XFOIL Version 6.96 Calculated polar for: GOE 419 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 ------ -------- --------- --------- -------- -------- -------- -9.750 -0.3907 0.11335 0.11173 -0.0174 1.0000 0.0019 -9.500 -0.3862 0.11045 0.10884 -0.0180 1.0000 0.0019 -9.000 -0.3784 0.10448 0.10289 -0.0192 1.0000 0.0019 -6.500 -0.2568 0.06772 0.06617 -0.0538 0.9509 0.0036 -6.250 -0.2172 0.06197 0.06036 -0.0641 0.9405 0.0036 -6.000 -0.1690 0.05579 0.05409 -0.0762 0.9330 0.0036 -5.500 -0.0759 0.02432 0.02226 -0.0875 0.8523 0.0038 -5.000 -0.0278 0.03649 0.03405 -0.1022 0.8637 0.0031 -4.750 -0.0042 0.03277 0.03015 -0.1035 0.8522 0.0029 -4.500 0.0202 0.02841 0.02556 -0.1043 0.8427 0.0026 -4.250 0.0440 0.02170 0.01841 -0.1040 0.8339 0.0022 -4.000 0.0631 0.01295 0.00878 -0.1016 0.8262 0.0019 -3.750 0.0863 0.01050 0.00573 -0.1004 0.8175 0.0019 -3.500 0.1104 0.00901 0.00383 -0.0994 0.8081 0.0020 -3.250 0.1355 0.00831 0.00294 -0.0988 0.7982 0.0023 -3.000 0.1609 0.00791 0.00240 -0.0982 0.7875 0.0027 -2.750 0.1866 0.00761 0.00198 -0.0977 0.7774 0.0035 -2.500 0.2121 0.00730 0.00156 -0.0972 0.7664 0.0050 -2.250 0.2378 0.00708 0.00122 -0.0967 0.7536 0.0078 -2.000 0.2630 0.00690 0.00104 -0.0961 0.7371 0.0214 -1.750 0.2884 0.00689 0.00094 -0.0956 0.7151 0.0299 -1.500 0.3134 0.00690 0.00083 -0.0951 0.6908 0.0348 -1.250 0.3385 0.00691 0.00074 -0.0945 0.6691 0.0377 -1.000 0.3639 0.00694 0.00066 -0.0941 0.6514 0.0393 -0.750 0.3896 0.00695 0.00059 -0.0937 0.6372 0.0447 -0.500 0.4154 0.00692 0.00057 -0.0933 0.6252 0.0661 -0.250 0.4415 0.00690 0.00056 -0.0931 0.6146 0.0850 0.000 0.4675 0.00688 0.00057 -0.0928 0.6049 0.1084 0.250 0.4934 0.00685 0.00059 -0.0925 0.5954 0.1495 0.500 0.5194 0.00678 0.00063 -0.0923 0.5860 0.2038 0.750 0.5452 0.00674 0.00068 -0.0920 0.5762 0.2544 1.000 0.5707 0.00665 0.00074 -0.0917 0.5657 0.3324 1.500 0.6480 0.00543 0.00100 -0.0975 0.5307 1.0000 1.750 0.6684 0.00581 0.00110 -0.0961 0.4678 1.0000 2.000 0.6889 0.00622 0.00128 -0.0947 0.4122 1.0000 2.250 0.7092 0.00669 0.00148 -0.0933 0.3494 1.0000 2.500 0.7179 0.00820 0.00204 -0.0900 0.1372 1.0000 2.750 0.7350 0.00911 0.00253 -0.0880 0.0117 1.0000 3.000 0.7590 0.00939 0.00285 -0.0873 0.0068 1.0000 3.250 0.7830 0.00965 0.00312 -0.0866 0.0046 1.0000 3.750 0.8292 0.01041 0.00401 -0.0847 0.0029 1.0000 4.000 0.8517 0.01085 0.00448 -0.0838 0.0022 1.0000 4.250 0.8734 0.01140 0.00514 -0.0825 0.0018 1.0000 4.500 0.8932 0.01214 0.00599 -0.0809 0.0015 1.0000 4.750 0.9126 0.01292 0.00687 -0.0793 0.0013 1.0000 5.000 0.9292 0.01407 0.00813 -0.0770 0.0012 1.0000 5.250 0.9446 0.01578 0.00996 -0.0744 0.0012 1.0000 5.500 0.9669 0.01886 0.01314 -0.0723 0.0013 1.0000 5.750 0.9946 0.02213 0.01661 -0.0709 0.0016 1.0000 6.000 1.0168 0.02476 0.01948 -0.0693 0.0020 1.0000 7.000 1.0772 0.03674 0.03244 -0.0606 0.0023 1.0000 7.250 1.0884 0.03974 0.03570 -0.0579 0.0023 1.0000 7.500 1.0975 0.04275 0.03896 -0.0552 0.0023 1.0000 7.750 1.1043 0.04586 0.04231 -0.0523 0.0023 1.0000 8.000 1.1089 0.04904 0.04571 -0.0494 0.0023 1.0000 8.250 1.1104 0.05233 0.04922 -0.0464 0.0023 1.0000 8.500 1.1092 0.05554 0.05263 -0.0433 0.0023 1.0000 8.750 1.1043 0.05882 0.05610 -0.0401 0.0023 1.0000 9.000 1.0963 0.06193 0.05936 -0.0369 0.0023 1.0000 9.250 1.0812 0.06455 0.06212 -0.0326 0.0023 1.0000 9.500 1.0646 0.06731 0.06500 -0.0293 0.0023 1.0000 9.750 1.0468 0.07055 0.06836 -0.0273 0.0023 1.0000