XFOIL Version 6.94 Calculated polar for: Eiffel 428 (Bleriot) 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2273 0.04490 0.03666 -0.0391 1.0000 0.2352 -2.750 -0.1984 0.04260 0.03380 -0.0411 1.0000 0.2172 -2.500 -0.1692 0.04093 0.03135 -0.0427 1.0000 0.2039 -2.250 -0.1486 0.03942 0.02968 -0.0426 1.0000 0.2005 -2.000 -0.1262 0.03830 0.02821 -0.0428 1.0000 0.1989 -1.750 -0.1041 0.03759 0.02711 -0.0429 1.0000 0.2017 -1.500 -0.0829 0.03705 0.02620 -0.0427 1.0000 0.2038 -1.250 -0.0624 0.03645 0.02537 -0.0424 1.0000 0.2049 -1.000 -0.0327 0.03612 0.02475 -0.0437 0.9972 0.2065 -0.750 0.0158 0.03624 0.02452 -0.0481 0.9871 0.2123 -0.500 0.0725 0.03658 0.02454 -0.0538 0.9724 0.2237 -0.250 0.1294 0.03680 0.02478 -0.0595 0.9560 0.2548 0.000 0.1974 0.03406 0.02452 -0.0667 0.9379 1.0000 0.250 0.2450 0.03512 0.02457 -0.0698 0.9207 1.0000 0.500 0.2775 0.03600 0.02507 -0.0713 0.9064 1.0000 0.750 0.3037 0.03694 0.02576 -0.0720 0.8950 1.0000 1.000 0.3403 0.03802 0.02661 -0.0743 0.8839 1.0000 1.250 0.3623 0.03898 0.02743 -0.0744 0.8739 1.0000 1.500 0.3893 0.04002 0.02835 -0.0753 0.8637 1.0000 1.750 0.4295 0.04105 0.02926 -0.0778 0.8519 1.0000 2.000 0.4408 0.04203 0.03018 -0.0763 0.8419 1.0000 2.500 0.4871 0.04426 0.03235 -0.0769 0.8239 1.0000 2.750 0.5197 0.04545 0.03352 -0.0784 0.8157 1.0000 3.000 0.5273 0.04672 0.03479 -0.0768 0.8086 1.0000 3.250 0.5526 0.04803 0.03616 -0.0776 0.8021 1.0000 3.500 0.5675 0.04936 0.03751 -0.0769 0.7941 1.0000 3.750 0.5985 0.05050 0.03873 -0.0779 0.7835 1.0000 4.000 0.6107 0.05174 0.04002 -0.0767 0.7720 1.0000 4.250 0.6383 0.05272 0.04113 -0.0769 0.7571 1.0000 4.500 0.6689 0.05360 0.04213 -0.0773 0.7418 1.0000 4.750 0.7062 0.05446 0.04316 -0.0785 0.7289 1.0000 5.000 0.7119 0.05593 0.04473 -0.0767 0.7161 1.0000