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.

A wind tunnel was designed and built to produce a pulsating flow. The pulsation was achieved by a series of shutters (rotating with constant angular velocity) placed upstream of the settling chamber inlet. The system was optimized to obtain nearly sinusoidal velocity waveforms with the highest obtainable amplitudes over the frequency range of 3.4 to 31.3 Hz. The velocity and pressure waveforms are given for different shutters settings and conditions. It was found that the velocity waveform shape and amplitude, obtained in the test section, are a strong function of pressure build-up in the diffuser upstream of the shutters box. An explanation is given, for various shutters settings and conditions, of how pressure release is achieved and how the pressure waveforms are generated. / M.S.

M.S.

The design of a cascade wind tunnel for testing of compressor blades at high angle of attack is described. Methods to insure uniform velocity profiles and control of inlet turbulence are discussed. The problem of maintaining two-dimensional flows at high angle of attack was addressed. A tunnel capable of testing cascades of compressor blades at angles of attack up to seventy-five degrees was constructed. Performance of the tunnel was evaluated and data were acquired for flow over double-circular-arc blades with angles of attack extending into the fully-stalled region. Comparisons were made with available data in the installed flow regime. Results showed that the tunnel had adequately uniform inlet velocities and low turbulence levels, and that two-dimensional flow was maintained over the center two-thirds of the high-aspect ratio blades. / Master of Science

M.S.