XFOIL Version 6.96 Calculated polar for: GOE 228 (MVA H.38) AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -17.250 -0.5829 0.09281 0.09026 -0.0960 0.9630 0.0175 -17.000 -0.6047 0.08585 0.08317 -0.1001 0.9557 0.0177 -16.750 -0.6237 0.07945 0.07666 -0.1039 0.9478 0.0179 -16.500 -0.6405 0.07360 0.07068 -0.1074 0.9393 0.0180 -16.250 -0.6545 0.06811 0.06508 -0.1109 0.9300 0.0181 -16.000 -0.6663 0.06311 0.05995 -0.1141 0.9209 0.0182 -15.750 -0.6769 0.05831 0.05504 -0.1173 0.9099 0.0183 -15.500 -0.6894 0.05337 0.04997 -0.1209 0.8979 0.0185 -15.250 -0.7002 0.04882 0.04530 -0.1244 0.8855 0.0187 -15.000 -0.7078 0.04466 0.04101 -0.1277 0.8737 0.0188 -14.750 -0.7123 0.04074 0.03698 -0.1311 0.8629 0.0190 -14.500 -0.7136 0.03707 0.03320 -0.1346 0.8535 0.0192 -14.250 -0.7131 0.03364 0.02966 -0.1380 0.8447 0.0194 -14.000 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0.3247 1.000 1.0114 0.01052 0.00367 -0.1919 0.3910 0.3403 1.250 1.0379 0.01067 0.00378 -0.1918 0.3850 0.3471 1.500 1.0648 0.01078 0.00389 -0.1917 0.3795 0.3541 1.750 1.0911 0.01092 0.00402 -0.1915 0.3744 0.3610 2.000 1.1169 0.01109 0.00416 -0.1913 0.3698 0.3666 2.250 1.1436 0.01119 0.00427 -0.1912 0.3665 0.3724 2.500 1.1699 0.01131 0.00440 -0.1910 0.3628 0.3805 3.000 1.2205 0.01163 0.00472 -0.1904 0.3546 0.3973 3.250 1.2458 0.01177 0.00487 -0.1900 0.3516 0.4063 3.500 1.2713 0.01190 0.00503 -0.1897 0.3485 0.4175 3.750 1.2959 0.01204 0.00520 -0.1893 0.3450 0.4302 4.000 1.3184 0.01221 0.00539 -0.1885 0.3414 0.4466 4.250 1.3394 0.01239 0.00561 -0.1874 0.3379 0.4681 4.500 1.3626 0.01255 0.00584 -0.1867 0.3351 0.4980 4.750 1.3865 0.01272 0.00608 -0.1862 0.3325 0.5251 5.000 1.4098 0.01293 0.00633 -0.1856 0.3295 0.5435 5.250 1.4325 0.01317 0.00658 -0.1849 0.3262 0.5564 5.500 1.4545 0.01345 0.00688 -0.1841 0.3228 0.5675 5.750 1.4761 0.01375 0.00718 -0.1832 0.3195 0.5760 6.000 1.4989 0.01398 0.00746 -0.1826 0.3169 0.5848 6.250 1.5209 0.01426 0.00775 -0.1818 0.3137 0.5928 6.500 1.5421 0.01458 0.00810 -0.1810 0.3099 0.6013 6.750 1.5622 0.01496 0.00847 -0.1800 0.3061 0.6085 7.000 1.5826 0.01532 0.00886 -0.1790 0.3030 0.6150 7.250 1.6042 0.01563 0.00921 -0.1783 0.3004 0.6218 7.500 1.6247 0.01600 0.00960 -0.1774 0.2968 0.6283 7.750 1.6436 0.01645 0.01007 -0.1764 0.2921 0.6364 8.000 1.6621 0.01693 0.01056 -0.1753 0.2883 0.6442 8.250 1.6822 0.01734 0.01102 -0.1744 0.2854 0.6530 8.500 1.7007 0.01783 0.01154 -0.1734 0.2799 0.6634 8.750 1.7158 0.01852 0.01223 -0.1719 0.2720 0.6767 9.000 1.7306 0.01925 0.01296 -0.1704 0.2613 0.6922 9.250 1.7394 0.02036 0.01402 -0.1682 0.2442 0.7127 9.500 1.7269 0.02286 0.01631 -0.1636 0.1955 0.7469 9.750 1.7179 0.02472 0.01837 -0.1592 0.1705 1.0000 10.000 1.7248 0.02614 0.01978 -0.1572 0.1623 1.0000 10.250 1.7362 0.02728 0.02093 -0.1557 0.1587 1.0000 10.500 1.7462 0.02854 0.02220 -0.1542 0.1552 1.0000 10.750 1.7549 0.02991 0.02358 -0.1526 0.1510 1.0000 11.000 1.7666 0.03108 0.02479 -0.1514 0.1485 1.0000 11.250 1.7789 0.03224 0.02600 -0.1502 0.1471 1.0000 11.500 1.7902 0.03350 0.02729 -0.1491 0.1451 1.0000 11.750 1.8001 0.03490 0.02873 -0.1478 0.1431 1.0000 12.000 1.8084 0.03644 0.03031 -0.1465 0.1405 1.0000 12.250 1.8144 0.03823 0.03213 -0.1451 0.1378 1.0000 12.500 1.8212 0.03998 0.03392 -0.1438 0.1351 1.0000 12.750 1.8298 0.04161 0.03561 -0.1427 0.1334 1.0000 13.000 1.8388 0.04324 0.03728 -0.1417 0.1311 1.0000 13.250 1.8444 0.04519 0.03928 -0.1406 0.1275 1.0000 13.500 1.8458 0.04762 0.04172 -0.1393 0.1234 1.0000 13.750 1.8474 0.05008 0.04422 -0.1381 0.1192 1.0000 14.000 1.8462 0.05288 0.04704 -0.1369 0.1116 1.0000 14.250 1.8250 0.05793 0.05201 -0.1350 0.0931 1.0000 14.500 1.7918 0.06461 0.05864 -0.1331 0.0744 1.0000 14.750 1.7592 0.07152 0.06556 -0.1316 0.0598 1.0000 15.000 1.7284 0.07838 0.07245 -0.1306 0.0465 1.0000 15.250 1.6991 0.08530 0.07939 -0.1300 0.0351 1.0000