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National Advisory Committee for Aeronautics, Technical Notes - A Flight Investigation of the Effect of Shape and Thickness of the Boundary Layer on the Pressure Distribution in the Presence of Shock

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 measure-
ments 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 106.

The results indicated very little difference in pressure distri-
bution 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 106) indicated large pressure
differences extending over an appreciable extent in the chordwise
direction.

The interaction of shock with laminar and turbulent boundary layers
at low Reynolds numbers (up to about 3 x 106) has been investigated in
detail in recent years (refs. 1 to 5). These investigations, and
particularly that of reference 1, indicated such a large difference in
pressure distribution with laminar and turbulent boundary layers that
an airfoil under these conditions would be expected to experience
appreciably different forces and moments. At high or full—scale

Reynolds numbers, no corresponding information was available on boundary-
layer——shock interaction. In order to provide some information at full—
scale Reynolds numbers up to about 20 x 106, an investigation, reported
herein, was initiated on a short-span airfoil built around the wing of a
fighter airplane.

The purpose of this paper is to present some measurements of
pressure distribution obtained in flight at Reynolds numbers from
17.5 X 106 to 21.2 x 106 with laminar and turbulent boundary layers
extending to the position of shock. These measurements were made in
dives up to a flight Mach number of 0.766 which was sufficiently high
to give extensive regions of local supersonic flow.