This thesis investigates the effects of camber change on the mission adaptive wing of a structurally designed unmanned aerial vehicle (UAV). The commercial computational fluid dynamics (CFD) software ANSYS/FLUENT is employed for the aerodynamic analyses. Several cambered airfoils are compared in terms of their aerodynamic coefficients and the effects of the camber change formed in specific sections of the wing on the spanwise pressure distribution are investigated. The mission adaptive wing is modeled structurally to observe the effect of spanwise pressure distribution on the wing structure. For the structural design and analysis of the UAV under this study, commercial software MSC/PATRAN and MSC/NASTRAN are used. The structural static and dynamic analyses of the unmanned aerial vehicle are also performed under specified flight conditions. The results of these analyses show that the designed structure is safe within the flight envelope. Having completed aero-structural design and analysis, the designed unmanned aerial vehicle is manufactured by TUSAS Aerospace Industries (TAI).

The scope of the thesis is to numerically investigate the aerodynamics of flapping airfoils in hover and forward flight. The flowfields around flapping airfoils are computed by solving the governing equations on moving and/or deforming grids. The effects of Reynolds number, reduced frequency and airfoil geometry on unsteady aerodynamics of flapping airfoils undergoing pure plunge and combined pitch-plunge motions in forward flight are investigated. It is observed that dynamic stall of the airfoil is the main mechanism of lift augmentation for both motions at all Reynolds numbers ranging from 10000 to 60000. However, the strength and duration of the leading edge vortex vary with airfoil geometry and reduced frequency. It is also observed that more favorable force characteristics are achieved at higher reduced frequencies and low plunging amplitudes while keeping the Strouhal number constant. The computed flowfields are compared with the wide range of experimental studies and high fidelity simulations thus it is concluded that the present approach is applicable for investigating the flapping wing aerodynamics in forward flight. The effects of vertical translation amplitude and Reynolds number on flapping airfoils in hover are also studied. As the vertical translation amplitude increases, the vortices become stronger and the formation of leading edge vortex is pushed towards the midstroke of the motion. The instantaneous aerodynamic forces for a given figure-of-eight motion do not alter significantly for Reynolds numbers ranging from 500 to 5500.

The aim of this study is to describe the conceptual design, performance analysis to validate the design and manufacturing steps of Middle East Technical University Tactical Unmanned Air Vehicle (METU TUAV). The system requirements are adopted from a market study and assumed as is. Utilizing competitor search and conceptual design methodology, the rough parameters of the aircraft are defined and a performance analysis is conducted to validate the requirements. After the design team is content that the design is meeting the requirements, material and production techniques are evaluated. The male and female molds of the aircraft are manufactured with glass fibre fabric and special mold resin. Using the female molds / with glass, carbon and aramid fibre materials and epoxy matrix / utilizing wet-layup and vacuum bagging techniques the METU TUAV is manufactured. Wing, tail and fuselage skins are manufactured first and the reinforcing structures are integrated and cured inside the skins. Then the skins are assembled and the separate components are obtained. The rear landing gear and tail booms are also manufactured from carbon fibre composites. The individual parts are assembled together in special alignment jigs and the METU TUAV is completed.

ABSTRACT A PARAMETRIC INVESTIGATION OF TIP INJECTION FOR ACTIVE TIP VORTEX CONTROL Dedekarginoglu, Riza Can M.Sc., Department of Aerospace Engineering Supervisor :Asst. Prof. Dr. Oguz Uzol December 2010, 79 pages Wing tip vortex is a challenging phenomenon that reduces the lift generation at the tip region of the wing. For aerial vehicles, several methodologies were presented for the sake of controlling vortices and alleviating effects of tip loss. In this study, the effect of wing tip injection on wing tip vortex structure was investigated computationally. A NACA0015 profile rectangular wing was employed with an aspect ratio of 3, at a free stream Reynolds number of 67000. 10 identical ejection holes along the wing were prepared chordwise to provide cross sectional air flow in order to determine the net effect of ejection over wing tip vortices and wake flow field. Study setup consists of a wind tunnel that is 1.6m long, 0.6m wide and 0.6m high, which the wing is attached to one side of it as a cantilever beam. Chord length of the wing is 0.1m and span is 0.3m. A constant free stream air flow is maintained with 10 m/s of velocity. Computer aided drawing (CAD) and grid generation were carried out using commercial tools. Whole setup was drawn using Rhinoceros. Surface mesh was created using ANSYS Gambit, ANSYS T-Grid software was used for generating the viscous mesh over the wing and finally for volume mesh ANSYS Gambit was utilized once more. FLUENT was chosen to be the flow solution tool with k-&omega / SST turbulence model. For 3 different angles of attack cases, respectively, 4&deg / , 8&deg / and 12&deg / , several injection scenarios were defined. There are 3 steady injection cases for each angle of attack case namely, no injection case, uniform injection case, triangular waveform injection case where there is no injection at the leading edge tip whereas there is injection which is equal to the uniform injection velocity at the trailing edge tip. Moreover there are 5 additional scenarios for 8&deg / angle of attack case that are, sinusoidal waveform injection case which consists of a chordwise velocity distribution shape that is a quarter sinus wave where maximum injection velocity is the same as the uniform velocity, reverse triangular waveform injection case where injection velocities were reversed with respect to triangular waveform case, two cases consisting of angled injections having both +15&deg / and -15&deg / with respect to the flapping axis of the wing. The effect of tube walls on the jet injection was neglected for all cases, therefore for the last case, in order to simulate pipe flow, a case is provided with uniform injection velocity. In that way, regardless of the solution method, a parametric study was performed. Considering each case, non-dimensional 3-axis velocity components, turbulent kinetic energy, vorticity magnitude, pressure, lift and drag values were computed and having the exactly same cases as an experimental study for 8&deg / angle of attack, a comparison of aerodynamic data series was presented. As results, it&rsquo / s observed that, vortex core locations were shifted upwards and away from the tip region. Increasing the turbulence level of the tip flow by tip injection, inherently the pressure difference became larger, however as the vortices ascend, tip loss decreases. In that way, a significant increase in the lift was observed while drag values are slightly increased, as well.

This study describes the design and implementation of the altitude and heading autopilot algorithms for a fixed wing unmanned air vehicle and navigation algorithm for attitude and heading reference outputs. Algorithm development is based on the nonlinear mathematical model of Middle East Technical University Tactical Unmanned Air Vehicle (METU TUAV), which is linearized at a selected trim condition. A comparison of nonlinear and linear mathematical models is also done. Based on the linear mathematical model of the METU TUAV, the classical control methods are applied during the design process of autopilot algorithms. For the confirmation purposes of the autopilot and navigation algorithms, a nonlinear simulation environment is developed in Matlab/Simulink including nonlinear model of the METU TUAV, altitude and heading autopilot loops, nonlinear actuator models, sensor models and navigation model. In the first part of the thesis, feedback signals for the controller are provided by IMU free measurements. In the second part, the feedback signals are provided by an attitude and heading reference mode, which incorporates the gyroscope solutions with the magnetic sensor and accelerometer sensor measurements by using a Kalman filter algorithm. The performance comparison of the controller is done for both cases where the effects of having different modes of the measurement sources are investigated.

The main purpose of this study is to aerodynamically design and optimize winglet, twist angle distribution and pitch angle for a wind turbine blade using CFD to produce more power. The RANS solver of Numeca Fine/Turbo was validated by two test cases, the NREL II and NREL VI blades. The results have shown a considerable agreement with measurements for both cases. Two different preconditioners have been implemented for the low Mach number flow. The results have shown the superiority of Merkle preconditioner over Hakimi one and Merkle was selected for further simulations. In addition to that, different turbulence models have been compared and the Launder &ndash / Sharma has shown the best agreement with measurements. Launder &ndash / Sharma was chosen for further simulations and for the design process. Before starting the design and optimization, different winglet configurations were studied. The winglets pointing towards the suction side of the blade have yielded higher power output. Genetic algorithm and artificial neural network were implemented in the design and optimization process. The optimized winglet has shown an increase in power of about 9.5 % where the optimized twist has yielded to an increase of 4%. Then the stall regulated blade has been converted into pitch regulated blade to yield more power output. The final design was produced by a combination of the optimized winglet, optimized twist andbest pitch angle for every wind speed. The final design has shown an increase in power output of about 38%.

The development and testing for a UAV testbed to be used in academic research and undergraduate education is proposed in this thesis. Analysis on commercial off-the-shelf UAV systems and autopilots lead to the development of a custom, open-architecture and modular UAV testbed. The main focus is to support research in UAV control field and education of the undergraduate students. The integration and use of commercial-off-the-shelf avionics and air vehicle are described in detail. System performance is examined both in flight and on the ground. Results of the system tests show that the developed system is a functional UAV testbed to be used in research of different flight control algorithms.

Nowadays, Global Positioning System (GPS) which provide global coverage, continuous tracking capability and high accuracy has been preferred as the primary tracking system for onboard real-time precision orbit determination of Low Earth Orbiters (LEO). In this work, real-time orbit determination algorithms are established on the basis of extended Kalman, unscented Kalman, regularized particle, extended Kalman particle and extended H-infinity filters. Particularly, particle filters which have not been applied to the real time orbit determination until now are also performed in this study and H-infinity filter is presented using all kinds of real GPS observations. Additionally, performance of unscented Kalman filter using GRAPHIC (Group and Phase Ionospheric Correction) measurements is investigated. To evaluate performances of all algorithms, comparisons are carried out using different types of GPS observations concerning C/A (Coarse/Acquisition) code pseudorange, GRAPHIC and navigation solutions. A software package for real time orbit determination is developed using recursive filters mentioned above. The software is implemented and tested in MATLAB&copy / R2010 programming language environment on the basis of the object oriented programming schema.

Helicopter flight dynamic, rotor aerodynamic and dynamic analyses activities have been a great dispute since the first helicopters, at both design and test stages. Predicting rotor aerodynamic and dynamic characteristics, helicopter dynamic behavior and trimmed flight conditions is a huge challenge to engineers as it involves the tradeoff between accuracy, fidelity, complexity and computational cost. Flight dynamic activities such as / predicting trim conditions, helicopter dynamic behavior and simulation of a flight condition or maneuver mostly require analysis tools with low computational cost and complexity. However this decreases accuracy and fidelity of the model. On the other hand, analyses at design stages, such as / blade geometric and structural design mostly requires accurate and higher fidelity aerodynamic load predictions over the rotor disk. Contrarily this brings high computational cost and complexity. Therefore separate analysis tools for each objective or one complete tool that can be used for all purposes are essential. Throughout this study a helicopter mathematical including trim model with a selective and modular structure is developed as a generic analysis tool. The selective structure enables the mathematical model to be used in both flight dynamic and comprehensive analysis while the modular structure plays a role as an infrastructure for further developments. The mathematical model developed is validated with flight test data of several helicopters. Besides, commercial helicopter comprehensive analysis tools are used to validate the mathematical model analyses. Results showed good agreement with the compared data.

In this thesis analysis and design of helicopter rotor blades were discussed for reduced vibrational level. For this purpose an optimization procedure was developed which involves coupling of the comprehensive rotorcraft analysis tool CAMRAD JA and the gradient based optimization algorithm. The main goal was to achieve favorable blade structural dynamics characteristics that would lead to reduction in vibrational level. For this purpose blade stiffness and mass distributions were considered as the design variables. In order to avoid likely occurrences of unrealistic results, the analyses were subjected to constraints which were sensitive to the design variables. The optimization procedure was applied on two isolated rotor blades and a full helicopter with main rotor, tail rotor and fuselage by using natural frequency separation and hub load minimization respectively. While the former approach relied on the blade natural frequencies, the latter approach involved higher harmonic aerodynamic and blade motion calculations. For both approaches, the improvement in vibration characteristics and blade mass and stiffness distributions of the initial design and the design after optimization analyses were compared and discussed.