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1
Ozkan, Bulent. "Dynamic Modeling, Guidance, And Control Of Homing Missiles." Phd thesis, METU, 2005. http://etd.lib.metu.edu.tr/upload/12606533/index.pdf.
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DYNAMIC MODELING, GUIDANCE, AND CONTROL OF HOMING MISSILES
Ö<br>ZKAN, Bü<br>lent
Ph. D., Department of Mechanical Engineering
Supervisor: Prof. Dr. M. Kemal Ö<br>ZGÖ<br>REN
Co-Supervisor: Dr. Gö<br>kmen MAHMUTYAZICIOgLU
September 2005, 236 pages
In this study, the dynamic modeling, guidance, and control of a missile with two relatively rotating parts are dealt with. The two parts of the missile are connected to each other by means of a roller bearing. In the first part of the study, the governing differential equations of motion of the mentioned missile are derived. Then, regarding the relative rotation between the bodies, the aerodynamic model of the missile is constructed by means of the Missile Datcom software available in TÜ<br>BiTAK-SAGE. After obtaining the required aerodynamic stability derivatives using the generated aerodynamic data, the necessary transfer functions are determined based on the equations of motion of the missile. Next, the guidance laws that are considered in this study are formulated. Here, the Linear Homing Guidance and the Parabolic Homing Guidance laws are introduced as alternatives to the Proportional Navigation Guidance law. On this occasion, the spatial derivation of the Proportional Navigation Guidance law is also done. Afterwards, the roll, pitch and yaw autopilots are designed using the determined transfer functions. As the roll autopilot is constructed to regulate the roll angle of the front body of the missile which is the controlled part, the pitch and yaw autopilots are designed to realize the command signals generated by the guidance laws. The guidance commands are in the form of either the lateral acceleration components or the flight path angles of the missile. Then, the target kinematics is modeled for a typical surface target. As a complementary part of the work, the design of a target state estimator is made as a first order fading memory filter. Finally, the entire guidance and control system is built by integrating all the models mentioned above. Using the entire system model, the computer simulations are carried out using the Matlab-Simulink software and the proposed guidance laws are compared with the Proportional Navigation Guidance law. The comparison is repeated for a selected single-body missile as well. Consequently, the simulation results are discussed and the study is evaluated.
2
Morgan, Robert W. "A New Paradigm in Optimal Missile Guidance." Diss., The University of Arizona, 2007. http://hdl.handle.net/10150/194121.
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This dissertation investigates advanced concepts in terminal missile guidance. The terminal phase of missile guidance usually lasts less than ten seconds and calls for very accurate maneuvering to ensure intercept. Technological advancements have produced increasingly sophisticated threats that greatly reduce the effectiveness of traditional approaches to missile guidance. Because of this, terminal missile guidance is, and will remain, an important and active area of research. The complexity of the problem and the desire for an optimal solution has resulted in researchers focusing on simplistic, usually linear, models. The fruit of these endeavors has resulted in some of the world's most advanced weapons systems. Even so, the resulting guidance schemes cannot possibly counter the evolving threats that will push the system outside the linear envelope for which they were designed. The research done in this dissertation greatly extends previous research in the area of optimal missile guidance. Herein it is shown that optimal missile guidance is fundamentally a pairing of an optimal guidance strategy and an optimal control strategy. The optimal guidance strategy is determined from a missile's information constraints, which are themselves largely determined from the missile's sensors. The optimal control strategy is determined by the missile's control constraints, and works to achieve a specified guidance strategy. This dichotomy of missile guidance is demonstrated by showing that missiles having different control constraints utilize the same guidance strategy so long as the information constraints are the same. This concept has hitherto been unrecognized because of the difficulty in developing an optimal control for the nonlinear set of equations that result from control constraints. Having overcome this difficulty by indirect means, evidence of the guidance strategy paradigm emerged. The guidance strategy paradigm is used to develop two advanced guidance laws. The new guidance laws are compared qualitatively and quantitatively with existing guidance laws.
3
Le, Voyer Damien. "Guidance and Control of a Naval Cruise Missile." Thesis, KTH, Reglerteknik, 2008. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-105890.
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Today the armed forces of many countries need to strike accurately potential enemies, wherever they might be, from a safe place. Since naval units can be deployed almost everywhere in the open sea, the idea of a naval cruise missile emerged in the 70’s. These missiles are designed to be launched from various naval vehicles such as frigates or submarines and strike deeply in the enemy territory. A program called Missile de Croisière Naval (MdCN Naval Cruise Missile) was therefore launched in 2006 by the DGA, the French procurement agency. MBDA is the industrial company appointed by the DGA to design and build the missile. Control aspects on a cruise missile are of primary interest since they impact the reliability, performance and availability of the weapon. In the aeronautics and weapon industry, gain scheduled controllers are used in most cases. However, many non-linear techniques have been developed in the literature and might improve the behaviour of the missile. The main objective of the present thesis is to apply non-linear techniques on the control and guidance loops of the MdCN too see whether of not they can improve such a system. Based on this report it should be easy for the engineers of the DGA to compare the controllers of the thesis and the classical gain scheduled controllers used in the industry. To achieve this task some basic knowledge of flight dynamics are recalled and a model of the MdCN is computed and divided into the control loop and the guidance loop. Then a non-linear controller for the launch phase using a Lyapunov based technique called back-stepping is designed and tested through a statistic analysis. During the cruise phase different anti-windup strategies are applied on the propulsion control loop of the missile and compared. Finally a software interface with a navigation-dedicated tool is coded and implemented in Simulink to analyse the complete Guidance-Navigation-Control loop and to see how navigation errors impact the control algorithms. The main contributions of this thesis are the controllers designed for the launch phase and the propulsion loop that will be compared with the controller that MBDA is going to deliver next year to see whether or not the non-linear techniques used in the thesis should be used on the missile. Furthermore, all the tools and procedure set up to interface the control and guidance laws with the navigation models and filters will give the possibility to the DGA to have a deeper understanding of the algorithms used by MBDA and to make sure that navigation and estimation issues are properly taken in account when designing the control and guidance laws.
4
Swee, John C. S. "Missile terminal guidance and control against evasive targets." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2000. http://handle.dtic.mil/100.2/ADA378653.
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Thesis (M.S. in Electrical Engineering) Naval Postgraduate School, March 2000.<br>Thesis advisor(s): Hutchins, Robert G. "March 2000." Includes bibliographical references (p. 83). Also available online.
5
McConnell, George. "Digital bank-to-turn control and guidance." Thesis, Queen's University Belfast, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.303013.
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Gunbatar, Yakup. "Varying Mass Missile Dynamics, Guidance &." Master's thesis, METU, 2007. http://etd.lib.metu.edu.tr/upload/2/12608977/index.pdf.
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The focus of this study is to be able to control the air-to-surface missile
throughout the entire flight, with emphasis on the propulsion phase to increase the
impact range of the missile.
A major difficulty in controlling the missile during the propulsion phase is the
important change in mass of the missile. This results in sliding the center of gravity
(cg) point and changing inertias. Moreover, aerodynamic coefficients and stability
derivatives are not assumed to be constant at predetermined ranges<br>conversely, they
depend on Mach number, angle of attack, and side slip angle. Consequently, as the
change of missile mass, cg point, inertia terms, and stability and aerodynamic
coefficients come together apart from flight operation stages, a great number of
points need to be taken into account when designing the controller. This makes
controlling the missile all the more complicated.
In this thesis, first the equations of motion are derived, in which, mass of the
missile is not assumed constant. Thus, not only the variation of mass but also the
variation of inertias is incorporated in the equations of motion. From the derived
v
equations of motion, a nonlinear inverse dynamics controller that can achieve desired
guidance for a conceptually developed air-to-surface missile has been designed,
tested and verified for a modeled missile with six degrees of freedom. For brevity of
the study, conceptual design and aerodynamic calculations are not given in detail.
Nevertheless, improvements for conceptual design are suggested.
As a result, it is shown that the controller works efficiently: the missile is able
to hit the target with less than 12 m circular error of probability (CEP). Finally,
studies and improvements are proposed.
7
Best, Robert Andrew. "Integrated tracking and guidance." Thesis, University of Birmingham, 1996. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.322491.
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Roddy, D. J. "Application of optimal control to bank-to-turn CLOS guidance." Thesis, Queen's University Belfast, 1985. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.373543.
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Vural, Ozgur Ahmet. "Fuzzy Logic Guidance System Design For Guided Missiles." Master's thesis, METU, 2003. http://etd.lib.metu.edu.tr/upload/1026715/index.pdf.
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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.
10
Tanner, Gwen Lettice. "Some topics in multiple hypothesis estimation and control using non-quadratic cost functions." Thesis, University of Warwick, 1997. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.362497.
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Tekin, Raziye. "Design, Modeling, Guidance And Control Of A Vertical Launch Surface To Air Missile." Master's thesis, METU, 2010. http://etd.lib.metu.edu.tr/upload/12612408/index.pdf.
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The recent interests in the necessity of high maneuverability and vertical launching triggered namely the unconventional control design techniques that are effective at high angle of attack flight regimes. For most of missile configurations, this interest required thrust vector control together with conventional aerodynamic control.
In this study, nonlinear modeling and dynamical analysis of a surface to air missile with both aerodynamic and thrust vector control is investigated. Aerodynamic force and moment modeling of the presented missile includes the challenging high angle of attack aerodynamics behavior and the so called hybrid control, which utilizes both tail fins and jet vanes as control surfaces. Thrust vector and aerodynamic control effectiveness is examined during flight envelope. Different autopilot designs are accomplished with hybrid control. Midcourse and terminal guidance algorithms are implemented and performed on target sets including maneuverable targets. A different initial turnover strategy is suggested and compared with standard skid-to-turn maneuver. Comparisons of initial roll with aerodynamic and thrust vector control are examined. Afterwards, some critical maneuvers and hybrid control ratio is studied with a real coded genetic algorithm. Rapid turnover for low altitude targets, intercept maneuver analysis with hybrid control ratio and lastly, engagement initiation maneuver optimization is fulfilled.
12
Sefastsson, Ulf. "Evaluation of Missile Guidance and Autopilot through a 6 DOF Simulation Model." Thesis, KTH, Optimeringslära och systemteori, 2016. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-188897.
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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.<br>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.
13
Evcimen, Cagdas. "Development And Comparison Of Autopilot And Guidance Algorithms For Missiles." Master's thesis, METU, 2007. http://etd.lib.metu.edu.tr/upload/2/12608677/index.pdf.
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In order to have an interception with a target, a missile should be guided with a successful guidance algorithm accompanied with a suitable autopilot structure. In this study, different autopilot and guidance designs for a canard-controlled missile
are developed. As a first step, nonlinear missile mathematical model is derived by using the equations of motion with aerodynamic coefficients found by Missile DATCOM program.
Autopilot design starts by the linearization of the nonlinear missile model around equilibrium flight conditions. Controllers based on the concepts of optimal control theory results and sliding mode control are designed. In all of the designs, angle
of attack command and roll angle command type autopilot structures are used. During the design process, variations in angle of attack, Mach number and altitude can lead to significant performance degradation. This problem is typically solved by applying gain-scheduling methodology according to these parameters.
There are different types of guidance methods in the literature. Throughout this study, proportional navigation guidance and its modified forms are selected as a base algorithm in the guidance system design. Other robust forms of guidance methods, such as an optimal guidance approach and sliding mode guidance, are also formed for performance comparison with traditional proportional navigation guidance approach. Finally, a new guidance method, optimal proportional-integral guidance, whose performance is the best among all of the methods included in the thesis against highly maneuvering targets, is introduced.
14
Sert, Muhammet. "A Rule Based Missile Evasion Method For Fighter Aircrafts." Master's thesis, METU, 2008. http://etd.lib.metu.edu.tr/upload/12609490/index.pdf.
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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.
15
Anisi, David A. "Online trajectory planning and observer based control." Licentiate thesis, Stockholm : Optimization and systems theory, Royal Institute of Technology, 2006. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-4153.
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Anisi, David A. "On Cooperative Surveillance, Online Trajectory Planning and Observer Based Control." Doctoral thesis, KTH, Optimeringslära och systemteori, 2009. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-9990.
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The main body of this thesis consists of six appended papers. In the first two, different cooperative surveillance problems are considered. The second two consider different aspects of the trajectory planning problem, while the last two deal with observer design for mobile robotic and Euler-Lagrange systems respectively.In Papers A and B, a combinatorial optimization based framework to cooperative surveillance missions using multiple Unmanned Ground Vehicles (UGVs) is proposed. In particular, Paper A considers the the Minimum Time UGV Surveillance Problem (MTUSP) while Paper B treats the Connectivity Constrained UGV Surveillance Problem (CUSP). The minimum time formulation is the following. Given a set of surveillance UGVs and a polyhedral area, find waypoint-paths for all UGVs such that every point of the area is visible from a point on a waypoint-path and such that the time for executing the search in parallel is minimized. The connectivity constrained formulation extends the MTUSP by additionally requiring the induced information graph to be kept recurrently connected at the time instants when the UGVs perform the surveillance mission. In these two papers, the NP-hardness of both these problems are shown and decomposition techniques are proposed that allow us to find an approximative solution efficiently in an algorithmic manner.Paper C addresses the problem of designing a real time, high performance trajectory planner for an aerial vehicle that uses information about terrain and enemy threats, to fly low and avoid radar exposure on the way to a given target. The high-level framework augments Receding Horizon Control (RHC) with a graph based terminal cost that captures the global characteristics of the environment. An important issue with RHC is to make sure that the greedy, short term optimization does not lead to long term problems, which in our case boils down to two things: not getting into situations where a collision is unavoidable, and making sure that the destination is actually reached. Hence, the main contribution of this paper is to present a trajectory planner with provable safety and task completion properties. Direct methods for trajectory optimization are traditionally based on a priori temporal discretization and collocation methods. In Paper D, the problem of adaptive node distribution is formulated as a constrained optimization problem, which is to be included in the underlying nonlinear mathematical programming problem. The benefits of utilizing the suggested method for online trajectory optimization are illustrated by a missile guidance example.In Paper E, the problem of active observer design for an important class of non-uniformly observable systems, namely mobile robotic systems, is considered. The set of feasible configurations and the set of output flow equivalent states are defined. It is shown that the inter-relation between these two sets may serve as the basis for design of active observers. The proposed observer design methodology is illustrated by considering a unicycle robot model, equipped with a set of range-measuring sensors. Finally, in Paper F, a geometrically intrinsic observer for Euler-Lagrange systems is defined and analyzed. This observer is a generalization of the observer proposed by Aghannan and Rouchon. Their contractivity result is reproduced and complemented by a proof that the region of contraction is infinitely thin. Moreover, assuming a priori bounds on the velocities, convergence of the observer is shown by means of Lyapunov's direct method in the case of configuration manifolds with constant curvature.<br>QC 20100622<br>TAIS, AURES
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Moon, Jongki. "Mission-based guidance system design for autonomous UAVs." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/31797.
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Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2010.<br>Committee Chair: Prasad, JVR; Committee Member: Costello, Mark; Committee Member: Johnson, Eric; Committee Member: Schrage, Daniel; Committee Member: Vela, Patricio. Part of the SMARTech Electronic Thesis and Dissertation Collection.
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Corban, J. Eric. "Real-time guidance and propulsion control for single-stage-to-orbit airbreathing vehicles." Diss., Georgia Institute of Technology, 1989. http://hdl.handle.net/1853/12889.
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Gramajo, German G. "Guidance control of small UAV with energy and maneuverability limitations for a search and coverage mission." Thesis, California State University, Long Beach, 2014. http://pqdtopen.proquest.com/#viewpdf?dispub=1526913.
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<p> This thesis presents an algorithm for a search and coverage mission that has increased autonomy in generating an ideal trajectory while explicitly considering the available energy in the optimization. Further, current algorithms used to generate trajectories depend on the operator providing a discrete set of turning rate requirements to obtain an optimal solution. This work proposes an additional modification to the algorithm so that it optimizes the trajectory for a range of turning rates instead of a discrete set of turning rates. This thesis conducts an evaluation of the algorithm with variation in turn duration, entry-heading angle, and entry point. Comparative studies of the algorithm with existing method indicates improved autonomy in choosing the optimization parameters while producing trajectories with better coverage area and closer final distance to the desired terminal point. </p>
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Steindorf, Lukas. "Constrained Low-Thrust Satellite Formation-Flying Using Relative Orbit Elements : Autonomous Guidance and Control for the NetSat Satellite Formation-Flying Mission." Thesis, Luleå tekniska universitet, Institutionen för system- och rymdteknik, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:ltu:diva-61599.
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This thesis proposes a continuous low-thrust guidance and control strategy for satellite formation-flying. Stabilizing feedback based on mean relative orbit elements and Lyapunov theory is used. A novel feedback gain matrix inspired by the fuel-optimal impulsive solution is designed to achieve near-optimal fuel consumption. A reference governor is developed to autonomously guide the spacecraft through the relative state-space in order to allow for arbitrarily constrained satellite formations. Constraints include desired thrust levels, time constraints, passive collision avoidance and locally constrained state-space areas. Keplerian dynamics are leveraged to further decrease fuel consumption. Simulations show fuel consumptions of only 4% higher delta-v than the fuel-optimal impulsive solution. The proposed control and guidance strategy is tested in a high-fidelity orbit propagation simulation using MATLAB/Simulink. Numerical simulations include orbit perturbations such as atmospheric drag, high-order geopotential, solar radiation pressure and third-body (Moon and Sun) effects. Test cases include reconfiguration scenarios with imposed wall, thrust and time constraints and a formation maintenance experiment as flown by TanDEM-X, the TanDEM-X Autonomous Formation-Flying (TAFF) experiment.
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Basso, Maik. "A framework for autonomous mission and guidance control of unmanned aerial vehicles based on computer vision techniques." reponame:Biblioteca Digital de Teses e Dissertações da UFRGS, 2018. http://hdl.handle.net/10183/179536.
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A computação visual é uma área do conhecimento que estuda o desenvolvimento de sistemas artificiais capazes de detectar e desenvolver a percepção do meio ambiente através de informações de imagem ou dados multidimensionais. A percepção visual e a manipulação são combinadas em sistemas robóticos através de duas etapas "olhar"e depois "movimentar-se", gerando um laço de controle de feedback visual. Neste contexto, existe um interesse crescimente no uso dessas técnicas em veículos aéreos não tripulados (VANTs), também conhecidos como drones. Essas técnicas são aplicadas para posicionar o drone em modo de vôo autônomo, ou para realizar a detecção de regiões para vigilância aérea ou pontos de interesse. Os sistemas de computação visual geralmente tomam três passos em sua operação, que são: aquisição de dados em forma numérica, processamento de dados e análise de dados. A etapa de aquisição de dados é geralmente realizada por câmeras e sensores de proximidade. Após a aquisição de dados, o computador embarcado realiza o processamento de dados executando algoritmos com técnicas de medição (variáveis, índice e coeficientes), detecção (padrões, objetos ou áreas) ou monitoramento (pessoas, veículos ou animais). Os dados processados são analisados e convertidos em comandos de decisão para o controle para o sistema robótico autônomo Visando realizar a integração dos sistemas de computação visual com as diferentes plataformas de VANTs, este trabalho propõe o desenvolvimento de um framework para controle de missão e guiamento de VANTs baseado em visão computacional. O framework é responsável por gerenciar, codificar, decodificar e interpretar comandos trocados entre as controladoras de voo e os algoritmos de computação visual. Como estudo de caso, foram desenvolvidos dois algoritmos destinados à aplicação em agricultura de precisão. O primeiro algoritmo realiza o cálculo de um coeficiente de reflectância visando a aplicação auto-regulada e eficiente de agroquímicos, e o segundo realiza a identificação das linhas de plantas para realizar o guiamento dos VANTs sobre a plantação. O desempenho do framework e dos algoritmos propostos foi avaliado e comparado com o estado da arte, obtendo resultados satisfatórios na implementação no hardware embarcado.<br>Cumputer Vision is an area of knowledge that studies the development of artificial systems capable of detecting and developing the perception of the environment through image information or multidimensional data. Nowadays, vision systems are widely integrated into robotic systems. Visual perception and manipulation are combined in two steps "look" and then "move", generating a visual feedback control loop. In this context, there is a growing interest in using computer vision techniques in unmanned aerial vehicles (UAVs), also known as drones. These techniques are applied to position the drone in autonomous flight mode, or to perform the detection of regions for aerial surveillance or points of interest. Computer vision systems generally take three steps to the operation, which are: data acquisition in numerical form, data processing and data analysis. The data acquisition step is usually performed by cameras or proximity sensors. After data acquisition, the embedded computer performs data processing by performing algorithms with measurement techniques (variables, index and coefficients), detection (patterns, objects or area) or monitoring (people, vehicles or animals). The resulting processed data is analyzed and then converted into decision commands that serve as control inputs for the autonomous robotic system In order to integrate the visual computing systems with the different UAVs platforms, this work proposes the development of a framework for mission control and guidance of UAVs based on computer vision. The framework is responsible for managing, encoding, decoding, and interpreting commands exchanged between flight controllers and visual computing algorithms. As a case study, two algorithms were developed to provide autonomy to UAVs intended for application in precision agriculture. The first algorithm performs the calculation of a reflectance coefficient used to perform the punctual, self-regulated and efficient application of agrochemicals. The second algorithm performs the identification of crop lines to perform the guidance of the UAVs on the plantation. The performance of the proposed framework and proposed algorithms was evaluated and compared with the state of the art, obtaining satisfactory results in the implementation of embedded hardware.
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Levis, Aaron. "Determining Feasibility of a Propulsionless Microsatellite Formation Flight Mission." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1882.
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Benefits of developing missions with multiple formation flying spacecraft as an alternative to a traditional monolithic vehicle are becoming apparent. In some cases, these missions can lower cost and increase flexibility among other situational advantages. However, there are various limitations that are imposed by these missions that are centered on the concept of maintaining the necessary formation. One such limitation is that of the propulsion system required for each spacecraft. To mitigate the complexity and mass of the onboard propulsion, the pairing of electromagnetic actuators and differential drag to replace the functionality of a propulsive system is investigated. By using COTS magnetorquer boards to command satellite orientation, a scenario in which two 3U CubeSats are initially deployed from the ISS NanoRacks at an altitude of 400 km. They are then commanded to achieve a relative separation of 1 km and hold the spacing to demonstrate the capability of formation flight. The scenario was simulated through the MATLAB/Simulink platform and the magnitude of the necessary command torques were determined. By comparison to the ISIS magnetorquer board, the necessary command torques seem relatively high than compared to what the actuator is capable of. The ISIS board may supply ~5e-6 Nm of torque while the mission requires as much as 3e-3 Nm at times. However, by extending the settling time of the control law at the expense of absolute orientation control, the control torques necessary to carry out the simulated mission are well within the bounds of the ISIS magnetorquer boards as well as other COTS boards. With this alteration, mission feasibility is determined. It should be noted that further analysis should be conducted regarding concerns with CubeSat detumble to further confirm feasibility.
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Rogers, Jonathan. "Applications of internal translating mass technologies to smart weapons systems." Diss., Atlanta, Ga. : Georgia Institute of Technology, 2009. http://hdl.handle.net/1853/31813.
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Thesis (Ph.D)--Aerospace Engineering, Georgia Institute of Technology, 2010.<br>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.
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Guzman, Esteban. "Generating Exploration Mission-3 Trajectories to a 9:2 NRHO using Machine Learning." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1953.
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The purpose of this thesis is to design a machine learning algorithm platform that provides expanded knowledge of mission availability through a launch season by improving trajectory resolution and introducing launch mission forecasting. The specific scenario addressed in this paper is one in which data is provided for four deterministic translational maneuvers through a mission to a Near Rectilinear Halo Orbit (NRHO) with a 9:2 synodic frequency. Current launch availability knowledge under NASA’s Orion Orbit Performance Team is established by altering optimization variables associated to given reference launch epochs. This current method can be an abstract task and relies on an orbit analyst to structure a mission based off an established mission design methodology associated to the performance of Orion and NASA's Space Launch System. Introducing a machine learning algorithm trained to construct mission scenarios within the feasible range of known trajectories reduces the required interaction of the orbit analyst by removing the needed step of optimizing the orbit to fit an expected translational response required of the spacecraft. In this study, k-Nearest Neighbor and Bayesian Linear Regression successfully predicted classical orbital elements for the launch windows observed. However both algorithms had limitations due to their approaches to model fitting. Training machine learning algorithms off of classical orbital elements introduced a repetitive approach to reconstructing mission segments for different arrival opportunities through the launch window and can prove to be a viable method of launch window scan generation for future missions.
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Gagliano, Joseph R. "Orbital Constellation Design and Analysis Using Spherical Trigonometry and Genetic Algorithms: A Mission Level Design Tool for Single Point Coverage on Any Planet." DigitalCommons@CalPoly, 2018. https://digitalcommons.calpoly.edu/theses/1877.
Full textAbstract:
Recent interest surrounding large scale satellite constellations has increased analysis efforts to create the most efficient designs. Multiple studies have successfully optimized constellation patterns using equations of motion propagation methods and genetic algorithms to arrive at optimal solutions. However, these approaches are computationally expensive for large scale constellations, making them impractical for quick iterative design analysis. Therefore, a minimalist algorithm and efficient computational method could be used to improve solution times. This thesis will provide a tool for single target constellation optimization using spherical trigonometry propagation, and an evolutionary genetic algorithm based on a multi-objective optimization function. Each constellation will be evaluated on a normalized fitness scale to determine optimization. The performance objective functions are based on average coverage time, average revisits, and a minimized number of satellites. To adhere to a wider audience, this design tool was written using traditional Matlab, and does not require any additional toolboxes.
To create an efficient design tool, spherical trigonometry propagation will be utilized to evaluate constellations for both coverage time and revisits over a single target. This approach was chosen to avoid solving complex ordinary differential equations for each satellite over a long period of time. By converting the satellite and planetary target into vectors of latitude and longitude in a common celestial sphere (i.e. ECI), the angle can be calculated between each set of vectors in three-dimensional space. A comparison of angle against a maximum view angle, , controlled by the elevation angle of the target and the satellite’s altitude, will determine coverage time and number of revisits during a single orbital period.
Traditional constellations are defined by an altitude (a), inclination (I), and Walker Delta Pattern notation: T/P/F. Where T represents the number of satellites, P is the number of orbital planes, and F indirectly defines the number of adjacent planes with satellite offsets. Assuming circular orbits, these five parameters outline any possible constellation design. The optimization algorithm will use these parameters as evolutionary traits to iterate through the solutions space. This process will pass down the best traits from one generation to the next, slowly evolving and converging the population towards an optimal solution. Utilizing tournament style selection, multi-parent recombination, and mutation techniques, each generation of children will improve on the last by evaluating the three performance objectives listed. The evolutionary algorithm will iterate through 100 generations (G) with a population (n) of 100.
The results of this study explore optimal constellation designs for seven targets evenly spaced from 0° to 90° latitude on Earth, Mars and Jupiter. Each test case reports the top ten constellations found based on optimal fitness. Scatterplots of the constellation design solution space and the multi-objective fitness function breakdown are provided to showcase convergence of the evolutionary genetic algorithm. The results highlight the ratio between constellation altitude and planetary radius as the most influential aspects for achieving optimal constellations due to the increased field of view ratio achievable on smaller planetary bodies. The multi-objective fitness function however, influences constellation design the most because it is the main optimization driver. All future constellation optimization problems should critically determine the best multi-objective fitness function needed for a specific study or mission.
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Kubik, Stephen T. "Application of Parent-Child UAV Tasking For Wildfire Detection and Response." DigitalCommons@CalPoly, 2008. https://digitalcommons.calpoly.edu/theses/28.
Full textAbstract:
In recent years, unmanned aerial vehicles (UAVs) have become a dominant force in the aerospace industry. Recent technological developments have moved these aircraft from remote operation roles to more active response missions. Of particular interest is the possibility of applying UAVs toward solving complex problems in long-endurance missions. Under that belief, the feasibility of utilizing UAVs for wildfire detection and response was investigated in a partnership that included NASA’s Aeronautics Research Mission Directorate and Science Mission Directorate, and the United States Forest Service. Under NASA’s Intelligent Mission Management (IMM) project, research was conducted to develop a mission architecture that would enable use of a high altitude UAV to search for reported wildfires with a separate low altitude UAV supporting ground assets.
This research proposes a “straw man” concept incorporating both a High Altitude Long Endurance (HALE) UAV and a Low Altitude Short Endurance (LASE) UAV in a loosely coupled, low cost solution tailored towards wildfire response. This report identifies the communications architecture, algorithms, and required system configuration that meets the outlined goals of the IMM project by mitigating wildfires and addressing the United States Forest Service immediate needs. The end product is a defined parent-child framework capable of meeting all wildfire mission goals. The concept has been implemented in simulation, the results of which are presented in this report.
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Mattei, Giovanni. "Robust nonlinear control : from continuous time to sampled-data with aerospace applications." Thesis, Paris 11, 2015. http://www.theses.fr/2015PA112025/document.
Full textAbstract:
La thèse porte sur le développement des techniques non linéaires robustes de stabilisation et commande des systèmes avec perturbations de model. D’abord, on introduit les concepts de base de stabilité et stabilisabilité robuste dans le contexte des systèmes non linéaires. Ensuite, on présente une méthodologie de stabilisation par retour d’état en présence d’incertitudes qui ne sont pas dans l’image de la commande («unmatched»). L’approche récursive du «backstepping» permet de compenser les perturbations «unmatched» et de construire une fonction de Lyapunov contrôlée robuste, utilisable pour le calcul ultérieur d’un compensateur des incertitudes dans l’image de la commande («matched»). Le contrôleur obtenu est appelé «recursive Lyapunov redesign». Ensuite, on introduit la technique de stabilisation par «Immersion & Invariance» comme outil pour rendre un donné contrôleur non linéaire, robuste par rapport à dynamiques non modelées. La première technique de contrôle non linéaire robuste proposée est appliquée au projet d’un autopilote pour un missile air-air et au développement d’une loi de commande d’attitude pour un satellite avec appendices flexibles. L’efficacité du «recursive Lyapunov redesign» est mis en évidence dans le deux cas d’étude considérés. En parallèle, on propose une méthode systématique de calcul des termes incertains basée sur un modèle déterministe d’incertitude. La partie finale du travail de thèse est relative à la stabilisation des systèmes sous échantillonnage. En particulier, on reformule, dans le contexte digital, la technique d’Immersion et Invariance. En premier lieu, on propose des solutions constructives en temps continu dans le cas d’une classe spéciale des systèmes en forme triangulaire «feedback form», au moyen de «backstepping» et d’arguments de domination non linéaire. L’implantation numérique est basée sur une loi multi-échelles, dont l’existence est garantie pour la classe des systèmes considérée. Le contrôleur digital assure la propriété d’attractivité et des trajectoires bornées. La loi de commande, calculée par approximation finie d’un développement asymptotique, est validée en simulation de deux exemples académiques et deux systèmes physiques, le pendule inversé sur un chariot et le satellite rigide<br>The dissertation deals with the problems of stabilization and control of nonlinear systems with deterministic model uncertainties. First, in the context of uncertain systems analysis, we introduce and explain the basic concepts of robust stability and stabilizability. Then, we propose a method of stabilization via state-feedback in presence of unmatched uncertainties in the dynamics. The recursive backstepping approach allows to compensate the uncertain terms acting outside the control span and to construct a robust control Lyapunov function, which is exploited in the subsequent design of a compensator for the matched uncertainties. The obtained controller is called recursive Lyapunov redesign. Next, we introduce the stabilization technique through Immersion \& Invariance (I\&I) as a tool to improve the robustness of a given nonlinear controller with respect to unmodeled dynamics. The recursive Lyapunov redesign is then applied to the attitude stabilization of a spacecraft with flexible appendages and to the autopilot design of an asymmetric air-to-air missile. Contextually, we develop a systematic method to rapidly evaluate the aerodynamic perturbation terms exploiting the deterministic model of the uncertainty. The effectiveness of the proposed controller is highlighted through several simulations in the second case-study considered. In the final part of the work, the technique of I\& I is reformulated in the digital setting in the case of a special class of systems in feedback form, for which constructive continuous-time solutions exist, by means of backstepping and nonlinear domination arguments. The sampled-data implementation is based on a multi-rate control solution, whose existence is guaranteed for the class of systems considered. The digital controller guarantees, under sampling, the properties of manifold attractivity and trajectory boundedness. The control law, computed by finite approximation of a series expansion, is finally validated through numerical simulations in two academic examples and in two case-studies, namely the cart-pendulum system and the rigid spacecraft
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Lin, Shih-Hsiu, and 林世修. "Midcourse Guidance and Control of Missile Interceptor." Thesis, 2009. http://ndltd.ncl.edu.tw/handle/53460677676045372325.
Full textAbstract:
碩士<br>國立交通大學<br>電機與控制工程系所<br>97<br>The thesis investigates three-dimensional midcourse guidance and control for missile interceptor. When missile enter the midcourse phase we makes the time-to-go prediction first, then estimates the predicted interception point (PIP) in use of the time-to-go prediction. According to the estimation of the PIP, midcourse guidance law computes the variation of velocity that missile required. Since the commanded guidance acceleration is achieved by thrust vector control (TVC) system in midcourse phase, we design two loops flight control system to stabilize the attitude and achieve the guidance command by TVC system. The midcourse guidance we used can reduce the amount of computation and the sensitivity of estimation error, be implemented easily and guide missile effectively. We simulate three-dimensional missile and target engagement, analyze zero effort miss (ZEM) caused by various yaw angle at launch and discuss the actuator time constant effects on control system.
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He, Wei-Te, and 何威德. "Guidance and Attitude Control of Missile Interceptor." Thesis, 2008. http://ndltd.ncl.edu.tw/handle/49151549804706335393.
Full textAbstract:
碩士<br>國立交通大學<br>電機與控制工程系所<br>96<br>The thesis investigates three-dimensional midcourse and terminal guidance laws for missile interceptor and the associated divert and attitude control system design. Since the commanded guidance acceleration, in both the midcourse phase and terminal phase, is achieved using divert thrusters and in the terminal phase the infrared seeker needs to continuously lock-on to the target, true proportional navigation guidance law seems to be the most appropriate choice among the various guidance laws. We derive the three-dimensional true proportional navigation guidance law and analyze, via simulations, various system parameters such as radar measurement errors, seeker lock-on range, and flight control system bandwidth, have on the performance of the interceptor. Divert and attitude control system use on-off thrusters as actuators. Divert control achieves the command guidance acceleration using pulse width modulation (PWM), in which the pulse duration is proportional to the magnitude of acceleration command. Attitude control system is linear and designed using linearized model, it uses angle and rate feedback, achieves three-axis stabilization and is implemented using PWM. Simulations for both guidance and control use Standard Missile III as a prototype. Results from this research should be useful for future system and hardware development.
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Jiang, Jian-Chiang, and 姜健強. "NONLINEAR CONTROL DESIGN FOR MISSILE TERMINAL GUIDANCE." Thesis, 2000. http://ndltd.ncl.edu.tw/handle/66094603515549287920.
Full textAbstract:
碩士<br>國立交通大學<br>電機與控制工程系<br>88<br>In this thesis, we study the interception problem of homing missiles. From the robust stabilization point of view, we propose two of robust control techniques and use them on the design of missile guidance law. When the missile pursuits the target, the acceleration of target is usually unknown and can be regarded as external disturbance. According to the robustness of sliding-mode control, we use Variable Structure Control scheme to design missile guidance law. Moreover, H-infinity control scheme is also proposed for missile interception. We wish the ratio of output energy and input energy will less than or equal to the prescribed value. We use the capture time and the cumulative velocity increment as two performance indices. For the purpose of performance comparison, we compare the performance indices of two missile guidance laws with those of TPN and IPN. From simulation results, we can find that VSC can achieve good performance against both non-maneuvering and maneuvering target and the performance can be controlled by the parameters chosen by the designer.
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Xu, Jia-Ming, and 許家銘. "Guidance and Control Systems Design for Anti-Ballistic-Missile Missiles with Lateral Jets." Thesis, 2000. http://ndltd.ncl.edu.tw/handle/01223011414979502661.
Full textAbstract:
碩士<br>國立臺灣大學<br>電機工程學研究所<br>88<br>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.
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ZHUANG, DONG-MING, and 莊棟明. "Guidance and nonlinear control for a highly maneuvable missile." Thesis, 1993. http://ndltd.ncl.edu.tw/handle/68688876247380373809.
Full textAbstract:
碩士<br>國立臺灣大學<br>電機工程研究所<br>81<br>Contents
1 簡介
1.1 動機
1.2 文獻探討
1.3 論文的組織
2 飛彈系統描述
2.1 符號
2.2 飛彈的動態方程式
3 自動駕駛儀設計
3.1 介紹
3.2 自動駕駛儀設計
4 非線性飛彈導航律
4.1 模擬用模型
4.2 飛彈攔截的數學表示式
4.3 終端引律
5 飛彈導引控制系統
6 結論
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Abdul, Saleem P. K. "Guidance Laws for Engagement Time Control." Thesis, 2016. http://hdl.handle.net/2005/2926.
Full textAbstract:
Autonomous aerial vehicles like missiles and unmanned aerial vehicles (UAVs) have attracted various military and civilian applications. The primary guidance objective of any autonomous vehicle is to reach the desired destination point (target or waypoint). However, many practical engagements impose additional constraints like minimum control effort, a desired final velocity direction or a predefined engagement time. This thesis addresses engagement time constrained guidance problems pertaining to missiles and UAVs.
The first part of the thesis discusses a nonlinear guidance law for impact time control of missiles against stationary target. The guidance law is designed with a particular choice of missile heading error variation as a function of ran to-target. The proposed heading error variation leads to an exact closed-form expression for the impact time. controlling the impact time, a closed-form relation is derived relating the control parameter to the desired impact time. A new Lyapunov based guidance law with a monotonically decreasing lateral acceleration is proposed in the next part of the thesis. An exact expression for impact time with minimum and maximum achievable impact times is derived. A control parameter is proposed with a closed-form relationship to the desired impact time.
Using the concept of predicted interception point, the two guidance laws are extended for impact time control against non-maneuvering and moving targets. The proposed guidance models are extended to three-dimensional engagements by deducing yaw and pitch lateral accelerations satisfying the desired heading error profile. Extensive simulation studies are carried out for single missile and salvo attack scenarios.
The last part of the thesis presents a guidance methodology governing the arrival time of a UAV at a waypoint. A specific arrival angle is considered as an additional constraint. The arrival constraints are satisfied by varying the navigation gain of the proportional navigation guidance law. The methodology is applied for simultaneous and sequential arrival of UAVs at a waypoint.
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Su, Huai-Wen, and 蘇懷文. "Application of Gain Scheduling and Fuzzy Control in Missile Guidance and Control System esign." Thesis, 1999. http://ndltd.ncl.edu.tw/handle/72561600792637564124.
Full textAbstract:
碩士<br>逢甲大學<br>自動控制工程學系<br>87<br>Actual aerodynamically controlled missiles are nonlinear: their dynamics changes with the operating points and may be essential nonlinearities in the plant. For the sack of simplification in designs, traditional aerodynamically controlled missiles are usually guided by the proportional navigation guidance law, and proportional-integral-derivative (PID) flight control system. However, these approaches cannot cope with highly nonlinear behavior and provide satisfactory engagement performance and performance robustness. In this thesis, we propose a fully fuzzy logic-based guidance and control design. Estimated line-of-sight angle rate, relative range and estimated target evasive acceleration are used to constitute the rule antecedent of the guidance law improving engagement performance. The novel design relaxes the interceptor acceleration requirements and also yields smaller miss distances. Tracking error and change of tracking error constitute the rule antecedent of the fuzzy autopilot to provide excellent tracking performance and robustness. Besides, a new approach by using Takagi-Sugeno fuzzy logic and gain scheduling is presented to design the terminal guidance law. The proposed method has two major advantages over the existing fuzzy logic design. First of all, it provides a general and formally motivated method for the guidance rule design. Secondly, the method for determining the guidance command is general and computationally efficient.
The concept of gain scheduling, originated in connection with the development of flight control systems, is an effective way of controlling systems whose dynamics change with the operating conditions. It is thus a very useful technique for reducing the effect of system parameter variations. Traditionally, the guidance and control gains are switched along the missile trajectories according to a function or a table lookup built in the computer. Unlike conventional designs, we propose here a fuzzy rule-based scheme for gain scheduling of guidance and control systems. Takagi-Sugeno fuzzy logic model is used as a fuzzy inference mechanism to interpolate the guidance and control parameters around boundaries of flight conditions based on known local guidance and control parameters.
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Chang, Chin-Liang, and 張欽亮. "NONLINEAR GUIDANCE CONTROL OF MISSILE INTERCEPTION AND SOFT- LANDING ON AN ASTEROID." Thesis, 1997. http://ndltd.ncl.edu.tw/handle/83688577641922919145.
Full textAbstract:
碩士<br>國立交通大學<br>控制工程系<br>85<br>In this thesis, we study the problems of missile interception
and soft landing on an asteroid. First we employ
linear stability theory to design guidance control
laws for missile interception problem. Lyapunov
approach is then used to enlarge the capture region. However,
for the case of large capture distance and/or maneuvering
target, such a guidance control law is found to
fail to intercept the target. Variable structure
control type guidance laws are proposed to
tackle the problem. This is achieved by
treating the maneuver motion of the target as unmodeled dynamics
and suitably selecting the sliding surfaces. Moreover, feedback
linearization type control laws are also proposed for
missile interception. These two control schemes are shown to
be effective in intercepting both non-maneuvering and
maneuvering target. Same approaches are then employed to design
the guidance control laws for the soft-landing on an asteroid.
The main difference between interception and landing problems
is that both aero-drag and gravitational force are concerned
in landing problem but not in missile interception.
Numerical simulations are presented to demonstrate the main
results, which show that feedback linearization type guidance
control laws provide better performance in both missile
interception problem and the soft landing on an
asteroid.
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Hsieh, Chao-Yi, and 謝昭夷. "Multi-objective Missile Guidance Control with Stochastic Continuous Wiener and Discontinuous Poisson Noises." Thesis, 2017. http://ndltd.ncl.edu.tw/handle/ytsj7a.
Full textAbstract:
碩士<br>國立清華大學<br>電機工程學系所<br>105<br>This study proposes a Multi-objective (MO) guidance law simultaneously for optimal H_2 missile interception with stochastic continuous Wiener noise and discontinuous Poisson jump
noise as well as optimal H_∞ external disturbance filtering of external disturbance on missile guidance. The first design objective of optimal H_2 missile interception is to minimize the
effect of intrinsic stochastic Wiener noise due to modeling uncertainty of the missile and the accumulated angle error of the gyroscope as well as the intrinsic stochastic Poisson jump noise
due to the inaccurate radar measurement of the missile because of the target suddenly side-step maneuver. The second design objective of H_∞ external disturbance filtering is to minimize
the effect of external disturbance due to the target’s acceleration on the missile guidance. An indirect method is proposed to solve the MO H_2/H_∞ guidance problem of missiles. In order to avoid solving a Hamilton-Jacobin inequality (HJI)-constrained MO H_2/H_∞ missile guidance problem based on Pareto optimal solution, fuzzy interpolation method is proposed to transform the HJI-constrained MO missile guidance problem to a linear matrix inequalities
(LMIs)-constrained MO missile guidance problem. An LMIs-based MO Evolutionary Algorithm (MOEA) is also proposed to solve the MO H_2/H_∞ missile guidance problem.
Finally, a simulation example is conducted to illustrate the design procedure and to validate the performance of the proposed MO H_2/H_∞ guidance law.
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Kumar, Shashi Ranjan. "Sliding Mode Control Based Guidance Strategies with Terminal Constraints." Thesis, 2015. http://etd.iisc.ernet.in/2005/3876.
Full textAbstract:
In the guidance literature, minimizing miss distance along with optimizing the energy usage had been an objective for several decades. In current day applications, additional terminal performance such as impact angle and impact time are of paramount importance. These terminal constraints increase warhead effectiveness and survivability of the interceptor. This thesis contributes to the design of guidance laws addressing terminal constraints such as impact angle, impact time, and both impact time as well as impact angle, in addition to interception of targets.
In the first part of the thesis, the guidance laws which ensure the alignment of the interceptor at a desired impact angle within a finite time is proposed using different variants of sliding mode control(SMC).The impact angle is first redefined in terms of line-of-sight angle and then the impact angle problem is converted to a simpler problem of controlling line-of-sight angle and their rates. The sliding mode capturability and interpretation of the guidance laws are presented. In order to cater to very large heading angle errors, which give rise to negative closing speed initially, modifications to the guidance laws are also suggested. The modifications to the guidance laws for avoiding singularities, which may be encountered during implementation, due to the inherent nature of terminal SMC, are suggested. However, the guidance laws, which alleviates the possibility of such singularities completely, are also designed by using non singular terminal SMC. The two loop guidance and control, for a skid-to-turn cruciform interceptor in the pitch plane, is also proposed with an autopilot designed using the concept of dynamic SMC. The guidance laws addressing impact angle constraint for three dimensional scenarios are also presented. Unlike the usual approach of decoupling the three dimensional engagement in to two mutually orthogonal planar engagements, the guidance laws are derived using coupled engagement dynamics. These guidance laws are designed using conventional and non singular terminal SMC and provide asymptotic and finite time alignment of the intercept or to the desired impact angles, respectively.
Next, the SMC based guidance laws which ensure the interception of targets at pre-specified impact times is proposed in this thesis. The guidance law is first designed for stationary targets and then extended to constant velocity targets using the notion of predicted interception point. A switching surface is designed using the concepts of collision course and time-to-go with non-linear engagement dynamics and its role in achieving the objectives is also discussed. In order to account for large heading angle errors and even for negative initial closing speeds, different methods of estimation of time-to-go, resulting in two different guidance laws, are used. Unlike the existing guidance laws, the proposed guidance laws achieve an impact time even less than its initially estimated value. The flexibility in selecting a desired impact time is also exploited using the maximum available acceleration information. A cooperative salvo attack strategy, based on the proposed impact time guidance law, with a desired impact time chosen in real time using a centralized coordination algorithm, is proposed for stationary targets. The coordination manager determines a common impact time based on time-to-goof the interceptors, by minimizing the total switching surface deviations which in turn reduces the control effort.
The thesis also proposes a SMC based guidance strategy which addresses impact angle and impact time constraints simultaneously. This guidance scheme is based on switching between impact time and impact angle guidance laws based on certain conditions. Unlike existing impact time guidance laws, the proposed guidance strategy takes into account the curvature of the trajectory due to the impact angle requirement. The interceptor first corrects its course to nullify the impact time error and then aims to achieve interception with desired impact angle. In order to reduce the transitions between the two guidance laws, a novel hysteresis loop is introduced in the switching conditions. Initially stationary targets are considered, and later the same guidance scheme is extended to constant velocity targets using the notion of predicted interception point.
Theclaimsofalltheguidancelawsarevalidatedwithextensivesimulationsandtheir performances are compared with existing guidance laws. Although all the guidance laws derived in the thesis are based on the assumption of constant speed interceptors, their performances are evaluated with a time-varying speed interceptor model, subjected to aerodynamic conditions, to validate their efficacy. The implementation of impact time guidance on time-varying speed interceptors is a formidable challenge in the guidance literature. Such implementations have also been presented in the thesis after introducing the notion of average speed and shown to yield satisfactory performance.
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Saroj, Kumar G. "An Integrated Estimation-Guidance Approach for Seeker-less Interceptors." Thesis, 2015. http://etd.iisc.ernet.in/2005/3828.
Full textAbstract:
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.
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Costa, António Rui Moreira Tinoco da. "Guidance of interceptor missiles based on Robust Control." Master's thesis, 2018. http://hdl.handle.net/10400.6/8555.
Full textAbstract:
Missiles development are constantly evolving. This is mainly due to the significantly increase in the performance of the missiles means of transportation (aircrafts, vessels, submarines, trucks and trains), allowing bigger and heavier armament, which results directly in much more precise control systems, with a capacity for different types of warheads, as well as an ability to store larger amounts of fuel.
Regarding the subject addressed in this thesis, it should be taking into consideration that a tactical missile has to be quite versatile, as it can either aim to shoot down an aircraft with high manoeuvrability or a cruise missile with a predefined trajectory, being thus necessary to withstand high speeds and g force.
A control system for a missile is responsible for its attitude, while missile guidance system is responsible for controlling its trajectories and, therefore, being able to detect that the missile is outside the interception trajectory, requiring an input signal to put it back on collision course. The focus of this dissertation is on the control of the trajectories of a tactical missile, which has to be capable of performing the basic function of detecting the signals received by the command, which in its turn will be applied to the control system.
An H8/LTR controller and the Artstein method applied on a Robust LQR controller were applied to the missile, where it’s concluded that the first one has a better performance for manoeuvrable or non-manoeuvrable targets. However, Robust LQR method reveals a strong potential when implemented to solve systems in which perturbations predominate, thus making the behaviour of both methods very similar.<br>O desenvolvimento dos mísseis está em constante evolução. Tal se deve principalmente ao aumento significativo do desempenho dos meios de transporte destes (aeronaves, embarcações, submarinos, camiões e comboios), permitindo assim transportar armamento de maiores dimensões e peso, o qual resulta diretamente em sistemas de controlo muito mais precisos, com uma capacidade para diferentes tipos de ogivas e armazenamento de maiores quantidades de combustível.
Relativamente ao assunto abordado neste trabalho, é preciso ter em conta que um míssil tático tem de ser bastante versátil, pois tanto pode ter como alvo a abater uma aeronave com elevada manobrabilidade ou um míssil de cruzeiro com uma trajetória pré-definida, sendo assim necessário suportar elevadas velocidades e força g.
Um sistema de controlo para um míssil é responsável pela sua atitude, enquanto o sistema de orientação deste é responsável pelo controlo das suas trajetórias, tendo assim de ser capaz de detetar que o míssil se encontra fora da trajetória de interceção com o alvo, necessitando de receber uma entrada que o volte a colocar na rota de colisão. O foco desta dissertação é no controlo das trajetórias de um míssil tático, tendo este de ser capaz de cumprir a função básica de detetar os sinais recebidos pelo comando, os quais por sua vez serão aplicados ao sistema de controlo, o que se resume em alterações do rumo do míssil.
Foi aplicado um sistema de orientação H8/LTR, bem como o método de Artstein a um LQR Robusto, onde se conclui que o primeiro apresenta um melhor desempenho tanto para alvos sem manobrabilidade como com manobrabilidade. Porém, é necessário ter em conta que o método do LQR Robusto revela um forte potencial quando implementado para solucionar sistemas nos quais predominem perturbações, fazendo assim com que o comportamento dos dois métodos seja bastante semelhante.
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(11068791), Athul Pradeepkumar Girija. "A Systems Framework and Analysis Tool for Rapid Conceptual Design of Aerocapture Missions." Thesis, 2021.
Find full textAbstract:
Aerocapture offers a near propellantless and quick method of orbit
insertion at atmosphere bearing planetary destinations. Compared to
conventional propulsive insertion, the primary advantage of using
aerocapture is the savings in propellant mass which could be used to
accommodate more useful payload. To protect the spacecraft from the
aerodynamic heating during the maneuver, the spacecraft must be enclosed
in a protective aeroshell or deployable drag device which also provides
aerodynamic control authority to target the desired conditions at
atmospheric exit. For inner planets such as Mars and Venus, aerocapture
offers a very attractive option for inserting small satellites or
constellations into very low circular orbits such as those used for
imaging or radar observations. The large amount of propellant required
for orbit insertion at outer planets such as Uranus and Neptune severely
limits the useful payload mass that can delivered to orbit as well as
the achievable flight time. For outer planet missions, aerocapture opens
up an entirely new class of short time of flight trajectories which are
infeasible with propulsive insertion. A systems framework for rapid
conceptual design of aerocapture missions considering the
interdependencies between various elements such as interplanetary
trajectory and vehicle control performance for aerocapture is presented.
The framework provides a step-by-step procedure to formulate an
aerocapture mission starting from a set of mission objectives. At the
core of the framework is the ``aerocapture feasibility chart", a
graphical method to visualize the various constraints arising from
control authority requirement, peak deceleration, stagnation-point peak
heat rate, and total heat load as a function of vehicle aerodynamic
performance and interplanetary arrival conditions. Aerocapture
feasibility charts have been compiled for all atmosphere-bearing Solar
System destinations for both lift and drag modulation control
techniques. The framework is illustrated by its application to
conceptual design of a Venus small satellite mission and a
Flagship-class Neptune mission using heritage blunt-body aeroshells.
The framework is implemented in the Aerocapture Mission Analysis Tool
(AMAT), a free and open-source Python package, to enable scientists and
mission designers perform rapid conceptual design of aerocapture
missions. AMAT can also be used for rapid Entry, Descent, and Landing
(EDL) studies for atmospheric probes and landers at any
atmosphere-bearing destination.
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(11014071), Vivek Muralidharan. "Stretching Directions in Cislunar Space: Stationkeeping and an application to Transfer Trajectory Design." Thesis, 2021.
Find full textAbstract:
<div>The orbits of interest for potential missions are stable or nearly stable to maintain long term presence for conducting scientific studies and to reduce the possibility of rapid departure. Near Rectilinear Halo Orbits (NRHOs) offer such stable or nearly stable orbits that are defined as part of the L1 and L2 halo orbit families in the circular restricted three-body problem. Within the Earth-Moon regime, the L1 and L2 NRHOs are proposed as long horizon trajectories for cislunar exploration missions, including NASA's upcoming Gateway mission. These stable or nearly stable orbits do not possess well-distinguished unstable and stable manifold structures. As a consequence, existing tools for stationkeeping and transfer trajectory design that exploit such underlying manifold structures are not reliable for orbits that are linearly stable. The current investigation focuses on leveraging stretching direction as an alternative for visualizing the flow of perturbations in the neighborhood of a reference trajectory. The information supplemented by the stretching directions are utilized to investigate the impact of maneuvers for two contrasting applications; the stationkeeping problem, where the goal is to maintain a spacecraft near a reference trajectory for a long period of time, and the transfer trajectory design application, where rapid departure and/or insertion is of concern.</div><div><br></div><div>Particularly, for the stationkeeping problem, a spacecraft incurs continuous deviations due to unmodeled forces and orbit determination errors in the complex multi-body dynamical regime. The flow dynamics in the region, using stretching directions, are utilized to identify appropriate maneuver and target locations to support a long lasting presence for the spacecraft near the desired path. The investigation reflects the impact of various factors on maneuver cost and boundedness. For orbits that are particularly sensitive to epoch time and possess distinct characteristics in the higher-fidelity ephemeris model compared to their CR3BP counterpart, an additional feedback control is applied for appropriate phasing. The effect of constraining maneuvers in a particular direction is also investigated for the 9:2 synodic resonant southern L2 NRHO, the current baseline for the Gateway mission. The stationkeeping strategy is applied to a range of L1 and L2 NRHOs, and validated in the higher-fidelity ephemeris model.</div><div><br></div><div>For missions with potential human presence, a rapid transfer between orbits of interest is a priority. The magnitude of the state variations along the maximum stretching direction is expected to grow rapidly and, therefore, offers information to depart from the orbit. Similarly, the maximum stretching in reverse time, enables arrival with a minimal maneuver magnitude. The impact of maneuvers in such sensitive directions is investigated. Further, enabling transfer design options to connect between two stable orbits. The transfer design strategy developed in this investigation is not restricted to a particular orbit but applicable to a broad range of stable and nearly stable orbits in the cislunar space, including the Distant Retrograde Orbit (DROs) and the Low Lunar Orbits (LLO) that are considered for potential missions. Examples for transfers linking a southern and a northern NRHO, a southern NRHO to a planar DRO, and a southern NRHO to a planar LLO are demonstrated.</div>
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Maity, Arnab. "Optimal Guidance Of Aerospace Vehicles Using Generalized MPSP With Advanced Control Of Supersonic Air-Breathing Engines." Thesis, 2012. http://etd.iisc.ernet.in/handle/2005/2550.
Full textAbstract:
A new suboptimal guidance law design approach for aerospace vehicles is proposed in this thesis, followed by an advanced control design for supersonic air-breathing engines. The guidance law is designed using the newly developed Generalized Model Predictive Static Programming (G-MPSP), which is based on the continuous time nonlinear optimal control framework. The key feature of this technique is one-time backward propagation of a small-dimensional weighting matrix dynamics, which is used to update the entire control history. This key feature, as well as the fact that it leads to a static optimization problem, lead to its computational efficiency. It has also been shown that the existing model predictive static programming (MPSP), which is based on the discrete time framework, is a special case of G-MPSP. The G-MPSP technique is further extended to incorporate ‘input inequality constraints’ in a limited sense using the penalty function philosophy. Next, this technique has been developed also further in a ‘flexible final time’ framework to converge rapidly to meet very stringent final conditions with limited number of iterations.
Using the G-MPSP technique in a flexible final time and input inequality constrained formulation, a suboptimal guidance law for a solid motor propelled carrier launch vehicle is successfully designed for a hypersonic mission. This guidance law assures very stringent final conditions at the injection point at the end of the guidance phase for successful beginning of the hypersonic vehicle operation. It also ensures that the angle of attack and structural load bounds are not violated throughout the trajectory. A second-order autopilot has been incorporated in the simulation studies to mimic the effect of the inner-loops on the guidance performance. Simulation studies with perturbations in the thrust-time behaviour, drag coefficient and mass demonstrate that the proposed guidance can meet the stringent requirements of the hypersonic mission.
The G-MPSP technique in a fixed final time and input inequality constrained formulation has also been used for optimal guidance of an aerospace vehicle propelled by supersonic air-breathing engine, where the resulting thrust can be manipulated by managing the fuel flow and nozzle area (which is not possible in solid motors). However, operation of supersonic air-breathing engines is quite complex as the thrust produced by the engine is a result of very complex nonlinear combustion dynamics inside the engine. Hence, to generate the desired thrust, accounting for a fairly detailed engine model, a dynamic inversion based nonlinear state feedback control design has been carried out. The objective of this controller is to ensure that the engine dynamically produces the thrust that tracks the commanded value of thrust generated from the guidance loop as closely as possible by regulating the fuel flow rate. Simultaneously, by manipulating throat area of the nozzle, it also manages the shock wave location in the intake for maximum pressure recovery with sufficient margin for robustness. To filter out the sensor and process noises and to estimate the states for making the control design operate based on output feedback, an extended Kalman filter (EKF) based state estimation design has also been carried out and the controller has been made to operate based on estimated states. Moreover, independent control designs have also been carried out for the actuators so that their response can be faster. In addition, this control design becomes more challenging to satisfy the imposed practical constraints like fuel-air ratio and peak combustion temperature limits. Simulation results clearly indicate that the proposed design is quite successful in assuring the desired performance of the air-breathing engine throughout the flight trajectory, i.e., both during the climb and cruise phases, while assuring adequate pressure margin for shock wave management.