Date of Award

Winter 1989

Document Type

Dissertation

Degree Name

Doctor of Philosophy (PhD)

Department

Mechanical & Aerospace Engineering

Program/Concentration

Mechanical Engineering

Committee Director

Surendra N. Tiwari

Committee Director

Ajay Kumar

Committee Member

E. Von Lavante

Committee Member

Oktay Baysal

Committee Member

J. Mark Dorrepaal

Abstract

A numerical study has been conducted to investigate the effects of blunt leading edges on the viscous flow field around a hypersonic vehicle such as the proposed National Aero-Space Plane. Attention is focused on two specific regions of the flow field. Analysis of these flow regions is required to accurately predict the overall flow field as well as to get necessary information on localized zones of high pressure and intense heating.

The forebody is modeled by slender cones and ogives with spherically blunted nose. A combination of Navier-Stokes and parabolized Navier-Stokes equations is used to compute the flow field. The influence of entropy layer thickness on the extent of the leading edge effects is also considered. The extent of downstream effects of leading edge thickness are determined at Mach numbers of 10 and 20 for cone angles of 5°, 10°, and 20°. Three values of nose bluntness are considered with the smallest nose blunting (0.0025m) representing the sharp cone/ogive.

For the flow region around the inlet the forebody shock can interact either with the blunt cowl leading edge shock or with the shock produced by the blunt leading edges of the swept sidewall compression surfaces of the inlet. For the interaction at the cowl leading edge, the forebody shock is assumed planar and the cowl is modeled by a two-dimensional cylindrically blunted wedge of infinite width. Use of the full Navier-Stokes equations is made on the cowl forebody and the thin-layer Navier-Stokes equations are suitably modified for space marching on the cowl afterbody. The results of the study show that the flow around the cowl is significantly altered by the impinging shock. The peak value of pressure is found to be nine times and heating rates eight times the stagnation point value for unimpinged case at Mach 8.03. The peak values were slightly lower for Mach 5.94 calculations. A three-dimensional thin-layer Navier-Stokes code has been used to calculate the flow field. The peak pressure for this case is found to be 2.25 times and the peak heating three times the unimpinged stagnation values. The results of the study are compared with the available experimental and numerical results.

DOI

10.25777/2ztc-x981

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