XFOIL Version 6.94 Calculated polar for: EPPLER 545 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.3017 0.01412 0.00777 -0.0622 0.6671 0.6969 0.500 0.3573 0.01393 0.00748 -0.0631 0.6570 0.7002 1.000 0.4132 0.01378 0.00725 -0.0639 0.6471 0.7036 1.500 0.4684 0.01365 0.00708 -0.0643 0.6377 0.7060 2.000 0.5220 0.01356 0.00704 -0.0644 0.6276 0.7085 2.500 0.5783 0.01359 0.00701 -0.0650 0.6178 0.7112 3.000 0.6303 0.01349 0.00699 -0.0648 0.6053 0.7143 3.500 0.6856 0.01354 0.00701 -0.0652 0.5945 0.7182 4.000 0.7377 0.01347 0.00698 -0.0650 0.5806 0.7218 4.500 0.7905 0.01348 0.00705 -0.0649 0.5684 0.7248 5.000 0.8425 0.01348 0.00713 -0.0646 0.5566 0.7279 5.500 0.8916 0.01355 0.00737 -0.0637 0.5432 0.7315 6.000 0.9410 0.01365 0.00754 -0.0628 0.5281 0.7355 6.500 0.9884 0.01376 0.00772 -0.0616 0.5106 0.7399 7.000 1.0284 0.01387 0.00793 -0.0590 0.4850 0.7447 7.500 1.0607 0.01401 0.00818 -0.0548 0.4518 0.7494 8.000 1.0812 0.01442 0.00857 -0.0485 0.4101 0.7546 8.500 1.0870 0.01537 0.00933 -0.0400 0.3584 0.7607 9.000 1.0916 0.01670 0.01050 -0.0321 0.3097 0.7667 9.500 1.0947 0.01830 0.01202 -0.0248 0.2730 0.7736 10.000 1.1039 0.02000 0.01371 -0.0192 0.2410 0.7817 10.500 1.1075 0.02213 0.01583 -0.0135 0.2145 0.7907 11.000 1.1148 0.02443 0.01819 -0.0089 0.1889 0.8018 11.500 1.1183 0.02715 0.02089 -0.0047 0.1410 0.8138 12.000 1.1076 0.03112 0.02476 -0.0002 0.1166 0.8296 12.500 1.1072 0.03460 0.02844 0.0033 0.0980 0.8533 13.000 1.1172 0.03838 0.03246 0.0039 0.0774 0.9414