XFOIL Version 6.94 Calculated polar for: GOA 1b 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1875 0.03173 0.01692 -0.0305 1.0000 0.2987 -2.750 -0.1613 0.02870 0.01468 -0.0289 1.0000 0.3751 -2.500 -0.1250 0.02385 0.01135 -0.0251 1.0000 1.0000 -2.250 -0.0981 0.02379 0.01015 -0.0238 1.0000 1.0000 -2.000 -0.0728 0.02376 0.00935 -0.0227 1.0000 1.0000 -1.750 -0.0481 0.02374 0.00876 -0.0217 1.0000 1.0000 -1.500 -0.0236 0.02374 0.00827 -0.0208 1.0000 1.0000 -1.250 0.0007 0.02376 0.00788 -0.0200 1.0000 1.0000 -1.000 0.0249 0.02380 0.00759 -0.0191 1.0000 1.0000 -0.750 0.0490 0.02385 0.00741 -0.0183 1.0000 1.0000 -0.500 0.0731 0.02392 0.00728 -0.0174 1.0000 1.0000 -0.250 0.0972 0.02401 0.00732 -0.0166 1.0000 1.0000 0.000 0.1215 0.02412 0.00748 -0.0157 1.0000 1.0000 0.250 0.1461 0.02423 0.00780 -0.0147 1.0000 1.0000 0.500 0.1717 0.02435 0.00832 -0.0137 1.0000 1.0000 0.750 0.2006 0.02441 0.00911 -0.0128 1.0000 1.0000 1.000 0.2406 0.02428 0.01034 -0.0134 1.0000 1.0000 1.250 0.3255 0.02897 0.01208 -0.0198 0.4390 1.0000 1.500 0.3547 0.03083 0.01359 -0.0193 0.4122 1.0000 1.750 0.3866 0.03260 0.01528 -0.0194 0.3953 1.0000 2.000 0.4198 0.03436 0.01716 -0.0200 0.3819 1.0000 2.250 0.4517 0.03625 0.01922 -0.0204 0.3674 1.0000 2.500 0.4829 0.03827 0.02144 -0.0210 0.3544 1.0000 2.750 0.5150 0.04037 0.02392 -0.0219 0.3478 1.0000 3.000 0.5482 0.04267 0.02678 -0.0234 0.3490 1.0000 3.250 0.5808 0.04525 0.02991 -0.0253 0.3528 1.0000 3.500 0.6119 0.04814 0.03331 -0.0272 0.3580 1.0000 3.750 0.6424 0.05122 0.03705 -0.0305 0.3670 1.0000 4.000 0.6711 0.05483 0.04103 -0.0331 0.3766 1.0000 4.250 0.6989 0.05879 0.04554 -0.0382 0.3921 1.0000 4.500 0.7239 0.06318 0.05034 -0.0435 0.4102 1.0000 4.750 0.7391 0.06794 0.05551 -0.0515 0.4351 1.0000 5.000 0.7560 0.07301 0.06079 -0.0579 0.4619 1.0000