Engineering Test Satellite VI (ETS-VI) is a three-axis-stabilized satellite launched by National Space Development Agency of Japan (NASDA) in August 1994. With a large and light solar paddles, ETS-VI is considered as a flexible satellite system from the control of view. Identification and attitude control experiments on orbit were planned as one of the bus experimental programs for ETS-VI, and carried out from December 1994 to March 1995. In this paper, some of the on-orbit attitude control experiments are presented. Attitude controllers were designed using I-PD control method based on the model obtained through analysis on the ground. According to the experiments, the controllers provided sufficient control performance which was perfectly predicted in advance through the simulation results.

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When a robot arm is mounted on a satellite to perform some tasks, the satellite attitude must be stabilized to retain the communication link and to generate electrical power from its solar panels. It is not realistic to control the total system as one dynamic system, since the number of degrees of freedom becomes too large to be handled by state-of-the-art satellite mounted computer. The proposed coordinated control concept is realized by the real-time compensation of the robot arm’s reaction by the satellite attitude control system and the robot arm motion control by the robot arm controller not to generate the excess reaction against the satellite attitude stability. Stability of the satellite attitude is guaranteed by the proposed coordinated control.

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The problem of impact dynamics of space robotic systems that consist of a rigid manipulator supported by a flexible deployable structure is addressed. Due to joint back-drivability and the dynamic coupling between the manipulator and its sup-porting structure, unknown motion of the system occurs after it makes impulsive contact with the environment. A method that uses the system’s dynamic model is proposed to estimate the motion of the system after impact and used to minimize the impact effect and vibrations of the supporting structure. This method is verified experimentally using the MIT Vehicle Emulation System (VES II). The experimental results show that the impact force and the system motion after impact can be reduced if the manipulator configuration prior to impact and the controller gains are properly selected.

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Effects of spherically asymmetric terms of the Phobos gravity en trajectories for Phobos exploration in close proximity such as hovering close to the surface of Phobos, horizontal movement in hovering condition, and trajectories connecting the L1 and L2 points of the Mars-Phobos system are analyzed in this paper. Harmonic coefficients of the gravitational potential of Phobos obtained by Chao et al. are used in this analysis. The required velocity increment ΔV through the hovering of 7.66 hours, which is the period of Phobos orbit around the Mars, is obtained over the whole regions of Phobos. The minimum value of ΔV was about 82 m/s at the sub-Mars and the anti-sub-Mars points in the point-mass approximation. For the gravity with asymmetric terms it is about 120 m/s at the anti-sub-Mars point. The maximum value of ΔV was about 301 m/s at poles in the point-mass approximation. For the gravity with asymmetric terms it ls about 340 m/s at the neighborhood of the Stickney crater. In the horizontal movement in hovering condition, it turns out that the optimum movement velocity is not so affected by the asymmetric terms of the Phobos gravity. A trajectory connecting the L1 and L2 points is obtained even in the gravity model with asymmetric terms. The obtained trajectory is also asymmetric. The minimum altitude of the orbit, the transit time, and the necessary velocity increment are about 1.5 km, 3.4 hours, and 21 m/s, respectively. From the error analysis, it is found that in the case of the symmetric trajectory of the point-mass model, the sensitivity to the initial velocity error ls very high, while not so high in the case of the asymmetric trajectory considering the harmonics of the gravity potential.

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Space exploration missions for heavenly bodies generally call a phase of orbital operations in close proximity to the body. Moons and asteroids have irregular shapes and may considerably disturb the periodic orbit of spacecraft. This paper aims at clarifying the relationships between the orbit stability and the irregular shapes of the body in a pure rotation. First, the general gravitational field is expanded in terms of spherical harmonics, whose 1st and 2nd coefficients are shown to correspond to the mass parameters of the body. Secondly, the analytic approximations for the equations of motion are derived for the planar motion using a proposed perturbation method. The true anomaly is chosen as an independent variable instead of time to utilize the periodicity of the motion. Finally, numerical simulations are given to validate the derived analytic approximations, where a dumbbell model is used as a representative gravitational model for an asteroid. The main results are : 1) in a direct orbit, when the orbital period is close to the rotation period, the perturbations become large and 2) in a retrograde orbit, the perturbations are always small.

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