XFOIL Version 6.96 Calculated polar for: GOE 437 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 ------ -------- --------- --------- -------- -------- -------- -8.000 -0.3211 0.09775 0.09438 -0.0286 1.0000 0.0315 -7.750 -0.3263 0.09623 0.09294 -0.0286 1.0000 0.0317 -7.500 -0.3340 0.09490 0.09170 -0.0279 1.0000 0.0319 -7.250 -0.3355 0.09285 0.08971 -0.0287 1.0000 0.0320 -7.000 -0.3356 0.09061 0.08751 -0.0296 1.0000 0.0321 -6.750 -0.3345 0.08820 0.08515 -0.0304 1.0000 0.0322 -6.500 -0.3330 0.08573 0.08272 -0.0310 1.0000 0.0323 -6.250 -0.2903 0.06827 0.06562 -0.0238 0.9954 0.0344 -6.000 -0.2689 0.06347 0.06080 -0.0283 0.9903 0.0358 -5.750 -0.2451 0.05839 0.05569 -0.0348 0.9849 0.0378 -5.500 -0.2200 0.05326 0.05052 -0.0421 0.9787 0.0402 -5.250 -0.1715 0.04771 0.04474 -0.0596 0.9720 0.0443 -5.000 -0.1512 0.04031 0.03732 -0.0643 0.9678 0.0459 -4.750 -0.1310 0.03700 0.03400 -0.0647 0.9609 0.0487 -4.500 -0.0949 0.03227 0.02914 -0.0712 0.9571 0.0527 -4.250 -0.1144 0.04653 0.04307 -0.0713 0.9651 0.0521 -4.000 -0.0604 0.04051 0.03654 -0.0808 0.9591 0.0588 -3.750 -0.0265 0.03727 0.03331 -0.0841 0.9559 0.0617 -3.500 0.0137 0.03370 0.02922 -0.0880 0.9474 0.0718 -3.250 0.0495 0.03070 0.02624 -0.0910 0.9437 0.0757 -3.000 0.0951 0.02172 0.01598 -0.0932 0.9392 0.0427 -2.750 0.1311 0.01873 0.01238 -0.0944 0.9327 0.0430 -2.500 0.1726 0.01647 0.00956 -0.0964 0.9286 0.0449 -2.250 0.2043 0.01496 0.00793 -0.0969 0.9184 0.0535 -2.000 0.2384 0.01355 0.00635 -0.0975 0.9095 0.0622 -1.750 0.2732 0.01294 0.00571 -0.0984 0.9009 0.0792 -1.500 0.3031 0.01257 0.00523 -0.0983 0.8890 0.0974 -1.250 0.3321 0.01197 0.00469 -0.0982 0.8768 0.1184 -1.000 0.3609 0.01147 0.00423 -0.0981 0.8641 0.1534 -0.750 0.3891 0.01110 0.00389 -0.0978 0.8505 0.1833 -0.500 0.4167 0.01081 0.00361 -0.0975 0.8360 0.2113 -0.250 0.4435 0.01047 0.00337 -0.0970 0.8208 0.2528 0.000 0.4887 0.00855 0.00317 -0.1005 0.8063 1.0000 0.250 0.5151 0.00864 0.00302 -0.0997 0.7895 1.0000 0.500 0.5404 0.00876 0.00295 -0.0989 0.7709 1.0000 0.750 0.5658 0.00889 0.00291 -0.0981 0.7524 1.0000 1.000 0.5913 0.00904 0.00288 -0.0973 0.7346 1.0000 1.250 0.6161 0.00919 0.00291 -0.0965 0.7150 1.0000 1.500 0.6410 0.00935 0.00294 -0.0957 0.6957 1.0000 1.750 0.6660 0.00952 0.00297 -0.0949 0.6769 1.0000 2.000 0.6908 0.00967 0.00303 -0.0941 0.6572 1.0000 2.250 0.7157 0.00984 0.00309 -0.0934 0.6387 1.0000 2.500 0.7406 0.01001 0.00319 -0.0927 0.6203 1.0000 2.750 0.7654 0.01020 0.00329 -0.0919 0.6017 1.0000 3.000 0.7902 0.01043 0.00342 -0.0912 0.5846 1.0000 3.250 0.8150 0.01070 0.00358 -0.0906 0.5689 1.0000 3.500 0.8399 0.01100 0.00379 -0.0899 0.5537 1.0000 3.750 0.8648 0.01132 0.00407 -0.0893 0.5387 1.0000 4.000 0.8898 0.01166 0.00436 -0.0888 0.5246 1.0000 4.250 0.9149 0.01201 0.00467 -0.0882 0.5113 1.0000 4.500 0.9398 0.01234 0.00501 -0.0877 0.4981 1.0000 4.750 0.9646 0.01265 0.00539 -0.0871 0.4852 1.0000 5.000 0.9885 0.01293 0.00566 -0.0863 0.4698 1.0000 5.250 1.0102 0.01302 0.00580 -0.0851 0.4457 1.0000 5.500 1.0324 0.01318 0.00596 -0.0839 0.4225 1.0000 5.750 1.0543 0.01339 0.00617 -0.0828 0.3983 1.0000 6.000 1.0737 0.01369 0.00639 -0.0812 0.3551 1.0000 6.250 1.0895 0.01442 0.00679 -0.0791 0.2779 1.0000 6.500 1.0863 0.01752 0.00840 -0.0750 0.0536 1.0000 6.750 1.1014 0.01883 0.00965 -0.0729 0.0360 1.0000 7.000 1.1176 0.01996 0.01097 -0.0709 0.0316 1.0000 7.250 1.1334 0.02102 0.01223 -0.0688 0.0298 1.0000 7.500 1.1463 0.02226 0.01362 -0.0664 0.0284 1.0000 7.750 1.1572 0.02363 0.01510 -0.0637 0.0275 1.0000 8.000 1.1672 0.02507 0.01664 -0.0611 0.0259 1.0000 8.250 1.1742 0.02696 0.01854 -0.0583 0.0239 1.0000 8.500 1.1870 0.02905 0.02065 -0.0561 0.0233 1.0000 8.750 1.2095 0.03160 0.02325 -0.0554 0.0230 1.0000 9.000 1.2351 0.03385 0.02560 -0.0549 0.0231 1.0000 9.250 1.2586 0.03563 0.02759 -0.0539 0.0239 1.0000 9.500 1.2869 0.03864 0.03111 -0.0527 0.0277 1.0000 9.750 1.3249 0.04384 0.03654 -0.0534 0.0337 1.0000