Thesis (Ph. D.)--Virginia Polytechnic Institute and State University, 1992. / Vita. Abstract. Includes bibliographical references (leaves 130-135). Also available via the Internet.

Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, September 1990. / Thesis Advisor(s): Ekaterinaris, John A. ; Platzer, Max. "September 1990." Description based on title screen as viewed on December 21, 2009. DTIC Identifier(s): Dynamics, leading edges, airfoils, dynamic stall, oscillating airfoil, pitching airfoil, leading edge geometry, pressure gradient, theses. Author(s) subject terms: Dynamic stall, oscillating airfoil, pitching airfoil, leading edge geometry, pressure gradient. Includes bibliographical references (p. 111-112). Also available in print.

Thesis (Ph. D.)--Ohio State University, 2004. / Title from first page of PDF file. Document formatted into pages; contains xvii, 147 p.; also includes graphics (some col.). Includes abstract and vita. Advisor: Gerald Gregorek, Aeronautical and Astronautical Engineering Graduate Program. Includes bibliographical references (p. 119-122).

No description available.

This dissertation deals with the instability, known as flutter, of the lifting and control surfaces of aircraft of advanced design at high altitudes and speeds. A simple model is used to represent the aerodynamics for flutter analysis of a two-degree-of-freedom airfoil system. Flutter solutions of this airfoil system are shown to be algebraically homomorphic in that solutions about different elastic axes can be found by mapping them to those about the mid-chord. Algebraic expressions for the flutter speed and frequency are thus obtained. For the prediction of flutter of a wing at transonic speeds, an accurate and efficient computer code is developed. The unique features of this code are the capability of accepting a steady mean flow regardless of its origin, a time dependent perturbation boundary condition for describing wing deformations on the mean surface, and a locally applied three-dimensional far-field boundary condition for minimizing wave reflections from numerical boundaries. Results for various test cases obtained using this code show good agreement with the experiments and other theories.

Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, September 2002. / Thesis advisor(s): Kevin D. Jones, Max F. Platzer. Includes bibliographical references (p. 61). Also available online.

Laser Doppler velocimetry, a non-intrusive flow measurement technique, was employed to experimentally investigate two-dimensional airfoil-vortex interaction. Vortices were generated by sinusoidally oscillating a NACA 0012 airfoil about its quarter-chord at a reduced frequency of k = 2.05 and an amplitude of ±10° angle of attack. The target airfoil, a NACA 63₂A015, was immersed in the wake, two chord lengths downstream of the vortex generators trailing edge. Phase-averaged velocity measurements of the flow around the target airfoil were made with the airfoil at angles of attack of α = 0° and α = 10°. A close encounter with a counterclockwise rotating vortex was observed for both angles of attack, and a head-on collision, which split the counterclockwise rotating vortex in two, was observed for α = 10°. Vorticity fields were constructed from the velocity measurements and the circulation of the vortex was evaluated throughout the interaction. The surface pressure fluctuations on the airfoil were determined by substituting the measured velocities into the Navier-Stokes equations and numerically integrating the resulting pressure gradients. Furthermore, an extensive investigation of the undisturbed wake of the oscillating airfoil was performed in order to determine the effect of oscillation frequency and amplitude on the wake development. / Ph. D.

Computational fluid dynamics (CFD) analysis was performed in order to compare the solutions of oscillating wings with Prandtl’s lifting-line theory. Quasi-steady and steady-periodic simulations were completed using the CFD software Star-CCM+. The simulations were performed for a number of frequencies in a pure plunging setup. Additional simulations were then completed using a setup of combined pitching and plunging at multiple frequencies. Results from the CFD simulations were compared to the quasi-steady lifting-line solution in the form of the axial-force, normal-force, power, and thrust coefficients, as well as the efficiency obtained for each simulation. The mean values were evaluated for each simulation and compared to the quasi-steady lifting-line solution. It was found that as the frequency of oscillation increased, the quasi-steady lifting-line solution was decreasingly accurate in predicting solutions.

A method is presented for calculating the shape sensitivity of a wing aeroelastic response with respect to changes in geometric shape. Yates’ modified strip method is used in conjunction with Giles' equivalent plate analysis to predict the flutter speed, frequency, and reduced frequency of the wing. Three methods are used to calculate the sensitivity of the eigenvalue. The first method is purely a finite difference calculation of the eigenvalue derivative directly from the solution of the flutter problem corresponding to the two different values of the shape parameters. The second method uses an analytic expression for the eigenvalue sensitivities of a general complex matrix, where the derivatives of the aerodynamic, mass, and stiffness matrices are computed using a finite difference approximation. The third method also uses an analytic expression for the eigenvalue sensitivities but the aerodynamic matrix is computed analytically. All three methods are found to be in good agreement with each other. The sensitivities of the eigenvalues were used to predict flutter speed, frequency , and reduced frequency. These approximations were found to be in good agreement with those obtained using a complete reanalysis. However, it is recommended that higher order terms be used in the calculations in order to assure greater accuracy. / Master of Science / incomplete_metadata

Delaying stall is always an attractive option in the aerospace industry. The major benefit of delaying stall is increased lift during takeoff and landings as well as during high angle of attack situations. Devices, such as fluidic oscillators, can be integrated into wing flaps to help delay the occurrence of stall by adding energized air to the airflow on the upper surface of the wing flap. The energized air from the oscillator allows the airflow to remain attached to the upper surface of the wing flap. The fluidic oscillator being integrated in this thesis is an active flow control device (AFC). One common method for integrating any device into a wing flap is to remove a section of the flap and mechanically secure the device. A current trend in the aerospace industry is the increased use of fiber-reinforced composites to replace traditional metal components on aircraft. The traditional methods of device integration cause additional complications when applied to composite components as compared to metal components. This thesis proposes an alternative method for integration of the AFC devices, which occurs before the fabrication of wing flaps is completed and they are attached to the aircraft wing. Seven design concepts are created to reduce the complications from using current methods of integration on composite wing flaps. The concepts are based on four design requirements: aerodynamics, manufacturing, maintenance, and structure. Four of the design concepts created are external designs, which place the AFC on the exterior surface of the wing flap in two types of grooved channels. The other three designs place the AFC inside the wing flap skin and are categorized as internal designs. In order for the air exiting the AFC to reach the upper surface of the wing flap, slots are created in the wing flap skin for the internal designs. Within each of the seven design concepts two design variants are created based on foam or ribbed core types. Prototypes were created for all of the external design AFC devices and the side inserted AFC and retaining pieces. Wing flap prototypes were created for the rounded groove straight AFC design, the semi-circular groove with straight AFC, and the side inserted AFC designs. The wing flaps were created using the VARTM process with a vertical layup for the external designs. The rounded groove and semi-circular groove prototypes each went through three generations of prototypes until an acceptable wing flap was created. The side inserted design utilized the lessons learned through each generation of the external design prototypes eliminating the need for multiple generations. The lessons learned through the prototyping process helped refine the designs and determine the ease of manufacturing to be used in the design evaluation. The evaluation of the designs is based on the four design requirements stated above. The assessment of the designs uses two levels of evaluation matrices to determine the most fitting design concept. As a result of the evaluation, all four of the external designs and one of the internal designs are eliminated. The two remaining internal designs' foam core and ribbed variants are compared to establish the final design selection. The vertically inserted AFC foam core design is the most fitting design concept for the integration of an AFC device into a composite wing flap.