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 = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.500 -0.4095 0.08762 0.08438 -0.0653 1.0000 0.0532 -10.250 -0.4330 0.08285 0.07961 -0.0671 1.0000 0.0532 -10.000 -0.4609 0.07913 0.07589 -0.0673 1.0000 0.0533 -9.750 -0.4877 0.07672 0.07354 -0.0652 1.0000 0.0532 -9.500 -0.5019 0.07183 0.06833 -0.0756 0.9844 0.0537 -9.000 -0.4691 0.05872 0.05528 -0.0837 0.9633 0.0579 -8.750 -0.4554 0.05461 0.05100 -0.0877 0.9430 0.0607 -8.500 -0.4535 0.04972 0.04540 -0.0938 0.9218 0.0683 -8.250 -0.4163 0.04564 0.04142 -0.0977 0.9112 0.0722 -8.000 -0.3854 0.04143 0.03666 -0.1043 0.8963 0.0831 -7.750 -0.3437 0.02953 0.02308 -0.1041 0.8849 0.0317 -7.500 -0.2992 0.02617 0.01929 -0.1079 0.8697 0.0311 -7.250 -0.2580 0.02446 0.01718 -0.1107 0.8528 0.0324 -7.000 -0.2227 0.02267 0.01512 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2.750 0.7649 0.01897 0.01185 -0.0980 0.6129 0.7672 3.000 0.7923 0.01903 0.01201 -0.0987 0.6067 0.7719 3.250 0.8177 0.01903 0.01204 -0.0984 0.6019 0.7744 3.500 0.8465 0.01907 0.01205 -0.0987 0.5981 0.7768 3.750 0.8685 0.01917 0.01227 -0.0979 0.5919 0.7799 4.000 0.8974 0.01917 0.01228 -0.0984 0.5863 0.7836 4.250 0.9336 0.01919 0.01224 -0.1006 0.5816 0.7873 4.500 0.9527 0.01922 0.01242 -0.0991 0.5745 0.7896 4.750 0.9795 0.01919 0.01241 -0.0990 0.5695 0.7921 5.000 1.0055 0.01926 0.01255 -0.0989 0.5639 0.7952 5.250 1.0318 0.01927 0.01263 -0.0990 0.5569 0.7987 5.500 1.0677 0.01920 0.01249 -0.1009 0.5517 0.8018 5.750 1.0868 0.01921 0.01269 -0.0995 0.5433 0.8043 6.000 1.1150 0.01906 0.01252 -0.0995 0.5373 0.8070 6.250 1.1354 0.01909 0.01270 -0.0984 0.5287 0.8101 6.500 1.1661 0.01896 0.01255 -0.0991 0.5221 0.8130 6.750 1.1897 0.01901 0.01276 -0.0988 0.5129 0.8163 7.000 1.2173 0.01891 0.01265 -0.0989 0.5057 0.8190 7.250 1.2364 0.01890 0.01280 -0.0974 0.4962 0.8217 7.500 1.2595 0.01889 0.01287 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12.500 1.3955 0.03893 0.03334 -0.0557 0.1669 0.9078 12.750 1.3891 0.04162 0.03606 -0.0544 0.1531 0.9168 13.000 1.3808 0.04442 0.03888 -0.0530 0.1397 0.9297 13.250 1.3690 0.04712 0.04164 -0.0513 0.1280 1.0000 13.500 1.3662 0.05073 0.04521 -0.0521 0.1128 1.0000 13.750 1.3636 0.05439 0.04886 -0.0529 0.0986 1.0000 14.000 1.3594 0.05830 0.05273 -0.0538 0.0861 1.0000 14.250 1.3552 0.06231 0.05672 -0.0548 0.0754 1.0000 14.500 1.3530 0.06618 0.06061 -0.0558 0.0660 1.0000 14.750 1.3502 0.07016 0.06461 -0.0569 0.0587 1.0000 15.000 1.3432 0.07477 0.06920 -0.0581 0.0528 1.0000 15.250 1.3438 0.07847 0.07301 -0.0593 0.0466 1.0000 15.500 1.3371 0.08321 0.07776 -0.0607 0.0420 1.0000 15.750 1.3349 0.08742 0.08206 -0.0621 0.0366 1.0000 16.000 1.3262 0.09266 0.08734 -0.0638 0.0319 1.0000 16.250 1.3167 0.09815 0.09288 -0.0658 0.0273 1.0000 16.500 1.3059 0.10391 0.09874 -0.0679 0.0235 1.0000 16.750 1.2977 0.10938 0.10428 -0.0701 0.0209 1.0000