XFOIL Version 6.94 Calculated polar for: EPPLER 548 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2187 0.01216 0.00598 -0.0293 0.5248 0.7326 0.500 0.2750 0.01196 0.00572 -0.0298 0.5191 0.7347 1.000 0.3316 0.01201 0.00570 -0.0304 0.5133 0.7370 1.500 0.3884 0.01207 0.00575 -0.0310 0.5078 0.7392 2.000 0.4450 0.01207 0.00576 -0.0316 0.5028 0.7413 2.500 0.5019 0.01210 0.00579 -0.0323 0.4974 0.7436 3.000 0.5589 0.01221 0.00585 -0.0330 0.4918 0.7460 3.500 0.6158 0.01244 0.00609 -0.0338 0.4863 0.7489 4.000 0.6718 0.01247 0.00619 -0.0343 0.4807 0.7512 4.500 0.7278 0.01249 0.00623 -0.0349 0.4739 0.7532 5.000 0.7836 0.01267 0.00640 -0.0354 0.4667 0.7556 5.500 0.8370 0.01270 0.00660 -0.0354 0.4604 0.7581 6.000 0.8906 0.01272 0.00671 -0.0354 0.4523 0.7607 6.500 0.9451 0.01298 0.00697 -0.0356 0.4431 0.7636 7.000 0.9960 0.01295 0.00714 -0.0351 0.4347 0.7673 7.500 1.0474 0.01303 0.00725 -0.0347 0.4236 0.7704 8.000 1.0968 0.01309 0.00749 -0.0340 0.4116 0.7734 8.500 1.1430 0.01317 0.00767 -0.0327 0.3967 0.7769 9.000 1.1855 0.01334 0.00795 -0.0307 0.3762 0.7807 9.500 1.2234 0.01368 0.00838 -0.0280 0.3471 0.7850 10.000 1.2423 0.01444 0.00905 -0.0221 0.3062 0.7899 10.500 1.2433 0.01574 0.01021 -0.0136 0.2637 0.7957 11.500 1.2273 0.01994 0.01432 0.0023 0.1968 0.8102 12.000 1.2113 0.02330 0.01768 0.0085 0.1709 0.8188 12.500 1.1983 0.02730 0.02172 0.0128 0.1482 0.8291 13.000 1.1842 0.03186 0.02635 0.0161 0.1278 0.8415 13.500 1.1690 0.03689 0.03146 0.0186 0.1101 0.8588 14.000 1.1550 0.04195 0.03670 0.0206 0.0942 0.8965