XFOIL Version 6.96 Calculated polar for: EPPLER 818 HYDROFOIL AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -7.500 -0.5497 0.08413 0.07817 -0.0386 1.0000 0.1009 -7.250 -0.5703 0.07620 0.07026 -0.0449 1.0000 0.0876 -7.000 -0.5804 0.07036 0.06432 -0.0480 1.0000 0.0799 -6.750 -0.5979 0.06254 0.05568 -0.0559 1.0000 0.0713 -6.500 -0.5865 0.05827 0.05152 -0.0548 1.0000 0.0687 -6.250 -0.5768 0.05323 0.04615 -0.0564 1.0000 0.0660 -6.000 -0.5599 0.04762 0.03973 -0.0593 1.0000 0.0610 -5.750 -0.5368 0.04351 0.03476 -0.0610 1.0000 0.0596 -5.500 -0.5133 0.03992 0.03071 -0.0616 1.0000 0.0600 -5.250 -0.4901 0.03688 0.02747 -0.0618 1.0000 0.0633 -5.000 -0.4638 0.03429 0.02445 -0.0618 1.0000 0.0666 -4.750 -0.4373 0.03206 0.02181 -0.0611 1.0000 0.0688 -4.500 -0.4122 0.03023 0.01968 -0.0595 1.0000 0.0710 -4.250 -0.3890 0.02824 0.01754 -0.0582 1.0000 0.0757 -4.000 -0.3618 0.02645 0.01563 -0.0579 1.0000 0.0869 -3.750 -0.3233 0.02291 0.01449 -0.0613 1.0000 0.4407 -3.500 -0.3458 0.02717 0.01958 -0.0442 1.0000 0.5734 -3.250 -0.3378 0.02764 0.01984 -0.0379 1.0000 0.6049 -3.000 -0.3150 0.02705 0.01887 -0.0370 1.0000 0.6279 -2.750 -0.2921 0.02649 0.01799 -0.0363 1.0000 0.6517 -2.500 -0.2742 0.02605 0.01736 -0.0342 1.0000 0.6722 -2.250 -0.2539 0.02558 0.01669 -0.0328 1.0000 0.6954 -2.000 -0.2338 0.02514 0.01607 -0.0315 1.0000 0.7199 -1.750 -0.2148 0.02472 0.01550 -0.0300 1.0000 0.7452 -1.500 -0.1971 0.02432 0.01481 -0.0281 1.0000 0.7714 -1.250 -0.1815 0.02391 0.01431 -0.0257 1.0000 0.7986 -1.000 -0.1680 0.02345 0.01379 -0.0229 1.0000 0.8281 -0.750 -0.1547 0.02295 0.01324 -0.0203 1.0000 0.8625 -0.500 -0.1433 0.02231 0.01261 -0.0172 1.0000 0.9032 -0.250 -0.1101 0.02160 0.01189 -0.0191 1.0000 0.9665 0.000 -0.0704 0.02121 0.01133 -0.0248 1.0000 1.0000 0.250 -0.0270 0.02141 0.01121 -0.0309 1.0000 1.0000 0.500 0.0125 0.02172 0.01132 -0.0356 1.0000 1.0000 0.750 0.0476 0.02210 0.01152 -0.0389 1.0000 1.0000 1.000 0.0791 0.02251 0.01180 -0.0413 1.0000 1.0000 1.250 0.1082 0.02295 0.01215 -0.0431 1.0000 1.0000 1.500 0.1355 0.02342 0.01256 -0.0443 1.0000 1.0000 1.750 0.1614 0.02393 0.01302 -0.0453 1.0000 1.0000 2.000 0.1862 0.02447 0.01354 -0.0460 1.0000 1.0000 2.250 0.2103 0.02505 0.01413 -0.0466 1.0000 1.0000 2.500 0.2335 0.02566 0.01479 -0.0471 1.0000 1.0000 2.750 0.2562 0.02632 0.01551 -0.0475 1.0000 1.0000 3.000 0.2782 0.02703 0.01631 -0.0478 1.0000 1.0000 3.250 0.2995 0.02779 0.01737 -0.0481 1.0000 1.0000 3.500 0.3202 0.02861 0.01835 -0.0483 1.0000 1.0000 3.750 0.3403 0.02950 0.01941 -0.0485 1.0000 1.0000 4.000 0.3597 0.03046 0.02056 -0.0488 1.0000 1.0000 4.250 0.3783 0.03151 0.02181 -0.0490 1.0000 1.0000 4.500 0.3962 0.03265 0.02318 -0.0493 1.0000 1.0000 4.750 0.4132 0.03391 0.02469 -0.0496 1.0000 1.0000 5.000 0.4433 0.03573 0.02688 -0.0526 0.9924 1.0000 5.250 0.5307 0.03828 0.03053 -0.0648 0.9447 1.0000 5.500 0.8403 0.03581 0.02495 -0.0748 0.0693 1.0000 5.750 0.8790 0.03940 0.02899 -0.0750 0.0700 1.0000 6.000 0.9101 0.04325 0.03332 -0.0741 0.0726 1.0000 6.250 0.9371 0.04778 0.03813 -0.0733 0.0757 1.0000 6.500 0.9552 0.05101 0.04233 -0.0700 0.0835 1.0000 6.750 0.9738 0.05534 0.04723 -0.0677 0.0929 1.0000 7.000 0.9858 0.06019 0.05281 -0.0648 0.1085 1.0000 7.250 0.9949 0.06573 0.05910 -0.0624 0.1347 1.0000 7.500 0.9973 0.07326 0.06748 -0.0621 0.1910 1.0000 7.750 0.9178 0.07208 0.06701 -0.0532 0.1921 1.0000 8.000 0.8876 0.07784 0.07300 -0.0521 0.2045 1.0000 8.250 0.8553 0.08252 0.07773 -0.0508 0.2030 1.0000 8.500 0.8220 0.08883 0.08410 -0.0523 0.2104 1.0000 8.750 0.7890 0.09598 0.09123 -0.0558 0.2122 1.0000