XFOIL Version 6.94 Calculated polar for: NACA 4413 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 0.0037 0.02532 0.01632 -0.0923 0.8984 0.1921 -2.750 0.0531 0.02450 0.01546 -0.0961 0.8946 0.2171 -2.500 0.0763 0.02412 0.01515 -0.0953 0.8819 0.2375 -2.250 0.1237 0.02328 0.01448 -0.0988 0.8775 0.2720 -2.000 0.1509 0.02284 0.01419 -0.0986 0.8661 0.3035 -1.750 0.1934 0.02197 0.01352 -0.1010 0.8604 0.3536 -1.500 0.2369 0.02085 0.01281 -0.1033 0.8567 0.4251 -1.250 0.2566 0.02030 0.01281 -0.1016 0.8435 0.5263 -1.000 0.2869 0.01888 0.01252 -0.0991 0.8392 0.8406 -0.750 0.3497 0.01862 0.01211 -0.1053 0.8299 1.0000 -0.500 0.3869 0.01832 0.01159 -0.1066 0.8221 1.0000 -0.250 0.4135 0.01834 0.01143 -0.1061 0.8107 1.0000 0.000 0.4491 0.01806 0.01098 -0.1070 0.8026 1.0000 0.250 0.4741 0.01816 0.01094 -0.1063 0.7909 1.0000 0.500 0.5089 0.01791 0.01054 -0.1070 0.7830 1.0000 0.750 0.5326 0.01809 0.01062 -0.1060 0.7709 1.0000 1.000 0.5671 0.01788 0.01026 -0.1067 0.7631 1.0000 1.250 0.5895 0.01815 0.01046 -0.1056 0.7509 1.0000 1.500 0.6236 0.01799 0.01016 -0.1061 0.7433 1.0000 1.750 0.6451 0.01834 0.01047 -0.1050 0.7310 1.0000 2.000 0.6788 0.01823 0.01023 -0.1055 0.7237 1.0000 2.250 0.6995 0.01865 0.01063 -0.1043 0.7116 1.0000 2.500 0.7328 0.01860 0.01045 -0.1048 0.7043 1.0000 2.750 0.7529 0.01908 0.01094 -0.1035 0.6926 1.0000 3.000 0.7855 0.01908 0.01083 -0.1039 0.6853 1.0000 3.250 0.8056 0.01959 0.01138 -0.1027 0.6741 1.0000 3.500 0.8372 0.01964 0.01134 -0.1030 0.6666 1.0000 3.750 0.8578 0.02017 0.01190 -0.1018 0.6559 1.0000 4.000 0.8881 0.02028 0.01194 -0.1019 0.6482 1.0000 4.250 0.9096 0.02080 0.01252 -0.1009 0.6381 1.0000 4.500 0.9386 0.02097 0.01264 -0.1008 0.6298 1.0000 4.750 0.9612 0.02147 0.01318 -0.1000 0.6203 1.0000 5.000 0.9887 0.02169 0.01340 -0.0997 0.6114 1.0000