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 = 0.022 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4702 0.06772 0.06008 0.0631 0.9999 0.7369 -2.750 -0.4612 0.06601 0.05831 0.0560 0.9999 0.6796 -2.500 -0.4216 0.06488 0.05696 0.0432 0.9999 0.6092 -2.250 -0.3680 0.06403 0.05576 0.0288 0.9999 0.5356 -2.000 -0.3217 0.06203 0.05359 0.0200 0.9999 0.4859 -1.750 -0.2759 0.06037 0.05174 0.0125 0.9999 0.4470 -1.500 -0.2357 0.05844 0.04971 0.0078 0.9999 0.4253 -1.000 -0.0232 0.04944 0.04026 -0.0238 0.8048 0.4002 -0.750 0.1309 0.04402 0.03219 -0.0441 0.5703 0.4026 -0.500 0.1991 0.04327 0.03030 -0.0519 0.5413 0.4225 -0.250 0.2726 0.04266 0.02885 -0.0612 0.5234 0.4507 0.000 0.3553 0.04153 0.02728 -0.0720 0.5100 0.5129 0.250 0.5215 0.03902 0.02555 -0.1010 0.4943 1.0001 0.500 0.5808 0.04046 0.02612 -0.1077 0.4874 1.0001 0.750 0.6236 0.04152 0.02678 -0.1109 0.4818 1.0001 1.000 0.6633 0.04264 0.02757 -0.1135 0.4766 1.0001 1.250 0.7013 0.04387 0.02852 -0.1157 0.4719 1.0001 1.500 0.7376 0.04514 0.02958 -0.1177 0.4692 1.0001 1.750 0.7720 0.04645 0.03079 -0.1194 0.4683 1.0001 2.000 0.8031 0.04779 0.03212 -0.1205 0.4684 1.0001 2.250 0.8300 0.04919 0.03360 -0.1210 0.4694 1.0001 2.500 0.8535 0.05070 0.03522 -0.1210 0.4708 1.0001 2.750 0.8738 0.05236 0.03701 -0.1206 0.4729 1.0001 3.000 0.8916 0.05422 0.03902 -0.1199 0.4754 1.0001 3.250 0.9075 0.05629 0.04123 -0.1190 0.4785 1.0001 3.500 0.9209 0.05860 0.04370 -0.1180 0.4819 1.0001 3.750 0.9349 0.06112 0.04630 -0.1171 0.4854 1.0001 4.000 0.9527 0.06380 0.04901 -0.1170 0.4887 1.0001 4.250 0.9144 0.06790 0.05353 -0.1103 0.4971 1.0001 4.500 0.8911 0.07263 0.05842 -0.1064 0.5056 1.0001 4.750 0.9017 0.07629 0.06205 -0.1064 0.5117 1.0001 5.000 0.7969 0.08611 0.07206 -0.0979 0.5329 1.0001