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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3732 0.05001 0.04038 -0.0014 0.9999 0.1681 -2.750 -0.3433 0.04655 0.03648 -0.0015 0.9999 0.1534 -2.500 -0.3123 0.04345 0.03287 -0.0014 0.9999 0.1417 -2.250 -0.2816 0.04050 0.02975 -0.0012 0.9999 0.1340 -2.000 -0.2492 0.03801 0.02693 -0.0008 0.9999 0.1274 -1.750 -0.2123 0.03578 0.02420 -0.0008 0.9999 0.1211 -1.500 -0.1722 0.03372 0.02203 -0.0016 0.9999 0.1173 -1.250 -0.1308 0.03191 0.02010 -0.0023 0.9999 0.1141 -1.000 -0.0878 0.03039 0.01860 -0.0034 0.9725 0.1127 -0.750 -0.0367 0.02970 0.01718 -0.0043 0.6271 0.1134 -0.500 -0.0156 0.03035 0.01632 -0.0012 0.4619 0.1156 -0.250 0.0153 0.03021 0.01580 -0.0009 0.4392 0.1213 0.000 0.0469 0.02999 0.01551 -0.0009 0.4269 0.1345 0.250 0.1297 0.02771 0.01582 -0.0114 0.4143 1.0001 0.500 0.1589 0.02821 0.01594 -0.0106 0.4098 1.0001 0.750 0.1880 0.02874 0.01619 -0.0100 0.4068 1.0001 1.000 0.2172 0.02929 0.01653 -0.0094 0.4049 1.0001 1.250 0.2466 0.02984 0.01692 -0.0090 0.4040 1.0001 1.500 0.2763 0.03039 0.01735 -0.0086 0.4039 1.0001 1.750 0.3062 0.03093 0.01779 -0.0084 0.4047 1.0001 2.000 0.3340 0.03158 0.01823 -0.0078 0.3914 1.0001 2.250 0.3554 0.03245 0.01858 -0.0066 0.3515 1.0001 2.500 0.3778 0.03317 0.01891 -0.0057 0.3194 1.0001 2.750 0.4010 0.03370 0.01913 -0.0051 0.2936 1.0001 3.000 0.4255 0.03379 0.01921 -0.0045 0.2737 1.0001 3.250 0.4501 0.03404 0.01935 -0.0040 0.2597 1.0001 3.500 0.4739 0.03416 0.01928 -0.0033 0.2435 1.0001 3.750 0.4989 0.03409 0.01925 -0.0027 0.2289 1.0001 4.000 0.5236 0.03418 0.01921 -0.0021 0.2163 1.0001 4.250 0.5494 0.03423 0.01932 -0.0014 0.2044 1.0001 4.500 0.5755 0.03444 0.01962 -0.0007 0.1917 1.0001 4.750 0.6018 0.03486 0.02015 0.0001 0.1782 1.0001 5.000 0.6281 0.03558 0.02098 0.0010 0.1630 1.0001