XFOIL Version 6.94 Calculated polar for: pp12 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0105 0.05077 0.04015 -0.0167 0.9999 0.9975 -2.750 -0.0447 0.05009 0.03969 -0.0072 0.9999 0.9824 -2.500 -0.0815 0.04953 0.03936 0.0023 0.9999 0.9684 -2.250 -0.1195 0.04880 0.03881 0.0117 0.9999 0.9538 -2.000 -0.1595 0.04791 0.03816 0.0214 0.9999 0.9414 -1.750 -0.2045 0.04684 0.03731 0.0320 0.9999 0.9311 -1.500 -0.2450 0.04577 0.03642 0.0414 0.9999 0.9203 -1.250 -0.2640 0.04504 0.03572 0.0455 0.9999 0.8982 -1.000 -0.2376 0.04585 0.03639 0.0380 0.9999 0.8505 -0.750 -0.1462 0.04870 0.03881 0.0145 0.9999 0.7635 -0.500 -0.0439 0.05122 0.04084 -0.0103 0.9999 0.6642 -0.250 0.0241 0.05235 0.04163 -0.0243 0.9999 0.6020 0.000 0.0790 0.05285 0.04186 -0.0340 0.9999 0.5586 0.250 0.1338 0.05337 0.04206 -0.0431 0.9999 0.5207 0.500 0.1781 0.05338 0.04186 -0.0490 0.9999 0.4943 0.750 0.2241 0.05361 0.04188 -0.0550 0.9999 0.4680 1.000 0.2690 0.05391 0.04193 -0.0605 0.9999 0.4470 1.250 0.3115 0.05440 0.04221 -0.0654 0.9999 0.4318 1.500 0.3534 0.05469 0.04237 -0.0699 0.9999 0.4271 1.750 0.3960 0.05514 0.04276 -0.0742 0.9999 0.4261 2.000 0.4341 0.05583 0.04345 -0.0780 0.9999 0.4224 2.250 0.4777 0.05650 0.04418 -0.0824 0.9999 0.4244 2.500 0.5184 0.05733 0.04515 -0.0862 0.9999 0.4312 2.750 0.5626 0.05812 0.04637 -0.0905 0.9999 0.4483 3.000 0.6093 0.05898 0.04781 -0.0949 0.9999 0.4693 3.500 1.0225 0.04771 0.03320 -0.1183 0.3133 1.0001 3.750 1.0664 0.05039 0.03534 -0.1195 0.2967 1.0001 4.000 1.1101 0.05263 0.03771 -0.1211 0.2911 1.0001 4.250 1.1528 0.05539 0.04073 -0.1227 0.2860 1.0001 4.500 1.1917 0.05835 0.04401 -0.1241 0.2842 1.0001 4.750 1.2254 0.06184 0.04792 -0.1249 0.2828 1.0001 5.000 1.2493 0.06513 0.05169 -0.1249 0.2833 1.0001