Future management of helicopter fleets will be more heavily based on individual component damage tracking and less on legacy usage monitoring (flight parameter-based) methods. This enhances health assessment capabilities by taking into account the actual loads on a component-by-component basis. However, accurate loads prediction in rotating frame components remains a challenge. Even with advanced computational fluid dynamics (CFD) techniques, prediction of the unsteady aerodynamic loads acting on the rotor blades is computationally intensive and problematic in terms of accurate loads prediction across the entire flight regime of the helicopter. High-speed flight can potentially introduce both shock and near-stall effects within a given rotor rotation. Low-speed flight can include blade-vortex interaction effects, wherein flow from a given blade (vorticity loading from tip vortices) impinges upon the preceding blade, causing unsteady aerodynamic loading that is difficult to quantity and predict numerically. Vehicle maneuvering can produce significantly higher blade pitching moments than steady flight. All of these regimes combine to represent the loading history of the rotor system. Therefore, accurate loads prediction methods, in terms of matching peak-to-peak, magnitude, phase, as well as vibratory/harmonic content, are required that capture all flight regimes for all critical structural components. This research focuses on the development of a loads prediction method, known as the Load Confluence Algorithm (LCA), and its application to the analysis of a large set of flight test data from the NASA/US Army UH-60A Airloads Program. The LCA combines measured response at a prescribed set of locations with a numerical model of the rotor system. For a given flight condition (steady flight, maneuvers, etc.) the numerical simulation's predicted loads distribution is iteratively incremented (by harmonic) until convergence with measured loads is reached at the prescribed locations (control points). Predicted loads response at non-instrumented locations is shown to be improved as well, thus enhancing fatigue lifing methods for these components. The procedure specifically investigates the harmonic content of the applied loads and the improved prediction of the harmonic components. The impact of the enhanced accuracy on loads predictions on component structural fatigue is illustrated by way of an example. Results show that, for a limited sensor set (two 3-axis sensors per blade), blade loads are accurately predicted across a full range of flight regimes. Hub loads are best modeled using the pushrod as the control point. Results also show that load magnitude has a tremendous influence on damage, with a 25% over-estimation of vibratory load resulting in a damage factor of nearly 3. This research highlights the importance of accurate loads prediction for a rotorcraft life tracking program. Small inaccuracies in loads lead to dramatic errors in damage assessment.

Increasing environmental concerns and fuel prices motivate the study of alternative, unconventional aircraft configurations. One such example is the blended-wing-body configuration, which has been shown to have several advantages over the conventional tube-and-wing aircraft configuration. In this thesis, a blended-wing-body aircraft is studied and optimized aerodynamically using a high-fidelity Euler-based flow solver, integrated geometry parameterization and mesh movement, adjoint-based gradient evaluation, and a sequential quadratic programming algorithm. Specifically, the aircraft is optimized at transonic conditions to minimize the sum of induced and wave drag. These optimizations are carried out with both fixed and varying airfoil sections. With varying airfoil sections and increased freedom, up to 52% drag reduction relative to the baseline geometry was achieved: at the target lift coefficient of 0.357, a drag coefficient of 0.01313 and an inviscid lift-to-drag ratio of 27.2 were obtained.

Increasing environmental concerns and fuel prices motivate the study of alternative, unconventional aircraft configurations. One such example is the blended-wing-body configuration, which has been shown to have several advantages over the conventional tube-and-wing aircraft configuration. In this thesis, a blended-wing-body aircraft is studied and optimized aerodynamically using a high-fidelity Euler-based flow solver, integrated geometry parameterization and mesh movement, adjoint-based gradient evaluation, and a sequential quadratic programming algorithm. Specifically, the aircraft is optimized at transonic conditions to minimize the sum of induced and wave drag. These optimizations are carried out with both fixed and varying airfoil sections. With varying airfoil sections and increased freedom, up to 52% drag reduction relative to the baseline geometry was achieved: at the target lift coefficient of 0.357, a drag coefficient of 0.01313 and an inviscid lift-to-drag ratio of 27.2 were obtained.

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The environmental and monetary cost of energy has renewed interest in horizontal-axis wind turbines (HAWT). One problem with HAWT design is turbulent winds, which cause cyclic loading and reduced life. Controlling short-term aerodynamic fluctuations with blade pitching or mechanical flaps is limited by the speed of actuation. The objective was to investigate using jet-flap-like fluidic actuators on the 'suction surface' of an aerofoil for rapid aerodynamic control. A NACA 0025 aerofoil was constructed for wind-tunnel experiments. The low Reynolds number (Re) flow was measured non-intrusively with particle image velocimetry (PIV). The jet showed limited effect compared to published work. The sharp trailing edge and distance to the jet were determined to be critical factors. At Re≈100000 the 'suction surface' jet sheet is less useful for control than the conventional 'pressure surface' sheet. The experiment suggests usage near the blade root on truncated aerofoils.

The environmental and monetary cost of energy has renewed interest in horizontal-axis wind turbines (HAWT). One problem with HAWT design is turbulent winds, which cause cyclic loading and reduced life. Controlling short-term aerodynamic fluctuations with blade pitching or mechanical flaps is limited by the speed of actuation. The objective was to investigate using jet-flap-like fluidic actuators on the 'suction surface' of an aerofoil for rapid aerodynamic control. A NACA 0025 aerofoil was constructed for wind-tunnel experiments. The low Reynolds number (Re) flow was measured non-intrusively with particle image velocimetry (PIV). The jet showed limited effect compared to published work. The sharp trailing edge and distance to the jet were determined to be critical factors. At Re≈100000 the 'suction surface' jet sheet is less useful for control than the conventional 'pressure surface' sheet. The experiment suggests usage near the blade root on truncated aerofoils.

This dissertation provides numerical studies to improve bladed disk assembly design for preventing blade high cycle fatigue failures. The analyses are divided into two major subjects. For the first subject presented in Chapter 2, the mechanisms of transonic fan flutter for tuned systems are studied to improve the shortcoming of traditional method for modern fans using a 3D time-linearized Navier-Stokes solver. Steady and unsteady flow parameters including local work on the blade surfaces are investigated. It was found that global local work monotonically became more unstable on the pressure side due to the flow rollback effect. The local work on the suction side significantly varied due to nodal diameter and flow rollback effect. Thus, the total local work for the least stable mode is dominant by the suction side. Local work on the pressure side appears to be affected by the shock on the suction side. For the second subject presented in Chapter 3, sensitivity studies are conducted on flutter and forced response due to frequency mistuning and aerodynamic asymmetry using the single family of modes approach by assuming manufacturing tolerance. The unsteady aerodynamic forces are computed using CFD methods assuming aerodynamic symmetry. The aerodynamic asymmetry is applied by perturbing the influence coefficient matrix. These aerodynamic perturbations influence both stiffness and damping while traditional frequency mistuning analysis only perturbs the stiffness. Flutter results from random aerodynamic perturbations of all blades showed that manufacturing variations that effect blade unsteady aerodynamics may cause a stable, perfectly symmetric engine to flutter. For forced response, maximum blade amplitudes are significantly influenced by the aerodynamic perturbation of the imaginary part (damping) of unsteady aerodynamic modal forces. This is contrary to blade frequency mistuning where the stiffness perturbation dominates. / Dissertation

Turbulent combustion processes are inherently unsteady and, thus, a source of acoustic radiation, which occurs due to the unsteady expansion of reacting gases. While prior studies have extensively characterized the total sound power radiated by turbulent flames, their spectral characteristics are not well understood. The objective of this research work is to measure the flow and acoustic properties of an open turbulent premixed jet flame and explain the spectral trends of combustion noise. The flame dynamics were characterized using high speed chemiluminescence images of the flame. A model based on the solution of the wave equation with unsteady heat release as the source was developed and was used to relate the measured chemiluminescence fluctuations to its acoustic emission. Acoustic measurements were performed in an anechoic environment for several burner diameters, flow velocities, turbulence intensities, fuels, and equivalence ratios. The acoustic emissions are shown to be characterized by four parameters: peak frequency (Fpeak), low frequency slope (beta), high frequency slope (alpha) and Overall Sound Pressure Level (OASPL). The peak frequency (Fpeak) is characterized by a Strouhal number based on the mean velocity and a flame length. The transfer function between the acoustic spectrum and the spectrum of heat release fluctuations has an f^2 dependence at low frequencies, while it converged to a constant value at high frequencies. Furthermore, the OASPL was found to be characterized by (Fpeak mfH)^2, which resembles the source term in the wave equation.

Morphing aircraft are envisioned to have multirole capability where the ability to change shape allows for adaptation to a changing mission environment. In order to calculate the properties of many wing configurations efficiently and rapidly, a model of a morphing aircraft is needed. This paper develops an aerodynamic model and a dynamic model of a morphing flying wing aircraft. The dynamic model includes realistic aerodynamic forces, consisting of lift, drag, and pitching moment about the leading edge, calculated using a constant strength source doublet panel method. The panel method allows for the calculation of aerodynamic forces due to large scale shape changing effects. The aerodynamic model allows for asymmetric configurations in order to generate rolling and yawing moments. The dynamic model calculates state information for the morphing wing based on the aerodynamic forces from the panel method. The model allows for multiple shape changing degrees-of-freedom for the wing, including thickness, sweep, dihedral angle, and chord length. Results show the model provides a versatile and computationally efficient tool for calculating the aerodynamic forces on the morphing aircraft and using these forces to show the associated states.

This thesis focuses on developing and examining the capabilities of a new in-house aerodynamic analysis tool, AeroSIM+, and investigating rotor-rotor aerodynamic interactions for two helicopters, one behind the other in forward flight. AeroSIM+ is a 3-D unsteady vortex panel method potential flow solver based on a free vortex wake methodology. Validation of the results with the experimental data is performed using the Caradonna-Tung hovering rotor test case. AeroSIM+ code is improved for forward flight conditions so that, the blades are allowed to move according to the rotor dynamics. In the simulations, blade airload prediction is seen to be sensitive to changes in vortex core size. Blade Vortex Interaction (BVI) locations differ depending on the relative position of the rear rotor with respect to the front rotor as well as on the forward flight speed. It was observed that the performance characteristics of the rear rotor alter depending on the relative positions of the rotors within the asymmetric wake flow field. The results of this thesis study such as the computed forces and moments on each rotor and the frequency characteristics of these loads can be also used in helicopter dynamics simulators.