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 = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.500 -0.4653 0.09282 0.08665 0.0002 1.0000 0.2481 -8.250 -0.4568 0.08797 0.08180 0.0003 1.0000 0.2508 -8.000 -0.5541 0.09491 0.08838 0.0060 1.0000 0.2535 -7.750 -0.5681 0.09210 0.08572 0.0046 1.0000 0.2621 -7.250 -0.6126 0.06946 0.06289 -0.0228 1.0000 0.1110 -7.000 -0.6140 0.06250 0.05570 -0.0248 1.0000 0.0999 -6.750 -0.6053 0.05772 0.05076 -0.0246 1.0000 0.0957 -6.500 -0.5987 0.05274 0.04552 -0.0245 1.0000 0.0927 -6.250 -0.5962 0.04658 0.03834 -0.0241 1.0000 0.0855 -6.000 -0.5829 0.04243 0.03373 -0.0227 1.0000 0.0854 -5.750 -0.5667 0.03855 0.02940 -0.0212 1.0000 0.0850 -5.500 -0.5482 0.03491 0.02520 -0.0195 1.0000 0.0852 -5.250 -0.5271 0.03151 0.02143 -0.0180 1.0000 0.0874 -5.000 -0.5046 0.02912 0.01877 -0.0165 1.0000 0.0965 -4.750 -0.4823 0.02692 0.01645 -0.0150 1.0000 0.1143 -4.500 -0.4593 0.02467 0.01428 -0.0132 1.0000 0.1484 -4.250 -0.4465 0.02237 0.01269 -0.0097 1.0000 0.2740 -4.000 -0.4365 0.02177 0.01250 -0.0059 1.0000 0.3710 -3.750 -0.4177 0.02086 0.01165 -0.0035 1.0000 0.4293 -3.500 -0.3941 0.01977 0.01052 -0.0021 1.0000 0.4800 -3.250 -0.3698 0.01861 0.00951 -0.0006 1.0000 0.5398 -3.000 -0.3484 0.01734 0.00864 0.0018 1.0000 0.6313 -2.750 -0.1548 0.01646 0.00740 -0.0255 1.0000 1.0000 -2.500 -0.1463 0.01594 0.00660 -0.0238 1.0000 1.0000 -2.250 -0.1314 0.01566 0.00605 -0.0223 1.0000 1.0000 -2.000 -0.1144 0.01549 0.00563 -0.0209 1.0000 1.0000 -1.750 -0.0966 0.01537 0.00524 -0.0194 1.0000 1.0000 -1.500 -0.0783 0.01529 0.00499 -0.0180 1.0000 1.0000 -1.250 -0.0597 0.01524 0.00480 -0.0166 1.0000 1.0000 -1.000 -0.0411 0.01522 0.00467 -0.0152 1.0000 1.0000 -0.750 -0.0228 0.01524 0.00459 -0.0137 1.0000 1.0000 -0.500 -0.0048 0.01529 0.00456 -0.0122 1.0000 1.0000 -0.250 0.0124 0.01538 0.00460 -0.0107 1.0000 1.0000 0.000 0.0286 0.01552 0.00472 -0.0091 1.0000 1.0000 0.250 0.0431 0.01573 0.00492 -0.0074 1.0000 1.0000 0.500 0.0549 0.01605 0.00526 -0.0054 1.0000 1.0000 0.750 0.0634 0.01653 0.00574 -0.0034 1.0000 1.0000 1.000 0.0696 0.01719 0.00639 -0.0015 1.0000 1.0000 1.250 0.0756 0.01798 0.00718 -0.0001 1.0000 1.0000 1.500 0.1217 0.01898 0.00826 -0.0065 0.9844 1.0000 1.750 0.1940 0.01971 0.00917 -0.0168 0.9556 1.0000 2.000 0.2624 0.02019 0.00993 -0.0258 0.9294 1.0000 2.250 0.3308 0.02045 0.01050 -0.0342 0.9054 1.0000 2.500 0.3872 0.02069 0.01103 -0.0399 0.8816 1.0000 2.750 0.4372 0.02096 0.01168 -0.0440 0.8597 1.0000 3.000 0.4794 0.02132 0.01234 -0.0465 0.8381 1.0000 3.250 0.5128 0.02186 0.01316 -0.0473 0.8167 1.0000 3.500 0.5463 0.02233 0.01396 -0.0477 0.7961 1.0000 3.750 0.5714 0.02299 0.01502 -0.0468 0.7742 1.0000 4.000 0.5985 0.02344 0.01586 -0.0453 0.7508 1.0000 4.250 0.6102 0.02113 0.01336 -0.0324 0.6566 1.0000 4.500 0.6064 0.01977 0.01157 -0.0204 0.5328 1.0000 4.750 0.5936 0.02233 0.01161 -0.0113 0.1628 1.0000 5.000 0.6077 0.02522 0.01400 -0.0087 0.1167 1.0000 5.250 0.6274 0.02723 0.01601 -0.0066 0.0970 1.0000 5.500 0.6532 0.02971 0.01836 -0.0051 0.0892 1.0000 5.750 0.6812 0.03190 0.02085 -0.0037 0.0830 1.0000 6.000 0.7070 0.03515 0.02401 -0.0030 0.0767 1.0000 6.250 0.7309 0.03795 0.02739 -0.0014 0.0759 1.0000 6.500 0.7535 0.04160 0.03135 -0.0001 0.0768 1.0000 6.750 0.7708 0.04417 0.03485 0.0025 0.0804 1.0000 7.000 0.7842 0.04819 0.03950 0.0045 0.0850 1.0000 7.250 0.7991 0.05289 0.04440 0.0058 0.0890 1.0000 7.500 0.8000 0.05699 0.04942 0.0078 0.0973 1.0000 7.750 0.8056 0.06182 0.05464 0.0088 0.1058 1.0000 8.000 0.8059 0.06763 0.06079 0.0090 0.1176 1.0000 8.250 0.7926 0.07363 0.06709 0.0080 0.1277 1.0000 8.500 0.7627 0.07952 0.07319 0.0049 0.1343 1.0000 8.750 0.7398 0.08652 0.08018 0.0006 0.1462 1.0000 9.000 0.7180 0.09553 0.08908 -0.0076 0.1666 1.0000 9.250 0.6225 0.08706 0.08093 0.0011 0.1319 1.0000