XFOIL Version 6.96 Calculated polar for: NACA-0009 9.0% smoothed 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.050 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -9.000 -0.5693 0.10116 0.09376 0.0165 1.0000 0.3639 -8.750 -0.7142 0.07868 0.07178 -0.0149 1.0000 0.1757 -8.500 -0.7137 0.07346 0.06654 -0.0151 1.0000 0.1724 -8.250 -0.7642 0.06347 0.05612 -0.0184 1.0000 0.1482 -8.000 -0.7681 0.05866 0.05100 -0.0173 1.0000 0.1481 -7.750 -0.7702 0.05414 0.04605 -0.0158 1.0000 0.1486 -7.500 -0.7678 0.04985 0.04126 -0.0139 1.0000 0.1489 -7.250 -0.7586 0.04570 0.03671 -0.0122 1.0000 0.1494 -7.000 -0.7428 0.04225 0.03304 -0.0107 1.0000 0.1522 -6.750 -0.7290 0.03964 0.03006 -0.0089 1.0000 0.1595 -6.500 -0.7134 0.03689 0.02699 -0.0071 1.0000 0.1662 -6.250 -0.6942 0.03431 0.02407 -0.0056 1.0000 0.1718 -6.000 -0.6744 0.03181 0.02110 -0.0040 1.0000 0.1782 -5.750 -0.6537 0.03001 0.01911 -0.0026 1.0000 0.1915 -5.500 -0.6311 0.02822 0.01724 -0.0014 1.0000 0.2060 -5.250 -0.6064 0.02638 0.01536 -0.0003 1.0000 0.2220 -5.000 -0.5831 0.02481 0.01382 0.0009 1.0000 0.2480 -4.750 -0.5614 0.02330 0.01254 0.0024 1.0000 0.2889 -4.500 -0.5414 0.02154 0.01133 0.0044 1.0000 0.3597 -4.250 -0.5291 0.01998 0.01068 0.0084 1.0000 0.4790 -4.000 -0.5177 0.01915 0.01049 0.0134 1.0000 0.5943 -3.750 -0.5027 0.01887 0.01051 0.0185 1.0000 0.6872 -3.500 -0.4803 0.01909 0.01089 0.0231 1.0000 0.7695 -3.250 -0.4247 0.02039 0.01199 0.0234 1.0000 0.8489 -3.000 -0.2970 0.02197 0.01277 0.0091 1.0000 0.9159 -2.750 -0.1843 0.02164 0.01183 -0.0069 1.0000 0.9649 -2.500 -0.0912 0.02030 0.01010 -0.0216 1.0000 1.0000 -2.250 -0.0830 0.01956 0.00931 -0.0204 1.0000 1.0000 -2.000 -0.0746 0.01894 0.00862 -0.0188 1.0000 1.0000 -1.750 -0.0661 0.01842 0.00807 -0.0170 1.0000 1.0000 -1.500 -0.0574 0.01800 0.00761 -0.0149 1.0000 1.0000 -1.250 -0.0485 0.01766 0.00724 -0.0126 1.0000 1.0000 -1.000 -0.0393 0.01739 0.00695 -0.0102 1.0000 1.0000 -0.750 -0.0297 0.01719 0.00674 -0.0077 1.0000 1.0000 -0.500 -0.0199 0.01705 0.00658 -0.0052 1.0000 1.0000 -0.250 -0.0100 0.01696 0.00649 -0.0026 1.0000 1.0000 0.000 0.0000 0.01694 0.00646 0.0000 1.0000 1.0000 0.250 0.0100 0.01696 0.00649 0.0026 1.0000 1.0000 0.500 0.0199 0.01704 0.00658 0.0052 1.0000 1.0000 0.750 0.0297 0.01718 0.00674 0.0077 1.0000 1.0000 1.000 0.0393 0.01739 0.00695 0.0102 1.0000 1.0000 1.250 0.0485 0.01765 0.00724 0.0126 1.0000 1.0000 1.500 0.0574 0.01800 0.00760 0.0149 1.0000 1.0000 1.750 0.0661 0.01842 0.00806 0.0170 1.0000 1.0000 2.000 0.0747 0.01894 0.00862 0.0188 1.0000 1.0000 2.250 0.0831 0.01956 0.00931 0.0204 1.0000 1.0000 2.500 0.0913 0.02030 0.01009 0.0216 1.0000 1.0000 2.750 0.1843 0.02163 0.01182 0.0070 0.9650 1.0000 3.000 0.2974 0.02197 0.01276 -0.0091 0.9158 1.0000 3.250 0.4247 0.02038 0.01198 -0.0234 0.8489 1.0000 3.500 0.4803 0.01909 0.01088 -0.0231 0.7695 1.0000 3.750 0.5026 0.01887 0.01052 -0.0185 0.6874 1.0000 4.000 0.5176 0.01915 0.01049 -0.0134 0.5942 1.0000 4.250 0.5291 0.01998 0.01068 -0.0083 0.4794 1.0000 4.500 0.5413 0.02154 0.01133 -0.0044 0.3598 1.0000 4.750 0.5613 0.02330 0.01253 -0.0024 0.2885 1.0000 5.000 0.5830 0.02481 0.01382 -0.0009 0.2479 1.0000 5.250 0.6063 0.02638 0.01535 0.0004 0.2220 1.0000 5.500 0.6311 0.02822 0.01724 0.0014 0.2060 1.0000 5.750 0.6536 0.03001 0.01911 0.0026 0.1915 1.0000 6.000 0.6744 0.03181 0.02109 0.0040 0.1782 1.0000 6.250 0.6942 0.03431 0.02406 0.0056 0.1718 1.0000 6.500 0.7133 0.03689 0.02698 0.0072 0.1662 1.0000 6.750 0.7289 0.03964 0.03006 0.0089 0.1595 1.0000 7.000 0.7428 0.04226 0.03304 0.0107 0.1522 1.0000 7.250 0.7590 0.04572 0.03670 0.0121 0.1493 1.0000 7.500 0.7679 0.04985 0.04125 0.0139 0.1489 1.0000 7.750 0.7707 0.05416 0.04606 0.0158 0.1486 1.0000 8.000 0.7677 0.05871 0.05106 0.0174 0.1481 1.0000 8.250 0.7649 0.06348 0.05611 0.0184 0.1482 1.0000 8.500 0.7260 0.07196 0.06501 0.0164 0.1658 1.0000 8.750 0.6992 0.08127 0.07438 0.0121 0.1896 1.0000 9.000 0.5669 0.10116 0.09375 -0.0169 0.3666 1.0000