Ephemeris Estimation of a Well-Defined Platform Using Satellite Laser Ranging from a Reduced Number of Ground Sites
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Abstract : This study presents the results of an analysis comparing ephemerides obtained using Satellite Laser Ranging (SLR) derived from a reduced number of ground sites. The study provides insight into the extent to which ephemeris can be determined for an extremely well-specified satellite. The study was conducted to determine the viability of using a single SLR site to provide an independent method of verifying onboard Global Positioning System navigational performance. Computational simulations included varying the number and distribution of sites as well as empirical modeling of nonconservative forces to determine the limitations of this SLR- based reduction strategy.Keywords:
Ephemeris
Satellite laser ranging
Ranging
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Basis (linear algebra)
Satellite laser ranging
Ranging
Realization (probability)
Data set
Accuracy and precision
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In order for celestial navigation techniques to provide accurate positioning estimates, precise ephemerides of the observed satellites are necessary. This work analyzed a method to correct for satellite ephemeris to be used in celestial navigation applications. This correction is the measured angle difference between the expected location of the satellite, which is given by propagating publicly available Two-Line Element sets (TLE), and their observed angles from a precisely known reference site. Therefore, the angle difference can be attributed completely to satellite ephemeris error assuming instrument error was accounted for. The intent is to calculate this correction from the reference site and relate it to remote sites that have visibility of the same satellite, but where its own location is known with some uncertainty. The effects of increased baseline distances from the reference site are studied, as well as time delays. Satellite observations were simulated and propagated using TLE. This simulated data was used to calculate the angle difference and project that angle to the viewpoint of the remote site. This corrected observed angle was integrated using an extended Kalman filter (EKF) with an inertial measurement unit (IMU) and a barometric altimeter. The performance of the position solution in the navigation filter was calculated as the error from simulated truth. The satellite ephemeris error measured at a reference location becomes less observable by a remote user according to the line-of-sight transformation due to the reference-satellite-remote geometry. A mathematical formula for calculating the applicability of projecting the remote site observation to other locations is developed and compared to simulated ephemeris errors. This formula allows a user to define geographic regions of validity through ephemeris error tolerance. Estimating the ephemeris error with regular updates from a reference site resulted in a reduction of IMU drift and a distance root mean squared (DRMS) error of 100 m.
Ephemeris
Dilution of precision
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Orbit Determination
Orbit (dynamics)
Satellite laser ranging
Empirical modelling
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Satellite laser ranging
Collocation (remote sensing)
Doris (gastropod)
Retroreflector
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Ranging
Satellite laser ranging
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Satellite laser ranging
Retroreflector
Orbit Determination
Ranging
Orbit (dynamics)
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In this paper predicting of position of satellite based on extended kalman filter with considering hardware implementation consideration and simultaneously maintaining desired accuracy is investigated. For this purpose, first, effective forces on orbital dynamic and nonlinear equation of orbital motion are presented. In order to increasing accuracy of prediction in position of satellite, J2, J3 and J4 harmonics of potentialfunction of the earth are considered and future position of satellite is predicted using linearized dynamic model and applying EKF on this model. Here Measurement data are position and velocity vector of satellite which are extracted by GPS receivers. Since in this paper systematic satellite design is considered, scenario of “ON TIME” of GPS receivers based on power consumption considerations is discussed. Finally simulation results for a LEO satellite and comparing these results with STK results, shows accuracy of presented modeling and equations.
Position (finance)
Orbital elements
Orbital mechanics
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Modern precise navigation services are creating increased applications for numerically generated state vectors for satellite operations. Traditional radar and optical techniques can achieve modest accuracy in orbit determination, but on-board GPS satellite receivers are changing the routine accuracy available. System requirements usually involve future locations, rather than past locations derived from OD techniques. This paper compares propagation of various satellite initial state vectors to independently produced Precision Orbit Ephemerides (POE’s). The initial state of each satellite is varied to reflect expected orbital accuracy achievable through existing orbit determination techniques. Satellite ephemerides are compared to known POE’s, and to precise reference ephemerides generated by state-of-the-art orbit determination techniques.
Ephemeris
State vector
Orbit Determination
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Abstract The autonomous navigation system characteristics are investigated for an artificial satellite around a planet. Navigation is based on the optical scanning of stars near the planet's horizon. The recognition of stars and the measurement of angles of star elevation above the horizon are carried out. The subsequent statistical processing of the measurements allows one to determine the satellite's orbital elements with an accuracy dependent on the measurement accuracy and procedure. The analysis employs an exact numerical algorithm and approximate numerical-analytical technique. The navigational accuracies are studied in relation to independent and correlated measurement errors. Comparison is made for navigational accuracies determined by the above two methods. It shows that the approximate technique allows rather well (with an error less than 10%) to find navigational errors even at a small number (3–6) of times of measurements per revolution. When this number is increased the methodical error of the approximate technique quickly diminishes. The navigational accuracies are obtained for the satellite in an orbit with altitude of 300 to 36,000 km. The navigational algorithm was tested in direct numerical simulation and its convergence bounds were determined. The latter show admissible deviations of initial values in orbital elements from the exact ones. The analysis shows that the algorithm is stable at rather large errors in giving the elements of initial approximation.
Orbit (dynamics)
Orbit Determination
Orbital elements
Elevation (ballistics)
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Ranging
Satellite laser ranging
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