XFOIL Version 6.94 Calculated polar for: AIRFOIL 3024 9.84% 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 0.4212 0.00633 0.00215 -0.0849 0.7661 1.0000 0.500 0.4711 0.00637 0.00203 -0.0835 0.7488 1.0000 1.000 0.5216 0.00643 0.00198 -0.0823 0.7299 1.0000 1.500 0.5724 0.00652 0.00198 -0.0811 0.7091 1.0000 2.000 0.6232 0.00666 0.00199 -0.0799 0.6856 1.0000 2.500 0.6740 0.00683 0.00207 -0.0788 0.6593 1.0000 3.000 0.7245 0.00706 0.00218 -0.0776 0.6289 1.0000 3.500 0.7747 0.00736 0.00236 -0.0764 0.5937 1.0000 4.000 0.8243 0.00773 0.00260 -0.0752 0.5547 1.0000 4.500 0.8727 0.00821 0.00293 -0.0738 0.5123 1.0000 5.000 0.9208 0.00873 0.00332 -0.0725 0.4663 1.0000 5.500 0.9674 0.00937 0.00380 -0.0709 0.4162 1.0000 6.000 1.0116 0.01017 0.00438 -0.0691 0.3595 1.0000 6.500 1.0529 0.01118 0.00508 -0.0670 0.2915 1.0000 7.000 1.0952 0.01212 0.00583 -0.0650 0.2411 1.0000 7.500 1.1355 0.01315 0.00668 -0.0628 0.1914 1.0000 8.000 1.1733 0.01432 0.00765 -0.0602 0.1436 1.0000 8.500 1.2044 0.01587 0.00887 -0.0568 0.0882 1.0000 9.000 1.2206 0.01807 0.01067 -0.0511 0.0276 1.0000 9.500 1.2408 0.01992 0.01251 -0.0459 0.0165 1.0000 10.000 1.2651 0.02159 0.01437 -0.0418 0.0148 1.0000 10.500 1.2835 0.02375 0.01671 -0.0375 0.0135 1.0000 11.000 1.2897 0.02691 0.02009 -0.0325 0.0124 1.0000 11.500 1.2999 0.03002 0.02340 -0.0287 0.0120 1.0000 12.000 1.3062 0.03372 0.02733 -0.0254 0.0118 1.0000 12.500 1.3113 0.03788 0.03171 -0.0226 0.0114 1.0000 13.000 1.3146 0.04255 0.03662 -0.0205 0.0111 1.0000 13.500 1.3166 0.04771 0.04205 -0.0190 0.0109 1.0000 14.000 1.3165 0.05352 0.04813 -0.0180 0.0108 1.0000 14.500 1.3150 0.05993 0.05484 -0.0178 0.0107 1.0000 15.000 1.3087 0.06742 0.06265 -0.0186 0.0106 1.0000 15.500 1.2967 0.07633 0.07191 -0.0206 0.0106 1.0000 16.000 1.2770 0.08711 0.08307 -0.0243 0.0106 1.0000 16.500 1.2514 0.09984 0.09619 -0.0301 0.0108 1.0000 17.000 1.2192 0.11490 0.11163 -0.0382 0.0110 1.0000 17.500 1.1843 0.13192 0.12900 -0.0483 0.0112 1.0000 18.000 1.1471 0.15096 0.14835 -0.0603 0.0114 1.0000 18.500 1.1019 0.17448 0.17214 -0.0751 0.0117 1.0000