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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0101 0.02678 0.01811 -0.0151 0.9994 1.0006 -2.750 0.0240 0.02793 0.01899 -0.0147 0.9994 1.0006 -2.500 0.0981 0.02765 0.01825 -0.0239 0.9644 1.0006 -2.250 0.1748 0.02723 0.01731 -0.0298 0.8714 1.0006 -2.000 0.1919 0.02813 0.01778 -0.0242 0.7884 1.0006 -1.750 0.2017 0.02912 0.01822 -0.0172 0.6967 1.0006 -1.500 0.2175 0.03024 0.01863 -0.0123 0.6045 1.0006 -1.250 0.2417 0.03130 0.01905 -0.0103 0.5490 1.0006 -1.000 0.2704 0.03236 0.01967 -0.0102 0.5152 1.0006 -0.750 0.3002 0.03341 0.02039 -0.0104 0.4929 1.0006 -0.500 0.3299 0.03444 0.02113 -0.0107 0.4767 1.0006 -0.250 0.3622 0.03573 0.02233 -0.0122 0.4620 1.0006 0.000 0.3910 0.03675 0.02303 -0.0122 0.4515 1.0006 0.250 0.4237 0.03827 0.02461 -0.0142 0.4400 1.0006 0.500 0.4529 0.03947 0.02562 -0.0147 0.4319 1.0006 0.750 0.4846 0.04132 0.02749 -0.0167 0.4261 1.0006 1.000 0.5173 0.04361 0.02991 -0.0196 0.4213 1.0006 1.250 0.5488 0.04602 0.03241 -0.0223 0.4177 1.0006 1.500 0.5789 0.04849 0.03493 -0.0249 0.4145 1.0006 1.750 0.6073 0.05077 0.03721 -0.0266 0.4110 1.0006 2.000 0.6340 0.05273 0.03903 -0.0270 0.4075 1.0006 2.250 0.6615 0.05675 0.04330 -0.0318 0.4059 1.0006 2.500 0.6863 0.06078 0.04748 -0.0356 0.4065 1.0006 2.750 0.6999 0.06873 0.05586 -0.0454 0.4134 1.0006 3.000 0.7035 0.07612 0.06343 -0.0524 0.4214 1.0006 3.250 0.7134 0.08116 0.06850 -0.0554 0.4260 1.0006 3.500 0.7279 0.08559 0.07297 -0.0574 0.4303 1.0006 4.250 0.5222 0.11466 0.10241 -0.0818 0.7404 1.0006 4.500 0.5255 0.11697 0.10465 -0.0807 0.7267 1.0006 5.000 0.5539 0.12327 0.11086 -0.0815 0.6983 1.0006