• Journal of Institute of Control, Robotics and Systems
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• v.9 no.10
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• pp.852-860
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• 2003

We introduce the error models for an attitude determination system(ADS) with gyroscopes and stellar sensor. The ADS error models are derived according to the definition of the reference frame and of the attitude error. The equivalent error models applicable to the attitude determination system with large attitude errors are presented. The simulation results show that the proposed error models improve performance of the attitude determination system.

• International Journal of Aeronautical and Space Sciences
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• v.1 no.1
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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.

• The Transactions of The Korean Institute of Electrical Engineers
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• v.63 no.9
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• pp.1259-1265
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• 2014

In this paper, the real-time attitude determination based Matlab using low-cost receivers was designed and evaluated. The GNSS attitude determination system was implemented to operation in real-time by TimerCallback in MATLAB. The TTM(Transmission Time Misalignment) of U-blox receiver was confirmed through zero baseline tests and this problem was revised. The computed attitude by the high-cost NovAtel receiver was compared to the computed attitude by the low-cost U-blox receiver. As a result of this, the performance of attitude determination systems by low-cost receiver was confirmed. To determine baseline, LAMBDA and BC-LAMBDA for integer ambiguities search methods were used. To confirm suitable integer ambiguity search method in real-time attitude determination algorithm, determined baselines by two methods were compared, and it was confirmed that BC-LAMBDA is more suitable. As a result of this, the operation of real-time attitude determination system was confirmed using 3 low-cost receivers.

• 제어로봇시스템학회:학술대회논문집
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• 2000.10a
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• pp.342-342
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• 2000

We present 3-axis stabilized spacecraft attitude determination algorithm using the magnetometer. The magnetometer has been used as a reliable, light-weight and inexpensive sensor in attitude determination and reaction wheel momentum dumping system. Recent studies have attempted to use the magnetometer when other attitude sensor, such as star tracker, fails. The differences between the measured and computed the Earth's magnetic field components are spacecraft attitude errors. In this paper, we propose extended Kalman filter(EKF) to determine spacecraft attitude with the magnetometer data and gyro-measured body rates. We develop and simulate this algorithm using MATLAB/SIMULINK. This algorithm can be used as a backup attitude determination system.

• 제어로봇시스템학회:학술대회논문집
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• 1996.10b
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• pp.1095-1098
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• 1996

In this work, attitude determination with Inertial Reference Unit as attitude sensor is considered. Usually, the attitude error from IRU increases because of gyro rate bias and noise. Therefore, other attitude sensors(sun sensor, horizon sensor, star tracker) are needed to compensate for error from IRU. In this paper, we use the extended Kalman filter for attitude estimation of spacecraft with IRU and star tracker.

• 제어로봇시스템학회:학술대회논문집
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• 2004.08a
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• pp.1966-1971
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• 2004

As the remote sensing satellite technology grows, the acquisition of accurate attitude and position information of the satellite has become more and more important. Due to the data processing limitation of the on-board orbit propagator and attitude determination algorithm, it is required to develop much more accurate orbit and attitude determination, which are so called POD (precision orbit determination) and PAD (precision attitude determination) techniques. The sensor and attitude dynamics simulation takes a great part in developing a PAD algorithm for two reasons: 1. when a PAD algorithm is developed before the launch, realistic sensor data are not available, and 2. reference attitude data are necessary for the performance verification of a PAD algorithm. A realistic attitude dynamics and sensor (IRU and star tracker) outputs simulation considering their physical characteristics are presented in this paper, which is planned to be used for a PAD algorithm development, test and performance verification.

• 제어로봇시스템학회:학술대회논문집
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• 1998.10a
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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.

• 제어로봇시스템학회:학술대회논문집
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• 2001.10a
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• pp.167.5-167
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• 2001

Single point attitude determination method provides an optimal attitude minimizing the Wahba loss function. However, for the insufficient number of measurement vectors, the conventional single point methods has no unique solution. Thus, we introduce the sequential method to and an optimal attitude minimizing the windowed loss function. In this paper, this function is de ned as the sum of square errors for all measurement vectors within the axed sliding window. For simple implementation, the proposed algorithm is rewritten as a recursive form. Moreover, the covariance matrix is derived and expressed as a recursive form. Finally, we apply this algorithm to the attitude determination system with three LOS measurement sensors.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.7
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• pp.609-618
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• 2015

This paper discusses the attitude determination of the CNUSAIL-1 cube-satellite. The primary mission of the CNUSAIL-1 is sail deployment and operation in low Earth orbit, and the secondary mission is to look into influence of the sail deployment on satellite attitude and orbit. The attitude determination strategy is proposed depending on three mission phases, and its performance and applicability are verified through numerical simulations. This study considers the following sensors: Sun sensors and a three-axis magnetometer as attitude reference sensors, and a three-axis MEMS gyroscope as an inertial attitude sensor. Because sensors used for cube satellites have relatively low performances and worse noise characteristics, an Extended Kalman filter (EKF) is applied to attitude determination. Additionally, it has the merits to deal with the Gaussian noises and to predict the attitude even with no measurements from reference attitude sensors, especially in the eclipse of the cube satellite. The performance of the EKF is compared to a deterministic attitude determination technique, QUEST(QUaternion ESTimation).

• International Journal of Aeronautical and Space Sciences
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• v.14 no.1
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• pp.91-98
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• 2013

The rigorous requirements of modern spacecraft missions necessitate a precise attitude determination strategy. This paper mainly researches that, based on three space-borne attitude sensors: 3-axis rate gyros, 3-antenna GPS receiver and star-sensor. To obtain global attitude estimation after an information fusion process, a feedback-involved Federated Kalman Filter (FKF), consisting of two subsystem Kalman filters (Gyros/GPS and Gyros/Star-sensor), is established. In these filters, the state equation is implemented according to the spacecraft's kinematic attitude model, while the residual error models of GPS and star-sensor observed attitude are utilized, to establish two observation equations, respectively. Taking the sensors' different update rates into account, these two subsystem filters are conducted under a variable step size state prediction method. To improve the fault tolerant capacity of the attitude determination system, this paper designs malfunction warning factors, based on the principle of ${\chi}^2$ residual verification. Mathematical simulation indicates that the information fusion strategy overwhelms the disadvantages of each sensor, acquiring global attitude estimation with precision at a 2-arcsecs level. Although a subsystem encounters malfunction, FKF still reaches precise and stable accuracy. In this process, malfunction warning factors advice malfunctions correctly and effectively.