XFOIL Version 6.94 Calculated polar for: EPPLER STF 863-615 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4741 0.02171 0.01701 -0.1261 0.8316 0.7063 0.500 0.5420 0.02139 0.01671 -0.1289 0.8302 0.7130 1.000 0.6094 0.02100 0.01635 -0.1318 0.8289 0.7148 1.500 0.6788 0.02050 0.01589 -0.1351 0.8274 0.7159 3.000 0.8461 0.01827 0.01400 -0.1362 0.8008 0.7204 4.000 1.0048 0.01191 0.00726 -0.1416 0.6637 0.7226 4.500 0.9682 0.01514 0.00937 -0.1250 0.4676 0.7238 5.000 0.9403 0.01909 0.01194 -0.1118 0.2184 0.7249 5.500 0.9405 0.02232 0.01403 -0.1039 0.0244 0.7259 6.000 0.9750 0.02354 0.01529 -0.1014 0.0174 0.7276 6.500 1.0087 0.02486 0.01669 -0.0988 0.0153 0.7287 7.000 1.0415 0.02624 0.01817 -0.0963 0.0137 0.7296 7.500 1.0687 0.02800 0.02002 -0.0929 0.0128 0.7306 8.000 1.1009 0.02944 0.02154 -0.0906 0.0109 0.7315 8.500 1.1331 0.03088 0.02312 -0.0883 0.0094 0.7324 9.000 1.1613 0.03258 0.02494 -0.0856 0.0085 0.7340 9.500 1.1850 0.03469 0.02722 -0.0823 0.0080 0.7356 10.000 1.2126 0.03654 0.02932 -0.0796 0.0071 0.7372 10.500 1.2436 0.03804 0.03097 -0.0778 0.0057 0.7390 11.000 1.2714 0.03991 0.03314 -0.0753 0.0042 0.7408 11.500 1.2908 0.04256 0.03605 -0.0721 0.0029 0.7427 12.000 1.3114 0.04512 0.03882 -0.0693 0.0023 0.7447 12.500 1.3218 0.04873 0.04271 -0.0656 0.0021 0.7471 13.000 1.3200 0.05383 0.04826 -0.0609 0.0019 0.7493 13.500 1.3072 0.06049 0.05547 -0.0561 0.0019 0.7509 14.000 1.2787 0.06949 0.06511 -0.0518 0.0019 0.7522 14.500 1.2361 0.08119 0.07742 -0.0496 0.0019 0.7532 15.000 1.1867 0.09538 0.09216 -0.0512 0.0019 0.7540 15.500 1.1415 0.11117 0.10837 -0.0571 0.0019 0.7552 16.000 1.1019 0.12868 0.12621 -0.0669 0.0020 0.7566 16.500 1.0617 0.15047 0.14828 -0.0804 0.0021 0.7575 17.500 0.6653 0.18245 0.18024 -0.0676 0.0049 0.7488 18.000 0.6654 0.18851 0.18631 -0.0698 0.0060 0.7504 18.500 0.6661 0.19470 0.19251 -0.0714 0.0081 0.7518 19.000 0.6669 0.20154 0.19935 -0.0738 0.0081 0.7532 20.000 0.6697 0.21518 0.21301 -0.0788 0.0081 0.7567 20.500 0.6702 0.22221 0.22006 -0.0817 0.0080 0.7589 21.000 0.6719 0.22918 0.22705 -0.0844 0.0077 0.7612 21.500 0.6757 0.23520 0.23310 -0.0866 0.0072 0.7640 22.000 0.6799 0.24096 0.23889 -0.0888 0.0068 0.7670 22.500 0.6843 0.24656 0.24451 -0.0910 0.0064 0.7703 23.000 0.6890 0.25195 0.24993 -0.0931 0.0061 0.7739 23.500 0.6942 0.25693 0.25495 -0.0949 0.0059 0.7781 24.000 0.6998 0.26253 0.26060 -0.0965 0.0058 0.7841 24.500 0.7026 0.26862 0.26673 -0.0990 0.0058 0.7905 25.500 0.7096 0.28082 0.27906 -0.1040 0.0056 0.8118 26.000 0.7136 0.28665 0.28501 -0.1061 0.0054 0.8455 26.500 0.7163 0.29149 0.28998 -0.1076 0.0051 0.9981 27.000 0.7204 0.29713 0.29563 -0.1099 0.0049 0.9981 27.500 0.7242 0.30269 0.30119 -0.1121 0.0046 0.9981 28.000 0.7278 0.30817 0.30668 -0.1144 0.0044 0.9981 28.500 0.7313 0.31362 0.31213 -0.1166 0.0043 0.9981 29.000 0.7345 0.31893 0.31745 -0.1188 0.0041 0.9981 29.500 0.7374 0.32411 0.32265 -0.1209 0.0040 0.9981 30.000 0.7401 0.32920 0.32775 -0.1231 0.0039 0.9981 30.500 0.7425 0.33414 0.33271 -0.1250 0.0038 0.9981