The Multiple Cooling Multi-Objective Simulated Annealing Algorithm is used for the conceptual design optimization of a supersonic Unmanned Combat Aerial Vehicle (UCAV). Single and multiobjective optimization problems are addressed while limiting performance requirements between desired bounds to obtain viable aircraft configurations. A conceptual aircraft design code was prepared for planned but flexible combat missions. The results demonstrate that the optimization technique employed is an effective tool for the conceptual design of aircrafts.

In this thesis, adaptive neural network controllers are designed for a high subsonic cruise missile. Two autopilot designs are included in the study using adaptive neural networks, namely an altitude hold autopilot designed for the longitudinal channel and a directional autopilot designed for heading control. Aerodynamic coefficients are obtained using missile geometry / a 5-Degree of Freedom (5-DOF) simulation model is obtained, and linearized at a single trim condition. An inverted model is used in the controller. Adaptive Neural Network (ANN) controllers namely, model inversion controllers with Sigma-Pi Neural Network, Single Hidden Layer Neural Network and Background Learning implemented Single Hidden Layer Neural Network, are deployed to cancel the modeling error and are applied for the longitudinal and directional channels of the missile. This approach simplifies the autopilot designing process by combining a controller with model inversion designed for a single flight condition with an on-line learning neural network to account for errors that are caused due to the approximate inversion. Simulations are performed both in the longitudinal and directional channels in order to demonstrate the effectiveness of the implemented control algorithms. The advantages and drawbacks of the implemented neural network based controllers are indicated.

Spray is an efficient tool in the usage whose primary objectives are to obtain droplets with increased liquid surface area and more dispersed liquid over a larger volume. The determination of spray characteristics has been a topic of extensive research recently. In the present investigation, the flow structure of a spray issuing from an oil burner nozzle was determined in a parametrical manner. The main tool in the experimental research is the Interferometric Particle Imaging (IPI) configuration. This method exploits the interference between light reflected from and refracted through individual transparent spray droplets which are illuminated by a laser light sheet in a wide angle forward-scatter region. Based on a scattering theory, the droplet diameter of spray particles can be related to the light pattern scattered from that particle. In addition, using double-framed images also enables the calculation of velocities associated with these particles. In this way, as a representation of spray structure, the droplet size and velocity distributions were obtained prior to a change in the primary parameters of liquid flow e.g. surface tension, viscosity, density and the injection pressure. The evolution of spray characteristics in space were also examined by conducting measurements in different radial and axial locations relative to spray centerline.

This thesis addresses the attitude control of satellites with similar and dissimilar actuators and control allocation methods on maneuvering. In addition, the control moment gyro (CMG) steering with gyroscopes having limited gimbal angle travel is also addressed. Full Momentum envelopes for a cluster of four CMG&#039 / s are obtained in a pyramid type mounting arrangement. The envelopes when gimbal travel is limited to plus-minus 90 degree are also obtained. The steering simulations using Moore Penrose (MP) pseudo inverse as well as blended inverse are presented and success of the pre planned blended inverse steering in avoiding gimbal angle limits is demonstrated through satellite slew maneuver simulations, showing the completion of the maneuver without violating gimbal angle travel restrictions. Dissimilar actuators, CMG and magnetic torquers are used as an approach of overactuated system. Steering simulations are carried out using different steering laws for constant torque and desired satellite slew maneuver scenarios. Success of the blended inverse steering algorithm over MP pseudo inverse is also demonstrated

This thesis describes the automatic flight control systems designed for a conventional and an over actuated unmanned air vehicle (UAV). A nonlinear simulation model including the flight mechanics equations together with the interpolated nonlinear aerodynamics, environmental effects, mass-inertia properties, thrust calculations and actuator dynamics is created / trim and linearization codes are developed. Automatic flight control system of the conventional UAV is designed by using both classical and robust control methods. Performances of the designs for full autonomous flight are tested through nonlinear simulations for different maneuvers in the presence of uncertainties and disturbances in the aircraft model. The fault tolerant control of an over actuated UAV is the main concern of the thesis. The flight control system is designed using classical control techniques. Two static control allocation methods are examined: Moore-Penrose pseudo inverse and blended inverse. For this purpose, an aircraft with three sets of ailerons is employed. It is shown that with redundant control surfaces, fault tolerant control is possible. Although both of the static control allocation methods are found to be quite successful to realize the maneuvers, the new blended inverse algorithm is shown to be more effective in controlling the aircraft when some of the control surfaces are lost. It is also demonstrated that, with redundant control surfaces it is possible to recover the aircraft during a maneuver even some of the control surfaces are damaged or got stuck at a particular deflection.

In this thesis, the performance of piezoelectric patches on investigation of vibration characteristics and active vibration control of a smart beam is presented. The smart beam is composed of eight surface-bonded piezoelectric patches symmetrically located on each side of a cantilever aluminium beam. At first, vibration characteristics of the smart beam is investigated by employment of piezoelectric patches as sensors and actuators. Smart beam is excited by either impact hammer or piezoelectric patch and the response of the smart beam particular to these excitations is measured by piezoelectric patches used as sensors. In order to investigate the performance of piezoelectric patches in sensing, the measurements are also conducted by commercially available sensing devices. Secondly, active vibration suppression of the smart beam via piezoelectric sensor/actuator pair is considered. For this purpose, system identification of the smart beam is conducted by using four piezoelectric patches as actuators and another piezoelectric patch as a sensor. The designed robust controller is experimentally implemented and active vibration suppression of the free and first resonance forced vibration is presented. Thirdly, active vibration control of the smart beam is studied by employment of piezoelectric patches as self-sensing actuators. Following the same approach used in the piezoelectric sensor/actuator pair case, system identification is conducted via self-sensing piezoelectric actuators and robust controller is designed for active vibration suppression of the smart beam. Finally, active vibration suppression via self-sensing piezoelectric actuators is experimentally presented.

This thesis concerns the design and development of neural network based predictive algorithms to predict approaching aircraft limits. Therefore, approximate dynamics of flight envelope parameters such as angle of attack and load factor are constructed using neural network augmented dynamic models. Then, constructed models are used to predict steady state responses. By inverting the models and solving for critical controls at the known envelope limits, critical control inputs are calculated as well. The performance of the predictor algorithm is then evaluated with a different neural network online adaptation law which uses a stack of recorded data. It is shown that using a stack of recorded data online, constructed models become much more representative of limit parameter dynamics compared to adaptation using instantaneous measured data only. The benefits of recording data online and using it for weight adaptation are presented in the scope of dynamic trim and control limit predictions.

Airfoils &deg / apping in pitch and plunge are studied, and the &deg / apping motion parameters are op- timized to maximize thrust generation and the e&plusmn / ciency of the thrust generation. Unsteady viscous &deg / ow&macr / elds over &deg / apping airfoils are computed on overset grids using a Navier-Stokes solver. Computations are performed in parallel using Parallel Virtual Machine library routines in a computer cluster. A single &deg / apping airfoil and dual airfoils &deg / apping in a biplane con- &macr / guration are considered. A gradient based optimization algorithm is employed. The thrust production and the e&plusmn / ciency of the thrust production are optimized with respect to &deg / apping parameters / the plunging and pitching amplitudes, the &deg / apping frequency, and the phase shift between the pitch and plunge motions. It is observed that thrust generation of &deg / apping airfoils strongly depends on the phase shift and high thrust values may be obtained at the expense of reduced e&plusmn / ciency. For a high e&plusmn / ciency in thrust generation, the e&reg / ective angle of attack of the airfoil is reduced and large scale vortex formations at the leading edge are prevented. At a &macr / xed reduced &deg / apping frequency of 1, a single &deg / apping airfoil in pitch and plunge motion produces the maximum average thrust coe&plusmn / cient of 1:41 at the plunge amplitude of 1:60, the pitch amplitude of 23:5o, and the phase shift of 103:4o whereas the maximum e&plusmn / ciency of 67:5% is obtained at the plunge amplitude of 0:83, the pitch amplitude of 35:5o and the phase shift of 86:5o.

In this thesis, the effect of semi-geodesic winding on the free vibration characteristics of filament wound composite shells of revolution with variable radii of curvature is studied. The analysis is performed by a semi-analytical solution method which is based on the numerical integration of the finite exponential Fourier transform of the fundamental shell of revolution equations. The governing equations for the free vibration analysis are initially obtained in terms of fundamental shell variables, and they are reduced to a system of first order ordinary differential equations by the application of finite exponential Fourier Transform, resulting in a two point boundary value problem. The boundary value problem is then reduced to a series of initial value problems, and the multisegment numerical integration technique is used in combination with the frequency trial method in order to extract the natural frequencies and determine the mode shapes within a given range of natural frequencies. Previous studies on geodesic winding is extended such that the effect of semi-geodesic winding which rely on the preset friction between the fiber and the mandrel surface on the stiffness and vibration characteristics of filament wound shells of revolution is investigated. Additionally, finite element analysis is employed to compare the results obtained from semi-analytical model solved by numerical integration and finite element model solved by finite element method. Sample results are obtained for filament wound truncated conical and spherical shells of revolution and the effect of the winding pattern on the vibration characteristics of shells of revolution is investigated thoroughly.

In this thesis, the structural design, structural analysis and producibility analysis of an unmanned aerial vehicle wing were performed. Three different wing models, made of different materials, were designed. The wings were aluminum wing model and composite wing models / made of prepreg and wet lay-up. All wings have the same aerodynamic geometry and structural configuration under the same flight conditions. The structural designs of three wings were done by using Unigraphics NX. The finite element modeling of the wings were built by using MSC Patran package program. After the application of the loads on models, structural analyses were performed by MSC Nastran. Finally, the producibility analysis of prepreg wing model was conducted by using FiberSIM package program. The prepreg wing model was selected as optimum design with studies conducted in the study considering weight, producibility, cruise and gust stress and displacement conditions.