XFOIL Version 6.94 Calculated polar for: E385 (8.41%) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.7031 0.00856 0.00332 -0.1579 0.7159 1.0000 0.500 0.7604 0.00882 0.00331 -0.1582 0.7042 1.0000 1.000 0.8168 0.00906 0.00342 -0.1583 0.6919 1.0000 1.500 0.8731 0.00935 0.00358 -0.1584 0.6805 1.0000 2.000 0.9295 0.00967 0.00376 -0.1585 0.6695 1.0000 2.500 0.9847 0.00993 0.00400 -0.1585 0.6568 1.0000 3.000 1.0395 0.01020 0.00427 -0.1583 0.6431 1.0000 3.500 1.0940 0.01048 0.00454 -0.1580 0.6292 1.0000 4.000 1.1479 0.01074 0.00484 -0.1576 0.6139 1.0000 4.500 1.2012 0.01101 0.00515 -0.1571 0.5976 1.0000 5.000 1.2537 0.01126 0.00551 -0.1564 0.5796 1.0000 5.500 1.3049 0.01151 0.00584 -0.1555 0.5573 1.0000 6.000 1.3513 0.01176 0.00608 -0.1535 0.5130 1.0000 6.500 1.3948 0.01239 0.00661 -0.1512 0.4549 1.0000 7.000 1.4296 0.01376 0.00759 -0.1477 0.3565 1.0000 8.000 1.4571 0.01993 0.01214 -0.1356 0.0897 1.0000 8.500 1.4549 0.02338 0.01515 -0.1269 0.0247 1.0000 9.000 1.4652 0.02620 0.01816 -0.1207 0.0195 1.0000 9.500 1.4769 0.02915 0.02138 -0.1155 0.0178 1.0000 10.000 1.4834 0.03289 0.02539 -0.1107 0.0167 1.0000 10.500 1.4872 0.03736 0.03009 -0.1066 0.0160 1.0000 11.000 1.4909 0.04234 0.03530 -0.1032 0.0155 1.0000 11.500 1.4970 0.04757 0.04072 -0.1003 0.0149 1.0000 12.000 1.5112 0.05347 0.04686 -0.0969 0.0140 1.0000 12.500 1.5307 0.05872 0.05251 -0.0947 0.0139 1.0000 13.000 1.5429 0.06544 0.05970 -0.0926 0.0138 1.0000 13.500 1.5416 0.07452 0.06932 -0.0910 0.0139 1.0000 14.000 1.5301 0.08361 0.07887 -0.0906 0.0141 1.0000 14.500 1.5124 0.09243 0.08812 -0.0918 0.0144 1.0000 15.000 1.4812 0.10466 0.10094 -0.0953 0.0150 1.0000 15.500 1.4160 0.12557 0.12263 -0.1054 0.0162 1.0000 16.000 1.3675 0.14613 0.14363 -0.1180 0.0170 1.0000 16.500 1.3239 0.16922 0.16705 -0.1342 0.0174 1.0000 17.000 1.2712 0.20269 0.20064 -0.1563 0.0188 1.0000