XFOIL Version 6.94 Calculated polar for: E212 (10.55%) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4903 0.00670 0.00209 -0.1237 0.8452 0.6600 0.500 0.5470 0.00671 0.00212 -0.1234 0.8123 0.7018 1.000 0.6008 0.00681 0.00220 -0.1225 0.7756 0.7431 1.500 0.6524 0.00698 0.00233 -0.1210 0.7353 0.7841 2.000 0.7013 0.00719 0.00251 -0.1191 0.6919 0.8241 2.500 0.7471 0.00745 0.00271 -0.1164 0.6468 0.8659 3.000 0.7877 0.00767 0.00289 -0.1126 0.5996 0.9146 3.500 0.8333 0.00791 0.00303 -0.1101 0.5483 1.0000 4.000 0.8855 0.00847 0.00336 -0.1095 0.4942 1.0000 4.500 0.9356 0.00911 0.00379 -0.1085 0.4368 1.0000 5.000 0.9823 0.01000 0.00431 -0.1069 0.3551 1.0000 5.500 1.0288 0.01097 0.00495 -0.1054 0.2811 1.0000 6.000 1.0744 0.01206 0.00569 -0.1038 0.2083 1.0000 6.500 1.1191 0.01323 0.00657 -0.1021 0.1468 1.0000 7.000 1.1612 0.01461 0.00762 -0.1000 0.0867 1.0000 7.500 1.1959 0.01669 0.00925 -0.0967 0.0234 1.0000 8.000 1.2348 0.01828 0.01089 -0.0937 0.0163 1.0000 8.500 1.2730 0.01974 0.01252 -0.0907 0.0146 1.0000 9.000 1.3044 0.02150 0.01445 -0.0867 0.0135 1.0000 9.500 1.3234 0.02379 0.01691 -0.0809 0.0127 1.0000 10.000 1.3351 0.02676 0.02008 -0.0745 0.0121 1.0000 10.500 1.3553 0.02918 0.02273 -0.0698 0.0117 1.0000 11.000 1.3702 0.03236 0.02617 -0.0650 0.0114 1.0000 11.500 1.3830 0.03606 0.03015 -0.0607 0.0112 1.0000 12.000 1.3924 0.04039 0.03479 -0.0569 0.0110 1.0000 12.500 1.3961 0.04554 0.04030 -0.0535 0.0109 1.0000 13.000 1.3916 0.05183 0.04699 -0.0510 0.0108 1.0000 13.500 1.3774 0.05969 0.05528 -0.0500 0.0108 1.0000 14.000 1.3532 0.06954 0.06558 -0.0512 0.0109 1.0000 14.500 1.3212 0.08178 0.07826 -0.0554 0.0110 1.0000 15.000 1.2841 0.09666 0.09356 -0.0629 0.0112 1.0000 15.500 1.2437 0.11424 0.11152 -0.0735 0.0113 1.0000 16.000 1.2004 0.13495 0.13257 -0.0872 0.0115 1.0000 16.500 1.1503 0.16026 0.15817 -0.1041 0.0119 1.0000