XFOIL Version 6.96 Calculated polar for: WORTMANN FX 08-S-176 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 -0.0931 0.10275 0.09713 -0.0855 0.6051 0.0162 -11.000 -0.0891 0.09940 0.09380 -0.0868 0.6042 0.0160 -10.750 -0.0918 0.09437 0.08881 -0.0893 0.6034 0.0164 -10.500 -0.0877 0.09167 0.08617 -0.0903 0.6024 0.0170 -10.250 -0.0752 0.09076 0.08526 -0.0905 0.6011 0.0183 -10.000 -0.0754 0.08687 0.08141 -0.0923 0.6003 0.0183 -9.750 -0.0738 0.08352 0.07809 -0.0940 0.5994 0.0193 -9.500 -0.0774 0.07911 0.07373 -0.0963 0.5986 0.0190 -9.250 -0.0855 0.07386 0.06854 -0.0994 0.5979 0.0185 -9.000 -0.0988 0.06782 0.06255 -0.1041 0.5971 0.0183 -8.750 -0.1231 0.06150 0.05621 -0.1070 0.5965 0.0178 -8.500 -0.1450 0.05737 0.05202 -0.1061 0.5958 0.0177 -8.250 -0.1691 0.05229 0.04677 -0.1040 0.5951 0.0173 -8.000 -0.1778 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1.0617 0.02231 0.01553 -0.0528 0.4990 0.9835 7.250 1.0809 0.02275 0.01606 -0.0527 0.4956 0.9881 7.500 1.1083 0.02276 0.01612 -0.0534 0.4935 0.9928 7.750 1.1383 0.02257 0.01599 -0.0542 0.4920 1.0000 8.250 1.1150 0.02615 0.01978 -0.0480 0.4794 1.0000 8.500 1.1478 0.02573 0.01941 -0.0488 0.4781 1.0000 8.750 1.0870 0.03271 0.02648 -0.0456 0.4606 1.0000 9.250 1.1303 0.03334 0.02721 -0.0461 0.4510 1.0000 9.750 1.1405 0.03666 0.03056 -0.0451 0.4262 1.0000 10.000 1.1698 0.03622 0.03016 -0.0455 0.4197 1.0000 10.250 1.1831 0.03722 0.03119 -0.0453 0.4109 1.0000 10.500 1.2117 0.03659 0.03045 -0.0454 0.3938 1.0000 10.750 1.2152 0.03823 0.03197 -0.0445 0.3721 1.0000 11.000 1.2221 0.03957 0.03320 -0.0438 0.3541 1.0000 11.250 1.2216 0.04160 0.03511 -0.0427 0.3349 1.0000 11.500 1.2255 0.04349 0.03702 -0.0422 0.3234 1.0000 11.750 1.2210 0.04621 0.03967 -0.0413 0.3021 1.0000 12.000 1.2154 0.04914 0.04251 -0.0406 0.2769 1.0000 12.250 1.1942 0.05363 0.04669 -0.0395 0.2295 1.0000 12.500 1.1606 0.05965 0.05233 -0.0383 0.1794 1.0000 12.750 1.1377 0.06490 0.05733 -0.0377 0.1357 1.0000 13.000 1.1139 0.07047 0.06261 -0.0374 0.0888 1.0000 13.250 1.0920 0.07606 0.06792 -0.0373 0.0449 1.0000 13.500 1.0844 0.08016 0.07190 -0.0375 0.0182 1.0000 13.750 1.0836 0.08354 0.07529 -0.0379 0.0135 1.0000 14.000 1.0863 0.08655 0.07836 -0.0383 0.0117 1.0000 14.250 1.0905 0.08941 0.08132 -0.0388 0.0110 1.0000 14.500 1.0941 0.09235 0.08437 -0.0393 0.0103 1.0000 14.750 1.0972 0.09543 0.08756 -0.0399 0.0098 1.0000 15.000 1.1002 0.09853 0.09079 -0.0405 0.0095 1.0000 15.250 1.1033 0.10167 0.09406 -0.0413 0.0093 1.0000 15.500 1.1067 0.10478 0.09729 -0.0420 0.0091 1.0000 15.750 1.1092 0.10804 0.10067 -0.0429 0.0089 1.0000 16.000 1.1117 0.11132 0.10407 -0.0438 0.0088 1.0000 16.250 1.1137 0.11471 0.10757 -0.0447 0.0086 1.0000 16.500 1.1157 0.11811 0.11108 -0.0458 0.0084 1.0000