XFOIL Version 6.94 Calculated polar for: GOE 195 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 0.5982 0.00829 0.00202 -0.1350 0.7132 0.2485 0.500 0.6521 0.00847 0.00213 -0.1343 0.6727 0.2988 1.000 0.7050 0.00866 0.00216 -0.1334 0.6208 0.3308 1.500 0.7586 0.00882 0.00225 -0.1328 0.5871 0.3648 2.000 0.8124 0.00882 0.00253 -0.1324 0.5631 0.4990 3.500 0.9633 0.00931 0.00318 -0.1286 0.4120 1.0000 4.000 1.0052 0.01097 0.00388 -0.1265 0.2474 1.0000 4.500 1.0449 0.01302 0.00503 -0.1242 0.0952 1.0000 5.000 1.0892 0.01444 0.00605 -0.1223 0.0269 1.0000 5.500 1.1387 0.01514 0.00672 -0.1211 0.0182 1.0000 6.000 1.1859 0.01606 0.00759 -0.1196 0.0037 1.0000 6.500 1.2335 0.01686 0.00849 -0.1180 0.0036 1.0000 7.000 1.2794 0.01776 0.00950 -0.1163 0.0037 1.0000 7.500 1.3236 0.01876 0.01063 -0.1142 0.0039 1.0000 8.000 1.3652 0.01992 0.01193 -0.1117 0.0042 1.0000 8.500 1.4027 0.02132 0.01351 -0.1086 0.0045 1.0000 9.000 1.4347 0.02300 0.01537 -0.1047 0.0048 1.0000 9.500 1.4629 0.02452 0.01705 -0.1002 0.0053 1.0000 10.000 1.4805 0.02675 0.01949 -0.0945 0.0058 1.0000 10.500 1.4866 0.02996 0.02291 -0.0882 0.0062 1.0000 11.000 1.4993 0.03297 0.02612 -0.0835 0.0068 1.0000 11.500 1.5016 0.03727 0.03071 -0.0786 0.0076 1.0000 12.000 1.4937 0.04309 0.03673 -0.0741 0.0082 1.0000 12.500 1.4993 0.04816 0.04219 -0.0702 0.0095 1.0000 14.500 1.2700 0.08568 0.08190 -0.0433 0.0155 1.0000 15.000 1.2135 0.09530 0.09186 -0.0444 0.0156 1.0000 15.500 1.1628 0.10505 0.10190 -0.0474 0.0156 1.0000