XFOIL Version 6.94 Calculated polar for: manu4/36 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 -0.2190 0.05568 0.05009 0.0227 0.9170 0.8246 -2.750 -0.3007 0.05525 0.04984 0.0384 0.9406 0.8087 -2.500 -0.4170 0.05598 0.05071 0.0627 0.9847 0.7958 -2.000 -0.2472 0.05267 0.04681 0.0100 0.6404 0.5265 -1.500 0.0297 0.04490 0.03347 -0.0401 0.4274 0.2854 -1.250 0.0675 0.04345 0.03157 -0.0420 0.4188 0.2825 -1.000 0.1113 0.04213 0.02976 -0.0451 0.4109 0.2815 -0.750 0.1670 0.04111 0.02799 -0.0505 0.4028 0.2844 -0.500 0.2261 0.04018 0.02663 -0.0567 0.3949 0.2956 -0.250 0.2944 0.03955 0.02546 -0.0649 0.3862 0.3209 0.000 0.3999 0.03848 0.02390 -0.0809 0.3782 0.3854 0.250 0.4886 0.03646 0.02290 -0.0932 0.3751 0.5751 0.500 0.6237 0.03660 0.02375 -0.1167 0.3721 1.0001 0.750 0.6680 0.03758 0.02442 -0.1201 0.3719 1.0001 1.000 0.7061 0.03858 0.02521 -0.1222 0.3720 1.0001 1.250 0.7403 0.03965 0.02613 -0.1235 0.3726 1.0001 1.500 0.7718 0.04078 0.02717 -0.1243 0.3736 1.0001 1.750 0.8014 0.04202 0.02836 -0.1247 0.3748 1.0001 2.000 0.8293 0.04335 0.02964 -0.1249 0.3763 1.0001 2.250 0.8561 0.04483 0.03106 -0.1249 0.3779 1.0001 2.500 0.8816 0.04643 0.03263 -0.1247 0.3792 1.0001 2.750 0.9024 0.04750 0.03392 -0.1236 0.3815 1.0001 3.000 0.9214 0.04883 0.03544 -0.1223 0.3839 1.0001 3.250 0.9392 0.05040 0.03719 -0.1210 0.3865 1.0001 3.500 0.9551 0.05217 0.03912 -0.1194 0.3889 1.0001 3.750 0.9695 0.05411 0.04121 -0.1177 0.3917 1.0001 4.000 0.9842 0.05624 0.04346 -0.1161 0.3954 1.0001 4.250 1.0019 0.05862 0.04593 -0.1152 0.3995 1.0001 4.500 1.0062 0.06069 0.04833 -0.1126 0.4071 1.0001 4.750 1.0043 0.06369 0.05161 -0.1097 0.4166 1.0001 5.000 1.0235 0.06677 0.05470 -0.1095 0.4228 1.0001