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 = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4521 0.07070 0.06170 0.0483 0.9999 0.6477 -2.750 -0.4237 0.06898 0.05975 0.0397 0.9999 0.6007 -2.500 -0.3892 0.06735 0.05788 0.0310 0.9999 0.5587 -2.250 -0.3549 0.06511 0.05550 0.0253 0.9999 0.5289 -2.000 -0.3207 0.06324 0.05346 0.0202 0.9999 0.5079 -1.750 -0.2850 0.06168 0.05171 0.0153 0.9999 0.4922 -1.500 -0.2508 0.05958 0.04956 0.0123 0.9999 0.4822 -1.250 -0.2098 0.05812 0.04798 0.0076 0.9999 0.4697 -1.000 -0.1687 0.05632 0.04623 0.0037 0.9999 0.4612 -0.750 -0.1247 0.05502 0.04497 -0.0011 0.9999 0.4548 -0.500 0.0436 0.05004 0.03974 -0.0298 0.8426 0.4698 -0.250 0.0980 0.04796 0.03747 -0.0351 0.7195 0.4852 0.000 0.2673 0.04152 0.02942 -0.0568 0.6225 0.5628 0.250 0.4942 0.03896 0.02550 -0.0967 0.5742 1.0001 0.500 0.5545 0.04042 0.02589 -0.1035 0.5625 1.0001 0.750 0.6026 0.04170 0.02651 -0.1079 0.5552 1.0001 1.000 0.6428 0.04291 0.02735 -0.1108 0.5498 1.0001 1.250 0.6810 0.04423 0.02835 -0.1133 0.5446 1.0001 1.500 0.7180 0.04566 0.02950 -0.1156 0.5399 1.0001 1.750 0.7536 0.04723 0.03080 -0.1176 0.5354 1.0001 2.000 0.7780 0.04874 0.03233 -0.1178 0.5321 1.0001 2.250 0.8008 0.05041 0.03403 -0.1178 0.5297 1.0001 2.500 0.8210 0.05224 0.03592 -0.1175 0.5277 1.0001 2.750 0.8387 0.05425 0.03800 -0.1170 0.5274 1.0001 3.000 0.8535 0.05651 0.04036 -0.1162 0.5284 1.0001 3.250 0.8665 0.05904 0.04299 -0.1155 0.5306 1.0001 3.500 0.8765 0.06188 0.04591 -0.1146 0.5332 1.0001 3.750 0.8869 0.06493 0.04903 -0.1140 0.5360 1.0001 4.000 0.8283 0.07109 0.05559 -0.1067 0.5451 1.0001 4.250 0.7989 0.07691 0.06149 -0.1036 0.5536 1.0001 4.500 0.8043 0.08099 0.06553 -0.1036 0.5593 1.0001 5.000 0.6711 0.09813 0.08271 -0.0975 0.5993 1.0001