XFOIL Version 6.96 Calculated polar for: WORTMANN FX 05-188 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.250 -0.6802 0.09838 0.09509 -0.0743 1.0000 0.0137 -17.000 -0.7046 0.08939 0.08594 -0.0787 0.9821 0.0137 -16.750 -0.7156 0.08031 0.07660 -0.0878 0.9481 0.0138 -16.500 -0.7217 0.07192 0.06793 -0.0969 0.9350 0.0138 -16.250 -0.7154 0.06564 0.06145 -0.1049 0.9234 0.0141 -16.000 -0.7043 0.05967 0.05518 -0.1135 0.9084 0.0141 -15.750 -0.6894 0.05542 0.05069 -0.1199 0.8905 0.0142 -15.500 -0.6808 0.05184 0.04687 -0.1239 0.8750 0.0142 -15.000 -0.6665 0.04673 0.04141 -0.1278 0.8540 0.0144 -14.500 -0.6532 0.04249 0.03689 -0.1296 0.8395 0.0147 -14.250 -0.6459 0.04061 0.03487 -0.1300 0.8330 0.0147 -13.750 -0.6281 0.03739 0.03140 -0.1304 0.8225 0.0150 -13.500 -0.6182 0.03597 0.02987 -0.1303 0.8175 0.0151 -13.250 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0.00528 -0.1095 0.4842 0.7080 6.000 1.0839 0.01151 0.00556 -0.1075 0.4600 0.7096 6.250 1.0938 0.01193 0.00588 -0.1044 0.4354 0.7114 6.500 1.1008 0.01247 0.00630 -0.1009 0.4084 0.7133 6.750 1.1042 0.01322 0.00691 -0.0969 0.3781 0.7152 7.000 1.1073 0.01405 0.00760 -0.0932 0.3477 0.7169 7.250 1.1044 0.01518 0.00854 -0.0888 0.3079 0.7187 7.500 1.0980 0.01660 0.00976 -0.0842 0.2638 0.7205 7.750 1.0941 0.01810 0.01108 -0.0804 0.2252 0.7222 8.000 1.0933 0.01959 0.01241 -0.0772 0.1904 0.7240 8.250 1.0960 0.02098 0.01369 -0.0747 0.1632 0.7259 8.500 1.0992 0.02241 0.01500 -0.0724 0.1367 0.7277 8.750 1.1012 0.02399 0.01643 -0.0700 0.1075 0.7294 9.000 1.1079 0.02533 0.01769 -0.0683 0.0903 0.7310 9.250 1.1169 0.02656 0.01888 -0.0668 0.0799 0.7324 9.500 1.1271 0.02775 0.02006 -0.0656 0.0723 0.7337 10.000 1.1500 0.03003 0.02236 -0.0635 0.0599 0.7359 10.250 1.1605 0.03126 0.02359 -0.0624 0.0518 0.7371 10.500 1.1691 0.03267 0.02498 -0.0612 0.0418 0.7384 10.750 1.1777 0.03412 0.02641 -0.0601 0.0348 0.7398 11.000 1.1866 0.03560 0.02790 -0.0591 0.0309 0.7412 11.250 1.1973 0.03696 0.02930 -0.0583 0.0288 0.7425 11.500 1.2068 0.03846 0.03083 -0.0575 0.0272 0.7437 11.750 1.2175 0.03990 0.03232 -0.0567 0.0260 0.7451 12.000 1.2282 0.04138 0.03386 -0.0561 0.0251 0.7464 12.250 1.2385 0.04292 0.03545 -0.0555 0.0243 0.7478 12.500 1.2483 0.04454 0.03712 -0.0549 0.0236 0.7491 12.750 1.2570 0.04630 0.03892 -0.0544 0.0230 0.7505 13.250 1.2727 0.05009 0.04284 -0.0533 0.0219 0.7531 13.500 1.2821 0.05188 0.04471 -0.0529 0.0216 0.7547 13.750 1.2905 0.05381 0.04672 -0.0526 0.0213 0.7564 14.000 1.2986 0.05580 0.04879 -0.0523 0.0209 0.7581 14.250 1.3061 0.05790 0.05097 -0.0521 0.0204 0.7599 14.500 1.3141 0.06001 0.05315 -0.0520 0.0200 0.7616 14.750 1.3207 0.06230 0.05551 -0.0519 0.0196 0.7634 15.000 1.3269 0.06468 0.05796 -0.0519 0.0192 0.7652 15.250 1.3319 0.06725 0.06059 -0.0520 0.0189 0.7669 15.500 1.3357 0.07000 0.06341 -0.0521 0.0186 0.7686 15.750 1.3381 0.07297 0.06646 -0.0523 0.0183 0.7701 16.000 1.3412 0.07591 0.06949 -0.0525 0.0181 0.7719 16.250 1.3474 0.07850 0.07218 -0.0529 0.0179 0.7739 16.500 1.3523 0.08128 0.07505 -0.0533 0.0177 0.7760 16.750 1.3572 0.08408 0.07795 -0.0538 0.0175 0.7782 17.000 1.3615 0.08704 0.08099 -0.0544 0.0173 0.7804 17.250 1.3660 0.08997 0.08401 -0.0550 0.0170 0.7827 17.500 1.3696 0.09308 0.08720 -0.0558 0.0167 0.7847 17.750 1.3737 0.09612 0.09034 -0.0566 0.0164 0.7868 18.000 1.3766 0.09936 0.09368 -0.0575 0.0162 0.7890 18.250 1.3797 0.10260 0.09701 -0.0585 0.0160 0.7914 18.500 1.3820 0.10598 0.10048 -0.0596 0.0158 0.7938 18.750 1.3843 0.10941 0.10400 -0.0609 0.0156 0.7964 19.000 1.3861 0.11292 0.10758 -0.0622 0.0154 0.7988 19.250 1.3872 0.11655 0.11129 -0.0636 0.0152 0.8010