XFOIL Version 6.96 Calculated polar for: FX 60-160 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.250 -0.4457 0.09044 0.08548 -0.0653 1.0000 0.0379 -11.000 -0.4761 0.08275 0.07783 -0.0683 1.0000 0.0375 -10.750 -0.5098 0.07693 0.07202 -0.0697 1.0000 0.0373 -10.500 -0.5370 0.06968 0.06464 -0.0754 0.9953 0.0370 -10.250 -0.5566 0.06262 0.05730 -0.0827 0.9853 0.0371 -10.000 -0.5723 0.05741 0.05176 -0.0873 0.9731 0.0374 -9.750 -0.5781 0.05244 0.04635 -0.0900 0.9611 0.0380 -9.500 -0.5838 0.04806 0.04129 -0.0911 0.9489 0.0395 -9.250 -0.5591 0.04582 0.03905 -0.0935 0.9449 0.0413 -9.000 -0.5454 0.04416 0.03726 -0.0933 0.9353 0.0427 -8.750 -0.5230 0.04173 0.03450 -0.0952 0.9305 0.0450 -8.500 -0.5114 0.03966 0.03200 -0.0943 0.9204 0.0472 -8.250 -0.4867 0.03730 0.02905 -0.0956 0.9158 0.0490 -8.000 -0.4704 0.03529 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