XFOIL Version 6.94 Calculated polar for: GOE 611 AIRFOIL 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.7021 0.01223 0.00401 -0.1174 0.6284 0.0428 1.000 0.7561 0.01220 0.00390 -0.1172 0.6074 0.0562 1.500 0.8086 0.01018 0.00401 -0.1168 0.5884 1.0000 2.000 0.8606 0.01050 0.00409 -0.1163 0.5686 1.0000 2.500 0.9126 0.01086 0.00426 -0.1159 0.5505 1.0000 3.000 0.9641 0.01128 0.00449 -0.1154 0.5334 1.0000 3.500 1.0153 0.01164 0.00475 -0.1149 0.5165 1.0000 4.000 1.0663 0.01203 0.00508 -0.1145 0.5006 1.0000 4.500 1.1091 0.01249 0.00530 -0.1123 0.4619 1.0000 5.000 1.1412 0.01338 0.00581 -0.1083 0.3845 1.0000 5.500 1.1652 0.01491 0.00676 -0.1033 0.2957 1.0000 6.000 1.1557 0.01813 0.00894 -0.0930 0.1366 1.0000 6.500 1.1771 0.01991 0.01053 -0.0881 0.1049 1.0000 7.500 1.2112 0.02447 0.01477 -0.0787 0.0064 1.0000 8.000 1.2404 0.02617 0.01661 -0.0760 0.0065 1.0000 8.500 1.2674 0.02816 0.01876 -0.0735 0.0070 1.0000 9.000 1.2903 0.03058 0.02139 -0.0708 0.0076 1.0000 9.500 1.3132 0.03313 0.02414 -0.0683 0.0086 1.0000 10.000 1.3294 0.03638 0.02763 -0.0657 0.0095 1.0000 10.500 1.3465 0.03971 0.03120 -0.0634 0.0108 1.0000 11.000 1.3596 0.04357 0.03531 -0.0612 0.0122 1.0000 11.500 1.3623 0.04873 0.04076 -0.0590 0.0136 1.0000 12.000 1.3690 0.05384 0.04615 -0.0576 0.0149 1.0000 12.500 1.3570 0.06145 0.05402 -0.0565 0.0159 1.0000 13.000 1.3555 0.06819 0.06101 -0.0560 0.0175 1.0000 13.500 1.3385 0.07699 0.07002 -0.0558 0.0187 1.0000 14.000 1.3400 0.08354 0.07679 -0.0555 0.0211 1.0000 14.500 1.3476 0.08871 0.08209 -0.0541 0.0244 1.0000 15.000 1.3946 0.08691 0.08022 -0.0482 0.0298 1.0000 17.000 1.5420 0.09591 0.09019 -0.0331 0.0283 1.0000 17.500 1.5090 0.10564 0.10040 -0.0352 0.0277 1.0000