A boundary-layer-transition and profile-drag investigation was conducted in flight by the National Advisory Committee for Aeronautics on an experimental low-drag wing installed on a P-47 airplane designated the XP-47F and supplied by the Army Air Forces. The wing incorporates airfoil section that vary from an NACA 66(215)-1(16.5), a = 1.0 at the plane of symmetry to an NACA 67(115)-213, a = 0.7 at the tip. The surface of the wing as constructed was found to have such a degree of waviness that it had to be refinished in order to obtain the performance generally expected of low-drag airfoils. Measurements were made at a section outside the propeller slipstream with smooth and with standard camouflage surfaces and on the upper surface of a section in the propeller slipstream with the surface smoothed. Tests were made in normal flight - that is, in level flight and in shallow dives - at indicated airspeeds ranging from about 150 to 300 miles per hour and in steady turns at 300 miles per hour with normal acceleration from 2g to 4g. These speed and acceleration limits were imposed by structural considerations. The tests in normal flight covered a range of section lift coefficient from about 0.58 to 0.15, of Reynolds number from about 9 x 106 to 18 x 106, and of Mach number from about 0.27 to 0.53. In the tests in turns at 300 miles per hour, the range of section lift coefficient was extended to 0.63.

Summary: A flight investigation was made of boundary-layer and profile-drag characteristics of smooth wing sections of a P-47D airplane. Measurements were made at three stations on the wing: boundary-layer measurements were made on the upper surface of the left wing in the slip-stream at 25 percent semispan; pressure-distribution measurements were made on the upper surface of the left wing at 63 percent semispan; and wake surveys were made at 63 percent semispan of the right wing. The tests were made in straight flight and in turns over a range of conditions in which airplane lift coefficients from 0.15 to 0.68, Reynolds numbers from 7.7 x 106 to 19.7 x 106, and Mach numbers from 0.25 to 0.69 were obtained. The results of the investigation indicated a minimum profile-drag coefficient of 0.0062 for the smooth section at 63 percent semispan. At the highest Mach number attained in the tests, the critical Mach number was exceeded by at least 0.04 with no evidence of compressibility shock losses appearing in the form of increased width of the wake or increased profile-drag coefficient. For flight conditions approaching the critical Mach number, variations in Mach number of as much as 0.17 appeared to have no effect on the profile-drag coefficient. In the slipstream, transition occurred at least as far back as 20 percent chord on the upper surface at low lift coefficients.

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Reproductions of reports, some declassified, of research done at Langley Memorial Aeronautical Laboratory during World War II. The order of reports does not represent when they were chronologically issued. Reference to the original version of each report is included.