• Journal of Institute of Control, Robotics and Systems
• /
• v.9 no.10
• /
• pp.852-860
• /
• 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
• /
• v.1 no.1
• /
• pp.70-80
• /
• 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
• /
• v.63 no.9
• /
• pp.1259-1265
• /
• 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
• /
• 1998.10a
• /
• pp.53-57
• /
• 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
• /
• 2000.10a
• /
• pp.342-342
• /
• 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.

• Journal of Electrical Engineering and Technology
• /
• v.7 no.4
• /
• pp.615-625
• /
• 2012

GPS attitude outputs or carrier phase observables can be effectively utilized to compensate the attitude error of the strapdown inertial navigation system. However, when the integer ambiguity is not correctly resolved and/or a cycle slip occurs, an erroneous GPS output can be obtained. If the erroneous GPS output is applied to the attitude determination GPS/INS (ADGPS/INS) integrated navigation system, the performance of the system can be degraded. This paper proposes an ADGPS/INS integration system using the triple difference carrier phase observables. The proposed integration system contains a cycle slip detection algorithm, in which the inertial information is combined. Computer simulations and flight test were performed to verify effectiveness of the proposed navigation system. Results show that the proposed system gives an accurate and reliable navigation solution even when the integer ambiguity is not correctly resolved and the cycle slip occurs.

• Journal of Advanced Navigation Technology
• /
• v.8 no.2
• /
• pp.112-119
• /
• 2004

This paper proposes an attitude determination GPS/INS integrated system scheme for a UAV and presents experimental flight test results. The proposed system is designed as a part of an autopilot system and comprises a GPS attitude determination receiver, an off-the-shelf inertial measurement unit (IMU), and a navigation computer unit (NCU). UAV requires accurate attitude information for stable automatic flight control. The proposed system can provide accurate attitude information for the flight control computer (FCC) so that stable automatic flight control can be achieved. In order to verify the performance of the proposed scheme, an integrated navigation system has been developed. In order to evaluate the developed navigation system, the flight test has been performed. In the flight test, the developed system was shown to provide the position, the velocity and the attitude satisfactorily enough for stable flight control. The accuracy of the attitude information of the developed system was confirmed by comparing attitude of vertical gyro.

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

• 제어로봇시스템학회:학술대회논문집
• /
• 2005.06a
• /
• pp.1211-1215
• /
• 2005

In the GPS attitude determination system, the baseline length constraints can be used efficiently to reduce the search space. It is possible by adopting the assumptions that the baseline length doesn't change and the true baseline length is precisely known. But in real situation, the baseline length might be changed by many reasons and it is impossible to measure the true baseline length because there exists measurement error and antenna phase centre movement. In order to analyze the effect of the baseline length accuracy, one needs to know the relation between the accuracy of the baseline length and success rates of the integer ambiguity resolution. In this paper, the effect of the baseline length accuracy to the integer ambiguity resolution in the attitude determination system is analyzed by empirical method. The results show that the margins in the baseline length accuracy is less than a few cm which implies that one should take great cares when applying the attitude determination system to the flexible structures.

• Journal of Institute of Control, Robotics and Systems
• /
• v.11 no.7
• /
• pp.633-643
• /
• 2005

An unmanned aerial vehicle (UAV) is an aircraft controlled by .emote commands from ground station and/o. pre-programmed onboard autopilot system. A navigation system in the UAV provides a navigation data for a flight control computer(FCC). The FCC requires accurate and reliable position, velocity and attitude information for guidance and control. This paper proposes an ADGPS/INS integrated navigation system for a UAV. The proposed navigation system comprises an attitude determination GPS (ADGPS) receive., a navigation computer unit, and a low-cost commercial MEMS inertial measurement unit(IMU). The navigation algorithm contains a fault detection and isolation (FDI) function fur integrity. In order to evaluate the performance of the proposed navigation system, two flight tests were preformed using a small aircraft. The first flight test was carried out to confirm fundamental operation of the proposed navigation system and to check the effectiveness of the FDI algorithm. In the second flight test, the navigation performance and the benefit of the GPS attitude information were checked in a high dynamic environment. The flight test results show that the proposed ADGPS/INS integrated navigation system gives a reliable performance even when anomalous GPS data is provided and better navigation performance than a conventional GPS/INS integration unit.