XFOIL Version 6.94 Calculated polar for: EPPLER 664 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.3906 0.02039 0.01509 -0.0635 0.7533 0.8552 1.000 0.4452 0.02013 0.01482 -0.0636 0.7458 0.8586 1.500 0.5038 0.01969 0.01433 -0.0646 0.7403 0.8621 2.000 0.5210 0.01951 0.01415 -0.0581 0.7322 0.8682 2.500 0.5382 0.01911 0.01376 -0.0521 0.7224 0.8735 3.000 0.6062 0.01840 0.01300 -0.0549 0.7169 0.8744 3.500 0.6355 0.01810 0.01280 -0.0502 0.7059 0.8772 4.000 0.6977 0.01740 0.01210 -0.0518 0.6979 0.8788 4.500 0.7401 0.01697 0.01175 -0.0499 0.6871 0.8807 5.000 0.7992 0.01626 0.01105 -0.0509 0.6765 0.8821 5.500 0.8351 0.01579 0.01070 -0.0477 0.6621 0.8853 6.000 0.8830 0.01521 0.01018 -0.0467 0.6473 0.8870 6.500 0.9194 0.01469 0.00975 -0.0435 0.6275 0.8889 7.000 0.9463 0.01439 0.00950 -0.0387 0.6012 0.8910 7.500 0.9780 0.01439 0.00947 -0.0349 0.5667 0.8931 8.000 0.9957 0.01472 0.00968 -0.0286 0.5241 0.8961 8.500 1.0091 0.01559 0.01041 -0.0224 0.4771 0.8986 9.000 1.0177 0.01691 0.01157 -0.0162 0.4306 0.9014 9.500 1.0254 0.01860 0.01311 -0.0106 0.3850 0.9043 10.000 1.0310 0.02068 0.01504 -0.0054 0.3348 0.9072 11.000 1.0432 0.02567 0.01964 0.0031 0.2368 0.9126 11.500 1.0514 0.02824 0.02207 0.0066 0.1950 0.9154 12.500 1.0740 0.03383 0.02745 0.0119 0.1253 0.9214 13.000 1.0875 0.03685 0.03042 0.0137 0.0998 0.9245 13.500 1.0994 0.04021 0.03373 0.0153 0.0781 0.9277 14.500 1.1227 0.04753 0.04112 0.0174 0.0471 0.9363 15.000 1.1320 0.05179 0.04547 0.0179 0.0366 0.9411 15.500 1.1379 0.05677 0.05056 0.0178 0.0283 0.9466 16.500 1.1297 0.07095 0.06509 0.0148 0.0159 0.9649 17.000 1.1225 0.07880 0.07317 0.0123 0.0142 1.0000 17.500 1.1187 0.08685 0.08140 0.0092 0.0133 1.0000 18.000 1.1126 0.09541 0.09010 0.0056 0.0128 1.0000 18.500 1.1123 0.10343 0.09838 0.0020 0.0121 1.0000 19.000 1.1140 0.11114 0.10629 -0.0019 0.0117 1.0000 19.500 1.1160 0.11888 0.11422 -0.0060 0.0113 1.0000 20.000 1.1191 0.12644 0.12196 -0.0103 0.0110 1.0000 20.500 1.1224 0.13387 0.12955 -0.0146 0.0107 1.0000 21.000 1.1256 0.14123 0.13708 -0.0191 0.0105 1.0000 21.500 1.1222 0.15000 0.14610 -0.0247 0.0105 1.0000 22.000 1.1090 0.16106 0.15748 -0.0322 0.0105 1.0000 22.500 1.0846 0.17498 0.17178 -0.0416 0.0107 1.0000