XFOIL Version 6.96 Calculated polar for: E61 (5.64%) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.250 -0.3700 0.10529 0.10141 -0.0167 1.0000 0.0431 -5.000 -0.3720 0.10100 0.09714 -0.0111 1.0000 0.0446 -4.750 -0.3693 0.09835 0.09452 -0.0104 1.0000 0.0458 -4.500 -0.3646 0.09581 0.09201 -0.0110 1.0000 0.0472 -4.250 -0.3572 0.09322 0.08944 -0.0125 1.0000 0.0489 -4.000 -0.3463 0.09050 0.08673 -0.0152 1.0000 0.0510 -3.750 -0.3253 0.08765 0.08383 -0.0214 1.0000 0.0539 -3.500 -0.2576 0.08376 0.07981 -0.0441 1.0000 0.0563 -3.250 -0.2657 0.07991 0.07603 -0.0380 1.0000 0.0573 -3.000 -0.2474 0.07634 0.07246 -0.0384 0.9958 0.0606 -2.750 -0.1559 0.07022 0.06613 -0.0612 0.9858 0.0717 -2.250 -0.0187 0.06150 0.05701 -0.0897 0.9737 0.0970 -2.000 0.0091 0.05743 0.05294 -0.0921 0.9684 0.1020 -1.750 0.0678 0.05367 0.04902 -0.1024 0.9637 0.1167 -1.500 0.1414 0.04993 0.04500 -0.1159 0.9607 0.1420 -1.250 0.1836 0.04757 0.04257 -0.1210 0.9553 0.1712 -0.500 0.4312 0.03505 0.02760 -0.1571 0.9507 0.0785 -0.250 0.4860 0.03374 0.02547 -0.1616 0.9482 0.0699 0.000 0.5236 0.03259 0.02420 -0.1642 0.9425 0.0746 0.250 0.5653 0.03197 0.02339 -0.1668 0.9378 0.0802 0.500 0.6118 0.03133 0.02270 -0.1704 0.9345 0.0874 0.750 0.6456 0.03106 0.02243 -0.1718 0.9268 0.1102 1.000 0.6858 0.02915 0.02227 -0.1740 0.9236 1.0000 1.250 0.7149 0.02954 0.02240 -0.1747 0.9146 1.0000 1.500 0.7581 0.02972 0.02235 -0.1776 0.9096 1.0000 1.750 0.7849 0.03008 0.02259 -0.1778 0.8997 1.0000 2.000 0.8186 0.03032 0.02274 -0.1791 0.8917 1.0000 2.250 0.8572 0.03036 0.02272 -0.1811 0.8846 1.0000 2.500 0.8865 0.03060 0.02292 -0.1815 0.8748 1.0000 2.750 0.9328 0.03027 0.02257 -0.1845 0.8697 1.0000 3.000 0.9609 0.03042 0.02273 -0.1846 0.8587 1.0000 3.250 0.9933 0.03039 0.02279 -0.1852 0.8489 1.0000 3.500 1.0412 0.02963 0.02210 -0.1881 0.8433 1.0000 3.750 1.0725 0.02944 0.02199 -0.1883 0.8321 1.0000 4.250 1.1613 0.02745 0.02033 -0.1920 0.8168 1.0000 4.500 1.1988 0.02659 0.01963 -0.1926 0.8059 1.0000 4.750 1.2646 0.02425 0.01754 -0.1973 0.8011 1.0000 5.000 1.3116 0.02289 0.01642 -0.1993 0.7882 1.0000 5.250 1.3487 0.02195 0.01576 -0.1995 0.7702 1.0000 5.500 1.3928 0.02070 0.01474 -0.2006 0.7487 1.0000 5.750 1.4279 0.01994 0.01416 -0.2003 0.7202 1.0000 6.000 1.4524 0.01967 0.01403 -0.1983 0.6845 1.0000 6.250 1.4760 0.01953 0.01401 -0.1961 0.6387 1.0000 6.500 1.4902 0.01979 0.01406 -0.1921 0.5673 1.0000 6.750 1.4950 0.02083 0.01451 -0.1867 0.4683 1.0000 7.000 1.4915 0.02255 0.01552 -0.1807 0.3680 1.0000 7.250 1.4846 0.02459 0.01686 -0.1746 0.2738 1.0000 7.500 1.4786 0.02703 0.01856 -0.1692 0.1743 1.0000 7.750 1.4635 0.03113 0.02157 -0.1626 0.0673 1.0000 8.000 1.4594 0.03421 0.02453 -0.1573 0.0472 1.0000 8.250 1.4611 0.03671 0.02710 -0.1532 0.0405 1.0000 8.500 1.4656 0.03945 0.02996 -0.1496 0.0371 1.0000 8.750 1.4811 0.04201 0.03267 -0.1473 0.0344 1.0000 9.000 1.5024 0.04465 0.03546 -0.1461 0.0311 1.0000 9.250 1.5427 0.04921 0.04019 -0.1479 0.0275 1.0000 9.500 1.5799 0.05392 0.04538 -0.1488 0.0270 1.0000 9.750 1.6001 0.05886 0.05084 -0.1477 0.0273 1.0000 10.000 1.6084 0.06406 0.05652 -0.1452 0.0277 1.0000 10.250 1.6162 0.07113 0.06404 -0.1434 0.0284 1.0000 10.500 1.6110 0.07308 0.06632 -0.1386 0.0288 1.0000 10.750 1.6000 0.07538 0.06903 -0.1334 0.0295 1.0000 11.000 1.5781 0.07924 0.07343 -0.1278 0.0306 1.0000 11.250 1.5507 0.08422 0.07892 -0.1233 0.0317 1.0000 11.500 1.5255 0.08970 0.08480 -0.1203 0.0325 1.0000 11.750 1.5014 0.09540 0.09083 -0.1187 0.0331 1.0000