XFOIL Version 6.96 Calculated polar for: FX 67-K-170/17 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 = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -11.250 -0.1032 0.09548 0.09074 -0.1043 0.7540 0.0138 -11.000 -0.1028 0.09149 0.08679 -0.1060 0.7522 0.0137 -10.750 -0.1018 0.08783 0.08317 -0.1076 0.7505 0.0136 -10.500 -0.1021 0.08402 0.07939 -0.1094 0.7488 0.0134 -10.250 -0.1029 0.08020 0.07560 -0.1112 0.7471 0.0133 -10.000 -0.1076 0.07549 0.07092 -0.1136 0.7454 0.0131 -9.750 -0.1166 0.06974 0.06520 -0.1170 0.7438 0.0130 -9.500 -0.1307 0.06276 0.05819 -0.1223 0.7422 0.0127 -9.250 -0.1506 0.05738 0.05275 -0.1251 0.7404 0.0125 -9.000 -0.1707 0.05339 0.04871 -0.1254 0.7376 0.0123 -8.750 -0.1931 0.05025 0.04545 -0.1232 0.7349 0.0124 -8.500 -0.2064 0.04701 0.04202 -0.1211 0.7327 0.0123 -8.250 -0.2154 0.04341 0.03814 -0.1189 0.7308 0.0124 -8.000 -0.2207 0.03948 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0.6571 0.8855 2.250 0.6288 0.01828 0.01140 -0.0814 0.6482 0.8977 2.500 0.6528 0.01834 0.01147 -0.0811 0.6465 0.9019 2.750 0.6796 0.01835 0.01146 -0.0812 0.6452 0.9064 3.000 0.7088 0.01829 0.01139 -0.0817 0.6442 0.9108 3.250 0.7392 0.01818 0.01127 -0.0824 0.6433 0.9143 3.750 0.7357 0.02004 0.01330 -0.0744 0.6320 0.9330 4.000 0.7670 0.01995 0.01322 -0.0754 0.6310 0.9379 4.250 0.8006 0.01983 0.01311 -0.0767 0.6301 0.9424 4.500 0.8364 0.01970 0.01301 -0.0784 0.6294 0.9466 4.750 0.8739 0.01953 0.01286 -0.0805 0.6287 0.9509 5.250 0.8926 0.02192 0.01542 -0.0787 0.6162 1.0000 5.500 0.9298 0.02139 0.01491 -0.0801 0.6154 1.0000 5.750 0.9774 0.02026 0.01380 -0.0828 0.6143 1.0000 6.000 0.9437 0.02317 0.01675 -0.0763 0.6022 1.0000 6.250 0.9940 0.02157 0.01517 -0.0788 0.6010 1.0000 6.750 1.0108 0.02333 0.01700 -0.0753 0.5876 1.0000 7.000 1.0482 0.02267 0.01639 -0.0766 0.5864 1.0000 7.500 1.0436 0.02595 0.01979 -0.0711 0.5694 1.0000 8.000 1.0855 0.02638 0.02033 -0.0702 0.5591 1.0000 8.250 1.0785 0.02847 0.02248 -0.0672 0.5479 1.0000 8.500 1.1135 0.02771 0.02179 -0.0680 0.5448 1.0000 9.000 1.1254 0.03021 0.02443 -0.0645 0.5247 1.0000 9.250 1.1575 0.02944 0.02369 -0.0649 0.5150 1.0000 9.500 1.1663 0.03047 0.02477 -0.0633 0.5018 1.0000 9.750 1.1822 0.03095 0.02530 -0.0624 0.4874 1.0000 10.000 1.1974 0.03145 0.02579 -0.0614 0.4670 1.0000 10.250 1.2210 0.03123 0.02547 -0.0608 0.4403 1.0000 10.500 1.2312 0.03185 0.02579 -0.0590 0.3977 1.0000 10.750 1.2277 0.03372 0.02740 -0.0564 0.3594 1.0000 11.000 1.2250 0.03575 0.02929 -0.0541 0.3300 1.0000 11.250 1.2210 0.03801 0.03141 -0.0520 0.3005 1.0000 11.500 1.2178 0.04034 0.03362 -0.0501 0.2728 1.0000 11.750 1.2138 0.04282 0.03597 -0.0483 0.2445 1.0000 12.000 1.2099 0.04541 0.03842 -0.0466 0.2162 1.0000 12.250 1.2032 0.04833 0.04116 -0.0449 0.1839 1.0000 12.500 1.1977 0.05123 0.04389 -0.0435 0.1519 1.0000 12.750 1.1929 0.05416 0.04664 -0.0422 0.1204 1.0000 13.250 1.1873 0.05987 0.05204 -0.0402 0.0702 1.0000 13.500 1.1886 0.06242 0.05454 -0.0395 0.0578 1.0000 13.750 1.1926 0.06476 0.05690 -0.0390 0.0501 1.0000 14.000 1.1962 0.06717 0.05931 -0.0386 0.0439 1.0000 14.250 1.2005 0.06956 0.06172 -0.0383 0.0387 1.0000 14.500 1.2044 0.07202 0.06422 -0.0380 0.0346 1.0000 14.750 1.2106 0.07426 0.06655 -0.0378 0.0312 1.0000 15.000 1.2145 0.07683 0.06919 -0.0377 0.0287 1.0000 15.250 1.2204 0.07918 0.07164 -0.0377 0.0257 1.0000 15.500 1.2281 0.08133 0.07391 -0.0377 0.0200 1.0000 15.750 1.2334 0.08380 0.07642 -0.0378 0.0147 1.0000 16.000 1.2352 0.08673 0.07938 -0.0379 0.0117 1.0000 16.250 1.2364 0.08981 0.08251 -0.0380 0.0102 1.0000 16.500 1.2370 0.09302 0.08583 -0.0383 0.0091 1.0000 16.750 1.2369 0.09637 0.08926 -0.0386 0.0083 1.0000 17.000 1.2365 0.09979 0.09277 -0.0391 0.0077 1.0000 17.250 1.2370 0.10314 0.09627 -0.0396 0.0070 1.0000 17.500 1.2377 0.10648 0.09972 -0.0403 0.0065 1.0000 17.750 1.2378 0.10996 0.10334 -0.0411 0.0062 1.0000 18.000 1.2364 0.11370 0.10717 -0.0420 0.0058 1.0000 18.250 1.2349 0.11751 0.11111 -0.0431 0.0057 1.0000