XFOIL Version 6.94 Calculated polar for: agora02 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.0322 0.05411 0.03943 -0.0053 0.9988 1.0012 -2.750 -0.0526 0.05222 0.03763 -0.0031 0.9988 1.0012 -2.500 -0.0719 0.05027 0.03556 -0.0017 0.9988 1.0012 -2.250 -0.0621 0.04909 0.03363 -0.0057 0.9988 1.0012 -2.000 -0.0333 0.04913 0.03263 -0.0102 0.9988 1.0012 -1.750 -0.0066 0.04960 0.03230 -0.0122 0.9988 1.0012 -1.500 0.0175 0.05020 0.03231 -0.0131 0.9988 1.0012 -1.250 0.0402 0.05089 0.03257 -0.0135 0.9988 1.0012 -1.000 0.0619 0.05166 0.03303 -0.0138 0.9988 1.0012 -0.750 0.0829 0.05253 0.03368 -0.0140 0.9988 1.0012 -0.500 0.1032 0.05352 0.03456 -0.0143 0.9988 1.0012 -0.250 0.1224 0.05469 0.03570 -0.0147 0.9988 1.0012 0.000 0.1400 0.05614 0.03722 -0.0154 0.9988 1.0012 0.250 0.1542 0.05814 0.03936 -0.0167 0.9988 1.0012 0.500 0.1618 0.06113 0.04250 -0.0191 0.9988 1.0012 0.750 0.1637 0.06499 0.04636 -0.0223 0.9988 1.0012 1.000 0.1669 0.06876 0.05006 -0.0253 0.9988 1.0012 1.250 0.1724 0.07226 0.05345 -0.0279 0.9988 1.0012 1.500 0.1795 0.07558 0.05666 -0.0302 0.9988 1.0012 1.750 0.1880 0.07883 0.05978 -0.0324 0.9988 1.0012 2.000 0.1974 0.08200 0.06283 -0.0344 0.9988 1.0012 2.250 0.2072 0.08513 0.06583 -0.0363 0.9988 1.0012 2.500 0.2175 0.08823 0.06881 -0.0381 0.9988 1.0012 2.750 0.2281 0.09132 0.07177 -0.0399 0.9988 1.0012 3.000 0.2390 0.09441 0.07474 -0.0416 0.9988 1.0012 3.250 0.2501 0.09751 0.07771 -0.0432 0.9988 1.0012 3.500 0.2613 0.10060 0.08070 -0.0449 0.9988 1.0012 3.750 0.2728 0.10370 0.08368 -0.0465 0.9988 1.0012 4.000 0.2844 0.10681 0.08667 -0.0480 0.9988 1.0012 4.250 0.2961 0.10992 0.08967 -0.0496 0.9988 1.0012 4.500 0.3079 0.11303 0.09268 -0.0511 0.9988 1.0012 4.750 0.3198 0.11616 0.09570 -0.0527 0.9988 1.0012 5.000 0.3317 0.11930 0.09873 -0.0542 0.9988 1.0012