XFOIL Version 6.96 Calculated polar for: GOE 424 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.500 -0.8208 0.09476 0.09208 -0.0771 1.0000 0.0185 -19.250 -0.8582 0.08499 0.08214 -0.0828 1.0000 0.0186 -19.000 -0.8826 0.07764 0.07466 -0.0871 1.0000 0.0187 -18.750 -0.8985 0.07181 0.06873 -0.0905 1.0000 0.0189 -18.500 -0.9093 0.06684 0.06367 -0.0934 1.0000 0.0190 -18.250 -0.9166 0.06251 0.05927 -0.0958 1.0000 0.0192 -18.000 -0.9204 0.05878 0.05547 -0.0979 1.0000 0.0193 -17.750 -0.9227 0.05535 0.05199 -0.0996 1.0000 0.0195 -17.500 -0.9242 0.05221 0.04878 -0.1011 1.0000 0.0197 -17.250 -0.9258 0.04933 0.04584 -0.1022 1.0000 0.0199 -17.000 -0.9289 0.04663 0.04307 -0.1028 1.0000 0.0200 -16.750 -0.9345 0.04418 0.04056 -0.1026 1.0000 0.0202 -16.500 -0.9427 0.04199 0.03830 -0.1015 1.0000 0.0204 -16.250 -0.9464 0.03984 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12.500 1.5795 0.02194 0.01688 -0.0530 0.2435 1.0000 12.750 1.5843 0.02345 0.01839 -0.0514 0.2362 1.0000 13.000 1.5916 0.02486 0.01983 -0.0502 0.2297 1.0000 13.250 1.5947 0.02663 0.02161 -0.0487 0.2227 1.0000 13.500 1.5980 0.02848 0.02347 -0.0475 0.2162 1.0000 13.750 1.6001 0.03052 0.02552 -0.0463 0.2095 1.0000 14.000 1.5988 0.03291 0.02793 -0.0451 0.2036 1.0000 14.250 1.6024 0.03495 0.03002 -0.0443 0.1982 1.0000 14.500 1.5987 0.03772 0.03281 -0.0432 0.1924 1.0000 14.750 1.5991 0.04017 0.03530 -0.0425 0.1876 1.0000 15.000 1.5991 0.04268 0.03786 -0.0418 0.1824 1.0000 15.250 1.5920 0.04589 0.04108 -0.0409 0.1767 1.0000 15.500 1.5907 0.04857 0.04380 -0.0403 0.1710 1.0000 15.750 1.5817 0.05203 0.04727 -0.0395 0.1638 1.0000 16.000 1.5761 0.05517 0.05044 -0.0389 0.1568 1.0000 16.250 1.5642 0.05900 0.05425 -0.0382 0.1487 1.0000 16.500 1.5559 0.06250 0.05775 -0.0377 0.1397 1.0000 16.750 1.5456 0.06623 0.06147 -0.0372 0.1315 1.0000 17.000 1.5321 0.07037 0.06560 -0.0367 0.1229 1.0000 17.250 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