XFOIL Version 6.94 Calculated polar for: mta1 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1492 0.06041 0.04179 -0.0063 1.0000 1.0000 -2.750 -0.1507 0.05841 0.04017 -0.0058 1.0000 1.0000 -2.500 -0.1542 0.05640 0.03855 -0.0051 1.0000 1.0000 -2.250 -0.1593 0.05437 0.03691 -0.0043 1.0000 1.0000 -2.000 -0.1625 0.05239 0.03515 -0.0043 1.0000 1.0000 -1.750 -0.1453 0.05062 0.03292 -0.0094 1.0000 1.0000 -1.500 -0.0933 0.04985 0.03025 -0.0195 1.0000 1.0000 -1.250 -0.0457 0.04987 0.02820 -0.0238 1.0000 1.0000 -1.000 -0.0100 0.04998 0.02672 -0.0243 1.0000 1.0000 -0.750 0.0201 0.05007 0.02561 -0.0235 1.0000 1.0000 -0.500 0.0475 0.05015 0.02478 -0.0225 1.0000 1.0000 -0.250 0.0738 0.05024 0.02418 -0.0214 1.0000 1.0000 0.000 0.0996 0.05034 0.02375 -0.0205 1.0000 1.0000 0.250 0.1251 0.05046 0.02352 -0.0196 1.0000 1.0000 0.500 0.1507 0.05061 0.02348 -0.0188 1.0000 1.0000 0.750 0.1765 0.05078 0.02361 -0.0180 1.0000 1.0000 1.000 0.2027 0.05098 0.02392 -0.0174 1.0000 1.0000 1.250 0.2297 0.05120 0.02448 -0.0170 1.0000 1.0000 1.500 0.2578 0.05147 0.02531 -0.0170 1.0000 1.0000 1.750 0.2871 0.05185 0.02651 -0.0177 1.0000 1.0000 2.000 0.3153 0.05259 0.02826 -0.0196 1.0000 1.0000 2.250 0.3341 0.05421 0.03076 -0.0225 1.0000 1.0000 2.500 0.3372 0.05696 0.03378 -0.0251 1.0000 1.0000 2.750 0.3339 0.06017 0.03679 -0.0270 1.0000 1.0000 3.000 0.3330 0.06331 0.03963 -0.0289 1.0000 1.0000 3.250 0.3350 0.06632 0.04235 -0.0307 1.0000 1.0000 3.500 0.3390 0.06924 0.04500 -0.0325 1.0000 1.0000 3.750 0.3443 0.07210 0.04760 -0.0342 1.0000 1.0000 4.000 0.3505 0.07493 0.05018 -0.0358 1.0000 1.0000 4.250 0.3576 0.07775 0.05277 -0.0374 1.0000 1.0000 4.500 0.3653 0.08058 0.05536 -0.0389 1.0000 1.0000 4.750 0.3734 0.08341 0.05796 -0.0403 1.0000 1.0000 5.000 0.3818 0.08624 0.06057 -0.0418 1.0000 1.0000