XFOIL Version 6.94 Calculated polar for: EPPLER 68 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 0.4707 0.00751 0.00264 -0.1140 0.7713 0.7358 0.500 0.5228 0.00751 0.00271 -0.1130 0.7559 0.7761 1.000 0.5743 0.00754 0.00280 -0.1118 0.7400 0.8130 1.500 0.6245 0.00760 0.00289 -0.1103 0.7230 0.8478 2.000 0.6724 0.00768 0.00298 -0.1082 0.7051 0.8820 2.500 0.7158 0.00774 0.00307 -0.1051 0.6858 0.9160 3.000 0.7606 0.00780 0.00314 -0.1024 0.6652 0.9543 4.000 0.8853 0.00816 0.00345 -0.1057 0.6165 1.0000 4.500 0.9305 0.00845 0.00367 -0.1038 0.5892 1.0000 5.000 0.9747 0.00881 0.00394 -0.1016 0.5593 1.0000 5.500 1.0166 0.00922 0.00429 -0.0989 0.5251 1.0000 6.000 1.0549 0.00972 0.00470 -0.0956 0.4865 1.0000 6.500 1.0843 0.01037 0.00518 -0.0905 0.4402 1.0000 7.000 1.1085 0.01119 0.00581 -0.0846 0.3849 1.0000 7.500 1.1292 0.01230 0.00665 -0.0784 0.3221 1.0000 8.000 1.1473 0.01366 0.00769 -0.0722 0.2583 1.0000 8.500 1.1650 0.01519 0.00893 -0.0663 0.1966 1.0000 9.000 1.1841 0.01678 0.01029 -0.0611 0.1486 1.0000 9.500 1.2033 0.01848 0.01181 -0.0561 0.1110 1.0000 10.000 1.2217 0.02031 0.01352 -0.0514 0.0802 1.0000 10.500 1.2361 0.02252 0.01559 -0.0466 0.0507 1.0000 11.000 1.2415 0.02549 0.01841 -0.0412 0.0266 1.0000 11.500 1.2536 0.02823 0.02122 -0.0371 0.0205 1.0000 12.000 1.2648 0.03121 0.02434 -0.0334 0.0181 1.0000 12.500 1.2698 0.03489 0.02814 -0.0298 0.0166 1.0000 13.000 1.2780 0.03861 0.03205 -0.0271 0.0156 1.0000 13.500 1.2819 0.04299 0.03659 -0.0249 0.0148 1.0000 14.000 1.2777 0.04854 0.04227 -0.0231 0.0141 1.0000 14.500 1.2827 0.05356 0.04751 -0.0222 0.0136 1.0000 15.000 1.2866 0.05901 0.05316 -0.0218 0.0129 1.0000 15.500 1.2893 0.06491 0.05923 -0.0221 0.0123 1.0000 16.000 1.2907 0.07122 0.06568 -0.0226 0.0119 1.0000 16.500 1.2931 0.07772 0.07237 -0.0234 0.0115 1.0000 17.000 1.2925 0.08505 0.08000 -0.0252 0.0113 1.0000 17.500 1.2886 0.09322 0.08848 -0.0278 0.0111 1.0000 18.000 1.2793 0.10261 0.09821 -0.0316 0.0108 1.0000 18.500 1.2666 0.11296 0.10888 -0.0363 0.0107 1.0000 19.000 1.2492 0.12454 0.12080 -0.0424 0.0105 1.0000 19.500 1.2280 0.13730 0.13389 -0.0499 0.0104 1.0000 20.000 1.1999 0.15212 0.14906 -0.0593 0.0104 1.0000 20.500 1.1620 0.17015 0.16746 -0.0712 0.0106 1.0000 21.000 1.1027 0.19574 0.19343 -0.0878 0.0112 1.0000 23.000 0.8698 0.29485 0.29328 -0.1405 0.0141 1.0000 23.500 0.8687 0.30902 0.30744 -0.1464 0.0130 1.0000 24.000 0.8765 0.31747 0.31591 -0.1499 0.0120 1.0000 24.500 0.8877 0.32389 0.32236 -0.1520 0.0117 1.0000 25.000 0.8888 0.33761 0.33607 -0.1573 0.0110 1.0000 25.500 0.8954 0.34711 0.34559 -0.1608 0.0101 1.0000