XFOIL Version 6.96 Calculated polar for: ONERA OA206 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- 0.250 0.0467 0.00418 0.00122 0.0057 0.8095 0.9246 0.500 0.0688 0.00423 0.00120 0.0074 0.7822 0.9432 0.750 0.0856 0.00420 0.00111 0.0105 0.7525 0.9641 1.000 0.1054 0.00416 0.00096 0.0127 0.7166 0.9825 1.250 0.1358 0.00429 0.00091 0.0122 0.6685 0.9883 1.500 0.1677 0.00448 0.00090 0.0112 0.6103 0.9927 1.750 0.2002 0.00479 0.00092 0.0100 0.5303 0.9969 2.000 0.2312 0.00512 0.00097 0.0090 0.4540 1.0000 2.250 0.2605 0.00542 0.00105 0.0084 0.3954 1.0000 2.500 0.2900 0.00563 0.00113 0.0078 0.3608 1.0000 2.750 0.3197 0.00581 0.00122 0.0072 0.3334 1.0000 3.000 0.3494 0.00600 0.00132 0.0067 0.3063 1.0000 3.250 0.3790 0.00626 0.00142 0.0061 0.2639 1.0000 3.500 0.4086 0.00654 0.00154 0.0054 0.2240 1.0000 3.750 0.4380 0.00684 0.00169 0.0048 0.1842 1.0000 4.000 0.4674 0.00716 0.00187 0.0042 0.1495 1.0000 4.250 0.4968 0.00745 0.00206 0.0037 0.1219 1.0000 4.500 0.5260 0.00778 0.00227 0.0031 0.0941 1.0000 4.750 0.5549 0.00829 0.00257 0.0025 0.0510 1.0000 5.000 0.5838 0.00886 0.00299 0.0020 0.0243 1.0000 5.250 0.6128 0.00923 0.00339 0.0017 0.0187 1.0000 5.500 0.6415 0.00992 0.00417 0.0013 0.0145 1.0000 5.750 0.6702 0.01019 0.00445 0.0010 0.0135 1.0000 6.000 0.6987 0.01055 0.00485 0.0007 0.0120 1.0000 6.250 0.7269 0.01105 0.00540 0.0004 0.0108 1.0000 6.500 0.7537 0.01245 0.00697 0.0002 0.0095 1.0000 6.750 0.7813 0.01304 0.00763 0.0000 0.0092 1.0000 7.000 0.8087 0.01366 0.00833 -0.0002 0.0088 1.0000 7.250 0.8357 0.01441 0.00919 -0.0003 0.0082 1.0000 7.500 0.8621 0.01533 0.01022 -0.0003 0.0077 1.0000 7.750 0.8879 0.01645 0.01146 -0.0003 0.0073 1.0000 8.000 0.9130 0.01771 0.01286 -0.0003 0.0071 1.0000 8.250 0.9380 0.01875 0.01402 -0.0003 0.0068 1.0000 8.500 0.9612 0.02039 0.01582 -0.0003 0.0065 1.0000 8.750 0.9816 0.02294 0.01867 0.0000 0.0063 1.0000 9.000 1.0000 0.02584 0.02191 0.0004 0.0063 1.0000 9.250 1.0154 0.02914 0.02557 0.0007 0.0063 1.0000 9.500 1.0273 0.03280 0.02960 0.0011 0.0062 1.0000 9.750 1.0342 0.03692 0.03410 0.0014 0.0062 1.0000 10.000 1.0354 0.04143 0.03896 0.0014 0.0063 1.0000 10.250 1.0292 0.04636 0.04420 0.0010 0.0063 1.0000 10.500 1.0157 0.05094 0.04901 -0.0001 0.0063 1.0000