XFOIL Version 6.94 Calculated polar for: manu4/38 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.012 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4265 0.06915 0.05677 0.0374 0.9999 0.6395 -2.750 -0.4038 0.06663 0.05408 0.0338 0.9999 0.6266 -2.500 -0.3796 0.06452 0.05171 0.0298 0.9999 0.6179 -2.250 -0.3521 0.06221 0.04917 0.0266 0.9999 0.6144 -2.000 -0.3217 0.06013 0.04688 0.0230 0.9999 0.6130 -1.750 -0.2880 0.05826 0.04475 0.0192 0.9999 0.6125 -1.500 -0.2513 0.05658 0.04281 0.0150 0.9999 0.6122 -1.250 -0.2121 0.05508 0.04103 0.0106 0.9999 0.6145 -1.000 -0.1721 0.05354 0.03926 0.0066 0.9999 0.6239 -0.750 -0.1313 0.05185 0.03750 0.0029 0.9999 0.6444 -0.500 -0.0866 0.05012 0.03579 -0.0013 0.9999 0.6739 -0.250 -0.0340 0.04777 0.03391 -0.0069 0.9999 0.7212 0.000 0.0635 0.04505 0.03259 -0.0226 0.9999 1.0001 0.250 0.1427 0.04645 0.03300 -0.0363 0.9999 1.0001 0.500 0.1838 0.04827 0.03436 -0.0423 0.9999 1.0001 0.750 0.1742 0.05225 0.03858 -0.0434 0.9999 1.0001 1.000 0.2905 0.05749 0.04253 -0.0703 0.8912 1.0001 1.250 0.3503 0.05920 0.04349 -0.0787 0.8372 1.0001 1.500 0.3967 0.06104 0.04472 -0.0841 0.8061 1.0001 1.750 0.4275 0.06321 0.04647 -0.0870 0.7856 1.0001 2.000 0.4609 0.06537 0.04823 -0.0901 0.7700 1.0001 2.250 0.4681 0.06832 0.05092 -0.0900 0.7605 1.0001 2.500 0.4951 0.07084 0.05314 -0.0923 0.7518 1.0001 2.750 0.5008 0.07401 0.05607 -0.0921 0.7475 1.0001 3.000 0.5057 0.07725 0.05910 -0.0919 0.7451 1.0001 3.250 0.5113 0.08049 0.06214 -0.0918 0.7443 1.0001 3.500 0.5168 0.08378 0.06525 -0.0918 0.7449 1.0001 3.750 0.5227 0.08711 0.06839 -0.0920 0.7468 1.0001 4.000 0.5306 0.09044 0.07157 -0.0924 0.7493 1.0001 4.250 0.5295 0.09377 0.07477 -0.0917 0.7536 1.0001 4.500 0.5239 0.09698 0.07785 -0.0905 0.7602 1.0001 4.750 0.5285 0.10024 0.08097 -0.0905 0.7659 1.0001 5.000 0.5332 0.10349 0.08410 -0.0906 0.7725 1.0001