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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0985 0.03746 0.02685 -0.0273 0.9992 0.3030 -2.750 -0.0696 0.03614 0.02596 -0.0271 0.9992 0.3757 -2.500 -0.0434 0.03373 0.02507 -0.0250 0.9992 0.5764 -2.250 -0.0323 0.03162 0.02444 -0.0196 0.9992 1.0008 -2.000 -0.0004 0.03221 0.02428 -0.0214 0.9992 1.0008 -1.750 0.2008 0.02964 0.02094 -0.0490 0.9040 1.0008 -1.500 0.2200 0.03093 0.02210 -0.0442 0.8635 1.0008 -1.250 0.2271 0.03271 0.02379 -0.0366 0.8208 1.0008 -1.000 0.2263 0.03498 0.02595 -0.0266 0.7688 1.0008 -0.750 0.2248 0.03746 0.02835 -0.0165 0.7008 1.0008 -0.500 0.2256 0.04012 0.03086 -0.0073 0.6234 1.0008 -0.250 0.2440 0.04008 0.03080 -0.0026 0.5440 1.0008 0.000 0.2881 0.03571 0.02619 -0.0039 0.4303 1.0008 0.250 0.3070 0.03544 0.02526 0.0009 0.4034 1.0008 0.500 0.3308 0.03593 0.02522 0.0034 0.3793 1.0008 0.750 0.3559 0.03671 0.02556 0.0052 0.3601 1.0008 1.000 0.3829 0.03760 0.02620 0.0064 0.3466 1.0008 1.250 0.4098 0.03869 0.02704 0.0075 0.3346 1.0008 1.500 0.4366 0.03994 0.02807 0.0085 0.3227 1.0008 1.750 0.4650 0.04132 0.02931 0.0091 0.3128 1.0008 2.000 0.4938 0.04283 0.03073 0.0096 0.3064 1.0008 2.250 0.5243 0.04454 0.03245 0.0094 0.2995 1.0008 2.500 0.5548 0.04645 0.03439 0.0089 0.2921 1.0008 2.750 0.5866 0.04865 0.03674 0.0077 0.2860 1.0008 3.000 0.6176 0.05106 0.03922 0.0067 0.2834 1.0008 3.250 0.6510 0.05413 0.04257 0.0040 0.2820 1.0008 3.500 0.6850 0.05826 0.04713 -0.0009 0.2803 1.0008 3.750 0.7165 0.06368 0.05298 -0.0075 0.2795 1.0008