Thesis (MScEng)-- Stellenbosch University, 2013. / ENGLISH ABSTRACT: In this thesis the design, analysis, implementation, and verification of a fault-tolerant unmanned aerial vehicle (UAV) flight control system which is robust to structural damage causing the natural flight dynamics of the vehicle to become asymmetric, is presented. The main purpose of the robust control architecture is to maintain flight stability after damage has occurred. The control system must be able to handle an abrupt change from an undamaged to a damaged state, and must also not depend on explicit knowledge of the damage. A robust control approach is therefore preferred above an adaptive control approach. As a secondary objective, the system must provide robust flight performance to ensure adequate response times and acceptable transients’ behaviour, both in normal flight, and after damage has occurred. An asymmetric six degrees of freedom equations of motion model is derived. The model accounts for the changes in the aerodynamic model of the aircraft as well as changes in the centre of gravity location. Vortex lattice techniques are used to determine the aerodynamic coefficients of the aircraft for damage to the main wing resulting in 0% to 40% spanwise lifting surface loss. A sequential quadratic programming optimisation algorithm is applied to the force and moment equations to find the trim flight state and actuator deflections of the asymmetric aircraft for constant airspeed and altitude. The trim flight state can be further constrained to force zero bank angle, zero sideslip angle or a desired relative weighting of nonzero bank angle and nonzero sideslip angle. The calculated trim actuator deflections are compared to the physical deflection limits to determine the feasibility of maintaining trim flight for different percentages of wing loss. Assuming that a valid trim condition exists, the relative stability of the aircraft’s natural modes is analysed as a function of percentage wing loss by tracing the locus of the open-loop poles. An acceleration-based flight control architecture is designed and implemented, and the robustness of the flight control stability and performance is analysed as a function of percentage wing loss. The robustness and performance of the flight control system is verified with a nonlinear simulation for spanwise wing loss from 0 to 40%. Practical flight tests are performed to verify the robustness and performance of the flight control systems to in-flight damage. A detachable wing with release mechanism is designed and manufactured to simulate 20% wing loss. The flight control system is implemented on a practical UAV and a successful flight test shows that it performs fully autonomous flight control, and is able to accommodate an in-flight partial wing loss. / AFRIKAANSE OPSOMMING: In hierdie tesis word die ontwerp, analise, implementasie en verifikasie van ’n fout-verdraende onbemande vliegtuig beheerstelsel wat robuust is tot strukturele skade wat die natuurlike vlug dinamika van die voertuig asimmetries maak, voorgestel. Die hoofdoel van hierdie robuuste beheer argitektuur is om stabiliteit te verseker na die skade aangerig is. Die beheerstelsel moet die skielike verandering van normale na beskadigde vlug hanteer sonder enige eksplisiete kennis daarvan. Dus word ’n robuuste beheer aanslag verkies bo ’n aanpassende beheer struktuur. Tweedens moet die vlugbeheerstelsel robuust genoeg wees om steeds die gewenste reaksietyd en aanvaarbare oorgangsverskynsels te kan hanteer, tydens beide normale en beskadigde vlug. ’n Asimmetriese ses grade van vryheid beweginsvergelykings model word afgelei. Die model het die vermoë om veranderinge in die aerodinamiese model van die vliegtuig, sowel as massamiddelpunt verskuiwing, voor te stel. “Vortex Lattice” metodes is gebruik om die aerodinamiese koëffisiënte van die beskadigde vlerk voor te stel tussen 0% en 40% verlies. ’n Sekwensiële kwadratiese programmering optimiserings algorithme is aangewend op die krag en moment vergelykings om die ekwilibrium vlug toestand en aktueerder defleksies te vind vir ’n asimmetriese vliegtuig met konstante lugspoed en hoogte. Die ekwilibrium vlug toestand word verder beperk deur ’n nul rolhoek, ’n nul sygliphoek of ’n relatiewe weging van die twee. Die bepaalde ekwilibrium defleksies word dan vergelyk met die fisiese limiete om hulle geldigheid te bepaal vir ekwilibrium vlug. As ’n geldige ekwilibrium toestand bestaan, kan die relatiewe stabiliteit van die vliegtuig se natuurlike modusse ontleed word as ’n persentasie van vlerkverlies deur die wortellokusse van die ooplus pole na te gaan. ’n Versnellings-gebaseerde vlug beheerstelsel argitektuur is ontwerp en geïmplementeer. Daarna is die robuustheid ontleed as ’n funksie van die persentasie vlerkverlies. Die robuustheid en gedrag van hierdie vlugbeheerstelsel is geverifieer met ’n nie-linêre simulasie vir 0 tot 40% vlerkverlies. Praktiese vlugtoetse is onderneem om die robuustheid en gedrag tydens/na skade gedurende ’n vlug, te verifeer. ’n Vlerkverlies meganisme is ontwerp en vervaardig om 20% vlerkverlies te simuleer. Die vlugbeheerstelsel is geïmplementeer op ’n onbemande vliegtuig en die daaropvolgende suksesvolle vlug lewer bewys dat die vlugbeheerstelsel wel skade, in die vorm van gedeeltelike vlerkverlies, tydens vlug kan hanteer.

This thesis deals with the design and development of a 75 mm fixed-wing nano-air vehicle (NAV). The NAV is designed to fit inside a cube with each side measuring 75 mm. The range and endurance of the NAV are 300 m and 2-3 minutes, respectively. The high-wing horizontal tailless NAV has a take-off weight of 19.5 g. The battery-powered single propeller NAV has two control surfaces in the form of elevator and rudder. This thesis contains a detailed account of the airfoil selection, selection of the configuration of NAV and the longitudinal, lateral and directional aerodynamic characterization of the NAV. The development of one of the lightweight autopilot hardware which weighs 1.8 g is also given in detail. The development of non-linear equations of motion of NAV including thrust and coupling effects is also discussed. The effects of the gyroscopic coupling and counter torque on the linear dynamics of the NAV are analyzed by conducting a parametric study about the variation of the eigenstructure attributable to the varying degree of coupling in the system matrix of the linear coupled model. A robust simultaneously stabilizing output feedback controller is synthesized for stabilizing the plants of the NAV. The synthesizing of the robust simultaneously stabilizing output feedback controller is based on a frequency-shaped central plant. A new procedure is developed to determine the frequency-shaped central plant utilizing the v-gap metric between the plants, the frequency-shaping of the plants with the pre and post compensators and the robust stabilization theory. An optimization problem is formulated to obtain these compensators. A novel iterative algorithm is developed to acquire the compensators by solving the optimization problem. Thereafter, an iterative algorithm is developed to find an output feedback controller for robust simultaneous stabilization by blending the existing features of robust stability condition of right co-prime uncertainty model of the frequency-shaped central plant, the maximum v-gap metric of the frequency-shaped central plant, H∞ loop-shaping and eigenstructure assignment algorithm for output feedback using the genetic algorithm. The six-degree-of-freedom numerical and hardware-in-loop simulations (HILS) of closed-loop non-linear and linear plants of NAV are performed to assess the performance of the controller and to validate the control algorithm implemented in the autopilot. The airworthiness of the aircraft is tested by conducting flight trials in radio-controlled (RC) mode without including the autopilot. The successful RC flight trial of the NAV indicates airworthiness of the aircraft which aided in freezing the configuration. This is one of the smallest fixed wing aerial vehicle that was successfully flown till date.

Cette thèse propose des solutions aux problématiques inhérentes au contrôle de formations aériennes de type leader­-suiveur pour des flottes de quadrirotors. Au regard des travaux existants, les stratégies qui sont proposés dans notre travail, considère que le(s) leader{s) a une interaction avec les suiveurs. En outre, les rôles de leader et de suiveur sont interchangeables lors de la formation. Dans un premier temps, la modélisation mathématique d'un seul quadrirotor et celle de la formation de quadrirotors est développée. Ensuite, le problème de suivi de trajectoire pour un seul quadrirotor est étudié. Au travers de l'analyse de 1, dynamique du système pour la conception d'une commande par platitude, il apparait que le suivi de trajectoire pour chaque quadrirotor équivaut à déterminer les sorties plates désirées. Un contrôleur pour système plats permettant l'asservissement des drones pour le suivi de trajectoire est donc proposé. Étant donné la propriété de double-boucle de la dynamique du quadrirotor en boucle fermée, un contrôleur d'attitude avec des grands gains est conçu, selon la théorie « singular perturbation system ». Puisque la dynamique du quadrirotor en boucle fermée fonctionne sur deux échelles de temps, la dynamique de rotation (boundary-layer mode) est contrôlée sur l'échelle de temps la plus rapide. La conception du contrôleur de formation dépend seulement de la dynamique de translation (modèle réduit dans une échelle de temps lente). Ce résultat a simplifié la conception du contrôleur de formation, de telle sorte que le modèle réduit du quadrirotor est utilisé au lieu du modèle complet. Étant donné que le modèle réduit du quadrirotor a une caractéristique de double-intégrateur, un algorithme de consensus pour des systèmes caractérisés par de multiple double-intégrateurs est proposé. Pour traiter le problème de la formation leader-suiveur, une matrice d'interaction est initialement proposée basée sur la matrice de Laplacienne. Nous montrons que la condition de convergence et la vitesse de convergence de l'erreur de formation dépendent de la plus petite valeur propre de la matrice d'interaction. Trois stratégies de contrôle de la formation avec une topologie fixe sont ensuite proposées. Le contrôle de formation par platitude est proposé pour obtenir une formation agressive, tandis que les dérivées de grands ordres de la trajectoire désirée pour chaque UAV sont estimées en utilisant un observateur; la méthode Lyapunov redesign est implémentée pour traiter les non-linéarités de la dynamique de la translation des quadrotors; une loi de commande bornée par l'utilisation, entre autre, de la fonction tangente hyperbolique est développée avec un feedback composite non linéaire, afin d'améliorer les performances de la formation. De plus, une commande de commutation saturée de la formation est étudiée, car la topologie de la formation est variable. La stabilité du système est obtenue grâce aux théories “convex hull » et « common Lyapunov function ». Cette stratégie de commande de commutation permet le changement des leaders dans la formation. Inspirée par certains travaux existants, tels que le contrôle de la formation avec des voisins anonymes, nous proposons, finalement, une loi de commande avec des voisins pondérés, qui montre une meilleure robustesse que le contrôle avec des voisins anonymes. Les résultats de simulation obtenus avec Matlab illustrent premièrement nos stratégies de contrôle que nous proposons De plus, en utilisant le langage de programmation C ++, nos stratégies sont mises en œuvre dans un framework de simulation et d'expérimentation développé au laboratoire Heudiasyc. Grâce aux nombreux tests variés que nous avons réalisés en simulation et en temps-réel, l'efficacité et les avantages de nos stratégies de contrôle de la formation proposées sont présentés. / In this thesis, we address a leader-follower (L-F) formation control problem for multiple UAVs, especially quadrotors. Different from existing works, the strategies, which are proposed in our work, consider that the leader(s) have interaction with the followers. Additionally, the leader(s) are changeable during the formation. First, the mathematical modeling of a single quadrotor and of the formation of quadrotors is developed. The trajectory tracking problem for a single quadrotor is investigated. Through the analysis of the flatness of the quadrotor dynamical model, the desired trajectory for each quadrotor is transferred to the design of the desired at outputs. A flatness-based trajectory tracking controller is, then, proposed. Considering the double-loop property of the closed-loop quadrotor dynamics, a high-gain attitude controller is designed, according to the singular perturbation system theory. Since the closed-loop quadrotor dynamics performs in two time scales, the rotational dynamics (boundary-layer model) is controlled in a fast time scale. The formation controller design is then only considered for the translational dynamics: reduced model in a slow time scale. This result has simplified the formation controller design such that the reduced model of the quadrotor is considered instead of the complete model. Since the reduced model of the quadrotor has a double-integrator characteristic, consensus algorithm for multiple double-integrator systems is proposed. Dealing with the leader-follower formation problem, an interaction matrix is originally proposed based on the Laplacian matrix. We prove that the convergence condition and convergence speed of the formation error are in terms of the smallest eigenvalue of the interaction matrix. Three formation control strategies with fixed formation topology are then proposed. The flatness-based formation control is proposed to deal with the aggressive formation problem, while the high-order derivatives of the desired trajectory for each UAV are estimated by using an observer; the Lyapunov redesign is developed to deal with the nonlinearities of the translational dynamics of the quadrotors; the hyperbolic tangent-based bounded control with composite nonlinear feedback is developed in order to improve the performance of the formation. In an additional way, a saturated switching control of the formation is investigated, where the formation topology is switching. The stability of the system is obtained by introducing the convex hull theory and the common Lyapunov function. This switching control strategy permits the change of the leaders in the formation. Inspired by some existing works, such as the anonymous neighbor-based formation control, we finally propose a weighted neighbor-based control, which shows better robustness than the anonymous neighbor-based control. Simulation results using Matlab primarily illustrate our proposed formation control strategies. Furthermore, using C++ programming, our strategies are implemented on the simulator-experiment framework, developed at Heudiasyc laboratory. Through a variety of tests on the simulator and real-time experiments, the efficiency and the advantages of our proposed formation control strategies are shown. Finally, a vision-based inter-distance detection system is developed. This system is composed by an on-board camera, infrared LEDs and an infrared filter. The idea is to detect the UAVs and calculate the inter-distance by calculating the area of the special LEDs patterns. This algorithm is validated on a PC, with a webcam and primarily implemented on a real quadrotor.

Le travail de recherche effectué dans cette thèse a été réalisé dans le cadre d'une convention CIFRE entre le laboratoire IMS de l'université Bordeaux 1 et la société Airbus Operations S.A.S. Cette thèse traite de la détection robuste et précoce des pannes oscillatoires de faible amplitude dans les systèmes de commandes de vol électriques. Une panne oscillatoire est une oscillation anormale d'une surface de contrôle due à un dysfonctionnement dans la chaîne d'asservissement de la servocommande d'une gouverne. Les pannes oscillatoires ont une influence sur la structure, l'aéroélasticité et la pilotabilité de l'avion, lorsqu'ils sont situés dans la bande passante de l'actionneur. La capacité à détecter ces pannes est très importante car elles ont un impact sur la conception structurale de l'avion. Au plan méthodologique, nous nous sommes focalisés sur l'estimation adaptative des paramètres et de l'état à base d'une technique de filtrage non linéaire local. Le mécanisme de filtrage opère sur un modèle non linaire de la chaine de contrôle-commande de l’actionneur hydraulique en amont des surfaces de contrôle. L'algorithme d'estimation est basé sur une interpolation polynomiale d'opérateurs linéaire, et offre l'avantage d'une implémentation relativement aisée. Un problème crucial et sous-jacent est la détermination des hyper-paramètres de réglage de cet algorithme. Nous avons proposé une démarche hors-ligne dédiée, en intégrant un critère de sensibilité vis-à-vis des pannes que nous devons détecter. La technique proposée a été implémentée et testée: les résultats expérimentaux obtenus sur banc essai et sur un simulateur A380 ont clairement mis en évidence l'apport de la nouvelle approche en termes de performances, tout en gardant le même niveau de robustesse. / The research work done in this PhD has been caried out in the frame of an industrial convention (CIFRE) between the IMS laboratory and Airbus Operations S.A.S. The thesis deals with robust and early detection of oscillatory failures (OFC: Oscillatory Failure Case) in the Electrical Flight Control System. An oscillatory failure is an abnormal oscillation of a control surface due to component malfunction in control surface servoloops. OFCs have an influence on structural loads, aeroelasticity and controllability when located within the actuator bandwidth. The ability to detect these failures is very important because they have an impact on the structural design of the aircraft. Usual monitoring techniques cannot always guarantee to remain within an envelope with acceptable robustness. In this work, we develop a model based strategy to detect such failures with small amplitude at a very early stage. The monitoring strategy is based on dedicated non linear local filtering for on-line joint parameter/state estimation, allowing for model parameter variations during A/C flight. This strategy is associated with the same decision making rules as currently used for in-service Airbus A380. We propose a method for adjusting the tuning parameters so that various design goals and trades-off can be easily formulated and managed. The performance of the proposed fault detection scheme is measured by its detection delay, its propensity to issue false alarms and whether it permits a failure to go undetected. The proposed technique has been implemented and tested with success on Airbus test facilities including an A380 flight simulator.

Le développement des logiciels avioniques les plus critiques, comme les commandes de vol électriques, présentent plusieurs contraintes qui peuvent être quasiment contradictoires – par exemple, performance et sûreté – et toutes ces contraintes doivent être respectées simultanément. L'objective de cette thèse est d'étudier et de proposer des évolutions dans le cycle de développement des logiciels de commande de vol chez Airbus afin d'améliorer leur performance, tout en respectant les contraintes industrielles existantes et en conservant des processus de vérification au moins aussi sûrs que ceux utilisés actuellement. Le critère principal d'évaluation de performance est le temps d'exécution au pire cas (WCET), vu qu'il est utilisé lors des analyses temporelles des logiciels de vol réels. Dans un premier temps, le DO-178, qui contient des considérations pour l'approbation des logiciels avioniques, est présenté. Le DO-178B et le DO-178C sont étudiés. Le DO-178B est la référence pour plusieurs logiciels de commande de vol développés chez Airbus et le DO-178C est la référence pour le développement des nouveaux logiciels à partir de 2012. Ensuite, l'étude de cas est présentée. Afin d'améliorer sa compréhension, le contexte historique est fourni à travers l'étude des autres logiciels de commande de vol, car plusieurs activités de son cycle de vie réutilisent des techniques qui ont été utilisées avec succès dans des projets précédents. Quelques activités qui présentent des causes potentielles de pertes de performance logicielle sont exposées et l'axe principal d'étude choisi pour le reste de la thèse est la phase de compilation. Ce choix se justifie dans le contexte des logiciels de commande de vol car la compilation est réalisée avec peu ou pas d'optimisations, son impact sur la performance des logiciels est donc important et des travaux de recherche récents permettent d'envisager un changement dans les paradigmes actuels de compilation sûre. / The development of safety-critical avionics software, such as aircraft flight control programs, presents many different constraints that are nearly contradictory, such as performance and safety requirements, and all must be met simultaneously. The objective of this Thesis is to propose modifications in the development cycle of Airbus flight control programs in order to improve their performance without weakening their verification processes or violating other industrial constraints. The main criterion for performance evaluation is the Worst-Case Execution Time (WCET), as it is used in the timing analysis that is performed in actual avionics software verification processes. In a first moment, the DO-178, which contains guidance for avionics software development approval, is presented. Both the DO-178B and the DO-178C are discussed, since the former was the reference for the development of many Airbus flight control programs and the latter shall be the reference for the development of new programs, starting from 2012. Then, the case study is presented. In order to better understand it, some historical context is provided by the study of other flight control programs - many of its life cycle activities reuse techniques that were successful in previous software projects. Each activity is evaluated in order to underline what are the performance bottlenecks in the flight control software development. Some potential underperforming activities are depicted and the main axis of study developed subsequently is the compilation phase: not only it is a well-known unoptimized activity that has important impacts over software performance, but it is also an activity that might undergo a paradigm change due to innovating compilers that are being developed by researchers. The CompCert compiler is presented and its use in the scope of this Thesis is justified - at the time of this Thesis, it was the compiler that was best prepared to perform meaningful experiments, such as compiling a large subset of the chosen case study. Its architecture is studied, together with its semantic preservation theorem, which is the backbone of its formally-verified part. Additional features that were developed in CompCert during this Thesis in order to meet Airbus's requirements - such as its annotation mechanism and its reference interpreter - are discussed in order to underline their usefulness in the development of flight control software. The evaluation of CompCert consists in a performance comparison with the current compilation strategy and an assessment of the impacts that its utilization might have over the verification strategy commonly employed in flight control software. The results of the performance comparison are promising, since CompCert-generated code has a WCET more than 10% lower than if it were compiled with a good quality non-optimizing compiler. As expected, the use of CompCert has impacts over some important verification activities but its formal development and increased verifiability helps in the development of new compiler verification activities that can keep the whole development process at least as safe as the current one. Some development strategy propositions are then presented, according to the certification credit that might be required by using CompCert.

Cette thèse s'intéresse à l'adaptation des lois de pilotage d'un avion de transport civil aux différentes incertitudes qui peuvent affecter sa dynamique. Le procédé de pilotage adaptatif est censé fonctionner en temps réel à bord de l'avion afin d'optimiser la performance boucle fermée en fonction des conditions dans lesquelles il évolue. Les incertitudes peuvent être liées à la méconnaissance des conditions de vol (par exemple la vitesse et l'altitude), à des non-linéarités aérodynamiques inconnues ou encore à la méconnaissance du pilote aux commandes. Les procédés adaptatifs qui répondent à ces problèmes se doivent d'être performants sur l'ensemble du domaine opérationnel de l'avion en présence de perturbations réalistes. D'autres contraintes spécifiques peuvent être ajoutées en fonction du contexte (par exemple des charges limites, la stabilité aéroélastique, etc.). Plusieurs méthodes adaptatives sont testées afin d'adapter le système aux larges incertitudes qui le composent. Elles associent en général un estimateur en ligne (aussi appelé loi de mise-à-jour) à une loi de commande structurée. La synthèse de ces deux éléments peut être réalisée simultanément pour les méthodes adaptatives dites " directes ", comme par exemple le Model Reference Adaptative Control qui utilise la stabilité au sens de Lyapounov. Mais cette synthèse peut aussi être découplée pour les méthodes adaptatives dites "indirectes", ce qui offre un large choix de techniques pour chaque élément (comme les Moindres Carrés pour l'estimation de paramètres physiques incertains et la synthèse sous forme LFR pour le correcteur). Le choix de la méthode dépend fortement du contexte applicatif et des nombreuses contraintes associées. Trois applications sont au cœur de ce mémoire. Elles traitent de l'ajustement de lois de guidage à un modèle pilote inconnu, du contrôle longitudinal de non-linéarités de l'avion, et de la mise au point de lois longitudinale et latérale de pilotage manuel qui s'adaptent à des conditions de vol inconnues. Des méthodes avancées d'analyse linéaire et non-linéaire (dérivées de la µ-analyse et d'algorithmes d'optimisation) sont aussi mises en place pour valider ces systèmes sophistiqués adaptatifs en temps réel. D'une façon générale, les méthodes adaptatives indirectes ont donné le plus de satisfaction. Leur performance est aussi bonne que celle des méthodes directes, mais le fait qu'elles estiment en ligne des paramètres physiques facilite la surveillance temps réel du procédé adaptatif et sa validation. / This thesis deals with adapting flight control laws of a civil transport aircraft to various incertainties which can affect its behaviour. The adaptive flight control system is supposed to run in real time onboard the airplane so that its closed-loop performance is optimized with respect to the current conditions. These incertainties may be linked to unknown flight conditions (e.g. unknown airspeed and altitude), or unknown aerodynamics non-linearities or even unknown behaviour of the pilot in command. The adaptive schemes that are derived to answer these problems must be valid on the whole flight envelope with realistic disturbances but other additional contraints may exist depending on the context (e.g. loads limits, aeroelastic stability, etc.). To accommodate for large uncertainties on the system, adaptive methods are investigated. They usually combine an online estimator (also called an update law) with a structured flight control law. The synthesis of both elements may be simultaneous on 'direct' adaptive methods, e.g. on Model Reference Adaptive Control, using Lyapunov's stability theory. But it can also be decoupled on 'indirect' adaptive methods, giving a full spectrum of techniques for both elements (such as Least-Squares for estimating unknown physical parameters and the LFR framework for designing controllers). The choice of a specific method really depends on the application context and the related constraints.Three applications are the core of this report. They deal with adjusting guidance law to the pilot's unknown behaviour, controlling a longitudinal non-linearity, and providing manual longitudinal and lateral flight control laws which adapt to unknown flight conditions. Advanced linear and non-linear analysis techniques (based on µ-analysis or on optimization algorithms) are also applied to validated these sophisticated real-time adaptive systems. Results showed that indirect adaptive schemes were generally the most satisfactory. Their performance is similar to the one of direct schemes but as indirect methods provide physical parameter estimates, real-time monitoring and offline validation seem quite easier.

This master thesis is focused on design of control stick grip and rudder pedals extension. These components are equipped with active elements, which provide pilot with haptic feedback. The purpose of the introduced design is to allow prototype to be built into the aeroplane so that the proposed concept of haptic feedback can be tested onboard. It shall verify used technical solutions as well to allow for their application on following development stages that aim at certification of the proposed haptic feedback system to be used in general aviation aeroplanes. The designed components are the successors of prototypes used for experiments carried on flight simulator. The design process follows the requirements of legislation and outcomes of the previous experiments. This thesis follows the design process from setting of the design requirements to mechanical test of 3D printed prototypes.

<p> This work focuses on solving the general optimal control problems with smart-learning-enabled and theory-supported optimal control (SET-OC) approaches. The proposed SET-OC includes two main directions. Firstly, according to the basic idea of the direct method, the smart-learning-enabled iterative optimization algorithm (SEIOA) is proposed for solving discrete optimal control problems. Via discretization and reformulation, the optimal control problem is converted into a general quadratically constrained quadratic programming (QCQP) problem. Then, the SEIOA is applied to solving QCQPs. To be specific, first, a structure-exploiting decomposition scheme is introduced to reduce the complexity of the original problem. Next, an iterative search, combined with an intersection-cutting plane, is developed to achieve global convergence. Furthermore, considering the implicit relationship between the algorithmic parameters and the convergence rate of the iterative search, deep learning is applied to design the algorithmic parameters from an appropriate amount of training data to improve convergence property. To demonstrate the effectiveness and improved computational performance of the proposed SEIOA, the developed algorithms have been implemented in extensive real-world application problems, including unmanned aerial vehicle path planning problems and general QCQP problems. According to the theoretical analysis of global convergence and the simulation results, the efficiency, robustness, and improved convergence rate of the optimization framework compared to the state-of-the-art optimization methods for solving general QCQP problems are analyzed and verified. Secondly, the onboard learning-based optimal control method (L-OCM) is proposed to solve the optimal control problems. Supported by the optimal control theory, the necessary conditions of optimality for optimal control of the optimal control problem can be derived, which leads to two two-point-boundary-value-problems (TPBVPs). Then, critical parameters are identified to approximate the complete solutions of the TPBVPs. To find the implicit relationship between the initial states and these critical parameters, deep neural networks are constructed to learn the values of these critical parameters in real-time with training data obtained from the offline solutions.  To demonstrate the effectiveness and improved computational performance of the proposed L-OCM approaches, the developed algorithms have been implemented in extensive real-world application problems, including two-dimensional human-Mars entry, powered-descent, landing guidance problems, and fuel-optimal powered descent guidance (PDG) problems. In addition, considering there is no thorough analysis of the properties of the optimal control profile for PDG when considering the state constraints, a rigid theoretical analysis of the fuel-optimal PDG problem with state constraints is further provided. According to the theoretical analysis and simulation results, the optimality, robustness, and real-time performance of the proposed L-OCM are analyzed and verified, which indicates the potential for onboard implementation. </p>

<p>Cyber-physical systems (CPS) are a class of complicated systems integrating cyber components with physical components. Although such a cyber-physical interaction improves the system performance and intelligence, it increases the system complexity and makes the system vulnerable to various types of faults, failures, and cyber-attacks. To assure the security and improve the resiliency of CPS, it is found that the hybrid system model can be a powerful tool in the domain of fault detection and isolation, cyber-attack diagnosis and containment, as well as resilient control and reconfiguration. Several problems are concerned in this dissertation. For situational awareness, \textit{mode discernibility}, which stands for whether the discrete state of a hybrid system can be correctly identified, is characterized and discussed with potential applications to monitoring system design. For CPS vulnerability analysis, the problem of stealthy attack design for systems with switching structures is investigated, which is motivated by the recent literature. To further understand and remedy for the vulnerabilities, the detectability and identifiability for severe cyber-attacks are defined and characterized, which are followed by the discussions on the methodologies for cyber-attack detection and identification. Last but not least, based on the understanding of identifiability, a framework of resilient control design is proposed to mitigate the impact of cyber-attacks, which can be generalized in future to account for additional design criteria.</p>

<p>The use of lower airspace for air taxi and cargo applications opens up exciting prospects for futuristic Unmanned Aircraft Systems (UAS). However, ensuring the safety and security of these UAS within densely populated urban areas presents significant challenges. Most modern aircraft systems, whether unmanned or otherwise, rely on the Global Navigation Satellite System (GNSS) as a primary sensor for navigation. From satellite navigations point of view, the dense urban environment compromises positioning accuracy due to signal interference, multipath effects, etc. Furthermore, civilian GNSS receivers are susceptible to spoofing attacks since they lack encryption capabilities. Therefore, in this thesis, we focus on examining the safety and cybersecurity assurance of UAS in dense urban environments, from both theoretical and experimental perspectives. </p> <p>To facilitate the verification and validation of the UAS, the first part of the thesis focuses on the development of a realistic GNSS sensor emulation using a Gazebo plugin. This plugin is designed to replicate the complex behavior of the GNSS sensor in urban settings, such as multipath reflections, signal blockages, etc. By leveraging the 3D models of the urban environments and the ray-tracing algorithm, the plugin predicts the spatial and temporal patterns of GNSS signals in densely populated urban environments. The efficacy of the plugin is demonstrated for various scenarios including routing, path planning, and UAS cybersecurity. </p> <p>Subsequently, a robust state estimation algorithm for dynamical systems whose states can be represented by Lie Groups (e.g., rigid body motion) is presented. Lie groups provide powerful tools to analyze the complex behavior of non-linear dynamical systems by leveraging their geometrical properties. The algorithm is designed for time-varying uncertainties in both the state dynamics and the measurements using the log-linear property of the Lie groups. When unknown disturbances are present (such as GNSS spoofing, and multipath effects), the log-linearization of the non-linear estimation error dynamics results in a non-linear evolution of the linear error dynamics. The sufficient conditions under which this non-linear evolution of estimation error is bounded are derived, and Lyapunov stability theory is employed to design a robust filter in the presence of an unknown-but-bounded disturbance. </p>