XFOIL Version 6.94 Calculated polar for: DORNIER A-5 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 ------- -------- --------- --------- -------- ------- ------- 1.000 0.5109 0.00915 0.00424 -0.0984 0.7707 0.7022 1.500 0.5573 0.00924 0.00437 -0.0961 0.7529 0.7082 2.000 0.5997 0.00932 0.00443 -0.0928 0.7271 0.7147 2.500 0.6375 0.00936 0.00433 -0.0889 0.6928 0.7224 3.000 0.6704 0.00960 0.00445 -0.0837 0.6506 0.7267 3.500 0.6964 0.01007 0.00469 -0.0773 0.5902 0.7325 4.000 0.7088 0.01091 0.00512 -0.0685 0.5064 0.7404 4.500 0.7106 0.01214 0.00586 -0.0581 0.4043 0.7451 5.000 0.7246 0.01336 0.00664 -0.0506 0.3244 0.7512 5.500 0.7598 0.01404 0.00716 -0.0471 0.2912 0.7580 6.500 0.8057 0.01622 0.00869 -0.0364 0.1614 0.7694 7.000 0.8394 0.01704 0.00941 -0.0332 0.1437 0.7757 7.500 0.8735 0.01767 0.01006 -0.0299 0.1296 0.7809 8.000 0.9077 0.01843 0.01076 -0.0268 0.1084 0.7870 8.500 0.9371 0.01956 0.01169 -0.0233 0.0892 0.7926 9.000 0.9662 0.02049 0.01258 -0.0198 0.0677 0.7973 9.500 0.9950 0.02162 0.01363 -0.0164 0.0536 0.8022 10.000 1.0231 0.02286 0.01485 -0.0132 0.0405 0.8076 10.500 1.0457 0.02438 0.01630 -0.0094 0.0268 0.8125 11.000 1.0563 0.02678 0.01863 -0.0046 0.0035 0.8169 11.500 1.0827 0.02829 0.02024 -0.0018 0.0028 0.8219 12.000 1.1073 0.02999 0.02205 0.0009 0.0027 0.8266 12.500 1.1303 0.03179 0.02400 0.0036 0.0026 0.8313 13.000 1.1510 0.03390 0.02627 0.0062 0.0025 0.8353 13.500 1.1696 0.03629 0.02882 0.0086 0.0025 0.8398 14.000 1.1860 0.03901 0.03172 0.0108 0.0025 0.8441 14.500 1.1995 0.04210 0.03502 0.0127 0.0025 0.8482 15.000 1.2099 0.04570 0.03883 0.0143 0.0025 0.8524 15.500 1.2170 0.04990 0.04325 0.0154 0.0025 0.8571 16.000 1.2217 0.05474 0.04832 0.0157 0.0025 0.8612 16.500 1.2228 0.06049 0.05434 0.0150 0.0025 0.8658 17.000 1.2192 0.06729 0.06141 0.0135 0.0025 0.8708 17.500 1.2108 0.07526 0.06966 0.0108 0.0025 0.8759 18.000 1.1957 0.08521 0.07994 0.0062 0.0025 0.8820 18.500 1.1757 0.09686 0.09192 0.0000 0.0025 0.8892 19.000 1.1493 0.11033 0.10575 -0.0079 0.0025 0.8973 19.500 1.1200 0.12536 0.12116 -0.0175 0.0026 0.9080 20.000 1.0910 0.14283 0.13901 -0.0305 0.0026 0.9243 20.500 1.0521 0.15946 0.15592 -0.0406 0.0026 1.0000 21.000 1.0197 0.17513 0.17180 -0.0501 0.0026 1.0000 21.500 0.9934 0.19014 0.18699 -0.0594 0.0027 1.0000 22.000 0.9780 0.20322 0.20020 -0.0677 0.0027 1.0000 22.500 0.9732 0.21394 0.21101 -0.0748 0.0027 1.0000 23.000 0.9775 0.22232 0.21943 -0.0807 0.0028 1.0000 23.500 0.9879 0.22886 0.22601 -0.0858 0.0028 1.0000 24.000 1.0009 0.23449 0.23165 -0.0906 0.0028 1.0000 24.500 1.0157 0.23939 0.23656 -0.0951 0.0028 1.0000 25.000 1.0302 0.24441 0.24159 -0.0998 0.0028 1.0000 25.500 1.0450 0.24924 0.24646 -0.1045 0.0029 1.0000 26.000 1.0568 0.25533 0.25260 -0.1099 0.0029 1.0000 26.500 1.0681 0.26175 0.25909 -0.1157 0.0029 1.0000 27.000 1.0784 0.26865 0.26606 -0.1219 0.0030 1.0000 27.500 1.0864 0.27681 0.27432 -0.1288 0.0031 1.0000