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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1577 0.06854 0.06052 0.0108 0.3892 0.6938 -2.750 -0.1239 0.06487 0.05661 0.0096 0.3824 0.7325 -2.500 -0.1116 0.06179 0.05346 0.0104 0.3796 0.7517 -2.250 -0.0959 0.05880 0.05047 0.0102 0.3761 0.7635 -2.000 -0.0903 0.05615 0.04781 0.0107 0.3741 0.7648 -1.750 -0.0794 0.05387 0.04546 0.0095 0.3719 0.7573 -1.500 -0.0384 0.05235 0.04373 0.0013 0.3682 0.7343 -1.250 0.2147 0.05775 0.04675 -0.0654 0.3619 0.3332 -1.000 0.2813 0.05584 0.04420 -0.0745 0.3607 0.2945 -0.750 0.3364 0.05430 0.04240 -0.0805 0.3599 0.2880 -0.500 0.3873 0.05338 0.04119 -0.0854 0.3597 0.2899 -0.250 0.4339 0.05268 0.04024 -0.0894 0.3600 0.2917 0.000 0.4814 0.05223 0.03951 -0.0934 0.3608 0.2942 0.250 0.5360 0.05197 0.03904 -0.0991 0.3622 0.3026 0.500 0.5922 0.05210 0.03889 -0.1051 0.3644 0.3191 0.750 0.6542 0.05225 0.03905 -0.1127 0.3673 0.3586 1.000 0.7343 0.05209 0.03920 -0.1242 0.3709 0.4692 1.250 0.7758 0.05281 0.04034 -0.1274 0.3733 0.5465 1.500 0.8077 0.05375 0.04165 -0.1288 0.3755 0.6197 1.750 0.9038 0.05579 0.04476 -0.1451 0.3773 1.0001 2.000 0.9292 0.05805 0.04691 -0.1454 0.3787 1.0001 2.250 0.9567 0.06076 0.04943 -0.1461 0.3807 1.0001 2.500 0.9583 0.06119 0.05034 -0.1418 0.3888 1.0001 2.750 0.9643 0.06365 0.05298 -0.1391 0.3971 1.0001 3.000 0.9806 0.06656 0.05581 -0.1382 0.4037 1.0001 3.250 0.9904 0.06884 0.05818 -0.1363 0.4124 1.0001 3.500 0.9584 0.07241 0.06217 -0.1287 0.4275 1.0001 3.750 0.9372 0.07615 0.06614 -0.1236 0.4425 1.0001 4.000 0.9265 0.08174 0.07180 -0.1213 0.4616 1.0001 4.500 0.2138 0.12022 0.11039 -0.0625 0.8659 0.3082 4.750 0.2302 0.12093 0.11086 -0.0643 0.8559 0.3204 5.000 0.2980 0.12676 0.11674 -0.0770 0.8465 0.3727