XFOIL Version 6.94 Calculated polar for: busher 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 0.2943 0.06324 0.05236 -0.0564 0.8326 0.1932 -2.750 0.3323 0.06297 0.05204 -0.0582 0.8208 0.1995 -2.500 0.3837 0.06204 0.05096 -0.0616 0.8079 0.2103 -2.250 0.4142 0.06173 0.05061 -0.0619 0.7944 0.2219 -2.000 0.4738 0.06088 0.04967 -0.0656 0.7820 0.3174 -1.750 0.4890 0.06183 0.05082 -0.0635 0.7659 0.3357 -1.500 0.5339 0.06161 0.05113 -0.0647 0.7494 0.3845 -1.250 0.5800 0.06012 0.05051 -0.0638 0.7312 0.4625 -1.000 0.6034 0.05960 0.05023 -0.0609 0.7097 0.4988 -0.750 0.7316 0.05208 0.04264 -0.0719 0.6983 0.5297 -0.500 0.7603 0.05089 0.04138 -0.0708 0.6737 0.5284 -0.250 0.8027 0.04895 0.03932 -0.0717 0.6490 0.5272 0.000 0.8480 0.04695 0.03716 -0.0730 0.6232 0.5263 0.250 0.8870 0.04551 0.03552 -0.0736 0.5959 0.5257 0.500 0.9194 0.04467 0.03449 -0.0735 0.5683 0.5253 0.750 0.9482 0.04423 0.03385 -0.0731 0.5421 0.5252 1.000 0.9754 0.04404 0.03346 -0.0727 0.5182 0.5252 1.250 1.0023 0.04395 0.03319 -0.0723 0.4968 0.5255 1.500 1.0304 0.04388 0.03287 -0.0721 0.4777 0.5260 1.750 1.0531 0.04422 0.03307 -0.0713 0.4596 0.5266 2.000 1.0744 0.04480 0.03354 -0.0704 0.4433 0.5274 2.250 1.0998 0.04522 0.03379 -0.0701 0.4288 0.5285 2.500 1.1296 0.04550 0.03386 -0.0704 0.4153 0.5300 2.750 1.1620 0.04591 0.03400 -0.0713 0.4025 0.5317 3.000 1.1937 0.04679 0.03471 -0.0723 0.3915 0.5336 3.250 1.2360 0.04757 0.03525 -0.0752 0.3818 0.5368 3.500 1.2781 0.04872 0.03626 -0.0781 0.3724 0.5415 3.750 1.3091 0.05054 0.03812 -0.0795 0.3648 0.5467 4.000 1.3381 0.05247 0.04011 -0.0805 0.3576 0.5528 4.250 1.3719 0.05476 0.04238 -0.0824 0.3518 0.5605 4.500 1.3786 0.05719 0.04524 -0.0799 0.3482 0.5658 4.750 1.3860 0.05978 0.04825 -0.0777 0.3448 0.5722 5.000 1.3950 0.06248 0.05135 -0.0759 0.3414 0.5810