achieving variation of the centre of solar pressure, and hence the solar moment, for satellite attitude control. Formulation Consider a satellite with an axisymmetric central rigid body carrying a pair of large but light rectangular panels in an elliptical orbit at geostationary altitudes (Fig. 1). In the nominal, straight and symmetrically deployed undeformed configuration of the panels, their centre of mass (c.m.) is assumed to coincide with the point SR representing the c.m. of the rigid body so that no net solar torque is generated. However, bending of the panels brought about by differential solar heating may cause significant shift in the center of solar pressure, thus giving rise to a resultant moment on the satellite. The principal coordinate axes of the satellite in the undeformed configuration are denoted by x, y, z, with panels deployed symmetrically about the .s-axis. As the panels undergo deformation due to heating, their c.m. Sp and hence the overall satellite c.m. represented by S moves along the a-axis. The principal coordinate axes for the satellite in the deformed configuration are denoted by X, Y, Z with origin at S, with the axis x parallel to the X axis and the z and Z axes coinciding. Using the well-known Lagrangian procedure, recognising that, in general, the natural frequency of appendage vibrations is likely to be much higher than the orbital frequency [7] and assuming relatively small panel deformations, the pitching motion can be modelled by the following differential equation:
RkJQdWJsaXNoZXIy MTU5NjU0Mg==