XFOIL Version 6.94 Calculated polar for: EPPLER 547 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 0.2387 0.01496 0.00875 -0.0362 0.6370 0.7576 0.500 0.2934 0.01491 0.00856 -0.0364 0.6268 0.7606 1.000 0.3451 0.01469 0.00834 -0.0365 0.6159 0.7634 1.500 0.4000 0.01454 0.00807 -0.0372 0.6060 0.7665 2.000 0.4537 0.01444 0.00796 -0.0378 0.5961 0.7701 2.500 0.5093 0.01433 0.00778 -0.0387 0.5862 0.7727 3.000 0.5632 0.01428 0.00773 -0.0391 0.5768 0.7748 3.500 0.6155 0.01423 0.00774 -0.0390 0.5667 0.7771 4.000 0.6704 0.01438 0.00787 -0.0394 0.5569 0.7801 4.500 0.7211 0.01434 0.00795 -0.0390 0.5457 0.7830 5.000 0.7764 0.01447 0.00804 -0.0396 0.5346 0.7859 5.500 0.8260 0.01445 0.00816 -0.0391 0.5227 0.7890 6.000 0.8797 0.01460 0.00831 -0.0394 0.5108 0.7920 6.500 0.9266 0.01455 0.00841 -0.0383 0.4971 0.7949 7.000 0.9721 0.01464 0.00865 -0.0368 0.4823 0.7986 7.500 1.0171 0.01477 0.00884 -0.0352 0.4660 0.8026 8.000 1.0570 0.01489 0.00906 -0.0328 0.4472 0.8068 8.500 1.0908 0.01507 0.00936 -0.0293 0.4251 0.8112 9.000 1.1121 0.01527 0.00967 -0.0234 0.3999 0.8158 9.500 1.1251 0.01581 0.01030 -0.0163 0.3676 0.8215 10.000 1.1301 0.01689 0.01132 -0.0087 0.3282 0.8287 10.500 1.1254 0.01856 0.01290 -0.0007 0.2854 0.8368 11.000 1.1165 0.02088 0.01513 0.0067 0.2455 0.8461 11.500 1.1070 0.02376 0.01793 0.0128 0.2098 0.8568 12.000 1.1018 0.02687 0.02106 0.0176 0.1779 0.8706 13.000 1.1022 0.03500 0.02937 0.0207 0.1209 0.9569 14.500 1.0943 0.05036 0.04462 0.0222 0.0665 1.0000 15.000 1.0912 0.05611 0.05040 0.0219 0.0551 1.0000 15.500 1.0868 0.06231 0.05663 0.0212 0.0466 1.0000 16.000 1.0856 0.06846 0.06287 0.0200 0.0394 1.0000 16.500 1.0838 0.07498 0.06950 0.0184 0.0338 1.0000 17.000 1.0807 0.08193 0.07656 0.0164 0.0292 1.0000 17.500 1.0775 0.08918 0.08394 0.0139 0.0255 1.0000