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 = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.2443 0.01139 0.00556 -0.0408 0.6093 0.7329 0.500 0.2992 0.01121 0.00531 -0.0412 0.5990 0.7351 1.000 0.3540 0.01114 0.00522 -0.0414 0.5885 0.7375 1.500 0.4098 0.01120 0.00520 -0.0419 0.5789 0.7396 2.000 0.4645 0.01116 0.00517 -0.0422 0.5692 0.7417 2.500 0.5203 0.01126 0.00518 -0.0427 0.5595 0.7440 3.000 0.5749 0.01127 0.00524 -0.0430 0.5504 0.7466 3.500 0.6295 0.01134 0.00528 -0.0433 0.5405 0.7495 4.000 0.6840 0.01145 0.00541 -0.0436 0.5303 0.7517 4.500 0.7371 0.01146 0.00544 -0.0437 0.5194 0.7539 5.000 0.7895 0.01155 0.00558 -0.0435 0.5086 0.7564 5.500 0.8405 0.01161 0.00573 -0.0430 0.4971 0.7589 6.000 0.8910 0.01177 0.00597 -0.0425 0.4850 0.7617 6.500 0.9393 0.01193 0.00617 -0.0416 0.4707 0.7648 7.000 0.9860 0.01207 0.00644 -0.0403 0.4556 0.7686 7.500 1.0304 0.01230 0.00673 -0.0387 0.4385 0.7717 8.000 1.0697 0.01251 0.00700 -0.0361 0.4177 0.7750 8.500 1.0997 0.01279 0.00736 -0.0317 0.3920 0.7789 9.000 1.1172 0.01340 0.00792 -0.0251 0.3583 0.7833 9.500 1.1276 0.01445 0.00886 -0.0180 0.3149 0.7884 10.000 1.1293 0.01598 0.01023 -0.0102 0.2695 0.7934 10.500 1.1267 0.01788 0.01205 -0.0028 0.2292 0.8003 11.000 1.1267 0.02012 0.01422 0.0033 0.1941 0.8074 11.500 1.1262 0.02274 0.01680 0.0085 0.1638 0.8147 12.000 1.1224 0.02582 0.01987 0.0133 0.1369 0.8238 12.500 1.1211 0.02919 0.02323 0.0169 0.1132 0.8343 13.000 1.1185 0.03283 0.02694 0.0202 0.0932 0.8483 13.500 1.1144 0.03686 0.03105 0.0229 0.0764 0.8702 14.000 1.1278 0.04159 0.03603 0.0209 0.0588 0.9472