INVESTIGATION OF BOUNDARY LAYER TRANSITION IN HYPERSONIC FLOW AT ANGLE OF ATTACK
Abstract
An analytical and experimental investigation has been made of boundary layer transition at angle of attack in hypersonic flow. A new method of calculating laminar flow stability has been developed that is more general than previous methods. The method employs the timewise integration of perturbation equations derived from the complete Navier-Stokes equations for two-dimensional flow of a compressible fluid with variable transport properties. The present calculations are for perturbations about a known steady flow. The perturbations are assumed to be sinusoidal in the streamwise direction, but no restriction is known that would prevent the streamwise wave form from being arbitrary, as are the vertical distributions and the time variations. The sinusoidal wave assumption greatly reduces the amount of computer time required and simplifies the interpretation of the results. Nonlinear terms are easily retained with the present method, but nonlinear calculations are not physically correct when perturbations are required to be sinusoidal. For this reason, nonlinear terms were deleted for all calculations presented. The experimental program was conducted at Mach numbers of 6, 8, and 10. The models included blunt and sharp flat plates and a 75-degree swept delta wing with sharp leading edges.
Document Details
- Document Type
- Technical Report
- Publication Date
- Aug 01, 1966
- Accession Number
- AD0803781
Entities
People
- A. L. Nagel
- R. T. Savage
- R. W. Blank
- R. Wanner
Organizations
- Boeing