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National Advisory Committee for Aeronautics, Technical Notes - An Experimental Study of the Lift and Pressure Distribution on a Double Wedge Profile at Mach Numbers Near Shock

An account is given of wind-tunnel measurements at low supersonic
speeds of the pressure distribution on a doubly symmetrical double-wedge
profile of approximately 8-percent thickness. The results cover the Mach
number range from 1.166 to 1.377, which brackets the value (1.221) given
by exact inviscid theory for attachment of the shock wave to the leading
edge at zero angle of attack. Data are given for angles of attack from
0° to 5° at a Reynolds number of 0.5% million. The results are discussed
in detail and compared with theoretical findings previously obtained on
the basis of the transonic small—disturbance theory.

As predicted by the theory, the experimental results show a large
increase in the initial lift-curve slope at Mach numbers near shock
attachment. On the front wedge, where viscous effects are small, the
numerical agreement between experiment and theory is good at the smaller
angles of attack. This agreement tends to deteriorate, however, as the
angle is increased. As might be expected from qualitative arguments
regarding the limitations of the theory, this deterioration proceeds more
rapidly the closer the Mach number is to the attachment value. As a
result, the increase in lift-curve slope at Mach numbers near shock
attachment disappears at the higher angles. On the rear wedge, where
viscous effects are large, the data at small angles of attack show an
unpredicted region of negative lift in the vicinity of the trailing edge.

In the case of the pressure drag due to angle of attack, agreement
between theory and experiment is Observed at small angles only when the
Mach number is above the attachment value. At Mach numbers below this
value, the drag rises less rapidly with angle of attack than is calculated
on the basis of the theoretical pressure differences between the top and
bottom of the airfoil. The measured drag and pressure distributions at
zero angle of attack agree well with existing theoretical and experimental
results throughout the Mach number range.