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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.6487 0.01074 0.00315 -0.1172 0.6119 0.0270 0.500 0.7033 0.01081 0.00315 -0.1171 0.5905 0.0334 1.000 0.7577 0.01086 0.00310 -0.1171 0.5707 0.0398 1.500 0.7987 0.00898 0.00333 -0.1148 0.5525 0.8731 2.000 0.8614 0.00916 0.00337 -0.1163 0.5331 1.0000 2.500 0.9139 0.00947 0.00352 -0.1160 0.5158 1.0000 3.000 0.9666 0.00977 0.00373 -0.1158 0.5000 1.0000 3.500 1.0182 0.01013 0.00397 -0.1154 0.4841 1.0000 4.000 1.0688 0.01054 0.00425 -0.1149 0.4670 1.0000 4.500 1.1062 0.01140 0.00461 -0.1119 0.3841 1.0000 5.000 1.1214 0.01376 0.00599 -0.1055 0.2281 1.0000 5.500 1.1325 0.01616 0.00770 -0.0984 0.1149 1.0000 6.000 1.1644 0.01722 0.00867 -0.0947 0.0954 1.0000 6.500 1.1746 0.01963 0.01071 -0.0880 0.0053 1.0000 7.000 1.2097 0.02071 0.01188 -0.0855 0.0053 1.0000 7.500 1.2429 0.02198 0.01323 -0.0828 0.0059 1.0000 8.000 1.2748 0.02341 0.01479 -0.0802 0.0069 1.0000 8.500 1.3042 0.02511 0.01663 -0.0775 0.0081 1.0000 9.000 1.3296 0.02722 0.01893 -0.0747 0.0088 1.0000 9.500 1.3486 0.02998 0.02191 -0.0716 0.0093 1.0000 10.000 1.3656 0.03308 0.02521 -0.0687 0.0098 1.0000 10.500 1.3766 0.03686 0.02921 -0.0658 0.0107 1.0000 11.000 1.3723 0.04232 0.03490 -0.0626 0.0113 1.0000 11.500 1.3829 0.04668 0.03947 -0.0607 0.0126 1.0000 12.000 1.3731 0.05356 0.04658 -0.0588 0.0132 1.0000 12.500 1.3594 0.06131 0.05452 -0.0576 0.0133 1.0000 13.000 1.3475 0.06913 0.06253 -0.0569 0.0132 1.0000 13.500 1.3396 0.07652 0.07009 -0.0564 0.0131 1.0000 14.000 1.3365 0.08315 0.07682 -0.0557 0.0123 1.0000 14.500 1.3423 0.08738 0.08102 -0.0531 0.0109 1.0000 15.000 1.1107 0.10109 0.09578 -0.0385 0.0142 1.0000 15.500 1.1235 0.10197 0.09659 -0.0350 0.0134 1.0000 16.000 1.1113 0.10582 0.10067 -0.0340 0.0127 1.0000 16.500 1.1209 0.10639 0.10130 -0.0313 0.0104 1.0000