XFOIL Version 6.94 Calculated polar for: EPPLER 580 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 0.6388 0.01233 0.00592 -0.1496 0.6918 0.6303 0.500 0.6920 0.01237 0.00595 -0.1492 0.6767 0.6354 1.000 0.7480 0.01249 0.00596 -0.1495 0.6627 0.6405 1.500 0.8042 0.01260 0.00597 -0.1499 0.6485 0.6459 2.000 0.8583 0.01273 0.00605 -0.1499 0.6342 0.6513 2.500 0.9124 0.01283 0.00615 -0.1498 0.6207 0.6563 3.000 0.9656 0.01300 0.00630 -0.1495 0.6065 0.6614 3.500 1.0171 0.01317 0.00651 -0.1489 0.5919 0.6672 4.000 1.0704 0.01345 0.00675 -0.1487 0.5774 0.6734 4.500 1.1221 0.01364 0.00694 -0.1482 0.5623 0.6791 5.000 1.1685 0.01383 0.00727 -0.1465 0.5463 0.6852 5.500 1.2158 0.01411 0.00761 -0.1452 0.5294 0.6921 6.000 1.2613 0.01439 0.00795 -0.1434 0.5117 0.6994 6.500 1.3035 0.01471 0.00836 -0.1410 0.4931 0.7072 7.000 1.3429 0.01510 0.00880 -0.1381 0.4731 0.7157 7.500 1.3727 0.01546 0.00925 -0.1334 0.4500 0.7243 8.000 1.3986 0.01604 0.00987 -0.1281 0.4242 0.7348 8.500 1.4209 0.01684 0.01067 -0.1224 0.3947 0.7467 9.000 1.4408 0.01787 0.01171 -0.1167 0.3609 0.7618 9.500 1.4598 0.01908 0.01298 -0.1111 0.3296 0.7789 10.000 1.4752 0.02061 0.01455 -0.1055 0.2993 0.8020 10.500 1.4836 0.02253 0.01654 -0.0994 0.2631 0.8402 11.000 1.4780 0.02501 0.01901 -0.0919 0.2232 1.0000 11.500 1.4831 0.02831 0.02218 -0.0876 0.1883 1.0000 12.000 1.4862 0.03206 0.02585 -0.0837 0.1577 1.0000 12.500 1.4859 0.03645 0.03016 -0.0803 0.1293 1.0000 13.500 1.4836 0.04668 0.04035 -0.0758 0.0846 1.0000 14.000 1.4800 0.05268 0.04634 -0.0745 0.0646 1.0000 14.500 1.4746 0.05931 0.05298 -0.0740 0.0479 1.0000 15.000 1.4668 0.06671 0.06042 -0.0741 0.0328 1.0000 15.500 1.4512 0.07565 0.06939 -0.0751 0.0181 1.0000 16.000 1.4342 0.08544 0.07930 -0.0770 0.0111 1.0000 16.500 1.4234 0.09480 0.08889 -0.0794 0.0092 1.0000 17.000 1.4133 0.10440 0.09872 -0.0825 0.0082 1.0000 17.500 1.4018 0.11453 0.10911 -0.0863 0.0076 1.0000 18.000 1.3918 0.12457 0.11941 -0.0906 0.0073 1.0000 18.500 1.3816 0.13474 0.12983 -0.0954 0.0070 1.0000 19.000 1.3715 0.14496 0.14026 -0.1007 0.0068 1.0000 19.500 1.3630 0.15484 0.15034 -0.1061 0.0066 1.0000 20.000 1.3565 0.16431 0.15998 -0.1116 0.0064 1.0000 20.500 1.3524 0.17317 0.16899 -0.1171 0.0063 1.0000 21.000 1.3512 0.18129 0.17722 -0.1222 0.0062 1.0000 21.500 1.3542 0.18824 0.18425 -0.1267 0.0061 1.0000 22.000 1.3579 0.19517 0.19134 -0.1313 0.0060 1.0000 22.500 1.3610 0.20221 0.19855 -0.1362 0.0060 1.0000 23.000 1.3622 0.20964 0.20616 -0.1415 0.0060 1.0000 23.500 1.3599 0.21800 0.21474 -0.1476 0.0060 1.0000 24.000 1.3546 0.22719 0.22414 -0.1545 0.0060 1.0000 24.500 1.3457 0.23753 0.23471 -0.1623 0.0060 1.0000