XFOIL Version 6.94 Calculated polar for: GOE 229 (MVA H.39) 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.7700 0.01171 0.00460 -0.1654 0.5652 0.4576 0.500 0.8259 0.01190 0.00472 -0.1657 0.5560 0.4650 1.000 0.8815 0.01208 0.00486 -0.1660 0.5469 0.4718 1.500 0.9362 0.01225 0.00502 -0.1661 0.5374 0.4778 2.000 0.9902 0.01246 0.00519 -0.1661 0.5275 0.4849 2.500 1.0440 0.01268 0.00543 -0.1661 0.5188 0.4918 3.000 1.0960 0.01289 0.00565 -0.1658 0.5083 0.4990 3.500 1.1266 0.01325 0.00580 -0.1613 0.4555 0.5065 4.000 1.1653 0.01371 0.00617 -0.1586 0.4217 0.5142 4.500 1.1633 0.01600 0.00775 -0.1495 0.3012 0.5209 5.000 1.1563 0.01936 0.01042 -0.1411 0.1998 0.5283 5.500 1.1490 0.02311 0.01365 -0.1338 0.1001 0.5355 6.000 1.1515 0.02668 0.01702 -0.1284 0.0058 0.5438 6.500 1.1834 0.02838 0.01887 -0.1266 0.0050 0.5558 7.000 1.2138 0.03027 0.02090 -0.1248 0.0049 0.5694 7.500 1.2423 0.03237 0.02316 -0.1230 0.0049 0.5850 8.000 1.2682 0.03476 0.02572 -0.1211 0.0050 0.6047 8.500 1.2915 0.03746 0.02862 -0.1192 0.0051 0.6279 9.000 1.3125 0.04050 0.03191 -0.1174 0.0054 0.6559 9.500 1.3308 0.04390 0.03558 -0.1156 0.0056 0.6885 10.000 1.3455 0.04775 0.03978 -0.1138 0.0059 0.7321 10.500 1.3511 0.05184 0.04448 -0.1111 0.0061 1.0000 11.000 1.3560 0.05715 0.04998 -0.1094 0.0064 1.0000 11.500 1.3628 0.06243 0.05540 -0.1082 0.0069 1.0000 12.000 1.3671 0.06814 0.06126 -0.1071 0.0075 1.0000 12.500 1.3695 0.07426 0.06760 -0.1063 0.0082 1.0000 13.000 1.3655 0.08134 0.07489 -0.1056 0.0093 1.0000 13.500 1.3541 0.08958 0.08336 -0.1054 0.0101 1.0000 14.000 1.3361 0.09889 0.09289 -0.1056 0.0107 1.0000 14.500 1.3113 0.10940 0.10362 -0.1064 0.0112 1.0000 15.000 1.3312 0.11372 0.10805 -0.1067 0.0129 1.0000 15.500 1.3265 0.12154 0.11607 -0.1076 0.0154 1.0000