353 results about "Orbit determination" patented technology

The invention relates to a single frequency high precision positioning method based on GNSS. The method comprises steps that 1, preparation work is carried out; 2, initial epoch position solution is carried out; 3, a carrier wave phase epoch difference observation model is constructed, and a relative position variance is solved; 4, the fuzziness information and the position information are transmitted; 5, the fuzziness information and the variance information are adjusted; 6, a fuzziness and position information virtual observation equation is constructed; and 7, least square estimation for a GRAPHIC combined observation equation and a virtual observation equation is commonly carried out, and a final positioning result is solved. Through the method, a GRAPHIC combined kinetics positioning algorithm aided by carrier wave phase epoch difference is utilized, kinetics navigation positioning precision is improved, precise ephemeris energy is employed to carry out high precise orbit determination for LEO satellites, and broadcast ephemeris energy is employed to realize real-time navigation positioning. The method is advantaged in that the method is not influenced by satellite change in a navigation positioning process, high reliability, low cost and simple data processing are realized, and the method is suitable for popularization and application.

The invention discloses a combined scheduling method of satellite-ground measurement and control resources of a low-mid-orbit satellite constellation based on orbit determination constraint satisfaction, and belongs to the field of combined scheduling of the satellite-ground measurement and control resources. The method comprises the steps of (1) building a visible geometry model, (2) building a visible link model, (3) building the constraint condition of orbit determination, (4) carrying out preprocessing, (5) determining a resource distribution principle, (6) building a combined scheduling model of the satellite-ground measurement and control resources, and (7) carrying out algorithm solution. Compared with the prior art, the method has the advantages that the scheduling result is reasonable, the scheduling process is efficient, and the scheduling method is high in universality. The air-and-ground-based integrated scheduling of the measurement and control resources is achieved. The transition from ground-based measurement and control to air-and-ground-based combined measurement and control in China is achieved.

The invention relates to a precise orbit determination method of a navigation satellite for assisting a clock error between stations, which is characterized by comprising the following steps: configuring an atomic clock and a nanosecond stage time synchronization system at each station of an area network, receiving a distance measurement signal of a navigation satellite by each station of the area network to obtain the carrier phase data of each station, detecting the periodic trip generation moment of the carrier phase data by a Blewitt method and carrying out time synchronization between the stations by a two-way satellite time frequency transmission method TWSTFT to obtain a clock error between the stations; resolving a differential equation under an inertial coordinate system to obtain an orbit of the navigation satellite. Before orbit determination data are processed, the method firstly realizes the time synchronization between the stations with high accuracy (i.e.: obtaining the clock error between the stations). During the processing of the orbit determination data, an improved non-error method is adopted, only the orbit needs to be resolved, and the clock error between the stations does not need to be resolved, which not only can improve geometrical factors but also is beneficial to separating a system error (the clock error between the stations). Thus, the orbit determination accuracy can be enhanced.

The invention discloses a space-based phased-array radar space multi-target orbit determination method, which mainly solves the problems that the space weak target cannot be evaluated effectively and the target orbit determination precision is low in the prior art. The method comprises the following steps of: processing target echo data by a zero-setting conformal algorithm, acquiring distance prior information by utilizing a distance pulse compression principle, and segmentally processing echo signals according to the distance prior information; segmenting the echo signals according to the number of the targets, wherein each segment of data is a target signal and an adjacent unit signal; performing multi-target two-dimensional angle evaluation on each segment of data by utilizing a sum-difference multi-beam angle-measuring principle; performing coordinate conversion based on a space target tracking result and detection satellite orbit information; and performing orbit determination on different space targets by a Laplace type iterative algorithm, and improving the orbit determination precision by a least square algorithm. According to the method, the influence of strong signals on weak targets can be reduced and the parameters of the weak targets can be evaluated accurately. The method can be applied in the actual application fields of space situation awareness, orbit resource management and the like.

The invention discloses a geosynchronous orbit constellation orbit determination method based on ground station/satellite link/GNSS combined measurement. The method comprises the steps of (1) establishing the state equation of a GSO constellation orbit determination system in a centralized structure, (2) establishing the measurement equation of the GSO constellation orbit determination system in a centralized structure, (3) carrying out combined optimization of a ground station for orbit determination and GNSS measurement combination, (4) determining the filtering method for realizing orbit parameter estimation according to the above established GSO constellation combined orbit determination system model based on ground station/satellite link/GNSS, and (5) carrying out the concrete realization of a GSO constellation orbit determination method based on ground station/satellite link/GNSS in the centralized structure. According to the method, the combined distribution optimization strategy of ground station and GNSS measurement is established based on a precision factor, and the calculation amount is reduced while the orbit determination precision is ensured.

The invention discloses an autonomous navigation signal enhancement method based on low-earth-orbit satellites. The method is implemented by use of a load carried on the low-earth-orbit satellites, wherein the load comprises a GNSS receiving module, a main control module, an emission module, a high-stability clock module, at least one set of GNSS signal receiving antennas and at least one set of enhancement information emission antennas; the GNSS receiving module is used for capturing and tracking navigation satellite signals and performing on-satellite autonomous orbit determination and autonomous time service; the main control module is used for interacting data with the GNSS receiving module and the emission module; the emission module is used for coding result data of on-satellite autonomous orbit determination and autonomous time service into telegraph texts and performing signal modulation and broadcasting; and the high-stability clock module is used for providing a time-frequency benchmark. According to the method, by uniting low-earth-orbit satellite navigation enhancement information and the navigation satellite signals to perform navigation positioning, the precision, availability and reliability of navigation positioning can be effectively improved. Besides, the method does not depend on a ground station, so that application is more flexible compared with a ground-based enhancement technology.

The invention relates to a satellite orbit maintenance and control method based on two lines of radicals, which comprises the following steps of: calculating two lines of radical data disclosed by a satellite by utilizing a SGP4 orbit determination model, calculating and obtaining a geographical latitude and a geographical longitude of a sub-satellite point of the satellite and comparing with an agreed ground nominal track. Through taking the duration of a first sample opposite to an orbit epoch moment as an independent variable and taking the ground track distance difference as a variable, all samples are subjected to quadratic curve fitting by using an average method; the difference delta at between an actual value of a satellite orbit semi-major axis and a nominal value by a latest sample moment as well as the average attenuation rate of the satellite orbit semi-major axis within the time interval of the sample are calculated and obtained through a fitting polynomial coefficient; and finally, the next orbit control time Tc as well as the control quantity delta af required to be used are predicted and obtained by utilizing a fitted secondary curve and an allowable drift range [-delta Lmax and delta Lmax] to ensure that the position of an actual ground track within the satellite next drift period opposite to a ground nominal track is within the allowrable drift range.

The invention discloses an error correction method of an autonomous navigation system. In the autonomous navigation system based on an ultraviolet earth sensor and a star sensor, the ultraviolet earth sensor is used for measuring a geocentric vector in a satellite local system, and the star sensor is used for measuring an attitude transformation matrix of the satellite local system. Since the sensors inevitably have relative installation error, and a filter cannot eliminate the system error, the navigation accuracy becomes poor. In the method provided by the invention, the relative installation error of the ultraviolet earth sensor and the star sensor is modeled, and the installation error is expanded into a state variable so as to perform navigation filtering. Assuming that the satellite can obtain high-accuracy position measurement information by a GPS receiver or ground orbit determination within certain period of time, the system error can be estimated and corrected in time by filtering. The method provided by the invention has simple operation and can obviously improve the navigation accuracy.

The invention discloses a method for tracking a lunar probe by using an earth station. The method comprises the following steps of: performing orbit determination on the lunar probe to generate a precise orbital element for the lunar probe to run on a lunar orbit; calculating a synthetic perturbation ephemeris according to the precise orbital element for the lunar probe to run on the lunar orbit and outputting a calculation result to a synthetic perturbation ephemeris file; calculating earth shadow forecast, moon shadow forecast and ascending node forecast, calculating substellar point forecast of the lunar probe relative to earth and moon and calculating the substellar point forecast of the sun relative to earth and moon; calculating the position of the lunar probe relative to the earth station; calculating shelter of the moon and forecasting the condition that the lunar probe is sheltered by the moon; generating an earth station observation forecast file; and tracking the lunar probe by the earth station according to the observation forecast file. By the method, the measurement and control of the lunar probe and the receiving of probe data are tracked by using the earth station.

The invention belongs to the technical field of navigation and positioning and discloses a low-orbit navigation constellation enhanced network RTK (real-time kinematic) technology. The low-orbit navigation constellation enhanced network RTK technology comprises the following steps: design of a global navigation and positioning constellation and multiple low-orbit constellation schemes, a low-orbitsatellite collecting orbit determination and inter-satellite link autonomous orbit determination method, design of a low-orbit satellite broadcast ephemeris and a high, medium and low-orbit satellitefusion processing unified mathematical model. The problem in the prior art that the network RTK method is low in positioning precision is solved, the baseline distance can be increased, the network density can be reduced, the cost can be reduced, the precision and the ambiguity fixing rate are increased, and the convergence time can be shortened. The sub-satellite point track coverage range of the low-orbit constellation is wide and even covers polar regions, the three-dimensional orbit errors of 60 and 192-star constellation are 46 cm and 18 cm correspondingly, the forecast precision is increased by about 30 percent compared with that of a klobuchar model, and the final positioning precision is at horizontal millimeter and height centimeter level.

Provided are a method and system for determining a precise orbit of a LEO satellite. The method includes: estimating a precise ephemeris of a global positioning system (GPS) satellite by fitting an orbit perturbation-based GPS dynamics model to observation data of the GPS satellite received from a GPS observatory and estimating a precise ephemeris of a Galileo satellite by fitting an orbit perturbation-based Galileo dynamics model to observation data of the Galileo satellite received from a Galileo observatory; determining an initial orbit value of a LEO satellite by fitting an orbit perturbation-based LEO satellite's basic dynamics model to navigation data received from the LEO satellite; and determining the precise orbit of the LEO satellite by calculating a difference between observation values, which are calculated based on a GPS and Galileo data received from the LEO satellite, the GPS observatory and the Galileo observatory, and calculated values, which are calculated based on an orbit perturbation-based LEO satellite's dynamics model that was calculated using the initial orbit value of the LEO satellite and the precise ephemeris of the GPS and Galileo satellites. Since both the GPS and Galileo data are received and used to determine the precise orbit of a LEO satellite, more precise orbit determination can be achieved.

The invention discloses an earth-Lagrange combined constellation autonomous orbit determination method based on inter-satellite ranging. The method includes following steps: step 1, establishing a state equation of an earth-Lagrange combined constellation autonomous orbit determination system; step 2, establishing a measurement equation of the earth-Lagrange combined constellation autonomous orbit determination system; step 3, determining a filtering method for realizing orbit parameter estimation; step 4, specifically realizing the earth-Lagrange combined constellation autonomous orbit determination method based on a selected filtering algorithm. By introducing a Lagrange satellite, the problem of deficient range when only inter-satellite ranging information is utilized for autonomous orbit determination is solved effectively, and complexity of system equipment is lowered; statistical characteristics of system noise are estimated online in real time through a self-adaptive nonlinear filtering algorithm, so that requirements on noise prior information are low, stability of the autonomous orbit determination filtering algorithm is improved, and autonomous orbit determination accuracy is improved.

The invention relates to a low earth orbit satellite autonomous orbit determination method which comprises the following steps of selecting a plurality of high earth orbit objects (including large space debris) as calibrated satellites to acquire a precise ephemeris of the high earth orbit objects; when a low earth orbit satellite cannot acquire the precise ephemeris, utilizing a boarded space-based visible light camera to point to the high earth orbit calibrated satellites, performing tracking for a period of time to acquire an only angle measurement quantity; utilizing an on-board computer,combining rough ephemeris of the low earth orbit satellite, the precise ephemeris of the high earth orbit objects with the measurement quantity, performing reversed determination to obtain a precise orbit of the low earth orbit satellite. The invention provides a novel satellite autonomous orbit determination way; and when global positioning navigational satellite information cannot be utilized, the novel satellite autonomous orbit determination method can be used as an alternative method. The low earth orbit satellite autonomous orbit determination method provided by the invention is relatively simple in calculation model, fully utilizes space resources and meanwhile, is low in space-based visible light observation cost, small in energy consumption and strong in engineering applicability.

A method for realizing precise orbit determination of low-medium orbit satellite utilizes Galileo positioning system, satellite laser distance measurement technique and integrated precise orbit determination new means to optimize space-ground orbit determination system for realizing precise orbit determination of low-medium orbit satellite equipped with laser reflector.

The invention belongs to the field of autonomous orbit determination of satellites and relates to an autonomous orbit determination method of a navigation satellite constellation. The autonomous orbit determination method of the navigation satellite constellation comprises the following steps: (1) starting autonomous orbit determination and initializing a system; (2) acquiring an inter-satellite ranging between satellites; (3) calculating an observation matrix of the inter-satellite ranging; (4) forecasting an orbit by using an orbit dynamical model; (5) updating measurement and estimating state by using a kalman filtering algorithm; and (6) judging whether the autonomous orbit determination is ended, operating the steps (2)-(6) again if the autonomous orbit determination is not ended, and conversely, ending and exiting an autonomous orbit determination program. Compared with the conventional autonomous orbit determination of navigation satellite constellation, the autonomous orbit determination method of the navigation satellite constellation has the advantages that the method is capable of avoiding the rank defect problem of autonomous orbit determination method of the navigation satellite constellation by using the inter-satellite ranging and is high in long-term autonomous orbit determination accuracy.

The invention discloses an autonomous orbit determination method for a satellite based on synthetic aperture radar, which belongs to the technical field of autonomous orbit determination of satellites and aims at achieving high-precision and real-time autonomous orbit determination of a low-orbiting satellite without being supported by a ground tracking telemetry and command station. The method comprises the following specific steps of: designing the shape and the material of a manual ground identification point; designing an arrangement mode of the ground identification point, arranging the ground identification point and measuring the position of the ground identification point in an earth-fixed coordinate system; storing information of the ground identification point in a space-borne computer; and after a space-borne synthetic aperture radar remotely senses the ground and identifies a ground identification, deriving the information of the ground identification point from a ground identification library on the satellite, applying an orbit determination equation to obtain the position and the speed of the satellite and further finish the real-time autonomous orbit determination of the satellite. The autonomous orbit determination technology for the satellite provides a novel autonomous orbit determination method for the satellite, which can be used for realizing high-precision real-time autonomous orbit determination of the low-orbiting satellite. The method disclosed by the invention is suitable for real-time autonomous orbit determination of a near earth satellite.

The embodiment of the invention discloses a satellite precise orbit determination method and device. The method comprises the steps that first observation data of a navigational satellite, determinedby a ground receiver and second observation data of the navigational satellite, determined by a low orbit satellite receiver are acquired; a first observation equation of the ground receiver and a second observation equation of the low orbit satellite receiver are determined; and the first observation equation and the second observation equation are solved according to the first observation data and the second observation data, and accordingly the orbital position of a navigational satellite is determined.

The invention relates to a coorbital satellite autonomous relative position keeping method based on laser loads. On-satellite autonomous orbit determination is achieved through a navigation constellation or orbit determination is measured through a ground station, the satellite position initial reference is obtained, determination of the relative positions between two space satellites and time unification of systems are achieved by combining inter-satellite laser ranging and communication integrated loads with rotary table pointing information, a target constellation is resolved through the current position of each star of the constellation, the existing position information and the target position information are compared, the relative position keeping requirement of each star is determined, the enveloping and pose deformation requirements of each star are resolved, the change rate of a satellite orbit is changed by solar wing position adjustment and deformation and adjustment of the perturbation influence, and accordingly relative position keeping of coorbital satellites is achieved, all the stars in the constellation can run in a relative good form after the orbit is affected by perturbation, and the orbit decay influence is reduced.

The invention relates to a low-orbit navigation enhancement precision correction data generation and uploading system and method. A ground data processing central station combines monitoring data acquired by a GNSS (Global Navigation Satellite System) monitoring station and low-orbit satellite monitoring data acquired by a satellite measurement and control station for processing to generate SSR (State Space Representation) data; the generated SSR data is evaluated and then transmitted to three service data uploading stations, the three service data uploading stations respectively transmit theSSR data to three geostationary orbit communication satellites as high elliptical orbit satellites through an uploading channel, a low-orbit navigation enhancement satellite receives the SSR data broadcasted by the geostationary orbit communication satellites, and processes the uploaded SSR data by adopting a navigation enhancement load t generate enhanced navigation signals so as to complete high-precision orbit determination and low-orbit navigation enhancement information generation. Compared to the prior art, the method for generating and uploading the SSR correction data of the navigationsystem required by the low-orbit navigation enhancement system is achieved without depending on low-orbit inter-satellite links, so that the complexity of the low-orbit system is reduced, the reliability of the low-orbit constellation system is improved, and the service life of the low-orbit constellation system is prolonged.