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 = 0.200 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -10.250 -0.4691 0.10438 0.10104 -0.0056 1.0000 0.0450 -10.000 -0.4768 0.09955 0.09624 -0.0088 1.0000 0.0470 -9.750 -0.5740 0.10450 0.10097 -0.0021 1.0000 0.0412 -9.500 -0.5730 0.10065 0.09714 -0.0038 1.0000 0.0421 -9.250 -0.5745 0.09634 0.09286 -0.0060 1.0000 0.0444 -9.000 -0.5787 0.09151 0.08810 -0.0095 1.0000 0.0461 -6.250 -0.6000 0.03460 0.02937 -0.0215 1.0000 0.0343 -6.000 -0.5883 0.03086 0.02519 -0.0192 1.0000 0.0337 -5.750 -0.5751 0.02636 0.02008 -0.0163 1.0000 0.0290 -5.500 -0.5512 0.02526 0.01843 -0.0139 1.0000 0.0250 -5.250 -0.5325 0.02192 0.01477 -0.0120 1.0000 0.0237 -5.000 -0.5115 0.01967 0.01224 -0.0103 1.0000 0.0231 -4.750 -0.4898 0.01790 0.01030 -0.0086 1.0000 0.0229 -4.500 -0.4686 0.01645 0.00873 -0.0070 1.0000 0.0230 -4.250 -0.4479 0.01529 0.00748 -0.0053 1.0000 0.0236 -4.000 -0.4269 0.01435 0.00642 -0.0036 1.0000 0.0247 -3.750 -0.4068 0.01315 0.00502 -0.0017 1.0000 0.0314 -3.500 -0.3898 0.01144 0.00416 0.0002 1.0000 0.1944 -3.250 -0.3684 0.01105 0.00410 0.0011 1.0000 0.2725 -3.000 -0.3468 0.01066 0.00386 0.0022 1.0000 0.3184 -2.750 -0.3256 0.01021 0.00360 0.0035 1.0000 0.3710 -2.500 -0.3056 0.00972 0.00341 0.0050 1.0000 0.4425 -2.250 -0.2872 0.00921 0.00325 0.0068 1.0000 0.5387 -2.000 -0.2708 0.00866 0.00320 0.0094 1.0000 0.6580 -1.750 -0.1716 0.00812 0.00335 -0.0044 1.0000 0.9263 -1.500 -0.0879 0.00834 0.00334 -0.0160 1.0000 0.9959 -1.250 -0.0577 0.00831 0.00322 -0.0169 1.0000 1.0000 -1.000 -0.0400 0.00831 0.00316 -0.0150 1.0000 1.0000 -0.750 -0.0230 0.00834 0.00315 -0.0132 1.0000 1.0000 -0.500 -0.0074 0.00842 0.00320 -0.0111 1.0000 1.0000 -0.250 0.0476 0.00839 0.00314 -0.0171 0.9874 1.0000 0.000 0.0964 0.00825 0.00299 -0.0217 0.9620 1.0000 0.250 0.1533 0.00811 0.00285 -0.0278 0.9352 1.0000 0.500 0.2226 0.00794 0.00263 -0.0365 0.8926 1.0000 0.750 0.2741 0.00794 0.00245 -0.0411 0.8290 1.0000 1.000 0.3008 0.00815 0.00243 -0.0405 0.7762 1.0000 1.250 0.3239 0.00838 0.00251 -0.0392 0.7362 1.0000 1.500 0.3467 0.00861 0.00259 -0.0380 0.7032 1.0000 1.750 0.3699 0.00884 0.00271 -0.0369 0.6760 1.0000 2.000 0.3935 0.00906 0.00286 -0.0360 0.6530 1.0000 2.250 0.4173 0.00929 0.00304 -0.0350 0.6328 1.0000 2.500 0.4414 0.00952 0.00328 -0.0342 0.6145 1.0000 2.750 0.4655 0.00973 0.00351 -0.0333 0.5967 1.0000 3.000 0.4893 0.00994 0.00374 -0.0324 0.5785 1.0000 3.250 0.5114 0.01011 0.00388 -0.0309 0.5498 1.0000 3.500 0.5290 0.01024 0.00379 -0.0285 0.4787 1.0000 3.750 0.5465 0.01068 0.00389 -0.0263 0.3704 1.0000 4.000 0.5499 0.01375 0.00528 -0.0223 0.0326 1.0000 4.250 0.5680 0.01496 0.00665 -0.0201 0.0272 1.0000 4.500 0.5871 0.01600 0.00781 -0.0182 0.0257 1.0000 4.750 0.6060 0.01726 0.00914 -0.0161 0.0245 1.0000 5.000 0.6265 0.01823 0.01012 -0.0146 0.0199 1.0000 5.250 0.6463 0.01990 0.01184 -0.0129 0.0189 1.0000 5.500 0.6684 0.02171 0.01379 -0.0113 0.0189 1.0000 5.750 0.6923 0.02383 0.01620 -0.0095 0.0205 1.0000 6.000 0.7144 0.02704 0.01991 -0.0073 0.0238 1.0000 12.500 0.6945 0.14496 0.14128 -0.0257 0.0321 1.0000 12.750 0.6982 0.14888 0.14518 -0.0270 0.0310 1.0000