XFOIL Version 6.94 Calculated polar for: AG455ct -02f rot. 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 0.1821 0.00566 0.00130 -0.0303 0.7054 1.0000 0.500 0.2375 0.00583 0.00127 -0.0299 0.6693 1.0000 1.000 0.2929 0.00600 0.00124 -0.0295 0.6275 1.0000 1.500 0.3480 0.00624 0.00125 -0.0290 0.5792 1.0000 2.000 0.4030 0.00655 0.00132 -0.0286 0.5270 1.0000 2.500 0.4579 0.00691 0.00146 -0.0283 0.4723 1.0000 3.000 0.5128 0.00732 0.00166 -0.0280 0.4170 1.0000 3.500 0.5675 0.00779 0.00194 -0.0277 0.3614 1.0000 4.000 0.6220 0.00831 0.00227 -0.0274 0.3067 1.0000 4.500 0.6763 0.00889 0.00271 -0.0272 0.2536 1.0000 5.000 0.7303 0.00955 0.00321 -0.0269 0.2039 1.0000 5.500 0.7838 0.01028 0.00383 -0.0267 0.1596 1.0000 6.000 0.8370 0.01108 0.00453 -0.0263 0.1212 1.0000 6.500 0.8893 0.01204 0.00539 -0.0259 0.0877 1.0000 7.000 0.9412 0.01306 0.00642 -0.0255 0.0601 1.0000 7.500 0.9912 0.01450 0.00785 -0.0248 0.0377 1.0000 8.000 1.0400 0.01618 0.00969 -0.0239 0.0259 1.0000 8.500 1.0835 0.01887 0.01265 -0.0225 0.0179 1.0000 9.000 1.1297 0.02078 0.01486 -0.0213 0.0152 1.0000 9.500 1.1717 0.02323 0.01756 -0.0200 0.0124 1.0000 10.500 1.2247 0.03317 0.02867 -0.0155 0.0098 1.0000 11.000 1.2348 0.03910 0.03519 -0.0132 0.0091 1.0000 11.500 1.2190 0.04612 0.04272 -0.0117 0.0089 1.0000 12.000 1.1901 0.05755 0.05461 -0.0192 0.0088 1.0000 12.500 1.1520 0.07395 0.07139 -0.0318 0.0090 1.0000 13.000 1.1037 0.09402 0.09174 -0.0453 0.0095 1.0000 13.500 1.0425 0.11810 0.11596 -0.0588 0.0102 1.0000