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In this paper, optimal guidance laws for multiple vehicles formation flight are addressed. The proposed laws are obtained as state-feedback solutions of linear quadratic optimal control problems to minimize the time varying energy cost with terminal constraints on position and velocity. Since the proposed laws are able to control the flight time, they are possible to make all vehicles join the formation concurrently at a specified time. A method to compensate the acceleration or maneuver of the formation leader is included in the guidance laws. And an inter-vehicle collision avoidance scheme is also considered. The performance of the guidance laws is tested via various numerical simulations.

The increased demand for high maneuverability of modern missiles requires excellent autopilot performance over a large flight envelope. The main difficulty in missile autopilot design is the influence of high angle-of-attack aerodynamic phenomena on stability characteristics (Hemsch, 1992). As a missile executes a high angle-of-attack maneuver, the forebody vortices can become asymmetric and give rise to significant lateral-forces and yawing moments. Classically, missile autopilots are designed using gain scheduling approaches in roll, pitch, and yaw channels (Blakelock, 1991;Shamma and Athans, 1992;Shamma and Cloutier, 1993;White et al., 2007). However, the performance of singleaxis autopilot design approaches are limited within some flight boundaries because aerodynamic coupling effects and their parameter dependencies in large angle-of-attack aerodynamics could not be considered. In this context, the design of a three-axis missile autopilot has been studied using linear and nonlinear control approaches (Choi et al., 2008;Devaud et al., 2001;Kim et al., 2008). These methods are based on gain scheduling techniques by local linearization, and demonstrate some improvements from a practical point of view.In this paper, a three-axis missile autopilot design with a multi-objective output-feedback control theory is presented. Because high angle-of-attack aerodynamics is highly nonlinear and estimates of the aerodynamic coefficients are very imprecise, a mix of H∞ and H2 criteria is employed to guarantee the performance of the three-axis missile autopilot. This control methodology, which uses linear matrix inequality (LMI) techniques, was motivated by the work in (Schere et al., 1997). This paper is organized as follows. In Section 2, we give an overview of multi-objective output-feedback control theory. Section 3 describes our formulation of the missile control problem. Section 4 describes strategies and numerical simulation results for the designed autopilot. Our conclusions are provided in Section 5. AbstractWe report on the design of a three-axis missile autopilot using multi-objective control synthesis via linear matrix inequality techniques. This autopilot design guarantees H 2 /H ∞ performance criteria for a set of finite linear models. These models are linearized at different aerodynamic roll angle conditions over the flight envelope to capture uncertainties that occur in the high-angle-of-attack regime. Simulation results are presented for different aerodynamic roll angle variations and show that the performance of the controller is very satisfactory.

In this paper, we propose a guidance synthesis method for enhancing anti-ship missiles' survivability against ship-borne CIWS (close-in weapon system). Using CEALM (coevolutionary augmented Lagrangian method), a direct optimization technique, an optimal control problem to minimize time-varying weighted sum of the inverse of aiming errors of CIWS is solved. The optimal evasive trajectory exhibits sinusoidal acceleration commands, which results in barrel-roll type evasive maneuvers. Inspired by the optimization results, a 3-dimensional biasedproportional navigation guidance (PNG) law to induce a barrel-roll maneuver during the homing phase is proposed.Capturability of the proposed guidance law is proved by using the Lyapunov stability theory. A proper choice of the barrel-roll direction also guarantees that the missile altitude can be made lower bounded. Performance of the proposed guidance laws is compared with conventional PNG via simulations.

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