Dissertations / Theses on the topic 'Guided missiles Guidance systems'

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.

A near-optimal guidance law has been developed using the direct method of calculus of variations that maximizes the kinetic energy transfer from a surface-launched missile upon interception to a ballistic missile target during the boost phase of flight. Mathematical models of a North Korean Taep'o-dong II (TD-2) medium-range ballistic missile and a Raytheon Standard Missile 6 (SM-6) interceptor are used to demonstrate the guidance lawâ s performance. This law will utilize the SM-6â s onboard computer and active radar sensors to independently predict an intercept point, solve the two-point boundary value problem, and determine a near-optimal flight path to that point. Determining a truly optimal flight path would require significant computing power and time, while a near-optimal flight path can be calculated onboard the interceptor and updated in real time without significant changes to the interceptorâ s hardware. That near-optimal guidance path is then converted into a set of command functions and fed back into the control computer of the interceptor. By modifying the second and third derivatives of the two-point boundary value problem, the intercept conditions can be varied to study their effects upon the optimal flight path regarding the maximization of kinetic energy upon impact.
US Navy (USN) author.

Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2010.
Committee Chair: Mark Costello; Committee Member: Eric Johnson; Committee Member: Frank Fresconi; Committee Member: Olivier Bauchau; Committee Member: Peter Plostins. Part of the SMARTech Electronic Thesis and Dissertation Collection.

Optimal midcourse guidance is examined for an air-to-air missile featuring boost-coast-sustain propulsion. A vertical plane, point-mass model is studied with load factor as a control variable. Time-range-energy optimal trajectories are computed, open-loop, via the usual necessary conditions and a multiple-shooting algorithm. A requirement on terminal velocity magnitude is examined for its effect on firing range. Next, a study of the optimal midcourse guidance problem with reduced-order models is presented. The models under study, in addition to the point-mass model, are: - Singularly perturbed model with y as fast variable; - Point mass model with approximation of the induced-drag; - Energy model. One of the major results in this study is that the reduced-order models are not accurate enough to approximate the optimal trajectories and so are of limited use as reference trajectories in an on-board scheme. Thus, optimal trajectories, computed by using the point-mass model, are selected as the reference trajectories for a closed-loop guidance scheme. Finally, an approach to on-board real-time calculations for an optimal guidance approximation is derived. Extremal fields and neighboring extremal theory ideas are used together with pre-calculated Euler solutions to construct a closed-loop guidance algorithm. The method is applied to the midcourse guidance of an air-to-air missile and was found to perform quite well.
Ph. D.

Optimal intercept trajectories for a boost-sustain-coast medium-range air-to-air missile are synthesized using optimal control theory. Optimality in time/range/energy at intercept of a target is the main objective. Attainable sets and their boundaries are obtained and used to generate optimal intercept points in a three-dimensional scenario. A three-phase closed-loop guidance scheme is used to generate an efficient guidance law against a maneuvering target. In the present study, target maneuvers are restricted to the horizontal plane. An initial boost-phase with near-optimal guidance in the presence of active control constraints and thrust switches is simulated. Target maneuvers are neglected during this phase. A new method of gain evaluation is detailed. A midcourse guidance scheme with neighboring guidance, transversal comparisons, and chasing center-of-attainability of target to augment performance is studied. Modifications in terminal guidance using proportional navigation, such as chasing the center-of-attainability of target, altitude shaping, and drag-resolution schemes are used to attempt better performance at intercept. A composite guidance strategy using a combination of neighboring guidance and proportional navigation for the midcourse guidance is introduced. The excellent performance of this guidance strategy and the improvement in storage requirements for on-board use make it a very special scheme.
Ph. D.

Guided missiles are among the most effective threats against air platforms. Aircraft and helicopter losses in the last decades were mostly due to guided missiles, 70% of which were infrared guided missiles. Today, there are as many as 500,000 shoulder-fired missiles in military arsenals around the world, whose guidance algorithms enable them to track the desired trajectories very precisely. In this thesis, main focus is on defining infrared missile guidance and control algorithms in order to study on various target-missile scenarios for the effectiveness of these algorithms. First, a mathematical model of a generic missile is given. Then the flight control system of the missile is created by using LQR and PID controllers. Different kinds of PNG algorithms applied to an infrared missile are presented. The seeker part of the infrared missile is also discussed. The effectiveness of guidance algorithms are studied based on different target &ndash
missile scenarios and the responses of them to IR countermeasures are also observed. This study shows that different guidance algorithms can be used for different scenarios. If suitable algorithms are combined and suitable constants are applied, the guided missile can track the target very precisely. In addition, the seeker part has to be improved with tracking algorithms in order to recognize IR-countermeasures and not to follow them.

An anti-tank guided missile is a weapon system primarily designed to hit and destroy armored tanks and other armored vehicles. Developed first-generation command-guided and second-generation semi-automatic command guided missiles had many disadvantages and lower hit rates. For that reason, third generation imaging infrared fire-and-forget missile concept is very popular nowadays. In this thesis, mainly, a mathematical model for a fire-and-forget anti-tank missile is developed and a flight control autopilot design is presented using PID and LQR techniques. For target tracking purposes, &ldquo
correlation&rdquo
, &ldquo
centroid&rdquo
and &ldquo
active contour&rdquo
algorithms are studied and these algorithms are tested over some scenarios for maximizing hit rate. Different target scenarios and countermeasures are discussed in an artificially created virtual environment.

The method of singular perturbations is applied to the determination of the optimal range-fuel-time trajectory for an air-breathing missile. This method is shown to lead to the reduced-order "cruise-dash" model, and this model is used in the optimization study. Earlier work in this area is extended by the inclusion of two not heretofore considered limits on the dynamical system. The results of the earlier work are shown to hold throughout much of the velocity regime in which the missile operates, but operation in the very high and very low velocity ranges is shown to be sharply curtailed, with the optimal operating points being changed drastically in some cases. Also, the effect of the non-zero minimum admissible throttle setting and the resultant throttle-chattering on the solution of the control problem is examined in some detail.
M.S.

This thesis describes the fuzzy logic based control system for an automated guided vehicle ( AGV ) designed to navigate from one position and orientation to another while avoiding obstacles. A vehicle with an onboard computer system and a beacon based location system has been used to provide experimental confirmation of the methods proposed during this research. A simulation package has been written and used to test control techniques designed for the vehicle. A series of navigation rules based upon the vehicle's current position relative to its goal produce a fuzzy fit vector, the entries in which represent the relative importance of sets defined over all the possible output steering angles. This fuzzy fit vector is operated on by a new technique called rule spreading which ensures that all possible outputs have some activation. An obstacle avoidance controller operates from information about obstacles near to the vehicle. A method has been devised for generating obstacle avoidance sets depending on the size, shape and steering mechanism of a vehicle to enable their definition to accurately reflect the geometry and dynamic performance of the vehicle. Using a set of inhibitive rules the obstacle avoidance system compiles a mask vector which indicates the potential for a collision if each one of the possible output sets is chosen. The fuzzy fit vector is multiplied with the mask vector to produce a combined fit vector representing the relative importance of the output sets considering the demands of both navigation and obstacle avoidance. This is operated on by a newly developed windowing technique which prevents any conflicts produced by this combination leading to an undesirable output. The final fit vector is then defuzzified to give a demand steering angle for the vehicle. A separate fuzzy controller produces a demand velocity. In tests carried out in simulation and on the research vehicle it has been shown that the control system provides a successful guidance and obstacle avoidance scheme for an automated vehicle.

A reduced order model describing the energy and heading angle dynamics of a scramjet missile is developed using a singular perturbation technique. The cruise analysis is briefly reviewed to determine the conditions at which the missile will cruise most efficiently. The turn and climb performance of the missile over the conditions of interest is then examined and a family of extremal trajectories is constructed which asymptotically approach the cruise at an intermediate altitude.
Master of Science

Thesis (M.S.)--Missouri University of Science and Technology, 2010.
Vita. The entire thesis text is included in file. Title from title screen of thesis/dissertation PDF file (viewed April 7, 2010) Includes bibliographical references (p. 70-77).

In this thesis, a multiple hypotheses tracking (MHT) algorithm is developed to successfully track multiple ballistic missiles within the boost phase. The success of previous work on the MHT algorithm and its application in other scientific fields enables this study to realize an efficient form of the algorithm and examine its feasibility in tracking multiple crossing ballistic missiles even though various accelerations due to staging are present. A framework is developed for the MHT, which includes a linear assignment problem approach used to search the measurement-to-contact association matrix for the set of exact N-best feasible hypotheses. To test the new MHT, an event in which multiple ballistic missiles have been launched and threaten the North American continent is considered. To aid in the interception and destruction of the threat far from their intended targets, the research focuses on the boost-phase portion of the missile flight. The near-simultaneous attacks are detected by a network of radar sensors positioned near the missile launch sites. Each sensor provides position reports or track files for the MHT routine to process. To quantify the performance of the algorithm, data from the National Air and Space Intelligence Center's IMPULSE ICBM model is used and demonstrates the feasibility of this approach. This is especially significant to the U.S. Missile Defense Agency since the IMPULSE model represents the cognizant analyst's accurate representation of the ballistic threats in a realistic environment. The results show that this new algorithm works exceptionally well in a realistic environment where complex interactions of missile staging, non-linear thrust profiles and sensor noise can significantly degrade the track algorithm performance especially in multiple target scenarios.

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

In this thesis, the problem of intercepting highly manoeuvrable threats using seeker-less interceptors that operate in the command guidance mode, is addressed. These systems are more prone to estimation errors than standard seeker-based systems. Several non-linear and optimal estimation and guidance concepts are presented in this thesis for interception of randomly maneuvering targets by seeker-less interceptors. The key contributions of the thesis can be broadly categorized into six groups, namely (i) an optimal selection of bank of lters in interactive multiple model (IMM) scheme to cater to various maneuvers that are expected during the end-game, (ii) an innovative algorithm to reduce chattering phenomenon and formulate effective guidance algorithm based on 'differential game guidance law' (modi ed DGL), (iii) IMM/DGL and IMM/modified DGL based integrated estimation/guidance (IEG) strategy, (iv) sensitivity and robustness analysis of Kalman lters and ne tuning of lters in filter bank using innovation covariance, (v) Performance of tuned IMM/PN, tuned IMM/DGL and tuned IMM/modi ed DGL against various target maneuvers, (vi) Performance comparison with realistic missile model. An innovative generalized state estimation formulation has been proposed in this the-sis for accurately estimating the states of incoming high speed randomly maneuvering targets. The IMM scheme and an optimal selection of lters, to cater to various maneu-vers that are expected during the end-game, is described in detail. The key advantage of this formulation is that it is generic and can capture evasive target maneuver as well as straight moving targets in a uni ed framework without any change of target model and tuning parameters. In this thesis, a game optimal guidance law is described in detail for 2D and 3D engagements. The performance of the differential game based guidance law (DGL) is compared with conventional Proportional Navigation (PN) guidance law, especially for 3D interception scenarios. An innovative chatter removal algorithm is introduced by modifying the differential game based guidance law (modified DGL). In this algorithm, chattering is reduced to the maximum extent possible by introducing a boundary layer around the switching surface and using a continuous control within the boundary layer. The thesis presents performance of the modified DGL algorithm against PN and DGL, through a comparison of miss distances and achieved accelerations. Simulation results are also presented for varying fiight path angle errors. Apart from the guidance logic, two novel ideas have been presented following the evolving "integrated estimation and guidance" philosophy. In the rst approach, an in-tegrated estimation/guidance (IEG) algorithm that integrates IMM estimator with DGL law (IMM/DGL), is proposed for seeker-less interception. In this interception scenario, the target performs an evasive bang-bang maneuver, while the sensor has noisy measure-ments and the interceptor is subject to an acceleration bound. The guidance parameters (i.e., the lateral acceleration commands) are computed with the help of zero e ort miss distance. The thesis presents the performance of the IEG algorithm against combined IMM with PN (IMM/PN), through a comparison of miss distances. In the second ap-proach, a novel modi ed IEG algorithm composed of IMM estimator and modi ed DGL guidance law is introduced to eliminate the chattering phenomenon. Results from both of these integrated approaches are quite promising. Monte Carlo simulation results re-veal that modi ed IEG algorithm achieves better homing performance, even if the target maneuver model is unknown to the estimator. These results and their analysis o er an insight to the interception process and the proposed algorithms. The selection of lter tuning parameters puts forward a major challenge for scien-tists and engineers. Two recently developed metrics, based on innovation covariance, are incorporated for determining the filter tuning parameters. For predicting the proper combination of the lter tuning parameters, the metrics are evaluated for a 3D interception problem. A detailed sensitivity and robustness analysis is carried out for each type of Kalman lters. Optimal and tuned Kalman lters are selected in the IMM con guration to cater to various maneuvers that are expected during the end-game. In the interception scenario examined in this thesis, the target performs various types of maneuvers, while the sensor has noisy measurements and the interceptor is subject to acceleration bound. The tuned IMM serves as a basis for synthesis of e cient lters for tracking maneuvering targets and reducing estimation errors. A numerical study is provided which demonstrates the performance and viability of tuned IMM/modi ed DGL based modi ed IEG strategy. In this thesis, comparison is also performed between tuned IMM/PN, tuned IMM/DGL and tuned IMM/modi ed DGL in integrated estimation/guidance scheme. The results are illustrated by an extensive Monte Carlo simulation study in the presence of estimation errors. Simulation results are also presented for end game maneuvers and varying light path angle errors . Numerical simulations to study the aerodynamic e ects on integrated estimation/ guidance structure and its e ect on performance of guidance laws are presented. A detailed comparison is also performed between tuned IMM/PN, tuned IMM/DGL and tuned IMM/modi ed DGL in integrated estimation/guidance scheme with realistically modelled missile against various target maneuvers. Though the time taken to intercept is higher when a realistic model is considered, the integrated estimation/guidance law still performs better. The miss distance is observed to be similar to the one obtained by considering simpli ed kinematic models.

A satellite launch vehicle not subjected to any perturbations, external or internal, could be guided along a trajectory by following a stored, pre-computed steering program. In practice, perturbations do occur, and in order to take account of them and to achieve an accurate injection, a closed loop guidance system is required. Guidance algorithm is developed by solving the optimal control problem. Closed form solution is difficult because the necessary conditions are in the form of Two Point Boundary Value Problems (TBVP) or Multi Point Boundary Value Problems (MPBVP). Development of non-iterative guidance algorithm is taken as a prime objective of this thesis to ensure reliable on-board implementation. If non-iterative algorithms are required, the usual practice is to approximate the system equations to derive closed form solutions. In the present work, approximations cannot be used because the algorithm has to cater to a wide variety of vehicles and missions. Present development adopts an alternate approach by splitting the reconfigurable algorithm development in to smaller sub-problems such that each sub-problem has closed form solution. The splitting is done in such a way that the solution of the sub-problems can be used as building blocks to construct the final solution. By adding or removing the building blocks, the algorithm can be configured to suit specific requirements. Chapter 1 discusses the motivation and objectives of the thesis and gives a literature survey. In chapter 2, Classical Flat Earth (CFE) guidance algorithm is discussed. The assumptions and the nature of solution are closely analyzed because CFE guidance is used as the baseline for further developments. New contribution in chapter 2 is the extension of CFE guidance for a generalized propulsion system in which liquid and solid engines are present. In chapter 3, CFE guidance is applied for a mission with large pitch steering angles. The result shows loss of optimality and performance. An algorithm based on regular perturbation is developed to compensate for the small angle approximation. The new contribution in chapter 3 is the development of Regular Perturbation based FE (RPFE) guidance as an extension of CFE guidance. RPFE guidance can be configured as CFE guidance and FEGP. Algorithms presented up to chapter 3 are developed to inject a satellite in to orbits with unspecified inertial orientation. Communication satellite missions demand injection in to an orbit with a specific inertial orientation defined by argument of perigee. This problem is formulated using Calculus of Variations in chapter 4. A non-iterative closed form solution (Predicted target Flat Earth or PFE guidance) is derived for this problem. In chapter 5, PFE guidance is extended to a multi-stage vehicle with a constraint on the impact point of spent lower stage. Since the problem is not analytically solvable, the original problem is split in to three sub-problems and solved. Chapter 6 has two parts. First part gives theoretical analysis of the sub-optimal strategies with special emphasis to guidance. Behavior of predicted terminal error and control commands in presence of plant approximations are theoretically analyzed for a class of optimal control problems and the results are presented as six theorems. Chapter 7 presents the conclusions and future works.

碩士
國立臺灣大學
電機工程學研究所
88
This thesis discusses the guidance and control system design problem for missiles with high controllability and accuracy. Assume that there is a target ballistic missile (the target) moving very fast inside the atmosphere, and an anti-ballistic-missile missile (the missile) with a lateral jet system is to intercept the target by making a direct collision. The lateral jet system of the missile can produce large lateral force instantaneously so that the missile can make lateral motion quickly. However, the number of lateral jets in the system is limited, so the aerodynamic control system of the missile must be used in the first stage of the intercept mission to bring the missile close to the target. How to integrate these two systems so that they can work together with the guidance system of the missile is the main research subject of this thesis. Our results show that in the first stage, the proportional navigation guidance law can be integrated with the aerodynamic control system, which is based on the linear quadratic optimal control method, to generate proper commands for the control surfaces of the missile. Then, just before the proportional navigation guidance law diverges, the lateral jet system and a zero-effort-miss concept based intelligent guidance law can take over, and accomplish the direct collision intercept mission. Our works include many simulation results for evaluating the performance of the proposed guidance and control system.

Automated Guided Vehicles (AGVs) depend on sensor systems to guide and navigate in manufacturing environments. The sensor system must ensure the correct path is followed and monitor the vehicle’s progress as it proceeds through the environment to complete assigned tasks. An AGV therefore needs an efficient guidance and navigation system, based on sensor technology that allows it to navigate safely and accurately. A variable sensor system will provide an AGV with the ability to interchange guidance and navigation techniques as the parameters in the environment change in order to complete tasks. This will maintain material handling operations and guarantee production rates. The standardization of sensor based guidance and navigation systems for AGVs and other autonomous robots will optimize performance. The project involved research, design, construction, assembly and testing of a variable sensor system for guidance and navigation of AGVs. Sensor technology facilitated AGV operation in changing environments and avoidance of other robots and automated machines in an Agile manufacturing environment. Routing algorithms and procedures enabled safe movement and AGV task completion. The project required mechanical, electronic and software integration for AGV navigation and guidance.
Thesis (M.Sc.Eng.)-University of KwaZulu-Natal, Durban, 2008.

This thesis deals with development of guidance laws for advanced applications. Two class of guidance problems, namely, impact angle constrained guidance and pulsed guidance for exo-atmospheric engagements, are considered here. Three impact angle constrained guidance schemes are developed using (i) Proportional navigation guidance (PNG), (ii) State Dependent Riccati Equation (SDRE) technique and (iii) geometric concepts, respectively. A collision course based pulsed guidance law is presented for exo-atmospheric interceptors. Proportional Navigation Guidance (PNG) law is the most widely used guidance law because of its ease of implementation and efficiency. However, in its original form, it achieves only a limited set of impact angles. A two stage PNG law is presented for achieving all impact angles against a stationary target. In the first phase of guidance, an orientation PNG command is used. The orientation navigation constant (N ) is a function of the initial engagement geometry and has a lower value (N less than 2). It is proved that following the orientation trajectory, the interceptor can switch to N = 2 and achieve the desired impact angle. Simulations, with a constant speed and with a realistic interceptor model, show successful interception of the target with all desired impact angles. Feedback implementation of the guidance law results in negligible errors in impact angle with uncompensated autopilot delays. The idea of a two-stage PNG law with impact angle constraint is further used to develop a guidance law for intercepting moving targets. Following the orientation trajectory, the interceptor can switch to N = 3 and achieve the desired impact angle. It is proved that the guidance achieves all impact angles in a surface-to-surface engagement scenario with receding and approaching targets, respectively. In a air-to-surface engagement scenario, it is proved that the guidance law achieves all impact angles in a deterministic set. Constant speed and realistic interceptor models are used for simulations. Results show negligible error in impact angle and miss distance for moving targets. The guidance law, in its feedback implementation form, achieves the desired impact angle for interceptors with delay and with a maneuvering target. The impact angle errors are low with negligible errors in miss distance. Next, the impact angle constrained guidance problem against a stationary target is solved as a non-linear regulator problem using the SDRE technique. The interceptor guidance problems are of finite time nature. As the main contribution of this part of the work, we solve a finite time interceptor guidance problem with infinite horizon SDRE formulation by choosing the state weighting matrix as a function of time-to-go. Numerical simulations are carried out both for a constant speed interceptor model and a realistic interceptor model. Simulations for both the models are carried out for various impact angles and firing angles. Robustness of the proposed guidance law with respect to autopilot lag is also verified by simulations. Results obtained show the efficiency of the SDRE approach for impact angle constrained missile guidance. A geometric guidance scheme is proposed for lateral interception of targets in a planar engagement scenario in the absence of line-of-sight rate information. A kill-band is defined for target initial positions capturable by an arc maneuver, followed by a straight line path by the interceptor. Guidance law for capturing targets inside the kill-band is presented and is further modified for targets outside the kill-band. Based on analytical studies on the kill-band, a guidance law is proposed for lateral interception of maneuvering targets. Simulations are carried with for typical low speed engagements. The concept of kill-band provides an inherent robustness to the proposed guidance law with respect to uncompensated system delays and target maneuver. As the final part of the work, an interceptor endgame pulsed guidance law for exoatmospheric engagements is derived by using the notion of collision heading. The proposed guidance law is derived in steps by (i) Obtaining the collision heading based on the collision triangle engagement geometry and then (ii) Computing the width of the pulse fired by the divert thruster to attain the collision heading. It is shown that this strategy is more effective than the existing zero effort miss (ZEM) based guidance laws for intercepting targets with higher heading angles off the nominal head-on collision course. A result on pulse firing sequence is also presented showing that firing pulses in quick succession results in minimum pulse widths and hence minimum control effort for a desired miss distance. Simulations are carried out for various engagement scenarios. Results show better miss-distance and divert thrust performance as compared to the existing ZEM based law.

Dissertação de Mestrado Integrado em Engenharia Electrotécnica e de Computadores apresentada à Faculdade de Ciências e Tecnologia
Veículos guiados automaticamente são um factor chave na otimização do transporte de materiais no setor industrial. O interesse destes veículos está diretamente relacionado com a sua capacidade de transportar cargas pesadas entre as diferentes etapas de uma linha de montagem, de modo repetitivo, e exigindo uma reduzida intervenção humana. A navegação destes veículos é frequentemente baseada num guiamento óptico ou magnético. No entanto, a inflexibilidade das linhas de guiamento combinada com a sua deterioração ao longo do tempo, promove o desenvolvimento de métodos alternativos ou complementares para a navegação destes veículos. O objetivo principal deste trabalho é desenvolver e validar um sistema de guiamento inercial, composto por acelerómetros e giroscópios, para complementar um sistema de guiamento magnético existente. O sistema guiado inercialmente deve ser capaz de orientar o veículo, permitindo a sua navegação ao longo de caminhos em linha reta nos quais a banda magnética não está presente. Para concretizar esta tarefa, o projeto e a assemblagem de um sistema de aquisição de dados inerciais é primeiramente realizado, seguindo-se um estudo de algoritmos de fusão sensorial bem como da sua implementação em sistemas embebidos. Os algoritmos de fusão mais precisos e normalmente escolhidos para fundir as medições dos sensores e mitigar erros existentes são computacionalmente pesados, exigindo uma grande capacidade computacional o que é incompatível com dispositivos de baixo custo. Deste modo, este trabalho foca-se em abordagens de fusão sensorial de baixa complexidade, capazes de fundir a informação de sensores inerciais de baixo custo e que possam ser implementados em sistemas embebidos com baixa capacidade computacional. Os resultados demonstram que o sistema de guiamento inercial permite emular adequadamente o comportamento do sistema de guiamento magnético existente, para caminhos em linha reta. Este trabalho faz parte do projeto AGVPOSYS (em inglês, Automated Guided Vehicle with innovative indoor POsitioning SYStem for the factory of the future) liderado pela empresa Active Space Technologies e foi realizado nas instalações desta mesma empresa.
Automated guided vehicles are a key factor in the optimization of the transportation of materials in the industry sector. The interest of these vehicles is directly related with their capability of carrying heavy loads among the different steps of an assembly line, repetitively, while demanding a reduced human intervention. The navigation of these vehicles is often based on optical or magnetic guidance. However, the inflexibility of guidelines combined with its deterioration over time leads to the development of alternative or complementary methods of navigation for these vehicles. The main objective of this work is to develop and validate an inertial-based guidance system, composed of accelerometers and gyroscopes, in order to complement an existing magnetic-based guidance system. The inertial-based guided system should be capable of guiding the vehicle, allowing its navigation along straight paths in which the magnetic tape is not available. For this task, the design and assembly of an inertial data acquisition system is firstly performed, followed by a study of sensor fusion algorithms as well as their implementation in embedded systems. Standard accurate fusion algorithms implemented to merge sensor measurements and mitigate existing errors are computationally intensive, demanding an high computational capability which is incompatible with low-cost devices. Therefore, this work focus on low-complexity fusion filter approaches, which are capable of merging low-cost inertial sensors data and which can be implemented in embedded systems with low-computational capabilities. The results demonstrate that the inertial-based guidance system properly emulate the behaviour of the magnetic-guided system, for straight paths. This work is included in the AGVPOSYS (Automated Guided Vehicle with innovative indoor POsitioning SYStem for the factory of the future) project led by the company Active Space Technologies and was developed in the facilities of this company.

My research provides a historical and archaeological context for this thesis, in which I argue the JB-2 missile is historically significant as a unique example of the rapid duplication of enemy technology for both physical and psychological retaliation, as a crucial link in the chain of development for America’s cruise missile program, and for its role in early Cold War deterrence. Jet Bomb model number 2 (JB-2), America’s first operationally successful, mass produced cruise missile, developed as a direct copy of the German V-1, with slight variation in manufacture due to differences between German and American components, machinery and tooling. Continuing modifications of the JB-2 during its service life led to improvements in performance, control, and accuracy. From 1944 to 1953, the JB-2 transitioned from a weapon quickly prepared for wartime deployment to an essential test vehicle for the United States Army, Air Force and Navy while supporting the U.S. policy of containment during the early Cold War.