XFOIL Version 6.94 Calculated polar for: ARA-D BL 13% 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1952 0.04291 0.03553 -0.0182 1.0000 0.2421 -2.750 -0.1733 0.04142 0.03399 -0.0182 1.0000 0.2484 -2.500 -0.1474 0.04009 0.03242 -0.0190 1.0000 0.2560 -2.250 -0.0144 0.03569 0.02773 -0.0359 0.9519 0.2868 -2.000 0.1252 0.02892 0.02075 -0.0397 0.6632 0.3408 -1.750 0.1401 0.02975 0.01956 -0.0338 0.4232 0.3609 -1.500 0.1663 0.02994 0.01917 -0.0325 0.3757 0.3891 -1.250 0.1938 0.02987 0.01895 -0.0314 0.3507 0.4277 -1.000 0.2220 0.02957 0.01885 -0.0303 0.3328 0.4892 -0.750 0.2400 0.02826 0.01895 -0.0261 0.3217 0.6725 -0.500 0.3020 0.02885 0.01892 -0.0298 0.3083 1.0000 -0.250 0.3329 0.02993 0.01966 -0.0296 0.3020 1.0000 0.000 0.3627 0.03104 0.02056 -0.0293 0.2966 1.0000 0.250 0.3918 0.03227 0.02159 -0.0290 0.2917 1.0000 0.500 0.4200 0.03374 0.02277 -0.0287 0.2867 1.0000 0.750 0.4477 0.03527 0.02423 -0.0286 0.2824 1.0000 1.000 0.4750 0.03671 0.02577 -0.0285 0.2783 1.0000 1.250 0.5020 0.03840 0.02754 -0.0284 0.2759 1.0000 1.500 0.5284 0.04033 0.02956 -0.0285 0.2749 1.0000 1.750 0.5537 0.04249 0.03184 -0.0286 0.2747 1.0000 2.000 0.5777 0.04492 0.03441 -0.0287 0.2754 1.0000 2.250 0.5999 0.04761 0.03725 -0.0290 0.2761 1.0000 2.500 0.6194 0.05059 0.04042 -0.0293 0.2767 1.0000 2.750 0.6361 0.05390 0.04392 -0.0297 0.2774 1.0000 3.000 0.6503 0.05759 0.04778 -0.0302 0.2790 1.0000 3.250 0.6654 0.06145 0.05169 -0.0306 0.2821 1.0000 4.250 0.4038 0.10940 0.10137 -0.0732 0.5864 1.0000 4.750 0.4400 0.11616 0.10786 -0.0733 0.5575 1.0000 5.000 0.4399 0.11817 0.10976 -0.0720 0.5406 1.0000