Document Type

Thesis - Open Access

Award Date

1971

Degree Name

Master of Science (MS)

Department / School

Mechanical Engineering

First Advisor

Edward Lumsdaine

Abstract

A supersonic wind-tunnel has been constructed using an 8 inch by 10 inch single stage jet vacuum ejector with a capacity of handling 0.4 lb/sec air at 4 inches Hg absolute suction pressure. A Toepler Schlieren system with 10-inch diameter parabolic mirrors has been set up for flow pattern visualization. The same apparatus was developed to obtain multi-colored views of the flow pattern. That was accomplished by placing an equilateral prism in front of the light source located at the focal plane of the first parabolic mirror. A spectrum was thus produced at the focal plane of the second parabolic mirror, and the color of the image on the screen could be adjusted by moving a slit (parallel to the bands of the spectrum) across the spectrum. An optical method was developed for visualization of the flow over axisymmetric bodies. A special cylindrical lens test section was designed such that a parallel beam of light coming to one side of the lens would be refracted parallel through its circular cavity and then refracted out parallel on the other side. Using the supersonic tunnel, experiments were conducted on shock boundary layer interaction at low Mach numbers (about 1.3) with two-dimensional as well as axisymmetric models. For the same area ratio the two-dimensional showed severe flow separation from one side but the axisymmetric model showed little or no flow separation both from the pressure data and the optical data. By inserting a plate horizontally, thus dividing the two-dimensional test section into two halves, the severe flow separation was almost eliminated. Although the shock wave is clearly visible when using the optical method for axisymmetric flow, minor flow separation could not be detected using this technique. Theoretical solutions to turbulent incompressible boundary layers for flow over axisymmetric or two-dimensional bodies are well known /1/*. Although these methods area fairly accurate for determining boundary layer growth, the accuracy deteriorates rapidly near the separation point. For the case of compressible flow, the situation in even worse. Thus experimental investigations of compressible boundary layers are needed. A detailed theoretical discussion of compressible boundary layers and shock boundary layer interaction for flow in an axisymmetric inlet is given in Reference 2. The construction of the supersonic tunnel described in this thesis and the testing of several two-dimensional as well as axisymmetric models using this tunnel were made to provide a facility where experimental tests can be conducted on various aspects of flow separation and shock boundary layer interaction.

Library of Congress Subject Headings

Boundary layer control
Wind tunnels
Shock waves

Format

application/pdf

Number of Pages

60

Publisher

South Dakota State University

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