XFOIL Version 6.94 Calculated polar for: manu01 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3927 0.06550 0.05877 0.0591 0.9999 0.8157 -2.750 -0.4499 0.06292 0.05631 0.0680 0.9999 0.7874 -2.500 -0.4691 0.06097 0.05434 0.0660 0.9999 0.7347 -2.250 -0.4180 0.06076 0.05393 0.0479 0.9999 0.6361 -2.000 -0.3434 0.06040 0.05326 0.0278 0.9999 0.5272 -1.750 -0.1097 0.05347 0.04532 -0.0209 0.8701 0.3772 -1.500 -0.0807 0.05039 0.04217 -0.0207 0.7749 0.3673 -1.250 0.0488 0.04610 0.03504 -0.0378 0.5198 0.3575 -1.000 0.0956 0.04512 0.03320 -0.0418 0.4995 0.3549 -0.750 0.1534 0.04430 0.03147 -0.0480 0.4846 0.3561 -0.500 0.2177 0.04337 0.02987 -0.0554 0.4728 0.3654 -0.250 0.2889 0.04267 0.02856 -0.0641 0.4638 0.3934 0.000 0.3653 0.04169 0.02728 -0.0738 0.4553 0.4383 0.250 0.4541 0.03946 0.02556 -0.0857 0.4456 0.6073 0.500 0.5902 0.03968 0.02604 -0.1092 0.4340 1.0001 0.750 0.6384 0.04077 0.02670 -0.1135 0.4304 1.0001 1.000 0.6820 0.04184 0.02747 -0.1169 0.4292 1.0001 1.250 0.7214 0.04294 0.02839 -0.1194 0.4287 1.0001 1.500 0.7573 0.04407 0.02942 -0.1213 0.4289 1.0001 1.750 0.7899 0.04527 0.03058 -0.1225 0.4296 1.0001 2.000 0.8196 0.04658 0.03189 -0.1233 0.4308 1.0001 2.250 0.8471 0.04798 0.03334 -0.1237 0.4325 1.0001 2.500 0.8726 0.04951 0.03491 -0.1238 0.4344 1.0001 2.750 0.8967 0.05118 0.03663 -0.1237 0.4367 1.0001 3.000 0.9194 0.05301 0.03850 -0.1235 0.4391 1.0001 3.250 0.9428 0.05502 0.04055 -0.1234 0.4415 1.0001 3.500 0.9507 0.05665 0.04251 -0.1210 0.4460 1.0001 3.750 0.9553 0.05891 0.04503 -0.1186 0.4516 1.0001 4.000 0.9611 0.06154 0.04782 -0.1164 0.4569 1.0001 4.250 0.9711 0.06434 0.05070 -0.1151 0.4614 1.0001 4.500 0.9882 0.06723 0.05360 -0.1147 0.4651 1.0001 4.750 0.9262 0.07236 0.05920 -0.1059 0.4777 1.0001 5.000 0.9198 0.07666 0.06354 -0.1039 0.4854 1.0001