XFOIL Version 6.96 Calculated polar for: GOE 434 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.000 -0.8750 0.09227 0.08847 -0.0783 1.0000 0.0500 -18.750 -0.8775 0.08864 0.08482 -0.0803 1.0000 0.0505 -18.250 -0.8760 0.08159 0.07769 -0.0852 0.9798 0.0515 -18.000 -0.8727 0.07853 0.07459 -0.0872 0.9694 0.0520 -17.750 -0.8732 0.07531 0.07131 -0.0888 0.9586 0.0525 -17.500 -0.8723 0.07249 0.06843 -0.0899 0.9470 0.0530 -17.250 -0.8705 0.06986 0.06573 -0.0909 0.9350 0.0535 -17.000 -0.8721 0.06687 0.06265 -0.0920 0.9191 0.0539 -16.750 -0.8698 0.06444 0.06010 -0.0926 0.8917 0.0543 -16.500 -0.8681 0.06203 0.05748 -0.0933 0.8518 0.0546 -16.250 -0.8702 0.05892 0.05419 -0.0947 0.8205 0.0550 -16.000 -0.8690 0.05615 0.05130 -0.0962 0.7979 0.0555 -15.750 -0.8706 0.05300 0.04803 -0.0979 0.7798 0.0560 -15.500 -0.8681 0.05037 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0.01157 0.00479 -0.1407 0.3395 0.3508 4.000 1.0988 0.01199 0.00511 -0.1394 0.3162 0.3563 4.250 1.1140 0.01260 0.00555 -0.1374 0.2855 0.3611 4.500 1.1237 0.01324 0.00603 -0.1343 0.2574 0.3644 4.750 1.1388 0.01381 0.00651 -0.1322 0.2385 0.3686 5.000 1.1569 0.01430 0.00695 -0.1307 0.2260 0.3742 5.250 1.1766 0.01472 0.00735 -0.1295 0.2181 0.3796 5.500 1.1956 0.01518 0.00779 -0.1282 0.2111 0.3847 5.750 1.2152 0.01561 0.00821 -0.1271 0.2061 0.3893 6.000 1.2357 0.01601 0.00862 -0.1261 0.2022 0.3957 6.250 1.2549 0.01646 0.00910 -0.1249 0.1983 0.4037 6.500 1.2729 0.01700 0.00963 -0.1237 0.1941 0.4109 6.750 1.2922 0.01748 0.01013 -0.1226 0.1912 0.4166 7.000 1.3114 0.01798 0.01066 -0.1216 0.1884 0.4253 7.250 1.3297 0.01855 0.01126 -0.1206 0.1853 0.4342 7.500 1.3466 0.01922 0.01194 -0.1194 0.1824 0.4419 7.750 1.3632 0.01992 0.01267 -0.1183 0.1796 0.4517 8.000 1.3803 0.02063 0.01342 -0.1173 0.1775 0.4627 8.250 1.3984 0.02130 0.01413 -0.1164 0.1759 0.4726 8.500 1.4155 0.02206 0.01495 -0.1155 0.1738 0.4867 8.750 1.4318 0.02289 0.01582 -0.1146 0.1718 0.4995 9.000 1.4471 0.02381 0.01680 -0.1137 0.1698 0.5150 9.250 1.4612 0.02486 0.01788 -0.1127 0.1678 0.5269 9.500 1.4749 0.02596 0.01902 -0.1117 0.1657 0.5401 9.750 1.4884 0.02711 0.02021 -0.1108 0.1640 0.5508 10.000 1.5035 0.02813 0.02130 -0.1101 0.1629 0.5625 10.250 1.5171 0.02930 0.02253 -0.1093 0.1615 0.5736 10.500 1.5304 0.03051 0.02378 -0.1085 0.1599 0.5839 10.750 1.5423 0.03186 0.02519 -0.1077 0.1582 0.5963 11.000 1.5537 0.03326 0.02664 -0.1068 0.1564 0.6083 11.250 1.5628 0.03489 0.02832 -0.1060 0.1544 0.6218 11.500 1.5722 0.03649 0.02997 -0.1051 0.1522 0.6348 11.750 1.5826 0.03805 0.03160 -0.1044 0.1508 0.6496 12.000 1.5937 0.03954 0.03317 -0.1037 0.1491 0.6669 12.250 1.6032 0.04119 0.03490 -0.1030 0.1470 0.6867 12.500 1.6116 0.04296 0.03674 -0.1022 0.1450 0.7120 12.750 1.6194 0.04475 0.03861 -0.1015 0.1427 0.7413 13.000 1.6246 0.04685 0.04079 -0.1007 0.1405 0.7685 13.250 1.6346 0.04846 0.04250 -0.1001 0.1389 0.7992 13.500 1.6428 0.05012 0.04438 -0.0993 0.1367 0.8688 13.750 1.6460 0.05200 0.04640 -0.0981 0.1343 1.0000 14.000 1.6519 0.05405 0.04845 -0.0974 0.1315 1.0000 14.250 1.6561 0.05632 0.05074 -0.0968 0.1289 1.0000 14.500 1.6616 0.05847 0.05290 -0.0962 0.1252 1.0000 14.750 1.6649 0.06081 0.05523 -0.0955 0.1208 1.0000 15.000 1.6673 0.06331 0.05775 -0.0949 0.1171 1.0000 15.250 1.6680 0.06597 0.06039 -0.0943 0.1115 1.0000 15.500 1.6678 0.06875 0.06316 -0.0937 0.1047 1.0000 15.750 1.6617 0.07220 0.06658 -0.0931 0.0963 1.0000 16.000 1.6562 0.07558 0.06993 -0.0925 0.0885 1.0000 16.250 1.6525 0.07887 0.07322 -0.0921 0.0844 1.0000 16.500 1.6522 0.08176 0.07612 -0.0917 0.0812 1.0000 16.750 1.6515 0.08479 0.07918 -0.0915 0.0793 1.0000 17.000 1.6538 0.08743 0.08184 -0.0913 0.0780 1.0000 17.250 1.6557 0.09015 0.08460 -0.0911 0.0768 1.0000 17.500 1.6556 0.09314 0.08763 -0.0911 0.0757 1.0000 17.750 1.6575 0.09588 0.09041 -0.0910 0.0747 1.0000