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 = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.7664 0.01409 0.00660 -0.1646 0.5932 0.4808 0.500 0.8214 0.01420 0.00666 -0.1648 0.5832 0.4874 1.000 0.8770 0.01425 0.00666 -0.1652 0.5734 0.4944 1.500 0.9322 0.01447 0.00685 -0.1655 0.5638 0.5013 2.000 0.9861 0.01465 0.00697 -0.1655 0.5535 0.5091 2.500 1.0414 0.01487 0.00716 -0.1659 0.5432 0.5167 3.000 1.0906 0.01498 0.00734 -0.1650 0.5304 0.5247 3.500 1.1241 0.01513 0.00723 -0.1609 0.4893 0.5327 4.000 1.1637 0.01540 0.00749 -0.1584 0.4626 0.5421 4.500 1.1881 0.01605 0.00794 -0.1531 0.4158 0.5512 5.000 1.2034 0.01756 0.00909 -0.1471 0.3472 0.5602 5.500 1.1928 0.02095 0.01189 -0.1384 0.2497 0.5690 6.000 1.2061 0.02347 0.01421 -0.1338 0.2098 0.5807 6.500 1.2226 0.02599 0.01663 -0.1300 0.1775 0.5940 7.500 1.1974 0.03623 0.02632 -0.1187 0.0066 0.6195 8.000 1.2219 0.03881 0.02908 -0.1169 0.0065 0.6396 8.500 1.2441 0.04172 0.03222 -0.1152 0.0065 0.6647 9.000 1.2640 0.04495 0.03572 -0.1135 0.0067 0.6959 9.500 1.2815 0.04850 0.03965 -0.1119 0.0069 0.7445 10.500 1.3032 0.05662 0.04842 -0.1080 0.0075 1.0000 11.000 1.3132 0.06153 0.05352 -0.1068 0.0080 1.0000 11.500 1.3175 0.06726 0.05946 -0.1057 0.0085 1.0000 12.000 1.3144 0.07401 0.06644 -0.1047 0.0089 1.0000 12.500 1.3127 0.08082 0.07346 -0.1042 0.0093 1.0000 13.000 1.3130 0.08753 0.08038 -0.1039 0.0101 1.0000 13.500 1.2993 0.09629 0.08936 -0.1040 0.0108 1.0000 14.000 1.2789 0.10620 0.09947 -0.1048 0.0112 1.0000 14.500 1.2820 0.11294 0.10639 -0.1054 0.0121 1.0000 15.000 1.2724 0.12158 0.11521 -0.1066 0.0131 1.0000 15.500 1.2677 0.12933 0.12303 -0.1079 0.0140 1.0000