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 -0.1942 0.05152 0.04428 0.0138 0.4356 0.7682 -2.750 -0.2034 0.04960 0.04230 0.0173 0.4262 0.7636 -2.500 -0.2193 0.04790 0.04057 0.0207 0.4219 0.7517 -2.250 -0.2094 0.04640 0.03886 0.0187 0.4134 0.7297 -2.000 -0.0248 0.04933 0.03895 -0.0387 0.3942 0.2944 -1.750 0.0109 0.04704 0.03619 -0.0408 0.3883 0.2715 -1.500 0.0508 0.04511 0.03356 -0.0435 0.3824 0.2569 -1.250 0.0955 0.04347 0.03143 -0.0468 0.3750 0.2530 -1.000 0.1468 0.04220 0.02959 -0.0514 0.3687 0.2555 -0.750 0.1939 0.04103 0.02808 -0.0551 0.3624 0.2611 -0.500 0.2504 0.04004 0.02672 -0.0608 0.3565 0.2676 -0.250 0.3358 0.03929 0.02532 -0.0725 0.3522 0.2828 0.000 0.4426 0.03872 0.02428 -0.0891 0.3487 0.3319 0.250 0.5326 0.03747 0.02352 -0.1021 0.3474 0.4731 0.500 0.5825 0.03732 0.02393 -0.1067 0.3470 0.5603 0.750 0.6977 0.03783 0.02545 -0.1262 0.3469 1.0001 1.000 0.7335 0.03906 0.02644 -0.1279 0.3475 1.0001 1.250 0.7659 0.04038 0.02754 -0.1289 0.3483 1.0001 1.500 0.7953 0.04146 0.02853 -0.1292 0.3496 1.0001 1.750 0.8212 0.04231 0.02946 -0.1287 0.3520 1.0001 2.000 0.8461 0.04346 0.03069 -0.1282 0.3546 1.0001 2.250 0.8695 0.04478 0.03207 -0.1275 0.3565 1.0001 2.500 0.8915 0.04621 0.03356 -0.1266 0.3584 1.0001 2.750 0.9123 0.04773 0.03515 -0.1255 0.3601 1.0001 3.000 0.9315 0.04937 0.03686 -0.1242 0.3619 1.0001 3.250 0.9497 0.05111 0.03871 -0.1229 0.3642 1.0001 3.500 0.9680 0.05307 0.04073 -0.1216 0.3671 1.0001 3.750 0.9888 0.05539 0.04302 -0.1209 0.3703 1.0001 4.000 0.9956 0.05674 0.04485 -0.1181 0.3792 1.0001 4.250 1.0038 0.05929 0.04762 -0.1159 0.3875 1.0001 4.500 1.0231 0.06215 0.05051 -0.1154 0.3933 1.0001 4.750 1.0085 0.06489 0.05377 -0.1110 0.4098 1.0001 5.000 1.0210 0.06796 0.05696 -0.1103 0.4217 1.0001