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 = 0.100 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -13.500 -0.2661 0.10210 0.09676 -0.0922 0.8289 0.0313 -13.250 -0.3083 0.08695 0.08157 -0.1005 0.8256 0.0304 -13.000 -0.3815 0.07129 0.06559 -0.1106 0.8216 0.0287 -12.750 -0.4087 0.06522 0.05924 -0.1139 0.8129 0.0286 -12.500 -0.4346 0.05989 0.05361 -0.1160 0.8040 0.0284 -12.250 -0.4449 0.05662 0.05013 -0.1170 0.7952 0.0286 -12.000 -0.4545 0.05351 0.04679 -0.1174 0.7868 0.0290 -11.750 -0.4626 0.05063 0.04368 -0.1175 0.7788 0.0293 -11.500 -0.4686 0.04793 0.04072 -0.1171 0.7715 0.0295 -11.250 -0.4713 0.04549 0.03801 -0.1165 0.7645 0.0298 -11.000 -0.4708 0.04326 0.03552 -0.1156 0.7579 0.0303 -10.750 -0.4663 0.04124 0.03321 -0.1147 0.7525 0.0307 -10.500 -0.4587 0.03931 0.03103 -0.1137 0.7462 0.0315 -10.250 -0.4465 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0.9248 0.07131 0.06435 -0.0512 0.4346 1.0000 11.000 0.9566 0.07057 0.06371 -0.0513 0.4324 1.0000 11.250 0.9523 0.07349 0.06671 -0.0508 0.4203 1.0000 11.500 0.9829 0.07276 0.06611 -0.0508 0.4177 1.0000 12.500 1.0307 0.07702 0.07079 -0.0495 0.3798 1.0000 12.750 1.0636 0.07530 0.06921 -0.0491 0.3736 1.0000 13.250 1.0861 0.07767 0.07178 -0.0485 0.3452 1.0000 13.500 1.1119 0.07687 0.07100 -0.0481 0.3274 1.0000 13.750 1.1287 0.07730 0.07140 -0.0476 0.3024 1.0000 14.000 1.1499 0.07699 0.07091 -0.0470 0.2685 1.0000 14.250 1.1580 0.07846 0.07213 -0.0464 0.2328 1.0000 14.500 1.1560 0.08139 0.07479 -0.0460 0.1986 1.0000 14.750 1.1496 0.08507 0.07824 -0.0459 0.1658 1.0000 15.000 1.1421 0.08903 0.08199 -0.0460 0.1379 1.0000 15.250 1.1345 0.09316 0.08590 -0.0463 0.1113 1.0000 15.500 1.1294 0.09705 0.08967 -0.0467 0.0896 1.0000 15.750 1.1247 0.10095 0.09343 -0.0472 0.0737 1.0000