XFOIL Version 6.96 Calculated polar for: GOE 483 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -6.500 -0.3117 0.08802 0.08509 -0.0230 1.0000 0.0279 -6.250 -0.3198 0.08603 0.08315 -0.0191 1.0000 0.0283 -6.000 -0.3106 0.08295 0.08010 -0.0197 0.9980 0.0291 -5.750 -0.2808 0.07859 0.07571 -0.0258 0.9931 0.0304 -5.500 -0.2490 0.07432 0.07140 -0.0329 0.9871 0.0321 -5.250 -0.2109 0.06983 0.06685 -0.0416 0.9827 0.0343 -5.000 -0.1697 0.06577 0.06272 -0.0514 0.9751 0.0372 -4.750 -0.1021 0.06185 0.05858 -0.0675 0.9704 0.0387 -4.500 -0.0765 0.05551 0.05220 -0.0723 0.9643 0.0397 -4.250 -0.0520 0.05191 0.04861 -0.0741 0.9595 0.0418 -4.000 0.0536 0.02716 0.02377 -0.0895 0.9416 0.0521 -3.750 0.0502 0.04330 0.03957 -0.0916 0.9504 0.0523 -3.500 0.0730 0.04077 0.03710 -0.0922 0.9450 0.0573 -3.250 0.1288 0.03628 0.03227 -0.1005 0.9413 0.0657 -3.000 0.1532 0.03402 0.02997 -0.1011 0.9324 0.0686 -2.750 0.1996 0.03045 0.02601 -0.1059 0.9274 0.0789 -2.500 0.2357 0.02260 0.01754 -0.1079 0.9186 0.0457 -2.250 0.2713 0.01854 0.01275 -0.1092 0.9125 0.0479 -2.000 0.3004 0.01597 0.00952 -0.1088 0.9029 0.0488 -1.750 0.3289 0.01436 0.00749 -0.1083 0.8937 0.0536 -1.500 0.3572 0.01374 0.00672 -0.1078 0.8847 0.0629 -1.250 0.3833 0.01301 0.00589 -0.1069 0.8736 0.0727 -1.000 0.4092 0.01258 0.00539 -0.1060 0.8622 0.0858 -0.750 0.4353 0.01221 0.00492 -0.1051 0.8510 0.1042 -0.500 0.4610 0.01178 0.00454 -0.1042 0.8397 0.1338 -0.250 0.4871 0.01167 0.00439 -0.1034 0.8281 0.1600 0.250 0.5388 0.01166 0.00430 -0.1018 0.8024 0.2025 0.500 0.5644 0.01158 0.00422 -0.1011 0.7885 0.2217 0.750 0.5902 0.01145 0.00404 -0.1003 0.7742 0.2420 1.000 0.6156 0.01127 0.00390 -0.0995 0.7588 0.2642 1.250 0.6409 0.01108 0.00375 -0.0988 0.7422 0.2885 1.500 0.6661 0.01089 0.00361 -0.0981 0.7251 0.3161 1.750 0.6913 0.01069 0.00349 -0.0973 0.7080 0.3553 2.000 0.7217 0.00949 0.00338 -0.0976 0.6909 1.0000 2.250 0.7476 0.00961 0.00336 -0.0969 0.6716 1.0000 2.500 0.7733 0.00975 0.00336 -0.0963 0.6531 1.0000 2.750 0.7990 0.00992 0.00339 -0.0956 0.6353 1.0000 3.000 0.8246 0.01010 0.00348 -0.0950 0.6164 1.0000 3.250 0.8500 0.01032 0.00361 -0.0944 0.5982 1.0000 3.500 0.8753 0.01057 0.00375 -0.0938 0.5809 1.0000 3.750 0.9006 0.01084 0.00393 -0.0932 0.5637 1.0000 4.000 0.9259 0.01112 0.00416 -0.0927 0.5472 1.0000 4.250 0.9512 0.01143 0.00442 -0.0921 0.5315 1.0000 4.500 0.9764 0.01174 0.00475 -0.0916 0.5162 1.0000 4.750 1.0016 0.01207 0.00507 -0.0911 0.5014 1.0000 5.000 1.0265 0.01238 0.00538 -0.0905 0.4857 1.0000 5.250 1.0496 0.01259 0.00557 -0.0896 0.4591 1.0000 5.500 1.0728 0.01282 0.00581 -0.0887 0.4317 1.0000 5.750 1.0964 0.01309 0.00610 -0.0879 0.4077 1.0000 6.000 1.1178 0.01345 0.00634 -0.0867 0.3643 1.0000 6.250 1.1362 0.01419 0.00675 -0.0853 0.2932 1.0000 6.500 1.1457 0.01624 0.00780 -0.0829 0.1367 1.0000 6.750 1.1539 0.01874 0.00957 -0.0804 0.0360 1.0000 7.000 1.1724 0.01988 0.01085 -0.0789 0.0299 1.0000 7.250 1.1907 0.02096 0.01218 -0.0773 0.0277 1.0000 7.500 1.2062 0.02224 0.01363 -0.0754 0.0262 1.0000 7.750 1.2192 0.02371 0.01523 -0.0732 0.0252 1.0000 8.000 1.2306 0.02535 0.01696 -0.0708 0.0245 1.0000 8.250 1.2407 0.02728 0.01892 -0.0685 0.0231 1.0000 8.500 1.2531 0.03055 0.02220 -0.0665 0.0216 1.0000 8.750 1.2772 0.03346 0.02519 -0.0658 0.0215 1.0000 9.000 1.3012 0.03569 0.02760 -0.0651 0.0218 1.0000 9.250 1.3229 0.03729 0.02938 -0.0639 0.0224 1.0000 9.500 1.3452 0.03906 0.03159 -0.0619 0.0250 1.0000 15.000 0.8849 0.17917 0.17645 -0.0933 0.0673 1.0000 15.250 0.9015 0.18080 0.17812 -0.0914 0.0653 1.0000