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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0254 0.02523 0.01728 -0.0179 0.9894 1.0006 -2.750 0.0984 0.02479 0.01625 -0.0254 0.9261 1.0006 -2.500 0.1335 0.02521 0.01617 -0.0241 0.8534 1.0006 -2.250 0.1498 0.02586 0.01640 -0.0194 0.7885 1.0006 -2.000 0.1643 0.02646 0.01654 -0.0142 0.7180 1.0006 -1.750 0.1804 0.02710 0.01656 -0.0094 0.6284 1.0006 -1.500 0.2018 0.02794 0.01660 -0.0065 0.5468 1.0006 -1.250 0.2289 0.02881 0.01684 -0.0057 0.4988 1.0006 -1.000 0.2588 0.02970 0.01733 -0.0059 0.4694 1.0006 -0.750 0.2883 0.03053 0.01776 -0.0060 0.4507 1.0006 -0.500 0.3193 0.03141 0.01846 -0.0067 0.4346 1.0006 -0.250 0.3500 0.03235 0.01922 -0.0074 0.4222 1.0006 0.000 0.3801 0.03326 0.01991 -0.0079 0.4115 1.0006 0.250 0.4106 0.03437 0.02089 -0.0087 0.4036 1.0006 0.500 0.4423 0.03572 0.02226 -0.0100 0.3971 1.0006 0.750 0.4730 0.03706 0.02356 -0.0111 0.3920 1.0006 1.000 0.5030 0.03845 0.02485 -0.0119 0.3880 1.0006 1.250 0.5337 0.04016 0.02654 -0.0133 0.3846 1.0006 1.500 0.5658 0.04236 0.02896 -0.0159 0.3808 1.0006 1.750 0.5968 0.04465 0.03140 -0.0184 0.3768 1.0006 2.000 0.6266 0.04706 0.03392 -0.0207 0.3736 1.0006 2.250 0.6560 0.05009 0.03713 -0.0238 0.3727 1.0006 2.500 0.6839 0.05386 0.04113 -0.0277 0.3730 1.0006 2.750 0.7088 0.05830 0.04580 -0.0321 0.3736 1.0006 3.000 0.7287 0.06357 0.05133 -0.0372 0.3749 1.0006 3.250 0.7429 0.06954 0.05749 -0.0425 0.3783 1.0006 3.500 0.7572 0.07471 0.06273 -0.0458 0.3831 1.0006