XFOIL Version 6.96 Calculated polar for: NACA-0009 9.0% smoothed 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.250 -0.9442 0.10031 0.09756 -0.0049 1.0000 0.0043 -15.000 -1.0016 0.08451 0.08164 -0.0131 1.0000 0.0040 -14.750 -1.0614 0.06779 0.06469 -0.0232 1.0000 0.0038 -14.500 -1.0959 0.05647 0.05314 -0.0314 1.0000 0.0036 -14.250 -1.1134 0.04947 0.04595 -0.0365 1.0000 0.0037 -14.000 -1.1313 0.04362 0.03989 -0.0400 1.0000 0.0036 -13.750 -1.1399 0.03993 0.03603 -0.0412 1.0000 0.0037 -13.500 -1.1486 0.03684 0.03276 -0.0409 1.0000 0.0038 -13.250 -1.1526 0.03457 0.03034 -0.0397 1.0000 0.0039 -13.000 -1.1568 0.03256 0.02817 -0.0376 1.0000 0.0039 -12.750 -1.1591 0.03086 0.02631 -0.0348 1.0000 0.0040 -12.500 -1.1592 0.02946 0.02477 -0.0316 1.0000 0.0041 -12.250 -1.1571 0.02819 0.02335 -0.0284 1.0000 0.0042 -12.000 -1.1498 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