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 = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.750 -0.0629 0.04499 0.03538 -0.0490 0.3405 0.2341 -2.500 -0.0564 0.04567 0.03631 -0.0442 0.3334 0.2972 -2.250 -0.0439 0.04479 0.03552 -0.0407 0.3280 0.3411 -2.000 -0.0356 0.04370 0.03461 -0.0356 0.3227 0.4090 -1.750 0.0437 0.03618 0.02445 -0.0481 0.3188 0.1880 -1.500 0.0894 0.03480 0.02241 -0.0515 0.3124 0.1876 -1.250 0.1346 0.03364 0.02076 -0.0547 0.3085 0.1917 -1.000 0.1869 0.03275 0.01960 -0.0595 0.3052 0.1997 -0.750 0.2462 0.03190 0.01850 -0.0659 0.3020 0.2073 -0.500 0.3081 0.03138 0.01776 -0.0729 0.2994 0.2200 -0.250 0.3829 0.03125 0.01750 -0.0829 0.2970 0.2475 0.000 0.4504 0.03110 0.01740 -0.0914 0.2954 0.3313 0.250 0.4998 0.03122 0.01781 -0.0963 0.2946 0.4034 0.500 0.5386 0.03164 0.01835 -0.0986 0.2944 0.4441 0.750 0.5759 0.03210 0.01899 -0.1007 0.2944 0.4820 1.000 0.6116 0.03231 0.01968 -0.1024 0.2948 0.5450 1.250 0.7235 0.03353 0.02166 -0.1215 0.2950 1.0001 1.500 0.7516 0.03447 0.02250 -0.1215 0.2950 1.0001 1.750 0.7784 0.03549 0.02344 -0.1212 0.2947 1.0001 2.000 0.8040 0.03650 0.02446 -0.1207 0.2949 1.0001 2.250 0.8287 0.03753 0.02551 -0.1199 0.2954 1.0001 2.500 0.8523 0.03861 0.02676 -0.1190 0.2976 1.0001 2.750 0.8754 0.03992 0.02822 -0.1181 0.3007 1.0001 3.000 0.8981 0.04138 0.02979 -0.1172 0.3037 1.0001 3.250 0.9202 0.04294 0.03145 -0.1162 0.3067 1.0001 3.500 0.9424 0.04470 0.03325 -0.1154 0.3094 1.0001 3.750 0.9660 0.04684 0.03533 -0.1150 0.3116 1.0001 4.000 0.9784 0.04800 0.03719 -0.1123 0.3235 1.0001 4.250 0.9984 0.05025 0.03950 -0.1114 0.3287 1.0001 4.500 1.0105 0.05248 0.04228 -0.1095 0.3451 1.0001 4.750 1.0341 0.05498 0.04480 -0.1095 0.3541 1.0001