XFOIL Version 6.94 Calculated polar for: pp13 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 ------- -------- --------- --------- -------- ------- ------- -2.750 -0.3225 0.06257 0.05502 0.0276 0.9998 0.6352 -2.500 -0.3264 0.05895 0.05152 0.0344 0.9998 0.6870 -2.250 -0.3275 0.05548 0.04814 0.0398 0.9998 0.7292 -2.000 -0.3268 0.05203 0.04479 0.0445 0.9998 0.7668 -1.750 -0.3240 0.04928 0.04211 0.0478 0.9998 0.7995 -1.500 -0.3155 0.04564 0.03855 0.0508 0.9998 0.8322 -1.250 -0.3072 0.04289 0.03586 0.0526 0.9998 0.8586 -1.000 -0.2945 0.04034 0.03335 0.0534 0.9998 0.8821 -0.750 -0.2705 0.03789 0.03095 0.0515 0.9998 0.9044 -0.500 -0.2300 0.03603 0.02912 0.0456 0.9998 0.9227 -0.250 -0.1791 0.03515 0.02823 0.0366 0.9998 0.9293 0.000 -0.1283 0.03569 0.02864 0.0272 0.9998 0.9146 0.250 -0.0444 0.03946 0.03205 0.0080 0.9998 0.8366 0.500 0.0517 0.04396 0.03619 -0.0154 0.9998 0.7058 0.750 0.1268 0.04690 0.03879 -0.0321 0.9998 0.5803 1.000 0.1873 0.04858 0.04007 -0.0435 0.9998 0.4749 1.250 0.2367 0.04915 0.04035 -0.0506 0.9998 0.4053 1.500 0.2820 0.04955 0.04049 -0.0561 0.9998 0.3589 1.750 0.3244 0.04995 0.04069 -0.0606 0.9998 0.3332 2.000 0.3656 0.05034 0.04099 -0.0647 0.9998 0.3180 2.250 0.4071 0.05097 0.04147 -0.0687 0.9998 0.3034 2.750 0.8515 0.03957 0.02629 -0.1094 0.2181 0.3395 3.000 0.9220 0.04052 0.02757 -0.1168 0.2077 0.4228 3.250 0.9909 0.04155 0.02849 -0.1218 0.2017 1.0002 3.500 1.0333 0.04299 0.02946 -0.1224 0.1967 1.0002 3.750 1.0835 0.04497 0.03112 -0.1246 0.1912 1.0002 4.000 1.1274 0.04699 0.03320 -0.1258 0.1897 1.0002 4.250 1.1662 0.04935 0.03566 -0.1265 0.1904 1.0002 4.500 1.1974 0.05134 0.03805 -0.1259 0.1925 1.0002 4.750 1.2246 0.05370 0.04098 -0.1249 0.1961 1.0002 5.000 1.2496 0.05675 0.04458 -0.1239 0.2008 1.0002