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 = 0.040 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.2032 0.05165 0.04458 0.0128 0.4470 0.7433 -2.750 -0.2179 0.04988 0.04272 0.0175 0.4363 0.7456 -2.500 -0.2180 0.04808 0.04075 0.0195 0.4226 0.7433 -2.250 -0.2113 0.04664 0.03909 0.0185 0.4132 0.7266 -2.000 -0.0254 0.04885 0.03840 -0.0390 0.3900 0.2764 -1.750 0.0090 0.04678 0.03576 -0.0408 0.3830 0.2596 -1.500 0.0450 0.04515 0.03369 -0.0426 0.3768 0.2571 -1.250 0.0926 0.04367 0.03160 -0.0465 0.3711 0.2576 -1.000 0.1406 0.04215 0.02967 -0.0504 0.3668 0.2561 -0.750 0.1933 0.04087 0.02793 -0.0552 0.3627 0.2580 -0.500 0.2560 0.03983 0.02630 -0.0619 0.3591 0.2648 -0.250 0.3347 0.03921 0.02532 -0.0724 0.3555 0.2874 0.000 0.4362 0.03843 0.02412 -0.0877 0.3525 0.3399 0.250 0.5375 0.03708 0.02335 -0.1032 0.3508 0.5016 0.500 0.5868 0.03705 0.02386 -0.1077 0.3498 0.6027 0.750 0.6990 0.03781 0.02532 -0.1266 0.3474 1.0001 1.000 0.7322 0.03886 0.02618 -0.1277 0.3464 1.0001 1.250 0.7623 0.03996 0.02713 -0.1282 0.3457 1.0001 1.500 0.7912 0.04116 0.02821 -0.1285 0.3460 1.0001 1.750 0.8193 0.04250 0.02945 -0.1286 0.3472 1.0001 2.000 0.8464 0.04397 0.03082 -0.1286 0.3485 1.0001 2.250 0.8723 0.04550 0.03229 -0.1284 0.3499 1.0001 2.500 0.8922 0.04622 0.03330 -0.1268 0.3537 1.0001 2.750 0.9121 0.04764 0.03492 -0.1255 0.3580 1.0001 3.000 0.9315 0.04931 0.03674 -0.1243 0.3622 1.0001 3.250 0.9506 0.05119 0.03869 -0.1232 0.3661 1.0001 3.500 0.9710 0.05330 0.04082 -0.1223 0.3695 1.0001 3.750 0.9876 0.05485 0.04258 -0.1208 0.3752 1.0001 4.000 0.9930 0.05698 0.04511 -0.1180 0.3849 1.0001 4.250 1.0095 0.05957 0.04774 -0.1170 0.3913 1.0001 4.500 1.0076 0.06189 0.05048 -0.1138 0.4042 1.0001 4.750 1.0198 0.06513 0.05385 -0.1128 0.4145 1.0001 5.000 0.9992 0.06911 0.05824 -0.1088 0.4354 1.0001