氏名　パルマン　セティヤマルタナ
学位の種類　博士（工学）
学位記番号　博甲第196号
学位授与の日付　平成11年6月30日
学位論文の題目　Attitude Maneuvers of a Flexible Satellite by Using Input Shapers（適切なトルク入力を用いた柔軟構造衛生の姿勢制御）
論文審査委員主査
　助教授　古口　日出男
　副査　教授　武藤　睦治
　副査　教授　長井　正嗣
　副査　教授　矢鍋　重夫
　副査　助教授　永澤　茂

• 論文目次
• 論文要旨

［平成11（1999）年度博士論文題名一覧］ ［博士論文題名一覧］に戻る．

1 Introduction　p.1
　1.1 The Area of Investigation　p.1
　1.2 The Purpose of the Thesis　p.4
　1.3 The Organization of the Thesis　p.4

2 Equations of Motion of a Spacecraft with Flexible Appendages　p.15
　2.1 Introduction　p.15
　2.2 Coordinate Reference Frame Definitions　p.16
　2.3 The Mathematical Modelling of Flexible Spacecraft Dynamics　p.22
　2.4 A Particular Model of a Satellite with Two Flexible Solar Panels　p.46
　2.5 Concluding Remarks　p.59

3 Response of System　p.63
　3.1 Model Analysis　p.63
　3.2 Numerical Integration　p.75
　3.3 Natural Frequencies, Mode Shapes, and Eigenvectors　p.94
　3.4 Concluding Remarks　p.101

4　Attitude Maneuvers of a Satellite with Flexible Solar Panels under Bang-Bang Commands　p.105
　4.1 Introduction　p.105
　4.2 Slew Maneuvers of the Flexible Satellite　p.105
　4.3 Rest-to-Rest Attitude Maneuvers of the Satellite under Bang-Bang Torque Inputs　p.107
　4.4 Discussion　p.117
　4.5 Concluding Remarks　p.117

5 Input Shaper to Reduce Residual Vibration in Rest-to-Rest Slew Maneuvers　p.119
　5.1 Introduction　p.119
　5.2 Residual Vibration and Vector Diagrams　p.121
　5.3 General Multi-Mode On-Off Input Shaper　p.135
　5.4 Multi-Mode On-Off Input Shaper for Attitude Angle Maneuvers　p.140
　5.5 Concluding Remarks　p.140

6 Rest-to-Rest Attitude Maneuvers of the Flexible Satellite under Shaped Torque Inputs　p.147
　6.1 Introduction　p.147
　6.2 Rest-to-Rest Attitude Maneuvers under Time-Optimal Shaped Commands　p.148
　6.3 Rest-to-Rest Attitude Maneuvers under Fuel-Efficient Shaped Commands　p.158
　6.4 Attitude Maneuvers of the Precise-Oriented Flexible Satellite by Using Shaped Torque Inputs　p.166
　6.5 Discussion and Concluding Remarks　p.172

7 Conclusion　p.177

　Appendix A　p.181

　List of Publications　p.187

　Attitude maneuvers of flexible spacecraft, from one rest condition to another rest condition, are studied. Equations of motion of the spacecraft are developed by employing a hybrid system of coordinates and Lagrange's formulation. The finite element method is used to discrete elastic deformations of the flexible substructures. For the particular study, a new finite element model of satellite consisting of a rigid main body and two symmetrical flexible solar panels, such as the Japan ETS-VI, is proposed. Each solar panel is modeled as a number collection of rectangular plate elements. Only out-of-plane deformations of the elements are considered. The satellite is supposed to be equipped with on-off reaction jets on the main body, so that it cannot produce the variable-amplitude actuation inputs. The satellite must be moved with constant amplitude force or torque pulses. Also, it is supposed that the solar panels have no damping properties and no controls on them.
　A satellite is designed to have a certain orientation, for example, with respect to the earth. For the correct orientation, the attitude of the satellite must be reoriented after reaching a geostationary orbit and needs frequent corrections during its operation. Attitude maneuvers of the flexible satellite without regarding to system flexibility can generate large amplitude vibrations, especially when the system is equipped with on-off jets. In the simulations under the bang-bang torque inputs, the satellite shows poor performances. The large-amplitude oscillations of the attitude angles and vibrations of the solar panels are resulted after the inputs were removed. The amplitude of the resultant residual oscillation in some cases can be larger than the desired angle displacement.
　In the operation, a communication satellite must have a certain precision of its antenna beam pointing. To maintain the precise pointing, when the flexible satellite is maneuvered, the oscillations of its attitude at the end of the maneuvers must be smaller than permissible values of pointing error of the antenna beam. To generate input profiles such that the satellite is rotated through a desired attitude angle, while the residual oscillation is within the permissible value, an input shaper method is proposed in this thesis. The method is implemented by convolving a sequence of impulses with a desired system command to produce a shaped input that then is used to drive the system. The amplitudes and time locations of the impulses are determined by solving a set of constraint equations that attempt to control the dynamic response of the system.
　The finite element model of satellite used in this thesis has a lot of natural frequencies. In the use of shaped inputs, when a vibration at one natural frequency is suppressed, other frequencies could be amplified at the end of maneuver. Then, the resultant residual oscillations of the attitude angles are still greater the required pointing precisions of satellite. So, to maneuver the satellite with residual attitude angle oscillations within the permissible levels, the shaped input applied must be selected by reducing large amplitude vibrations potentially resulted by the system.