• Journal of Astronomy and Space Sciences
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• 제20권1호
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• pp.11-20
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• 2003

A method for estimating the NORAD SGP4 atmospheric drag term from minimum osculating orbit states, i.e., two osculating orbits, is developed. The first osculating orbit state is converted into the NORAD TLE-type mean orbit state by iterative procedure. Then the converted TLE is propagated to the second orbit state using the SGP4 model with the incremental SGP4 drag term. The iterative orbit propagation procedure is finished when the difference of the two osculating semi-major axes between the propagated orbit and the given second orbit is minimized. In order to minimize the effect of the short-term variations of the osculating semi-major axis, the osculating argument of latitude of the second orbit is propagated to the same argument of latitude of the first orbit. The method is applied to the estimation of the NORAD-type TLE for the KOMPSAT-1 spacecraft. The SGP4 drag terms are estimated from both NORAD SGP4 orbit propagation and the numerical orbit propagation results. Variations of the estimated drag terms are analyzed for the KOMPSAT-1 satellite orbit determination results.

• Journal of Astronomy and Space Sciences
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• 제22권3호
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• pp.243-248
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• 2005

NORAD Two Line Element (TLE) is very useful to simplify the ground station antenna pointing and mission operations. When a satellite operations facility has the capability to determine NORAD type TLE which is independent of NORAD, it is important to analyze the applicable tracking data arcs for obtaining the best possible orbit. The applicable tracking data arcs for NORAD independent TLE orbit determination of the KOMPSAT-1 using GPS navigation solutions was analyzed for the best possible orbit determination and propagation results. Data spans of the GPS navigation solutions from 1 day to 5 days were used for TLE orbit determination and the results were used as Initial orbit for SGP4 orbit propagation. The operational orbit determination results using KOMPSAT-1 Mission Analysis and Planning System(MAPS) were used as references for the comparisons. The best-matched orbit determination was obtained when 3 days of GPS navigation solutions were used. The resulting 4 days of orbit propagation results were within 2 km of the KOMPSAI-1 MAPS results.

• Journal of Astronomy and Space Sciences
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• 제16권2호
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• pp.159-166
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• 1999

As the spatial resolution of remote sensing satellites becomes higher, very accurate determination of the position of a LEO (Low Earth Orbit) satellite is demanding more than ever. Non-symmetric Earth gravity is the major perturbation force to LEO satellites. Since the orbit propagation is performed in the celestial frame while Earth gravity is defined in the terrestrial frame, it is required to convert the coordinates of the satellite from one to the other accurately. Unless the coordinate conversion between the two frames is performed accurately the orbit propagation calculates incorrect Earth gravitational force at a specific time instant, and hence, causes errors in orbit prediction. The coordinate conversion between the two frames involves precession, nutation, Earth rotation and polar motion. Among these factors, unpredictability and uncertainty of Earth rotation, called UTI-UTC, is the largest error source. In this paper, the effect of UTI-UTC on the accuracy of the LEO propagation is introduced, tested and analzed. Considering the maximum unpredictability of UTI-UTC, 0.9 seconds, the meaningful order of non-spherical Earth harmonic functions is derived.

• International Journal of Aerospace System Engineering
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• 제3권1호
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• pp.5-9
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• 2016

This paper briefly introduces efforts into space debris modeling towards a better understanding of space debris environment. Space debris modeling mainly consists of debris generation and orbit propagation. Debris generation can characterize and predict physical properties of fragments originating from explosions or collisions. Orbit propagation can characterize, track, and predict the behavior of individual or groups of space objects. Therefore, space debris modeling can build evolutionary models as essential tools to predict the stability of the future space debris populations. Space debris modeling is also useful and effective to improve the efficiency of measurements to be aware of the present environment.

• Journal of Astronomy and Space Sciences
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• 제31권1호
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• pp.41-50
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• 2014

The Earth is not perfectly spherical and its rotational axis is not fixed in space, and these geophysical and kinematic irregularities work as dominant perturbations in satellite orbit propagation. The International Earth Rotation Service (IERS) provides the Conventions as guidelines for using the Earth's model and the reference time and coordinate systems defined by the International Astronomical Union (IAU). These guidelines are directly applied to model orbital dynamics of Earth satellites. In the present work, the effects of the latest conventions released in 2010 on orbit propagation are investigated by comparison with cases of applying the previous guidelines, IERS Conventions (2003). All seven major updates are tested, i.e., for the models of the precession/nutation, the geopotential, the ocean tides, the ocean pole tides, the free core nutation, the polar motion, and the solar system ephemeris. The resultant position differences for one week of orbit propagation range from tens of meters for the geopotential model change from EGM96 to EGM2008 to a few mm for the precession/nutation model change from IAU2000 to IAU2006. The along-track differences vary secularly while the cross-track components show periodic variation. However, the radial-track position differences are very small compared with the other components in all cases. These phenomena reflect the variation of the ascending node and the argument of latitude. The reason is that the changed models tested in the current study can be regarded as small fluctuations of the geopotential model from the point of view of orbital dynamics. The ascending node and the argument of latitude are more sensitive to the geopotential than the other elements. This study contributes to understanding of the relation between the Earth's geophysical properties and orbital motion of satellites as well as satellite-based observations.

• Journal of Astronomy and Space Sciences
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• 제34권2호
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• pp.127-138
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• 2017

The current study designs the mission orbit of the lunar CubeSat spacecraft to measure the lunar local magnetic anomaly. To perform this mission, the CubeSat will impact the lunar surface over the Reiner Gamma swirl on the Moon. Orbit analyses are conducted comprising ${\Delta}V$ and error propagation analysis for the CubeSat mission orbit. First, three possible orbit scenarios are presented in terms of the CubeSat's impacting trajectories. For each scenario, it is important to achieve mission objectives with a minimum ${\Delta}V$ since the CubeSat is limited in size and cost. Therefore, the ${\Delta}V$ needed for the CubeSat to maneuver from the initial orbit toward the impacting trajectory is analyzed for each orbit scenario. In addition, error propagation analysis is performed for each scenario to evaluate how initial errors, such as position error, velocity error, and maneuver error, that occur when the CubeSat is separated from the lunar orbiter, eventually affect the final impact position. As a result, the current study adopts a CubeSat release from the circular orbit at 100 km altitude and an impact slope of $15^{\circ}$, among the possible impacting scenarios. For this scenario, the required ${\Delta}V$ is calculated as the result of the ${\Delta}V$ analysis. It can be used to practically make an estimate of this specific mission's fuel budget. In addition, the current study suggests error constraints for ${\Delta}V$ for the mission.

• International Journal of Aeronautical and Space Sciences
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• 제8권1호
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• pp.122-128
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• 2007

In this paper, a method of station keeping strategy using relative orbital motion and numerical optimization technique is presented for geostationary satellite. Relative position vector with respect to an ideal geostationary orbit is generated using high precision orbit propagation, and compressed in terms of polynomial and trigonometric function. Then, this relative orbit model is combined with optimization scheme to propose a very efficient and flexible method of station keeping planning. Proper selection of objective and constraint functions for optimization can yield a variety of station keeping methods improved over the classical ones. Nonlinear simulation results have been shown to support such concept.

• Journal of Astronomy and Space Sciences
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• 제21권3호
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• pp.221-232
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• 2004

TLE로부터 SGP4/SDP4 모델을 이용하여 인공위성의 가상의 위치 정보를 얻은 후 Gauss 방법을 사용하여 인공위성의 예비궤도를 결정해보았다. 예비궤도 결정에 필요한 임의의 세 점 사이의 시간간격을 변화시켜 얻은 결과를 위성의 위치 참값과 비교하여 최소의 차이를 가지는 관측 시간 간격을 찾아보았으며, Gauss 예비궤도 결정법의 성능을 비교, 분석하였다. 실제 인공위성 관측 결과와의 비교를 위해서 한국천문연구원의 광시야 망원경을 사용하여 GPS위성(PRN 26)과 무궁화 2호의 광학관측 데이터를 얻은 후 같은 방법으로 예비궤도를 결정해 보았다. 인공위성의 정밀궤도결정을 위하여 시뮬레이션에서 얻어진 가상의 광학관측 데이터를 가지고 정밀케도결정을 수행하였으며, 관측 데이터의 오차와 관측 시간 간격에 따라 정밀궤도결정을 수행하여 원하는 정밀도를 얻기 위한 관측 시스템의 조건에 대해서 알아보았다.

• Journal of Astronomy and Space Sciences
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• 제28권3호
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• pp.193-201
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• 2011

A strategy for geostationary orbit (or geostationary earth orbit [GEO]) surveillance based on optical angular observations is presented in this study. For the dynamic model, precise analytical orbit model developed by Lee et al. (1997) is used to improve computation performance and the unscented Kalman filer (UKF) is applied as a real-time filtering method. The UKF is known to perform well under highly nonlinear conditions such as surveillance in this study. The strategy that combines the analytical orbit propagation model and the UKF is tested for various conditions like different level of initial error and different level of measurement noise. The dependencies on observation interval and number of ground station are also tested. The test results shows that the GEO orbit determination based on the UKF and the analytical orbit model can be applied to GEO orbit tracking and surveillance effectively.

• Journal of Astronomy and Space Sciences
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• 제14권1호
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• pp.136-141
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• 1997

임의의 순간 인공위성의 위치와 속도를 정밀하게 계산하기 위해서는 섭동력을 일으키는 우주공간의 환경을 정확하게 이해하고 분석하여 정량화함으로써 섭동력에 대한 수리적인 모형을 만들어야 한다. 이들 우주환경모델에 의해서 인공위성이 받는 총가속도는 2계 미분방정식으로 표현되며, 이 방정식을 두 번 적분함으로써 원하는 시각에서의 인공위성의 위치와 속도를 얻는 코웰방법을 사용하여 궤도예측 알고리즘을 완성하였다. 정지위성의 궤도에 미치는 주요한 섭동력으로는 지구의 비대칭 중력 포텐셜에 의한 섭동력, 태양과 달의 중력에 의한 섭동력, 태양의 복사압에 의한 섭동력들이 있는데, 그것들의 정밀성을 최대한 높이기 위해 spherical harmonic 계수들을 40 $\times$ 40까지 적용할 수 있도록 했으며, JPL DE 403 ephemeris의 polynomial 내삽을 통해 지구로부터 태양과 달까지의 거리를 정밀하게 계산하였다. 그리고 수치적분 방법으로는 적분간격을 자동으로 조절하는 8계 Runge-Kutta single step 방법을 사용하였다.