XFOIL Version 6.94 Calculated polar for: GOE 440 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.6619 0.00885 0.00266 -0.1233 0.5397 0.5721 1.000 0.7165 0.00922 0.00286 -0.1229 0.4863 0.6187 1.500 0.7706 0.00963 0.00311 -0.1224 0.4395 0.6672 2.000 0.8246 0.01001 0.00342 -0.1220 0.4068 0.7156 2.500 0.8783 0.01034 0.00376 -0.1214 0.3854 0.7676 3.000 0.9295 0.01057 0.00410 -0.1202 0.3666 0.8407 3.500 0.9752 0.01064 0.00430 -0.1177 0.3505 1.0000 4.000 1.0307 0.01110 0.00466 -0.1178 0.3354 1.0000 4.500 1.0856 0.01157 0.00506 -0.1177 0.3207 1.0000 5.000 1.1402 0.01202 0.00548 -0.1175 0.3060 1.0000 5.500 1.1934 0.01255 0.00595 -0.1171 0.2907 1.0000 6.000 1.2466 0.01299 0.00634 -0.1168 0.2658 1.0000 6.500 1.2978 0.01361 0.00687 -0.1162 0.2410 1.0000 7.000 1.3456 0.01453 0.00751 -0.1151 0.1950 1.0000 7.500 1.3876 0.01599 0.00857 -0.1133 0.1358 1.0000 8.000 1.4204 0.01826 0.01028 -0.1103 0.0663 1.0000 8.500 1.4424 0.02127 0.01300 -0.1056 0.0056 1.0000 9.500 1.5106 0.02400 0.01606 -0.0992 0.0050 1.0000 10.000 1.5343 0.02575 0.01802 -0.0948 0.0045 1.0000 10.500 1.5541 0.02792 0.02042 -0.0903 0.0050 1.0000 11.000 1.5669 0.03081 0.02356 -0.0859 0.0047 1.0000 11.500 1.5652 0.03524 0.02829 -0.0815 0.0043 1.0000 12.000 1.5719 0.03946 0.03275 -0.0788 0.0046 1.0000 12.500 1.5605 0.04623 0.03979 -0.0772 0.0052 1.0000 13.000 1.5370 0.05564 0.04952 -0.0779 0.0044 1.0000 13.500 1.4828 0.07057 0.06482 -0.0817 0.0041 1.0000 14.000 1.4530 0.08304 0.07761 -0.0854 0.0038 1.0000 14.500 1.4554 0.09031 0.08505 -0.0875 0.0043 1.0000 15.000 1.3889 0.10956 0.10461 -0.0944 0.0039 1.0000 16.000 1.3634 0.13002 0.12542 -0.1018 0.0047 1.0000 16.500 1.3407 0.14224 0.13780 -0.1071 0.0038 1.0000 17.500 1.3465 0.15796 0.15381 -0.1142 0.0055 1.0000 18.000 1.3486 0.16562 0.16158 -0.1178 0.0058 1.0000