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.

This thesis gives a comprehensive study on the effect of using different structural idealizations on the design, analysis and optimization of thin walled semi-monocoque wing structures in the preliminary design phase. In the design part, wing structures are designed by employing two different structural idealizations that are typically used in the preliminary design phase. In the structural analysis part, finite element analysis of one of the designed wing configurations is performed using six different one and two dimensional element pairs which are typically used to model the sub-elements of semi-monocoque wing structures. The effect of using different finite element types on the analysis results of the wing structure is investigated. During the analysis study, depending on the mesh size used, conclusions are also inferred with regard to the deficiency of certain element types in handling the true external load acting on the wing structure. Finally in the optimization part, wing structure is optimized for minimum weight by using finite element models which have the same six different element pairs used in the analysis phase. The effect of using different one and two dimensional element pairs on the final optimized configurations of the wing structure is investigated, and conclusions are inferred with regard to the sensitivity of the optimized wing configurations with respect to the choice of different element types in the finite element model. Final optimized wing structure configurations are also compared with the simplified method based designs which are also optimized iteratively. Based on the results presented in the thesis, it is concluded that with the simplified methods, preliminary sizing of the wing structures can be performed with enough confidence, as long as the simplified method based designs are also optimized. Results of the simplified method of analysis showed that simplified method is applicable to be used as an analysis tool in performing the preliminary sizing of the wing structure before moving on to more refined finite element based analysis.

This study investigates the effect of chordwisely modulated tip injection on the flow and turbulence characteristics of the tip vortex through experimental measurements downstream of a rectangular half-wing that has an aspect ratio of three. This injection technique involves spanwise jets at the tip that are issued from a series of holes along the chord line normal to the freestream flow direction. The injection mass flow rate from each hole is individually controlled using computer driven solenoid valves and therefore the flow injection geometrical pattern at the tip can be adjusted to any desired waveform shape, with any proper injection velocity. The measurements are performed in a blow-down wind tunnel using Constant Temperature Anemometry and Kiel probe traverses as well as Stereoscopic Particle Image Velocimetry. Current data show consistent trends with v previously observed effects of steady uniform tip injection such as the upward and outward motion of the vortex as well as increased levels of turbulence within the vortex core. The vortex size gets bigger with injection and the total pressure levels get reduced significantly near the vortex core. The injection pattern also seems to affect the size of the wing wake as well as the wake entrainment characteristics of the tip vortex. Depending on the injection waveform pattern and injection momentum coefficient the helicoidal shape of the tip vortex also seems to get affected.

This thesis presents a finite element method based comparative study of linear and geometrically non-linear progressive failure analysis of thin walled composite aerospace structures, which are typically subjected to combined in-plane and out-of-plane loadings. Different ply and constituent based failure criteria and material property degradation schemes have been included in a PCL code to be executed in MSC Nastran. As case studies, progressive failure analyses of sample composite laminates with cut-outs under combined loading are executed to study the effect of geometric non-linearity on the first ply failure and progression of failure. Ply and constituent based failure criteria and different material property degradation schemes are also compared in terms of predicting the first ply failure and failure progression. For mode independent failure criteria, a method is proposed for the determination of separate material property degradation factors for fiber and matrix failures which are assumed to occur simultaneously. The results of the present study show that under combined out-of-plane and in-plane loading, linear analysis can significantly underestimate or overestimate the failure progression compared to geometrically non-linear analysis even at low levels of out-of-plane loading.