XFOIL Version 6.94 Calculated polar for: EPPLER 403 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.4997 0.01250 0.00777 -0.1282 0.8580 0.6507 1.000 0.5873 0.01185 0.00714 -0.1347 0.8562 0.6560 2.500 0.7643 0.01037 0.00586 -0.1362 0.8244 0.6690 3.000 0.8223 0.00996 0.00553 -0.1364 0.8077 0.6739 3.500 0.8781 0.00964 0.00520 -0.1361 0.7799 0.6785 4.000 0.9324 0.00953 0.00496 -0.1356 0.7256 0.6828 4.500 0.9595 0.01012 0.00512 -0.1295 0.6308 0.6876 5.000 0.9649 0.01140 0.00589 -0.1196 0.5268 0.6921 5.500 0.9820 0.01265 0.00674 -0.1126 0.4401 0.6970 6.000 1.0064 0.01390 0.00762 -0.1075 0.3603 0.7016 6.500 1.0318 0.01529 0.00859 -0.1028 0.2708 0.7063 7.000 1.0629 0.01657 0.00960 -0.0992 0.2042 0.7113 7.500 1.0941 0.01800 0.01074 -0.0958 0.1403 0.7163 8.000 1.1240 0.01963 0.01205 -0.0925 0.0840 0.7213 8.500 1.1513 0.02158 0.01368 -0.0889 0.0332 0.7256 9.000 1.1768 0.02356 0.01560 -0.0849 0.0122 0.7310 10.000 1.2363 0.02705 0.01942 -0.0786 0.0093 0.7427 10.500 1.2622 0.02917 0.02171 -0.0753 0.0088 0.7477 11.000 1.2825 0.03164 0.02440 -0.0716 0.0084 0.7537 11.500 1.2985 0.03457 0.02753 -0.0678 0.0083 0.7600 12.000 1.3098 0.03811 0.03125 -0.0641 0.0081 0.7666 12.500 1.3184 0.04210 0.03548 -0.0607 0.0080 0.7731 13.000 1.3254 0.04662 0.04023 -0.0577 0.0080 0.7799 13.500 1.3332 0.05152 0.04536 -0.0555 0.0079 0.7871 14.000 1.3413 0.05670 0.05083 -0.0539 0.0079 0.7946 14.500 1.3474 0.06242 0.05687 -0.0529 0.0080 0.8036 15.000 1.3494 0.06898 0.06377 -0.0526 0.0080 0.8134 15.500 1.3447 0.07673 0.07196 -0.0532 0.0081 0.8247 16.000 1.3313 0.08639 0.08211 -0.0552 0.0083 0.8368 16.500 1.3036 0.09920 0.09551 -0.0596 0.0085 0.8501 17.000 1.2651 0.11487 0.11178 -0.0672 0.0088 0.8673