XFOIL Version 6.94 Calculated polar for: eif385mod5 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 0.0794 0.03961 0.03140 -0.0238 0.4499 0.7155 -2.750 0.1148 0.03952 0.03049 -0.0278 0.4048 0.7151 -2.500 0.1549 0.03922 0.02966 -0.0327 0.3841 0.7124 -2.250 0.1989 0.03881 0.02894 -0.0381 0.3686 0.7101 -2.000 0.2444 0.03874 0.02842 -0.0437 0.3567 0.7070 -1.750 0.3011 0.03898 0.02835 -0.0520 0.3478 0.6929 -1.500 0.3530 0.03917 0.02823 -0.0585 0.3422 0.6748 -1.250 0.4080 0.03990 0.02855 -0.0655 0.3377 0.6489 -1.000 0.4590 0.04067 0.02896 -0.0712 0.3343 0.6261 -0.750 0.5074 0.04155 0.02953 -0.0759 0.3318 0.6084 -0.500 0.5539 0.04274 0.03042 -0.0801 0.3302 0.5893 -0.250 0.5941 0.04360 0.03113 -0.0828 0.3294 0.5788 0.000 0.6336 0.04478 0.03217 -0.0853 0.3291 0.5658 0.250 0.6701 0.04585 0.03321 -0.0872 0.3291 0.5534 0.500 0.7057 0.04727 0.03454 -0.0889 0.3289 0.5420 0.750 0.7383 0.04838 0.03574 -0.0901 0.3285 0.5364 1.000 0.7700 0.04965 0.03708 -0.0911 0.3279 0.5320 1.250 0.8009 0.05113 0.03863 -0.0921 0.3275 0.5263 1.500 0.8313 0.05298 0.04050 -0.0931 0.3276 0.5190 1.750 0.8601 0.05488 0.04250 -0.0938 0.3287 0.5136 2.000 0.8887 0.05643 0.04426 -0.0945 0.3311 0.5081 2.250 0.9132 0.05813 0.04648 -0.0953 0.3366 0.5033 2.500 0.9350 0.06095 0.04968 -0.0962 0.3429 0.4991 2.750 0.9562 0.06409 0.05303 -0.0970 0.3488 0.4954 3.000 0.9772 0.06732 0.05639 -0.0976 0.3541 0.4924 3.250 0.9854 0.07134 0.06096 -0.0986 0.3653 0.4908 3.500 0.9838 0.07741 0.06741 -0.0999 0.3802 0.4899 4.500 0.5924 0.13281 0.12414 -0.1199 0.7697 0.5151 4.750 0.6029 0.13513 0.12641 -0.1196 0.7555 0.5112 5.000 0.6166 0.13788 0.12908 -0.1199 0.7414 0.5068