes are mutually shifted in an east-west direction. The plurality of artificial satellites include eight or more artificial satellites for each orbital plane. Each of the artificial satellites includes a fore-and-aft communication device. For each orbital plane, each of the artificial satellites forms a communication network with front and rear artificial satellites by the fore-and-aft communication device. Each of the artificial satellites on each orbital plane crosses southern and northern edges of the orbital plane in synchronization with artificial satellites on other orbital planes, and forms a communication network with that artificial satellites by the fore-and-aft communication device.

This invention provides a method, apparatus, and medium for launching satellites in emergency scenarios, comprising: if the satellite launch point is a fixed value and the orbital inclination is an unknown value, determining the first latitude argument and the geographical longitude of the first ascending node when the satellite passes through the launch point at the first current orbital inclination, and the second latitude argument and the geographical longitude of the second ascending node when passing through the observation point, based on the coordinates of the observation point and the coordinates of the launch point; determining the target orbital inclination of the satellite and the corresponding launch window based on the first latitude argument, the geographical longitude of the first ascending node, the second latitude argument, and the geographical longitude of the second ascending node; if the satellite launch point is an unknown value and the orbital inclination is a fixed value, determining the longitude and latitude of the nadir point corresponding to each launch window based on the coordinates of the observation point, the target observation time, and the observation time step; thus, in emergency launch scenarios, the launch window can be determined according to the launch requirements, thereby quickly and accurately launching the satellite to a preset position to achieve fixed-point observation of emergency events.

The application discloses a GNSS-based extrapolation method for a large-ellipse frozen orbit, which comprises the following steps: step one: acquiring the position and speed of a navigation receiver orbit according to the effective state of navigation data; step two: calculating satellite orbit elements (distance and speed values, moment of momentum, semi-major axis, orbital angular velocity, eccentricity, orbital inclination, ascending node right ascension, latitude amplitude angle, and perigee amplitude angle) by using t GNSS moment of momentum, semi-major axis, orbital angular velocity, eccentricity, orbital inclination, ascending node right ascension, latitude amplitude angle, and perigee amplitude angle); step three: judging the effectiveness of navigation data according to the calculation results of the satellite orbit elements; step four: extrapolating and calculating the orbit parameter true anomaly and latitude amplitude angle at the current satellite time; and step five: calculating the instantaneous orbital angular velocity, thereby completing the GNSS-based orbit extrapolation design for a large-ellipse frozen orbit. The application only uses the orbiting information of the GNSS time to extrapolate the orbit elements of the satellite, thereby achieving satellite autonomous positioning and autonomous attitude control, reducing the positioning requirements of the satellite to the ground station, and improving the autonomous ability of the satellite.

A satellite constellation (110) includes a plurality of artificial satellites flying on an inclined circular orbit for each of six or more orbital planes having a common orbital inclination and arranged so that south and north axes are mutually shifted in an east-west direction. The plurality of artificial satellites include eight or more artificial satellites for each orbital plane. Each of the artificial satellites includes a fore-and-aft communication device. For each orbital plane, each of the artificial satellites forms a communication network with front and rear artificial satellites by the fore-and-aft communication device. Each of the artificial satellites on each orbital plane crosses southern and northern edges of the orbital plane in synchronization with artificial satellites on other orbital planes, and forms a communication network with that artificial satellites by the fore-and-aft communication device.

The present application relates to the technical field of spacecraft attitude orbit control, and particularly relates to a kind of coplanar installation electric thruster autonomous position keeping control method and device.Method includes: when receiving autonomous position keeping control instruction, based on current star time and pre-determined position keeping control period, the orbit inclination vector variation in position keeping control period is determined;Based on the orbit inclination vector variation in position keeping control period, the total firing duration in position keeping control period is determined to determine the firing duration corresponding to each operating day;Based on the orbit inclination vector variation in position keeping control period, the firing point right ascension is determined;Based on the firing duration corresponding to each operating day and the firing point right ascension, the coplanar installation electric thruster is used to autonomously carry out position keeping control in each operating day.This scheme can predict the position keeping parameters of multiple days, and the calculation is simple and easy to implement.

The application discloses a rotating tether system maneuvering method for out-of-plane inter-orbit load transmission, which systematically designs a load throwing mechanism by constructing a double-body rotating tether system framework. Firstly, a load release mode is planned based on the characteristics of the double-body rotating tether system, and then the required velocity increment and position increment parameters for realizing the release mode are derived. By calculating the total mass and length of the tether, and combining the relative position vector and velocity vector between the end-body satellites required for load throwing, the key parameters of the rotating plane for throwing are accurately calculated. Finally, the rotating plane maneuvering control strategy for the whole task cycle is constructed based on the above parameters. Compared with the traditional coplanar orbit transfer scheme, the application breaks through the out-of-plane orbit transmission task of effectively loading from the initial inclined low earth orbit to the target equatorial geostationary orbit. In the orbit inclination adjustment maneuvering task in the satellite launching stage, it has significant technical reference value and engineering application prospect.

The invention discloses a calculation method for determining a satellite and a satellite visible time period through TLE data, and relates to the technical field of satellite communication and inter-satellite link management. In a low-orbit constellation networking scene, through joint analysis of an orbit inclination angle difference and a BSTAR parameter, a scaling factor evolved along with time is generated, a communication distance threshold value is corrected in real time, and the communication time period of the satellite and the satellite visible time period is determined. Error transmission in the track direction is restrained; a dynamic tolerance model for sensing a relative motion state is introduced, so that the step length is automatically refined in a satellite high-speed approaching stage in a window search process, a millisecond-level visible interval is accurately captured, and meanwhile, missing detection in high-dynamic relative motion is avoided; in addition, for the problem of boundary jitter in a visible state, a stable output strategy of hysteresis filtering and double-threshold linkage is adopted; the dynamic tolerance is linearly mapped to the hysteresis bias, a self-adaptive buffer zone entering / exiting the threshold is formed, it is guaranteed that transient distance fluctuation does not trigger state jump, and a continuous and complete visible time period is output.

The application discloses a kind of through TLE data determination satellite and satellite visible time period calculation method, it is related to satellite communication and inter-satellite link management technical field, in low-orbit constellation networking scene, the application is generated by the joint analysis of orbit inclination difference and BSTAR parameter, scaling factor evolving with time, real-time correction communication distance threshold, inhibit error transmission along the direction of trajectory;Introduce the dynamic tolerance model of relative motion state perception, make the window search process automatically refine step in the high-speed approach phase of satellite, accurately capture millisecond-level visible interval, while avoiding missing detection in high dynamic relative motion;In addition, for the visible state boundary jitter problem, adopt the stabilization output strategy of hysteresis filtering and double threshold linkage;By linear mapping dynamic tolerance to hysteresis bias, form the adaptive buffer zone of entry / exit threshold, ensure that short distance fluctuation does not trigger state jump, output continuous complete visible time period.

This invention provides a polar-orbiting satellite navigation constellation system, a navigation method, and a user terminal. The polar-orbiting satellite navigation constellation system includes: four orbital planes, each with its intersection point with the equator evenly distributed along the equator, with a longitude interval of 45 degrees between adjacent orbital planes; nine satellites evenly distributed on each orbital plane, with a phase interval of 40 degrees between adjacent satellites; each satellite orbits at an altitude of 80,000–90,000 km and an orbital inclination of 90 degrees. It also employs an isotope power supply system and a redundant cold / hot backup system in the attitude and orbit control system. This invention utilizes a high-orbit scheme that avoids the influence of the Van Allen radiation belts, effectively solving the radiation effects caused by the Earth's radiation belts and increasing the reliability of the navigation system; the use of an isotope power supply system improves the operational lifespan of the navigation system; and the cold / hot redundant backup of the attitude and orbit control system enhances the reliability of the control system.

The invention provides a giant constellation multi-objective optimization design method based on knowledge migration, and the method comprises the steps: firstly building a giant constellation system key subject analysis model which is used for providing a model basis for the subsequent giant constellation analysis; constructing a mathematical optimization model which takes the orbit height, the orbit inclination angle, the orbital plane number and the like as design variables, considers constellation ground coverage constraints, and realizes the minimization of the total mass of the giant constellation system and the maximization of the inter-satellite capacity; on the basis, in order to avoid the problem of cold start in the design process, a population screening strategy and a sequence sampling mechanism are customized, a multi-objective optimization design method based on knowledge migration is constructed to optimize pre-selected design variables, and a design scheme that the total mass of a constellation system is as small as possible and the inter-satellite capacity is as large as possible is obtained. And then the giant constellation scheme design is completed. According to the method, the optimization efficiency of the overall design stage can be improved, the design period can be shortened, and the design, modification and optimization of the giant constellation scheme can be quickly completed.

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 invention relates to a fuel-saving double-pulse transfer method and device between space orbits and computer equipment, the method expands a fuel-saving double-pulse solution between elliptical orbits in a plane to the space elliptical orbits, and provides a continuation hybrid optimization algorithm for solving by continuously changing the inclination angles of initial and final orbits. The continuation hybrid optimization algorithm comprises the following steps: constructing an initial auxiliary solution, and constructing a discrete continuation sequence of which the orbit inclination angle coefficient is changed from 0 to 1; in each step, the optimal solution corresponding to the orbit inclination angle coefficient in the previous step is used as an initial value, and optimization is carried out by adopting an adaptive step length, so that a fuel-saving two-pulse solution meeting constraint conditions between space orbits is obtained. According to the method, the accuracy and effectiveness of solving the fuel-saving two-pulse orbit transfer problem between the space orbits are improved.

A satellite constellation includes a plurality of artificial satellites flying on an inclined circular orbit for each of six or more orbital planes having a common orbital inclination and arranged so that south and north axes are mutually shifted in an east-west direction. The plurality of artificial satellites include eight or more artificial satellites for each orbital plane. Each of the artificial satellites includes a fore-and-aft communication device. For each orbital plane, each of the artificial satellites forms a communication network with front and rear artificial satellites by the fore-and-aft communication device. Each of the artificial satellites on each orbital plane crosses southern and northern edges of the orbital plane in synchronization with artificial satellites on other orbital planes, and forms a communication network with that artificial satellites by the fore-and-aft communication device.

This invention discloses an optimization method for low-thrust transfer between circular orbits based on piecewise constant yaw angle guidance, belonging to the field of spacecraft orbit control technology. The implementation method of this invention is as follows: the low-thrust transfer process is divided into a transfer segment, a drift segment, and an adjustment segment; a piecewise constant yaw angle guidance strategy is proposed, using latitude argument angles α, β, γ as switching points, dividing each orbital period into four arc segments, maintaining a constant yaw angle and thrust switching state within each arc segment; based on the Gaussian perturbation equation, analytical solutions are derived for the long-term changes of the orbital semi-major axis, orbital inclination, and right ascension of the ascending node under this strategy and perturbation influence; the parameters of the optimal drift orbit are estimated based on the impulse assumption; a mixed-integer nonlinear programming model is established with the objective of minimizing the total velocity increment, and based on the dynamic relationship that "inclination changes are only generated by the thrust of arc segment 1," the model is transformed into several nonlinear programming sub-problems for solution, obtaining near-optimal guidance parameters. This invention avoids tedious numerical integration, significantly improves the efficiency of orbit design and optimization, and the resulting transfer scheme has fuel consumption close to the theoretical optimum. Moreover, the algorithm is robust and suitable for rapid orbit design tasks such as large-scale constellation deployment.

The invention discloses a low-orbit communication satellite constellation design method and device, and belongs to the technical field of satellite constellations, and the method comprises the steps: obtaining a target coverage latitude range, a main constellation orbit eccentricity rate and a ground receiver elevation angle range; calculating a geocentric angle half-angle range of a single satellite according to an elevation angle range of a ground receiver, and calculating a total satellite number range of a main constellation; determining the orbit inclination angle of the main constellation according to the orbit eccentricity rate of the main constellation; the surface area of the target coverage latitude range is calculated according to the target coverage latitude range, the ground coverage area of a single satellite is calculated according to the earth center angle half-angle range, and then the total satellite number range of the supplementary constellation is calculated; calculating an orbit inclination angle range of the supplementary constellation according to the target coverage latitude range and the geocentric angle half-angle range; and designing a low-orbit communication satellite constellation according to the total satellite number range and the orbit inclination angle range of the main constellation and the supplementary constellation. According to the invention, the problem of low constellation coverage of communication satellites in the prior art can be solved.