Academic literature on the topic 'Guidance, control, missile'

DYNAMIC MODELING, GUIDANCE, AND CONTROL OF HOMING MISSILES &Ouml<br>ZKAN, B&uuml<br>lent Ph. D., Department of Mechanical Engineering Supervisor: Prof. Dr. M. Kemal &Ouml<br>ZG&Ouml<br>REN Co-Supervisor: Dr. G&ouml<br>kmen MAHMUTYAZICIOgLU September 2005, 236 pages In this study, the dynamic modeling, guidance, and control of a missile with two relatively rotating parts are dealt with. The two parts of the missile are connected to each other by means of a roller bearing. In the first part of the study, the governing differential equations of motion of the mentioned missile are derived. Then, regarding the relative rotation between the bodies, the aerodynamic model of the missile is constructed by means of the Missile Datcom software available in T&Uuml<br>BiTAK-SAGE. After obtaining the required aerodynamic stability derivatives using the generated aerodynamic data, the necessary transfer functions are determined based on the equations of motion of the missile. Next, the guidance laws that are considered in this study are formulated. Here, the Linear Homing Guidance and the Parabolic Homing Guidance laws are introduced as alternatives to the Proportional Navigation Guidance law. On this occasion, the spatial derivation of the Proportional Navigation Guidance law is also done. Afterwards, the roll, pitch and yaw autopilots are designed using the determined transfer functions. As the roll autopilot is constructed to regulate the roll angle of the front body of the missile which is the controlled part, the pitch and yaw autopilots are designed to realize the command signals generated by the guidance laws. The guidance commands are in the form of either the lateral acceleration components or the flight path angles of the missile. Then, the target kinematics is modeled for a typical surface target. As a complementary part of the work, the design of a target state estimator is made as a first order fading memory filter. Finally, the entire guidance and control system is built by integrating all the models mentioned above. Using the entire system model, the computer simulations are carried out using the Matlab-Simulink software and the proposed guidance laws are compared with the Proportional Navigation Guidance law. The comparison is repeated for a selected single-body missile as well. Consequently, the simulation results are discussed and the study is evaluated.

This dissertation investigates advanced concepts in terminal missile guidance. The terminal phase of missile guidance usually lasts less than ten seconds and calls for very accurate maneuvering to ensure intercept. Technological advancements have produced increasingly sophisticated threats that greatly reduce the effectiveness of traditional approaches to missile guidance. Because of this, terminal missile guidance is, and will remain, an important and active area of research. The complexity of the problem and the desire for an optimal solution has resulted in researchers focusing on simplistic, usually linear, models. The fruit of these endeavors has resulted in some of the world's most advanced weapons systems. Even so, the resulting guidance schemes cannot possibly counter the evolving threats that will push the system outside the linear envelope for which they were designed. The research done in this dissertation greatly extends previous research in the area of optimal missile guidance. Herein it is shown that optimal missile guidance is fundamentally a pairing of an optimal guidance strategy and an optimal control strategy. The optimal guidance strategy is determined from a missile's information constraints, which are themselves largely determined from the missile's sensors. The optimal control strategy is determined by the missile's control constraints, and works to achieve a specified guidance strategy. This dichotomy of missile guidance is demonstrated by showing that missiles having different control constraints utilize the same guidance strategy so long as the information constraints are the same. This concept has hitherto been unrecognized because of the difficulty in developing an optimal control for the nonlinear set of equations that result from control constraints. Having overcome this difficulty by indirect means, evidence of the guidance strategy paradigm emerged. The guidance strategy paradigm is used to develop two advanced guidance laws. The new guidance laws are compared qualitatively and quantitatively with existing guidance laws.

Today the armed forces of many countries need to strike accurately potential enemies, wherever they might be, from a safe place. Since naval units can be deployed almost everywhere in the open sea, the idea of a naval cruise missile emerged in the 70’s. These missiles are designed to be launched from various naval vehicles such as frigates or submarines and strike deeply in the enemy territory. A program called Missile de Croisière Naval (MdCN Naval Cruise Missile) was therefore launched in 2006 by the DGA, the French procurement agency. MBDA is the industrial company appointed by the DGA to design and build the missile. Control aspects on a cruise missile are of primary interest since they impact the reliability, performance and availability of the weapon. In the aeronautics and weapon industry, gain scheduled controllers are used in most cases. However, many non-linear techniques have  been developed in the literature and might improve the behaviour of the missile. The main objective of the present thesis is to apply non-linear techniques on the control and guidance loops of the MdCN too see whether of not they can improve such a system. Based on this report it should be easy for the engineers of the DGA to compare the controllers of the thesis and the classical gain scheduled controllers used in the industry. To achieve this task some basic knowledge of flight dynamics are recalled and a model of the MdCN is computed and divided into the control loop and the guidance loop. Then a non-linear controller for the launch phase using a Lyapunov based technique called back-stepping is designed and tested through a statistic analysis. During the cruise phase different anti-windup strategies are applied on the propulsion control loop of the missile and compared. Finally a software interface with a navigation-dedicated tool is coded and implemented in Simulink to analyse the complete Guidance-Navigation-Control loop and to see how navigation errors impact the control algorithms. The main contributions of this thesis are the controllers designed for the launch phase and the propulsion loop that will be compared with the controller that MBDA is going to deliver next year to see whether or not the non-linear techniques used in the thesis should be used on the missile. Furthermore, all the tools and procedure set up to interface the control and guidance laws with the navigation models and filters will give the possibility to the DGA to have a deeper understanding of the algorithms used by MBDA and to make sure that navigation and estimation issues are properly taken in account when designing the control and guidance laws.

Thesis (M.S. in Electrical Engineering) Naval Postgraduate School, March 2000.<br>Thesis advisor(s): Hutchins, Robert G. "March 2000." Includes bibliographical references (p. 83). Also available online.

The focus of this study is to be able to control the air-to-surface missile throughout the entire flight, with emphasis on the propulsion phase to increase the impact range of the missile. A major difficulty in controlling the missile during the propulsion phase is the important change in mass of the missile. This results in sliding the center of gravity (cg) point and changing inertias. Moreover, aerodynamic coefficients and stability derivatives are not assumed to be constant at predetermined ranges<br>conversely, they depend on Mach number, angle of attack, and side slip angle. Consequently, as the change of missile mass, cg point, inertia terms, and stability and aerodynamic coefficients come together apart from flight operation stages, a great number of points need to be taken into account when designing the controller. This makes controlling the missile all the more complicated. In this thesis, first the equations of motion are derived, in which, mass of the missile is not assumed constant. Thus, not only the variation of mass but also the variation of inertias is incorporated in the equations of motion. From the derived v equations of motion, a nonlinear inverse dynamics controller that can achieve desired guidance for a conceptually developed air-to-surface missile has been designed, tested and verified for a modeled missile with six degrees of freedom. For brevity of the study, conceptual design and aerodynamic calculations are not given in detail. Nevertheless, improvements for conceptual design are suggested. As a result, it is shown that the controller works efficiently: the missile is able to hit the target with less than 12 m circular error of probability (CEP). Finally, studies and improvements are proposed.

This thesis involves modeling, guidance, control, and flight simulations of a canard controlled guided missile. The autopilot is designed by a pole placement technique. Designed autopilot is used with the guidance systems considered in the thesis. Five different guidance methods are applied in the thesis, one of which is the famous proportional navigation guidance. The other four guidance methods are different fuzzy logic guidance systems designed considering different types of guidance inputs. Simulations are done against five different target types and the performances of the five guidance methods are compared and discussed.