XFOIL Version 6.96 Calculated polar for: EPPLER STF 863-615 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 ------ -------- --------- --------- -------- -------- -------- -8.750 -0.5760 0.09700 0.09409 -0.0394 0.9988 0.0295 -8.500 -0.5929 0.09212 0.08919 -0.0424 0.9962 0.0293 -8.250 -0.6129 0.08775 0.08477 -0.0445 0.9915 0.0290 -8.000 -0.6238 0.08252 0.07940 -0.0486 0.9885 0.0292 -7.750 -0.6333 0.07459 0.07084 -0.0549 0.9827 0.0302 -7.500 -0.6224 0.07170 0.06809 -0.0551 0.9813 0.0308 -7.250 -0.6156 0.06839 0.06468 -0.0554 0.9774 0.0319 -7.000 -0.5989 0.06322 0.05909 -0.0590 0.9747 0.0345 -6.750 -0.5768 0.05953 0.05502 -0.0621 0.9734 0.0382 -6.500 -0.5670 0.05691 0.05238 -0.0613 0.9697 0.0401 -6.250 -0.5399 0.05334 0.04831 -0.0645 0.9670 0.0470 -6.000 -0.5128 0.05086 0.04569 -0.0670 0.9655 0.0531 -5.750 -0.4813 0.04858 0.04310 -0.0704 0.9645 0.0637 -5.500 -0.4646 0.04641 0.04066 -0.0704 0.9610 0.0757 -5.250 -0.4366 0.04446 0.03865 -0.0727 0.9580 0.0923 -4.250 -0.2922 0.03519 0.02735 -0.0744 0.9508 0.0452 -4.000 -0.2602 0.03399 0.02597 -0.0745 0.9478 0.0400 -3.750 -0.2251 0.03346 0.02523 -0.0755 0.9459 0.0379 -3.500 -0.1885 0.03312 0.02500 -0.0778 0.9446 0.0408 -3.250 -0.1463 0.03322 0.02510 -0.0811 0.9436 0.0416 -3.000 -0.1290 0.03223 0.02412 -0.0800 0.9359 0.0421 -2.750 -0.0811 0.03226 0.02414 -0.0850 0.9333 0.0445 -2.500 -0.0530 0.03203 0.02385 -0.0859 0.9252 0.0496 -2.250 0.0416 0.03026 0.02430 -0.1003 0.9094 0.6221 -2.000 0.0613 0.03129 0.02534 -0.0972 0.8993 0.6678 -1.750 0.0852 0.03227 0.02628 -0.0954 0.8945 0.6929 -1.500 0.1034 0.03286 0.02691 -0.0924 0.8874 0.7056 -1.250 0.1197 0.03327 0.02731 -0.0897 0.8814 0.7172 -1.000 0.1479 0.03387 0.02788 -0.0890 0.8782 0.7327 -0.750 0.1601 0.03426 0.02828 -0.0855 0.8704 0.7459 -0.500 0.1802 0.03438 0.02839 -0.0836 0.8650 0.7537 -0.250 0.2095 0.03469 0.02869 -0.0829 0.8620 0.7693 0.000 0.2196 0.03487 0.02887 -0.0793 0.8527 0.7820 0.250 0.2473 0.03477 0.02876 -0.0792 0.8482 0.7878 0.500 0.2890 0.03474 0.02868 -0.0823 0.8457 0.7886 0.750 0.3312 0.03479 0.02870 -0.0854 0.8442 0.7900 1.000 0.3408 0.03479 0.02870 -0.0832 0.8329 0.7912 1.250 0.3809 0.03470 0.02859 -0.0861 0.8302 0.7918 1.500 0.4225 0.03463 0.02852 -0.0892 0.8285 0.7925 1.750 0.4639 0.03464 0.02854 -0.0922 0.8274 0.7938 2.000 0.4712 0.03480 0.02873 -0.0898 0.8153 0.7947 2.250 0.5101 0.03473 0.02871 -0.0924 0.8134 0.7952 2.500 0.5509 0.03461 0.02862 -0.0954 0.8120 0.7957 2.750 0.5623 0.03491 0.02896 -0.0936 0.8008 0.7964 3.000 0.6002 0.03474 0.02885 -0.0960 0.7984 0.7975 3.250 0.6421 0.03438 0.02859 -0.0989 0.7967 0.7984 3.500 0.6578 0.03457 0.02885 -0.0977 0.7860 0.7989 3.750 0.6972 0.03414 0.02850 -0.1001 0.7832 0.7994 4.000 0.7395 0.03353 0.02800 -0.1028 0.7813 0.7999 4.250 0.7585 0.03350 0.02809 -0.1019 0.7705 0.8003 4.500 0.8021 0.03249 0.02722 -0.1045 0.7675 0.8009 4.750 0.8638 0.02969 0.02459 -0.1087 0.7647 0.8018 5.000 0.9002 0.02722 0.02228 -0.1085 0.7503 0.8028 5.500 1.0168 0.01874 0.01354 -0.1123 0.6113 0.8043 6.000 0.9844 0.02272 0.01553 -0.0978 0.3027 0.8053 6.250 0.9632 0.02567 0.01737 -0.0911 0.1248 0.8058 6.500 0.9602 0.02782 0.01894 -0.0870 0.0507 0.8063 6.750 0.9722 0.02896 0.02009 -0.0851 0.0428 0.8070 7.000 0.9843 0.03009 0.02128 -0.0832 0.0388 0.8078 7.250 0.9958 0.03124 0.02250 -0.0814 0.0354 0.8086 7.500 1.0062 0.03252 0.02378 -0.0794 0.0331 0.8093 7.750 1.0184 0.03373 0.02506 -0.0776 0.0315 0.8103 8.000 1.0309 0.03509 0.02636 -0.0759 0.0302 0.8115 8.250 1.0500 0.03622 0.02759 -0.0747 0.0292 0.8128 8.500 1.0756 0.03730 0.02871 -0.0744 0.0279 0.8140 8.750 1.1040 0.03844 0.02990 -0.0747 0.0256 0.8151 9.000 1.1454 0.03983 0.03143 -0.0763 0.0243 0.8162 9.250 1.1004 0.03486 0.02705 -0.0639 0.0244 0.8173 9.500 1.2139 0.04185 0.03447 -0.0751 0.0264 0.8186 13.500 0.9610 0.13812 0.13553 -0.0592 0.0554 0.8312 13.750 0.9442 0.15014 0.14752 -0.0677 0.0553 0.8319 14.000 0.9368 0.15817 0.15552 -0.0731 0.0537 0.8330