XFOIL Version 6.94 Calculated polar for: EPPLER 668 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 0.5307 0.01136 0.00689 -0.1245 0.7873 0.8275 0.500 0.5847 0.01123 0.00676 -0.1237 0.7814 0.8403 1.000 0.6458 0.01108 0.00653 -0.1245 0.7765 0.8526 1.500 0.6908 0.01105 0.00658 -0.1221 0.7683 0.8644 2.000 0.7451 0.01074 0.00629 -0.1216 0.7605 0.8747 2.500 0.8047 0.01038 0.00590 -0.1220 0.7542 0.8841 3.000 0.8512 0.01015 0.00578 -0.1202 0.7425 0.8928 3.500 0.9077 0.00967 0.00529 -0.1201 0.7326 0.8998 4.000 0.9542 0.00944 0.00515 -0.1183 0.7164 0.9076 4.500 1.0017 0.00921 0.00496 -0.1166 0.6965 0.9141 5.000 1.0458 0.00909 0.00481 -0.1142 0.6695 0.9217 5.500 1.0831 0.00926 0.00491 -0.1106 0.6309 0.9301 6.000 1.1005 0.00956 0.00508 -0.1030 0.5846 0.9426 6.500 1.1109 0.01018 0.00556 -0.0944 0.5357 0.9623 7.000 1.1372 0.01122 0.00642 -0.0901 0.4807 1.0000 7.500 1.1574 0.01264 0.00762 -0.0853 0.4284 1.0000 8.000 1.1770 0.01423 0.00903 -0.0807 0.3763 1.0000 8.500 1.1935 0.01608 0.01065 -0.0760 0.3247 1.0000 9.000 1.2120 0.01798 0.01235 -0.0717 0.2753 1.0000 9.500 1.2312 0.01995 0.01413 -0.0677 0.2280 1.0000 10.000 1.2491 0.02210 0.01606 -0.0638 0.1830 1.0000 10.500 1.2678 0.02432 0.01811 -0.0602 0.1452 1.0000 11.000 1.2870 0.02662 0.02030 -0.0570 0.1138 1.0000 11.500 1.3060 0.02906 0.02266 -0.0539 0.0878 1.0000 12.000 1.3222 0.03182 0.02537 -0.0509 0.0674 1.0000 12.500 1.3370 0.03486 0.02839 -0.0480 0.0522 1.0000 13.000 1.3509 0.03814 0.03170 -0.0454 0.0409 1.0000 13.500 1.3631 0.04177 0.03541 -0.0431 0.0329 1.0000 14.000 1.3731 0.04581 0.03956 -0.0412 0.0275 1.0000 14.500 1.3811 0.05030 0.04418 -0.0396 0.0234 1.0000 15.000 1.3881 0.05518 0.04920 -0.0386 0.0200 1.0000 15.500 1.3944 0.06040 0.05460 -0.0381 0.0169 1.0000 16.000 1.3969 0.06643 0.06084 -0.0381 0.0144 1.0000 16.500 1.3932 0.07364 0.06827 -0.0389 0.0125 1.0000 17.000 1.3826 0.08233 0.07714 -0.0408 0.0110 1.0000 17.500 1.3794 0.09038 0.08546 -0.0431 0.0093 1.0000 18.000 1.3639 0.10080 0.09613 -0.0469 0.0083 1.0000 18.500 1.3532 0.11082 0.10642 -0.0512 0.0073 1.0000 19.000 1.3351 0.12240 0.11820 -0.0568 0.0068 1.0000 19.500 1.3201 0.13368 0.12977 -0.0627 0.0062 1.0000 20.000 1.3070 0.14468 0.14102 -0.0689 0.0057 1.0000 20.500 1.2935 0.15590 0.15244 -0.0757 0.0053 1.0000 21.000 1.2799 0.16715 0.16386 -0.0828 0.0050 1.0000 21.500 1.2658 0.17832 0.17518 -0.0900 0.0048 1.0000 22.000 1.2548 0.18918 0.18627 -0.0972 0.0046 1.0000 22.500 1.2429 0.20047 0.19778 -0.1048 0.0044 1.0000 23.000 1.2296 0.21228 0.20979 -0.1128 0.0043 1.0000 23.500 1.2125 0.22555 0.22328 -0.1216 0.0042 1.0000 24.000 1.1821 0.24377 0.24175 -0.1328 0.0043 1.0000