XFOIL Version 6.96 Calculated polar for: WORTMANN FX 62-K-131 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 ------ -------- --------- --------- -------- -------- -------- -11.750 -0.1813 0.11879 0.11558 -0.0704 0.9776 0.0276 -11.500 -0.1691 0.11519 0.11197 -0.0721 0.9761 0.0287 -11.250 -0.1598 0.11148 0.10825 -0.0744 0.9749 0.0299 -11.000 -0.1508 0.10751 0.10427 -0.0771 0.9738 0.0311 -10.750 -0.1481 0.10420 0.10097 -0.0783 0.9693 0.0325 -10.500 -0.1429 0.10022 0.09699 -0.0808 0.9668 0.0342 -10.250 -0.1408 0.09586 0.09264 -0.0849 0.9647 0.0364 -9.250 -0.1242 0.07777 0.07460 -0.0945 0.9551 0.0406 -9.000 -0.1148 0.07379 0.07061 -0.0966 0.9528 0.0424 -8.750 -0.1079 0.06909 0.06591 -0.0999 0.9511 0.0442 -8.500 -0.1023 0.06374 0.06056 -0.1046 0.9495 0.0462 -8.250 -0.1046 0.05861 0.05545 -0.1088 0.9458 0.0477 -7.500 -0.1423 0.04467 0.04138 -0.1207 0.9255 0.0489 -7.250 -0.1562 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0.7713 0.8724 3.000 0.8645 0.01794 0.01215 -0.1325 0.7682 0.8757 3.250 0.8852 0.01807 0.01239 -0.1316 0.7596 0.8801 3.500 0.9235 0.01775 0.01209 -0.1337 0.7556 0.8834 3.750 0.9448 0.01770 0.01213 -0.1326 0.7479 0.8871 4.000 0.9744 0.01736 0.01187 -0.1327 0.7425 0.8906 4.250 1.0023 0.01726 0.01186 -0.1329 0.7353 0.8942 4.500 1.0371 0.01703 0.01170 -0.1344 0.7289 0.8977 4.750 1.0588 0.01685 0.01162 -0.1332 0.7207 0.9013 5.000 1.0894 0.01643 0.01129 -0.1335 0.7137 0.9046 5.250 1.1164 0.01603 0.01098 -0.1331 0.7011 0.9084 5.500 1.1460 0.01554 0.01053 -0.1332 0.6862 0.9119 5.750 1.1669 0.01523 0.01031 -0.1316 0.6721 0.9155 6.000 1.1912 0.01504 0.01020 -0.1308 0.6572 0.9191 6.250 1.2134 0.01501 0.01030 -0.1298 0.6369 0.9232 6.500 1.2332 0.01489 0.01018 -0.1281 0.6098 0.9270 6.750 1.2498 0.01494 0.01019 -0.1258 0.5790 0.9310 7.000 1.2660 0.01525 0.01038 -0.1237 0.5410 0.9350 7.250 1.2758 0.01578 0.01074 -0.1206 0.4955 0.9394 7.500 1.2781 0.01645 0.01122 -0.1162 0.4476 0.9447 7.750 1.2823 0.01749 0.01205 -0.1127 0.3940 0.9497 8.000 1.2791 0.01873 0.01303 -0.1082 0.3400 0.9554 8.250 1.2724 0.02048 0.01441 -0.1038 0.2742 0.9619 8.500 1.2627 0.02246 0.01598 -0.0993 0.2048 0.9690 8.750 1.2550 0.02468 0.01778 -0.0956 0.1424 0.9773 9.000 1.2498 0.02684 0.01965 -0.0924 0.0920 0.9998 9.250 1.2484 0.02959 0.02204 -0.0904 0.0461 0.9999 9.500 1.2432 0.03283 0.02516 -0.0877 0.0240 0.9999 9.750 1.2431 0.03567 0.02814 -0.0858 0.0194 0.9999 10.000 1.2486 0.03803 0.03065 -0.0844 0.0171 0.9999 10.250 1.2536 0.04044 0.03318 -0.0832 0.0158 0.9999 10.500 1.2575 0.04298 0.03581 -0.0818 0.0149 0.9999 10.750 1.2604 0.04571 0.03863 -0.0800 0.0142 0.9999 11.000 1.2699 0.04832 0.04133 -0.0778 0.0136 0.9999 11.250 1.2910 0.05054 0.04369 -0.0764 0.0133 0.9999 11.500 1.3118 0.05293 0.04630 -0.0754 0.0132 0.9999 11.750 1.3239 0.05540 0.04904 -0.0745 0.0129 0.9999 12.000 1.3313 0.05795 0.05185 -0.0736 0.0123 0.9999 12.250 1.3367 0.06085 0.05502 -0.0727 0.0118 0.9999 12.500 1.3385 0.06386 0.05832 -0.0718 0.0110 0.9999 12.750 1.3374 0.06769 0.06245 -0.0708 0.0109 0.9999 13.000 1.3313 0.07214 0.06723 -0.0698 0.0110 0.9999 13.250 1.3212 0.07693 0.07233 -0.0691 0.0111 0.9999 13.500 1.3082 0.08206 0.07777 -0.0688 0.0113 0.9999 13.750 1.2926 0.08767 0.08367 -0.0691 0.0114 0.9999 14.000 1.2759 0.09360 0.08987 -0.0701 0.0115 0.9999 14.250 1.2581 0.10004 0.09656 -0.0719 0.0117 0.9999 14.500 1.2405 0.10680 0.10356 -0.0746 0.0118 0.9999 14.750 1.2226 0.11411 0.11109 -0.0781 0.0119 0.9999 15.000 1.2043 0.12213 0.11933 -0.0826 0.0121 0.9999 15.250 1.1867 0.13049 0.12787 -0.0879 0.0123 0.9999 15.500 1.1692 0.13963 0.13718 -0.0942 0.0124 0.9999