XFOIL Version 6.94 Calculated polar for: CLARK V 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.4997 0.00880 0.00424 -0.1151 0.8599 1.0000 0.500 0.5520 0.00857 0.00387 -0.1142 0.8424 1.0000 1.000 0.6010 0.00851 0.00368 -0.1127 0.8224 1.0000 1.500 0.6506 0.00850 0.00358 -0.1112 0.8010 1.0000 2.000 0.6971 0.00855 0.00352 -0.1090 0.7730 1.0000 2.500 0.7441 0.00865 0.00353 -0.1069 0.7443 1.0000 3.000 0.7909 0.00883 0.00361 -0.1048 0.7163 1.0000 3.500 0.8360 0.00907 0.00380 -0.1024 0.6854 1.0000 4.000 0.8696 0.00951 0.00388 -0.0973 0.6122 1.0000 4.500 0.9090 0.01002 0.00418 -0.0938 0.5604 1.0000 5.000 0.9281 0.01125 0.00465 -0.0865 0.4277 1.0000 5.500 0.9460 0.01298 0.00552 -0.0797 0.2885 1.0000 6.000 0.9675 0.01486 0.00656 -0.0740 0.1588 1.0000 6.500 0.9801 0.01741 0.00825 -0.0671 0.0242 1.0000 7.000 1.0168 0.01847 0.00926 -0.0639 0.0057 1.0000 7.500 1.0554 0.01940 0.01028 -0.0612 0.0050 1.0000 8.000 1.0917 0.02051 0.01156 -0.0581 0.0049 1.0000 8.500 1.1256 0.02180 0.01307 -0.0548 0.0048 1.0000 9.000 1.1560 0.02337 0.01495 -0.0511 0.0049 1.0000 9.500 1.1810 0.02536 0.01723 -0.0470 0.0049 1.0000 10.000 1.1982 0.02796 0.02012 -0.0422 0.0050 1.0000 10.500 1.2088 0.03114 0.02358 -0.0373 0.0051 1.0000 11.000 1.2192 0.03451 0.02717 -0.0331 0.0053 1.0000 11.500 1.2280 0.03824 0.03113 -0.0293 0.0055 1.0000 12.000 1.2354 0.04235 0.03546 -0.0261 0.0058 1.0000 12.500 1.2409 0.04689 0.04031 -0.0233 0.0060 1.0000 13.000 1.2443 0.05198 0.04569 -0.0209 0.0065 1.0000 13.500 1.2442 0.05789 0.05198 -0.0185 0.0071 1.0000 14.000 1.2377 0.06501 0.05952 -0.0168 0.0077 1.0000 14.500 1.2210 0.07390 0.06888 -0.0166 0.0083 1.0000 15.000 1.1971 0.08443 0.07984 -0.0185 0.0087 1.0000 15.500 1.1677 0.09677 0.09258 -0.0228 0.0090 1.0000 16.000 1.1364 0.11071 0.10690 -0.0297 0.0091 1.0000 16.500 1.1052 0.12620 0.12272 -0.0387 0.0092 1.0000 17.000 1.0748 0.14325 0.14007 -0.0496 0.0091 1.0000 17.500 1.0430 0.16272 0.15980 -0.0624 0.0090 1.0000 18.000 1.0011 0.18926 0.18650 -0.0783 0.0089 1.0000 18.500 0.9848 0.20873 0.20586 -0.0883 0.0099 1.0000 19.500 0.7099 0.20100 0.19839 -0.0591 0.0096 1.0000 20.000 0.7005 0.21244 0.20982 -0.0619 0.0156 1.0000