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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.4245 0.02954 0.01672 0.0288 1.0000 0.1637 -2.750 -0.4148 0.02987 0.01691 0.0322 1.0000 0.1788 -2.500 -0.4089 0.03015 0.01718 0.0360 1.0000 0.2002 -2.250 -0.4001 0.03032 0.01735 0.0392 1.0000 0.2208 -2.000 -0.3921 0.03049 0.01733 0.0426 1.0000 0.2441 -1.750 -0.3880 0.03053 0.01772 0.0462 1.0000 0.2873 -1.500 -0.3824 0.03046 0.01796 0.0497 1.0000 0.3390 -1.250 -0.3512 0.03029 0.01737 0.0487 1.0000 0.3378 -1.000 -0.3157 0.03007 0.01683 0.0466 1.0000 0.3367 -0.750 -0.2796 0.02989 0.01643 0.0444 1.0000 0.3358 -0.500 -0.2441 0.02975 0.01616 0.0423 1.0000 0.3353 -0.250 -0.2099 0.02965 0.01601 0.0404 1.0000 0.3354 0.000 -0.1765 0.02957 0.01598 0.0388 1.0000 0.3360 0.250 0.1694 0.02577 0.01433 -0.0203 0.8575 0.3527 0.500 0.2234 0.02382 0.01298 -0.0212 0.7596 0.3616 0.750 0.2243 0.02521 0.01225 -0.0136 0.4126 0.3671 1.000 0.2410 0.02610 0.01289 -0.0122 0.3378 0.3764 1.250 0.2643 0.02633 0.01326 -0.0115 0.2855 0.3937 1.500 0.2909 0.02602 0.01354 -0.0113 0.2344 0.4599 1.750 0.4195 0.02760 0.01567 -0.0291 0.0816 1.0000 2.000 0.4450 0.02829 0.01645 -0.0278 0.0792 1.0000 2.250 0.4703 0.02902 0.01730 -0.0264 0.0785 1.0000 2.500 0.4955 0.02984 0.01823 -0.0250 0.0799 1.0000 2.750 0.5204 0.03085 0.01926 -0.0235 0.0825 1.0000 3.000 0.5470 0.03150 0.02001 -0.0220 0.0862 1.0000 3.250 0.5759 0.03209 0.02070 -0.0206 0.0921 1.0000 3.500 0.6120 0.03335 0.02188 -0.0200 0.0986 1.0000 3.750 0.6425 0.03341 0.02231 -0.0182 0.1073 1.0000 4.000 0.6789 0.03417 0.02343 -0.0171 0.1170 1.0000 4.250 0.7164 0.03569 0.02530 -0.0162 0.1282 1.0000 4.500 0.7515 0.03802 0.02849 -0.0150 0.1412 1.0000 4.750 0.7800 0.04302 0.03455 -0.0139 0.1553 1.0000 5.000 0.7993 0.04893 0.04113 -0.0126 0.1691 1.0000