XFOIL Version 6.94 Calculated polar for: manu04 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 -0.3962 0.05977 0.05320 0.0582 0.9999 0.8207 -2.750 -0.4557 0.05731 0.05087 0.0676 0.9999 0.7930 -2.500 -0.4691 0.05556 0.04909 0.0628 0.9999 0.7252 -2.250 -0.3591 0.05738 0.05008 0.0246 0.9999 0.5167 -2.000 -0.2834 0.05567 0.04764 0.0067 0.9999 0.3888 -1.750 -0.2406 0.05337 0.04504 0.0014 0.9999 0.3470 -1.500 -0.0622 0.04522 0.03640 -0.0268 0.8477 0.3233 -1.250 0.0563 0.04058 0.02917 -0.0407 0.5080 0.3154 -1.000 0.0971 0.04008 0.02773 -0.0433 0.4824 0.3173 -0.750 0.1485 0.03945 0.02634 -0.0479 0.4647 0.3252 -0.500 0.2081 0.03880 0.02504 -0.0542 0.4520 0.3472 -0.250 0.2823 0.03800 0.02370 -0.0634 0.4411 0.3819 0.000 0.3733 0.03606 0.02197 -0.0759 0.4303 0.5076 0.250 0.5132 0.03521 0.02179 -0.0997 0.4158 1.0001 0.500 0.5603 0.03602 0.02217 -0.1036 0.4110 1.0001 0.750 0.6051 0.03686 0.02268 -0.1071 0.4080 1.0001 1.000 0.6473 0.03773 0.02334 -0.1100 0.4066 1.0001 1.250 0.6869 0.03865 0.02414 -0.1124 0.4057 1.0001 1.500 0.7232 0.03963 0.02504 -0.1142 0.4055 1.0001 1.750 0.7566 0.04068 0.02608 -0.1155 0.4058 1.0001 2.000 0.7874 0.04182 0.02724 -0.1163 0.4066 1.0001 2.250 0.8160 0.04304 0.02853 -0.1167 0.4079 1.0001 2.500 0.8430 0.04438 0.02993 -0.1169 0.4095 1.0001 2.750 0.8688 0.04583 0.03146 -0.1169 0.4115 1.0001 3.000 0.8931 0.04741 0.03311 -0.1167 0.4136 1.0001 3.250 0.9171 0.04912 0.03488 -0.1165 0.4159 1.0001 3.500 0.9385 0.05081 0.03675 -0.1159 0.4185 1.0001 3.750 0.9514 0.05233 0.03859 -0.1142 0.4227 1.0001 4.000 0.9638 0.05434 0.04084 -0.1125 0.4264 1.0001 4.250 0.9757 0.05654 0.04323 -0.1109 0.4301 1.0001 4.500 0.9868 0.05898 0.04581 -0.1093 0.4332 1.0001 4.750 1.0004 0.06155 0.04852 -0.1082 0.4367 1.0001 5.000 1.0089 0.06420 0.05134 -0.1066 0.4409 1.0001