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 = 0.060 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -2.500 -0.0393 0.04542 0.03542 -0.0478 0.3216 0.2125 -2.250 -0.0096 0.04259 0.03196 -0.0485 0.3169 0.1982 -2.000 0.0233 0.04040 0.02939 -0.0496 0.3121 0.1925 -1.750 0.0608 0.03770 0.02582 -0.0511 0.3080 0.1858 -1.500 0.1105 0.03604 0.02366 -0.0553 0.3043 0.1858 -1.250 0.1673 0.03514 0.02260 -0.0612 0.3010 0.1909 -1.000 0.2298 0.03427 0.02120 -0.0681 0.2981 0.1982 -0.750 0.2920 0.03370 0.02024 -0.0752 0.2958 0.2050 -0.500 0.3441 0.03339 0.01959 -0.0800 0.2946 0.2163 -0.250 0.4028 0.03326 0.01946 -0.0866 0.2937 0.2356 0.000 0.4607 0.03338 0.01969 -0.0931 0.2932 0.2813 0.250 0.5234 0.03326 0.01986 -0.1008 0.2929 0.3809 0.500 0.5641 0.03376 0.02050 -0.1035 0.2930 0.4276 0.750 0.6022 0.03434 0.02125 -0.1058 0.2934 0.4659 1.000 0.6379 0.03488 0.02206 -0.1076 0.2936 0.5073 1.250 0.6724 0.03513 0.02291 -0.1092 0.2935 0.5939 1.500 0.7790 0.03681 0.02515 -0.1272 0.2927 1.0001 1.750 0.8059 0.03795 0.02621 -0.1269 0.2928 1.0001 2.000 0.8322 0.03923 0.02740 -0.1266 0.2936 1.0001 2.250 0.8582 0.04077 0.02884 -0.1264 0.2949 1.0001 2.500 0.8809 0.04135 0.02964 -0.1250 0.2979 1.0001 2.750 0.9018 0.04264 0.03124 -0.1235 0.3032 1.0001 3.000 0.9230 0.04428 0.03300 -0.1224 0.3075 1.0001 3.250 0.9443 0.04607 0.03484 -0.1214 0.3112 1.0001 3.500 0.9667 0.04821 0.03693 -0.1208 0.3141 1.0001 3.750 0.9802 0.04959 0.03890 -0.1183 0.3272 1.0001 4.000 1.0002 0.05206 0.04141 -0.1175 0.3334 1.0001 4.250 1.0129 0.05446 0.04429 -0.1158 0.3531 1.0001 4.750 0.5348 0.09815 0.09065 -0.1016 0.7345 1.0001 5.000 0.5605 0.09955 0.09188 -0.1012 0.7034 1.0001