XFOIL Version 6.94 Calculated polar for: manu4/28 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1965 0.09203 0.07895 0.0255 0.9999 0.8684 -2.750 -0.2115 0.08946 0.07665 0.0292 0.9999 0.8577 -2.500 -0.2271 0.08678 0.07420 0.0329 0.9999 0.8481 -2.250 -0.2575 0.08401 0.07170 0.0393 0.9999 0.8378 -2.000 -0.2712 0.08118 0.06913 0.0426 0.9999 0.8297 -1.750 -0.2892 0.07833 0.06645 0.0465 0.9999 0.8218 -1.500 -0.2901 0.07568 0.06405 0.0473 0.9999 0.8160 -1.250 -0.2833 0.07324 0.06177 0.0468 0.9999 0.8128 -1.000 -0.2617 0.07118 0.05998 0.0437 0.9999 0.8139 -0.750 -0.2313 0.06958 0.05876 0.0391 0.9999 0.8195 -0.500 -0.1959 0.06876 0.05851 0.0330 0.9999 0.8298 -0.250 -0.1756 0.06945 0.05979 0.0280 0.9999 0.8387 0.000 -0.1843 0.07228 0.06292 0.0260 0.9999 0.8401 0.250 -0.1894 0.07538 0.06608 0.0233 0.9999 0.8436 0.500 -0.1772 0.07807 0.06886 0.0179 0.9999 0.8568 0.750 -0.1510 0.08066 0.07158 0.0098 0.9999 0.8859 1.000 -0.1324 0.08178 0.07284 0.0031 0.9999 0.9604 1.250 -0.1109 0.08487 0.07529 -0.0025 0.9999 1.0001 1.500 -0.0957 0.08674 0.07672 -0.0065 0.9999 1.0001 1.750 -0.0758 0.08902 0.07851 -0.0113 0.9999 1.0001 2.000 -0.0514 0.09163 0.08057 -0.0168 0.9999 1.0001 2.250 -0.0241 0.09453 0.08286 -0.0227 0.9999 1.0001 2.500 0.0058 0.09772 0.08543 -0.0288 0.9999 1.0001 2.750 0.0377 0.10116 0.08820 -0.0352 0.9999 1.0001 3.000 0.0721 0.10489 0.09117 -0.0419 0.9999 1.0001 3.250 0.1066 0.10878 0.09424 -0.0484 0.9999 1.0001 3.500 0.1390 0.11266 0.09726 -0.0540 0.9999 1.0001 3.750 0.1667 0.11631 0.10008 -0.0583 0.9999 1.0001 4.000 0.1904 0.11972 0.10275 -0.0616 0.9999 1.0001 4.250 0.2111 0.12294 0.10529 -0.0641 0.9999 1.0001 4.500 0.2298 0.12601 0.10774 -0.0660 0.9999 1.0001 4.750 0.2469 0.12899 0.11015 -0.0676 0.9999 1.0001 5.000 0.2629 0.13190 0.11253 -0.0690 0.9999 1.0001