XFOIL Version 6.94 Calculated polar for: nabac 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3524 0.05363 0.04131 0.0056 0.9999 0.3179 -2.750 -0.3149 0.05098 0.03841 0.0039 0.9999 0.3187 -2.500 -0.2756 0.04858 0.03589 0.0023 0.9999 0.3224 -2.250 -0.2376 0.04610 0.03372 0.0016 0.9999 0.3321 -2.000 -0.1968 0.04394 0.03177 0.0005 0.9999 0.3482 -1.750 -0.1570 0.04160 0.03010 -0.0001 0.9999 0.3748 -1.500 -0.1169 0.03902 0.02862 -0.0007 0.9999 0.4227 -1.000 0.0460 0.03483 0.02587 -0.0149 0.7803 1.0001 -0.750 0.0814 0.03655 0.02571 -0.0130 0.5984 1.0001 -0.500 0.1084 0.03786 0.02574 -0.0112 0.5486 1.0001 -0.250 0.1348 0.03919 0.02593 -0.0095 0.5234 1.0001 0.000 0.1639 0.04051 0.02632 -0.0088 0.5061 1.0001 0.250 0.1936 0.04188 0.02698 -0.0085 0.4935 1.0001 0.500 0.2258 0.04326 0.02788 -0.0092 0.4845 1.0001 0.750 0.2583 0.04473 0.02899 -0.0102 0.4775 1.0001 1.000 0.2904 0.04634 0.03025 -0.0114 0.4725 1.0001 1.250 0.3224 0.04810 0.03172 -0.0128 0.4693 1.0001 1.500 0.3547 0.05002 0.03345 -0.0148 0.4679 1.0001 1.750 0.3872 0.05213 0.03546 -0.0173 0.4681 1.0001 2.000 0.4198 0.05452 0.03785 -0.0208 0.4700 1.0001 2.250 0.4510 0.05724 0.04057 -0.0248 0.4736 1.0001 2.500 0.4795 0.06027 0.04364 -0.0290 0.4786 1.0001 2.750 0.5047 0.06347 0.04683 -0.0327 0.4842 1.0001 3.000 0.5281 0.06671 0.04997 -0.0354 0.4896 1.0001 3.250 0.5431 0.07075 0.05415 -0.0407 0.4990 1.0001 3.500 0.5554 0.07468 0.05804 -0.0442 0.5085 1.0001 3.750 0.5653 0.07856 0.06187 -0.0472 0.5180 1.0001 4.000 0.5670 0.08270 0.06591 -0.0501 0.5302 1.0001 4.250 0.5627 0.08678 0.06990 -0.0524 0.5436 1.0001 4.500 0.5587 0.09072 0.07371 -0.0541 0.5577 1.0001 4.750 0.5537 0.09453 0.07738 -0.0554 0.5731 1.0001 5.000 0.5463 0.09812 0.08081 -0.0561 0.5903 1.0001