XFOIL Version 6.94 Calculated polar for: GOE 223 (MVA H.34) AIRFOIL 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.7840 0.01209 0.00480 -0.1579 0.5077 0.4505 1.000 0.8394 0.01219 0.00491 -0.1581 0.4988 0.4580 1.500 0.8940 0.01249 0.00512 -0.1582 0.4902 0.4648 2.000 0.9497 0.01266 0.00529 -0.1585 0.4840 0.4711 2.500 1.0044 0.01274 0.00546 -0.1586 0.4765 0.4798 3.000 1.0571 0.01299 0.00569 -0.1584 0.4678 0.4881 3.500 1.1010 0.01328 0.00581 -0.1566 0.4294 0.4956 4.000 1.1097 0.01502 0.00682 -0.1492 0.3012 0.5041 4.500 1.1214 0.01743 0.00872 -0.1431 0.2137 0.5132 5.000 1.1509 0.01891 0.01005 -0.1399 0.1832 0.5235 5.500 1.1565 0.02189 0.01265 -0.1340 0.1023 0.5342 6.000 1.1886 0.02338 0.01419 -0.1317 0.0955 0.5480 6.500 1.2196 0.02501 0.01587 -0.1295 0.0821 0.5654 7.000 1.2230 0.02876 0.01948 -0.1249 0.0043 0.5837 7.500 1.2513 0.03084 0.02172 -0.1229 0.0039 0.6098 8.000 1.2771 0.03316 0.02422 -0.1209 0.0038 0.6403 8.500 1.3003 0.03578 0.02705 -0.1189 0.0038 0.6753 9.000 1.3207 0.03873 0.03024 -0.1168 0.0039 0.7155 10.000 1.3493 0.04544 0.03769 -0.1121 0.0041 1.0000 10.500 1.3608 0.04982 0.04222 -0.1104 0.0043 1.0000 11.000 1.3681 0.05486 0.04743 -0.1089 0.0044 1.0000 11.500 1.3736 0.06027 0.05300 -0.1076 0.0046 1.0000 12.000 1.3810 0.06563 0.05850 -0.1067 0.0048 1.0000 12.500 1.3841 0.07166 0.06470 -0.1059 0.0051 1.0000 13.000 1.3813 0.07850 0.07173 -0.1053 0.0054 1.0000 13.500 1.3744 0.08603 0.07943 -0.1049 0.0057 1.0000 14.000 1.3613 0.09448 0.08807 -0.1049 0.0059 1.0000 14.500 1.3482 0.10315 0.09692 -0.1053 0.0061 1.0000 15.000 1.3588 0.10860 0.10252 -0.1056 0.0067 1.0000 15.500 1.3495 0.11695 0.11106 -0.1066 0.0072 1.0000 16.000 1.3363 0.12599 0.12024 -0.1081 0.0077 1.0000 16.500 1.3403 0.13256 0.12694 -0.1094 0.0082 1.0000