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1
Montalvo, Carlos. "Meta aircraft flight dynamics and controls." Diss., Georgia Institute of Technology, 2014. http://hdl.handle.net/1853/51854.
Full textAbstract:
The field of mobile robotic systems has become a rich area of research and design. These systems can navigate difficult terrain using multiple actuators with conventional ambulation, by hopping, jumping, or for aerial vehicles, using flapping wings, propellers, or engines to maintain aerial flight. Unmanned Aerial Systems(UAS) have been used extensively in both military and civilian applications such as reconnaissance or search and rescue missions. For air vehicles, range and endurance is a crucial design parameter as it governs which missions can be performed by a particular vehicle. In addition, when considering the presence of external disturbances such as atmospheric winds, these missions can be even more challenging. Meta aircraft technologies is one area of research that can increase range and endurance by taking advantage of an increase in L/D. A meta aircraft is an aircraft composed of smaller individual aircraft connected together through a similar connection mechanism that can potentially transfer power, loads, or information. This dissertation examines meta aircraft flight dynamics and controls for a variety of different configurations. First, the dynamics of meta aircraft systems are explored with a focus on the changes in fundamental aircraft modes and flexible modes of the system. Specifically, when aircraft are connected, the fundamental modes change, can become overdamped or even unstable. In addition, connected aircraft exhibit complex flexible modes and mode shapes that change based on the parameters of the connection joint and the number of connected aircraft. Second, the connection dynamics are explored for meta aircraft where the vehicles are connected wing tip to wing tip using passive magnets with a particular focus on modeling the connection event between aircraft in a practical environment. It is found that a multi-stage connection control law with position and velocity feedback from GPS and connection point image feedback from a camera yields adequate connection performance in the presence of realistic sensor errors and atmospheric winds. Furthermore, atmosphericwinds with low frequency gusts at the intensity normally found in a realistic environment pose the most significant threat to the success of connection. The frequency content of the atmospheric disturbance is an
important variable to determine success of connection. Finally, the geometry of magnets that create the connection force field can alter connection rates. Finally, the performance of a generic meta aircraft system are explored. Using a simplified rigid body model to approximate any meta aircraft configuration, adequate connection is achieved in the presence of realistic winds. Using this controller overall performance is studied. In winds, there is an overall decrease in outer loop performance for meta aircraft. However, inner loop performance increases for meta aircraft. In addition, the aerodynamic benefit of different configurations are investigated. Wing to wing tip connected flight provides the most benefit in terms of average increased Lift to Drag ratio while tip to tail configurations drop the Lift to Drag ratio as trailing aircraft fly in the downwash of the leading aircraft.
2
Bhardwaj, Pradeep. "Aircraft cruise performance optimization using chattering controls." Thesis, Virginia Tech, 1986. http://hdl.handle.net/10919/45750.
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Aircraft Cruise Performance is examined by using energy-state modelling to investigate fuel-range optimal trajectories. Chattering controls are considered appropriate when the hodograph is non-convex.
Classical steady-state cruise, simple chattering-cruise and the extended chattering-cruise models are studied as constrained parameter-optimization problems. The term "extended chattering" refers to vehicle system modelling extended to maintain vertical equilibrium only on the average. Numerical solution is obtained using a variable-metric gradient-protection algorithm and computational results are presented for three different aircraft.
This study shows that simple chattering cruise for certain specific energies can result in substantial fuel savings over classical steady-state cruise. However extended chattering cruise results in only marginal fuel savings when compared to simple chattering cruise.Master of Science
3
Cadwell, John Andres Jr. "Control of Longitudinal Pitch Rate as Aircraft Center of Gravity Changes." DigitalCommons@CalPoly, 2010. https://digitalcommons.calpoly.edu/theses/426.
Full textAbstract:
In order for an aircraft to remain in stable flight, the center of gravity (CG) of an aircraft must be located in front of the center of lift (CL). As the center of gravity moves rearward, pitch stability decreases and the sensitivity to control input increases. This increase in sensitivity is known as pitch gain variance. Minimizing the pitch gain variance results in an aircraft with consistent handling characteristics across a broad range of center of gravity locations.
This thesis focuses on the development and testing of an open loop computer simulation model and a closed loop control system to minimize pitch axis gain variation as center of gravity changes. DATCOM and MatLab are used to generate the open loop aircraft flight model; then a closed loop PD (proportional-derivate) controller is designed based on Ziegler-Nichols closed loop tuning methods. Computer simulation results show that the open loop control system exhibited unacceptable pitch gain variance, and that the closed loop control system not only minimizes gain variance, but also stabilizes the aircraft in all test cases. The controller is also implemented in a Scorpio Miss 2 radio controlled aircraft using an onboard microprocessor. Flight testing shows that performance is satisfactory.
4
Ur, Rahman Naveed. "Propulsion and flight controls integration for the blended wing body aircraft." Thesis, Cranfield University, 2009. http://hdl.handle.net/1826/4095.
Full textAbstract:
The Blended Wing Body (BWB) aircraft offers a number of aerodynamic perfor-
mance advantages when compared with conventional configurations. However, while
operating at low airspeeds with nominal static margins, the controls on the BWB
aircraft begin to saturate and the dynamic performance gets sluggish. Augmenta-
tion of aerodynamic controls with the propulsion system is therefore considered in
this research. Two aspects were of interest, namely thrust vectoring (TVC) and flap
blowing. An aerodynamic model for the BWB aircraft with blown flap effects was
formulated using empirical and vortex lattice methods and then integrated with a
three spool Trent 500 turbofan engine model. The objectives were to estimate the
effect of vectored thrust and engine bleed on its performance and to ascertain the
corresponding gains in aerodynamic control effectiveness.
To enhance control effectiveness, both internally and external blown flaps were sim-
ulated. For a full span internally blown flap (IBF) arrangement using IPC flow, the
amount of bleed mass flow and consequently the achievable blowing coefficients are
limited. For IBF, the pitch control effectiveness was shown to increase by 18% at low
airspeeds. The associated detoriation in engine performance due to compressor bleed
could be avoided either by bleeding the compressor at an earlier station along its ax-
ial length or matching the engine for permanent bleed extraction. For an externally
blown flap (EBF) arrangement using bypass air, high blowing coefficients are shown
to be achieved at 100% Fan RPM. This results in a 44% increase in pitch control
authority at landing and take-off speeds. The main benefit occurs at take-off, where
both TVC and flap blowing help in achieving early pitch rotation, reducing take-off
field lengths and lift-off speeds considerably. With central flap blowing and a lim-
ited TVC of 10◦, the lift-off range reduces by 48% and lift-off velocity by almost 26%.
For the lateral-directional axis it was shown that both aileron and rudder control
powers can be almost doubled at a blowing coefficient of Cu = 0.2. Increased roll
authority greatly helps in achieving better roll response at low speeds, whereas the
increased rudder power helps in maintaining flight path in presence of asymmetric
thrust or engine failure, otherwise not possible using the conventional winglet rudder.
5
Nelson, Mark David. "A Comparison of Two Methods Used to Deal with Saturation of Multiple, Redundant Aircraft Control Effectors." Thesis, Virginia Tech, 2001. http://hdl.handle.net/10919/34673.
Full textAbstract:
A comparison of two methods to deal with allocating controls for unattainable moments in an aircraft was performed using a testbed airframe that resembled an F/A-18 with a large control effector suite. The method of preserving the desired moment direction to deal with unattainable moments is currently used in a specific control allocator. A new method of prioritizing the pitch axis is compared to the moment-direction preservation. Realtime piloted simulations are completed to evaluate the characteristics and performance of these methods.
A direct comparison between the method of preserving the moment direction by scaling the control solution vector and prioritizing the pitching moment axis is performed for a specific case. Representative maneuvers are flown with a highly unstable airframe to evaluate the ability to achieve the specific task. Flight performance and pilot interpretation are used to evaluate the two methods.
Pilot comments and performance results favored the method of pitch-axis prioritization. This method provided favorable flight characteristics compared to the alternative method of preserving the moment direction for the specific tasks detailed in this paper.
NOTE: An updated copy of this ETD was added on 09/28/2010.Master of Science
6
Johnson, Bruce, and John Smith. "CAN BUS USED FOR DATA ACQUISITION SYSTEM CONTROLS (AUTOMOTIVE SOLUTION FOR AIRCRAFT PROBLEM)." International Foundation for Telemetering, 2005. http://hdl.handle.net/10150/604882.
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ITC/USA 2005 Conference Proceedings / The Forty-First Annual International Telemetering Conference and Technical Exhibition / October 24-27, 2005 / Riviera Hotel & Convention Center, Las Vegas, NevadaThis paper discusses using the CAN (Control Area Network) Bus protocol for control and status of flight test data acquisition systems. The application of the CAN (Control Area Network) on an F/A-18 aircraft will be discussed in detail.
7
ALONSO, ELENA. "CONTROL DESIGN AND IMPLEMENTATION FOR THE SELF-SEPARATION OF IN-TRAIL AIRCRAFT." University of Cincinnati / OhioLINK, 2005. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1116261083.
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Schmollgruber, Peter. "Enhancement of the conceptual aircraft design process through certification constraints management and full mission simulations." Thesis, Toulouse, ISAE, 2018. http://www.theses.fr/2018ESAE0036.
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La conception d'un nouvel avion est initiée durant la phase avant-projet. Dans un premier temps, lesconcepteurs d’aéronefs identifient un ensemble de concepts potentiels pouvant répondre aux exigencesdu client en s’appuyant sur des informations fournies par les spécialistes disciplinaires et expertssystème. Ensuite, les solutions sont évaluées via un processus de dimensionnement basé sur uneanalyse multidisciplinaire. Dans le domaine des avions de transport civil, les objectifs ambitieux entermes de consommation de carburant amènent à étudier des configurations innovantes incluant denouvelles technologies. Cependant, peu de données sur de telles architectures sont disponibles dans lesphases amont de la conception. Ainsi, afin d'éviter une sélection ou élimination erronée d'unesolution, un objectif clé de la recherche en conception d'aéronefs est l’ajout de connaissances dansl'analyse multidisciplinaire.Aujourd’hui, cet objectif est atteint avec différentes approches: application d’optimisationsmultidisciplinaires, ajout de précision grâce aux analyses haute fidélité, introduction de nouvellesdisciplines ou systèmes et enfin, gestion de l'incertitude. Le rôle du concepteur est alors de combinerces options dans un processus de conception multidisciplinaire afin de converger vers le concept leplus performant tout en répondant aux contraintes de certification. Afin d’illustrer ce processus,l’optimisation d'un avion de transport avec assistance au sol pour le décollage qui a mis en évidencel'impact des contraintes de certification sur la conception du véhicule a été effectuée. La revuesuccessive des textes réglementaires et de recherches associées de la gestion du trafic aérien ont concluà la nécessité d’inclure des simulations au sein de l’analyse multidisciplinaire. Tenant compte de cesconclusions, la recherche effectuée dans le cadre de cette thèse propose alors d’ajouter desconnaissances en développant l’analyse et l’optimisation de la conception multidisciplinaire avec unnouveau module de contrainte de certification et des fonctionnalités de simulation complètes.Développé dans le cadre de la thèse, le module de contraintes de certification (CCM) a été utilisé pourrésoudre quatre problèmes d’optimisation associés à un avion de transport civil classique basé surl’outil de dimensionnement ONERA / ISAE-SUPAERO appelé FAST. Grâce à l'interface utilisateurdu CCM, un gain de temps au niveau de la mise en place de ces optimisations a constaté. De plus, lesrésultats ont confirmé la nécessité de définir au mieux et dès que possible les contraintes decertification.Pour atteindre des capacités de simulation complètes, l'analyse multidisciplinaire au sein de FAST aété améliorée. Premièrement, l'outil d'analyse aérodynamique a été modifié afin de générer la base dedonnées complète pour alimenter un modèle à 6 degrés de liberté. Ensuite, un nouveau module decalcul des propriétés d'inertie a été ajouté. Enfin, le simulateur open source JSBSim a été utilisé avecdifférentes lois de contrôle pour augmenter la stabilité et permettre la navigation automatisée. Lacomparaison entre les trajectoires de vol obtenues avec FAST et les données réelles sur les avionsenregistrées avec une antenne ADS-B a confirmé la validité de l'approcheThe design of a new aircraft is initiated at the conceptual design phase. In an initial step, aircraftdesigners, disciplinary and subsystems experts identify a set of potential concepts that could fulfill thecustomer requirements. To select the most promising candidates, aircraft designers carry out the sizingprocess through a Multidisciplinary Design Analysis. Nowadays, in the field of civil transport aircraft,environmental constraints set challenging goals in terms of fuel consumption for the next generationsof airplanes. With the “tube and wing” configuration offering low expectations on furtherimprovements, disruptive vehicle concepts including new technologies are investigated. However,little information on such architectures is available in the early phases of the design process. Thus, inorder to avoid mistakenly selecting or eliminating a wrong concept, a key objective in Aircraft Designresearch is to add knowledge in the Multidisciplinary Design Analysis.Nowadays, this objective is achieved with different approaches: implementation of MultidisciplinaryDesign Optimization, addition of accuracy through high fidelity analyses, introduction of newdisciplines or systems and uncertainty management. The role of the aircraft designer is then tocombine these options in a multidisciplinary design process to converge to the most promising conceptmeeting certification constraints. To illustrate this process, the optimization of a transport aircraftfeaturing ground based assistance has been performed. Using monolithic optimization architecture andadvanced structural models for the wing and fuselage, this study emphasized the impact ofcertification constraints on final results. Further review of the regulatory texts concluded that aircraftsimulation capabilities are needed to assess some requirements. The same need has been identified inthe field of Air Traffic Management that provides constraints for aircraft operations. This researchproposes then to add knowledge through an expansion of the Multidisciplinary Design Analysis andOptimization with a new Certification Constraint Module and full simulation capabilities.Following the development of the Certification Constraint Module (CCM), its capabilities have beenused to perform four optimization problems associated to a conventional civil transport aircraft basedon the ONERA / ISAE-SUPAERO sizing tool called FAST. Facilitated by the Graphical UserInterface of the CCM, the setup time of these optimizations has been reduced and the results clearlyconfirmed the necessity to consider certification constraints very early in the design process in order toselect the most promising concepts.To achieve full simulation capabilities, the multidisciplinary analysis within FAST had to beenhanced. First, the aerodynamics analysis tool has been modified so that necessary coefficients for a6 Degrees-of-Freedom model could be generated. Second, a new module computing inertia propertieshas been added. Last, the open source simulator JSBSim has been used including different controllaws for stability augmentation and automated navigation. The comparison between flight trajectoriesobtained with FAST and real aircraft data recorded with ADS-B antenna confirmed the validity of theapproach
9
Ledet, Jeffrey H. "Simulation and Performance Evaluation of Algorithms for Unmanned Aircraft Conflict Detection and Resolution." ScholarWorks@UNO, 2016. http://scholarworks.uno.edu/td/2168.
Full textAbstract:
The problem of aircraft conflict detection and resolution (CDR) in uncertainty is addressed in this thesis. The main goal in CDR is to provide safety for the aircraft while minimizing their fuel consumption and flight delays. In reality, a high degree of uncertainty can exist in certain aircraft-aircraft encounters especially in cases where aircraft do not have the capabilities to communicate with each other. Through the use of a probabilistic approach and a multiple model (MM) trajectory information processing framework, this uncertainty can be effectively handled. For conflict detection, a randomized Monte Carlo (MC) algorithm is used to accurately detect conflicts, and, if a conflict is detected, a conflict resolution algorithm is run that utilizes a sequential list Viterbi algorithm. This thesis presents the MM CDR method and a comprehensive MC simulation and performance evaluation study that demonstrates its capabilities and efficiency.
10
Hejkalová, Anna. "Návrh UL letounu pro piloty s pohybovým omezením." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2020. http://www.nusl.cz/ntk/nusl-417581.
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The diploma thesis deals with the issue of operating aircrafts by people with physical disabi-lities. Firstly, it formulates medical eligibility requirements for pilots with mobility restricti-ons. Because there is a large number of movement restrictions and each of them is very indi-vidual, an adaptation for only one of these categories was chosen for the conceptual design of the aircraft in this diploma thesis: limitation of the lower half of the body. The aircraft is designed to best meet the needs of pilots with this limitation of movement, with emphasis on ergonomics, the possibility of independent use without any help of other people and espe-cially on the safety of the user. Furthermore, the work examines possibilities of adjusting the control of aircraft for the disabled, tries to choose the most suitable one which is then desci-bed. In the end, an examination of the UL2 regulation requirements for the designed aircraft is done.
11
Lewis, Benjamin Paul. "A Visual Return-to-Home System for GPS-Denied Flight." BYU ScholarsArchive, 2016. https://scholarsarchive.byu.edu/etd/6254.
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Unmanned aerial vehicle technology is rapidly maturing. In recent years, the sight of hobbyist aircraft has become more common. Corporations and governments are also interested in using drone aircraft for applications such as package delivery, surveillance and communications. These autonomous UAV technologies demand robust systems that perform under any circumstances. Many UAV applications rely on GPS to obtain information about their location and velocity. However, the GPS system has known vulnerabilities, including environmental signal degradation, terrestrial or solar weather, or malicious attacks such as GPS spoofing. These conditions occur with enough frequency to cause concern. Without a GPS signal, the state estimation in many autopilots quickly degrades. In the absence of a reliable backup navigation scheme, this loss of state will cause the aircraft to drift off course, and in many cases the aircraft will lose power or crash. While no single approach can solve all of the issues with GPS signal degradation, individual events can be addressed and solved. In this thesis, we present a system which will return an aircraft to its launch point upon the loss of GPS. This functionality is advantageous because it allows recovery of the UAV in circumstances which the lack of GPS information would make difficult. The system presented in this thesis accomplishes the return of the aircraft by means of onboard visual navigation, which removes the dependence of the aircraft on external sensors and systems. The system presented here uses an downward-facing onboard camera and computer to capture a string of overlapping images (keyframes) of the ground as the aircraft travels on its outbound journey. When a signal is received, the aircraft switches into return-to-home mode. The system uses the homography matrix and other vision processing techniques to produce information about the location of the current keyframe relative to the aircraft. This information is used to navigate the aircraft to the location of each saved keyframe in reverse order. As each keyframe is reached, the system programmatically loads the next target keyframe. By following the chain of keyframes in reverse, the system reaches the launch location. Contributions in this thesis include the return-to-home visual flight system for UAVs, which has been tested in simulation and with flight tests. Features of this system include methods for determining new keyframes and switching keyframes on the inbound flight, extracting data between images, and flight navigation based on this information. This system is a piece of the wider GPS-denied framework under development in the BYU MAGICC lab.
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Maharaj, Davendra Yukteshwar. "The application of nonlinear control theory to robust helicopter flight control." Thesis, Imperial College London, 1994. http://hdl.handle.net/10044/1/7420.
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Hopper, David John Frederick. "Active control of V/STOL aircraft." Thesis, University of Salford, 1990. http://usir.salford.ac.uk/14698/.
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Vertical/Short Take-Off arid Landing (V/STOL) fighter aircraft are characterised by increased control complexity caused by the extra degree ol freedom. This can result in a high pilot workload which may be alleviated with the careful application of active flight control. However, the advent of control configured vehicles demands that the controller design must be part of a fully integrated and iterative aircraft design; hence it must allow the two-way flow of design information. In this thesis a suitable controller design method is developed to solve this two-fold problem. The method is based upon a singular perturbation analysis which is used to expose the underlying dynamics of a closed-loop state-space system. developments are described which allow high-order, dynamically complex parasitics, such as actuators, to be included in the design. Furthermore, the method gives the designer insight into the problem allowing tuning and engineering trade-offs to be performed intelligently with a two-way flow of design information. The end result is a robust high-gain multivariable controller. In order fully to develop arid analyse the method it has been applied to a representative non-linear time-varying aircraft simulation model. This LS supplied by the Royal Aerospace Establishment, Bedford. The necessary slate-space matrices are otitairted by lirLearisirig the model at several different flight cases. This occurs over a wide flight envelope, from hover to 300 Kts, and consequently the multivariable control laws are implemented using gain scheduling. Finally, task tailored control and handling qualities requirements are derived for a V/STOL aircraft in the form of a design brief. This design brief is then fulfilled by designing a controller which alleviates pilot workload during transitions from jet-borne to fully wing-borne flight (and vice versa).
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Hinson, Brian Thomas. "Adaptive control of elastic aircraft." Thesis, Wichita State University, 2010. http://hdl.handle.net/10057/3722.
Full textAbstract:
This thesis documents the development of an adaptive controller designed to control the elastic aircraft
dynamics of a generic general aviation aircraft. The elastic aircraft equations of motion are derived using Lagrange’s
equation and the principal of virtual work. A minimum kinetic energy axis system is chosen as the body reference
axis, which results in structural equations of motion that are decoupled from the rigid body equations of motion. An
aerodynamic strip method is utilized to develop closed-form expressions for the longitudinal generalized structural
forces. The adaptive controller is designed using a model reference adaptive control scheme, modified for general
aviation to use an “E-Z fly” decoupled control architecture, which tracks vertical flight path angle and true airspeed.
The adaptive control signal is computed using a weighted least mean square optimization, which gives the control
designer more influence on the behavior of the adaptation. A notch filter is designed to decouple the controller and
adaptation from the structural modes.
The controller is implemented in the MATLAB®/Simulink® environment, and the equations of motion are
integrated in simulation for a range of structural flexibility and plant failures. Results show that the controller is
capable of handling the uncertainties associated with unmodeled aeroelastic modes. Additionally, the controller
shows resilience to “A” and “B” matrix failures, such as 25% loss in elevator and throttle effectiveness. Actuator
speed is found to limit the amount of failure the system can recover from, where a fast actuator facilitates adaptation
to much larger failures. The notch filter is shown to be successful at decoupling the controller from the structural
modes, even for a highly flexible aircraft. Performance without the notch filter is not degraded when the structural
modes are outside the controller bandwidth; however, when structural modes fall within the controller bandwidth,
the notch filter is required to damp excessive control activity. The proposed controller shows balance between good
tracking performance and time delay margin, which is a measure of robustness in the system. This is attributed to the
weighted least mean square optimization procedure that gives the control designer more influence over the behavior
of the adaption.Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Aerospace Engineering.
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Copley-Woods, Djuna S. (Djuna Sunlight) 1977. "Aircraft interior acoustic noise control." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/9330.
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Thesis (S.B.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering, 1999.Includes bibliographical references (p. 45).
An experimental study was perfonned to determine which materials are best suited for internal aircraft noise reduction. An impedance tube with dimensions of a scaled aircraft was constructed and evaluated, and eleven materials were tested and compared based on their noise reduction properties, weight, and thickness. Polyvinylidene Fluoride was tested for use in active noise control for a large space.
by Djuna S. Copley-Woods.
S.B.
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Liceaga-Castro, Jesus U. "Helicopter flight control by individual channel design." Thesis, University of Glasgow, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.247303.
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Worsfold, M. "The control of corrosion on ageing aircraft." Thesis, Cranfield University, 1996. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.309724.
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Von, Klemperer Nicholas. "Dual-axis tilting quadrotor aircraft: Dynamic modelling and control of dual-axis tilting quadrotor aircraft." Master's thesis, Faculty of Engineering and the Built Environment, 2018. http://hdl.handle.net/11427/30156.
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This dissertation aims to apply non-zero attitude and position setpoint tracking to a quadrotor aircraft, achieved by solving the problem of a quadrotor’s inherent underactuation. The introduction of extra actuation aims to mechanically accommodate for stable tracking of non-zero state trajectories. The requirement of the project is to design, model, simulate and control a novel quadrotor platform which can articulate all six degrees of rotational and translational freedom (6-DOF) by redirecting and vectoring each propeller’s individually produced thrust. Considering the extended articulation, the proposal is to add an additional two axes (degrees) of actuation to each propeller on a traditional quadrotor frame. Each lift propeller can be independently pitched or rolled relative to the body frame. Such an adaptation, to what is an otherwise well understood aircraft, produces an over-actuated control problem. Being first and foremost a control engineering project, the focus of this work is plant model identification and control solution of the proposed aircraft design. A higher-level setpoint tracking control loop designs a generalized plant input (net forces and torques) to act on the vehicle. An allocation rule then distributes that virtual input in solving for explicit actuator servo positions and rotational propeller speeds. The dissertation is structured as follows: First a schedule of relevant existing works is reviewed in Ch:1 following an introduction to the project. Thereafter the prototype’s design is detailed in Ch:2, however only the final outcome of the design stage is presented. Following that, kinematics associated with generalized rigid body motion are derived in Ch:3 and subsequently expanded to incorporate any aerodynamic and multibody nonlinearities which may arise as a result of the aircraft’s configuration (changes). Higher-level state tracking control design is applied in Ch:4 whilst lower-level control allocation rules are then proposed in Ch:5. Next, a comprehensive simulation is constructed in Ch:6, based on the plant dynamics derived in order to test and compare the proposed controller techniques. Finally a conclusion on the design(s) proposed and results achieved is presented in Ch:7. Throughout the research, physical tests and simulations are used to corroborate proposed models or theorems. It was decided to omit flight tests of the platform due to time constraints, those aspects of the project remain open to further investigation. The subsequent embedded systems design stemming from the proposed control plant is outlined in the latter of Ch:2, Sec:2.4. Such implementations are not investigated here but design proposals are suggested. The primary outcome of the investigation is ascertaining the practicality and feasibility of such a design, most importantly whether or not the complexity of the mechanical design is an acceptable compromise for the additional degrees of control actuation introduced. Control derivations and the prototype design presented here are by no means optimal nor the most exhaustive solutions, focus is placed on the whole system and not just a single aspect of it.
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Chun, Louis Hoo Loung. "The design and test rig evaluation of advanced control laws for primary flight control actuators." Thesis, University of Bristol, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.333962.
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Kellett, Martin Gerard. "Scheduled multivariable control of battlefield helicopters." Thesis, Cranfield University, 1993. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.333663.
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Daoud, Younis Sharif. "Control and optimization of aircraft trajectories." Thesis, University of Hertfordshire, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.303524.
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Farah, Hassan. "The fuzzy logic control of aircraft." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1999. http://www.collectionscanada.ca/obj/s4/f2/dsk1/tape8/PQDD_0003/MQ43339.pdf.
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Ashworth, Anthony Ian. "Active control of V/STOL aircraft." Thesis, Lancaster University, 1995. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.296678.
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Farah, Hassan (Hassan Kahiye) Carleton University Dissertation Engineering Mechanical and Aerospace. "The Fuzzy logic control of aircraft." Ottawa, 1999.
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Öhrn, Philip, and Markus Åstrand. "Direct Lift Control of Fighter Aircraft." Thesis, Linköpings universitet, Reglerteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-157464.
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Direct lift control for aircraft has been around in the aeronautical industry for decades but is mainly used in commercial aircraft with dedicated direct lift control surfaces. The focus of this thesis is to investigate if direct lift control is feasible for a fighter aircraft, similar to Saab JAS 39 Gripen, without dedicated control surfaces. The modelled system is an aircraft that is inherently unstable and contains nonlinearities both in its aerodynamics and in the form of limited control surface deflection and deflection rates. The dynamics of the aircraft are linearised around a flight case representative of a landing scenario. Direct lift control is then applied to give a more immediate relation from pilot stick input to change in flight path angle while also preserving the pitch attitude. Two different control strategies, linear quadratic control and model predictive control, were chosen for the implementation. Since fighter aircraft are systems with fast dynamics it was important to limit the computational time. This constraint motivated the use of specialised methods to speed up the optimisation of the model predictive controller. Results from simulations in a nonlinear simulation environment supplied by Saab, as well as tests in high-fidelity flight simulation rigs with a pilot, proved that direct lift control is feasible for the investigated fighter aircraft. Sufficient control authority and performance when controlling the flight path angle were observed. Both developed controllers have their own advantages and which strategy is the most suitable depends on what the user prioritises. Pilot workload during landing as well as precision at touch down were deemed similar to conventional control.
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O'Sullivan, Donald Quinn 1970. "Aircraft interior structural-acoustic control design." Thesis, Massachusetts Institute of Technology, 1998. http://hdl.handle.net/1721.1/9888.
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Thesis (S.M.)--Massachusetts Institute of Technology, Dept. of Mechanical Engineering; and, (S.M.)--Massachusetts Institute of Technology, Technology and Policy Program, 1998.Includes bibliographical references (p. 177-184).
by Donald Quinn O'Sullivan.
S.M.
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Go, Tiauw Hiong. "Aircraft wing rock dynamics and control." Thesis, Massachusetts Institute of Technology, 1999. http://hdl.handle.net/1721.1/50081.
Full textAbstract:
Thesis (Sc.D.)--Massachusetts Institute of Technology, Dept. of Aeronautics and Astronautics, 1999.Includes bibliographical references (p. 232-236).
The dynamics of wing rock on rigid aircraft having single, two, and three rotational degrees-of-freedom are analyzed. For the purpose of the analysis, nonlinear mathematical models of the aircraft are developed. The aerodynamic expressions contained in the models can be built by fitting the appropriate aerodynamic data into the model. The dynamic analysis is performed analytically using a technique combining the Multiple Time Scales method, Center Manifold Reduction principle, and bifurcation theory. The technique yields solutions in parametric forms and leads to the separation of fast and slow dynamics, and a great insight into the system behavior. Further, a unified framework for the investigation of wing rock dynamics and control of aircraft is developed. Good agreement between the analytical results and the numerical simulations is demonstrated. Based on the results of the dynamic analysis, appropriate control strategies for the wing rock alleviation are developed. The control power limitation of the conventional aerodynamics control surfaces is considered and its effects on the alleviation of wing rock are investigated. Finally, the potential use of advanced controls to overcome the conventional controls limitation is discussed.
by Tiauw Hiong Go.
Sc.D.
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Shwan, Kurdi Mir. "Nonlinear Attitude Control ofa Generic Aircraft." Thesis, KTH, Flygdynamik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-261696.
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Determining suitable controllers for the process of evaluating dynamic per-formance of multiple versions of an aircraft’s aerodynamical, geometric and propulsive properties in its conceptual stage is an expensive task.In this report a proposition is made to utilize a generalized feedback lin-earizing controller that o˙ers the aircraft designer valuable insight into the manoeuvre performance of their aircraft. This is carried out by first estab-lishing fundamental requirements for a controller capable of treating a generic airframe, and formulating the resulting control laws.It is shown in this report, that with a suÿciently simple aerodynamic and propulsive model explicit feedback linearization is possible with satisfactory performance and robustness. Whereas it would be necessary to implement INDI if explicit inverse mappings are not obtainable. Which in turn would introduce additional tuning parameters.Robustness verification is performed in two stages, firstly by introducing a high model uncertainty within the flight control system and showing, via simulation, that the control system successfully performs desired multi-axial manoeuvres whilst managing to maintain the induced side slip below 0.1◦. Secondly by disturbing the aircraft with a discrete side slip. Critical side slip disturbance angle was found to be considerably larger than that for regular aircraft entailing that the used case study may be somewhat over dimensioned with respect to yaw control authority.
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Seigler, Thomas Michael. "Dynamics and Control of Morphing Aircraft." Diss., Virginia Tech, 2005. http://hdl.handle.net/10919/28681.
Full textAbstract:
The following work is directed towards an evaluation of aircraft that undergo structural shape change for the purpose of optimized flight and maneuvering control authority. Dynamical equations are derived for a morphing aircraft based on two primary representations; a general non-rigid model and a multi-rigid-body. A simplified model is then proposed by considering the altering structural portions to be composed of a small number of mass particles. The equations are then extended to consider atmospheric flight representations where the longitudinal and lateral equations are derived. Two aspects of morphing control are considered. The first is a regulation problem in which it is desired to maintain stability in the presence of large changes in both aerodynamic and inertial properties. From a baseline aircraft model various wing planform designs were constructed using Datcom to determine the required aerodynamic contributions. Based on nonlinear numerical evaluations adequate stabilization control was demonstrated using a robust linear control design. In maneuvering, divergent characteristics were observed at high structural transition rates. The second aspect considered is the use of structural changes for improved flight performance. A variable span aircraft is then considered in which asymmetric wing extension is used to effect the rolling moment. An evaluation of the variable span aircraft is performed in the context of bank-to-turn guidance in which an input-output control law is implemented.Ph. D.
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Tuzcu, Ilhan. "Dynamics and Control of Flexible Aircraft." Diss., Virginia Tech, 2001. http://hdl.handle.net/10919/25958.
Full textAbstract:
This dissertation integrates in a single mathematical formulation the disciplines pertinent to the
flight of flexible aircraft, namely, analytical dynamics, structural dynamics, aerodynamics
and controls. The unified formulation is based on fundamental principles and incorporates
in a natural manner both rigid body motions of the aircraft as a whole and elastic deformations
of the flexible components (fuselage, wing and empennage), as well as the aerodynamic,
propulsion, gravity and control forces. The aircraft motion is described in terms of three
translations (forward motion, sideslip and plunge) and three rotations (roll, pitch and yaw)
of a reference frame attached to the undeformed fuselage, and acting as aircraft body axes,
and elastic displacements of each of the flexible components relative to corresponding body
axes. The mathematical formulation consists of six ordinary differential equations for the
rigid body motions and one set of ordinary differential equations for each elastic displacement.
A perturbation approach permits division of the problem into a nonlinear "zero-order Problem"
for the rigid body motions, corresponding to flight dynamics, and a linear "first-order
problem" for the elastic deformations and perturbations in the rigid body translations and
rotations, corresponding to "extended aeroelasticity." Due to computational speed advantages,
the aerodynamic forces are derived by means of strip theory. The control forces for the flight
dynamics problem are obtained by an "inverse" process. On the other hand, the feedback control
forces for the extended aeroelasticity problem are derived by means of LQG theory. A numerical
example corresponding to steady level flight and steady level turn maneuver is included.Ph. D.
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Kumar, Abhishek. "Convex Modeling Techniques for Aircraft Control." Thesis, Virginia Tech, 2000. http://hdl.handle.net/10919/33530.
Full textAbstract:
The need to design a controller that self-schedules itself during the flight of an aircraft has been an active area of research. New methods have been developed beyond the traditional gain-scheduling approach. One such design method leads to a linear parameter varying (LPV) controller that changes based on the real-time variation of system dynamics. Before such a controller can be designed, the system has to also be represented as an LPV system. The current effort proposes a LPV modeling technique that is inspired by an affine LPV modeling techniques found in recent research. The properties of the proposed modeling method are investigated and compared to the affine modeling technique. It is shown that the proposed modeling technique represents the actual system behavior more closely than the existing affine modeling technique.
To study the effect of the two LPV modeling techniques on controller design, a linear quadratic regulator (LQR) controller using linear matrix inequality (LMI) formulation is designed. This control design method provides a measure of conservatism that is used to compare the controllers based on the different modeling techniques. An F-16 short-period model is used to implement the modeling techniques and design the controllers. It was found that the controller based on the proposed LPV modeling method is less conservative than the controller based on the existing LPV method. Interesting features of LMI formulation for multiple plant models were also discovered during the exercise.
A stability robustness analysis was also conducted as an additional comparison of the performance of the controllers designed using the two modeling methods. A scalar measure, called the probability of instability, is used as a measure of robustness. It was found that the controller based on the proposed modeling technique has the necessary robustness properties even though it is less conservative than the controller designed based on the existing modeling approach.Master of Science
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Sun, Xiao-Dong. "Active control of aircraft using spoilers." Thesis, Imperial College London, 1993. http://hdl.handle.net/10044/1/11324.
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Simpson, Mark N. "The application of semi-active control technology to aircraft landing gear." Thesis, Loughborough University, 1988. https://dspace.lboro.ac.uk/2134/6764.
Full textAbstract:
The purpose of the research investigation was to study the application of semi-active control technology to the design of a suspension system to be used in a landing gear of a high speed military aircraft. A semi-active system was used because it will allow a system to be driven from the hydraulic systems already existing in the aircraft without extensive modification. The research work involved establishing a theoretical mathematical model for the semi-active damping system. This model involved a large number of non-linear dynamic phenomena and elements including a two-stage gas spring, lever geometry, break out friction, square law damping and the switching function needed to achieve the semi-active control. Validation of the model was carried out by means of an extensive study of the dynamic responses obtained from digital simulation. An extended programme of laboratory experiments was also carried out to confirm the theoretical and simulated results, and to demonstrate the potential benefits in performance which can be achieved with those obtained from standard and optimized passive suspension system. The experimental rig involved a physical model which used hydraulic elements of a general industry standard, but not specially approved for aircraft use. The apparatus was arranged to permit a considerable degree of freedom for implementing the control laws which facilitated the assessment of different control schemes and allowed, at the same time, the ready simulation of various passive damping arrangements. An extensive series of trials was carried out on the final design and involved frequency response tests and subjecting the experimental suspension to inputs obtained from a simulated runway profile. The profile simulation was a discrete representation of a particular runway chosen for its roughness which was characteristic of runways from which high-speed military aircraft operate. From the research investigation and these trials it was established that semi-active control of the damping function is superior to standard techniques and achieves a substantial reduction in the energy transmitted to the airframe during ground manoeuvres.
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Aslin, P. P. "Aircraft simulation and robust flight control system design." Thesis, University of York, 1985. http://etheses.whiterose.ac.uk/9821/.
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Elramlawy, Abdelbaset Abdelgaied. "Multivariable flight control systems for agile combat aircraft." Thesis, University of Salford, 1992. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.305206.
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MacCormac, J. K. M. "Investigations on flight trajectory optimisation and adaptive control." Thesis, University of Bath, 1994. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.238734.
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Sangwian, Sirirat. "Multivariable Sliding Mode Control for Aircraft Engines." Cleveland State University / OhioLINK, 2011. http://rave.ohiolink.edu/etdc/view?acc_num=csu1315587541.
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Ebel, Kathryn C. "Adaptive Sliding Mode Control for Aircraft Engines." Cleveland State University / OhioLINK, 2011. http://rave.ohiolink.edu/etdc/view?acc_num=csu1323882562.
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Aslam-Mir, Shahzad. "Reconfigurable flight control systems for a generic fighter aircraft." Thesis, University of Southampton, 1992. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.357503.
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Broadley, Jonathan I. "The control of trailing edge separation on highly swept wings using vortex generators." Thesis, Cranfield University, 1998. http://hdl.handle.net/1826/3966.
Full textAbstract:
The results from a series of low speed wind tunnel tests on two half model highly swept
wings (a symmetrical aerofoil section and a highly cambered aerofail section) are
presented in order to examine the trailing edge flow separation mechanism and its
development with wing sweep between 30' and 60'. The tests involved surface oil flow
visualisation, smoke flow visualisation, surface static pressure and force balance
measurements at streamwise chord Reynolds numbers from 1.5 x 105 to 5.2 x 106 and
Mach number from 0.09 to 0.17. These results are used to assess two viscous-inviscid
interaction CFD methods (BVGK and VFP) and two boundary layer methods (TAPERBL
and WAKELAG) used to predict the flow over the highly cambered wing.
A parametric study using cropped delta vane vortex generators in a co-rotating array was
conducted on the 40' swept wing to investigate the effect of vane chordwise position,
vane orientation, vane height relative to the boundary layer thickness and vane spacing on
the prevention of the trailing edge separation. The performance of these flow control
devices is assessed in terms of changes in; the wing surface flowfield, lift curve slope and
the lift-dependant drag factor. In addition comparisons are made between the clean wing
and flow control wing measured pressure distributions.
The results and analysis show that the performance of the vortex generators is improved
when the height of the vortex generator is approximately equal to that of the local
boundary layer thickness and when the vane angular deflection to the local upstream flow
direction is between 14' and 21'. The performance is also seen to depend on the vanes
position ahead of separation and on the adverse pressure gradient to be restored and may
also depend on a vane spacing made non-dimensional on the wing normal chord rather
than the vane height.
Similar performance improvements are observed with the wing swept to 50' using the
positioning guidelines from this optimisation study. The performance of concave slats,
canted cropped delta vanes, 'bent'wires and sub-boundary layer wires as vortex generating
devices are seen to be not as effective as upright cropped delta vane vortex generators.
To assist in the interpretation of the parametric vortex generator study a low speed wind
tunnel technique is developed using shear stress sensitive liquid crystals to investigate the
downstream development of vortices from cropped delta vane vortex generators. The
results show that --
i) submerged vortices have less influence on the surface flow with downstream
distance than vortices closer to the edge of the boundary layer, and
ii) the primary increase in skin ffiction arises in the flow adjacent to the upflow
side of the vortex. This area increases with vortex size.
The results from this research programme are finally shown to be applicable in two market
areas. The first is as a performance improvement on current highly swept winged military
aircraft and the second is as flight controls on future aircraft from making the vortex
generating devices active. The possible customers in these two areas are identified and
marketing strategies developed for each case.
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Beeton, Wiaan. "Fault tolerant flight control of a UAV with asymmetric damage to its primary lifting surface." Stellenbosch : Stellenbosch University, 2013. http://hdl.handle.net/10019.1/85625.
Full textAbstract:
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.
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Bartak, John R. "Mitigating the MANPADS threat : International Agency, U.S., and Russian efforts /." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2005. http://library.nps.navy.mil/uhtbin/hyperion/05Mar%5FBartak.pdf.
Full textAbstract:
Thesis (M.A. in National Security Affairs)--Naval Postgraduate School, March 2005.Thesis Advisor(s): Mikhail Tsypkin, Edward J. Laurance. Includes bibliographical references (p. 73-79). Also available online.
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Jones, Natasha Ruth. "Designing shock control bumps for transonic commercial aircraft." Thesis, University of Cambridge, 2017. https://www.repository.cam.ac.uk/handle/1810/275993.
Full textAbstract:
Shock control bumps (SCBs) are considered promising flow control devices for transonic commercial aircraft. By generating a λ-shock structure, 2D SCBs offer large drag savings, but perform poorly when that structure breaks down off-design. Milder-performing 3D devices produce weak vortices, that may offer some boundary layer control, and SCBs also affect buffet via direct impact on shock motions and separation. To date however, design studies have largely ignored complications from the swept, spanwise-varying flows, so this thesis tackles the question of whether SCB arrays can offer useful benefits to the performance of transonic commercial aircraft. Using a numerical infinite-wing model, a simple rotation adaptation is shown to redress deficient on-design drag performance of 3D SCBs under swept flows. With the correct rotation (dependent on height, planform and spacing) bumps follow performance-design trends similar to those in unswept flow. With this knowledge, an array design method is developed to tailor 2D and 3D devices to local flow conditions on an aircraft model, aiming to maximise on-design drag performance. Careful infinite-wing setup means the influence of rotation and array height on performance is replicated on the aircraft. Predicted array designs achieve 74-87% of their estimated local drag savings. However, with wave drag being a smaller percentage of the total, the influence of arrays on lift is more significant and makes the optimal designs shorter than predicted. Strategies for improving off-design drag performance are then evaluated. Stagger, an alternating chordwise translation applied to 3D arrays, broadens operating range and lowers drag penalties by better accommodating off-design shock movements, but offers a less favourable trade-off against on-design drag than simply reducing the array height. However, a 2D array can always outperform a 3D on drag objectives. Lastly, buffet performance is inferred using steady indicators based on trailing edge pressure and shock location. These disagree regarding the impact on buffet onset, unresolvably due to a lack of validation data, but agree that arrays could alleviate flow development post-onset. Optimal array designs depend on prioritised objectives: considering buffet severity and on-design drag, tall 2D (or 3D) arrays; for buffet and minimum off-design drag penalties (similar to the motivation behind vortex generator application), 3D arrays of varying height and stagger. A simple flight fuel consumption model utilising the computed drag data shows that many arrays are neutral or offer small savings (up to 0.3%) across a range of mission profiles. While likely too small to merit application for solely drag purposes, this implies buffet benefits without cost to efficiency. Unsteady tests and proper assessment of buffet onset are needed to confirm this.
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Hyde, Richard Alden. "The application of robust control to VSTOL aircraft." Thesis, University of Cambridge, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.333330.
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Gittens, Simon Nevis. "Microprocessor-based digital flight control system design for an R.P.V." Thesis, University of Bath, 1985. https://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.304272.
Full textAbstract:
The development of a microprocessor based digital flight control system for a particular R.P.V. is described. The tasks required of this system are defined, and thereafter, the hardware circuits and the software structure necessary to implement a prototype are presented. The autopilot control laws are inferred from z-plane root loci, and then confirmed using digital simulations of the de-coupled roll and pitch attitude loops. The problems of the finite wordlength implementation of the control laws are discussed, and then both hybrid simulation and actual flight results are used to prove the performance of the prototype. To exploit the adaptive capabilities of a software based system, a sliding mode variable structure control law is developed for the roll attitude loop. Digital simulations are used to show that significant improvements in sensitivity reduction can be achieved under some conditions. These improvements are lost if a realistic servo-actuator model is employed. Another objective, namely the reduction of the disturbance error induced by trim imbalance, is maintained provided a reduced order switching function is used.
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Simon, Daniel. "Fighter Aircraft Maneuver Limiting Using MPC : Theory and Application." Doctoral thesis, Linköpings universitet, Reglerteknik, 2017. http://urn.kb.se/resolve?urn=urn:nbn:se:liu:diva-139945.
Full textAbstract:
Flight control design for modern fighter aircraft is a challenging task. Aircraft are dynamical systems, which naturally contain a variety of constraints and nonlinearities such as, e.g., maximum permissible load factor, angle of attack and control surface deflections. Taking these limitations into account in the design of control systems is becoming increasingly important as the performance and complexity of the aircraft is constantly increasing. The aeronautical industry has traditionally applied feedforward, anti-windup or similar techniques and different ad hoc engineering solutions to handle constraints on the aircraft. However these approaches often rely on engineering experience and insight rather than a theoretical foundation, and can often require a tremendous amount of time to tune. In this thesis we investigate model predictive control as an alternative design tool to handle the constraints that arises in the flight control design. We derive a simple reference tracking MPC algorithm for linear systems that build on the dual mode formulation with guaranteed stability and low complexity suitable for implementation in real time safety critical systems. To reduce the computational burden of nonlinear model predictive control we propose a method to handle the nonlinear constraints, using a set of dynamically generated local inner polytopic approximations. The main benefit of the proposed method is that while computationally cheap it still can guarantee recursive feasibility and convergence. An alternative to deriving MPC algorithms with guaranteed stability properties is to analyze the closed loop stability, post design. Here we focus on deriving a tool based on Mixed Integer Linear Programming for analysis of the closed loop stability and robust stability of linear systems controlled with MPC controllers. To test the performance of model predictive control for a real world example we design and implement a standard MPC controller in the development simulator for the JAS 39 Gripen aircraft at Saab Aeronautics. This part of the thesis focuses on practical and tuning aspects of designing MPC controllers for fighter aircraft. Finally we have compared the MPC design with an alternative approach to maneuver limiting using a command governor.
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Stenfelt, Gloria. "Aerodynamics and lateral control of tailless aircraft." Doctoral thesis, KTH, Flygdynamik, 2012. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-91407.
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Zaveri, Viral Shailesh. "H2 control of singularly perturbed aircraft system." Thesis, Wichita State University, 2011. http://hdl.handle.net/10057/5025.
Full textAbstract:
The objective of this research is to develop an analytical approach to control two-time-scale systems operating under certain noise parameters. This approach addresses two important design criteria: augmentation of large-scale system with disturbance model and its two-time-scale representation, and order reduction of the large-scale systems for reduced controller design complexity. The problem of large-scale system with Gaussian noises is solved as the stochastic system implementing linear-quadratic Gaussian control. Order reduction method uses singular perturbation techniques for the simplicity of control algorithms. Control law design process for a singularly perturbed stochastic system includes implementation and comparative analysis of optimal, composite, and reduced controller techniques. Practical model, longitudinal dynamics of digital fly-by-wire F-8C fighter aircraft, illustrates the validation of the proposed concepts.Thesis (M.S.)--Wichita State University, College of Engineering, Dept. of Electrical Engineering and Computer Science.
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Roughen, Kevin Michael. "Active aeroelastic control of supersonic transport aircraft." Diss., Restricted to subscribing institutions, 2009. http://proquest.umi.com/pqdweb?did=1998391981&sid=1&Fmt=2&clientId=1564&RQT=309&VName=PQD.
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Cetek, Cem. "Aircraft control with nonlinear indicial response model." Ohio : Ohio University, 1999. http://www.ohiolink.edu/etd/view.cgi?ohiou1175888353.
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