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National Advisory Committee for Aeronautics, Research Memorandum - Aerodynamic Study of a Wing Fuselage Combination Employing a Wing Swept Back 63° - Investigation at a Lift Coefficient of 0.25

Tests have been performed at a Mach number of 1.53 with a wing-
fuselage combination having a wing with 63° leading—edge sweep,
an aspect ratio of 3.16, and a taper ratio of 0.25. ‘Jhis wing had an
NACA 61+AOO5 thickness distribution parallel to the plane of symmetry
and was cambered and twisted. The principal object of the investi—
gation was to determine the effects of camber and. twist on the maximum
lift—drag ratio and pitching-moment characteristics. The results
obtained from these tests are compared with those of a preceding ,
investigation employing the same fuselage and wing plan form but with
a wing having an NACA 6lmoo6 thickness distribution and no camber or
twist. Tests were also performed to determine the characteristics
of configurations obtained by rotating the cambered, twisted. wing
panels to 67° and 70°. leading—edge sweep. The effects of Reynolds
number were also investigated.

At a Reynolds number of 0.81L million, the thinner, cambered, and
twisted—wing configuration with 63° leading-edge sweep had a maxinmm
lift-drag ratio of 8.3 as compared with 7.2 for the similar configu-
ration of the earlier investigation. This increase resulted from a
decrease in minimum drag coefficient, a displacement of the minimum
of the drag curve to a positive lift coefficient and. a decrease in
the rate of drag rise with increased lift coefficient. Although the
total center—of—lift travel was greater for the cambered-wing configu—
ration, the change in center-of-lift location with lift coefficient
near that for maximum lift—drag ratio was reduced.

As in the earlier investigation, the sweep angle for maximum
lift—drag ratio at a Mach number of 1.53 Was found to be approximately
67°. The magnitude of the maximum lift—drag ratio increased, with
increased Reynolds number. values of 7.3 and 8.3 were obtained at
Reynolds numbers of 0.62 and 0.8% million, respectively, with the 63°
wing configuration, and values of 6.6, 7.7, and 9.0 were obtained at
Reynolds numbers of 0.31, 0.62, and 0.95 million, respectiVely, with
the 67° configuration. These results indicate that further unprove—
ment may be expected at Reynolds numbers beyond the range of the
small-scale tests.