XFOIL Version 6.94 Calculated polar for: manu02 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 0.0155 0.06090 0.05220 -0.0126 0.3828 0.8433 -2.750 0.0193 0.05860 0.04990 -0.0118 0.3808 0.8393 -2.500 -0.0373 0.05667 0.04820 -0.0002 0.3828 0.8211 -2.250 -0.1352 0.05473 0.04659 0.0186 0.3871 0.7986 -2.000 -0.1530 0.05238 0.04426 0.0222 0.3868 0.7837 -1.750 -0.1397 0.05041 0.04218 0.0191 0.3842 0.7601 -1.250 0.1629 0.05363 0.04223 -0.0570 0.3713 0.3052 -1.000 0.2338 0.05197 0.03984 -0.0664 0.3685 0.2888 -0.750 0.2968 0.05069 0.03812 -0.0739 0.3671 0.2892 -0.500 0.3514 0.04955 0.03668 -0.0794 0.3664 0.2893 -0.250 0.4094 0.04878 0.03550 -0.0855 0.3661 0.2916 0.000 0.4749 0.04826 0.03476 -0.0935 0.3661 0.3021 0.250 0.5391 0.04813 0.03435 -0.1011 0.3659 0.3227 0.500 0.6078 0.04787 0.03402 -0.1098 0.3649 0.3724 0.750 0.6833 0.04723 0.03378 -0.1200 0.3635 0.4892 1.000 0.7217 0.04754 0.03459 -0.1224 0.3632 0.5694 1.250 0.7616 0.04775 0.03545 -0.1254 0.3633 0.6704 1.500 0.8585 0.05033 0.03848 -0.1413 0.3657 1.0001 1.750 0.8904 0.05244 0.04041 -0.1427 0.3676 1.0001 2.000 0.9094 0.05294 0.04116 -0.1411 0.3727 1.0001 2.250 0.9275 0.05459 0.04294 -0.1398 0.3783 1.0001 2.500 0.9456 0.05660 0.04498 -0.1387 0.3835 1.0001 2.750 0.9638 0.05886 0.04720 -0.1376 0.3883 1.0001 3.000 0.9855 0.06152 0.04975 -0.1373 0.3924 1.0001 3.250 0.9863 0.06285 0.05145 -0.1334 0.4017 1.0001 3.500 0.9832 0.06566 0.05449 -0.1296 0.4116 1.0001 3.750 1.0019 0.06891 0.05764 -0.1293 0.4188 1.0001 4.000 0.9564 0.07235 0.06166 -0.1206 0.4367 1.0001 4.250 0.6707 0.09588 0.08665 -0.1038 0.5380 1.0001 4.500 0.7669 0.09422 0.08449 -0.1071 0.5297 1.0001 5.000 0.3620 0.12272 0.11382 -0.0883 0.8792 0.5791