XFOIL Version 6.94 Calculated polar for: AG44ct -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 0.2071 0.00579 0.00139 -0.0361 0.7078 1.0000 0.500 0.2616 0.00595 0.00137 -0.0355 0.6775 1.0000 1.000 0.3161 0.00608 0.00134 -0.0349 0.6417 1.0000 1.500 0.3704 0.00628 0.00134 -0.0343 0.5991 1.0000 2.000 0.4244 0.00656 0.00140 -0.0337 0.5486 1.0000 2.500 0.4784 0.00692 0.00153 -0.0332 0.4931 1.0000 3.000 0.5322 0.00734 0.00174 -0.0327 0.4367 1.0000 3.500 0.5859 0.00781 0.00200 -0.0322 0.3844 1.0000 4.000 0.6397 0.00831 0.00235 -0.0318 0.3354 1.0000 4.500 0.6933 0.00885 0.00274 -0.0314 0.2906 1.0000 5.000 0.7467 0.00942 0.00319 -0.0310 0.2486 1.0000 5.500 0.7998 0.01004 0.00374 -0.0306 0.2084 1.0000 6.000 0.8525 0.01074 0.00434 -0.0301 0.1694 1.0000 6.500 0.9045 0.01153 0.00507 -0.0297 0.1314 1.0000 7.000 0.9556 0.01248 0.00592 -0.0291 0.0942 1.0000 7.500 1.0052 0.01370 0.00702 -0.0284 0.0576 1.0000 8.000 1.0528 0.01528 0.00861 -0.0274 0.0323 1.0000 8.500 1.0986 0.01714 0.01062 -0.0261 0.0213 1.0000 9.000 1.1375 0.02004 0.01380 -0.0241 0.0159 1.0000 9.500 1.1783 0.02240 0.01646 -0.0223 0.0137 1.0000 10.000 1.2158 0.02492 0.01922 -0.0205 0.0115 1.0000 10.500 1.1735 0.02093 0.01605 -0.0115 0.0102 1.0000 11.000 1.1705 0.02616 0.02176 -0.0068 0.0098 1.0000 11.500 1.1545 0.03137 0.02736 -0.0040 0.0095 1.0000 12.000 1.1222 0.04064 0.03704 -0.0053 0.0095 1.0000 12.500 1.0849 0.05302 0.04978 -0.0091 0.0095 1.0000 13.000 1.0406 0.06821 0.06527 -0.0146 0.0096 1.0000 13.500 0.9926 0.08483 0.08214 -0.0211 0.0098 1.0000