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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0221 0.02608 0.01762 0.0024 0.4743 1.0006 -2.750 0.0496 0.02687 0.01722 0.0031 0.3975 1.0006 -2.250 0.1088 0.02826 0.01717 0.0030 0.3512 1.0006 -2.000 0.1388 0.02896 0.01738 0.0028 0.3407 1.0006 -1.750 0.1690 0.02967 0.01771 0.0027 0.3320 1.0006 -1.500 0.1988 0.03048 0.01807 0.0026 0.3250 1.0006 -1.250 0.2298 0.03125 0.01867 0.0022 0.3186 1.0006 -1.000 0.2605 0.03211 0.01931 0.0019 0.3129 1.0006 -0.750 0.2910 0.03306 0.02002 0.0016 0.3087 1.0006 -0.500 0.3214 0.03417 0.02093 0.0013 0.3062 1.0006 -0.250 0.3526 0.03538 0.02207 0.0007 0.3048 1.0006 0.000 0.3842 0.03671 0.02343 -0.0001 0.3039 1.0006 0.250 0.4158 0.03823 0.02498 -0.0010 0.3035 1.0006 0.500 0.4476 0.03994 0.02675 -0.0022 0.3035 1.0006 0.750 0.4793 0.04186 0.02875 -0.0036 0.3040 1.0006 1.000 0.5107 0.04399 0.03097 -0.0052 0.3044 1.0006 1.250 0.5419 0.04635 0.03345 -0.0070 0.3046 1.0006 1.500 0.5724 0.04894 0.03618 -0.0091 0.3047 1.0006 1.750 0.6020 0.05181 0.03917 -0.0112 0.3054 1.0006 2.000 0.6336 0.05878 0.04704 -0.0205 0.3159 1.0006 2.250 0.6582 0.06354 0.05194 -0.0243 0.3225 1.0006