• International Journal of Aeronautical and Space Sciences
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• v.8 no.2
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• pp.17-27
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• 2007

Minimum-time attitude maneuvers of three-axis stabilized spacecraft are presented to study the feasibility of using three magnetic torquers perform large angle maneuvers. Previous applications of magnetic torquers have been limited to spin-stabilized satellites or supplemental actuators of three axis stabilized satellites because of the capability of magnetic torquers to produce torques about a specific axes. The minimum-time attitude maneuver problem is solved by applying a parameter optimization method for orbital cases to verify that the magnetic torque system can perform as required. Direct collocation and a nonlinear programming method with a constraining method by Simpson's rule are used to convert the minimum-time maneuver problems into parameter optimization problems. An appropriate number of nodes is presented to find a bang-bang type solution to the minimum-time problem. Some modifications in the boundary conditions of final attitude are made to solve the problem more robustly and efficiently. The numerical studies illustrate that the presented method can provide a capable and robust attitude reorientation by using only magnetic torquers. However, the required maneuver times are relatively longer than when thrusters or wheels are used. Performance of the system in the presence of errors in the magnetometer as well as the geomagnetic field model still good.

• Journal of Astronomy and Space Sciences
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• v.26 no.4
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• pp.651-658
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• 2009

Nonlinear sliding surface design in variable structure systems for spacecraft attitude control problems is studied. A robustness analysis is performed for regular form of system, and calculation of actuator bandwidth is presented by reviewing sliding surface dynamics. To achieve non-singular attitude description and minimal parameterization, spacecraft attitude control problems are considered based on modified Rodrigues parameters (MRP). It is shown that the derived controller ensures the sliding motion in pre-determined region irrespective of unmodeled effects and disturbances.

• Journal of the Korea Institute of Military Science and Technology
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• v.4 no.2
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• pp.147-156
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• 2001

In this paper, a robust attitude control scheme based on Eigenaxis rotation for the spacecraft is proposed. Eigenaxis rotation transforms the attitude of spacecraft to the shortest path and is represented by quaternion. The control law consists of PD-type control part for the nominal system and the robust control part for compensating inertia uncertainty. For the proposed controller, stability analysis is performed and the performance is shown via computer simulation.

• International Journal of Control, Automation, and Systems
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• v.4 no.4
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• pp.480-485
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• 2006

In this paper, the error investigation of a 3-dimensional GPS attitude determination system using the error covariance analysis is given. New efficient formulas for computing the Euler Angle Dilution of Precision (EADOP) are also derived. The formulas are easy to compute and represent the attitude error as a function of the nominal attitude of a vehicle, the baseline configuration and the receiver noise. Using the formula, the accuracy of the Euler angle can be analytically predicted without the use of computer simulations. Applications to some configurations reveal the effectiveness of the proposed method.

• Journal of Institute of Control, Robotics and Systems
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• v.9 no.9
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• pp.736-744
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• 2003

The GPS attitude output 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 information is directly applied to the AGPS/INS integration system, the performance of the system can be rapidly degraded. This paper proposes an AGPS/INS integration system using the triple difference carrier phase observables. The proposed integration system contains a cycle slip detection algorithm, in which inertial information is combined. Computer simulations and van test were performed to verify the proposed integration system. The results show that the proposed system gives an accurate and reliable navigation solution even when the integer ambiguity is not correct and the cycle slip occurs.

• Proceedings of the Korean Institute of Intelligent Systems Conference
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• 1998.06a
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• pp.438-442
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• 1998

The helicopter system is non-linear and complex. Futhermore, because of absence of an accurate mathematical model, it is difficult accurately to control its attitude. But we can control the non-modeled system with the uncertainty and unstructure using the fuzzy control algorithm. Therefore, we apply optimized fuzzy controllers for the control of its elevation angle and azimuth one using expert's intuitions and knowledges. The simulation and experimental results of the hellicopter simulator CE150 with MATLAB shall be introduced.

• Journal of Mechanical Science and Technology
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• v.17 no.12
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• pp.1912-1921
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• 2003

This paper presents a global mode modeling of space structures and a control scheme from the practical point of view. Since the size of the satellite has become bigger and the accuracy of attitude control more strictly required, it is necessary to consider the structural flexibility of the spacecraft. Although it is well known that the finite element (FE) model can accurately model the flexibility of the satellite, there are associated problems : FE model has the system matrix with high order and does not provide any physical insights, and is available only after all structural features have been decided. Therefore, it is almost impossible to design attitude and orbit controller using FE model unless the structural features are in place. In order to deal with this problem, the control design scheme with the global mode (GM) model is suggested. This paper describes a flexible structure modeling and three-axis controller design process and demonstrates the adequate performance of the design with respect to the maneuverability by applying it to a large flexible spacecraft model.

• Journal of Advanced Marine Engineering and Technology
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• v.38 no.10
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• pp.1318-1326
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• 2014

The purpose of a tracking control system is to track a moving target and to find the exact information of the target. If the platform of the tracking control system is equipped on a moving vehicle such as a ship, the tracking control system will treat even the additional platform motion. In order to avoid the complexity comprising the tracking control system, a process to treat the platform motion, named stabilizing system, must be separated from the tracking control system. In this paper, a method to comprise an attitude control system for the platform stabilization is proposed using an adaptive fuzzy control which is applicable to the system with structural and parametric uncertainty. The suggested adaptive fuzzy control algorithm is the 2nd/1st-type indirect adaptive fuzzy control algorithm using the advantages of 1st-type and 2nd-type indirect adaptive fuzzy control algorithm. Several experiments using the implemented stabilizing system are executed for verifying the effectiveness of the suggested method.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.7-10
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• 2006

In order to inject satellites into a target orbit, launch vehicles should have a precise attitude and control system capable of controlling three axises of pitch, yaw and roll. For launch vehicles, there are two types of attitude control system currently in popular use; the first one is a cold gas method, and the other is a liquid propulsion system using a single and dual property propellant. The purpose of this paper is to analyze the characteristics of thrust control system using said propellant, thereby providing for a rationale for its application to the upper stages of launch vehicles, in terms of the simplicity of the system, economics of structure and operation.

• Journal of Advanced Information Technology and Convergence
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• v.8 no.2
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• pp.101-110
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• 2018

This paper introduces feedback linearization (FL) based adaptive sliding mode control (ASMC) effective against ground effects of the quadrotor UAV. The proposed control has the capability of estimation and effective rejection of those effects by adaptive mechanism, which resulting stable attitude and positioning of the quadrotor. As output variables of quadrotor, x-y-z position and yaw angle are chosen. Dynamic extension of the quadrotor dynamics is obtained for terms of roll and pitch control input to be appeared explicitly in x-y-z dynamics, and then linear feedback control including a ground effect is designed. A sliding mode control (SMC) is designed with a class of FL including higher derivative terms, sliding surfaces for which is designed as a class of integral type of resulting closed loop dynamics. The asymptotic stability of the overall system was assured, based on Lyapunov stability methods. It was evaluated through some simulation that attitude control capability is stable under excessive estimation error for unknown ground effect and initial attitude of roll, pitch, and yaw angle of $30^{\circ}$ in all. Effectiveness of the proposed method was shown for quadrotor system with ground effects.