The authors have been developing unique low gravity devices applying yo-yo principle. The paper reports a new model which improves drawbacks clarified by a series of experiments of the original model named Ez-Space. The basic ideas of the design are described with analyses of its dynamics. Its performances are verified by preliminary experiments of the new model Ez-Space II .

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The authors have developed a new low gravity device applying yo-yo principle Ez-Space II which is an improved version of the original device. Its design principle and basic characteristics are presented at HASTIC Academic Symposium in Japanese. Also, more details are reported in a paper included in this volume of JSTS. This paper explains results of miscellaneous educational experiments conducted principally at TAMAROKUTO Science Center in Tokyo. Some additional information of the device is also stated.

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KAGUYA, which was called the SELenological and ENgineering Explorer (SELENE) in its development phase, is Japan’s first full-scale lunar explorer. KAGUYA was launched on September14, 2007, reached to a lunar orbit on October 3, 2007, and established a circular polar lunar observation orbit at an altitude of 100 km on October 18, 2007. KAGUYA necessitates a reliable attitude and orbit control system (AOCS) for a reliable lunar polar orbit insertion and an accurate lunar-center-pointing three-axis attitude control for high precision observations from lunar orbit. In addition, KAGUYA has to realize various different control mode before reaching the lunar orbit: sun acquisition, inertial-frame-pointing attitude control in lunar transfer orbit, attitude control at orbital maneuvers, attitude control for daughter satellite separations, and lunar-center-pointing attitude control for lunar observation. In fact, the AOCS must realize these different requirements for different phases using the limited resources of KAGUYA satellite. Considering these system requirements and technical challenges for the AOCS, we designed and developed the simple but reliable and accurate KAGUYA AOCS. This paper first shows an outline of KAGUYA’s system and mission profile, then describes the KAGUYA AOCS and its control modes, and finally presents AOCS operation results during the lunar transfer phase and the lunar orbit injection phase.

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This paper reports results of a preliminary study of an on-orbit advanced robust attitude control experiment using a Japanese large flexible communication satellite: the Engineering Test Satellite VIII (ETS-VIII). The on-orbit attitude control experiment is designed to develop a baseline technology for a class of future large flexible spacecraft and earth observation satellites which require attitude maneuvers using advanced control technology. An important issue is the design of a reduced order multi-input multi-output (MIMO) robust controller for a linear parameter varying (LPV) spacecraft system. We therefore prepared three control laws after tradeoff studies. Among them, in this paper, a linear time invariant (LTI)-μ controller is presented; it is designed parametrically by decomposing the s/c model into control modes and residual modes. In μ-synthesis, the varying parameters are treated as structured uncertainty that is stabilized robustly with unstructured uncertainty by setting the μ-value to less than one. The final controller is LTl-designed through D-K iteration. The candidate experiment controllers were evaluated using preliminary MATLAB-based laboratory design and synthesis and full software simulation of onboard algorithms with emulation of the AOCS hardware. This report presents the outline of this experiment and some spacecraft model features, along with a case study of the proposed method for ETS-VIII and numerical simulation results.

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A feedforward algorithm for flexible spacecraft maneuvers is presented. We call this preshaping profiler the nil-mode-exciting (NME) profiler-improved. This algorithm is designed particularly for single-axis rest-to-rest rotational maneuvers (switching maneuvers) with linear actuators. Generally spacecraft with large flexible structure has a lot of large-mass flexible-modes. Therefore, uncertainty of high-order modes needs to be considered for high-accuracy controller design. This paper presents an extra-insensitive maneuvering method which overcomes above-mentioned difficulties. This algorithm includes a preshaping profiler formulated from sampling function (also known as sine function), consequently feedforward control inputs generated from the preshaping profiler have no frequency response above a certain designed frequency. Therefore residual vibration at the end-point of maneuver can be highly reduced with minimum loss of maneuver agility.

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A Magnetic Bearing Wheel (MBW) with inclined magnetic poles has been developed, which allows the fabrication of a 5-DOF active magnetic bearing with six electromagnets and six displacement sensors, leading to a reduction in the size and weight of the MBW system. The MBW controller consists of two modules: a magnetic bearing controller, which controls the rotor displacements by PI-D control, and a disturbance feedback controller, which reduces the MBW disturbance. A design of the magnetic bearing controller is based on the geometric analysis of the electromagnetic forces and magnetic bearing forces/moments. The magnetic bearing controller utilizes the redundancy of electromagnetic forces in order to satisfy certain restrictions on the magnetically suspended system and to optimize the electromagnetic forces used for biasing. On the other hand, the disturbance feedback controller generates rotor displacement commands that are capable of ensuring a reduction in the MBW disturbance. The magnetic bearing control system with the disturbance feedback controller has extremely good stability as the stability can be controlled by applying an appropriate phase lead to the displacement commands. Our control system enables tP,e reduction in the MBW disturbance to 0.1 Nrms/0.05 Nmrms or less for all rotor speeds up to 6000 rpm.

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Recently, JAXA initiated the study of a new type of earth observation satellite which can select observation areas by changing the attitude of the satellite. Because this kind of satellite requires a high-torque attitude control system, a control moment gyro (CMG) is considered as an actuator candidate. The maximum torque produced by a CMG can be dozens of times the torque produced by a reaction wheel. Because the torque vector of a CMG changes according to the gimbal angle, careful consideration must be given to singularity configurations. In the singularity configurations, the compound torque vector from CMGs is restricted to two dimensions even in typical four-skew redundant CMG configurations. For that reason, a satellite loses tri-axial attitude control capability. Many researchers of CMG singularity avoidance problems have studied this problem and created various CMG steering logics. We must verify CMG control logics not only through numerical simulation but also through experimentation by using CMG hardware to develop satellite-mounted CMGs. Since 2005 we have studied CMG singularity avoidance logic in numerical simulations using Matlab/Simulink. Then in 2006 we started developing CMG Software Evaluation Equipment . We developed an air bearing that comprises a ball and a saucer to demonstrate tri-axial free motion like that of a satellite in orbit. This ball floats in the saucer with air and rotates freely on three axes. An aluminum square pipe is attached over the air bearing as a balance. Four CM Gs are mounted on one side of the balance. The MPU, the gyro, and the battery are mounted on the other side of the balance. Furthermore, we are developing a CMG Test Bed for use as ground test equipment to perform the Dynamic Close Loop Test with CMG EM and Attitude and Orbit Control System (AOCS) Hardware. An especially ambitious undertaking is to mount experimental electronic devices in a ball to eliminate the limit (roll, pitch, yaw). The ball must be large. And we produced aluminum hemispheres with a radius of2000 mm.

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