XFOIL Version 6.96 Calculated polar for: WORTMANN FX 3 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 ------ -------- --------- --------- -------- -------- -------- -19.750 -0.8174 0.08386 0.07972 -0.1217 0.7922 0.0058 -19.500 -0.8366 0.07733 0.07306 -0.1251 0.7922 0.0059 -19.250 -0.8498 0.07193 0.06753 -0.1279 0.7921 0.0059 -19.000 -0.8591 0.06731 0.06279 -0.1302 0.7921 0.0059 -18.750 -0.8652 0.06330 0.05868 -0.1320 0.7921 0.0060 -18.500 -0.8688 0.05976 0.05503 -0.1335 0.7920 0.0060 -18.250 -0.8697 0.05663 0.05181 -0.1347 0.7920 0.0061 -18.000 -0.8690 0.05381 0.04889 -0.1357 0.7920 0.0062 -17.750 -0.8667 0.05125 0.04624 -0.1366 0.7919 0.0062 -17.500 -0.8626 0.04895 0.04386 -0.1373 0.7919 0.0063 -17.250 -0.8570 0.04687 0.04170 -0.1378 0.7918 0.0064 -17.000 -0.8504 0.04495 0.03969 -0.1383 0.7918 0.0064 -16.750 -0.8466 0.04284 0.03750 -0.1385 0.7917 0.0066 -16.500 -0.8402 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0.0081 0.6778 13.500 1.4216 0.04854 0.04155 -0.0911 0.0078 0.6779 13.750 1.4338 0.04975 0.04281 -0.0900 0.0075 0.6779 14.000 1.4441 0.05112 0.04423 -0.0887 0.0071 0.6779 14.250 1.4555 0.05243 0.04558 -0.0876 0.0069 0.6780 14.500 1.4669 0.05374 0.04695 -0.0866 0.0067 0.6780 14.750 1.4778 0.05511 0.04838 -0.0855 0.0066 0.6781 15.000 1.4875 0.05661 0.04994 -0.0844 0.0064 0.6781 15.250 1.4974 0.05813 0.05151 -0.0834 0.0062 0.6781 15.500 1.5064 0.05973 0.05317 -0.0823 0.0060 0.6782 15.750 1.5146 0.06144 0.05495 -0.0813 0.0059 0.6782 16.000 1.5227 0.06317 0.05673 -0.0803 0.0057 0.6783 16.250 1.5290 0.06510 0.05872 -0.0792 0.0056 0.6783 16.500 1.5348 0.06708 0.06077 -0.0782 0.0054 0.6783 16.750 1.5382 0.06937 0.06314 -0.0771 0.0053 0.6784 17.000 1.5427 0.07158 0.06542 -0.0762 0.0052 0.6784 17.250 1.5470 0.07386 0.06779 -0.0753 0.0051 0.6784 17.500 1.5513 0.07616 0.07017 -0.0746 0.0051 0.6785 17.750 1.5550 0.07857 0.07267 -0.0740 0.0050 0.6785 18.000 1.5573 0.08117 0.07535 -0.0734 0.0049 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