Date of Award

Spring 1973

Document Type

Thesis

Department

Mechanical & Aerospace Engineering

Program/Concentration

Mechanical Engineering

Committee Director

G. L. Goglia

Call Number for Print

Special Collections; LD4331.E56G67

Abstract

A numerical study of the turbulent boundary layer in a hypersonic nozzle is presented. The boundary layer equations are solved by a second order, implicit finite difference scheme developed at Langley Research Center. The technique is applicable to laminar, transitional or turbulent compressible or incompressible boundary layer flows for planar or axisymmetric geometries. A technique is presented for utilizing experimental data to start the finite difference solution.

Turbulent boundary-layer measurements on axisymmetric nozzle walls at Mach number 6 are presented and compared with numerical predictions obtained by starting the solution in the stagnation chamber and from a downstream experimental station. Accurate numerical predictions were obtained for all cases when the classical boundary layer equations were valid. The numerical solution utilizing local data to start the finite difference solution improves agreement with measured velocity and temperature profiles and displacement thickness.

Boundary layer measurements and numerical results indicate that the Crocco — velocity relation relaxes from a quadratic-type variation toward a more linear variation as the pressure gradient relaxes to zero.

For low Reynolds number flows modification of the eddy viscosity model utilized in the present finite difference scheme improves the accuracy of the predictions. Local application of an approximate low Reynolds number eddy viscosity model indicates improved agreement in boundary layer thickness, displacement thickness, and momentum thickness.

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DOI

10.25777/a76t-ns67

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