XFOIL Version 6.94 Calculated polar for: EPPLER 397 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.7085 0.00936 0.00349 -0.1702 0.7112 0.5244 1.000 0.8212 0.00934 0.00372 -0.1706 0.6966 0.6283 1.500 0.8760 0.00939 0.00397 -0.1705 0.6889 0.6858 2.000 0.9295 0.00936 0.00412 -0.1699 0.6803 0.7478 2.500 0.9827 0.00942 0.00433 -0.1692 0.6717 0.8175 3.000 1.0219 0.00923 0.00444 -0.1653 0.6620 0.9130 3.500 1.0785 0.00928 0.00446 -0.1655 0.6518 1.0000 4.000 1.1305 0.00939 0.00463 -0.1649 0.6394 1.0000 4.500 1.1816 0.00950 0.00471 -0.1639 0.6226 1.0000 5.000 1.2317 0.00968 0.00488 -0.1628 0.6057 1.0000 5.500 1.2783 0.00989 0.00511 -0.1611 0.5852 1.0000 6.000 1.3236 0.01020 0.00543 -0.1591 0.5645 1.0000 6.500 1.3656 0.01057 0.00584 -0.1565 0.5404 1.0000 7.000 1.3999 0.01109 0.00634 -0.1524 0.5074 1.0000 7.500 1.4199 0.01186 0.00700 -0.1456 0.4664 1.0000 8.000 1.4306 0.01306 0.00801 -0.1375 0.4166 1.0000 8.500 1.4327 0.01480 0.00950 -0.1286 0.3569 1.0000 9.000 1.4297 0.01713 0.01155 -0.1199 0.2960 1.0000 9.500 1.4273 0.01991 0.01406 -0.1124 0.2397 1.0000 10.000 1.4284 0.02293 0.01686 -0.1061 0.1912 1.0000 10.500 1.4313 0.02622 0.01996 -0.1008 0.1471 1.0000 11.000 1.4318 0.03003 0.02353 -0.0960 0.1036 1.0000 11.500 1.4342 0.03402 0.02735 -0.0919 0.0691 1.0000 12.000 1.4410 0.03797 0.03124 -0.0887 0.0525 1.0000 12.500 1.4475 0.04221 0.03554 -0.0860 0.0444 1.0000 13.000 1.4581 0.04632 0.03977 -0.0839 0.0370 1.0000 13.500 1.4726 0.05023 0.04378 -0.0823 0.0279 1.0000 14.000 1.4826 0.05483 0.04839 -0.0809 0.0192 1.0000 15.000 1.4815 0.06750 0.06136 -0.0792 0.0116 1.0000 15.500 1.4784 0.07475 0.06881 -0.0793 0.0098 1.0000 16.000 1.4692 0.08328 0.07752 -0.0802 0.0087 1.0000 16.500 1.4686 0.09091 0.08541 -0.0815 0.0078 1.0000 17.000 1.4627 0.09966 0.09437 -0.0836 0.0071 1.0000 17.500 1.4486 0.10996 0.10486 -0.0867 0.0067 1.0000 18.000 1.4440 0.11892 0.11407 -0.0898 0.0064 1.0000 18.500 1.4388 0.12801 0.12341 -0.0934 0.0061 1.0000 19.000 1.4338 0.13709 0.13272 -0.0975 0.0058 1.0000 19.500 1.4291 0.14609 0.14194 -0.1019 0.0056 1.0000 20.000 1.4242 0.15511 0.15115 -0.1068 0.0054 1.0000 20.500 1.4190 0.16414 0.16037 -0.1120 0.0052 1.0000 21.000 1.4125 0.17332 0.16975 -0.1175 0.0051 1.0000 21.500 1.4029 0.18326 0.17993 -0.1239 0.0050 1.0000 22.000 1.3899 0.19438 0.19132 -0.1315 0.0049 1.0000 22.500 1.3743 0.20654 0.20375 -0.1401 0.0049 1.0000 23.000 1.3560 0.22003 0.21751 -0.1498 0.0050 1.0000 25.000 1.2487 0.31397 0.31188 -0.1998 0.0100 1.0000 25.500 1.2598 0.31946 0.31738 -0.2039 0.0092 1.0000 26.000 1.2720 0.32316 0.32110 -0.2071 0.0088 1.0000 26.500 1.2804 0.33190 0.32980 -0.2125 0.0081 1.0000 27.000 1.2910 0.33724 0.33515 -0.2166 0.0072 1.0000 27.500 1.3018 0.34118 0.33910 -0.2202 0.0069 1.0000 28.500 1.3212 0.35399 0.35190 -0.2291 0.0060 1.0000 29.000 1.3310 0.35897 0.35688 -0.2332 0.0055 1.0000 29.500 1.3403 0.36326 0.36118 -0.2371 0.0052 1.0000 30.000 1.3492 0.36703 0.36496 -0.2409 0.0050 1.0000 30.500 1.3589 0.37044 0.36840 -0.2438 0.0049 1.0000 31.000 1.3663 0.37640 0.37434 -0.2489 0.0049 1.0000 31.500 1.3754 0.38255 0.38049 -0.2534 0.0046 1.0000 32.000 1.3836 0.38700 0.38495 -0.2575 0.0043 1.0000 32.500 1.3912 0.39110 0.38907 -0.2615 0.0040 1.0000 33.000 1.3984 0.39499 0.39297 -0.2655 0.0038 1.0000 33.500 1.4051 0.39864 0.39663 -0.2695 0.0037 1.0000 34.000 1.4114 0.40209 0.40010 -0.2734 0.0036 1.0000 34.500 1.4169 0.40499 0.40301 -0.2773 0.0035 1.0000 35.000 1.4223 0.40789 0.40594 -0.2805 0.0034 1.0000 35.500 1.4276 0.41161 0.40967 -0.2848 0.0034 1.0000 36.000 1.4326 0.41514 0.41321 -0.2890 0.0034 1.0000 36.500 1.4375 0.41865 0.41673 -0.2933 0.0033 1.0000 37.000 1.4422 0.42184 0.41994 -0.2974 0.0033 1.0000 37.500 1.4465 0.42472 0.42283 -0.3015 0.0033 1.0000