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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0685 0.02321 0.01547 -0.0225 0.9021 1.0006 -2.750 0.0933 0.02372 0.01548 -0.0199 0.8421 1.0006 -2.500 0.1102 0.02424 0.01554 -0.0157 0.7862 1.0006 -2.250 0.1286 0.02464 0.01551 -0.0119 0.7282 1.0006 -2.000 0.1473 0.02502 0.01537 -0.0081 0.6564 1.0006 -1.750 0.1681 0.02553 0.01516 -0.0050 0.5666 1.0006 -1.500 0.1940 0.02627 0.01516 -0.0038 0.4935 1.0006 -1.250 0.2226 0.02702 0.01533 -0.0036 0.4549 1.0006 -1.000 0.2522 0.02774 0.01562 -0.0038 0.4314 1.0006 -0.750 0.2826 0.02847 0.01606 -0.0042 0.4138 1.0006 -0.500 0.3128 0.02922 0.01654 -0.0047 0.4006 1.0006 -0.250 0.3429 0.02997 0.01705 -0.0050 0.3897 1.0006 0.000 0.3732 0.03087 0.01777 -0.0056 0.3819 1.0006 0.250 0.4043 0.03187 0.01874 -0.0064 0.3754 1.0006 0.500 0.4346 0.03290 0.01966 -0.0070 0.3701 1.0006 0.750 0.4643 0.03400 0.02060 -0.0075 0.3660 1.0006 1.000 0.4952 0.03535 0.02193 -0.0086 0.3625 1.0006 1.250 0.5269 0.03688 0.02359 -0.0101 0.3585 1.0006 1.500 0.5578 0.03844 0.02524 -0.0116 0.3543 1.0006 1.750 0.5879 0.04001 0.02684 -0.0128 0.3505 1.0006 2.000 0.6179 0.04179 0.02867 -0.0142 0.3482 1.0006 2.250 0.6479 0.04392 0.03091 -0.0159 0.3469 1.0006 2.500 0.6771 0.04618 0.03327 -0.0176 0.3456 1.0006 2.750 0.7051 0.04843 0.03559 -0.0190 0.3439 1.0006 3.000 0.7328 0.05145 0.03879 -0.0214 0.3426 1.0006 3.250 0.7590 0.05521 0.04282 -0.0247 0.3423 1.0006 3.750 0.7802 0.07364 0.06234 -0.0433 0.3596 1.0006 4.000 0.7891 0.07983 0.06866 -0.0475 0.3663 1.0006 5.000 0.4268 0.11481 0.10492 -0.0516 0.5553 1.0006