68 results about "Orbital inclination" patented technology

The method serves to place a space vehicle, such as a satellite, on a target orbit such as the orbit adapted to normal operation of the space vehicle and starting from an elliptical initial orbit that is significantly different from, and in particular more eccentric than the target orbit. The space vehicle is caused to describe a spiral trajectory made up of a plurality of intermediate orbits while a set of high specific impulse thrusters mounted on the space vehicle are fired continuously and without interruption, thereby causing the spiral trajectory to vary so that on each successive revolution, at least during a first stage of the maneuver, perigee altitude increases, apogee altitude varies in a desired direction, and any difference in inclination between the intermediate orbit and the target orbit is decreased, after which, at least during a second stage of the maneuver, changes in perigee altitude and in apogee altitude are controlled individually in predetermined constant directions, while any difference in inclination between the intermediate orbit and the target orbit continues to be reduced until the apogee altitude, the perigee altitude, and the orbital inclination of an intermediate orbit of the space vehicle have substantially the values of the target orbit.

The invention discloses an efficient multi-objective optimization method for a satellite constellation and belongs to the field of constellation system of a spacecraft. The method comprises the stepsof determining an initial condition based on a Walker-delta constellation configuration, and establishing a constellation orbit kinetic equation, an earth coverage analysis module and an earth observation resolution model; optimizing an earth orbit height, an orbital inclination and an ascending node right ascension with a coverage percentage and a ground pixel resolution by use of a sequential radial basis function multi-objective optimization strategy; and constructing an objective function based on l2 weighting and an improved Pareto fitness function, replacing a high time-consuming constellation performance simulation model with an RBF (Radial Basis Function) agent model to optimize design, and updating and managing the RBF agent model through sequence sampling in an interest interval.Therefore, a Pareto non-inferior solution set satisfying engineering requirements is acquired as a satellite constellation design scheme, the coverage percentage of a constellation for a target observation region is realized as high as possible, the pixel resolution of an effective load is realized as low as possible, the calculation cost and the design cost of the satellite constellation are reduced, and the Pareto leading edge search capability is improved.

The invention relates to a gravity satellite formation orbital stability optimization design and an earth gravity field precision inversion method on the basis of disturbing inter-satellite distance principle, in particular to an orbital stability optimization design method for a four-satellite system (FSS). In order to guarantee the stability of the four-satellite system, the quantity of satellite orbits is optimally designed, orbital semi-major axes, orbital eccentricities, orbital inclinations and right ascensions of ascending nodes keep unchanged, the difference of arguments of perigees each pair of satellites and the difference of mean anomalies of the pair of satellites are 180 degrees respectively, an initial argument of perigee of each satellite is arranged at the equator, an initial mean anomaly of each satellite is arranged at a pole, and the ratio of the semi-major axis of each elliptical orbit of the four-satellite system to a semi-minor axis of the elliptical orbit of the four-satellite system is 2:1. The gravity satellite formation orbital stability optimization design and the earth gravity field precision inversion method have the advantages that an earth gravity field is precisely and quickly inverted on the basis of a disturbing inter-satellite distance process; and the orbits are high in stability owing to the method, the earth gravity field computation precision is effectively improved, the gravity field inversion speed is increased to a great extent, and requirements on the performance of a computer are low.

The invention provides an optimization design method of a common track emergency reconnaissance constellation based on a sun synchronous return orbit. Firstly, the longitude and latitude coordinates of a target center position, the satellite overhead time and the reconnaissance frequency requirements are inputted. Secondly, the orbital altitude, eccentricity and orbital inclination of the satellite are calculated. Thirdly, the Greenwich star time angle is calculated according to the satellite overhead time. Then, the true nearest point angle, the right ascension of the reference satellite andthe number of satellites in the constellation are calculated. Then, the right ascension of each satellite in the constellation is optimized according to the right ascension of the ascending intersection point of the reference satellite, the true nearest point angle and the reconnaissance frequency. Finally, the orbital altitude, eccentricity, orbital inclination, perigee amplitude, right ascensionand true proximity of each satellite in the constellation are output. By using the method, the emergency reconnaissance constellation which meets the requirement of target area coverage can be designed quickly.

The invention discloses an application of low-thrust control in geostationary satellite orbit inclination maintenance. A simulation procedure includes the steps that an initial satellite status vectorand an estimated thruster work period are input, an objective function and a constraint function are written according to an inclination control objective and an optimal objective, numerical integration calculation is conducted, the optimal thruster work period is solved according to a parameter optimization method, and a post-control objective and an optimal control vector are output. The application is scientific and reasonable, use is safe and convenient, adaptability analysis is conducted through establishment of an inclination maintenance optimized mathematical model based on three methods of the minimum principle, phase plane control and parameter optimization, simulation verification is conducted on the parameter optimization method, a simulation result indicates that a low-thrustcontrol strategy meets an orbit inclination maintenance requirement, the low-thrust control strategy can be used as a reference in following satellite management, and the simulation result indicates that a low-thrust optimization control method is adaptable in geostationary satellite inclination maintenance.

The invention discloses a calculation method of subsatellite points and photographic point trajectory self-intersection points of a near-earth regression orbit satellite. Based on an orbital perturbation model of the earth oblateness (J2), the mean influence of the earth oblateness on the long-term perturbation of the right ascension of the ascending node, the perigee argument and the aplanatic angle is utilized, so that the mean drift period and the orbital intersection period of the satellite orbital plane relative to the earth resonate, and revisit regression of the orbital relative to thesubsatellite point and the photographic point trajectory of the earth surface is formed. After the orbit inclination angle, the eccentricity ratio, the regression days and the period number are given,the orbit semi-major axis can be solved, then the longitude of the sub-satellite point trajectory self-intersection point is analyzed and solved, the latitude of the self-intersection point under thecorresponding precision is screened out through numerical screening, and the distribution situation of the sub-satellite point trajectory self-intersection point in one regression period can be obtained. The invention provides a practical near-earth regression orbit sub-satellite point trajectory self-intersection point calculation method, and the method has important application value in near-earth surveying and mapping, remote sensing, reconnaissance, InSAR and other regression orbit satellite measurement tasks.

The embodiment of the invention relates to an observation method and system for minute-level rapid traversal of a high-orbit target, and the method comprises the steps: setting the observation capability of an observation system according to the combination of an observation region and a coverage demand for an observation target; sending the observation system to a quasi GEO orbit as an initial observation point; constructing a multi-satellite observation system with the total number of satellites being N by taking the initial observation point as the standard, wherein the system is a constellation evenly distributed on the quasi GEO orbit; if the track inclination angle i is equal to 0 degree, completing the system construction; and if the inclination angle i is not equal to 0, carrying out adaptive maneuvering of the platform attitude, so that the visual axis of each observation point points to a 0 degree synchronous belt area. The observation view field of the multi-satellite observation system forms a gap-free observation screen covering the GEO zone by 360 degrees in real time in the GEO area, the track of the GEO target can be completely covered by the gap-free observation screen, and the observation method enables the traversal observation period of high-orbit space debris and the target to enter the minute-level magnitude.

The invention discloses a nonsingular orbital elements-based 14 parameter broadcast ephemeris satellite position estimating method. The method can be applied to satellite tracks small in eccentricityratios and track inclination angles; via adoption of the method, MEO, IGSO and GEO satellite broadcast ephemeris user algorithms can be unified. The method comprises the following steps: a new secondtype nonsingular orbital elements-based 14 parameter broadcast ephemeris model is built; based on the model, a new broadcast ephemeris satellite position estimating method is put forward; compared with technologies of the prior art, the method disclosed in the invention is advantageous in that ephemeris parameter are reduced by 2, and navigation satellite communication resources can be saved. In the model put forward in the invention, eccentricity and track inclination angle parameters are directly adopted; a problem of matrix singularity which occurs during broadcast ephemeris parameter fitting processes when satellites run in orbits small in eccentricity ratios and track inclination angles, and the method can be applied to the satellite tracks small in eccentricity ratios and track inclination angles; MEO, IGSO and GEO satellite broadcast ephemeris user algorithm models can be unified.

The invention discloses a simple orbit extrapolation method on a satellite. The method comprises the following steps of determining the type of orbit elements according to the orbit determination result of a measurement and control system and with the combination of the orbit inclination angle and the eccentricity ratio characteristics; 2, performing orbit extrapolation by using the measurement and control system and taking the orbit determination result as an input, and outputting satellite positions and velocity vectors obtained through extrapolation according to the type of the orbit elements to obtain respective change sequences of the six orbit elements; determining the orbit fitting coefficients according to the respective change sequences of the orbit elements; and transmitting theground orbit fitting coefficients, the type of the orbit elements and the epoch t<0> to the satellite through a remote control instruction, and realizing simple orbit extrapolation on the satellite according to the calculated orbit elements at the time after the epoch t<0> on the satellite. The method greatly reduces the calculation amount on the satellite, reduces the complexity of an orbit extrapolation model, achieves the moderate orbit precision and meets the precision requirement.

The invention provides a small geostationary orbit satellite orbit transfer method and system. The method is characterized in that a small geostationary orbit satellite is transferred from an earth transfer orbit to a geostationary orbit through three stages, and specifically comprises the following steps: in the first stage, an orbit transfer chemical thruster ignites, and orbit transfer is executed by taking rapid lifting of the perigee height of the orbit to be higher than a first height and taking a domestic measurement and control arc section not smaller than a threshold measurement and control arc section as a control target; in the second stage, the electric thruster adjusts the track inclination angle to the threshold inclination angle while adjusting the track perigee height to the second height, and track transfer is executed with the track eccentricity ratio adjusted to the threshold eccentricity ratio as the control target; and in the third stage, the attitude control chemical thruster is ignited to enter the target orbit position precision range to serve as a target to execute fixed-point capture.

The invention belongs to the technical field of satellite network design, and discloses a Walker constellation-oriented constellation topological configuration characterization method and system and application. According to the implementation scheme, T/P/F parameters, orbit height and orbit inclination angle of Walker are initialized, and all satellites are numbered according to positions; calculating a value set of a distortion coefficient U of the constellation according to Walker constellation parameters and link establishment constraints of an inter-satellite link; selecting a distortioncoefficient U from the value set, and designing an adjacent matrix corresponding to Walker constellation topological configuration. T/P/F/U is used for representing a constellation topological structure formed by a Walker constellation of T/P/F according to the distortion coefficient U. The invention provides a concept of a Walker constellation distortion coefficient, so as to distinguish different inter-satellite network topologies formed by Walker constellations with parameters of T/P/F, describe the space segment information bearing capacity of the Walker constellation, and help to design aWalker constellation with comprehensively optimal coverage and communication capacity.

The invention provides a method for calculating the position of a pseudo-satellite, comprising steps of: determining whether it is a pseudo-satellite system; if so, reading a coordinate constant of the pseudo-satellite system; calculating the normalized time; calculating the average angular velocity of the satellite; separately calculating the mean anomaly, the eccentric anomaly, the true anomaly,the ascending node angular distance, and the perturbation correction term at the signal transmission moment; calculating the ascending node angular distance, the satellite vector length and the orbital inclination angle after the perturbation correction; calculating the position of the pseudo-satellite in an orbital plane; calculating the ascending node right ascension at the signal transmissionmoment; calculating the position of the pseudo-satellite in a earth-fixed coordinate system. The method only modifies the basic geodetic parameter in a satellite position algorithm, so that the pseudo-satellite spatial position can be calculated by using four ephemeris parameters, and solves problem that the original broadcast ephemeris cannot represent the coordinates of a ground stationary object.

The invention discloses a formation configuration reconstruction control method based on J2 perturbation active utilization. The method comprises the following steps: calculating configuration parameter deviation in a formation plane; when it is judged that the angle change direction of the relative eccentricity rate vector before and after reconstruction is the same as the rotation direction of the relative eccentricity rate vector caused by J2 perturbation influence, the amplitude of the relative eccentricity rate vector is controlled, and angle change is completed based on J2 perturbation; when it is judged that the angle change direction of the relative eccentricity rate vector before and after reconstruction is opposite to the rotation direction of the relative eccentricity rate vector caused by J2 perturbation influence, in-plane configuration reconstruction is completed with time as priority; determining a strategy for performing fuel-saving out-of-plane configuration reconstruction by using J2 perturbation; and pulse control causes a master-slave satellite orbit inclination angle difference of active bias, a relative inclination angle vector y-axis component is changed based on J2 perturbation, and out-of-plane configuration reconstruction is completed. By using the invention, formation configuration reconstruction control which saves more fuel consumption than traditional analysis pulse reconstruction is realized. The method can be widely applied to the field of satellite formation configuration control.

An airship is disclosed that stores air and uses it in a propulsion system that requires oxygen. This allows the airship to travel above the atmosphere, where oxygen is not available. The airship maximizes known mature technologies such as the buoyancy of a lighter than air balloon and the aerodynamic lift and propulsion of an airplane, until the altitude renders these technologies ineffective. The modified airship uses the stored air to extend transportation operations above natural limits and permit high-speed travel beyond traditional atmospheric speeds. Entering the edge of space at high tangential velocity permits placement of payloads into orbit. Entering the edge of space at low tangential velocity presents opportunities for low energy plane changes in orbital inclination. The modified airship may also alter the aerodynamic shape as desired in flight. The airship takes off, leaves and re-enters the atmosphere and lands without specialized launch facilities or long landing strips. The reusable vehicle provides transportation services from point to point on the earth surface and from the earth to points in space.

The invention provides a space-based radar efficiency optimization method. The method includes the following steps: determining an inclination angle of a track on the basis of a latitude zone of a target area and the loading distance of space-based radar; calculating the height of the track according to the loading distance of space-based radar and the inclination angle of the track; and arranging multiple space-based radar satellites on the same track surface by enabling the phase angles of the satellites to be equal or arranging the space-based radar satellites on N track surfaces. A zero-degree inclination angle of a non-zero-degree inclination angle of a track can be selected on the basis of a latitude zone of a target area and the loading distance of space-based radar, that is to say, a track with a low inclination angle can be selected. Therefore, the target area can be detected many times and the high system efficiency can be achieved by means of relatively less satellites, and a high temporal resolution detection requirement of an airplane target can be met.

The invention discloses an orbit planning method for a sun-synchronous circular regression orbit. The design of the sun-synchronous circular regression orbit needs to meet two index requirements of ascending node right ascension derivative constraint and regression coefficient constraint of the sun-synchronous orbit. According to the two index requirements, the two important parameters of the orbit semi-major axis and the orbit inclination angle of the sun-synchronous circular regression orbit can be obtained through iterative calculation; the position of a launching point near the central position of a launching area is traversed and selected according to the calculated semi-major axis of the orbit, the calculated inclination angle of the orbit, the central position of the designated launching area and other parameters; and deviation distribution of subsatellite point geocentric longitude and target point geocentric longitude are calculated when the sun-synchronous circular regression orbit subsatellite point passes through the target latitude circle for multiple times corresponding to different launching point positions, the orbit design with the minimum deviation value is selected to carry out threshold judgment, and the effective orbit design is stored. According to the invention, rapid and effective detection of the target area can be realized.

A satellite orbit forecasting method based on an artificial neural network algorithm comprises the steps: reading TLE two-line orbit data containing six elements, and generating a simulation curve; processing the discontinuous curve by adopting a splicing method to form a continuous curve; on the basis of the six elements, adding a TLE data acquisition timestamp T to construct seven elements of TLE prediction suitable for TLE data prediction; training the orbit inclination angle i, the orbit eccentricity e and the average motion n in the six elements of the orbit by adopting a prediction algorithm based on LSTM; performing training by adopting an element prediction algorithm based on linear regression to obtain an optimal function; and repeating the training and correcting to obtain an executable prediction model. When the method provided by the invention is used for forecasting a satellite orbit, the overall position error precision of the model can be controlled within 5km in a short period and a middle period, and the overall position error precision of the model can be controlled within 23km in a long-term forecasting process.

The invention discloses a method and device for determining a moon-earth transfer orbit. The method comprises the steps of according to the approximate reentry time of a detector reentering a preset reentry point of the earth atmosphere from the moon, a preset orbit inclination angle when the detector arrives at the preset reentry point, the latitude and longitude of a drop point, and the moon-earth transfer time of the detector, determining a reentry time initial value of a preset reentry point; determining a moon-earth transfer orbit number of the preset reentry point, and determining a moon-earth transfer orbit under a two-body model according to the reentry time initial value and the moon-earth transfer orbit number; correcting the moon-earth transfer orbit under the two-body model to obtain a target moon-earth transfer orbit under the high-precision model; according to a preset intersection line strategy corresponding to a moon-earth transfer orbit and a moon-surrounding orbit, correcting a perilune height and a lunar section moon-earth transfer inclination angle in the target moon-earth transfer orbit under the high-precision model; and when the corrected lunar section moon-earth transfer inclination angle enables the perilune of the moon-earth transfer orbit to be in the lunar orbit plane, determining moon-earth incidence control parameters.

The invention discloses a constellation design method, and the method comprises the steps: determining the area of an observed fixed region, and determining a corresponding circular coverage region when a single satellite covers the observed fixed region according to the area; determining a view field angle or a beam angle of the observed fixed region according to a preset view field pointing constraint condition, and calculating the orbit height of the single satellite according to the circular coverage region and the view field angle or the beam angle; determining the number of satellites required for covering one latitude zone region, an orbit inclination angle and a difference value of true near-point angles of two adjacent satellites in the same orbit plane according to the circular coverage region, and determining the number of orbit planes according to the number of satellites; and calculating the number of satellites in the same orbital plane according to the difference, calculating the total number of satellites according to the number of orbital planes and the number of satellites in the same orbital plane, and designing a satellite constellation. The technical problem that in the prior art, constellations are not suitable for serving ground fixed areas and observing fixed areas in celestial spheres is solved.

The invention provides a Mars global remote sensing orbit design method and system based on near fire point drift, and relates to the technical field of orbit design of deep space exploration spacecrafts, and the method comprises the steps: S1, determining a near fire point height upper limit, and enabling the near fire point height at the end of life to be greater than the fire atmosphere height; S2, calculating a track inclination angle adjusting range; S3, calculating the argument change rate of the near fire point; S4, calculating the right ascension change rate of the ascending node; and S5, analyzing and calculating orbit semi-major axis parameters according to parameters of a near fire point argument change rate and an ascending node right ascension change rate required by a remote sensing detection task. According to the invention, the requirements of low-orbit target imaging and high-orbit ground communication can be met; the emission weight of the deep space detector is saved; and the engineering task requirements of global coverage, full-range coverage detection of different orbit heights and detailed detection of key areas can be met.

The embodiment of the invention discloses a method for realizing beam-to-ground coverage through low-orbit satellite beam position design. The method comprises the following steps: S10, acquiring an orbit height, an orbit inclination angle, an equivalent half-cone angle of a ground coverage area and a spot beam half-cone angle of a low-orbit communication satellite; S12, according to the orbit height, the orbit inclination angle, the equivalent half-cone angle of the ground coverage area and the spot beam half-cone angle of the low-orbit communication satellite, obtaining the number of spot beam positions for realizing equivalent coverage of the equivalent half-cone angle area; S14, arranging the spot beam positions according to an equilateral overlapping mode or a honeycomb overlapping mode; S16, calculating an azimuth angle and a pitch angle of a center direction vector of a spot beam corresponding to each arranged spot beam position on a given reference coordinate system; S18, driving an antenna of the satellite to enable a center direction vector of a spot beam emitted by the antenna of the satellite to point to the earth according to the calculated azimuth angle and pitch angle.

A satellite constellation (200) comprises three artificial satellites (210A to 210C) that monitor a target area of the Earth (101). Each artificial satellite circulates on elliptical orbits having sun-synchronization and an orbit inclination angle. A long axis of each elliptical orbit forms an equal angle with each long axis of two adjacent elliptical orbits in a latitude direction.