XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC-SC2 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.250 -1.0743 0.08216 0.07950 -0.0329 1.0000 0.0131 -16.000 -1.1334 0.06520 0.06226 -0.0447 1.0000 0.0131 -15.750 -1.1627 0.05531 0.05214 -0.0522 1.0000 0.0131 -15.500 -1.1807 0.04878 0.04543 -0.0566 1.0000 0.0131 -15.250 -1.1925 0.04405 0.04053 -0.0590 1.0000 0.0131 -15.000 -1.2004 0.04044 0.03676 -0.0599 1.0000 0.0131 -14.750 -1.2051 0.03756 0.03374 -0.0598 1.0000 0.0131 -14.500 -1.2073 0.03520 0.03125 -0.0590 1.0000 0.0132 -14.250 -1.2072 0.03325 0.02916 -0.0577 1.0000 0.0132 -14.000 -1.2134 0.03094 0.02669 -0.0554 1.0000 0.0133 -13.750 -1.2147 0.02920 0.02482 -0.0528 1.0000 0.0134 -13.500 -1.2128 0.02787 0.02338 -0.0500 1.0000 0.0135 -13.250 -1.2097 0.02685 0.02227 -0.0467 1.0000 0.0136 -13.000 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0.04750 0.04362 0.0336 0.0136 1.0000 16.500 1.3432 0.05180 0.04807 0.0323 0.0136 1.0000 16.750 1.3305 0.05704 0.05346 0.0298 0.0136 1.0000 17.000 1.3126 0.06372 0.06032 0.0259 0.0136 1.0000 17.250 1.2880 0.07232 0.06911 0.0204 0.0136 1.0000 17.500 1.2524 0.08361 0.08062 0.0132 0.0136 1.0000