XFOIL Version 6.94 Calculated polar for: EPPLER 335 BL AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0546 0.02256 0.01508 -0.0166 0.8351 1.0006 -2.750 0.0736 0.02297 0.01501 -0.0131 0.7862 1.0006 -2.500 0.0930 0.02329 0.01484 -0.0098 0.7340 1.0006 -2.250 0.1146 0.02354 0.01463 -0.0070 0.6742 1.0006 -2.000 0.1367 0.02383 0.01435 -0.0044 0.5973 1.0006 -1.750 0.1616 0.02434 0.01410 -0.0029 0.5060 1.0006 -1.500 0.1895 0.02502 0.01406 -0.0024 0.4494 1.0006 -1.250 0.2191 0.02567 0.01423 -0.0026 0.4177 1.0006 -1.000 0.2489 0.02631 0.01447 -0.0029 0.3981 1.0006 -0.750 0.2785 0.02695 0.01472 -0.0031 0.3841 1.0006 -0.500 0.3091 0.02760 0.01519 -0.0036 0.3723 1.0006 -0.250 0.3384 0.02834 0.01561 -0.0038 0.3646 1.0006 0.000 0.3693 0.02915 0.01636 -0.0044 0.3582 1.0006 0.250 0.3998 0.03002 0.01714 -0.0050 0.3528 1.0006 0.500 0.4298 0.03094 0.01793 -0.0055 0.3484 1.0006 0.750 0.4595 0.03194 0.01876 -0.0059 0.3447 1.0006 1.000 0.4901 0.03309 0.01988 -0.0067 0.3412 1.0006 1.250 0.5212 0.03432 0.02121 -0.0078 0.3372 1.0006 1.500 0.5518 0.03561 0.02254 -0.0089 0.3334 1.0006 1.750 0.5822 0.03701 0.02398 -0.0099 0.3305 1.0006 2.000 0.6126 0.03862 0.02566 -0.0112 0.3289 1.0006 2.250 0.6426 0.04036 0.02748 -0.0124 0.3274 1.0006 2.500 0.6720 0.04213 0.02930 -0.0136 0.3255 1.0006 2.750 0.7004 0.04399 0.03118 -0.0145 0.3234 1.0006 3.000 0.7295 0.04656 0.03399 -0.0167 0.3220 1.0006 3.250 0.7577 0.04950 0.03718 -0.0191 0.3218 1.0006 3.500 0.7845 0.05271 0.04059 -0.0214 0.3224 1.0006 3.750 0.8095 0.05636 0.04445 -0.0240 0.3230 1.0006 4.000 0.8068 0.07151 0.06073 -0.0397 0.3368 1.0006 4.250 0.8141 0.07860 0.06798 -0.0446 0.3456 1.0006 5.000 0.6185 0.11920 0.10904 -0.0765 0.5340 1.0006