• 제어로봇시스템학회:학술대회논문집
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• 제어로봇시스템학회 2003년도 ICCAS
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• pp.2331-2334
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• 2003

The basic difference between the EKF(Extended Kalman Filter) and UKF(Unscented Kalman Filter) stems from the manner in which Gaussian random variables(GRV) are represented for propagating through system dynamics. In the EKF, the state distribution is approximated by a GRV, which is then propagated analytically through the first-order linearization of the nonlinear system. This can possibly introduce large errors in the true posterior mean and covariance of the transformed GRV, which may lead to sub-optimal performance and sometimes divergence of the filter. However, the UKF addresses this problem by using a deterministic sampling approach. The state distribution is also approximated by a GRV, but is now represented using a minimal set of carefully chosen sample points. These sample points completely capture the true mean and covariance of the GRV, and UKF captures the posterior mean and covariance accurately up to the 2nd order(Taylor series expansion) for any nonlinearity. This paper utilizes the UKF to determine spacecraft orbit when only magnetometer is available. Several catastrophic failures of spacecraft in orbit have been attributed to failures of the spacecraft mission. Recently studies on contingency-major sensor failure cases- have been performed. For mission success, contingency design or plan should be implemented in case of a major sensor failure. Therefore the algorithm presented in this paper can be used for a spacecraft without GPS or contingency design in case of GPS failure.

• International Journal of Aeronautical and Space Sciences
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• 제10권1호
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• pp.67-74
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• 2009

In general, the lunar landing stage can be divided into two distinct phases: de-orbit and descent, and the descent phase usually comprises two sub-phases: braking and approach. And many optimization problems of minimal energy are usually focused on descent phases. In these approaches, the energy of de-orbit burning is not considered. Therefore, a possible low perilune altitude can be chosen to save fuel for the descent phase. Perilune altitude is typically specified between 10 and 15km because of the mountainous lunar terrain and possible guidance errors. However, it requires more de-orbit burning energy for the lower perilune altitude. Therefore, in this paper, the perilune altitude of the intermediate orbit is also considered with optimal thrust programming for minimal energy. Furthermore, the perilune altitude and optimal thrust programming can be expressed by a function of the radius of a parking orbit by using continuation method and co-state estimator.

• 한국측량학회지
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• 제28권6호
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• pp.605-612
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• 2010

광범위한 지역을 관측하기 위한 많은 종류의 위성들이 발사되어 지구를 관측하고 있다. 이러한 위성들은 영상정보 이외에 천체력 자료, RPC 계수 등과 같은 위성궤도와 관련 정보들을 제공하고 있다. 위성에서 제공하는 이러한 궤도정보를 활용할 경우 영상의 가하보정에 요구되는 기준점을 줄일 수 있다. 본 연구에서는 RADARSAT-l SAR 위성영상을 대상으로 동일궤도에서 촬영된 다중 위성영상들의 효과적인 기하보정을 위하여 기준영상에서 단일기준점 및 천제력자료를 활용하여 위성궤도를 모델링하고, 이를 기반으로 동일궤도상에서 취득된 인접영상의 기하보정기법을 기술하였다. 정확도 평가를 위해서 본 연구에서 제시한 기법으로 생성된 기하보정영상을 Erdas Imagine에서 처리한 기하보정영상과 비교하여 정확도를 평가하였다.

• 한국항행학회논문지
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• 제13권2호
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• pp.178-186
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• 2009

SBAS(Satellite Based Augmentation System) 시스템에서는 GNSS 사용자들의 위치 정확도 향상을 위해 위성궤도 및 시계보정정보를 제공하고 있는데, 본 논문에서는 이러한 보정정보의 정확도에 대해 분석하였다. IGS(International GNSS Service)에서 제공하는 GPS 위성의 정밀궤도력을 참값으로 가정하고, 그에 대한 오차를 이용하여 정확도를 분석/수행하였다. 이때 IGS 정밀궤도력과의 정확한 비교를 위해 GPS 위성에 대한 안테나 위상중심 편차와 P1-C1 편이를 고려하였다. SBAS 위성궤도 및 시계보정 정보로는 미국의 WAAS와 일본의 MSAS 보정정보를 이용하였다. 정확도 분석을 통해 SBAS에서 제공하는 위성궤도 보정정보와 위성시계 보정정보가 상당한 상관관계를 가지고 있음을 확인하였다. 또한 보정정보의 정확도는 SBAS 시스템의 지상 네트워크 크기와 위성의 궤적에 영향을 받는 것을 확인하였다.

• 천문학회보
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• 제35권2호
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• pp.54-54
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• 2010

Interplanetary space is filled with dust particles originating mainly from comets and asteroids. Such interplanetary dust particles lose their angular momentum by olar radiation pressure, causing the dust grains to slowly spiral inward Poynting-Robertson effect). As dust particles move into the Sun under the influence of Poynting-Robertson drag force, they may encounter regions of resonance just outside planetary orbits, and be trapped by their gravities, forming the density enhancements in the dust cloud (circumsolar resonance ring). The circumsolar resonance ring near the Earth orbit was detected in the zodiacal cloud through observations of infrared space telescopes. So far, there is no observational evidence other than Earth because of the detection difficulty from Earth bounded orbit. A Venus Climate Orbiter, AKATSUKI, will provide a unique opportunity to study the Venusian resonance ring. It equips a near-infrared camera for the observations of the zodiacal light during the cruising phase. Here we consider whether Venus gravity produces the circumsolar resonance ring around the orbit. We thus perform the dynamical simulation of micron-sized dust particles released outside the Earth orbit. We consider solar radiation pressure, solar gravity, and planetary perturbations. It is found that about 40 % of the dust particles passing through the Venus orbit are trapped by the gravity. Based on the simulation, we estimate the brightness of the Venusian resonance ring from AKATSUKI's locations.

• Journal of Astronomy and Space Sciences
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• 제20권1호
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• pp.1-10
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• 2003

A precise kinematic orbit determination (P-KOD) procedure for Low Earth Orbiter(LEO) using the GPS ion-free triple differenced carrier phases is presented. Because the triple differenced observables provide only relative information, the first epoch's positions of the orbit should be held fixed. Then, both forward and backward filtering was executed to mitigate the effect of biases of the first epoch's position. p-KOD utilizes the precise GPS orbits and ground stations data from International GPS Service (IGS) so that the only unknown parameters to be solved are positions of the satellite at each epoch. Currently, the 3-D accuracy off-KOD applied to CHAMP (CHAllenging Min-isatellite Payload) shows better than 35 cm compared to the published rapid scientific orbit (RSO) solution from GFZ (GeoForschungsZentrum Potsdam). The data screening for cycle slips is a particularly challenging procedure for LEO, which moves very fast in the middle of the ionospheric layer. It was found that data screening using SNR (signal to noise ratio) generates best results based on the residual analysis using RSO. It is expected that much better accuracy are achievable with refined prescreening procedure and optimized geometry of the satellites and ground stations.

• 한국통신학회논문지
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• 제25권5A호
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• pp.641-650
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• 2000

톤 방식을 사용하는 저궤도 위성 거리 측정에서 링크 잡음과 도플러 영향을 분석하고, 다목적 실용위성과 거리측정 시스템에 적합한 거리 측정 검출부의 운용 잡음 대역폭을 제안하였다. 저궤도 위성의 운동 특성을 통한 궤도 파라미터별 영향과 거리 측정 신호의 신호대 잡음 스펙트럼 밀도를 구하여 잡음과 도플러의 영향을 분석하였다. 위성 링크 잡음의 영향은 거리 신호 검출부 PLL 대역폭이 낮을수록 작으나, 도플러에 의한 영향은 PLL 대역폭이 클수록 작아진다. 다목적 실용위성의 거리 측정 시스템에서 적용하여 측정한 결과와 위성 궤도 특성을 모델링하여 모의 실험한 거리 측정 시스템의 분석 결과를 통하여 저궤도 위성의 거리 측정에 대한 도플러와 잡음의 영향 분석에 대한 타당성을 확인하였다.

• Journal of Astronomy and Space Sciences
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• 제35권1호
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• pp.31-37
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• 2018

Pc1 pulsations are geomagnetic fluctuations in the frequency range of 0.2 to 5 Hz. There have been several observations of Pc1 pulsations in low earth orbit by MAGSAT, DE-2, Viking, Freja, CHAMP, and SWARM satellites. However, there has been a clear limitation in resolving the spatial and temporal variations of the pulsation by using a single-point observation by a single satellite. To overcome such limitations of previous observations, a new space mission was recently initiated, using the concept of multi-satellites, named the Small scale magNetospheric and Ionospheric Plasma Experiments (SNIPE). The SNIPE mission consists of four nanosatellites (~10 kg), which will be launched into a polar orbit at an altitude of 600 km (TBD) in 2020. Four satellites will be deployed in orbit, and the distances between each satellite will be controlled from 10 to 1,000 km by a high-end formation-flying algorithm. One of the possible science targets of the SNIPE mission is observing electromagnetic ion cyclotron (EMIC) waves. In this paper, we report on examples of observations, showing the limitations of previous EMIC observations in low earth orbit, and suggest possibilities to overcome those limitations through a new mission.

• Journal of Power Electronics
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• 제6권1호
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• pp.18-28
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• 2006

The Electric Power Subsystem (EPS) is one of the most critical systems on any satellite because nearly every subsystem requires power. This makes the choice of power systems the most important task facing satellite designers. The main purpose of the Satellite EPS is to provide continuous, regulated and conditioned power to all the satellite subsystems. It has to withstand radiation, thermal cycling and vacuums in hostile space environments, as well as subsystem degradation over time. The EPS power characteristics are determined by both the parameters of the system itself and by the satellite orbit. After satellite separation from the launch vehicle (LV) to its orbit, in almost all situations, the satellite subsystems (attitude determination and control, communication and onboard computer and data handling (OBC&DH)), take their needed power from a storage battery (SB) and solar arrays (SA) besides the consumed power in the EPS management device. At this point (separation point, detumbling mode), the satellite's angular motion is high and the orientation of the solar arrays, with respect to the Sun, will change in a non-uniform way, so the amount of power generated by the solar arrays will be affected. The objective of this research is to select satellite EPS component types, to estimate solar array illumination parameters and to determine the efficiency of solar arrays during both detumbling and normal operation modes.

• 항공우주기술
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• 제10권2호
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• pp.65-73
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• 2011

위성의 발사 후 정지궤도위성을 운용궤도에 진입시키기 위해서는 전이궤도 원지점 (Apogee)에서 위성에 장착된 액체원지점엔진을 발사하여 근지점(Perigee) 고도를 정지궤도 고도에 이르도록 높여준다. 이 과정에서 궤도결정 결과를 피드백하여 정밀하게 궤도조정을 수행하고 종료 후 원하는 경도에 안착시키기 위해 상황에 따라 엔진발사를 3회에서 4회로 나누어 수행하게 된다. 본 논문에서는 먼저 임펄스 형태의 기동을 가정하여 각 액체원지점 엔진 발사시점과 ${\Delta}V$ 벡터를 결정하기 위한 알고리즘을 수립하였고, 추가적으로 연속점화에 따른 오차를 보정하기 위한 기법을 제안하였다. 또한 시뮬레이션을 통하여 제안된 기법의 타당성을 분석하였다.