XFOIL Version 6.96 Calculated polar for: NASA/LANGLEY RC10-64C 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.750 -0.9354 0.06077 0.05871 -0.0439 1.0000 0.0109 -13.500 -0.9594 0.05207 0.04980 -0.0511 1.0000 0.0108 -13.250 -0.9838 0.04623 0.04376 -0.0534 1.0000 0.0108 -13.000 -1.0093 0.04163 0.03894 -0.0526 1.0000 0.0108 -12.750 -1.0339 0.03798 0.03507 -0.0494 1.0000 0.0109 -12.500 -1.0522 0.03549 0.03239 -0.0446 1.0000 0.0110 -12.250 -1.0613 0.03367 0.03043 -0.0400 1.0000 0.0111 -12.000 -1.0582 0.03228 0.02895 -0.0371 1.0000 0.0113 -11.750 -1.0510 0.03115 0.02773 -0.0346 1.0000 0.0114 -11.500 -1.0415 0.03015 0.02664 -0.0322 1.0000 0.0116 -11.250 -1.0315 0.02911 0.02549 -0.0299 1.0000 0.0118 -11.000 -1.0213 0.02805 0.02431 -0.0274 1.0000 0.0121 -10.750 -1.0112 0.02698 0.02312 -0.0249 1.0000 0.0123 -10.500 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18.750 1.1084 0.13320 0.13114 -0.0144 0.0091 1.0000