XFOIL Version 6.96 Calculated polar for: AH 79-100 B AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -4.500 0.3884 0.00978 0.00464 -0.2070 0.7792 0.0226 -4.000 0.4451 0.00961 0.00442 -0.2074 0.7692 0.0253 -3.500 0.5032 0.00896 0.00363 -0.2081 0.7583 0.0268 -3.000 0.5629 0.00823 0.00273 -0.2093 0.7485 0.0287 -2.500 0.6223 0.00778 0.00221 -0.2103 0.7396 0.0314 -2.000 0.6805 0.00757 0.00192 -0.2109 0.7303 0.0348 -1.500 0.7386 0.00738 0.00164 -0.2115 0.7200 0.0460 -1.000 0.7978 0.00701 0.00153 -0.2125 0.7092 0.1489 -0.500 0.8586 0.00649 0.00161 -0.2143 0.6996 0.4190 0.000 0.9154 0.00646 0.00173 -0.2147 0.6896 0.5284 0.500 0.9716 0.00649 0.00182 -0.2148 0.6790 0.5800 1.000 1.0272 0.00652 0.00195 -0.2149 0.6668 0.6368 1.500 1.0823 0.00660 0.00209 -0.2148 0.6535 0.6771 2.000 1.1367 0.00669 0.00225 -0.2146 0.6383 0.7287 2.500 1.1889 0.00682 0.00245 -0.2140 0.6157 0.8036 3.000 1.2339 0.00673 0.00259 -0.2116 0.5943 1.0000 3.500 1.2866 0.00704 0.00282 -0.2111 0.5740 1.0000 4.000 1.3387 0.00738 0.00309 -0.2106 0.5493 1.0000 4.500 1.3872 0.00792 0.00346 -0.2094 0.5103 1.0000 5.000 1.4357 0.00847 0.00388 -0.2082 0.4734 1.0000 5.500 1.4798 0.00930 0.00447 -0.2062 0.4206 1.0000 6.000 1.5233 0.01016 0.00511 -0.2042 0.3718 1.0000 6.500 1.5593 0.01150 0.00605 -0.2009 0.2942 1.0000 7.000 1.5861 0.01337 0.00738 -0.1961 0.2011 1.0000 7.500 1.6036 0.01544 0.00896 -0.1895 0.1162 1.0000 8.000 1.6181 0.01750 0.01068 -0.1825 0.0590 1.0000 8.500 1.6455 0.01888 0.01203 -0.1780 0.0436 1.0000 9.000 1.6661 0.02075 0.01380 -0.1727 0.0230 1.0000 9.500 1.6824 0.02306 0.01610 -0.1671 0.0123 1.0000 10.000 1.7045 0.02509 0.01823 -0.1627 0.0102 1.0000