XFOIL Version 6.94 Calculated polar for: nabad 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 -0.3910 0.05034 0.04315 0.0027 0.9999 0.2400 -2.750 -0.3504 0.04579 0.03790 0.0002 0.9999 0.1798 -2.500 -0.3162 0.04221 0.03368 -0.0001 0.9999 0.1519 -2.250 -0.2849 0.03898 0.03011 0.0004 0.9999 0.1360 -2.000 -0.2396 0.03602 0.02632 -0.0006 0.8678 0.1210 -1.750 -0.2178 0.03530 0.02390 0.0042 0.4592 0.1139 -1.500 -0.1866 0.03426 0.02170 0.0047 0.3806 0.1076 -1.250 -0.1488 0.03283 0.01965 0.0042 0.3615 0.1019 -1.000 -0.1076 0.03170 0.01796 0.0031 0.3500 0.0973 -0.750 -0.0705 0.03074 0.01673 0.0026 0.3433 0.0941 -0.500 -0.0370 0.03006 0.01584 0.0027 0.3386 0.0922 -0.250 -0.0050 0.02961 0.01526 0.0029 0.3353 0.0916 0.000 0.0272 0.02933 0.01490 0.0030 0.3331 0.0919 0.250 0.0596 0.02916 0.01475 0.0031 0.3318 0.0953 0.500 0.0925 0.02900 0.01477 0.0029 0.3311 0.1202 0.750 0.1836 0.02737 0.01581 -0.0095 0.3310 1.0001 1.000 0.2132 0.02779 0.01613 -0.0089 0.3318 1.0001 1.250 0.2425 0.02825 0.01650 -0.0083 0.3314 1.0001 1.500 0.2689 0.02885 0.01679 -0.0077 0.3102 1.0001 1.750 0.2930 0.02936 0.01683 -0.0070 0.2757 1.0001 2.000 0.3185 0.02975 0.01681 -0.0062 0.2506 1.0001 2.250 0.3440 0.03013 0.01692 -0.0058 0.2346 1.0001 2.500 0.3689 0.03046 0.01693 -0.0053 0.2196 1.0001 2.750 0.3950 0.03034 0.01681 -0.0050 0.2065 1.0001 3.000 0.4205 0.03031 0.01666 -0.0046 0.1939 1.0001 3.250 0.4458 0.03014 0.01636 -0.0043 0.1798 1.0001 3.500 0.4719 0.02994 0.01608 -0.0039 0.1679 1.0001 3.750 0.4988 0.02976 0.01587 -0.0034 0.1573 1.0001 4.000 0.5263 0.02966 0.01579 -0.0029 0.1476 1.0001 4.250 0.5540 0.02970 0.01591 -0.0022 0.1396 1.0001 4.500 0.5819 0.02977 0.01606 -0.0015 0.1218 1.0001 4.750 0.6088 0.03015 0.01605 -0.0011 0.0823 1.0001 5.000 0.6341 0.03202 0.01756 -0.0008 0.0744 1.0001