XFOIL Version 6.94 Calculated polar for: manu01 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1255 0.05038 0.04329 0.0043 0.4945 0.8070 -2.000 -0.0344 0.05002 0.03980 -0.0365 0.4355 0.3164 -1.750 -0.0010 0.04820 0.03739 -0.0384 0.4252 0.2956 -1.500 0.0344 0.04643 0.03511 -0.0402 0.4172 0.2863 -1.250 0.0767 0.04491 0.03303 -0.0433 0.4099 0.2804 -1.000 0.1215 0.04363 0.03139 -0.0466 0.4030 0.2851 -0.750 0.1779 0.04258 0.02968 -0.0523 0.3954 0.2884 -0.500 0.2358 0.04158 0.02814 -0.0581 0.3887 0.2934 -0.250 0.3073 0.04077 0.02679 -0.0670 0.3806 0.3071 0.000 0.3994 0.04008 0.02562 -0.0802 0.3742 0.3492 0.250 0.5136 0.03837 0.02432 -0.0981 0.3709 0.5117 0.500 0.5701 0.03745 0.02435 -0.1041 0.3700 0.6587 0.750 0.6875 0.03881 0.02590 -0.1238 0.3691 1.0001 1.000 0.7259 0.03995 0.02680 -0.1261 0.3695 1.0001 1.250 0.7604 0.04116 0.02782 -0.1275 0.3701 1.0001 1.500 0.7922 0.04245 0.02895 -0.1284 0.3710 1.0001 1.750 0.8207 0.04370 0.03014 -0.1286 0.3723 1.0001 2.000 0.8448 0.04461 0.03118 -0.1279 0.3747 1.0001 2.250 0.8676 0.04579 0.03250 -0.1271 0.3780 1.0001 2.500 0.8898 0.04722 0.03403 -0.1264 0.3810 1.0001 2.750 0.9114 0.04878 0.03567 -0.1256 0.3837 1.0001 3.000 0.9309 0.05047 0.03744 -0.1245 0.3861 1.0001 3.250 0.9489 0.05226 0.03933 -0.1232 0.3882 1.0001 3.500 0.9660 0.05419 0.04134 -0.1218 0.3902 1.0001 3.750 0.9835 0.05631 0.04350 -0.1206 0.3925 1.0001 4.000 1.0027 0.05873 0.04591 -0.1197 0.3947 1.0001 4.250 0.9984 0.06033 0.04805 -0.1157 0.4034 1.0001 4.500 1.0026 0.06317 0.05109 -0.1133 0.4114 1.0001 4.750 1.0218 0.06617 0.05410 -0.1130 0.4172 1.0001 5.000 0.9862 0.07003 0.05848 -0.1068 0.4339 1.0001