Scale-effect tests in a turbulent tunnel of the NACA 65(sub 3)-418, a = 1.0 airfoil section with 0.20-airfoil-chord split flap Page: 9 of 22
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NACA ACR No. 14122 7.
.. - For the airfoi-l-with flap defl66ted, t-results show
no consistent variation of section profile-drag coefficient
with Reynolds number. In fact, it may be concluded from
these results that the section profile-drag coefficient
with flap deflected is, to a first approximation, inde-
pendent of Reynolds number.
Pitohing moment.- The somewhat irregular curves of
section pit cing-moment coefficient at the lowest
Reynolds numbers appear to be caused by the inaccuracy
of the tunnel balance system at the low speeds. This
inaccuracy is also shown by the large difference between
the original and check tests at R = 0.19 x 106 (fig. 2(a)).
Accuracy at Reynolds numbers higher than 0.19 x 10 is
much better, as shown by the table in the section
entitled "Precision." The slope of the pitching-moment-
coefficient curve of the plain airfoil becomes slightl
more negative with increase in Reynolds number (fig. 6).
The pitching-moment-coefficient slope for the airfoil with
flap deflected varied with lift coefficient in such a way
that presentation of the slopes was not practicable.
Scale-effect tests of the NACA 653-418, a = 1.0 air-
foil section with a split flap having a chord 20 percent
of the airfoil chord have been made in the LMAL 7- by 10-
foot tunnel. The Reynolds number ranged from 0.19 to
2.99 x 106; the Mach number attained was never greater
than 0.10. From these tests, the following conclusions
have been drawn:
1. The maximum lift coefficient increased with
Reynolds number. Deflecting the flap added an increment
of maximum lift coefficient that seemed to be .almost
constant at all Reynolds numbers.
2. The slope of the section lift curve with flap
deflected showed no consistent variation with Reynolds
number, although the slope of the section lift curve for
the plain airfoil increased up to a Reynolds number of
about 1.0 x 106 and then remained nearly constant up to
a Reynolds number of about 3.0 x 10 ,the limit of the
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Tucker, Warren A. & Wallace, Arthur R. Scale-effect tests in a turbulent tunnel of the NACA 65(sub 3)-418, a = 1.0 airfoil section with 0.20-airfoil-chord split flap, report, September 1944; (digital.library.unt.edu/ark:/67531/metadc61404/m1/9/: accessed November 15, 2018), University of North Texas Libraries, Digital Library, digital.library.unt.edu; crediting UNT Libraries Government Documents Department.