No description available.

No description available.

No description available.

Mode of access: Internet via World Wide Web. Available at http://hdl.handle.net/1947/8028. / "February 2007"

Wall correction theories for two-dimensional airfoils in wind tunnels with partly open walls are examined. Conventional wall correction theories are linearized theories, valid only for small thin models of slight camber at low angles of attack. Such theories are shown to be useless for the prediction of the required wall corrections for large models, models at high angles of attack, or models developing high lift. An exact numerical theory is presented in which it is not necessary to make these assumptions. The airfoil and any solid wall sections are represented by surface source and vortex singularities as in the method of A.M.O. Smith. Aerodynamic lift is determined by numerical integration of the calculated pressure distributions around the airfoil contour. The theory indicates that certain wall configurations will require small or negligible wall corrections for tests on lifting airfoils. / Applied Science, Faculty of / Mechanical Engineering, Department of / Graduate

This thesis deals with a new approach to reduce wall corrections in high-lift airfoil testing, by employing two similar non-uniform transversely slotted walls. The solid elements of the slotted wall are symmetrical airfoils at zero incidence, and the spaces between the slats are non-uniform, increasing linearly towards the rear. This wall configuration provides the flow conditions close to the free air test environment which lead to negligible or small wall corrections. The theory uses the potential flow surface vortex-element method, with "Full Load" Kutta Conditions satisfied on the test airfoil and wall slats. This method is very well supported by physical evidence and it is simple to use. The surface velocities can be calculated directly and the aerodynamic lift and pitching moment are determined by numerical integration of the calculated pressure distributions around the airfoil contour. This method can be developed to include a simulation of the flow in the plenum chambers in the analysis. Also, the performance of this new wall configuration was examined experimentally. Two different sizes of NACA-0015 airfoil were tested in the existing low speed wind tunnel after modifying both the configuration of the side walls and the test section to accommodate the new test. Pressure distributions about the test airfoils were measured using pressure taps around their contours. Also the lifts and the pitching moments were obtained by integrating the measured surface pressures. The experimental, results show that the use of the new wall configuration with AOAR = 59% would produce wind tunnel test data very close to the free air values. / Applied Science, Faculty of / Mechanical Engineering, Department of / Graduate

No description available.

No description available.

Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, June 2003. / Thesis advisor(s): E. Roberts Wood, Richard M. Howard. Includes bibliographical references (p. 53-54). Also available online.

The purpose of this research was to demonstrate the viability of a method, due to Professor W. R. Sears, for obtaining wind-tunnel wall-corrections from measurements of near-field flow parameters by an interative procedure. A case is made for the improved accuracy of this method over the standard method of images. The wall-correction method was applied to an actual wind-tunnel test of a slightly oversized wing model at low subsonic speeds (Mach number ≈ 0.1). The wind tunnel facility and experimental setup and method are described and discussed. The wall-correction method consists of iterating between the region of space exterior to the test section boundary and the one interior to it. The flow fields in both regions are defined in terms of plane singularity elements each with an unknown, constant strength distribution. The method for expressing these flow fields as a linear system and for obtaining the associated matrices is described. The boundary conditions for the inner flow are slightly different from those of the outer flow because of the presence of the wing. There are actually two different but consistent sets of boundary conditions at the wing which lead to two different but compatible calculations for the wall-correction. The near-field flow parameter measured during the wind-tunnel test is the wing perturbation velocity potential, obtained from the quantity p ͚ - pᵢ. Here, i represents any of the 46 static taps distributed over the test section walls. It was decided to use 140 singularity elements for the outer flow description; therefore, a method was devised for fitting a least-squares surface to the measured p̂ᵢ's and integrating to obtain 140φᵢ's. The procedure for the iterations is described and the criterion for convergence to unconfined flow is presented. Test cases consisting of known, simple flows are used along the way to verify the computational methods. Finally, the wall correction to the lift curve of the wing model is presented as well as the correction at a typical tail position and the correction to the induced drag of the wing.