XFOIL Version 6.94 Calculated polar for: EPPLER STE 871-514 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.1758 0.02349 0.01867 -0.0391 0.8559 0.8816 1.000 0.2211 0.02305 0.01818 -0.0383 0.8408 0.8846 2.000 0.3512 0.02212 0.01722 -0.0431 0.8234 0.8879 3.500 0.5537 0.01874 0.01401 -0.0492 0.7922 0.8908 4.000 0.6198 0.01707 0.01242 -0.0504 0.7741 0.8914 5.000 0.7628 0.01374 0.00913 -0.0544 0.7049 0.8934 5.500 0.8265 0.01340 0.00810 -0.0552 0.5676 0.8938 6.000 0.8340 0.01528 0.00920 -0.0476 0.4202 0.8952 6.500 0.8482 0.01735 0.01052 -0.0420 0.2781 0.8964 7.000 0.8719 0.01933 0.01184 -0.0383 0.1618 0.8976 7.500 0.9033 0.02093 0.01315 -0.0356 0.1149 0.8995 8.000 0.9383 0.02238 0.01453 -0.0336 0.0972 0.9006 8.500 0.9728 0.02389 0.01601 -0.0315 0.0867 0.9017 9.000 1.0102 0.02526 0.01746 -0.0299 0.0795 0.9028 10.000 1.0835 0.02846 0.02079 -0.0265 0.0689 0.9053 10.500 1.1226 0.03001 0.02242 -0.0253 0.0646 0.9065 11.000 1.1627 0.03173 0.02431 -0.0240 0.0608 0.9081 11.500 1.2068 0.03349 0.02606 -0.0235 0.0569 0.9100 12.000 1.2289 0.03517 0.02808 -0.0202 0.0531 0.9126 12.500 1.2512 0.03709 0.03019 -0.0174 0.0490 0.9156 13.000 1.2774 0.03903 0.03220 -0.0152 0.0460 0.9195 13.500 1.2954 0.04166 0.03523 -0.0123 0.0434 0.9230 14.500 1.3206 0.04724 0.04136 -0.0067 0.0386 0.9327 15.000 1.3313 0.05004 0.04423 -0.0045 0.0362 0.9414 15.500 1.3294 0.05422 0.04894 -0.0022 0.0337 0.9647 16.500 1.3441 0.06381 0.05903 -0.0036 0.0280 0.9980 17.000 1.3576 0.06896 0.06457 -0.0057 0.0219 0.9980 17.500 1.3501 0.07684 0.07267 -0.0078 0.0150 0.9980 18.000 1.3225 0.08801 0.08414 -0.0114 0.0130 0.9980 18.500 1.2855 0.10165 0.09813 -0.0174 0.0121 0.9980 19.000 1.2428 0.11735 0.11419 -0.0255 0.0121 0.9980 19.500 1.1945 0.13520 0.13235 -0.0356 0.0127 0.9980 20.000 1.1443 0.15451 0.15203 -0.0470 0.0137 0.9980 20.500 1.0902 0.17621 0.17397 -0.0604 0.0150 0.9980