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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1404 0.05677 0.05084 0.0124 0.5892 0.8241 -2.750 -0.2123 0.05941 0.05421 0.0303 0.6358 0.8383 -2.250 -0.2851 0.05309 0.04765 0.0342 0.7517 0.7251 -2.000 -0.0563 0.05422 0.04468 -0.0284 0.4685 0.3882 -1.750 -0.0119 0.05293 0.04255 -0.0341 0.4557 0.3522 -1.500 0.0321 0.05144 0.04034 -0.0385 0.4465 0.3340 -1.250 0.0735 0.04944 0.03802 -0.0415 0.4382 0.3260 -1.000 0.1275 0.04808 0.03589 -0.0472 0.4298 0.3162 -0.750 0.1829 0.04693 0.03410 -0.0528 0.4219 0.3141 -0.500 0.2389 0.04578 0.03259 -0.0585 0.4134 0.3226 -0.250 0.3088 0.04509 0.03125 -0.0671 0.4057 0.3401 0.000 0.3948 0.04437 0.03000 -0.0790 0.4009 0.3669 0.250 0.4916 0.04297 0.02861 -0.0931 0.3982 0.4540 0.500 0.5701 0.04107 0.02772 -0.1033 0.3966 0.6466 0.750 0.6993 0.04207 0.02915 -0.1257 0.3956 1.0001 1.000 0.7420 0.04333 0.03013 -0.1290 0.3962 1.0001 1.250 0.7790 0.04463 0.03123 -0.1311 0.3972 1.0001 1.500 0.8122 0.04601 0.03245 -0.1324 0.3985 1.0001 1.750 0.8429 0.04749 0.03381 -0.1333 0.4001 1.0001 2.000 0.8721 0.04913 0.03532 -0.1339 0.4017 1.0001 2.250 0.8945 0.05036 0.03666 -0.1331 0.4042 1.0001 2.500 0.9121 0.05165 0.03812 -0.1316 0.4072 1.0001 2.750 0.9292 0.05328 0.03987 -0.1303 0.4103 1.0001 3.000 0.9451 0.05511 0.04181 -0.1288 0.4133 1.0001 3.250 0.9590 0.05710 0.04390 -0.1271 0.4161 1.0001 3.500 0.9717 0.05929 0.04616 -0.1253 0.4186 1.0001 3.750 0.9843 0.06165 0.04860 -0.1236 0.4216 1.0001 4.000 0.9991 0.06425 0.05121 -0.1223 0.4245 1.0001 4.250 0.9871 0.06674 0.05407 -0.1177 0.4325 1.0001 4.500 0.9696 0.07053 0.05810 -0.1131 0.4422 1.0001 4.750 0.9881 0.07389 0.06141 -0.1133 0.4491 1.0001 5.000 0.9038 0.08103 0.06898 -0.1034 0.4700 1.0001