XFOIL Version 6.94 Calculated polar for: GOA 3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1726 0.02857 0.01684 -0.0282 0.9999 0.5224 -2.750 -0.1269 0.02532 0.01360 -0.0271 0.9999 1.0001 -2.500 -0.1012 0.02531 0.01274 -0.0261 0.9999 1.0001 -2.250 -0.0767 0.02532 0.01216 -0.0251 0.9999 1.0001 -2.000 -0.0525 0.02535 0.01170 -0.0240 0.9999 1.0001 -1.750 -0.0285 0.02539 0.01130 -0.0230 0.9999 1.0001 -1.500 -0.0044 0.02546 0.01104 -0.0221 0.9999 1.0001 -1.250 0.0197 0.02553 0.01089 -0.0211 0.9999 1.0001 -1.000 0.0440 0.02562 0.01086 -0.0201 0.9999 1.0001 -0.750 0.0687 0.02573 0.01097 -0.0191 0.9999 1.0001 -0.500 0.0942 0.02583 0.01125 -0.0180 0.9999 1.0001 -0.250 0.1220 0.02589 0.01168 -0.0169 0.9999 1.0001 0.000 0.1551 0.02583 0.01234 -0.0164 0.9999 1.0001 0.250 0.2664 0.02908 0.01324 -0.0267 0.4322 1.0001 0.500 0.2938 0.03076 0.01433 -0.0259 0.4047 1.0001 0.750 0.3246 0.03232 0.01555 -0.0256 0.3872 1.0001 1.000 0.3575 0.03397 0.01697 -0.0258 0.3756 1.0001 1.250 0.3922 0.03549 0.01858 -0.0265 0.3677 1.0001 1.500 0.4254 0.03732 0.02036 -0.0269 0.3587 1.0001 1.750 0.4580 0.03901 0.02228 -0.0276 0.3490 1.0001 2.000 0.4893 0.04101 0.02439 -0.0281 0.3405 1.0001 2.250 0.5211 0.04297 0.02671 -0.0291 0.3367 1.0001 2.500 0.5532 0.04520 0.02932 -0.0303 0.3391 1.0001 2.750 0.5843 0.04771 0.03214 -0.0315 0.3430 1.0001 3.000 0.6145 0.05011 0.03517 -0.0339 0.3511 1.0001 3.250 0.6429 0.05308 0.03849 -0.0355 0.3586 1.0001 3.500 0.6699 0.05619 0.04216 -0.0390 0.3707 1.0001 3.750 0.6949 0.05973 0.04612 -0.0427 0.3842 1.0001 4.000 0.7163 0.06366 0.05045 -0.0469 0.4005 1.0001 4.250 0.7324 0.06799 0.05508 -0.0520 0.4202 1.0001 4.500 0.7468 0.07265 0.05993 -0.0571 0.4426 1.0001 4.750 0.7427 0.07760 0.06503 -0.0638 0.4708 1.0001 5.000 0.7351 0.08265 0.07009 -0.0701 0.5025 1.0001