• 한국항공우주학회지
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• 제38권3호
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• pp.243-248
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• 2010

본 연구에서는 인공위성의 관성 모멘트 추정을 위해 MRP(Modified Rodrigues Parameter) 자세 변수기반의 추정기를 설계하였다. MRP는 인공위성 자세 결정시 쿼터니 언(Quaternion) 파라미터의 구속 조건으로부터 발생하는 필터의 오차 공분산 행렬의 특이(Singularity) 현상을 피할 수 있는 장점이 있다. 한편 MRP의 경우 자세각 변위가 클 경우에 역시 특이현상이 발생할 수 있어 이를 피하기 위해 적절한 자세각 범위에서 인위적인 기준 운동을 생성하여 필터 설계에 적용하였다. 쿼터니언 파라미터의 단점을 극복하여 보다 안정된 오차 공분산 갱신 결과의 필터의 개선된 성능을 예상할 수 있다.

• Transactions on Control, Automation and Systems Engineering
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• 제3권4호
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• pp.257-261
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• 2001

Nonlinear pulse width modulation (PWM) controlled system is considered to achieve control performance of thruster controlled spacecraft. The actual PWM controlled motions occur, very closely, around the average model trajectory. Furthermore nonlinear PWM controller design can be directly applied to thruster controlled spacecraft to determine thruster on-time. Sliding mode control for attitude tracking of three-axis thruster-controlled spacecraft is presented. Simulation results are shown which use modified Rodrigues parameters and sliding mode control law to achieve attitude tracking of a three-axis spacecraft with thrusters.

• 제어로봇시스템학회:학술대회논문집
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• 제어로봇시스템학회 1999년도 제14차 학술회의논문집
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• pp.41-44
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• 1999

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 Astronomy and Space Sciences
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• 제26권4호
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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.

• International Journal of Aeronautical and Space Sciences
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• 제3권2호
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• pp.1-12
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• 2002

This paper considers the robust and optimal three-axis attitude stabilization of rigid spacecraft with inertia uncertainties. The attitude motion of rigid spacecraft described in terms of either the Cayley-Rodrigues parameters or the Modified Rodrigues parameters is considered. A class of robust nonlinear control laws with relaxed feedback gain structures is proposed for attitude stabilization of rigid spacecraft with inertia uncertainties. Global asymptotic stability of the proposed control laws is shown by using the LaSalle Invariance Principle. The optimality properties of the proposed control laws are also investigated by using the Hamilton-Jacobi theory. A numerical example is given to illustrate the theoretical results presented in this paper.

• 제어로봇시스템학회:학술대회논문집
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• 제어로봇시스템학회 2000년도 제15차 학술회의논문집
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• pp.461-461
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• 2000

Nonlinear pulse width modulation(PWM) controlled system is considered to achieve control performance of thruster-controlled spacecraft. The actual PWM controlled motions occurs, very closely, around the average model rajectory. Furthermore nonlinear PWM controller design can be directly applied to thruster controlled spacecraft to determine thruster on-time. Sliding mode control for attitude tracking of three-axis thruster-controlled spacecraft is presented. Simulation results are shown which use modified Rodrigues parameters and sliding mode control law to achieve attitude tracking of a three-axis spacecraft with thrusters.

• International Journal of Aeronautical and Space Sciences
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• 제8권1호
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• pp.10-20
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• 2007

For spacecraft attitude control, reaction wheel (RW) steering laws with more than three wheels for three-axis attitude control can be derived by using a control allocation (CA) approach.1-2 The CA technique deals with a problem of distributing a given control demand to available sets of actuators.3-4 There are many references for CA with applications to aerospace systems. For spacecraft, the control torque command for three body-fixed reference frames can be constructed by a combination of multiple wheels, usually four-wheel pyramid sets. Multi-wheel configurations can be exploited to satisfy a body-axis control torque requirement while satisfying objectives such as minimum control energy.1-2 In general, the reaction wheel steering laws determine required torque command for each wheel in the form of matrix pseudo-inverse. In general, the attitude control command is generated in the form of a feedback control. The spacecraft body angular rate measured by gyros is used to estimate angular displacement also.⁵ Combination of the body angular rate and attitude parameters such as quaternion and MRPs(Modified Rodrigues Parameters) is typically used in synthesizing the control command which should be produced by RWs.¹ The attitude sensor signals are usually corrupted by noise; gyros tend to contain errors such as drift and random noise. The attitude determination system can estimate such errors, and provide best true signals for feedback control.⁶ Even if the attitude determination system, for instance, sophisticated algorithm such as the EKF(Extended Kalman Filter) algorithm⁶, can eliminate the errors efficiently, it is quite probable that the control command still contains noise sources. The noise and/or other high frequency components in the control command would cause the wheel speed to change in an undesirable manner. The closed-loop system, governed by the feedback control law, is also directly affected by the noise due to imperfect sensor characteristics. The noise components in the sensor signal should be mitigated so that the control command is isolated from the noise effect. This can be done by adding a filter to the sensor output or preventing rapid change in the control command. Dynamic control allocation(DCA), recently studied by Härkegård, is to distribute the control command in the sense of dynamics⁴: the allocation is made over a certain time interval, not a fixed time instant. The dynamic behavior of the control command is taken into account in the course of distributing the control command. Not only the control command requirement, but also variation of the control command over a sampling interval is included in the performance criterion to be optimized. The result is a control command in the form of a finite difference equation over the given time interval.⁴ It results in a filter dynamics by taking the previous control command into account for the synthesis of current control command. Stability of the proposed dynamic control allocation (CA) approach was proved to ensure the control command is bounded at the steady-state. In this study, we extended the results presented in Ref. 4 by adding a two-step dynamic CA term in deriving the control allocation law. Also, the strict equality constraint, between the virtual and actual control inputs, is relaxed in order to construct control command with a smooth profile. The proposed DCA technique is applied to a spacecraft attitude control problem. The sensor noise and/or irregular signals, which are existent in most of spacecraft attitude sensors, can be handled effectively by the proposed approach.