XFOIL Version 6.94 Calculated polar for: EPPLER 66 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5431 0.00864 0.00407 -0.1353 0.8617 0.8604 0.500 0.5917 0.00832 0.00383 -0.1326 0.8451 0.9597 1.000 0.6597 0.00822 0.00363 -0.1351 0.8281 1.0000 1.500 0.7194 0.00824 0.00352 -0.1358 0.8083 1.0000 2.000 0.7764 0.00830 0.00348 -0.1358 0.7857 1.0000 2.500 0.8327 0.00841 0.00352 -0.1356 0.7605 1.0000 3.000 0.8857 0.00860 0.00364 -0.1348 0.7321 1.0000 3.500 0.9366 0.00884 0.00381 -0.1335 0.6988 1.0000 4.000 0.9850 0.00916 0.00405 -0.1317 0.6611 1.0000 4.500 1.0313 0.00958 0.00438 -0.1296 0.6184 1.0000 5.000 1.0742 0.01012 0.00477 -0.1268 0.5687 1.0000 5.500 1.1060 0.01100 0.00528 -0.1219 0.4839 1.0000 6.000 1.1316 0.01224 0.00600 -0.1162 0.3847 1.0000 6.500 1.1613 0.01342 0.00686 -0.1115 0.3086 1.0000 7.000 1.1848 0.01490 0.00789 -0.1059 0.2252 1.0000 7.500 1.2034 0.01689 0.00924 -0.1000 0.1280 1.0000 8.000 1.2323 0.01836 0.01052 -0.0958 0.0861 1.0000 8.500 1.2603 0.01988 0.01192 -0.0917 0.0562 1.0000 9.000 1.2869 0.02151 0.01352 -0.0874 0.0348 1.0000 9.500 1.3111 0.02333 0.01532 -0.0831 0.0201 1.0000 10.000 1.3357 0.02516 0.01727 -0.0791 0.0146 1.0000 10.500 1.3581 0.02720 0.01948 -0.0750 0.0107 1.0000 11.000 1.3767 0.02964 0.02210 -0.0709 0.0064 1.0000 11.500 1.3913 0.03255 0.02524 -0.0668 0.0039 1.0000 12.000 1.4012 0.03603 0.02897 -0.0628 0.0031 1.0000 12.500 1.4075 0.04007 0.03334 -0.0593 0.0028 1.0000 13.000 1.4075 0.04504 0.03865 -0.0563 0.0026 1.0000 13.500 1.4031 0.05093 0.04488 -0.0542 0.0025 1.0000 14.000 1.3938 0.05799 0.05230 -0.0533 0.0024 1.0000 14.500 1.3799 0.06644 0.06112 -0.0539 0.0024 1.0000 15.000 1.3618 0.07649 0.07154 -0.0562 0.0024 1.0000 15.500 1.3409 0.08803 0.08347 -0.0603 0.0024 1.0000 16.000 1.3170 0.10120 0.09702 -0.0662 0.0024 1.0000 16.500 1.2906 0.11592 0.11211 -0.0738 0.0024 1.0000 17.000 1.2634 0.13176 0.12831 -0.0829 0.0024 1.0000 17.500 1.2363 0.14845 0.14532 -0.0932 0.0025 1.0000 18.000 1.2095 0.16600 0.16316 -0.1042 0.0026 1.0000 18.500 1.1813 0.18507 0.18248 -0.1162 0.0027 1.0000 19.000 1.1469 0.20827 0.20585 -0.1298 0.0030 1.0000