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 = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.000 -0.1543 0.10177 0.09779 -0.0947 0.7728 0.0748 -12.750 -0.1291 0.09933 0.09527 -0.0921 0.7666 0.0767 -12.500 -0.3803 0.06579 0.06128 -0.1232 0.8073 0.0421 -12.250 -0.3747 0.06277 0.05812 -0.1233 0.7995 0.0408 -12.000 -0.3920 0.05834 0.05358 -0.1240 0.7912 0.0402 -11.750 -0.4106 0.05456 0.04959 -0.1241 0.7847 0.0396 -11.500 -0.4289 0.05119 0.04606 -0.1237 0.7770 0.0392 -11.250 -0.4464 0.04780 0.04242 -0.1225 0.7708 0.0383 -11.000 -0.4642 0.04469 0.03906 -0.1206 0.7643 0.0376 -10.750 -0.4818 0.04183 0.03587 -0.1176 0.7583 0.0368 -10.500 -0.4926 0.03906 0.03268 -0.1144 0.7537 0.0361 -10.250 -0.4918 0.03663 0.02994 -0.1124 0.7478 0.0359 -10.000 -0.4819 0.03465 0.02769 -0.1110 0.7426 0.0362 -9.750 -0.4674 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0.02246 0.01533 -0.0873 0.6337 0.6829 0.000 0.4052 0.02249 0.01527 -0.0880 0.6323 0.6848 0.250 0.4339 0.02254 0.01523 -0.0889 0.6310 0.6869 0.500 0.4623 0.02263 0.01527 -0.0896 0.6298 0.6884 0.750 0.4792 0.02321 0.01593 -0.0886 0.6272 0.6895 1.000 0.4943 0.02388 0.01668 -0.0875 0.6240 0.6907 1.250 0.5152 0.02427 0.01710 -0.0872 0.6213 0.6919 1.500 0.5389 0.02456 0.01740 -0.0871 0.6193 0.6933 1.750 0.5646 0.02479 0.01761 -0.0874 0.6178 0.6947 2.000 0.5925 0.02494 0.01774 -0.0879 0.6166 0.6964 2.250 0.6216 0.02504 0.01781 -0.0885 0.6153 0.6984 2.500 0.6515 0.02517 0.01789 -0.0894 0.6142 0.7001 2.750 0.6167 0.02836 0.02133 -0.0822 0.6055 0.7014 3.000 0.6352 0.02898 0.02195 -0.0817 0.6032 0.7026 3.250 0.6603 0.02929 0.02226 -0.0818 0.6018 0.7043 3.500 0.6886 0.02941 0.02241 -0.0822 0.6007 0.7061 3.750 0.7201 0.02943 0.02243 -0.0831 0.5999 0.7076 4.000 0.7541 0.02936 0.02237 -0.0843 0.5991 0.7094 4.250 0.5144 0.04506 0.03825 -0.0588 0.5701 0.7105 4.500 0.5500 0.04473 0.03793 -0.0596 0.5691 0.7123 4.750 0.5664 0.04571 0.03892 -0.0590 0.5657 0.7139 5.000 0.6146 0.04456 0.03778 -0.0608 0.5666 0.7159 5.250 0.6584 0.04371 0.03695 -0.0623 0.5672 0.7179 5.500 0.5992 0.04963 0.04289 -0.0568 0.5532 0.7197 5.750 0.6357 0.04921 0.04251 -0.0575 0.5525 0.7216 6.000 0.6703 0.04900 0.04236 -0.0582 0.5519 0.7239 7.500 0.6922 0.05978 0.05345 -0.0520 0.5114 0.7417 7.750 0.7227 0.05975 0.05350 -0.0525 0.5094 0.7465 8.000 0.7560 0.05947 0.05329 -0.0531 0.5081 0.7519 8.250 0.7448 0.06260 0.05651 -0.0517 0.4972 0.7561 8.500 0.7741 0.06250 0.05651 -0.0521 0.4947 0.7630 8.750 0.8101 0.06179 0.05591 -0.0526 0.4932 0.7709 9.000 0.8479 0.06091 0.05515 -0.0533 0.4921 0.7807 9.250 0.8870 0.05992 0.05433 -0.0541 0.4913 0.7928 9.500 0.9296 0.05871 0.05330 -0.0552 0.4908 0.8093 10.500 1.2854 0.03456 0.02999 -0.0690 0.5053 1.0000 10.750 1.0640 0.05808 0.05343 -0.0590 0.4633 1.0000 11.000 1.1172 0.05460 0.05003 -0.0596 0.4629 1.0000 11.250 1.1821 0.04942 0.04497 -0.0604 0.4632 1.0000 11.500 1.2718 0.04165 0.03732 -0.0621 0.4634 1.0000 11.750 1.2679 0.04391 0.03964 -0.0609 0.4503 1.0000 12.000 1.2787 0.04481 0.04060 -0.0603 0.4372 1.0000 12.250 1.3035 0.04427 0.04010 -0.0601 0.4192 1.0000 12.500 1.3078 0.04600 0.04184 -0.0592 0.3966 1.0000 13.000 1.3402 0.04595 0.04103 -0.0568 0.3024 1.0000 13.250 1.3224 0.04990 0.04473 -0.0551 0.2664 1.0000 13.500 1.3035 0.05421 0.04881 -0.0537 0.2334 1.0000 13.750 1.2836 0.05883 0.05319 -0.0526 0.1997 1.0000 14.000 1.2651 0.06348 0.05761 -0.0516 0.1656 1.0000 14.250 1.2470 0.06825 0.06213 -0.0509 0.1323 1.0000 14.500 1.2300 0.07298 0.06661 -0.0504 0.1008 1.0000 14.750 1.2137 0.07777 0.07114 -0.0499 0.0742 1.0000 15.000 1.2024 0.08213 0.07537 -0.0496 0.0587 1.0000 15.250 1.1946 0.08615 0.07932 -0.0495 0.0484 1.0000 15.500 1.1899 0.08989 0.08305 -0.0494 0.0428 1.0000 15.750 1.1856 0.09359 0.08679 -0.0494 0.0383 1.0000