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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.250 -0.4328 0.08400 0.08184 -0.0451 1.0000 0.0186 -9.000 -0.4370 0.07885 0.07672 -0.0478 1.0000 0.0187 -8.750 -0.4480 0.07354 0.07147 -0.0498 1.0000 0.0187 -8.000 -0.4818 0.05603 0.05366 -0.0640 0.9929 0.0187 -7.750 -0.4680 0.05049 0.04778 -0.0696 0.9877 0.0188 -7.500 -0.4486 0.04555 0.04246 -0.0739 0.9847 0.0189 -7.250 -0.4453 0.03501 0.03139 -0.0778 0.9798 0.0201 -7.000 -0.4204 0.03199 0.02824 -0.0796 0.9765 0.0209 -6.750 -0.3917 0.02934 0.02537 -0.0817 0.9742 0.0220 -6.500 -0.3605 0.02679 0.02252 -0.0838 0.9724 0.0238 -6.250 -0.3288 0.02133 0.01627 -0.0841 0.9706 0.0170 -6.000 -0.3012 0.01865 0.01323 -0.0842 0.9667 0.0162 -5.750 -0.2722 0.01720 0.01158 -0.0846 0.9625 0.0167 -5.500 -0.2410 0.01612 0.01035 -0.0853 0.9594 0.0177 -5.250 -0.2102 0.01465 0.00872 -0.0859 0.9566 0.0178 -5.000 -0.1845 0.01373 0.00770 -0.0854 0.9509 0.0183 -4.750 -0.1566 0.01306 0.00695 -0.0853 0.9460 0.0190 -4.500 -0.1306 0.01159 0.00536 -0.0852 0.9419 0.0206 -4.250 -0.1056 0.01106 0.00480 -0.0847 0.9349 0.0234 -4.000 -0.0782 0.01059 0.00423 -0.0845 0.9297 0.0260 -3.750 -0.0517 0.01019 0.00374 -0.0841 0.9234 0.0286 -3.500 -0.0249 0.00969 0.00318 -0.0838 0.9171 0.0407 -3.250 -0.0006 0.00845 0.00276 -0.0838 0.9107 0.2374 -3.000 0.0256 0.00800 0.00257 -0.0837 0.9036 0.3281 -2.750 0.0522 0.00763 0.00245 -0.0836 0.8971 0.4085 -2.500 0.0790 0.00736 0.00235 -0.0835 0.8896 0.4748 -2.250 0.1056 0.00706 0.00227 -0.0833 0.8825 0.5457 -2.000 0.1322 0.00682 0.00219 -0.0831 0.8748 0.6051 -1.750 0.1593 0.00668 0.00215 -0.0829 0.8668 0.6550 -1.500 0.1865 0.00657 0.00209 -0.0826 0.8590 0.6938 -1.250 0.2137 0.00649 0.00205 -0.0824 0.8497 0.7253 -1.000 0.2408 0.00641 0.00199 -0.0821 0.8411 0.7536 -0.750 0.2677 0.00632 0.00194 -0.0818 0.8316 0.7809 -0.500 0.2940 0.00623 0.00191 -0.0814 0.8212 0.8118 -0.250 0.3186 0.00612 0.00190 -0.0804 0.8108 0.8524 0.000 0.3412 0.00604 0.00190 -0.0789 0.8000 0.9012 0.250 0.3664 0.00601 0.00187 -0.0779 0.7886 0.9467 0.500 0.4017 0.00600 0.00183 -0.0795 0.7761 0.9793 0.750 0.4419 0.00601 0.00179 -0.0823 0.7629 1.0000 1.000 0.4682 0.00607 0.00177 -0.0821 0.7487 1.0000 1.250 0.4948 0.00615 0.00177 -0.0819 0.7340 1.0000 1.500 0.5215 0.00625 0.00179 -0.0818 0.7185 1.0000 1.750 0.5484 0.00635 0.00184 -0.0816 0.7018 1.0000 2.000 0.5752 0.00647 0.00189 -0.0815 0.6849 1.0000 2.250 0.6021 0.00660 0.00196 -0.0814 0.6689 1.0000 2.500 0.6291 0.00674 0.00205 -0.0813 0.6536 1.0000 2.750 0.6561 0.00688 0.00215 -0.0813 0.6388 1.0000 3.000 0.6831 0.00704 0.00227 -0.0812 0.6244 1.0000 3.250 0.7100 0.00719 0.00240 -0.0812 0.6100 1.0000 3.500 0.7365 0.00738 0.00255 -0.0810 0.5924 1.0000 3.750 0.7624 0.00758 0.00269 -0.0807 0.5669 1.0000 4.000 0.7871 0.00786 0.00284 -0.0802 0.5314 1.0000 4.250 0.8114 0.00819 0.00302 -0.0797 0.4928 1.0000 4.500 0.8350 0.00860 0.00326 -0.0790 0.4450 1.0000 4.750 0.8565 0.00922 0.00358 -0.0781 0.3790 1.0000 5.000 0.8697 0.01071 0.00426 -0.0761 0.2248 1.0000 5.250 0.8760 0.01311 0.00562 -0.0730 0.0304 1.0000 5.500 0.8982 0.01381 0.00637 -0.0719 0.0236 1.0000 5.750 0.9217 0.01430 0.00692 -0.0713 0.0207 1.0000 6.000 0.9433 0.01498 0.00767 -0.0703 0.0185 1.0000 6.250 0.9589 0.01623 0.00904 -0.0682 0.0168 1.0000 6.500 0.9764 0.01722 0.01013 -0.0665 0.0161 1.0000 6.750 0.9952 0.01807 0.01106 -0.0650 0.0154 1.0000 7.000 1.0142 0.01888 0.01193 -0.0636 0.0144 1.0000 7.250 1.0327 0.01972 0.01283 -0.0622 0.0134 1.0000 7.500 1.0495 0.02081 0.01399 -0.0605 0.0129 1.0000 7.750 1.0664 0.02199 0.01523 -0.0589 0.0123 1.0000 8.000 1.0838 0.02340 0.01670 -0.0574 0.0119 1.0000 8.250 1.1037 0.02543 0.01879 -0.0563 0.0115 1.0000 8.500 1.1277 0.02779 0.02127 -0.0557 0.0114 1.0000 19.000 0.6850 0.20740 0.20547 -0.0662 0.0109 1.0000 19.250 0.6870 0.21093 0.20901 -0.0677 0.0109 1.0000