XFOIL Version 6.94 Calculated polar for: manu4/33 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1498 0.06009 0.05187 -0.0220 0.3471 0.4329 -2.000 -0.0900 0.05016 0.04163 -0.0114 0.3292 0.5856 -1.750 -0.0613 0.04762 0.03902 -0.0122 0.3264 0.6104 -1.250 0.1861 0.04219 0.03008 -0.0642 0.3184 0.2143 -1.000 0.2455 0.04081 0.02819 -0.0704 0.3164 0.2187 -0.750 0.3036 0.03968 0.02649 -0.0764 0.3147 0.2216 -0.500 0.3558 0.03907 0.02578 -0.0814 0.3135 0.2284 -0.250 0.4128 0.03867 0.02504 -0.0873 0.3129 0.2403 0.000 0.4707 0.03880 0.02511 -0.0937 0.3126 0.2629 0.250 0.5399 0.03874 0.02500 -0.1025 0.3130 0.3189 0.500 0.6058 0.03860 0.02523 -0.1108 0.3140 0.4182 0.750 0.6457 0.03915 0.02603 -0.1135 0.3146 0.4705 1.000 0.6828 0.03973 0.02694 -0.1156 0.3148 0.5165 1.250 0.7160 0.04028 0.02794 -0.1169 0.3150 0.5732 1.500 0.8190 0.04178 0.03041 -0.1342 0.3150 1.0001 1.750 0.8463 0.04314 0.03175 -0.1342 0.3172 1.0001 2.000 0.8717 0.04464 0.03324 -0.1338 0.3204 1.0001 2.250 0.8959 0.04625 0.03484 -0.1333 0.3237 1.0001 2.500 0.9188 0.04803 0.03657 -0.1326 0.3267 1.0001 2.750 0.9422 0.05012 0.03856 -0.1321 0.3296 1.0001 3.000 0.9593 0.05100 0.03978 -0.1300 0.3376 1.0001 3.250 0.9736 0.05311 0.04212 -0.1281 0.3463 1.0001 3.500 0.9936 0.05566 0.04465 -0.1273 0.3519 1.0001 3.750 1.0005 0.05745 0.04694 -0.1245 0.3691 1.0001 4.000 1.0187 0.05982 0.04943 -0.1239 0.3830 1.0001 4.250 1.0364 0.06359 0.05335 -0.1240 0.4034 1.0001 4.750 0.4774 0.11006 0.10252 -0.1025 0.7888 1.0001 5.000 0.5050 0.11265 0.10488 -0.1037 0.7675 1.0001