XFOIL Version 6.96 Calculated polar for: FX 66-17AII-182 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 ------ -------- --------- --------- -------- -------- -------- -12.250 -0.2701 0.09676 0.09176 -0.0802 0.8432 0.0285 -12.000 -0.3759 0.07056 0.06533 -0.0963 0.8545 0.0252 -11.750 -0.3921 0.06404 0.05846 -0.1022 0.8266 0.0251 -11.500 -0.4066 0.05929 0.05339 -0.1054 0.8042 0.0250 -11.250 -0.4252 0.05507 0.04881 -0.1072 0.7862 0.0251 -10.750 -0.4460 0.04904 0.04229 -0.1078 0.7577 0.0257 -10.500 -0.4501 0.04685 0.03989 -0.1071 0.7463 0.0260 -10.250 -0.4494 0.04528 0.03818 -0.1060 0.7365 0.0267 -10.000 -0.4491 0.04374 0.03643 -0.1044 0.7276 0.0274 -9.750 -0.4473 0.04170 0.03409 -0.1029 0.7203 0.0282 -9.500 -0.4417 0.03958 0.03164 -0.1017 0.7131 0.0289 -9.250 -0.4320 0.03740 0.02906 -0.1005 0.7073 0.0296 -9.000 -0.4191 0.03535 0.02663 -0.0993 0.7007 0.0306 -8.750 -0.4022 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0.01299 -0.0723 0.5706 0.6949 1.000 0.4651 0.02315 0.01310 -0.0722 0.5685 0.6980 1.250 0.4920 0.02326 0.01315 -0.0724 0.5667 0.7014 1.500 0.5200 0.02335 0.01317 -0.0727 0.5650 0.7051 1.750 0.5389 0.02377 0.01365 -0.0718 0.5616 0.7088 2.000 0.5592 0.02412 0.01407 -0.0710 0.5583 0.7117 2.250 0.5824 0.02440 0.01438 -0.0707 0.5556 0.7148 2.500 0.6074 0.02463 0.01462 -0.0706 0.5535 0.7184 2.750 0.6341 0.02482 0.01478 -0.0708 0.5516 0.7225 3.000 0.6619 0.02495 0.01489 -0.0711 0.5500 0.7258 3.250 0.6807 0.02541 0.01545 -0.0701 0.5470 0.7292 3.500 0.6938 0.02612 0.01628 -0.0685 0.5430 0.7328 3.750 0.7135 0.02660 0.01681 -0.0677 0.5401 0.7370 4.000 0.7370 0.02691 0.01717 -0.0675 0.5378 0.7410 4.250 0.7638 0.02709 0.01738 -0.0676 0.5359 0.7449 4.500 0.7929 0.02723 0.01752 -0.0681 0.5343 0.7492 4.750 0.7989 0.02836 0.01881 -0.0657 0.5299 0.7539 5.000 0.8063 0.02942 0.02001 -0.0635 0.5256 0.7581 5.250 0.8269 0.02988 0.02054 -0.0630 0.5228 0.7630 5.500 0.8533 0.03011 0.02082 -0.0631 0.5208 0.7688 5.750 0.8821 0.03025 0.02103 -0.0635 0.5193 0.7743 6.250 0.8554 0.03428 0.02537 -0.0548 0.5076 0.7881 6.500 0.8808 0.03451 0.02571 -0.0548 0.5056 0.7953 6.750 0.9111 0.03458 0.02588 -0.0554 0.5042 0.8040 7.000 0.9446 0.03453 0.02594 -0.0564 0.5031 0.8145 7.500 0.8912 0.04164 0.03345 -0.0488 0.4870 0.8577 8.000 0.8840 0.04772 0.03978 -0.0483 0.4725 1.0000 8.500 0.8813 0.05316 0.04529 -0.0470 0.4585 1.0000 9.000 0.8802 0.05848 0.05070 -0.0460 0.4447 1.0000 9.250 0.9076 0.05849 0.05079 -0.0460 0.4435 1.0000 9.750 0.9095 0.06353 0.05594 -0.0450 0.4297 1.0000 10.250 0.9171 0.06800 0.06057 -0.0443 0.4163 1.0000 10.500 0.9446 0.06779 0.06045 -0.0441 0.4146 1.0000 11.000 0.9552 0.07190 0.06472 -0.0435 0.4010 1.0000 11.500 0.9698 0.07552 0.06855 -0.0429 0.3875 1.0000 12.000 0.9880 0.07867 0.07191 -0.0423 0.3738 1.0000 12.500 1.0147 0.08043 0.07390 -0.0416 0.3595 1.0000 12.750 1.0304 0.08048 0.07408 -0.0410 0.3491 1.0000 13.000 1.0592 0.07841 0.07210 -0.0400 0.3376 1.0000 13.250 1.0668 0.07969 0.07346 -0.0397 0.3214 1.0000 13.500 1.0762 0.08109 0.07495 -0.0396 0.3069 1.0000 13.750 1.0826 0.08303 0.07697 -0.0395 0.2901 1.0000 14.000 1.1045 0.08263 0.07658 -0.0390 0.2687 1.0000 14.250 1.1279 0.08179 0.07550 -0.0381 0.2309 1.0000 14.500 1.1328 0.08358 0.07696 -0.0377 0.1930 1.0000 14.750 1.1263 0.08724 0.08038 -0.0378 0.1619 1.0000 15.000 1.1171 0.09146 0.08440 -0.0382 0.1346 1.0000 15.250 1.1078 0.09583 0.08861 -0.0387 0.1107 1.0000