XFOIL Version 6.94 Calculated polar for: Curtis C-72 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.6683 0.00917 0.00337 -0.0983 0.6418 0.6509 0.500 0.7344 0.00893 0.00356 -0.1003 0.5971 1.0000 1.000 0.7856 0.00906 0.00368 -0.0997 0.5847 1.0000 1.500 0.8347 0.00940 0.00369 -0.0986 0.5389 1.0000 2.000 0.8825 0.00991 0.00393 -0.0975 0.4862 1.0000 2.500 0.9268 0.01064 0.00420 -0.0958 0.4256 1.0000 3.000 0.9758 0.01124 0.00467 -0.0951 0.3855 1.0000 3.500 1.0239 0.01193 0.00514 -0.0944 0.3593 1.0000 4.000 1.0701 0.01279 0.00580 -0.0934 0.3303 1.0000 4.500 1.1199 0.01374 0.00679 -0.0931 0.3038 1.0000 5.500 1.2201 0.01419 0.00736 -0.0926 0.2863 1.0000 6.000 1.2639 0.01495 0.00794 -0.0914 0.2498 1.0000 6.500 1.2954 0.01643 0.00897 -0.0884 0.1832 1.0000 7.000 1.3219 0.01826 0.01025 -0.0849 0.1374 1.0000 7.500 1.3019 0.02327 0.01462 -0.0748 0.0216 1.0000 8.500 1.3166 0.02962 0.02145 -0.0644 0.0156 1.0000 9.000 1.3249 0.03314 0.02521 -0.0606 0.0153 1.0000 9.500 1.3242 0.03759 0.02991 -0.0570 0.0152 1.0000 10.000 1.3276 0.04224 0.03480 -0.0541 0.0155 1.0000 10.500 1.3233 0.04806 0.04093 -0.0517 0.0160 1.0000 11.000 1.3204 0.05404 0.04733 -0.0485 0.0172 1.0000 11.500 1.3063 0.06177 0.05567 -0.0447 0.0191 1.0000 12.000 1.2716 0.07301 0.06751 -0.0441 0.0207 1.0000 12.500 1.2290 0.08659 0.08165 -0.0463 0.0219 1.0000 13.000 1.1816 0.10212 0.09761 -0.0525 0.0224 1.0000 13.500 1.1362 0.11895 0.11479 -0.0615 0.0224 1.0000 14.000 1.0923 0.13696 0.13311 -0.0723 0.0222 1.0000 14.500 1.0465 0.15717 0.15358 -0.0831 0.0224 1.0000