A Flight Investigation of the Effect of Shape and Thickness of the Boundary Layer on the Pressure Distribution in the Presence of Shock
Abstract
An investigation was made in flight at free-stream Mach numbers up to about 0.77 to determine the effect of a laminar boundary layer and thin and thick turbulent boundary layers on the chordwise pressure distribution over an airfoil in the presence of shock at full-scale Reynolds numbers. Boundary-layer and pressure-distribution measurements were made on a short-span airfoil built around the wing of a fighter airplane. Boundary-layer Reynolds numbers (based on momentum thickness and flow parameters at the outer edge of the boundary layer) were about 3,000 for the laminar boundary layer and 10,000 for the thickest turbulent boundary layer with local Mach numbers ranging up to 1.3 and chord Reynolds numbers up to about 21 x 10(exp 6). The results indicated very little difference in pressure distribution with laminar and turbulent boundary layers extending up to the position of shock. The principal difference was a 2- to 3-percent-chord more forward position of the pressure rise at the surface with the turbulent boundary layers. Other investigations made at low Reynolds numbers (of the order of 3 x 10(exp 6) indicated large pressure differences extending over an appreciable extent in the chordwise direction.
Document Details
- Document Type
- Technical Report
- Publication Date
- Sep 01, 1952
- Accession Number
- ADA377069
Entities
People
- Eziaslav N. Harrin
Organizations
- National Aeronautics and Space Administration