• Transactions of the Korean Society of Automotive Engineers
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• v.22 no.1
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• pp.65-70
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• 2014

An Integrated Dynamics Control system with four wheel Steering (IDCS) is proposed and analysed in this study. It integrates and controls steer angle of front and rear wheel simultaneously to enhance lateral stability and steerability. An active front steer (AFS) system and an active rear steer (ARS) system are also developed to compare their performances. The systems are evaluated during brake maneuver and several road conditions are used to test the performances. The results showed that IDCS vehicle follows the reference yaw rate and reduces side slip angle very well. AFS and ARS vehicles track the reference yaw rate but they can not reduce side slip angle. On split-${\mu}$ road, IDCS controller forces the vehicle to go straight ahead but AFS and ARS vehicles show lateral deviation from centerline.

• International Journal of Aeronautical and Space Sciences
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• v.3 no.2
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• pp.13-23
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• 2002

In this paper, we present a sliding mode control strategy for the re-orientation maneuver of rigid spacecraft containing rotating wheels. The wheels are considered as internal devices, and external inputs are employed for generation of control commands. The formulation is developed for a general case while particular example is applied to pitch bias momentum spacecraft with a single momentum wheel. The resultant control commands are used to take the gyroscopic effects into account which are caused by the rotating wheels. The controller designed demonstrates that the nutational motion of the pitch bias momentum spacecraft is effectively controlled. It is also assumed that the external control torque device is of on-off nature, and pulse width modulation technique is applied to construct proper control torque history.

• Journal of Astronomy and Space Sciences
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• v.1 no.1
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• pp.23-34
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• 1984

A bubble trapped in the liquid manifold of INTELSAT IV F-7 spacecraft caused a mass imbalance between the System 1 propellant tanks and a wobble half angle of 0.38 degree to 0.48 degree. A maneuver on May 14, 1980 passed the bubble through the axial jet and allowed propellant to redistribute. A 0.2 rpm change in sin rate was observed with an exponential decay time constant of 6 minutes. In this paper, moment of inertia, tank geometry and hydrodynamic models are derived to match the observed spin rate data. The values of the total mass of propellant considered were 16, 19 and 20 kgs with corresponding mass imbalances of 14.3, 15 and 15.1 Kgs, respectively. The result shows excellent agreement with observed spin rate data but it was necessary to assume a greater mass of hydrazine in the tanks than propellant accounting indicated.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.18 no.3
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• pp.27-33
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• 2010

The goal for the CD&R system is to predict that a conflict is going to occur in the future, communicate the detected conflict to a human operator, and, in some cases, assist in the resolution of the conflict situation. To provide insight into different methods of conflict detection and resolution, a literature review of previous research models and current developmental and operational systems was performed. This paper focuses only on the specific attributes of each model, not on the depth to which a model has been analyzed, validated, or accepted. Thus, care should be taken to remember that a model that seems to be simple according to our categorization scheme may be significantly more viable than an apparently sophisticated model.

• International Journal of Aeronautical and Space Sciences
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• v.7 no.2
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• pp.14-25
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• 2006

This paper is focused on designing an implementable control law to perform spacecraft various missions using momentum exchange devices such as reaction wheels(RWs) and control moment gyros(CMGs). A compact equation of motion of a spacecraft installed with various momentum exchange devices is derived in this paper. A hybrid control law is proposed for precision attitude control of agile spacecraft. The control law proposed in this paper allocates control torque to the CMGs and the RWs adequately to satisfy the precision attitude control and large angle maneuver simultaneously. The saturation problem of reaction wheels and the singularity problem of control moment gyros are considered. The problems are successfully resolved by using the proposed hybrid closed loop control law. Finally, the proposed hybrid control law is demonstrated by numerical simulations.

• International Journal of Aeronautical and Space Sciences
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• v.11 no.2
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• pp.87-97
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• 2010

A fault tolerant satellite attitude control scheme with a modified iterative learning law is proposed for dealing with actuator faults. The actuator fault is modeled to reflect the degradation of actuation effectiveness, and the solar array-induced disturbance is considered as an external disturbance. To estimate the magnitudes of the actuator fault and the external disturbance, a modified iterative learning law using only the information associated with the state error is applied. Stability analysis is performed to obtain the gain matrices of the modified iterative learning law using the Lyapunov theorem. The proposed fault tolerant control scheme is applied to the rest-to-rest maneuver of a large satellite system, and numerical simulations are performed to verify the performance of the proposed scheme.

• International Journal of Aeronautical and Space Sciences
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• v.18 no.2
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• pp.300-314
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• 2017

This paper examines finite rotation angle (FRA) applications for spacecraft attitude control. The coordinate transformation matrix and the attitude kinematics represented by FRAs are introduced. The interpolation techniques for the angular orientations are thoroughly investigated using the FRAs and the results are compared to those using traditional methods. The paper proposes trajectory description techniques by using extremely smooth polynomial functions of time, which can describe point-to-point attitude maneuvers in a realizable and accurate manner with the help of unique FRA features. In addition, new controller design techniques using the FRAs are developed by combining the proposed interpolation techniques with a model predictive control framework. The proposed techniques are validated through their attitude control applications for an aggressive point-to-point maneuver. Conclusively, the FRAs provide much more flexibility than quaternions and Euler angles when describing kinematics, generating trajectories, and designing attitude controllers for spacecraft.

• Bulletin of the Korean Space Science Society
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• 2003.10a
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• pp.47-47
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• 2003

Analytical formulae are presented to approximate the evolution of the semi major axis, the maneuver time, and the final mass fraction for low thrust orbital transfers with circular initial orbit, circular target orbit, and constant thrust directed either always along or always opposite the velocity vector. For comparison, the associated results for high-thrust transfers, i.e. the two-impulse Hohmann transfer, are summarized. All results are implemented in a computer code designed to analyze planar planetary and interplanetary space missions. This implementation yields fast and reasonably accurate approximations to trajectory performance boundaries. Consequently, the approach can provide trajectory analysis for each spacecraft configuration during the conceptual space mission design phase. As an example, a mission from Low-Earth Orbit (LEO) to Jupiter's moon Europa is analyzed.

• Journal of Astronomy and Space Sciences
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• v.15 no.1
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• pp.245-250
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• 1998

Attitude determination and control of satellite is important component which determines the accomplish satellite missions. In this study, attitude control using reaction wheels and momentum dumping of wheels are considered. Attitude control law is designed by Sliding control and LQR. Attitude maneuver control law is obtained by Shooting method. Wheels momentum dumping control law is designed by Bang-Bang control. Four reaction wheels are configurated for minimized the electric power consumption. Wheels control torque and magnetic moment of magnetic torquer are limited.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.82-83
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• 2002

궤도상에 올려진 위성들은 인형의 궤도 운행을 하게 된다. 그러나 지구가 완전한 구형이 아니고 태양과 달의 인력이 작용하여 위성에 섭동이 발생하게 된다. 그리고 무중력 상태의 우주이므로 태양풍이나 미세 운석 그리고 위성체 내부의 가스 누출이나 내부의 토크 변화에 의해 위성 자세에 조금의 변동을 야기한다. 통신 위성의 경우 지상의 한 지점을 계속 향하고 있어야 하므로 정기적인 자세제어가 필요하다. 위성의 섭동에 의해 EWSK(East-West station keeping)나 NSSK(North-South station keeping)를 하기 위해 추력 모델은 단일 $\Delta$$\upsilon$기동이나 회전 세차 운동(spin precession maneuver)을 지원해야 한다. 위성은 주어진 임무를 수행하는데 필요한 $\Delta$$\upsilon$기동을 위해 적절한 성능의 추력기와 임무기간 동안 사용할 적절한 양의 추진제를 탑재하고 있다. 지상에서 필요한 임무를 수행하기 위해 위성에 지령을 하였을 때, 추력기가 정상작동을 하였는지 그리고 잔류 추진제가 어느 정도 인지를 정확히 알 수 있어야 한다.