Thesis (M.S.)--Brigham Young University. Dept. of Mechanical Engineering, 2008. / Includes bibliographical references (p. 79-82).

Thesis: B.S., Massachusetts Institute of Technology, Department of Mechanical Engineering, 1976 / MICROFICHE COPY AVAILABLE IN ARCHIVES AND ENGINEERING. / Includes bibliographical references. / B.S. / B.S. Massachusetts Institute of Technology, Department of Mechanical Engineering

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This thesis presents the results obtained so far in an investigation concerning viscous effects in incompressible flows over airfoils. These effects are taken into account by assuming the existence of a boundary layer which interacts with an external inviscid flow. Numerical methods for solving the inviscid and boundary-layer flows are briefly described. The main objective of the investigation is the development of an interaction technique between both regions of flow. The method chosen for the interaction is the so-called semi-inverse procedure. This procedure is derived from a perturbation analysis of the linearized versions of the governing equations. The resulting method is subjected to a stability analysis, which shows that it will break down when used in conjunction with separated-flow boundary-layer solvers. The semi-inverse procedure is tested on several airfoils, using an attached-flow boundary-layer solver. Numerical results show that the method is sufficiently accurate for engineering purposes in the low-to-medium range of angles of attack, but its applicability is questionable when there is a large separation region. Finally, recommendations are made regarding future work to overcome the limitations of the present technique. / Master of Science

From the preceding investigation, it may be concluded that the method of calculating subsonic compressible flows proposed by Gelbart is valid and practical for arbitrary airfoil shapes for which a convenient mapping function exists, It was found that the method can be carried out with comparative ease as compared to numerical methods of solution and therefore lends itself to engineering use. A particular advantage of the method is that it allows the approximate solution of the direct problem, that is, the flow about a given body at a prescribed angle of attack and freestream Mach number. / Master of Science

The panel method is reviewed, in which linear polynomials are used to approximate the vorticity distribution on the surface of a body. Two new panel methods are developed, in which quadratic and cubic polynomials are used in an attempt to make the derivative of the velocity continuous along the surface of the body. But the results are not better than those obtained by choosing the linear polynomials as the interpolations. After the three interpolation schemes are critically evaluated, numerical examples, based on the linear scheme, are presented to illustrate some applications of the method. The flow around an airfoil in a wind tunnel is calculated, the separation of the laminar boundary layer is determined by solving the boundary-layer equations with a finite-difference scheme, and the stability of the boundary layer is investigated. It is found that the walls of the wind tunnel do not affect the separation and stability of the boundary layer significantly. Finally, the interaction of a fee vortex core with an airfoil near the ground is modeled. When the free vortex core passes the airfoil along a lower trajectory the airfoil experiences a very large thrust and suction. When along a upper trajectory, the lift and drag vary in a much smaller range. No significant effect of the ground on the airfoil is observed. / M.S.

The dependence of an airfoil stall behavior upon the acoustic l environment was experimentally investigated at a Reynolds number of 200,000. The Wortmann FX-63-137-ESM airfoil section was used for the model with an aspect ratio of 4. Some acoustic disturbances could alter the transition process in the shear layer of the separation bubble on the upper surface of the airfoil. These disturbances could delay the deep leading edge stall or hasten stall recovery in some cases. A good agreement was found with the Crabtree criterion to predict the leading edge stall. / Master of Science

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.

Approved for public release; distribution is unlimited / A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and non-uniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12)10 airfoil, coating profiles of varying leading edge non-uniformity were added. This non-uniformity is typical of that expected due to fluid being drawn away from the leading edge during the coating process. The CFD code, RVCQ3D, is a steady, quasi-three-dimensional Reynolds Averaged Navier-Stokes (RANS) solver. A k-omega turbulence model was used for the Reynolds' Stress closure. The code predicted that these changes in leading edge shape can lead to alternating pressure gradients in the first few percent of chord that create small separation bubbles and possibly early transition to turbulence. The change in total pressure loss and trailing edge deviation are presented as a function of the coating non-uniformity parameter. Results are presented for six leading edge profiles over a range of incidences and inlet Mach numbers from 0.6 to 0.8. Reynolds number was 600,000 and free-stream turbulence was 6%. A two-dimensional map is provided that shows the allowable degree of coating non-uniformity as a function of incidence and inlet Mach number. / Lieutenant Commander, United States Navy