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 = 0.500 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.500 -0.6783 0.07466 0.07251 0.0094 1.0000 0.0155 -7.250 -0.6769 0.06813 0.06590 0.0065 1.0000 0.0159 -7.000 -0.6667 0.06389 0.06159 0.0047 1.0000 0.0162 -6.750 -0.6535 0.05994 0.05755 0.0031 1.0000 0.0165 -6.500 -0.6381 0.05602 0.05353 0.0015 1.0000 0.0170 -6.250 -0.6206 0.05204 0.04941 0.0000 1.0000 0.0176 -6.000 -0.6011 0.04799 0.04519 -0.0013 1.0000 0.0185 -5.750 -0.5787 0.04388 0.04086 -0.0023 1.0000 0.0200 -5.500 -0.5457 0.04209 0.03873 -0.0020 1.0000 0.0224 -3.750 -0.3666 0.01696 0.01135 -0.0037 1.0000 0.0217 -3.500 -0.3373 0.01513 0.00926 -0.0038 1.0000 0.0218 -3.250 -0.3079 0.01410 0.00808 -0.0040 1.0000 0.0223 -3.000 -0.2796 0.01202 0.00585 -0.0041 1.0000 0.0240 -2.750 -0.2510 0.01093 0.00472 -0.0043 1.0000 0.0252 -2.500 -0.2222 0.01024 0.00402 -0.0046 1.0000 0.0269 -2.250 -0.1934 0.00969 0.00344 -0.0050 1.0000 0.0291 -2.000 -0.1636 0.00930 0.00304 -0.0055 0.9968 0.0317 -1.750 -0.1281 0.00869 0.00239 -0.0072 0.9860 0.0373 -1.500 -0.0970 0.00846 0.00215 -0.0077 0.9739 0.0441 -1.250 -0.0724 0.00819 0.00191 -0.0068 0.9612 0.0655 -1.000 -0.0499 0.00582 0.00166 -0.0071 0.9512 0.6598 -0.750 -0.0353 0.00534 0.00190 -0.0033 0.9401 0.8630 -0.500 -0.0207 0.00541 0.00203 0.0006 0.9287 0.9176 -0.250 -0.0112 0.00538 0.00202 0.0057 0.9168 0.9527 0.000 0.0082 0.00524 0.00185 0.0086 0.9056 0.9816 0.250 0.0612 0.00525 0.00181 0.0036 0.8942 1.0000 0.500 0.0846 0.00519 0.00168 0.0046 0.8784 1.0000 0.750 0.1087 0.00513 0.00157 0.0055 0.8597 1.0000 1.000 0.1329 0.00510 0.00147 0.0064 0.8381 1.0000 1.250 0.1578 0.00511 0.00139 0.0071 0.8119 1.0000 1.500 0.1833 0.00517 0.00132 0.0078 0.7790 1.0000 1.750 0.2099 0.00529 0.00129 0.0081 0.7381 1.0000 2.000 0.2373 0.00547 0.00129 0.0082 0.6868 1.0000 2.250 0.2653 0.00576 0.00131 0.0081 0.6158 1.0000 2.500 0.2938 0.00615 0.00137 0.0078 0.5310 1.0000 2.750 0.3226 0.00661 0.00149 0.0072 0.4442 1.0000 3.000 0.3516 0.00696 0.00165 0.0067 0.3917 1.0000 3.250 0.3806 0.00726 0.00180 0.0062 0.3536 1.0000 3.500 0.4095 0.00756 0.00196 0.0057 0.3144 1.0000 3.750 0.4384 0.00789 0.00212 0.0052 0.2688 1.0000 4.000 0.4672 0.00822 0.00231 0.0048 0.2259 1.0000 4.250 0.4959 0.00860 0.00253 0.0043 0.1834 1.0000 4.750 0.5530 0.00944 0.00307 0.0034 0.1076 1.0000 5.000 0.5809 0.01031 0.00360 0.0028 0.0433 1.0000 5.250 0.6093 0.01104 0.00432 0.0025 0.0281 1.0000 5.500 0.6373 0.01190 0.00524 0.0023 0.0221 1.0000 5.750 0.6652 0.01264 0.00608 0.0021 0.0202 1.0000 6.000 0.6928 0.01341 0.00695 0.0020 0.0184 1.0000 6.250 0.7200 0.01415 0.00774 0.0018 0.0165 1.0000 6.500 0.7449 0.01600 0.00973 0.0018 0.0146 1.0000 6.750 0.7707 0.01750 0.01140 0.0019 0.0141 1.0000 7.000 0.7969 0.01869 0.01275 0.0021 0.0135 1.0000 7.250 0.8223 0.02033 0.01461 0.0023 0.0130 1.0000 7.500 0.8464 0.02258 0.01718 0.0026 0.0126 1.0000 7.750 0.8682 0.02581 0.02084 0.0030 0.0127 1.0000 8.000 0.8847 0.03064 0.02622 0.0035 0.0134 1.0000 13.750 0.6176 0.15277 0.15070 -0.0380 0.0172 1.0000 14.000 0.6185 0.15550 0.15343 -0.0390 0.0166 1.0000