XFOIL Version 6.94 Calculated polar for: AG45ct -02f 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.1848 0.00577 0.00135 -0.0307 0.7001 1.0000 0.500 0.2398 0.00591 0.00131 -0.0302 0.6673 1.0000 1.000 0.2947 0.00605 0.00127 -0.0297 0.6297 1.0000 1.500 0.3495 0.00626 0.00128 -0.0292 0.5850 1.0000 2.000 0.4042 0.00654 0.00134 -0.0287 0.5356 1.0000 2.500 0.4588 0.00689 0.00148 -0.0283 0.4843 1.0000 3.000 0.5133 0.00728 0.00167 -0.0279 0.4321 1.0000 3.500 0.5678 0.00771 0.00194 -0.0276 0.3816 1.0000 4.000 0.6220 0.00821 0.00226 -0.0273 0.3321 1.0000 4.500 0.6763 0.00873 0.00267 -0.0270 0.2849 1.0000 5.000 0.7300 0.00934 0.00313 -0.0267 0.2360 1.0000 5.500 0.7834 0.01003 0.00371 -0.0263 0.1875 1.0000 6.000 0.8361 0.01085 0.00439 -0.0260 0.1426 1.0000 6.500 0.8884 0.01174 0.00521 -0.0256 0.1041 1.0000 7.000 0.9398 0.01278 0.00617 -0.0251 0.0700 1.0000 7.500 0.9900 0.01407 0.00739 -0.0245 0.0421 1.0000 8.000 1.0371 0.01604 0.00946 -0.0234 0.0260 1.0000 8.500 1.0824 0.01821 0.01186 -0.0222 0.0180 1.0000 9.000 1.1261 0.02051 0.01448 -0.0207 0.0152 1.0000 9.500 1.1679 0.02287 0.01708 -0.0193 0.0126 1.0000 10.000 1.1410 0.01897 0.01402 -0.0119 0.0107 1.0000 10.500 1.1635 0.02266 0.01818 -0.0094 0.0100 1.0000 11.000 1.1599 0.02809 0.02409 -0.0055 0.0096 1.0000 11.500 1.1304 0.03512 0.03157 -0.0043 0.0094 1.0000 12.000 1.0919 0.04639 0.04323 -0.0080 0.0094 1.0000 12.500 1.0474 0.06126 0.05841 -0.0139 0.0096 1.0000 13.000 1.0012 0.07785 0.07527 -0.0207 0.0097 1.0000 13.500 0.9476 0.09630 0.09391 -0.0285 0.0102 1.0000 14.000 0.8783 0.11774 0.11553 -0.0380 0.0107 1.0000