Kenji Kitamura
(University of Tokyo, Japan)

In order to determine the attitude of interplanetary spacecraft, it is popular to use sun-sensor and star-sensor. However, star-sensor is not a feasible solution for low cost missions. With respect to the interplanetary spinner spacecraft, this paper describes a simpler attitude determination system utilizing Sun angle and Earth angel information. This system is adopted by IKAROS, Japanese interplanetary solar sail demonstration mission. The Sun angle is provided by the sun sensor, while the Earth angle is provided by Doppler modulation of the RF signal frequency measured at ground stations. Hence the transmitting antenna should be off-centered from the spin axis to realize this system. In this method, ground system and orbit determination system are closely related to each other, and this causes some problems peculiar to the method. In this study, theoretical analysis is conducted so as to evaluate such problems. Moreover, on-orbit data of IKAROS is used and compared to the analysis.

Motoshi Kawasaki
(Yokohama National University, Japan)

Spacecrafts are required to land precisely on a specified area in order to acquire important data for scientific research from the moon. In this paper, high-accuracy "Polynomial Guidance Law" is introduced to realize safe and pinpoint lunar landing. The purpose of the guidance system is to make a soft landing automatically. The severe requirement on weight is given. The fuel consumption must be minimized. The solution of optimal control problem is usually given by solving two-point boundary value problem. A large amount of calculation is necessary to solve the problem using iterative calculation for optimization. The Polynomial Guidance Law does not need iterative calculation, thus computational load can be saved significantly. In addition, the flight path designed by the guidance law accurately satisfies the assigned location of landing, because the guidance law is derived from polynomial equations, which are defined as directional distance function respectively. The robustness is also required, because there are a lot of uncertainties in system parameters and initial conditions. The results of numerical simulation show the optimality and robustness of the guidance law.

Misuzu Haruki
(Yokohama National University, Japan)

CMG (Control moment Gyro) system is able to generate high torque compared to other actuators and the system can control attitude of satellite for the minimum-time maneuver. But the optimum-solutions for minimum-time maneuver are calculated by iterative calculation, which has two problems that initial assumption is needed and calculation cost is high. Therefore the object of the study is to design an on-line and optimum control law; it gives input for minimum-time maneuver with an on-line processing without any initial assumption. The optimum-solutions have bang-bang input for minimum-time with one singular input to keep large angular momentum. The new control law is designed using rule-based feed-forward control and feed-back control. In the feed-forward control the attitude is controlled using rule of three steps. Step1 chooses a CMG of singular input using relation between an initial state of the system and maneuver-axis. Step2 gives amplitude of singular input using maneuver-axis, and Step3 gives switching-time of the other bang-bang input. Some error of final attitude is modified by feedback control. The on-line control law based on rules achieves to control the satellite with CMG system in minimum-time.

Yoshinobu Okano
(Tokyo Metropolitan University, Japan)

The spacecraft using solar radiation pressure for navigation recently gets a lot of attention. Interplanetary Kite-Craft Accreted by Radiation Of the Sun, IKAROS is one of the these spacecrafts. IKAROS launched on May 23, demonstrate solar sailing which mean accelerated and navigated by solar radiation pressure. IKAROS already demonstrated acceleration of solar radiation and also experimented in spin axis attitude control by using Reflectivity Control Device, RCD which is capable of changing its reflectance according to applied voltage. Because spacecraft usually receives sun light on orbit, it is enable to navigate by using solar radiation pressure without propellant and high power consumption. This method is efficient to guidance spacecraft and satellites. In this paper, we propose spin axis control of general spinning spacecraft and satellite. IKAROS usually spinning and receiving solar radiation pressure, have equilibrium point on spin axis which means spin axis was navigated some point. This equilibrium point is depending on shape and geometry of spacecraft so that IKAROS which capable of changing reflectance by RCD, can control balance point of spin axis. We study controlling this equilibrium point not only IKAROS but general spinning spacecrafts and satellites using reflectivity control devices.