Possibility of orbit control using gravity gradient (GG) effects without any mass expulsion is discussed. For simplicity, a dumb-bell type satellite and circular orbits are mainly considered. It is shown that the GG effects can be applied to convert attitude torques into orbital torques and vice versa. In central gravitational force fields, maximum orbital torques or thrusts are available from the GG force when roll or pitch angle is ±π/4 provided that attitude angle is null when the dumb-bell axis coincides with local vertical. Such external torques as geomagnetic or solar wind pressure can be utilized to maintain the ±π/4 attitude, then orbital torques are availableforever. In non-central gravitational fields, without any external torque, the orbital radii of circular orbits can be increased by controlling the satellite attitude using electric energy. The use of the earth's oblateness effects and the exterior Lunar potential is discussed. Numerical examples are presented.
As a method of active flutter suppression, controlling a leading-edge control surface is analytically studied for a two-dimensional wing in the use of the oscillating incompressible aerodynamic force. Four quantities, i.e. rotation and translation of the wing and their velocities, are fedback so as to move the leading-edge control surface. The control law proportional to wing rotation is most effective in flutter suppression independently of the location of the center of gravity. Therefore this feed-back system is only one method of flutter suppression in the case where the center of gravity is behind a certain point, for example, 45% chord. The mechanism of the present flutter suppression is clearly explained by an approximate analysis in the use of the steady aerodynamic force.
The homing missile is fired after the target is locked on by the seeker, but the "fire and forget" missile is fired toward the collision point and during flight the seeker locks on the target, and then provides its own homing guidance. The flight between firing and the seeker looks on is called midcourse guidance. Midcourse flight path can be taken in any curve but the straight line is preferable for it gives the shortest flight time before lock on, thus minimizing the errors caused by dispersion, bias and target maneuver. This paper describes details of a straight line autopilot. The direct rate gyro output is used as the stabilization of the missile, and the integrated output is used as the attitude reference which keeps the attitude angle constant thus keeping the flight path straight. This autopilot can be modified from the present inventory by replacing rate gyros with ones having better static balance etc. and the addition of an integrating circuit to the rate gyro output.