XFOIL Version 6.96 Calculated polar for: FX 61-168 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.750 -0.3495 0.09962 0.09534 -0.0486 1.0000 0.1500 -9.500 -0.3513 0.09672 0.09253 -0.0475 1.0000 0.1587 -9.250 -0.3670 0.09390 0.08986 -0.0475 1.0000 0.1657 -9.000 -0.4139 0.09152 0.08776 -0.0465 1.0000 0.1672 -8.750 -0.4071 0.09190 0.08825 -0.0407 0.9983 0.1734 -7.750 -0.4693 0.04915 0.04337 -0.0866 0.9252 0.0830 -7.500 -0.4335 0.04337 0.03668 -0.0891 0.9199 0.0647 -7.250 -0.4083 0.03960 0.03256 -0.0901 0.9098 0.0626 -7.000 -0.3681 0.03539 0.02778 -0.0928 0.9058 0.0586 -6.750 -0.3368 0.03273 0.02463 -0.0934 0.8971 0.0574 -6.500 -0.2926 0.03023 0.02188 -0.0959 0.8926 0.0583 -6.250 -0.2501 0.02847 0.01995 -0.0983 0.8876 0.0635 -6.000 -0.2159 0.02720 0.01847 -0.0987 0.8793 0.0665 -5.750 -0.1740 0.02514 0.01667 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0.5583 0.03458 0.02691 -0.0403 0.6300 0.8920 4.250 0.5284 0.03691 0.02932 -0.0345 0.6170 0.8970 4.500 0.5797 0.03594 0.02836 -0.0372 0.6148 0.9005 4.750 0.6409 0.03484 0.02725 -0.0413 0.6132 0.9040 5.000 0.5546 0.04071 0.03326 -0.0319 0.5900 0.9107 5.250 0.5947 0.03993 0.03251 -0.0329 0.5861 0.9146 5.500 0.6482 0.03849 0.03112 -0.0351 0.5843 0.9184 5.750 0.7058 0.03696 0.02961 -0.0379 0.5829 0.9218 6.000 0.6237 0.04417 0.03694 -0.0315 0.5584 0.9294 6.250 0.6674 0.04315 0.03597 -0.0326 0.5552 0.9338 6.500 0.6869 0.04370 0.03661 -0.0322 0.5486 0.9392 6.750 0.7572 0.04053 0.03351 -0.0344 0.5503 0.9432 7.000 0.8447 0.03697 0.03002 -0.0393 0.5517 0.9459 7.250 0.8017 0.04107 0.03424 -0.0341 0.5349 0.9555 7.500 0.8914 0.03706 0.03033 -0.0387 0.5361 0.9589 7.750 0.8034 0.04585 0.03924 -0.0339 0.5098 0.9746 8.000 0.8694 0.04275 0.03629 -0.0359 0.5091 0.9856 8.250 0.9541 0.03851 0.03217 -0.0395 0.5087 1.0000 8.500 1.0564 0.03412 0.02788 -0.0462 0.5070 1.0000 8.750 1.0474 0.03614 0.03005 -0.0437 0.4946 1.0000 9.000 1.0550 0.03758 0.03163 -0.0431 0.4835 1.0000 9.250 1.1563 0.03299 0.02714 -0.0496 0.4789 1.0000 9.500 1.1733 0.03340 0.02771 -0.0492 0.4674 1.0000 9.750 1.2195 0.03229 0.02672 -0.0514 0.4571 1.0000 10.000 1.2719 0.03078 0.02528 -0.0543 0.4449 1.0000 10.250 1.3048 0.03017 0.02473 -0.0551 0.4303 1.0000 10.500 1.3286 0.03006 0.02469 -0.0549 0.4144 1.0000 10.750 1.3457 0.03026 0.02496 -0.0541 0.3969 1.0000 11.000 1.3615 0.03057 0.02532 -0.0533 0.3780 1.0000 11.250 1.3791 0.03088 0.02556 -0.0526 0.3579 1.0000 11.500 1.3795 0.03233 0.02710 -0.0509 0.3394 1.0000 11.750 1.3809 0.03395 0.02878 -0.0495 0.3207 1.0000 12.000 1.3825 0.03569 0.03051 -0.0484 0.3017 1.0000 12.250 1.3833 0.03767 0.03243 -0.0475 0.2828 1.0000 12.500 1.3783 0.04031 0.03511 -0.0468 0.2645 1.0000 12.750 1.3733 0.04310 0.03790 -0.0462 0.2462 1.0000 13.000 1.3685 0.04600 0.04072 -0.0457 0.2283 1.0000 13.250 1.3632 0.04907 0.04368 -0.0454 0.2108 1.0000 13.500 1.3545 0.05267 0.04732 -0.0453 0.1937 1.0000 13.750 1.3460 0.05637 0.05097 -0.0454 0.1770 1.0000 14.000 1.3373 0.06024 0.05478 -0.0456 0.1607 1.0000 14.250 1.3286 0.06426 0.05874 -0.0459 0.1457 1.0000 14.500 1.3212 0.06828 0.06270 -0.0464 0.1318 1.0000 14.750 1.3158 0.07222 0.06659 -0.0469 0.1193 1.0000 15.000 1.3112 0.07615 0.07049 -0.0476 0.1084 1.0000 15.250 1.3076 0.07998 0.07425 -0.0485 0.0990 1.0000 15.500 1.3017 0.08436 0.07871 -0.0498 0.0909 1.0000 15.750 1.2986 0.08841 0.08282 -0.0509 0.0833 1.0000 16.000 1.2965 0.09222 0.08655 -0.0520 0.0763 1.0000 16.250 1.2902 0.09703 0.09157 -0.0538 0.0702 1.0000 16.500 1.2887 0.10081 0.09521 -0.0551 0.0637 1.0000