XFOIL Version 6.94 Calculated polar for: pp14 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 -0.1986 0.05788 0.04511 0.0211 0.9996 0.9064 -2.750 -0.2143 0.05736 0.04478 0.0242 0.9996 0.8850 -2.500 -0.1978 0.05701 0.04448 0.0190 0.9996 0.8513 -2.250 -0.1345 0.05736 0.04472 0.0018 0.9996 0.8024 -2.000 -0.0437 0.05765 0.04478 -0.0207 0.9996 0.7525 -1.750 0.0511 0.05791 0.04483 -0.0428 0.9996 0.7047 -1.500 0.1426 0.05828 0.04498 -0.0632 0.9996 0.6602 -1.250 0.2222 0.05856 0.04509 -0.0796 0.9996 0.6308 -1.000 0.2831 0.05842 0.04488 -0.0903 0.9996 0.6152 -0.750 0.3433 0.05860 0.04497 -0.1006 0.9996 0.6023 -0.500 0.3982 0.05883 0.04513 -0.1092 0.9996 0.5910 -0.250 0.4493 0.05922 0.04546 -0.1167 0.9996 0.5865 0.000 0.4923 0.05941 0.04569 -0.1218 0.9996 0.5900 0.250 0.5333 0.05974 0.04607 -0.1263 0.9996 0.5972 0.500 0.5744 0.06032 0.04669 -0.1307 0.9996 0.6064 0.750 0.6106 0.06072 0.04729 -0.1338 0.9996 0.6225 1.000 0.6461 0.06119 0.04806 -0.1365 0.9996 0.6504 1.250 0.6792 0.06148 0.04883 -0.1386 0.9996 0.6926 1.500 0.7059 0.06143 0.04963 -0.1393 0.9996 0.7752 1.750 0.7233 0.06260 0.05102 -0.1395 0.9996 1.0004 2.000 0.7628 0.06460 0.05287 -0.1432 0.9996 1.0004 2.250 0.7975 0.06645 0.05453 -0.1452 0.9996 1.0004 2.500 0.8277 0.06814 0.05596 -0.1457 0.9996 1.0004 2.750 0.8538 0.06971 0.05749 -0.1453 0.9996 1.0004 3.000 0.8773 0.07149 0.05970 -0.1448 0.9996 1.0004 3.500 1.2431 0.06308 0.04943 -0.1621 0.4289 1.0004 3.750 1.2691 0.06697 0.05337 -0.1623 0.4150 1.0004 4.000 1.2988 0.07089 0.05716 -0.1625 0.4029 1.0004 4.250 1.3104 0.07563 0.06234 -0.1623 0.3974 1.0004 4.500 1.3186 0.08076 0.06782 -0.1621 0.3937 1.0004 4.750 1.3199 0.08647 0.07386 -0.1619 0.3924 1.0004 5.000 1.3115 0.09294 0.08065 -0.1619 0.3937 1.0004