XFOIL Version 6.94 Calculated polar for: manu4/28 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.1841 0.06384 0.05442 -0.0407 0.4012 0.8969 -2.750 0.1316 0.06279 0.05354 -0.0295 0.4024 0.8694 -2.500 0.0776 0.06100 0.05194 -0.0187 0.4050 0.8413 -2.250 0.0057 0.05899 0.05015 -0.0044 0.4079 0.8183 -2.000 -0.0710 0.05668 0.04822 0.0107 0.4123 0.7992 -1.750 -0.0953 0.05488 0.04642 0.0132 0.4126 0.7606 -1.500 0.0237 0.05868 0.04916 -0.0215 0.4030 0.5765 -1.250 0.1549 0.05960 0.04872 -0.0522 0.3944 0.3823 -1.000 0.2230 0.05820 0.04660 -0.0622 0.3914 0.3383 -0.750 0.2774 0.05646 0.04463 -0.0681 0.3895 0.3324 -0.500 0.3333 0.05530 0.04315 -0.0743 0.3882 0.3297 -0.250 0.3926 0.05438 0.04188 -0.0811 0.3874 0.3266 0.000 0.4563 0.05385 0.04094 -0.0887 0.3871 0.3291 0.250 0.5176 0.05368 0.04033 -0.0958 0.3876 0.3387 0.500 0.5771 0.05353 0.04004 -0.1026 0.3889 0.3622 0.750 0.6443 0.05332 0.03988 -0.1112 0.3909 0.4089 1.000 0.7226 0.05267 0.03966 -0.1220 0.3938 0.5334 1.250 0.7641 0.05275 0.04040 -0.1253 0.3970 0.6404 1.500 0.8709 0.05459 0.04314 -0.1437 0.4017 1.0001 1.750 0.9010 0.05675 0.04516 -0.1450 0.4043 1.0001 2.000 0.9286 0.05905 0.04731 -0.1458 0.4063 1.0001 2.250 0.9558 0.06157 0.04965 -0.1465 0.4081 1.0001 2.500 0.9598 0.06265 0.05100 -0.1428 0.4125 1.0001 2.750 0.9603 0.06479 0.05334 -0.1391 0.4175 1.0001 3.000 0.9580 0.06751 0.05621 -0.1353 0.4235 1.0001 3.250 0.9611 0.07055 0.05927 -0.1325 0.4288 1.0001 3.500 0.9800 0.07390 0.06245 -0.1324 0.4340 1.0001 3.750 0.8940 0.07870 0.06797 -0.1184 0.4500 1.0001 4.000 0.7940 0.08675 0.07649 -0.1062 0.4718 1.0001 4.250 0.8020 0.09251 0.08213 -0.1079 0.4901 1.0001