XFOIL Version 6.96 Calculated polar for: GOE 673 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.500 -0.3228 0.09741 0.09290 -0.0379 1.0000 0.1233 -9.250 -0.3379 0.09456 0.09014 -0.0392 1.0000 0.1285 -9.000 -0.3753 0.09234 0.08808 -0.0422 1.0000 0.1299 -8.750 -0.3377 0.08746 0.08313 -0.0375 1.0000 0.1347 -8.500 -0.3436 0.08460 0.08034 -0.0367 1.0000 0.1398 -8.250 -0.3835 0.08267 0.07857 -0.0379 1.0000 0.1432 -8.000 -0.3778 0.07852 0.07446 -0.0355 1.0000 0.1461 -7.750 -0.3727 0.07580 0.07176 -0.0328 1.0000 0.1502 -7.500 -0.5144 0.08150 0.07727 -0.0414 1.0000 0.1443 -7.250 -0.4842 0.07910 0.07494 -0.0332 1.0000 0.1487 -5.500 -0.4742 0.03780 0.03063 -0.0448 1.0000 0.0637 -5.250 -0.4531 0.03356 0.02623 -0.0448 1.0000 0.0601 -5.000 -0.4287 0.03040 0.02254 -0.0445 1.0000 0.0570 -4.750 -0.4035 0.02835 0.02006 -0.0442 1.0000 0.0583 -4.500 -0.3774 0.02659 0.01792 -0.0438 1.0000 0.0597 -4.250 -0.3507 0.02491 0.01592 -0.0434 1.0000 0.0599 -4.000 -0.3246 0.02358 0.01441 -0.0430 1.0000 0.0613 -3.750 -0.2991 0.02253 0.01325 -0.0426 1.0000 0.0638 -3.500 -0.2736 0.02138 0.01222 -0.0427 1.0000 0.0726 -3.250 -0.2472 0.02050 0.01132 -0.0429 1.0000 0.0822 -3.000 -0.2176 0.01940 0.01054 -0.0437 1.0000 0.1285 -2.750 -0.1863 0.01759 0.01055 -0.0454 1.0000 0.5052 -2.500 -0.1641 0.01740 0.01097 -0.0442 0.9991 0.6500 -2.250 -0.1326 0.01739 0.01132 -0.0436 0.9927 0.7852 -2.000 -0.1039 0.01737 0.01143 -0.0425 0.9848 0.8873 -1.750 -0.0523 0.01739 0.01118 -0.0479 0.9746 1.0000 -1.500 -0.0044 0.01769 0.01112 -0.0531 0.9670 1.0000 -1.250 0.0401 0.01793 0.01107 -0.0573 0.9582 1.0000 -1.000 0.0802 0.01815 0.01108 -0.0605 0.9485 1.0000 -0.750 0.1293 0.01836 0.01108 -0.0653 0.9420 1.0000 -0.500 0.1650 0.01852 0.01108 -0.0674 0.9309 1.0000 -0.250 0.2026 0.01868 0.01112 -0.0698 0.9208 1.0000 0.000 0.2507 0.01871 0.01105 -0.0740 0.9144 1.0000 0.250 0.2845 0.01882 0.01107 -0.0755 0.9031 1.0000 0.500 0.3209 0.01889 0.01108 -0.0774 0.8929 1.0000 0.750 0.3692 0.01873 0.01089 -0.0813 0.8869 1.0000 1.000 0.4011 0.01876 0.01091 -0.0821 0.8755 1.0000 1.250 0.4348 0.01874 0.01088 -0.0832 0.8651 1.0000 1.500 0.4776 0.01844 0.01057 -0.0856 0.8586 1.0000 1.750 0.5071 0.01842 0.01057 -0.0857 0.8467 1.0000 2.000 0.5374 0.01836 0.01056 -0.0859 0.8352 1.0000 2.250 0.5692 0.01822 0.01046 -0.0862 0.8245 1.0000 2.500 0.6063 0.01784 0.01011 -0.0872 0.8164 1.0000 2.750 0.6344 0.01776 0.01012 -0.0868 0.8036 1.0000 3.000 0.6628 0.01767 0.01009 -0.0864 0.7907 1.0000 3.250 0.6914 0.01756 0.01005 -0.0860 0.7775 1.0000 3.500 0.7200 0.01746 0.01001 -0.0855 0.7639 1.0000 3.750 0.7483 0.01737 0.01004 -0.0851 0.7499 1.0000 4.000 0.7766 0.01732 0.01006 -0.0846 0.7354 1.0000 4.250 0.8047 0.01731 0.01014 -0.0841 0.7206 1.0000 4.500 0.8327 0.01733 0.01027 -0.0835 0.7055 1.0000 4.750 0.8600 0.01744 0.01052 -0.0830 0.6897 1.0000 5.000 0.8844 0.01708 0.01018 -0.0809 0.6604 1.0000 5.250 0.9036 0.01654 0.00952 -0.0775 0.6115 1.0000 5.500 0.9217 0.01647 0.00936 -0.0745 0.5608 1.0000 5.750 0.9308 0.01676 0.00930 -0.0699 0.4610 1.0000 6.000 0.9166 0.01975 0.01030 -0.0629 0.1535 1.0000 6.250 0.9194 0.02245 0.01227 -0.0591 0.0787 1.0000 6.500 0.9305 0.02406 0.01396 -0.0564 0.0673 1.0000 6.750 0.9416 0.02553 0.01550 -0.0538 0.0599 1.0000 7.000 0.9500 0.02733 0.01732 -0.0507 0.0568 1.0000 7.250 0.9674 0.02887 0.01891 -0.0488 0.0544 1.0000 7.500 0.9926 0.03071 0.02075 -0.0480 0.0520 1.0000 7.750 1.0207 0.03284 0.02282 -0.0480 0.0479 1.0000 8.000 1.0614 0.03654 0.02651 -0.0498 0.0462 1.0000 8.250 1.0925 0.03925 0.02946 -0.0496 0.0465 1.0000 8.500 1.1180 0.04145 0.03211 -0.0482 0.0482 1.0000 8.750 1.1375 0.04490 0.03625 -0.0458 0.0522 1.0000 9.000 1.1559 0.04913 0.04087 -0.0442 0.0550 1.0000 9.250 1.1724 0.05450 0.04650 -0.0431 0.0568 1.0000 9.500 1.2049 0.06179 0.05499 -0.0394 0.0965 1.0000 9.750 1.1004 0.05152 0.04548 -0.0248 0.0774 1.0000