XFOIL Version 6.96 Calculated polar for: FX 66-S-196 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -15.500 -0.3303 0.10358 0.10071 -0.0945 0.8487 0.0172 -15.250 -0.3727 0.09016 0.08719 -0.1014 0.8415 0.0165 -15.000 -0.4739 0.06799 0.06461 -0.1147 0.8427 0.0151 -14.750 -0.4809 0.06416 0.06057 -0.1168 0.8199 0.0153 -14.500 -0.5024 0.05881 0.05497 -0.1192 0.8046 0.0155 -14.250 -0.5402 0.05199 0.04779 -0.1212 0.7927 0.0152 -14.000 -0.5322 0.05102 0.04675 -0.1216 0.7786 0.0159 -13.750 -0.5538 0.04658 0.04201 -0.1221 0.7685 0.0157 -13.500 -0.5597 0.04409 0.03934 -0.1222 0.7586 0.0160 -13.250 -0.5717 0.04077 0.03576 -0.1218 0.7504 0.0159 -13.000 -0.5731 0.03890 0.03371 -0.1214 0.7419 0.0164 -12.750 -0.5763 0.03664 0.03122 -0.1207 0.7347 0.0166 -12.500 -0.5750 0.03467 0.02906 -0.1199 0.7275 0.0166 -12.250 -0.5726 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0.6408 7.000 1.2404 0.01396 0.00812 -0.1074 0.5363 0.6432 7.250 1.2599 0.01389 0.00821 -0.1061 0.5325 0.6457 7.500 1.2835 0.01380 0.00820 -0.1055 0.5282 0.6487 7.750 1.3104 0.01361 0.00800 -0.1054 0.5236 0.6521 8.000 1.3322 0.01360 0.00807 -0.1046 0.5186 0.6555 8.250 1.3508 0.01345 0.00802 -0.1030 0.5119 0.6594 8.500 1.3753 0.01339 0.00795 -0.1026 0.5059 0.6638 8.750 1.3871 0.01337 0.00811 -0.0998 0.4984 0.6689 9.000 1.4050 0.01337 0.00814 -0.0982 0.4916 0.6742 9.250 1.4160 0.01344 0.00837 -0.0954 0.4836 0.6805 9.750 1.4322 0.01377 0.00885 -0.0887 0.4625 0.6960 10.000 1.4387 0.01417 0.00930 -0.0855 0.4489 0.7058 10.250 1.4395 0.01486 0.00999 -0.0817 0.4314 0.7176 10.500 1.4339 0.01602 0.01115 -0.0776 0.4024 0.7315 10.750 1.4165 0.01814 0.01319 -0.0728 0.3712 0.7495 11.000 1.3988 0.02110 0.01610 -0.0694 0.3361 0.7765 11.250 1.3798 0.02511 0.02010 -0.0677 0.3014 0.8182 11.500 1.3754 0.03139 0.02641 -0.0729 0.2490 0.9271 11.750 1.3629 0.03585 0.03075 -0.0727 0.2169 1.0000 12.000 1.3454 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