The ability to gather and process information regarding the condition of forest fires is essential to cost-effective, safe, and efficient fire fighting. Advances in sensory and autopilot technology have made miniature unmanned aerial systems (UASs) an important tool in the acquisition of information. This thesis addresses some of the challenges faced when employing UASs for forest-fire perimeter surveillance; namely, perimeter tracking, cooperative perimeter surveillance, and path planning. Solutions to the first two issues are presented and a method for understanding path planning within the context of a forest-fire environment is demonstrated. Both simulation and hardware results are provided for each solution.

In this thesis, a new guidance method for fighter aircrafts and a new guidance method for missiles are developed. Also, guidance and control systems of the aircraft and the missile used are designed to simulate the generic engagement scenarios between the missile and the aircraft. Suggested methods have been tested under excessive simulation studies. The aircraft guidance method developed here is a rule based missile evasion method. The main idea to develop this method stems from the maximization of the miss distance for an engagement scenario between a missile and an aircraft. To do this, an optimal control problem with state and input dependent inequality constraints is solved and the solution method is applied on different problems that represent generic scenarios. Then, the solutions of the optimal control problems are used to extract rules. Finally, a method that uses the interpolation of the extracted rules is given to guide the aircraft. The new guidance method developed for missiles is formulated by modifying the classical proportional navigation guidance method using the position estimates. The position estimation is obtained by utilization of a Kalman based filtering method, called interacting multiple models.

An autonomous system consisting of an unmanned aerial vehicle (UAV) in cooperation with an unmanned ground vehicle (UGV) is of interest in applications for military reconnaissance, surveillance and target acquisition (RSTA). The basic idea of such a system is to take advantage of the vehicles strengths and counteract their weaknesses. The cooperation aspect suggests that the UAV is capable of autonomously landing on the UGV. A fundamental part of the landing is to localise the UAV with respect to the UGV. Traditional navigation systems utilise global navigation satellite system (GNSS) receivers for localisation. GNSS receivers have many advantages, but they are sensitive to interference and spoofing. Therefore, this thesis investigates the feasibility of autonomous landing in a GNSS-denied scenario. The proposed landing system is divided into a control and an estimation system. The control system uses a proportional navigation (PN) control law to approach the UGV. When sufficiently close, a proportional-integral-derivative (PID) controller is used to match the movements of the UGV and perform a controlled descent and landing. The estimation system comprises an extended Kalman filter that utilises measurements from a camera, an imu and ultra-wide band (UWB) impulse radios. The landing system is composed of various results from previous research. First, the sensors used by the landing system are evaluated experimentally to get an understanding of their characteristics. The results are then used to determine the optimal sensor placements, in the design of the EKF, as well as, to shape the simulation environment and make it realistic. The simulation environment is used to evaluate the proposed landing system. The combined system is able to land the UAV safely on the moving UGV, confirming a fully-functional landing system. Additionally, the estimation system is evaluated experimentally, with results comparable to those obtained in simulation. The overall results are promising for the possibility of using the landing system with the presented hardware platform to perform a successful landing.

No / This paper presents an investigation into the challenges in implementing a hard real-time optimal non-stationary system using general regression neural network (GRNN). This includes investigation into the dynamics of the problem domain, discretisation of the problem domain to reduce the computational complexity, parameters selection of the optimization algorithm, convergence guarantee for real-time solution and off-line optimization for real-time solution. In order to demonstrate these challenges, this investigation considers a real-time optimal missile guidance algorithm using GRNN to achieve an accurate interception of the maneuvering targets in three-dimension. Evolutionary Genetic Algorithms (GAs) are used to generate optimal guidance training data set for a large missile defense space to train the GRNN. The Navigation, Constant of the Proportional Navigation Guidance and the target position at launching are considered for optimization using GAs. This is achieved by minimizing the. miss distance and missile flight time. Finally, the merits of the proposed schemes for real-time accurate interception are presented and discussed through a set of experiments. (C) 2012 Elsevier Ltd. All rights reserved.

Missile guidance and autopilot have been active fields of research since the second world war. There are lots of literature on the subjects, but the bulk of which are confined to overly simplified models, and therefore the publications of the methods applied to more realistic models are scarce. In this report a nonlinear 6 DOF simulation model of a tail-controlled air-to-air missile is considered. Through several assumptions and simplifications a linearized approximation of the plant is obtained, which then is used in the implementation of 5 guidance laws and 2 autopilots. The guidance laws are all based on a linearized collision geometry, and the autopilots are based on model predictive control (MPC). Both autopilots use linear quadratic MPC (LQMPC), and one is more robust to modelling errors than the conventional LQMPC. The guidance laws and autopilots are then evaluated with respect to performance in terms of miss distance in 4 interception scenarios with a moving target. The results show that the in this model the autopilots perform equally well, and that the guidance laws with more information about the target generally exhibit smaller miss distances, but at the cost of a considerably larger flight time for some scenarios. The conclusions are that the simplifying assumptions in the modelling are legitimate and that the challenges of missile control probably does not lie in the guidance or autopilot, but rather in the target tracking. Therefore it is suggested that future work include measurement noise and process disturbances in the model. / Det har forskats kring styrlagarna och styrautomaterna för robotar sedan an-dra världskrigets. Det finns mycket litteratur på områdena, men merparten av de publicerade resultaten behandlar enbart grovt förenklade modeller, och därför är tillgången på publikationer där metoderna applicerats i en mer realistisk modell begränsat. I denna rapport behandlas en olinjär simuleringsmodell av en jaktrobot som styrs med stjärtfenor och har sex frihetsgrader. Genom en rad antaganden och förenklingar erhålls en linjäriserad modell av missilen, vilket sedan används för implementering av fem styrlagar och två styrautomater. Styr-lagarna är alla baserade på en linjäriserad kollisionsgeometri och styrautomaterna är baserade på modellprediktiv styrning (MPC). Båda styrautomaterna använder linjärkvadratisk MPC, där den ena påstås vara mer robust gentemot modellfel. Styrlagarna och -automaterna utvärderas ur ett prestandaperspektiv med fokus på bomavstånd i fyra realistiska genskjutningsscenarier med ett rörligt mål. Resultaten visar att båda styrautomaterna presterar lika bra, och att de styrlagar med mer information om målets position/hastighet/acceleration generellt presterar bättre, men att de för vissa skjutfall får en väsentligt längre flygtid. Slutsatserna är att förenklingarna och antagandena i linjäriseringen är välgrundade, och att utmaningarna i missilstyrning inte ligger i utformning av styrlag/-automat, utan förmodligen i målsökningen. Därför föreslås det slutligen att framtida arbete bl. a. inkluderar mätbrus och störningar.

Guidance is defined as the determination of a strategy for following a nominal path in the presence of o-nominal conditions, disturbances and uncertainties, and the strategy employed is called a guidance law. Variants of Proportional Navigation (PN), such as True Proportional Navigation (TPN) and Pure Proportional Navigation (PPN), have been studied extensively in the literature on tactical missile guidance. In the absence of target maneuvers, in a linear interceptor guidance problem, TPN was shown to be optimal. However, the standard PN class of guidance laws per se does not show good performance against maneuvering targets, and was found to be eective in intercepting a maneuvering target only from a restrictive set of initial geometries. Also, since these guidance laws were eectively designed for lower speed targets, they show a degraded performance when applied against higher speed targets. However, in the current defense scenario, two classes of agile targets, which are capable of continuous maneuver, and/or of much higher speed than the interceptor, are a reality. This thesis presents analysis of several variants of PN class of guidance laws against these two classes of agile targets. In the literature, an augmentation of the TPN guidance law, termed as Augmented Proportional Navigation (APN), was shown to be optimal in linearized engagement framework. The present work proposes an augmentation of the PPN guidance law, which is more realistic than TPN for an aerodynamically controlled interceptor, and an-alyzes its capturability in fully nonlinear framework, and develops sauciest conditions on speed ratio, navigation gain and augmentation parameter to ensure that all possible initial engagement geometries are included in the capture zone when applied against a target executing piecewise continuous maneuver. The thesis also obtains the capture zone in the relative velocity space for augmented PPN guidance law. In the literature, a novel guidance law was proposed for the interception of higher speed targets in planar engagement by using a negative navigation gain instead of the standard positive one, and was termed as Retro-PN. It was shown that even though the Retro-PN guided interceptor takes more time than PN guided one in achieving successful interception, Retro-PN performs significantly better than the classical PN law, in terms of capturability, lateral acceleration demand, and closing velocity, when used against higher speed targets. The thesis analyzes Retro-PN guidance law in 3-D engagement geometries to yield the complete capture zone of interceptors guided by Retro-PN guidance philosophy, and derives necessary and sucient conditions for the capture of higher speed non-maneuvering targets with and without a constraint on finiteness of lateral acceleration. Terminal impact angle control is crucial for enhancement of warhead eectiveness. In the literature, this problem has been addressed mostly in the context of targets with lower speeds than the interceptor. The thesis analyzes the performance of a composite PN guidance law, that uses standard PPN and the Retro-PN guidance laws based on initial engagement geometry and requirement of impact angle, against higher speed non-maneuvering targets. Then, to expand the set of achievable impact angles, it proposes a modified composite PN guidance scheme, and analyzes the same. For implementation of many modern guidance laws, a good estimate of time-to-go is essential. This requirement is especially severe in case of impact time constrained en-gagement scenarios. To this end, an ecient and fast time-to-go estimation algorithm for generic 3-D engagement is required. Two time-to-go estimation algorithms are presented and analyzed in this work for the engagement of a PPN or Retro-PN guided interceptor and a higher speed target. The first one is a closed form approximation of time-to-go in terms of range, nominal closing speed and an indicator of heading error, and the second one is a numerical recursive time-to-go estimation algorithm. To improve the odds of intercepting an intelligent target and destroying it, a salvo attack of two or more interceptors could be considered as a viable option. Moreover, this simultaneous salvo attack can also be further improved in eciency by incorporating the shoot-look-shoot approach in making a decision about launching interceptors. This can be considered as the first step towards a layered defense system, which has been described in the literature as a potentially eective strategy against short range or long range ballistic threat. To this end, the present work proposes two PPN and Retro-PN based guidance strategies for achieving simultaneous salvo attack on a higher speed non-maneuvering target. For the implementation of the same the numerical recursive time-to-go estimation technique proposed in this work is utilized

Ce mémoire de thèse traite du développement des dynamiques et des lois de commande de vol d’un projectile d’artillerie gyrostabilisé guidé par une tête découplée. Un modèle non linéaire du projectile est proposé, et sert à calculer un modèle linéarisé de la dynamique de roulis du nez et un modèle q-LPV des chaînes de tangage et de lacet à paramètres fortement variants. Les incertitudes de modélisation sont prises en compte pour concevoir l’autopilote. Des propriétés importantes des projectiles gyrostabilisés, qui sont liées au couplage dynamique tangage/lacet, aux modes internes et à la stabilité, sont mises en valeur grâce au modèle q-LPV. En vue de l’utiliser pour calculer une loi de commande, la dimension de son vecteur des paramètres est réduite et la position des capteurs intégrés dans le nez est considérée. Un seul correcteur linéaire est suffisant pour la dynamique de l’angle de roulis du nez alors qu’une stratégie systématique de commande par séquencement de gains basée sur une linéarisation est élaborée séparément pour générer un correcteur séquencé des facteurs de charge de tangage et de lacet. Des structures de commande fixées d’ordre réduit sont conçues en appliquant la même approche de synthèse linéaire H∞ par façonnage de gain de boucle pour les axes de roulis et de tangage/lacet. De très bonnes propriétés de performance et de robustesse en boucle fermée, comparables à celles fournies par des correcteurs d’ordre plein, sont obtenues. Finalement, l’efficacité de l’autopilote augmenté d’une loi de guidage par navigation proportionnelle pure est vérifiée via de nombreuses simulations non linéaires de trajectoires. Ces dernières correspondent à des scénarios de vol nominaux d’interception de cibles balistiques, non balistiques immobiles, ou manœuvrantes, ainsi qu’à des scénarios considérant des perturbations sur les conditions de tir ou sur les dynamiques du projectile guidé / This thesis addresses the development of the flight dynamics and control laws for an artillery spin-stabilized projectile equipped with a decoupled guidance nose. A projectile nonlinear model is discussed, and it is used for computing a linearized model of the nose roll dynamics along with a q-LPV model of the highly parameter-varying pitch/yaw-dynamics. Modeling uncertainty is taken into account for autopilot design. Important properties specific to spin-stabilized projectiles, which are relevant to pitch/yaw-channel cross-coupling, internal modes and stability, are highlighted using the q-LPV model. In order to use the latter for calculating a control law, the dimension of its parameter vector is reduced and the position of the nose-embedded sensors is considered. A single linear controller is sufficient for the nose roll angle dynamics whereas a systematic linearization-based gain-scheduled control strategy is separately devised to provide a pitch/yaw-axis load factor gain-scheduled controller. Controllers of reduced-order fixed structures are computed by applying the same H∞ linear design loop-shaping approach for the roll and pitch/yaw-axes. Very good closed-loop performance and robustness properties, which are similar to those provided by full order controllers, are obtained. Finally, the effectiveness of the autopilot augmented by a pure proportional navigation guidance law is verified through a variety of nonlinear trajectory simulations. The latter correspond to nominal flight scenarios with ballistic, non-ballistic stationary, and maneuvering interception points, and to scenarios with perturbed launch conditions or guided projectile dynamics

DYNAMIC MODELING, GUIDANCE, AND CONTROL OF HOMING MISSILES &Ouml / ZKAN, B&uuml / lent Ph. D., Department of Mechanical Engineering Supervisor: Prof. Dr. M. Kemal &Ouml / ZG&Ouml / REN Co-Supervisor: Dr. G&ouml / 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 / 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 thesis proposes and evaluates a new cooperative guidance law called General Vector Explicit - Impact Time and Angle Control Guidance (GENEX-ITACG). The motivation for GENEX-ITACG came from an explicit trajectory shaping guidance law called General Vector Explicit Guidance (GENEX). GENEX simultaneously achieves design specifications on miss distance and terminal missile approach angle while also providing a design parameter that adjusts the aggressiveness of this approach angle. Encouraged by the applicability of this user parameter, GENEX-ITACG is an extension that allows a salvo of missiles to cooperatively achieve the same objectives of GENEX against a stationary target through the incorporation of a cooperative trajectory shaping guidance law called Impact Time and Angle Control Guidance (ITACG). ITACG allows a salvo of missile to simultaneously hit a stationary target at a prescribed impact angle and impact time. This predetermined impact time is what allows each missile involved in the salvo attack to simultaneously arrived at the target with unique approach angles, which greatly increases the probability of success against well defended targets. GENEX-ITACG further increases this probability of kill by allowing each missile to approach the target with a unique approach angle rate through the use of a user design parameter. The incorporation of ITACG into GENEX is accomplished through the use of linear optimal control by casting the cost function of GENEX into the formulation of ITACG. The feasibility GENEXITACG is demonstrated across three scenarios that demonstrate the ITACG portion of the guidance law, the GENEX portion of the guidance law, and finally the entirety of the guidance law. The results indicate that GENEX-ITACG is able to successfully guide a salvo of missiles to simultaneously hit a stationary target at a predefined terminal impact angle and impact time, while also allowing the user to adjust the aggressiveness of approach.