XFOIL Version 6.94 Calculated polar for: kut1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4217 0.03113 0.01797 0.0280 1.0000 0.1899 -2.750 -0.4166 0.03147 0.01816 0.0321 1.0000 0.2107 -2.500 -0.4039 0.03159 0.01814 0.0347 1.0000 0.2262 -2.250 -0.3997 0.03174 0.01840 0.0386 1.0000 0.2573 -2.000 -0.3944 0.03172 0.01860 0.0422 1.0000 0.2976 -1.500 -0.3610 0.03133 0.01826 0.0453 1.0000 0.3594 -1.250 -0.3263 0.03110 0.01763 0.0435 1.0000 0.3579 -1.000 -0.2919 0.03091 0.01710 0.0416 1.0000 0.3567 -0.750 -0.2579 0.03075 0.01672 0.0398 1.0000 0.3562 -0.500 -0.2236 0.03060 0.01647 0.0380 1.0000 0.3564 -0.250 -0.1896 0.03047 0.01635 0.0362 1.0000 0.3572 0.000 -0.1562 0.03036 0.01634 0.0345 1.0000 0.3588 0.250 -0.1234 0.03028 0.01646 0.0330 1.0000 0.3611 0.500 -0.0916 0.03021 0.01667 0.0317 1.0000 0.3642 0.750 0.2463 0.02437 0.01249 -0.0179 0.4853 0.4191 1.750 0.4155 0.02815 0.01559 -0.0284 0.0951 1.0000 2.000 0.4383 0.02890 0.01640 -0.0267 0.0902 1.0000 2.250 0.4615 0.02968 0.01729 -0.0250 0.0875 1.0000 2.500 0.4855 0.03058 0.01832 -0.0235 0.0863 1.0000 2.750 0.5098 0.03144 0.01928 -0.0220 0.0869 1.0000 3.000 0.5346 0.03224 0.02018 -0.0203 0.0895 1.0000 3.250 0.5607 0.03312 0.02114 -0.0188 0.0934 1.0000 3.500 0.5927 0.03440 0.02241 -0.0178 0.0979 1.0000 3.750 0.6234 0.03463 0.02298 -0.0162 0.1045 1.0000 4.000 0.6661 0.03601 0.02452 -0.0162 0.1129 1.0000 4.250 0.6995 0.03643 0.02566 -0.0145 0.1227 1.0000 4.500 0.7385 0.03836 0.02835 -0.0137 0.1342 1.0000 4.750 0.7710 0.04263 0.03365 -0.0128 0.1468 1.0000 5.000 0.7913 0.04815 0.04004 -0.0115 0.1587 1.0000