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Helicopters in forward flight experience complex aerodynamic phenomena to various degrees. In low speed level flight, the vortex wake remains close to the rotor disk and interacts with the rotor blades to give rise to blade vortex interaction phenomena. In high speed flight, compressibility effects dominate leading to the formation of shocks. If the required thrust is high, the combination of high collective pitch and cyclic pitch variations give rise to three-dimensional dynamic stall phenomena. Maneuvers further exacerbate the unsteady airloads and affect rotor and hub design. The strength and durability of the rotor blades and hub components is dependent on accurate estimates of peak-to-peak structural loads. Accurate knowledge of control loads is important for sizing the expensive swash-plate components and assuring long fatigue life. Over the last two decades, computational tools have been developed for modeling rotorcraft aeromechanics. In spite of this progress, loads prediction in unsteady maneuvers which is critical for peak design loads continues to be a challenging task. The primary goal of this research effort is to investigate important physical phenomena that cause severe loads on the rotor in steady flight and in extreme maneuvers. The present work utilizes a hybrid Navier-Stokes/free-wake CFD methodology coupled to a finite element based multi-body dynamics analysis to systematically study steady level and maneuvering flight conditions. Computational results are presented for the UH-60A rotor for a parametric sweep of speed and thrust conditions and correlated with test data at the NFAC Wind Tunnel. Good agreement with test data has been achieved using the current methodology for trim settings and integrated hub loads, torque, and power. Two severe diving turn maneuvers for the UH-60A recorded in the NASA/Army Airloads Flight Tests Database have also been investigated. These maneuvers are characterized by high load factors and high speed flight. The helicopter experiences significant vibration during these maneuvers. Mean and peak-to-peak structural loads and extensive stall phenomena including an advancing side stall phenomena have been captured by the present analyses.

The two-noise source model for predicting jet noise claims that the radiated jet noise is composed of two distinct sources one associated with the small-scale turbulence and another associated with the large-scale turbulence. The former source is claimed to radiate noise predominantly at larger angles with respect to the downstream jet axis, whereas the large-scale turbulence radiates predominantly at the shallower angles. A key objective of this effort is to experimentally validate this model using correlation and coherence measurements. Upon the successful validation of the two-noise source model for jets exhausting from multiple nozzle geometries driven at Mach numbers ranging from subsonic to supersonic, a three-microphone signal enhancement technique is employed to separate the contribution of the small-scale turbulence from that of the large-scale turbulence in the far-field. This is the first-ever quantitative separation of the contributions of the turbulence scales in far-field jet noise measurements. Furthermore, by suitable selection of far-field microphone positions, the separation of the contribution of any internal or core noise from that of the jet-mixing noise is achieved. Using coherence-based techniques to separate the contributions of the small-scale turbulence, large-scale turbulence, and any internal or core noise from far-field exhaust noise measurements forms the backbone of this effort. In the application of coherence-based multiple-microphone signal processing techniques to separate the contributions of the small-scale turbulence, large-scale turbulence, and any internal or core noise in the far-field, research efforts focus on three techniques (1) the coherent output power spectrum using two microphones, (2) an ordinary coherence method using the three-microphone technique, and (3) the partial-coherence method using five microphones. The assumption of jet noise incoherence between correlating microphone is included in each of these methods. In light of the noise radiation mechanisms described within the framework of the two-noise source model and their spatial characteristics as experimentally determined in the far-field, the assumption of jet noise incoherence is evaluated through a series of experiments designed to study jet noise coherence across a variety of nozzle geometries and jet Mach numbers ranging from subsonic to supersonic. Guidelines for the suitable selection of far-field microphone locations are established.

This dissertation describes three aspects of the comprehensive rotorcraft analysis. First, a physics-based methodology for the modeling of hydraulic devices within multibody-based comprehensive models of rotorcraft systems is developed. This newly proposed approach can predict the fully nonlinear behavior of hydraulic devices, and pressure levels in the hydraulic chambers are coupled with the dynamic response of the system. The proposed model evaluates relevant hydraulic quantities such as chamber pressures, orifice flow rates, and pressure relief valve displacements. This model could be used to design lead-lag dampers with desirable force and damping characteristics. The second part of this research is in the area of computational aeroelasticity, in which an interface between computational fluid dynamics (CFD) and computational structural dynamics (CSD) is established. This interface enables data exchange between CFD and CSD with the goal of achieving accurate airloads predictions. In this work, a loose coupling approach based on the delta-airload method is developed in a finite-element method based multibody dynamics formulation, DYMORE. A loose coupling analysis between a CFD code, OVERFLOW-2, and a CSD program, DYMORE, is performed to validate this aerodynamic interface. The ability to accurately capture the wake structure around a helicopter rotor is crucial for rotorcraft performance analysis. In the third part of this thesis, a new representation of the wake vortex structure based on Non-Uniform Rational B-Spline (NURBS) curves and surfaces is proposed to develop an efficient model for prescribed and free wakes. The proposed formulation has the potential to reduce the computational cost associated with the use of the Helmholtz¡¯s law and the Biot-Savart law when calculating the induced flow field around the rotor. An efficient free wake analysis will considerably decrease the computational cost of comprehensive rotorcraft analysis, making the approach more attractive to routine use in industrial settings.

To decrease the computational cost of genetic algorithm optimizations, surrogate models are used during optimization. Online update of surrogate models and repeated exchange of surrogate models with exact model during genetic optimization converts static optimization problems to dynamic ones. However, genetic algorithms fail to converge to the global optimum in dynamic optimization problems. To address these problems, a multiploid genetic algorithm optimization method is proposed. Multi-fidelity surrogate models are assigned to corresponding levels of fitness values to sustain the static optimization problem. Low fidelity fitness values are used to decrease the computational cost. The exact/highest-fidelity model fitness value is used for converging to the global optimum. The algorithm is applied to single and multi-objective turbine blade aerodynamic optimization problems. The design objectives are selected as maximizing the adiabatic efficiency and torque so as to reduce the weight, size and the cost of the gas turbine engine. A 3-D steady Reynolds-Averaged Navier-Stokes solver is coupled with an automated unstructured grid generation tool. The solver is validated by using two well known test cases. Blade geometry is modelled by 37 design variables. Fine and coarse grid solutions are respected as high and low fidelity surrogate models, respectively. One of the test cases is selected as the baseline and is modified in the design process. The effects of input parameters on the performance of the multiploid genetic algorithm are studied. It is demonstrated that the proposed algorithm accelerates the optimization cycle while providing convergence to the global optimum for single and multi-objective problems.

Accuracy of the finite difference sensitivity calculations are improved by calculating the optimum finite difference interval sizes. In an aerodynamic inverse design algorithm, a compressor cascade geometry is perturbed by shape functions and finite differences sensitivity derivatives of the flow variables are calculated with respect to the base geometry flow variables. Sensitivity derivatives are used in an optimization code and a new airfoil is designed verifying given design characteristics. Accurate sensitivities are needed for optimization process. In order to find the optimum finite difference interval size, a method is investigated. Convergence error estimation techniques in iterative solutions and second derivative estimations are investigated to facilitate this method. For validation of the method, analytical sensitivity calculations of Euler equations are used and several applications are performed. Efficiency of the finite difference sensitivity calculations is improved by parallel computing. Finite difference sensitivity calculations are independent tasks in an inverse aerodynamic design algorithm and can be computed separately. Sensitivity calculations are performed on parallel processors and computing time is decreased.

This study presents the theoretical analysis of an active flutter suppression methodology applied to a smart fin. The smart fin consists of a cantilever aluminum plate-like structure with surface bonded piezoelectric (PZT, Lead- Zirconate-Titanate) patches. A thermal analogy method for the purpose of modeling of piezoelectric actuators in MSC&reg / /NASTRAN based on the analogy between thermal strains and piezoelectric strains was presented. The results obtained by the thermal analogy were compared with the reference results and very good agreement was observed. The unsteady aerodynamic loads acting on the structure were calculated by using a linear two-dimensional Doublet-Lattice Method available in MSC&reg / /NASTRAN. These aerodynamic loads were approximated as rational functions of the Laplace variable by using one of the aerodynamic approximation schemes, Roger&amp / #8217 / s approximation, with least-squares method. These approximated aerodynamic loads together with the structural matrices obtained by the finite element method were used to develop the aeroelastic equations of motion of the smart fin in state-space form. The Hinf robust controllers were then designed for the state-space aeroelastic model of the smart fin by considering both SISO (Single-Input Single-Output) and MIMO (Multi-Input Multi-Output) system models. The verification studies of the controllers showed satisfactory flutter suppression performance around the flutter point and a significant improvement in the flutter speed of the smart fin was also observed.

Long span bridges are susceptible to wind. Hence it is important to study their wind-induced vibrations to avoid any probable structural failures. In this thesis, the results of an experimental and computational investigation of the aerodynamic characteristics of trapezoid bridge deck cross-sections with three different aspect ratios (10, 12 and 15) and four different side angles (75&ordm / , 60 &ordm / , 45 &ordm / and 30 &ordm / ) are analyzed and presented. The flow around rigid fixed bridge deck models is investigated to obtain the relevant aerodynamic coefficients and the vortex shedding frequency and Strouhal number. Two dimensional unsteady Reynolds Averaged Navier-Stokes equations are solved using commercial CFD software at different Reynolds numbers. The numerical results are compared with the experimental data obtained by testing the model bridge decks geometries in a low speed wind tunnel. The results of this study demonstrate that the models aerodynamic parameters except their lift coefficient are almost dependent on the aspect ratio. In addition, the influence of side angle on all aerodynamic parameters has to be taken in account.

This thesis explores one possible method of preventing helicopter crashes caused by brownout using the noise generated by the helicopter rotor as an altimeter. The hypothesis under consideration is that the helicopter's height, velocity, and obstacle locations with respect to the helicopter, can be determined by comparing incident and reflected rotor noise signals, provided adequate bandwidth and signal to noise ratio. Heights can be determined by measuring the cepstrum of the reflected helicopter noise. The velocity can be determined by measuring small amounts of Doppler distortion using the Mellin-Scale Transform. Height and velocity detection algorithms are developed, optimized for this application, and tested using a microphone array. The algorithms and array are tested using a hemianechoic chamber and outside in Georgia Tech's Burger Bowl. Height and obstacle detection are determined to be feasible with the existing array. Velocity detection and surface mapping are not successfully accomplished.

The unsteady interaction between trailing edge aerodynamic flow control and airfoil motion in pitch and plunge is investigated in wind tunnel experiments using a two degree-of-freedom traverse which enables application of time-dependent external torque and forces by servo motors. The global aerodynamic forces and moments are regulated by controlling vorticity generation and accumulation near the trailing edge of the airfoil using hybrid synthetic jet actuators. The dynamic coupling between the actuation and the time-dependent flow field is characterized using simultaneous force and particle image velocimetry (PIV) measurements that are taken phase-locked to the commanded actuation waveform. The effect of the unsteady motion on the model-embedded flow control is assessed in both trajectory tracking and disturbance rejection maneuvers. The time-varying aerodynamic lift and pitching moment are estimated from a PIV wake survey using a reduced order model based on classical unsteady aerodynamic theory. These measurements suggest that the entire flow over the airfoil readjusts within 2-3 convective time scales, which is about two orders of magnitude shorter than the characteristic time associated with the controlled maneuver of the wind tunnel model. This illustrates that flow-control actuation can be typically effected on time scales that are commensurate with the flow's convective time scale, and that the maneuver response is primarily limited by the inertia of the platform.