XFOIL Version 6.96 Calculated polar for: FX 66-182 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 ------ -------- --------- --------- -------- -------- -------- -12.250 -0.2280 0.10379 0.10006 -0.0912 0.8236 0.0628 -12.000 -0.2198 0.09868 0.09480 -0.0922 0.7989 0.0639 -11.750 -0.2004 0.09757 0.09350 -0.0904 0.7773 0.0654 -11.500 -0.1907 0.09524 0.09104 -0.0906 0.7628 0.0671 -11.250 -0.1879 0.09195 0.08768 -0.0919 0.7524 0.0699 -11.000 -0.1395 0.08100 0.07681 -0.0881 0.7184 0.0787 -10.750 -0.1321 0.07828 0.07404 -0.0882 0.7128 0.0814 -10.500 -0.1395 0.07322 0.06895 -0.0904 0.7096 0.0841 -10.000 -0.3659 0.03720 0.03217 -0.1071 0.7160 0.0407 -9.750 -0.4241 0.04670 0.04118 -0.1084 0.7207 0.0434 -9.500 -0.4477 0.04033 0.03389 -0.1042 0.7158 0.0342 -9.250 -0.4409 0.03750 0.03071 -0.1028 0.7105 0.0343 -9.000 -0.4286 0.03478 0.02781 -0.1020 0.7047 0.0350 -8.750 -0.4106 0.03307 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0.9870 0.02705 0.02048 -0.0781 0.5305 0.7881 6.250 1.0056 0.02764 0.02118 -0.0772 0.5275 0.7935 6.500 1.0340 0.02778 0.02138 -0.0777 0.5255 0.7997 6.750 1.0644 0.02773 0.02140 -0.0782 0.5239 0.8058 7.000 1.0952 0.02781 0.02154 -0.0789 0.5225 0.8137 7.250 1.0458 0.03167 0.02575 -0.0699 0.5133 0.8234 7.500 1.0703 0.03184 0.02603 -0.0697 0.5110 0.8341 7.750 0.9996 0.03663 0.03103 -0.0591 0.5005 0.8542 8.000 1.0651 0.03455 0.02911 -0.0629 0.5018 0.8787 8.250 1.1653 0.03115 0.02585 -0.0717 0.5034 0.9608 8.500 1.0371 0.04027 0.03505 -0.0576 0.4886 1.0000 8.750 1.0923 0.03816 0.03298 -0.0596 0.4882 1.0000 9.000 1.1549 0.03545 0.03033 -0.0623 0.4865 1.0000 9.250 0.9707 0.05368 0.04850 -0.0529 0.4602 1.0000 9.500 1.0208 0.05106 0.04594 -0.0532 0.4605 1.0000 9.750 1.0739 0.04767 0.04264 -0.0535 0.4601 1.0000 10.000 1.1252 0.04418 0.03923 -0.0535 0.4583 1.0000 10.250 1.0766 0.05166 0.04672 -0.0516 0.4447 1.0000 10.500 1.1258 0.04801 0.04316 -0.0513 0.4415 1.0000 10.750 1.0514 0.05876 0.05389 -0.0496 0.4236 1.0000 11.000 1.4377 0.02326 0.01831 -0.0635 0.4237 1.0000 11.250 1.4392 0.02394 0.01907 -0.0604 0.4130 1.0000 11.500 1.4343 0.02511 0.02030 -0.0572 0.3991 1.0000 11.750 1.4260 0.02685 0.02207 -0.0543 0.3833 1.0000 12.000 1.4158 0.02905 0.02426 -0.0518 0.3635 1.0000 12.250 1.4018 0.03185 0.02701 -0.0496 0.3393 1.0000 12.500 1.3836 0.03520 0.03022 -0.0475 0.3107 1.0000 12.750 1.3599 0.03925 0.03410 -0.0455 0.2784 1.0000 13.000 1.3329 0.04381 0.03846 -0.0437 0.2493 1.0000 13.250 1.3081 0.04849 0.04298 -0.0424 0.2221 1.0000 13.500 1.2820 0.05358 0.04788 -0.0414 0.1929 1.0000 13.750 1.2572 0.05882 0.05293 -0.0408 0.1628 1.0000 14.000 1.2347 0.06407 0.05798 -0.0406 0.1332 1.0000 14.250 1.2119 0.06955 0.06322 -0.0405 0.1009 1.0000 14.500 1.1932 0.07477 0.06822 -0.0406 0.0746 1.0000 14.750 1.1766 0.07990 0.07318 -0.0409 0.0563 1.0000 15.000 1.1662 0.08442 0.07763 -0.0413 0.0471 1.0000 15.250 1.1622 0.08821 0.08145 -0.0417 0.0417 1.0000 15.500 1.1565 0.09227 0.08551 -0.0422 0.0385 1.0000