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Prof Hiroshi Yamakawa 

Hiroshi Yamakawa was born in Geneva, Switzerland in 1965. He received master's degree of engineering in 1990 and Ph. D in 1993 both from the University of Tokyo. He became a research associate and an associate professor at the Institute of Space and Astronautical Science in 1993-2003, and an associate professor at Japan Aerospace Exploration Agency in 2003-2006. He engaged in mission design of numerous Earth-orbiting scientific satellite projects as well as in lunar and interplanetary missions. He also engaged in the navigation, guidance, and control systems of the solid propellant rocket, M-V, and liquid propellant reusable sounding rocket, RVT. He was a study manager and a project manager of the Euro-Japan collaborative mission to Mercury "BepiColombo" from 2000 through 2006. He was a visiting scientist at NASA JPL in 1997-1998 and at ESA ESTEC in 2002. He moved to Kyoto University in 2006 as a professor of the Research Institute of Sustainable Humanosphere, a professor of the Graduate School of Engineering (cooperating chair) and a deputy director of the Unit for Synergetic Studies of Space. He was appointed as secretary general at the Secretariat of Strategic Headquarters for Space Policy, Cabinet Secretary, Government of Japan in July 2010 through July 2012. He was assigned member of the Committee for National Space Policy, Cabinet Office in July 2012. His academic interest lies in orbital mechanics (spacecraft formation dynamics, solar sail dynamics, Halo orbits), trajectory optimization (interplanetary trajectory design, low-thrust trajectory optimization), and space propulsion (magnetic sail, solar sail, Coulomb force, Lorentz force orbit control).

Recent Advances in Asteroid Manipulation Technologies
Some kinds of asteroid deflection methods are proposed and investigated. An electric solar wind sail, which uses the natural solar wind dynamic pressure to generate the propulsive acceleration, is employed to accelerate the projectile to impact against the asteroid. The relation between the attitude of the spacecraft and the thrust force is modelled and the dynamics is formulated. In addition, locally optimal control laws are generated. The problem is studied in an optimal framework by maximizing the achievable deflection distance of the asteroid, based on the global optimization procedure. The result shows that the kinetic impact of the spacecraft using electric solar wind sail can achieve the larger impact velocity and the deflection distance. The electrostatic tractor, which tows the asteroid by means of electrostatic force produced by artificially charging of the asteroid and the spacecraft. The electric potential distribution is computed and the results shows the effective shielding length is more than 10 times larger than the Debye length. The dynamics of the charged spacecraft around the charged asteroid is also formulated based on the rotating reference frame. The results of the numerical simulation show that the electrostatic force can extend the achievable deflection distance. It is also shown that towing the asteroid by multiple spacecraft can extend the deflection distance.
Keywords: Near Earth Asteroid Deflection Missions, Electric Solar Wind Sail, Electrostatic Tractor

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