XFOIL Version 6.94 Calculated polar for: kut1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4201 0.03290 0.01928 0.0273 1.0000 0.2235 -2.750 -0.4126 0.03311 0.01933 0.0309 1.0000 0.2444 -2.500 -0.4084 0.03319 0.01960 0.0348 1.0000 0.2782 -2.000 -0.4013 0.03293 0.01989 0.0431 1.0000 0.3749 -1.750 -0.3750 0.03267 0.01925 0.0430 1.0000 0.3861 -1.500 -0.3422 0.03244 0.01854 0.0416 1.0000 0.3849 -1.250 -0.3074 0.03216 0.01791 0.0396 1.0000 0.3840 -1.000 -0.2728 0.03193 0.01740 0.0377 1.0000 0.3837 -0.750 -0.2384 0.03174 0.01703 0.0358 1.0000 0.3840 -0.500 -0.2040 0.03158 0.01679 0.0340 1.0000 0.3851 -0.250 -0.1692 0.03139 0.01668 0.0320 1.0000 0.3872 0.000 -0.1358 0.03123 0.01670 0.0303 1.0000 0.3904 0.250 -0.1038 0.03112 0.01683 0.0289 1.0000 0.3947 0.500 -0.0730 0.03105 0.01705 0.0278 1.0000 0.4003 0.750 -0.0417 0.03088 0.01742 0.0267 1.0000 0.4077 1.250 0.3682 0.02789 0.01568 -0.0308 0.2921 1.0000 1.500 0.3905 0.02830 0.01584 -0.0294 0.2475 1.0000 1.750 0.4146 0.02855 0.01593 -0.0280 0.2021 1.0000 2.000 0.4344 0.02939 0.01629 -0.0261 0.1087 1.0000 2.250 0.4553 0.03031 0.01724 -0.0241 0.1014 1.0000 2.500 0.4768 0.03129 0.01826 -0.0222 0.0985 1.0000 2.750 0.4992 0.03220 0.01928 -0.0204 0.0974 1.0000 3.000 0.5216 0.03319 0.02035 -0.0185 0.0971 1.0000 3.250 0.5454 0.03433 0.02157 -0.0169 0.0979 1.0000 3.500 0.5716 0.03514 0.02256 -0.0152 0.1002 1.0000 3.750 0.6006 0.03584 0.02350 -0.0137 0.1049 1.0000 4.000 0.6362 0.03698 0.02484 -0.0129 0.1108 1.0000 4.250 0.6742 0.03778 0.02624 -0.0120 0.1179 1.0000 4.500 0.7197 0.03963 0.02855 -0.0121 0.1275 1.0000 4.750 0.7574 0.04237 0.03244 -0.0113 0.1387 1.0000 5.000 0.7837 0.04739 0.03849 -0.0104 0.1495 1.0000