XFOIL Version 6.94 Calculated polar for: EPPLER 335 BL2 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0608 0.02813 0.01999 -0.0102 0.7628 1.0006 -2.750 0.0724 0.02922 0.02005 -0.0036 0.6151 1.0006 -2.500 0.0933 0.03035 0.02000 -0.0006 0.5099 1.0006 -2.250 0.1206 0.03137 0.02010 0.0003 0.4610 1.0006 -2.000 0.1496 0.03232 0.02034 0.0005 0.4342 1.0006 -1.750 0.1802 0.03329 0.02086 0.0002 0.4156 1.0006 -1.500 0.2116 0.03434 0.02157 -0.0004 0.4026 1.0006 -1.250 0.2422 0.03542 0.02229 -0.0007 0.3930 1.0006 -1.000 0.2742 0.03667 0.02333 -0.0016 0.3858 1.0006 -0.750 0.3069 0.03805 0.02464 -0.0028 0.3790 1.0006 -0.500 0.3384 0.03941 0.02583 -0.0037 0.3729 1.0006 -0.250 0.3687 0.04083 0.02699 -0.0042 0.3676 1.0006 0.000 0.4024 0.04275 0.02907 -0.0065 0.3634 1.0006 0.250 0.4353 0.04492 0.03134 -0.0088 0.3607 1.0006 0.500 0.4680 0.04747 0.03400 -0.0115 0.3601 1.0006 0.750 0.5003 0.05042 0.03707 -0.0147 0.3607 1.0006 1.000 0.5316 0.05375 0.04051 -0.0181 0.3623 1.0006 1.250 0.5610 0.05737 0.04423 -0.0217 0.3647 1.0006 1.500 0.5883 0.06097 0.04785 -0.0245 0.3672 1.0006 1.750 0.6088 0.06915 0.05659 -0.0361 0.3786 1.0006 2.000 0.6253 0.07442 0.06191 -0.0407 0.3850 1.0006 2.250 0.6396 0.07965 0.06719 -0.0448 0.3922 1.0006 2.500 0.6380 0.08750 0.07514 -0.0522 0.4100 1.0006 2.750 0.5781 0.09917 0.08704 -0.0644 0.4726 1.0006