• 제어로봇시스템학회:학술대회논문집
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• 1994.10a
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• pp.44-48
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• 1994

A new method of quaternion feedback control for the attitude acquisition of spacecraft is suggested to limit the angular rates of rigid body which are not desirable and make a control algorithm complicate. New attitude acquisition control algorithm is evaluated and compared with the existing quaternion feedback control method for the large slewing maneuvers through simulations. The simulation results reveal that a new method is effective on limiting the angular rates of spacecraft.

• Journal of Mechanical Science and Technology
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• v.18 no.2
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• pp.203-210
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• 2004

Attitude control law design for spacecraft large angle maneuvers is investigated in this paper. The feedback linearization technique is applied to the design of a nonlinear tracking control law. The output function to be tracked is the quaternion attitude parameter. The designed control law turns out to be a combination of attitude and attitude rate tracking commands. The attitude-only output function, therefore, leads to a stable closed-loop system following the given reference trajectory. The principal advantage of the proposed method is that it is relatively easy to produce reference trajectories and associated controller.

• Proceedings of the KIEE Conference
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• 2007.10a
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• pp.70-71
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• 2007

This paper presents an application of LQR(Linear Quadratic Regulator) for experimental control moment gyroscope. To be specific, mathematical model is first derived based on the quaternion and Lagrange's equation, state feedback controller using LQR scheme is designed, and to show the stability of the scheme, experimental results are given.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.50 no.6
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• pp.393-400
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• 2022

The Divert and Attitude Control System(DACS) used in high-altitude engagements is expensive and complex. In this paper, we design a high-altitude terminal guidance and control loop of guided-missile equipped with a Thrust Vector Control(TVC) that is less expensive and simpler than DACS. The proposed system utilizes a quaternion feedback control technique to track the thrust attitude command converted from the acceleration command of true proportional navigation guidance. The performance analysis of the proposed terminal guidance and control loop is conducted through engagement simulations against ballistic targets at a high altitude.

• Journal of Institute of Control, Robotics and Systems
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• v.8 no.4
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• pp.313-318
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• 2002

A robust attitude controller using Lyapunov redesign technique for spacecraft is proposed. In this controller, qua- ternion feedback is considered to have the attitude maneuver capability very close to the eigen-axis rotation. The controller consists of three parts: the nominal feedback parts which is a PD-type controller for the nominal system without uncertainties, the additional term compensating for the gyroscopic motion, and the third part for ensuring robustness to uncertainties. Lyapunov stability criteria is applied to stability analysis. The performance of the proposed controller is demonstrated via computer simulation.

• 제어로봇시스템학회:학술대회논문집
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• 1988.10a
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• pp.633-638
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• 1988

In this paper, a zero-velocity update technique to improve navigation accuracy of a SDINS(Strapdown Inertial Navigation System) has been studied. An indirect feedback Kalman filter which includes SDINS error equations based on a quaternion between body-fixed frame and local level navigation frame is employed for processing zero-velocity updates in an on-board navigation filter. Simulation results for land-mobile vehicle show that the zerovelocity update technique make a significant contribution to improving SDINS performance without any external aids.

• 제어로봇시스템학회:학술대회논문집
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• 1995.10a
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• pp.115-118
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• 1995

This paper present an attitude control using quaternions as feedback attitude errors. The Euler's eigenaxis rotation provides the shortest angular path between two attitudes. This eigenaxis rotation can be achieved by using quaternions since quaternions are related with the eigenaxis. The suggested controller uses error quaternions and body angular rates and generates a decoupling control torque that counteracts the natural gyroscopic coupling torque. The momentum dumping strategy using the earth magnetic field is also applied in this paper to unload the angular momentum of the reaction wheels used in the attitude control.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.2
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• pp.129-138
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• 2021

Due to the active progress in space programs for various types of ground and space missions, the high agile spacecraft maneuverability is also required. To meet the requirement of the given space missions, the Control Moment Gyros (CMG) for the alternatives of the classical reaction wheels can release the attitude maneuverability restrictions. In addition, the angular rates of the spacecraft is constrained due to the limited actuator characteristics. In this paper, a sliding mode control technique for the attitude control of the spacecraft equipped with the pyramid type of CSCMG(Constant Speed CMG) is designed, and the stability of the control system is guaranteed by using the Lyapunov stability theory. Finally, the control law proposed is analyized by numertical simulations.

• International Journal of Aeronautical and Space Sciences
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• v.8 no.1
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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.