XFOIL Version 6.94 Calculated polar for: rt 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 -0.2130 0.04423 0.03289 0.0341 0.9639 0.5691 -2.750 -0.1351 0.04413 0.03275 0.0241 0.9362 0.5857 -2.500 -0.0672 0.04371 0.03235 0.0162 0.9116 0.5955 -2.250 0.0069 0.04260 0.03126 0.0082 0.8839 0.6137 -2.000 0.1058 0.04005 0.02891 -0.0032 0.8561 0.6320 -1.750 0.1799 0.03737 0.02644 -0.0095 0.8234 0.6499 -1.500 0.2828 0.03258 0.02197 -0.0190 0.7746 0.6777 -1.250 0.3520 0.02907 0.01855 -0.0223 0.6816 0.7092 -1.000 0.4093 0.02813 0.01692 -0.0259 0.5361 0.7445 -0.750 0.4707 0.02867 0.01714 -0.0334 0.4618 0.7964 -0.250 0.6372 0.03220 0.01976 -0.0602 0.4051 1.0009 0.000 0.6739 0.03283 0.01996 -0.0626 0.3954 1.0009 0.250 0.7074 0.03347 0.02041 -0.0642 0.3867 1.0009 0.500 0.7418 0.03413 0.02079 -0.0659 0.3803 1.0009 0.750 0.7753 0.03494 0.02141 -0.0674 0.3743 1.0009 1.000 0.8045 0.03573 0.02219 -0.0680 0.3703 1.0009 1.250 0.8336 0.03661 0.02304 -0.0686 0.3672 1.0009 1.500 0.8618 0.03755 0.02399 -0.0691 0.3647 1.0009 1.750 0.8892 0.03859 0.02502 -0.0694 0.3622 1.0009 2.000 0.9158 0.03970 0.02614 -0.0695 0.3605 1.0009 2.250 0.9414 0.04092 0.02740 -0.0696 0.3592 1.0009 2.500 0.9659 0.04226 0.02880 -0.0695 0.3581 1.0009 2.750 0.9889 0.04372 0.03035 -0.0691 0.3576 1.0009 3.000 1.0102 0.04529 0.03204 -0.0685 0.3572 1.0009 3.250 1.0293 0.04700 0.03392 -0.0676 0.3572 1.0009 3.500 1.0457 0.04889 0.03604 -0.0663 0.3574 1.0009 3.750 1.0580 0.05102 0.03843 -0.0645 0.3582 1.0009 4.000 1.0673 0.05346 0.04113 -0.0623 0.3587 1.0009 4.250 1.0652 0.05665 0.04468 -0.0588 0.3601 1.0009 4.500 1.0476 0.06108 0.04951 -0.0538 0.3622 1.0009 4.750 1.0041 0.06779 0.05662 -0.0466 0.3656 1.0009 5.000 0.9153 0.07844 0.06753 -0.0369 0.3717 1.0009