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 = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0661 0.06120 0.05214 -0.0209 0.4178 0.8713 -2.750 -0.0016 0.05984 0.05101 -0.0073 0.4207 0.8450 -2.500 -0.0640 0.05774 0.04925 0.0051 0.4236 0.8201 -2.250 -0.1396 0.05548 0.04734 0.0196 0.4279 0.8010 -2.000 -0.1745 0.05376 0.04566 0.0233 0.4291 0.7571 -1.500 0.0591 0.05789 0.04711 -0.0399 0.4109 0.3792 -1.250 0.1149 0.05529 0.04410 -0.0467 0.4065 0.3483 -1.000 0.1758 0.05397 0.04212 -0.0544 0.4023 0.3319 -0.750 0.2316 0.05240 0.04018 -0.0604 0.3994 0.3285 -0.500 0.2952 0.05117 0.03848 -0.0679 0.3968 0.3242 -0.250 0.3727 0.05043 0.03712 -0.0784 0.3947 0.3259 0.000 0.4460 0.04988 0.03607 -0.0879 0.3938 0.3365 0.250 0.5173 0.04946 0.03548 -0.0970 0.3936 0.3646 0.500 0.6028 0.04893 0.03482 -0.1091 0.3937 0.4263 0.750 0.6760 0.04798 0.03451 -0.1185 0.3936 0.5648 1.250 0.8227 0.04958 0.03728 -0.1393 0.3921 1.0001 1.500 0.8549 0.05127 0.03880 -0.1407 0.3922 1.0001 1.750 0.8847 0.05307 0.04045 -0.1416 0.3929 1.0001 2.000 0.9139 0.05518 0.04240 -0.1424 0.3944 1.0001 2.250 0.9264 0.05591 0.04339 -0.1399 0.3991 1.0001 2.500 0.9396 0.05764 0.04526 -0.1379 0.4043 1.0001 2.750 0.9538 0.05974 0.04742 -0.1363 0.4094 1.0001 3.000 0.9680 0.06208 0.04979 -0.1348 0.4143 1.0001 3.250 0.9856 0.06478 0.05241 -0.1340 0.4186 1.0001 3.500 0.9884 0.06684 0.05466 -0.1308 0.4251 1.0001 3.750 0.9604 0.07016 0.05839 -0.1240 0.4360 1.0001 4.000 0.9651 0.07369 0.06192 -0.1221 0.4440 1.0001 4.250 0.8800 0.08003 0.06879 -0.1103 0.4622 1.0001 4.500 0.8542 0.08558 0.07441 -0.1069 0.4787 1.0001