XFOIL Version 6.96 Calculated polar for: EPPLER 904 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.250 -0.4723 0.09151 0.08929 -0.0367 1.0000 0.0062 -9.000 -0.4757 0.08799 0.08580 -0.0376 1.0000 0.0065 -8.750 -0.4814 0.08373 0.08159 -0.0392 0.9997 0.0062 -8.500 -0.4757 0.07741 0.07527 -0.0466 0.9965 0.0060 -8.250 -0.4714 0.07139 0.06924 -0.0552 0.9904 0.0065 -8.000 -0.4700 0.06540 0.06318 -0.0612 0.9809 0.0064 -7.750 -0.4673 0.06019 0.05788 -0.0641 0.9676 0.0065 -7.500 -0.4590 0.05481 0.05234 -0.0671 0.9576 0.0065 -7.250 -0.4460 0.04957 0.04691 -0.0696 0.9499 0.0068 -7.000 -0.4303 0.04440 0.04149 -0.0715 0.9435 0.0068 -5.500 -0.3318 0.02167 0.01657 -0.0636 0.9039 0.0033 -5.250 -0.3134 0.01912 0.01364 -0.0617 0.8993 0.0031 -5.000 -0.2914 0.01699 0.01116 -0.0603 0.8955 0.0029 -4.750 -0.2672 0.01513 0.00900 -0.0592 0.8924 0.0027 -4.500 -0.2429 0.01370 0.00735 -0.0583 0.8895 0.0026 -4.250 -0.2200 0.01252 0.00602 -0.0571 0.8864 0.0025 -4.000 -0.1974 0.01163 0.00500 -0.0560 0.8835 0.0025 -3.750 -0.1745 0.01088 0.00412 -0.0549 0.8810 0.0024 -3.500 -0.1503 0.01033 0.00342 -0.0542 0.8788 0.0025 -3.250 -0.1253 0.00997 0.00292 -0.0536 0.8767 0.0026 -3.000 -0.1000 0.00972 0.00255 -0.0531 0.8742 0.0028 -2.750 -0.0741 0.00955 0.00225 -0.0528 0.8718 0.0030 -2.500 -0.0486 0.00932 0.00204 -0.0524 0.8697 0.0199 -2.250 -0.0268 0.00870 0.00186 -0.0517 0.8676 0.1295 -1.750 -0.0088 0.00613 0.00162 -0.0451 0.8617 0.7344 -1.500 0.0147 0.00608 0.00178 -0.0439 0.8593 0.8086 -1.250 0.0410 0.00611 0.00180 -0.0436 0.8571 0.8321 -1.000 0.0673 0.00615 0.00179 -0.0433 0.8551 0.8492 -0.750 0.0935 0.00622 0.00186 -0.0428 0.8530 0.8674 -0.500 0.1187 0.00631 0.00195 -0.0421 0.8504 0.8831 -0.250 0.1434 0.00637 0.00203 -0.0414 0.8468 0.8942 0.000 0.1683 0.00642 0.00207 -0.0407 0.8429 0.9036 0.250 0.1954 0.00649 0.00212 -0.0404 0.8389 0.9106 0.500 0.2192 0.00650 0.00215 -0.0395 0.8329 0.9168 0.750 0.2467 0.00652 0.00215 -0.0393 0.8269 0.9203 1.000 0.2721 0.00651 0.00216 -0.0389 0.8193 0.9235 1.250 0.2987 0.00650 0.00213 -0.0386 0.8126 0.9263 1.500 0.3234 0.00649 0.00215 -0.0381 0.8061 0.9290 1.750 0.3511 0.00649 0.00218 -0.0382 0.8007 0.9305 2.000 0.3779 0.00650 0.00232 -0.0381 0.7944 0.9322 2.250 0.4049 0.00651 0.00238 -0.0380 0.7879 0.9339 2.500 0.4311 0.00651 0.00246 -0.0377 0.7799 0.9358 2.750 0.4561 0.00650 0.00251 -0.0372 0.7660 0.9380 3.000 0.4733 0.00654 0.00235 -0.0346 0.6880 0.9411 3.250 0.4712 0.00763 0.00255 -0.0283 0.4960 0.9457 3.500 0.4715 0.00898 0.00297 -0.0233 0.2836 0.9509 3.750 0.4752 0.01058 0.00352 -0.0194 0.0483 0.9563 4.000 0.4967 0.01135 0.00422 -0.0181 0.0060 0.9590 4.250 0.5210 0.01181 0.00485 -0.0175 0.0049 0.9615 4.500 0.5439 0.01231 0.00538 -0.0169 0.0024 0.9646 4.750 0.5651 0.01293 0.00609 -0.0158 0.0022 0.9684 5.000 0.5905 0.01363 0.00689 -0.0156 0.0018 0.9704 5.250 0.6140 0.01472 0.00811 -0.0149 0.0016 0.9731 5.500 0.6378 0.01606 0.00960 -0.0142 0.0015 0.9759 5.750 0.6636 0.01784 0.01156 -0.0136 0.0015 0.9784 6.000 0.6906 0.02001 0.01398 -0.0132 0.0015 0.9806 6.250 0.7182 0.02252 0.01680 -0.0129 0.0015 0.9823 6.500 0.7435 0.02576 0.02043 -0.0120 0.0016 0.9841 6.750 0.7653 0.02907 0.02413 -0.0109 0.0017 0.9871 7.000 0.7827 0.03277 0.02822 -0.0092 0.0017 0.9914 7.250 0.7962 0.03700 0.03285 -0.0074 0.0019 0.9963 7.500 0.8020 0.04065 0.03686 -0.0044 0.0020 1.0000 7.750 0.7976 0.04353 0.03995 0.0005 0.0020 1.0000 8.000 0.7889 0.04707 0.04372 0.0051 0.0021 1.0000 8.250 0.7751 0.05079 0.04761 0.0096 0.0022 1.0000 8.750 0.7706 0.05484 0.05192 0.0161 0.0024 1.0000 9.000 0.7681 0.05678 0.05397 0.0192 0.0025 1.0000 9.250 0.7487 0.06010 0.05742 0.0227 0.0024 1.0000 9.500 0.7515 0.06283 0.06027 0.0244 0.0027 1.0000 9.750 0.7270 0.06646 0.06400 0.0256 0.0025 1.0000 10.000 0.7095 0.07087 0.06851 0.0252 0.0025 1.0000 10.250 0.6997 0.07557 0.07330 0.0236 0.0025 1.0000 10.500 0.6771 0.08340 0.08120 0.0183 0.0024 1.0000