XFOIL Version 6.96 Calculated polar for: NACA 66-206 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.250 -0.5480 0.11102 0.10394 -0.0105 1.0000 0.0851 -9.000 -0.5556 0.10869 0.10174 -0.0147 1.0000 0.0882 -8.750 -0.5665 0.10617 0.09935 -0.0197 1.0000 0.0890 -8.500 -0.5444 0.09999 0.09313 -0.0150 1.0000 0.0945 -8.250 -0.5443 0.09672 0.08993 -0.0167 1.0000 0.0982 -8.000 -0.5600 0.09428 0.08762 -0.0235 1.0000 0.1020 -7.750 -0.5501 0.08945 0.08285 -0.0217 1.0000 0.1057 -7.500 -0.5461 0.08581 0.07923 -0.0226 1.0000 0.1107 -7.250 -0.5505 0.08225 0.07570 -0.0266 1.0000 0.1179 -6.250 -0.5047 0.06362 0.05618 -0.0354 1.0000 0.0483 -6.000 -0.4950 0.05914 0.05174 -0.0348 1.0000 0.0470 -5.750 -0.4837 0.05540 0.04781 -0.0347 1.0000 0.0467 -5.500 -0.4708 0.05177 0.04393 -0.0343 1.0000 0.0465 -5.250 -0.4566 0.04819 0.04010 -0.0338 1.0000 0.0456 -5.000 -0.4408 0.04469 0.03630 -0.0331 1.0000 0.0439 -4.750 -0.4229 0.04135 0.03259 -0.0324 1.0000 0.0420 -4.500 -0.4031 0.03813 0.02891 -0.0314 1.0000 0.0405 -4.250 -0.3819 0.03519 0.02547 -0.0303 1.0000 0.0396 -4.000 -0.3596 0.03248 0.02228 -0.0292 1.0000 0.0395 -3.750 -0.3365 0.03006 0.01942 -0.0280 1.0000 0.0400 -3.500 -0.3132 0.02812 0.01705 -0.0269 1.0000 0.0447 -3.250 -0.2891 0.02624 0.01460 -0.0257 1.0000 0.0502 -3.000 -0.2654 0.02439 0.01258 -0.0243 1.0000 0.0524 -2.750 -0.2419 0.02293 0.01090 -0.0227 1.0000 0.0556 -2.500 -0.2189 0.02178 0.00946 -0.0212 1.0000 0.0595 -2.250 -0.1973 0.02070 0.00821 -0.0200 1.0000 0.0670 -2.000 -0.1756 0.01981 0.00721 -0.0190 1.0000 0.0885 -1.500 -0.1085 0.01550 0.00587 -0.0182 1.0000 1.0000 -1.250 -0.0933 0.01547 0.00539 -0.0162 1.0000 1.0000 -1.000 -0.0764 0.01548 0.00508 -0.0147 1.0000 1.0000 -0.750 -0.0581 0.01551 0.00485 -0.0134 1.0000 1.0000 -0.500 -0.0388 0.01557 0.00469 -0.0124 1.0000 1.0000 -0.250 -0.0188 0.01565 0.00459 -0.0114 1.0000 1.0000 0.000 0.0017 0.01576 0.00454 -0.0106 1.0000 1.0000 0.250 0.0226 0.01589 0.00450 -0.0099 1.0000 1.0000 0.500 0.0438 0.01605 0.00455 -0.0093 1.0000 1.0000 0.750 0.0652 0.01622 0.00466 -0.0087 1.0000 1.0000 1.000 0.0867 0.01642 0.00483 -0.0081 1.0000 1.0000 1.250 0.1082 0.01665 0.00505 -0.0076 1.0000 1.0000 1.500 0.1298 0.01689 0.00533 -0.0072 1.0000 1.0000 1.750 0.1513 0.01716 0.00565 -0.0067 1.0000 1.0000 2.000 0.1728 0.01746 0.00602 -0.0063 1.0000 1.0000 2.250 0.1942 0.01778 0.00645 -0.0059 1.0000 1.0000 2.500 0.2310 0.01830 0.00718 -0.0086 0.9915 1.0000 2.750 0.2689 0.01885 0.00809 -0.0115 0.9824 1.0000 3.000 0.3075 0.01938 0.00897 -0.0144 0.9722 1.0000 3.250 0.3474 0.01991 0.00994 -0.0174 0.9601 1.0000 3.500 0.4695 0.02082 0.00800 -0.0234 0.0932 1.0000 3.750 0.4912 0.02215 0.00941 -0.0223 0.0740 1.0000 4.000 0.5133 0.02364 0.01101 -0.0212 0.0658 1.0000 4.250 0.5400 0.02514 0.01285 -0.0202 0.0606 1.0000 4.500 0.5691 0.02707 0.01489 -0.0199 0.0514 1.0000 4.750 0.5988 0.02898 0.01712 -0.0194 0.0444 1.0000 5.000 0.6277 0.03159 0.02001 -0.0189 0.0421 1.0000 5.250 0.6539 0.03485 0.02360 -0.0183 0.0406 1.0000 5.500 0.6774 0.03869 0.02799 -0.0173 0.0400 1.0000 5.750 0.6990 0.04150 0.03149 -0.0158 0.0389 1.0000 6.000 0.7182 0.04440 0.03511 -0.0141 0.0357 1.0000 6.250 0.7344 0.04788 0.03913 -0.0127 0.0337 1.0000 6.500 0.7483 0.05182 0.04353 -0.0115 0.0340 1.0000 6.750 0.7596 0.05595 0.04808 -0.0104 0.0346 1.0000 7.000 0.7685 0.06021 0.05269 -0.0095 0.0353 1.0000 7.250 0.7751 0.06454 0.05730 -0.0088 0.0362 1.0000 7.500 0.7793 0.06901 0.06199 -0.0085 0.0372 1.0000 7.750 0.7819 0.07357 0.06670 -0.0083 0.0382 1.0000 8.000 0.7850 0.07784 0.07123 -0.0084 0.0423 1.0000 8.250 0.7774 0.08283 0.07640 -0.0099 0.0451 1.0000 8.500 0.7700 0.08753 0.08116 -0.0111 0.0474 1.0000 8.750 0.7650 0.09213 0.08576 -0.0124 0.0497 1.0000 9.000 0.6560 0.08801 0.08203 -0.0100 0.0439 1.0000 9.250 0.6464 0.09407 0.08806 -0.0137 0.0456 1.0000