XFOIL Version 6.96 Calculated polar for: FX 61-168 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 ------ -------- --------- --------- -------- -------- -------- -12.500 -0.5716 0.06537 0.06341 -0.0666 1.0000 0.0059 -12.250 -0.6019 0.05827 0.05615 -0.0699 1.0000 0.0059 -12.000 -0.6250 0.05319 0.05096 -0.0714 1.0000 0.0058 -11.750 -0.6650 0.04470 0.04209 -0.0752 0.9990 0.0058 -11.500 -0.6750 0.03845 0.03551 -0.0802 0.9958 0.0057 -11.250 -0.6719 0.03431 0.03104 -0.0832 0.9913 0.0059 -10.750 -0.6487 0.02915 0.02544 -0.0867 0.9750 0.0060 -10.500 -0.6563 0.02462 0.02048 -0.0853 0.9511 0.0063 -10.250 -0.6052 0.02250 0.01819 -0.0926 0.9348 0.0066 -10.000 -0.5228 0.02078 0.01625 -0.1058 0.9060 0.0070 -9.750 -0.4502 0.01954 0.01464 -0.1166 0.8525 0.0074 -9.500 -0.4268 0.01882 0.01362 -0.1169 0.8036 0.0077 -9.250 -0.4106 0.01823 0.01282 -0.1157 0.7721 0.0079 -9.000 -0.3962 0.01764 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