46 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 present invention discloses an aerodynamic layout of a large-lift-surface ultra-low-orbit aircraft, which includes an aircraft body, a group of aerodynamic sails, and a group of solar cell wings. The body of the aircraft provides a space for equipment installation, and its cross-section is an equilateral triangle structure with a side length of W and a length of L. The large-lift-surface aerodynamic sails are located directly below the aircraft body, providing the main mechanical load and the area for solar cell placement. The solar cell wings are attached to the upper surface of the aerodynamic sails. The designed large-lift-surface aerodynamic layout can effectively increase the lift-to-drag ratio of the aircraft under the design requirements of a fixed antenna angle, and change the orbital plane of the aircraft through an aerodynamic assist method, so as to achieve a large-range orbital inclination maneuver for the aircraft in the ultra-low orbit with the minimum fuel consumption.

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 invention relates to the technical field of satellite communication, and provides a satellite network automatic generation method and system of a multi-shell satellite-ground-satellite dynamic virtual topological structure. The method comprises the following steps: acquiring TLE data to calculate parameters such as an orbit inclination angle and height of an on-orbit satellite, layering the on-orbit satellite and determining satellites of different shell layers; screening out networking satellites and backup satellites based on parameters of different shell satellites; after the networking satellites are logically numbered, inter-satellite links are constructed based on the parameters of the networking satellites of all the shell layers; the satellite ground stations are numbered, and a satellite-ground link is constructed based on the relative elevation angles of the satellite ground stations and the networking satellites; the network topology connection relationship of each time slice is determined according to the construction of the inter-satellite link and the satellite-ground link, the dynamic topology is converted into a static sequence by using a discretization means, and the dynamic topology structure of the satellite network is determined. The method can effectively adapt to the dynamic change characteristics of the satellite network, and improves the management efficiency of the satellite network.

A method for node sensitivity analysis affecting the degradation of the communication performance of a giant constellation. The present invention relates to the field of spacecraft system simulation. It solves the problem that in the prior art, there is a lack of a method for measuring the sensitivity of the constellation to changes in the number of satellites, which brings changes in performance such as communication delay. The method includes: reading the parameter values under different two-line element data segments according to the format of the two-line element data record and saving them to the SGP4 orbit model structure variable; dividing different sub-constellations within the giant constellation according to the orbital inclination or the semi-major axis of the orbit; setting that the satellites within each orbital plane form a routing domain, setting the gateway on / off condition, setting that the satellite nodes can communicate within the domain, and restricting the connectivity of the nodes in different domains; updating the connectivity between the nodes, outputting the on / off situation and delay between the nodes; and performing routing and outputting the routing result to complete the node sensitivity analysis. It is also applicable to the field of measuring the communication delay of giant constellations.

The invention relates to a pathfinder test satellite suitable for a deep space exploration flight mission, and the satellite orbit of the pathfinder test satellite is a non-near earth orbit which avoids the influence of a normal Airy radiation band, and is a low orbit inclination angle which is free from the influence of strong space radiation. The method has the beneficial effects that the satellite orbit is not influenced by the Nabelian radiation band and is consistent with the actual deep space orbit environment.

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 provides a satellite portable station hand-automatic integrated azimuth and elevation angle driving adjustment device, which comprises a supporting tripod, an adjustment seat, an azimuth shaft, an elevation shaft, a first driving mechanism, a second driving mechanism and an antenna pot surface, the azimuth shaft and the elevation shaft are assembled in the same adjustment seat, the traditional split type layout is changed, the installation space is greatly saved, the hollow lightweight design not only reduces the weight of the whole machine, meets the carrying requirements of the satellite portable station, but also protects the transmission core components by using the internal preset space, improves the adaptability of the outdoor environment, and the azimuth shaft and the elevation shaft work cooperatively in the main adjustment seat and do not interfere with each other, and the worm drive can realize 360 continuous horizontal rotation of the azimuth angle and 0-90 vertical rotation of the elevation angle, the independent double-shaft design ensures that the antenna beam can point to any airspace position, and meets the satellite pointing requirements of different orbit inclination satellites.

This invention discloses a multifunctional satellite constellation configuration design optimization method based on a large language model, belonging to the field of spacecraft system design technology. The method includes: setting design variables and fixed variables, where the design variables include the orbital inclination and orbital altitude of satellites in the satellite constellation; constructing an evaluation function based on coverage percentage, ground pixel resolution, average number of visible satellites, and geometric dilution of precision; iteratively generating solutions based on the design variables and fixed variables using the large language model, evaluating the solutions using the evaluation function, and obtaining an optimal satellite constellation configuration design scheme based on the evaluation results. Through the above technical solution, the present invention can provide a satellite constellation configuration design optimization method that takes into account multiple optimization aspects.

The invention discloses a rotating tether system maneuvering method for load transmission between different-plane orbits, and the method systematically designs a load throwing mechanism by constructing a double-body rotating tether system architecture: firstly, planning a load release mode based on the characteristics of a double-body rotating tether system; deducing speed increment and position increment parameters required for realizing the release mode; by calculating the total mass and length of the tether and combining the relative position vector and the velocity vector between the end body satellites required by load throwing, the key parameters of the throwing rotation surface are accurately calculated; and finally, constructing a rotating surface maneuvering control strategy of a full task period based on the parameters. Compared with a traditional coplanar orbit transfer scheme, the non-coplanar orbit transfer task of the effective load from the initial inclined low earth orbit to the target equator geostationary orbit is achieved in a breakthrough mode. The method has obvious technical reference value and engineering application prospect in orbit inclination angle adjustment maneuvering tasks in a satellite launching stage.

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 present invention discloses an orbital design method for large-range inclination change using multiple gravity assists. The orbital design method for large-range inclination change using multiple gravity assists has the following input information: assuming that the orbital inclination of the gravity assist small body around the planet is 0, at a certain moment, the spacecraft has achieved coincidence of its position with that of the gravity assist body through orbital control, the positions and velocities of the spacecraft and the small body relative to the planet are known, the orbital inclination of the spacecraft at this moment is i0, and the target inclination to be achieved is i f , then the inclination adjustment is achieved in the shortest time through multiple gravity assists. The present invention solves the problems in the prior art that there are too many variables to solve during the orbital inclination change process and large-range inclination change cannot be achieved.

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 present invention provides a method, device and electronic device for determining the inter-satellite distance of satellites, which obtains the first orbital inclination, the first latitude argument and the first ascending node right ascension of the first satellite, and obtains the second orbital inclination, the second latitude argument and the second ascending node right ascension of the second satellite; wherein the first satellite and the second satellite belong to the same satellite shell; based on the obtained information and the pre-constructed target distance expression, the maximum distance threshold and the minimum distance threshold between the first satellite and the second satellite are determined. In this method, it is only necessary to obtain the orbital inclination, latitude argument and ascending node right ascension corresponding to the first satellite and the second satellite, and the maximum distance threshold and the minimum distance threshold corresponding to the first satellite and the second satellite can be calculated by the target distance expression, and the calculation is simpler and more efficient.

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 invention provides an orbit design method for space-based observation of a low-orbit satellite constellation, belongs to the field of spacecraft orbit design, and solves the problem that a space-based observation platform is low in observation efficiency of the low-orbit satellite constellation. Comprising the following steps: combining an earth gravitational potential function with an earth non-spherical perturbative force function to obtain a J2 perturbative force function; a Kepler orbital element is used for representing a J2 perturbation function, and the J2 perturbation function is combined with a Lagrange type perturbation motion equation to construct a Lagrange perturbation model; a lagrangian perturbation model, the influence of J2 perturbation force on the number of low-orbit satellite orbits and the drift law of the right ascension of an ascending node are utilized to construct complementary angles of the orbit inclination angle of a space-based observation platform and the orbit inclination angle of a low-orbit satellite; the space-based observation platform is a retrograde orbit configuration or a plurality of retrograde orbit configurations, and the eccentricity rate and the semi-major axis of the space-based observation platform are the same as those of the low-orbit satellites, so that observation of all the low-orbit satellites in the constellation is completed; according to the invention, the observation efficiency of the space-based observation platform on the low-orbit satellite constellation is improved.

A satellite system may have a constellation of communications satellites that provides services to users with electronic devices such as portable electronic devices and home and office equipment. A network operations center may use gateways to communicate with the satellite constellation. The satellite constellation may include sets of satellites with different orbits such as circular orbits with different inclinations, sets of satellites with elliptic orbits, sets of satellites with circular orbits of different altitudes including low earth orbits, medium earth orbits, and / or geosynchronous orbits, and / or sets of satellites with other orbits. The satellite orbits of the satellites in the satellite constellation may be selected to provide coverage to desired user population concentrations at different locations on the earth without using an excessive number of satellites.

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 present invention relates to the field of electrical digital data processing, and provides a satellite visibility evaluation method, system, and electronic device. The method includes: obtaining a first visibility result of a target satellite, where the target satellite is at least one of a plurality of satellites used for shore-based satellite positioning measurement; determining a second visibility result of the target satellite based on a built simulation scenario, where the simulation scenario is built based on orbital parameters, and the orbital parameters include the semi-major axis of the orbit of the target satellite, the orbital inclination, the right ascension of the ascending node, and the argument of perigee; evaluating the first visibility result based on the second visibility result to determine the reliability of the first visibility result. To solve the defect in the prior art that in shore-based GNSS reflection measurement, the visibility of satellites cannot be clearly obtained, and thus reliable measurement results cannot be obtained. The solution of the present application can efficiently and accurately determine the visibility of satellites in shore-based GNSS reflection signal measurement under complex sea conditions.

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.

The present invention provides a global carbon inventory satellite, comprising: a frozen orbit unit, configured to set an orbital inclination, so that the global carbon inventory satellite operates on a frozen orbit, with the apogee of the frozen orbit frozen above the latitude of an area with intensive human activities; and a sun-synchronous orbit unit, configured to enable the global carbon inventory satellite to continue operating on the sun-synchronous orbit, so that the global carbon inventory satellite is always in an illuminated area when it operates to the apogee; wherein the relationship between the perigee orbit altitude and the apogee orbit altitude is obtained based on the value of the ascending node of the sun-synchronous orbit, the value of the orbital inclination, a first function, and a second function; the first function represents the relationship between the semi-major axis of the orbit and the perigee orbit altitude and the apogee orbit altitude; the second function represents the relationship between the orbital eccentricity and the perigee orbit altitude and the apogee orbit altitude.

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.

The invention discloses a gravity satellite constellation design method for eliminating sea tide mixing, and relates to the technical field of satellite gravity. Comprising the following steps: calculating an ascending node angular distance difference of each sea tide component; determining the satellite orbital plane number, the orbital height, the orbital inclination angle and the inversion period of the gravity satellite constellation; determining a sea tide frequency mixing control strategy; and determining an optimal satellite constellation design scheme according to the sea tide mixing elimination strategy. According to the invention, a sea tide mixing control strategy is formulated according to the sea tide mixing characteristic of each sea tide component, and a practical method is provided for eliminating the influence of sea tide mixing on the inversion gravity field by using satellite constellation design.

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.