This study addresses a sliding-mode-based contour-following controller design for guidance and autopilot systems of launch vehicles with highly maneuverable actuators, mainly consisting of the thrust vector control and side jet systems, to perform the trajectory-tracking task given a predetermined trajectory. In particular, the trajectory is pre-designed by the method of calculus of variations with focus on maximizing the final velocity transferring to the orbit to obtain the optimal trajectory for a launch vehicle throughout the whole course. Chiefly, an integrated guidance/autopilot controller is designed to achieve the main goal of robust tracking issues for launch vehicles in real time such that the relative motion between a vehicle and the predetermined trajectory is minimized. As for the attitude control, an autopilot system is designed not only to stabilize the attitude of the launch vehicle but also to realize the guidance law of the translational motion control. In order to establish a complete system model, besides the motion dynamics of launch vehicles, we also take into account several influential factors such as the propellant effort of a movable nozzle thrust vector control and side jet systems, aerodynamic influence, the Earth’s gravitational field, and wind gusts/shears. The overall stability of the integrated guidance/autopilot system is assured via rigorous analysis by Lyapunov stability theory, and its corresponding performance is verified by numerical simulations.
This investigation addresses a nonlinear terminal guidance/autopilot controller with pulse-type control inputs for intercepting a theater ballistic missile in the exoatmospheric region. Appropriate initial conditions on the terminal phase are assumed to apply after the end of the midcourse operation. Accordingly, the terminal controller seeks to minimize the distance between the commanded missile and the target missile to ensure a hit-to-kill interception. In particular, a 3D terminal guidance law is initially developed to eliminate the so-called "sliding velocity," thus, constraining the relative motion between the missile and the target along the line of sight. Sliding mode control is adopted to design stable pulse-type control systems. Then, a quaternion-based attitude controller is used to orient appropriately the commanded missile, taking into account the fact that the missile is a rigid body, to realize interceptability. The stability of the overall integrated terminal guidance/autopilot system is then analyzed thoroughly, based on Lyapunov stability theory. Finally, extensive simulations are conducted to verify the validity and effectiveness of the integrated controller with the pulse type inputs developed herein.
In this brief, we propose a variable structure based nonlinear missile guidance/autopilot system with highly maneuverable actuators, mainly consisting of thrust vector control and divert control system, for the task of intercepting of a theater ballistic missile. The aim of the present work is to achieve bounded target interception under the mentioned 5 degree-of-freedom (DOF) control such that the distance between the missile and the target will enter the range of triggering the missile's explosion. First, a 3-DOF sliding-mode guidance law of the missile considering external disturbances and zero-effort-miss (ZEM) is designed to minimize the distance between the center of the missile and that of the target. Next, a quaternion-based sliding-mode attitude controller is developed to track the attitude command while coping with variation of missile's inertia and uncertain aerodynamic force/wind gusts. The stability of the overall system and ZEM-phase convergence are analyzed thoroughly via Lyapunov stability theory. Extensive simulation results are obtained to validate the effectiveness of the proposed integrated guidance/autopilot system by use of the 5-DOF inputs.
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