XFOIL Version 6.94 Calculated polar for: agora05 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1099 0.04242 0.03048 -0.0103 0.9992 1.0008 -2.750 -0.0830 0.04216 0.02898 -0.0145 0.9992 1.0008 -2.500 -0.0577 0.04247 0.02848 -0.0158 0.9992 1.0008 -2.250 -0.0341 0.04290 0.02832 -0.0162 0.9992 1.0008 -2.000 -0.0111 0.04339 0.02836 -0.0163 0.9992 1.0008 -1.750 0.0116 0.04393 0.02854 -0.0163 0.9992 1.0008 -1.500 0.0340 0.04453 0.02886 -0.0163 0.9992 1.0008 -1.250 0.0563 0.04517 0.02931 -0.0162 0.9992 1.0008 -1.000 0.0784 0.04589 0.02988 -0.0162 0.9992 1.0008 -0.750 0.1002 0.04667 0.03059 -0.0162 0.9992 1.0008 -0.500 0.1217 0.04757 0.03148 -0.0163 0.9992 1.0008 -0.250 0.1425 0.04863 0.03263 -0.0166 0.9992 1.0008 0.000 0.1615 0.05003 0.03421 -0.0173 0.9992 1.0008 0.250 0.1747 0.05239 0.03687 -0.0191 0.9992 1.0008 0.500 0.1720 0.05714 0.04184 -0.0234 0.9992 1.0008 0.750 0.1704 0.06177 0.04646 -0.0273 0.9992 1.0008 1.000 0.3569 0.07235 0.05707 -0.0647 0.8476 1.0008 1.250 0.3853 0.07699 0.06162 -0.0686 0.8176 1.0008 1.750 0.4128 0.08524 0.06968 -0.0720 0.7899 1.0008 2.000 0.4167 0.08896 0.07330 -0.0728 0.7909 1.0008 2.250 0.4223 0.09271 0.07695 -0.0739 0.7933 1.0008 2.500 0.4249 0.09619 0.08035 -0.0747 0.8008 1.0008 2.750 0.4143 0.09885 0.08290 -0.0739 0.8207 1.0008 3.000 0.4018 0.10123 0.08519 -0.0728 0.8498 1.0008 3.250 0.3718 0.10202 0.08590 -0.0686 0.8952 1.0008 3.500 0.3354 0.10233 0.08610 -0.0618 0.9493 1.0008 3.750 0.2892 0.10166 0.08530 -0.0511 0.9992 1.0008 4.000 0.3009 0.10489 0.08844 -0.0526 0.9992 1.0008 4.250 0.3128 0.10813 0.09159 -0.0542 0.9992 1.0008 4.500 0.3247 0.11138 0.09475 -0.0557 0.9992 1.0008 4.750 0.3366 0.11464 0.09793 -0.0572 0.9992 1.0008 5.000 0.3486 0.11791 0.10112 -0.0587 0.9992 1.0008