XFOIL Version 6.96 Calculated polar for: EPPLER 874 HYDROFOIL 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 = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.750 -0.5864 0.09553 0.09396 -0.0030 1.0000 0.0076 -9.500 -0.5867 0.09116 0.08960 -0.0049 1.0000 0.0077 -9.250 -0.5922 0.08584 0.08430 -0.0074 1.0000 0.0075 -5.500 -0.5837 0.01397 0.00910 -0.0046 1.0000 0.0053 -5.250 -0.5624 0.01328 0.00836 -0.0033 1.0000 0.0060 -5.000 -0.5412 0.01249 0.00748 -0.0019 1.0000 0.0064 -4.750 -0.5208 0.01154 0.00642 -0.0002 1.0000 0.0065 -4.500 -0.4946 0.01100 0.00581 0.0001 0.9995 0.0071 -4.250 -0.4604 0.01027 0.00499 -0.0013 0.9975 0.0074 -4.000 -0.4256 0.00942 0.00399 -0.0027 0.9953 0.0072 -3.750 -0.3917 0.00882 0.00325 -0.0039 0.9929 0.0071 -3.500 -0.3579 0.00839 0.00272 -0.0051 0.9896 0.0073 -3.250 -0.3222 0.00809 0.00233 -0.0067 0.9863 0.0079 -3.000 -0.2857 0.00769 0.00198 -0.0086 0.9832 0.0281 -2.750 -0.2610 0.00688 0.00162 -0.0082 0.9736 0.1374 -2.500 -0.2334 0.00649 0.00143 -0.0082 0.9613 0.1891 -2.250 -0.1985 0.00617 0.00126 -0.0098 0.9486 0.2322 -2.000 -0.1476 0.00583 0.00109 -0.0151 0.9345 0.2867 -1.750 -0.0988 0.00535 0.00091 -0.0201 0.8950 0.3998 -1.500 -0.0723 0.00531 0.00082 -0.0197 0.8403 0.4511 -1.250 -0.0498 0.00529 0.00077 -0.0185 0.7941 0.5008 -1.000 -0.0274 0.00523 0.00074 -0.0173 0.7537 0.5562 -0.750 -0.0049 0.00514 0.00071 -0.0161 0.7190 0.6174 -0.500 0.0170 0.00499 0.00069 -0.0148 0.6888 0.6863 -0.250 0.0380 0.00481 0.00068 -0.0132 0.6621 0.7656 0.000 0.0592 0.00457 0.00069 -0.0116 0.6381 0.8596 0.250 0.1372 0.00464 0.00085 -0.0229 0.6071 0.9433 0.500 0.1767 0.00486 0.00098 -0.0253 0.5831 0.9615 0.750 0.2065 0.00503 0.00107 -0.0257 0.5633 0.9757 1.000 0.2380 0.00521 0.00118 -0.0265 0.5448 0.9843 1.250 0.2810 0.00534 0.00125 -0.0299 0.5258 0.9892 1.750 0.3610 0.00557 0.00137 -0.0354 0.4877 0.9969 2.000 0.4001 0.00571 0.00136 -0.0381 0.4497 0.9997 2.250 0.4253 0.00590 0.00140 -0.0376 0.4062 1.0000 2.500 0.4491 0.00607 0.00146 -0.0368 0.3720 1.0000 2.750 0.4726 0.00630 0.00153 -0.0360 0.3276 1.0000 3.000 0.4880 0.00761 0.00191 -0.0340 0.1011 1.0000 3.250 0.5081 0.00840 0.00233 -0.0326 0.0094 1.0000 3.500 0.5321 0.00863 0.00260 -0.0317 0.0076 1.0000 3.750 0.5558 0.00890 0.00293 -0.0307 0.0067 1.0000 4.000 0.5767 0.00964 0.00383 -0.0291 0.0051 1.0000 4.250 0.5989 0.01014 0.00441 -0.0278 0.0049 1.0000 4.500 0.6193 0.01088 0.00526 -0.0261 0.0048 1.0000 4.750 0.6377 0.01192 0.00642 -0.0240 0.0048 1.0000 5.000 0.6525 0.01388 0.00849 -0.0211 0.0053 1.0000 5.250 0.6765 0.01413 0.00878 -0.0202 0.0056 1.0000 6.000 0.7238 0.02581 0.02143 -0.0116 0.0096 1.0000 6.250 0.7398 0.02782 0.02369 -0.0094 0.0095 1.0000 6.500 0.7575 0.02892 0.02500 -0.0074 0.0093 1.0000 6.750 0.7818 0.02872 0.02489 -0.0062 0.0084 1.0000 7.000 0.7984 0.03046 0.02683 -0.0041 0.0076 1.0000 7.250 0.8113 0.03279 0.02937 -0.0018 0.0071 1.0000 7.500 0.8221 0.03523 0.03202 0.0005 0.0067 1.0000 7.750 0.8307 0.03774 0.03473 0.0028 0.0064 1.0000 8.000 0.8371 0.04035 0.03753 0.0051 0.0062 1.0000 8.250 0.8417 0.04294 0.04029 0.0073 0.0061 1.0000 8.500 0.8429 0.04583 0.04335 0.0096 0.0060 1.0000 8.750 0.8411 0.04854 0.04620 0.0117 0.0058 1.0000 9.000 0.8355 0.05141 0.04921 0.0139 0.0057 1.0000 9.250 0.8237 0.05441 0.05234 0.0166 0.0057 1.0000 9.500 0.8047 0.05730 0.05535 0.0194 0.0057 1.0000 9.750 0.7873 0.06113 0.05928 0.0186 0.0057 1.0000 10.000 0.7669 0.06714 0.06541 0.0141 0.0057 1.0000 10.250 0.7437 0.07795 0.07637 0.0051 0.0059 1.0000