XFOIL Version 6.94 Calculated polar for: EPPLER 540 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 0.2878 0.01952 0.01396 -0.0522 0.7616 0.7893 0.500 0.3462 0.01931 0.01371 -0.0531 0.7560 0.7917 1.000 0.3879 0.01911 0.01351 -0.0513 0.7492 0.7948 1.500 0.4168 0.01866 0.01310 -0.0475 0.7404 0.7991 2.000 0.4634 0.01805 0.01244 -0.0475 0.7335 0.8034 2.500 0.5021 0.01764 0.01205 -0.0462 0.7241 0.8064 3.000 0.5533 0.01705 0.01150 -0.0460 0.7149 0.8081 3.500 0.6046 0.01677 0.01127 -0.0456 0.7055 0.8101 4.000 0.6539 0.01633 0.01093 -0.0449 0.6940 0.8116 4.500 0.7045 0.01598 0.01066 -0.0445 0.6822 0.8133 5.000 0.7608 0.01549 0.01021 -0.0451 0.6686 0.8150 5.500 0.8032 0.01508 0.00995 -0.0431 0.6515 0.8172 6.000 0.8459 0.01467 0.00966 -0.0409 0.6310 0.8203 6.500 0.8865 0.01427 0.00933 -0.0385 0.6046 0.8225 7.000 0.9120 0.01404 0.00916 -0.0331 0.5666 0.8248 7.500 0.9227 0.01414 0.00910 -0.0249 0.5079 0.8271 8.000 0.9171 0.01503 0.00967 -0.0143 0.4347 0.8299 8.500 0.9072 0.01653 0.01087 -0.0041 0.3658 0.8334 9.000 0.9008 0.01839 0.01248 0.0044 0.3011 0.8372 9.500 0.9013 0.02045 0.01431 0.0108 0.2413 0.8403 10.000 0.9060 0.02270 0.01635 0.0159 0.1870 0.8434 10.500 0.9130 0.02514 0.01858 0.0201 0.1380 0.8461 11.000 0.9197 0.02767 0.02094 0.0242 0.0974 0.8492 11.500 0.9256 0.03048 0.02364 0.0280 0.0683 0.8525 12.000 0.9335 0.03346 0.02660 0.0311 0.0502 0.8561 12.500 0.9435 0.03651 0.02970 0.0336 0.0389 0.8605 13.000 0.9546 0.03976 0.03302 0.0354 0.0305 0.8647 13.500 0.9629 0.04343 0.03679 0.0369 0.0238 0.8685 15.500 0.9776 0.06243 0.05656 0.0394 0.0104 0.8894 16.000 0.9796 0.06805 0.06235 0.0390 0.0097 0.8965 16.500 0.9817 0.07438 0.06908 0.0378 0.0089 0.9050 17.000 0.9815 0.08132 0.07638 0.0357 0.0084 0.9179 17.500 0.9804 0.08955 0.08499 0.0316 0.0080 0.9434 19.500 0.9105 0.13624 0.13299 0.0049 0.0077 1.0000