• Journal of the Korean Society of Surveying, Geodesy, Photogrammetry and Cartography
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• v.27 no.4
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• pp.437-444
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• 2009

A bi-directional, multi-step numerical integrator is developed to determine the GPS (Global Positioning System) orbit based on a dynamic approach, which shows micrometer-level accuracy at GPS altitude. The acceleration due to the planets other than the Moon and the Sun is so small that it is replaced by the empirical forces in the Solar Radiation Pressure (SRP) model. The satellite orbit parameters are estimated with the least-squares adjustment method using both the integrated orbit and the published IGS (International GNSS Service) precise orbit. For this estimation procedure, the integration should be applied to the partial derivatives of the acceleration with respect to the unknown parameters as well as the acceleration itself. The accuracy of the satellite orbit is evaluated by the RMS (Root Mean Squares error) of the residuals calculated from the estimated orbit parameters. The overall RMS of orbit error during March 2009 was 5.2 mm, and there are no specific patterns in the absolute orbit error depending on the satellite types and the directions of coordinate frame. The SRP model used in this study includes only the direct and once-per-revolution terms. Therefore there is errant behavior regarding twice-per-revolution, which needs further investigation.

• Aerospace Engineering and Technology
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• v.2 no.1
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• pp.54-62
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• 2003

The satellite in space has a tiny size but is subject to the disturbance torques caused by various sources. The presence of environmental disturbance torques significantly affects the orient as well as the orbital motion of satellite. The sources of environmental effects on LEO Satellite attitude dynamics are various. Four of these, gravity gradient, Earth's magnetic field, solar radiation pressure and aerodynamic are dominant and deterministic. In this study, we describe the model of environmental disturbance torques acting on LEO Satellite and the effects of environmental disturbance torques on LEO Satellite attitude dynamics in detail.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.41 no.10
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• pp.770-778
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• 2013

Solar sail propulsion uses solar radiation pressure to propel spacecraft without propellant, and it is useful for deep-space missions and continuos orbit maneuver missions. After a brief introduction of solar sail dynamics, locally optimal trajectories in Sun-centered and Earth-centered orbits are analyzed. Numerical simulations for the optimal trajectories are performed. Trajectory for the rendezvous with Halley comet is generated, and different planet escape methods are compared.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.9
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• pp.60-66
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• 2006

Now, on developing COMS(Communication, Ocean and Meteorological Satellite) has solar panel on the south panel only. Therefore, the wheel off-loading has to be performed periodically to reduce a induced momentum energy by a asymmetric solar panel. One of two East/West station keeping maneuver to correct simultaneously longitude and eccentricity, orbit corrections may be performed during one of the two wheel off-loading manoeuvres per day to get enough observation time for meteorological and ocean sensor. In this paper, we applied a linearized orbit maneuver equation to acquire maneuver time and delta-V. Nonlinear simulation for the station keeping is performed and compared with general station keeping strategy for fuel reduction.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.65-70
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• 2003

In this work variations of orbital parameters are first derived from the perturbation equations using difference equation method under Earth oblateness and atmospheric drag. A simple and effective scheme is proposed to compute the required delta v and fuel consumption to compensate for atmospheric drag. The scheme is applied to KOMPSAT example. And by means of numerical simulations we quantitatively analyze influences due to each perturbation source, i.e., nonspherical Earth, atmospheric drag, third body gravities (Sun, Moon), and solar radiation.

• Bulletin of the Korean Space Science Society
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• 1993.10a
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• pp.9-9
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• 1993

GPS는 지구 중심으로부터 GPS 위성의 거리와 위성과 관측자사이의 의사거리(pseudorange)를 이용해서 위치를 결정하는 시스템이다. 1993년 6월 12일 연세대학교에서 3시간동안 관측하여 연세대학교의 위치를 구하였다. 이 위치는 WGS-84 타원체이므로 Bessel타원체로 좌표변환하였다. 위치를 결정하기 위해서는 정확한 위성의 위치와 의사거리에 미치는 잡음(noise)을 제거해야 한다. GPS 위성의 위치 결정에는 지구 비대칭중력항에 의한 섭동, 태양, 달에 의한 섭동, 태앙 복사압에 의한 섭동, 지각, 해양의 조석력에 의한 섭동, 태양빛의 지구 반사도(albedo)에 의한 섭동을 고려해야하며 이를 위해서 위성의 Telemetry를 분석하여 구해 보았다 의사거리의 잡음중 가장 큰 요소인 이온층, 대류층에 의한 지연(delay)에 대해 연구 하였고 각각 Kiobuchar모델, Hopfield모델을 써서 보정을 하였다. 자료 처리를 P모델, PV모델을 만들어 칼만 필터에 적용하였고 RV모델이 P모델보다 더 정확하였나, 위치 결정의 정확도를 알아 보기위해서 국립 천문대부설 GPS관측소에서 결정한 위치와 비교,분석하였다.

• Journal of Astronomy and Space Sciences
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• v.17 no.1
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• pp.107-116
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• 2000

This paper presents the preliminary survey and simulation results of the prediction of Leonid stream's orbital motion. Based on the model survey on eject velocity and perturbation of meteoroid particles, a simulation program was developed and applied to orbital motion of Leonid stream. The Jones ejection distribution model was used to describe the particle's eject velocity and the orbital dynamic model includes perturbations of major planet's gravity. DE405 ephemeris file generated by Solar System Dynamics Group at Jet Propulsion Laboratory in NASA was used for the planet's ephemeris calculations. Solar radiation pressure were also considered in the simulation and 8th order Runge-Kutta algorithm was used a numerical integration method.

• Journal of Astronomy and Space Sciences
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• v.14 no.2
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• pp.381-385
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• 1997

A long-term prediction algorithm of geostationary orbit was developed using the analytical method. The perturbation force models include geopotential upto fifth order and degree and luni-solar gravitation, and solar radiation pressure. All of the perturbation effects were analyzed by secular variations, short-period variations, and long-period variations for equinoctial elements such as the semi-major axis, eccentricity vector, inclination vector, and mean longitude of the satellite. Result of the analytical orbit propagator was compared with that of the cowell orbit propagator for the KOREASAT. The comparison indicated that the analytical solution could predict the semi-major axis with an accuracy of better than $pm35$ meters over a period of 3 month.

• Bulletin of the Korean Space Science Society
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• 2009.10a
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• pp.39.4-40
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• 2009

이 연구의 목적은 고해상도 합성개구레이더 센서를 탑재한 관측위성의 운용요구사항에 맞춰 임무기간 동안 관측 목표지역을 주기적으로 반복하고 지상궤적을 $\pm2km$ 범위 내에서 안정성을 갖도록 유지 조정하는 궤도제어 알고리즘 연구를 수행하는데 있다. 기존에 수행되어 왔던 지상궤적에 대한 오차를 해석적으로 계산하여 궤도를 유지 조정하는 방법이 아닌 기준궤도에 대하여 상대좌표계에서 표현된 위성의 실제 접촉궤도를 기준궤도와 직접적으로 비교하여 목표궤적을 유지 조정하는 알고리즘을 연구하였다. 이를 위해 첫째, 고해상도 관측위성의 운용요구사항을 만족하는 계획된 목표궤도인 기준궤도를 설계하였다. 기본적으로 기준궤도는 임무 설계 시 완전한 주기성이 고려된 최대한 실제에 가까운 궤도이기 때문에 지구중력장 모델만을 고려하여 간략하게 설계하였다. 둘째, 실제의 인공위성의 궤도는 계획된 기준궤도를 유지해야 하지만 시간에 따라 섭동력의 영향을 받아 계획된 궤도로부터 벗어나게 된다. 기준궤도로부터 실제궤도가 얼마나 벗어나는지에 대한 정량적 분석을 위해 지구 중력장, 달-태양 중력, 대기저항력, 태양복사압, 조석력 등과 같은 다양한 섭동력의 영향에 대한 분석을 수행하였다. 셋째, 반경방향(radial), 진행방향(along-track), 교차방향(cross-track)의 세 방향의 성분으로 구성된 우주공간오차(Space Error) 개념을 적용하여, 투영된 지상궤적에 상응하는 오차를 계산하는 것 보다 안정적으로 오차를 계산하였다. 또한 운용요구사항에 따라 허용된 범위 내에서 궤도를 유지하기 위해 GVE(Gauss Variation Equation)을 이용한 궤도조정을 수행하였다. 섭동력의 분석 결과로부터 지구대기저항력, 달-태양 중력으로 인해 가장 두드러지는 장반경과 궤도이심률의 변화를 조정하기 위해, 임무에 사용되는 추력기의 연료 효율을 고려하여 동결궤도가 유지될 수 있는 최적의 위도이각에서 In-plane에 대한 궤도조정만을 수행하여 장반경과 이심률을 동시에 조정하였다. 지구대기와 태양활동의 영향으로 시간에 따른 장반경의 변화율에 따라 궤도조정 주기를 가지는 것을 알 수 있었고, 이 변화율 때문에 생기는 우주공간오차의 증가를 보정하여 위성의 지상궤적을 목표범위 안에서 유지할 수 있었다.

• Journal of Advanced Navigation Technology
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• v.16 no.6
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• pp.900-909
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• 2012

Low Earth orbit satellites with satellite navigation receiver use onboard navigation filters for filtering measurement signals and for orbit prediction under signal loss. Precision satellite dynamic models, core of the navigation filter, are studied and a computation program is developed. Gravity acceleration, precision coordinate transform, third-body gravity, atmospheric drag, and solar radiation pressure models are combined into an orbit prediction algorithm, and a proven precision orbit determination software is used to validate the program. Orbit prediction accuracy is analyzed with simulated and flight orbit data. The program meets an accuracy level for onboard real-time navigation filter.