XFOIL Version 6.94 Calculated polar for: nacc632 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0169 0.02619 0.01365 -0.0245 0.9994 1.0006 -2.750 0.0044 0.02666 0.01389 -0.0242 0.9994 1.0006 -2.500 0.0255 0.02719 0.01420 -0.0238 0.9994 1.0006 -2.250 0.0466 0.02778 0.01462 -0.0235 0.9994 1.0006 -2.000 0.0677 0.02843 0.01511 -0.0232 0.9994 1.0006 -1.750 0.0889 0.02913 0.01567 -0.0230 0.9994 1.0006 -1.500 0.1101 0.02988 0.01629 -0.0228 0.9994 1.0006 -1.250 0.1313 0.03067 0.01698 -0.0227 0.9994 1.0006 -1.000 0.1523 0.03150 0.01774 -0.0226 0.9994 1.0006 -0.750 0.1732 0.03238 0.01857 -0.0226 0.9994 1.0006 -0.500 0.1939 0.03331 0.01946 -0.0226 0.9994 1.0006 -0.250 0.2143 0.03429 0.02044 -0.0227 0.9994 1.0006 0.000 0.2345 0.03534 0.02150 -0.0228 0.9994 1.0006 0.250 0.2542 0.03646 0.02266 -0.0230 0.9994 1.0006 0.500 0.2734 0.03769 0.02394 -0.0233 0.9994 1.0006 0.750 0.2919 0.03904 0.02539 -0.0237 0.9994 1.0006 1.000 0.3091 0.04060 0.02707 -0.0243 0.9994 1.0006 1.250 0.3240 0.04253 0.02917 -0.0252 0.9994 1.0006 1.500 0.5355 0.03950 0.02685 -0.0590 0.8923 1.0006 1.750 0.6350 0.03627 0.02389 -0.0629 0.7698 1.0006 2.000 0.6583 0.03603 0.02347 -0.0568 0.6936 1.0006 2.250 0.6796 0.03616 0.02336 -0.0520 0.6359 1.0006 2.500 0.7028 0.03672 0.02370 -0.0489 0.5893 1.0006 2.750 0.7269 0.03731 0.02408 -0.0462 0.5510 1.0006 3.000 0.7528 0.03844 0.02507 -0.0450 0.5186 1.0006 3.250 0.7792 0.03982 0.02638 -0.0443 0.4907 1.0006 3.500 0.8055 0.04099 0.02737 -0.0431 0.4661 1.0006 3.750 0.8327 0.04291 0.02940 -0.0432 0.4445 1.0006 4.000 0.8591 0.04494 0.03150 -0.0433 0.4248 1.0006 4.250 0.8847 0.04701 0.03363 -0.0433 0.4063 1.0006 4.500 0.9104 0.04973 0.03651 -0.0438 0.3929 1.0006 4.750 0.9371 0.05230 0.03911 -0.0439 0.3826 1.0006 5.000 0.9566 0.05678 0.04418 -0.0461 0.3735 1.0006