LARGE ANGLE MANOEUVRE OF AN UNDERACTUATED SMALL SATELLITE USING TWO WHEELS Nadjim Mehdi Horri Stephen Hodgart Surrey Space Centre, University of Surrey, Guildford, GU2 7XH, UK Phone:+44 1483 682692 , Fax: +44 1483689503, email: n.horri@eim.surrey.ac.uk ABSTRACT Failure of mechanical controllers onboard a satellite is a well-known phenomenon that has already been disastrous for several space missions. In the case studied here, we take as an example the mini-satellite UoSat-12 built by SSTL at University of Surrey. The Z- axis reaction wheel of this satellite has failed and the 3-axis control performance using magnetorquing is being very limited. The use of thrusters would involve undesired fuel consumption. As an alternative, we present here some latest theory, which shows how full 3-axis control can still be achieved, using the two remaining reaction wheels from a standard orthogonal 3-wheel configuration. Using a novel nonlinear time invariant and discontinuous approach, we show that the attitude is precisely and rapidly altered, without transient oscillations, to the required earth pointing. The case of a small non- zero total momentum is illustrated. One consequence of these results is that a fully redundant 3-axis control can be envisaged using a 3-wheel configuration. 1. INTRODUCTION It is now well known that only non-smooth control laws can be developed to stabilize nonholonomic systems, including underactuated satellites [2]. In this paper, we focus on the case of a satellite controlled using only two reaction wheels in the event of one wheel failure. Using two wheels, the only 3-axis stabilizing control law in the literature is a time varying control law that ensures the stabilization for a zero total momentum satellite, but the system goes through important transient oscillations [1]. In this paper , we show how a discontinuous control approach can be used to achieve stability and large angle maneuvers for underactuated satellites. NOTATIONS: I B A : Direction cosine matrix from inertial reference to body frame I = [ ] T I I I 3 2 1 , , : Inertia tensor of the body of the satellite about it’s centre of mass.