XFOIL Version 6.94 Calculated polar for: manu04 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4205 0.06703 0.05722 0.0306 0.9999 0.5621 -2.750 -0.3913 0.06457 0.05453 0.0252 0.9999 0.5361 -2.500 -0.3612 0.06243 0.05215 0.0202 0.9999 0.5159 -2.250 -0.3287 0.06052 0.04996 0.0154 0.9999 0.4976 -2.000 -0.2936 0.05879 0.04791 0.0106 0.9999 0.4798 -1.750 -0.2606 0.05668 0.04563 0.0078 0.9999 0.4681 -1.500 -0.2230 0.05523 0.04381 0.0038 0.9999 0.4574 -1.250 -0.1868 0.05353 0.04195 0.0009 0.9999 0.4524 -1.000 -0.1485 0.05207 0.04034 -0.0020 0.9999 0.4543 -0.750 -0.1072 0.05072 0.03892 -0.0053 0.9999 0.4609 -0.500 -0.0631 0.04920 0.03750 -0.0089 0.9999 0.4707 -0.250 -0.0141 0.04824 0.03662 -0.0140 0.9999 0.4805 0.000 0.1568 0.04270 0.03153 -0.0421 0.8219 0.5607 0.250 0.4299 0.03596 0.02310 -0.0857 0.6283 1.0001 0.500 0.5020 0.03696 0.02245 -0.0937 0.6001 1.0001 0.750 0.5516 0.03815 0.02277 -0.0980 0.5846 1.0001 1.000 0.5963 0.03936 0.02339 -0.1015 0.5739 1.0001 1.250 0.6364 0.04057 0.02425 -0.1042 0.5666 1.0001 1.500 0.6718 0.04178 0.02530 -0.1063 0.5612 1.0001 1.750 0.7066 0.04313 0.02648 -0.1082 0.5559 1.0001 2.000 0.7417 0.04463 0.02780 -0.1101 0.5507 1.0001 2.250 0.7707 0.04617 0.02929 -0.1111 0.5468 1.0001 2.500 0.7934 0.04781 0.03104 -0.1112 0.5435 1.0001 2.750 0.8144 0.04960 0.03292 -0.1111 0.5408 1.0001 3.000 0.8331 0.05158 0.03500 -0.1107 0.5390 1.0001 3.250 0.8485 0.05376 0.03730 -0.1101 0.5391 1.0001 3.500 0.8597 0.05627 0.03996 -0.1092 0.5406 1.0001 3.750 0.8682 0.05914 0.04298 -0.1083 0.5432 1.0001 4.000 0.8745 0.06236 0.04630 -0.1073 0.5462 1.0001 4.250 0.8681 0.06627 0.05041 -0.1054 0.5507 1.0001 4.500 0.8097 0.07336 0.05766 -0.1000 0.5610 1.0001 4.750 0.8010 0.07817 0.06244 -0.0991 0.5683 1.0001