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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -16.000 -1.1667 0.04954 0.04673 -0.0598 0.9454 0.0076 -15.750 -1.1867 0.04376 0.04076 -0.0627 0.9457 0.0076 -15.500 -1.1968 0.04004 0.03689 -0.0633 0.9461 0.0077 -15.250 -1.2014 0.03728 0.03401 -0.0629 0.9464 0.0077 -15.000 -1.2016 0.03514 0.03176 -0.0620 0.9467 0.0078 -14.750 -1.1988 0.03339 0.02990 -0.0607 0.9470 0.0079 -14.500 -1.1937 0.03189 0.02830 -0.0592 0.9473 0.0080 -14.250 -1.1866 0.03059 0.02689 -0.0576 0.9476 0.0082 -14.000 -1.1781 0.02940 0.02560 -0.0559 0.9478 0.0083 -13.750 -1.1681 0.02832 0.02442 -0.0542 0.9480 0.0084 -13.500 -1.1574 0.02732 0.02332 -0.0523 0.9481 0.0086 -13.250 -1.1441 0.02633 0.02223 -0.0509 0.9483 0.0087 -13.000 -1.1288 0.02538 0.02117 -0.0496 0.9484 0.0089 -12.750 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