• 대한원격탐사학회:학술대회논문집
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• 대한원격탐사학회 2002년도 Proceedings of International Symposium on Remote Sensing
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• pp.553-558
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• 2002

Physical sensor model in pushbroom satellite images can be made from sensor modeling by rotation parameters and attitude parameters on the satellite track. These parameters are determined by the information obtained from GPS, INS, or star tracker. Provided from satellite image, an auxiliary data error is connected directly with an error of rotation parameters and attitude parameters. This paper analyzed how obtaining satellite images influenced errors of rotation parameters and attitude parameters. furthermore, for detailed analysis, this paper generated simulated satellite image, which was changed variously by rotation parameters and attitude parameters of satellite sensor model. Simulated satellite image is generated by using high-resolution digital aerial image and DEM (Digital Elevation Model) data. Moreover, this paper determined correlation of rotation parameter and attitude parameters through error analysis of simulated satellite image that was generated by various rotation parameters and attitude parameters.

• International Journal of Aeronautical and Space Sciences
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• 제1권1호
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• pp.70-80
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• 2000

The attitude knowledge error model is formulated for specifically KOMPSAT attitude determination system using the Lefferts/Markley/Shuster method, and the attitude determination(AD) error analysis is performed so as to investgate the on-board attitude determination capability of KOrea Multi-Purpose SATellite(KOMPSAT) using the covariance analysis method. Analysis results show there is almost no initial value effect on Attitude Determination (AD) error and the sensor noise effects on AD error are drastically decreased as is predicted because of the inherent characteristic of Kalman filter structure. However, it shows that the earth radiance effect of IR-sensor(earth sensor) and the bias effects of both IR-sensor and fine sun sensor are the dominant factors degrading AD error and gyro rate bias estimate error in AD system. Analysis results show that the attitude determination errors of roll, pitch and yaw axes are 0.056, 0.092 and 0.093 degrees, respectively. These numbers are smaller than the required values for the normal mission of KOMPSAT. Also, the selected on-orbit data of KOMPSAT is presented to demonstrate the designed AD system.

• Journal of Mechanical Science and Technology
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• 제15권6호
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• pp.743-751
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• 2001

A Prototype reaction wheel, named the Hankuk Aviation University (HAU) reaction wheel, has been developed for KAISTSAT satellite series. Torque and force disturbances are inherent in reaction wheels, and thus the force and torque characteristics should be examined for every newly developed reaction wheel. The torque and force disturbance noises in the prototype HAU reaction wheel are measured with a torque-measuring table developed for the present study. A new measuring procedure based on a simple principle is applied for the measurements. The frequency characteristics of the torque and force noises are analyzed by examining the power spectral density. The effect of the torque noise on the attitude stability is also examined through numerical simulations with a single-axis attitude model. The noise-induced attitude error and jitter and found to be well below the specified error limits for the KAISTSAT satellite series.

• 제어로봇시스템학회:학술대회논문집
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• 제어로봇시스템학회 1998년도 제13차 학술회의논문집
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• pp.53-57
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• 1998

Star tracker placement configuration is proposed and the properness of the placement configuration is verified for star tracker's sun avoidance angle requirement. Precision attitude determination system is successfully designed using a gyro-star tracker inertial reference system for a candidate LEO spacecraft. Elaborate kalman filter formulation for a spacecraft is proposed for covariance analysis. The covariance analysis is performed to verify the capability of the proposed attitude determination system. The analysis results show that the attitude determination error and drift rate error are good enough to satisfy the mission of a candidate spacecraft.

• 한국항공우주학회지
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• 제45권11호
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• pp.967-974
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• 2017

반작용휠은 저속구간에서 마찰로 인해서 비선형 토크 응답을 보인다. 따라서 이 구간에서는 위성의 정밀한 자세제어를 달성하기 어렵다. 기존 연구들은 마찰력 보상이나 디더명령을 인가하는 방법을 사용하여 본 문제를 해결하려 하였다. 하지만 마찰력 모델링의 어려움이나 휠속도의 빈번한 영점 교차 때문에, 이러한 방법을 실제 위성 자세제어에 적용하기에는 어려움이 있다. 이를 해결하기 위해서, 자세오차에 따라서 자세제어기의 이득값을 조절하는 방법을 제안한다.

• 대한전기학회:학술대회논문집
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• 대한전기학회 1991년도 하계학술대회 논문집
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• pp.829-832
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• 1991

Attitude and position parameter of satellite is needed in producing digital elevation model. This paper propose a method which determines the order of polynomial and initial values of satellite exteria parameters using header file of computer compatible tape. Experiments on a SPOT level 1A full science show that planimetric error is 11m and the altimetric error is about 18.3m.

• 한국항공우주학회지
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• 제48권12호
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• pp.987-995
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• 2020

고성능 탑재체들과 자세제어 센서들 간의 정밀정렬은 인공위성의 정확한 자세지향 및 높은 지향 안정성을 위해 필수적이다. 위성 개발사들은 조립 및 시험기간 동안 위성 정렬을 위해 데오드라이트 측정 시스템을 주로 사용한다. 데오드라이트 측정 시 시선 방향 오차, 수평축의 오차, 수직방향 인덱스 오차 그리고 수직축 오차로 인해 측정오차가 발생할 수 있다. 이러한 오차들 뿐 아니라 다수의 데오드라이트를 사용한 측정 시 발생할 수 있는 오차들을 정렬큐브 측정실험을 통해 분석하였다. 정렬큐브 측정실험을 기반으로 정렬측정 정확도를 향상시킬 수 있는 방법이 제안되었고, 측정 결과 위성의 설계 요구조건도 만족시킬 수 있었다.

• 한국군사과학기술학회지
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• 제10권2호
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• pp.160-169
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• 2007

In this paper, we have described and simulated the effect analysis and correction of phase errors in the SAR rawdata induced by satellite attitude errors such as drift, jitter. This simulation is based on the FSA(Frequency Scaling Algorithm) for high resolution image formation of the Spotlight SAR. Phase errors produce the degradation of SAR image quality such as loss of resolution, geometric distortion, loss of contrast, spurious targets, and decrease in SNR. To resolve this problem, this paper presents method for correction of phase errors using the PGA(Phase Gradient Algorithm) in connection with the FSA. Several results of the phase errors correction are presented for Spotlight SAR rawdata.

• 전기학회논문지
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• 제57권1호
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• pp.98-101
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• 2008

In this paper, we describe the RLG current stabilizer circuit for attitude control in the satellite. The RLG makes use of the Sagnac effect within a resonant of a HeNe laser. The difference between two discharge currents causes one of the gyro bias errors. The theoretical background and current stabilizer are introduced. It is verified that the circuit designed is applicable to the test of input voltage and temperature.

• 항공우주기술
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• 제10권2호
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• pp.87-93
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• 2011

고정밀 태양센서는 인공위성의 자세제어에 중요 센서로서, 위성으로 입사되는 태양 빛의 방향을 측정하거나 위성이 태양을 보지 못하는 상태에 있는지를 판단하기 위해서 사용되고 있다. 또한 정지궤도 위성에서는 전이궤도 및 임무궤도 상에서 기준 자세로 부터 벗어난 자세오차 정보를 획득하기 위해서 또는 이상 발생 시 태양벡터를 획득하기 위해서 고정밀 태양센서를 사용하고 있다. 본 논문에서는 저궤도 위성과 정지궤도 위성용 고정밀 태양센서의 형상에 대한 이해를 바탕으로 태양의 입사각에 대한 출력 전류 관계를 나타내는 전달 함수를 이용하여 고정밀 태양센서 운용 원리를 설명한다.