XFOIL Version 6.94 Calculated polar for: S4083 (8%) 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 0.4301 0.00669 0.00172 -0.0889 0.5816 1.0000 0.500 0.4810 0.00700 0.00172 -0.0875 0.5447 1.0000 1.000 0.5323 0.00731 0.00178 -0.0863 0.5141 1.0000 1.500 0.5843 0.00760 0.00188 -0.0853 0.4879 1.0000 2.000 0.6363 0.00791 0.00201 -0.0843 0.4644 1.0000 2.500 0.6887 0.00820 0.00220 -0.0833 0.4434 1.0000 3.000 0.7410 0.00850 0.00241 -0.0824 0.4241 1.0000 3.500 0.7931 0.00882 0.00268 -0.0815 0.4059 1.0000 4.000 0.8449 0.00917 0.00296 -0.0805 0.3872 1.0000 4.500 0.8967 0.00952 0.00328 -0.0796 0.3692 1.0000 5.000 0.9486 0.00985 0.00366 -0.0786 0.3506 1.0000 5.500 0.9999 0.01020 0.00404 -0.0776 0.3293 1.0000 6.000 1.0502 0.01061 0.00446 -0.0765 0.3014 1.0000 6.500 1.0998 0.01112 0.00493 -0.0753 0.2680 1.0000 7.000 1.1469 0.01186 0.00554 -0.0738 0.2192 1.0000 7.500 1.1905 0.01301 0.00645 -0.0718 0.1579 1.0000 8.000 1.2268 0.01492 0.00787 -0.0690 0.0757 1.0000 8.500 1.2547 0.01768 0.01024 -0.0646 0.0120 1.0000 10.000 1.3433 0.02383 0.01720 -0.0523 0.0078 1.0000 10.500 1.3488 0.02669 0.02033 -0.0450 0.0076 1.0000 11.000 1.3455 0.03002 0.02393 -0.0374 0.0075 1.0000 11.500 1.3410 0.03401 0.02819 -0.0318 0.0075 1.0000 12.000 1.3370 0.03880 0.03327 -0.0283 0.0076 1.0000 12.500 1.3305 0.04485 0.03965 -0.0267 0.0076 1.0000 13.000 1.3249 0.05153 0.04666 -0.0267 0.0077 1.0000 13.500 1.3130 0.05968 0.05513 -0.0281 0.0077 1.0000 14.000 1.2978 0.06893 0.06474 -0.0310 0.0078 1.0000 14.500 1.2752 0.07975 0.07584 -0.0354 0.0076 1.0000 15.000 1.2527 0.09157 0.08798 -0.0412 0.0076 1.0000 15.500 1.2263 0.10507 0.10180 -0.0484 0.0076 1.0000 16.000 1.1998 0.11982 0.11684 -0.0570 0.0076 1.0000 16.500 1.1561 0.14088 0.13835 -0.0698 0.0083 1.0000