XFOIL Version 6.94 Calculated polar for: manu4/30 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0293 0.09151 0.07180 -0.0178 0.9999 1.0001 -2.750 -0.0287 0.08951 0.07006 -0.0169 0.9999 1.0001 -2.500 -0.0288 0.08749 0.06834 -0.0159 0.9999 1.0001 -2.250 -0.0299 0.08544 0.06659 -0.0147 0.9999 1.0001 -2.000 -0.0321 0.08338 0.06483 -0.0134 0.9999 1.0001 -1.750 -0.0357 0.08126 0.06305 -0.0119 0.9999 1.0001 -1.500 -0.0408 0.07909 0.06122 -0.0102 0.9999 1.0001 -1.250 -0.0462 0.07694 0.05939 -0.0085 0.9999 1.0001 -1.000 -0.0496 0.07489 0.05754 -0.0073 0.9999 1.0001 -0.750 -0.0470 0.07317 0.05588 -0.0072 0.9999 1.0001 -0.500 -0.0355 0.07199 0.05458 -0.0087 0.9999 1.0001 -0.250 -0.0134 0.07146 0.05376 -0.0121 0.9999 1.0001 0.000 0.0177 0.07160 0.05356 -0.0170 0.9999 1.0001 0.250 0.0533 0.07243 0.05401 -0.0228 0.9999 1.0001 0.500 0.0868 0.07398 0.05523 -0.0284 0.9999 1.0001 0.750 0.1089 0.07642 0.05747 -0.0329 0.9999 1.0001 1.000 0.1138 0.07987 0.06079 -0.0357 0.9999 1.0001 1.250 0.1119 0.08382 0.06442 -0.0378 0.9999 1.0001 1.500 0.1154 0.08757 0.06768 -0.0403 0.9999 1.0001 1.750 0.1242 0.09107 0.07059 -0.0431 0.9999 1.0001 2.000 0.1357 0.09440 0.07324 -0.0458 0.9999 1.0001 2.250 0.1484 0.09758 0.07576 -0.0483 0.9999 1.0001 2.500 0.1614 0.10066 0.07818 -0.0506 0.9999 1.0001 2.750 0.1746 0.10364 0.08052 -0.0525 0.9999 1.0001 3.000 0.1878 0.10653 0.08278 -0.0542 0.9999 1.0001 3.250 0.2007 0.10936 0.08502 -0.0558 0.9999 1.0001 3.500 0.2137 0.11212 0.08722 -0.0571 0.9999 1.0001 3.750 0.2265 0.11485 0.08938 -0.0584 0.9999 1.0001 4.000 0.2393 0.11754 0.09156 -0.0595 0.9999 1.0001 4.250 0.2519 0.12020 0.09371 -0.0605 0.9999 1.0001 4.500 0.2645 0.12284 0.09590 -0.0615 0.9999 1.0001 4.750 0.2770 0.12545 0.09806 -0.0625 0.9999 1.0001 5.000 0.2895 0.12805 0.10023 -0.0633 0.9999 1.0001