XFOIL Version 6.94 Calculated polar for: M6 (85%) 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 0.1876 0.00828 0.00358 -0.0211 0.6869 0.9672 0.500 0.2619 0.00850 0.00366 -0.0246 0.6659 0.9744 1.000 0.3177 0.00843 0.00347 -0.0250 0.6459 0.9744 1.500 0.3734 0.00840 0.00333 -0.0252 0.6250 0.9744 2.000 0.4287 0.00838 0.00322 -0.0253 0.5999 0.9744 2.500 0.4840 0.00841 0.00318 -0.0254 0.5779 0.9744 3.000 0.5385 0.00846 0.00320 -0.0254 0.5554 0.9744 3.500 0.5920 0.00857 0.00325 -0.0251 0.5284 0.9744 4.000 0.6448 0.00872 0.00336 -0.0246 0.4996 0.9744 4.500 0.6965 0.00895 0.00351 -0.0240 0.4589 0.9744 5.000 0.7467 0.00934 0.00374 -0.0232 0.4025 0.9744 5.500 0.7948 0.01001 0.00412 -0.0222 0.3246 0.9744 6.000 0.8401 0.01106 0.00476 -0.0209 0.2277 0.9744 6.500 0.8801 0.01271 0.00580 -0.0191 0.1074 0.9744 7.000 0.9181 0.01434 0.00706 -0.0166 0.0470 0.9744 7.500 0.9559 0.01573 0.00839 -0.0138 0.0324 0.9744 8.000 0.9920 0.01705 0.00975 -0.0108 0.0281 0.9744 8.500 1.0241 0.01845 0.01124 -0.0071 0.0260 0.9744 9.000 1.0508 0.02005 0.01291 -0.0027 0.0248 0.9744 9.500 1.0748 0.02166 0.01464 0.0021 0.0238 0.9744 10.000 1.0936 0.02345 0.01649 0.0074 0.0232 0.9744 10.500 1.1055 0.02549 0.01857 0.0136 0.0228 0.9744 11.000 1.1227 0.02753 0.02083 0.0186 0.0224 0.9744 11.500 1.1420 0.02990 0.02340 0.0226 0.0220 0.9744 12.000 1.1628 0.03260 0.02631 0.0260 0.0217 0.9744 12.500 1.1814 0.03573 0.02970 0.0293 0.0215 0.9744 13.000 1.1931 0.03945 0.03374 0.0324 0.0214 0.9744 13.500 1.1945 0.04413 0.03881 0.0354 0.0215 0.9744 14.000 1.1830 0.05013 0.04526 0.0375 0.0217 0.9744 14.500 1.1592 0.05767 0.05324 0.0382 0.0220 0.9744 15.000 1.1280 0.06657 0.06254 0.0367 0.0222 0.9744 15.500 1.0952 0.07656 0.07287 0.0334 0.0224 0.9744 16.000 1.0588 0.08821 0.08483 0.0280 0.0226 0.9744 16.500 1.0242 0.10103 0.09790 0.0210 0.0227 0.9744