XFOIL Version 6.94 Calculated polar for: NACA M24 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 ------- -------- --------- --------- -------- ------- ------- 0.000 0.3382 0.01315 0.00501 -0.0244 0.5163 0.0283 0.500 0.3954 0.01325 0.00514 -0.0250 0.5082 0.0369 1.000 0.4516 0.01294 0.00475 -0.0252 0.5003 0.0419 1.500 0.5060 0.01246 0.00433 -0.0251 0.4929 0.0450 2.000 0.5606 0.01209 0.00402 -0.0249 0.4850 0.0495 2.500 0.6149 0.01194 0.00385 -0.0248 0.4767 0.0595 3.000 0.6684 0.01170 0.00398 -0.0246 0.4697 0.1724 3.500 0.7899 0.01085 0.00496 -0.0386 0.4602 0.9818 4.000 0.9242 0.01128 0.00524 -0.0560 0.4497 1.0000 4.500 0.9762 0.01118 0.00501 -0.0558 0.4234 1.0000 5.000 1.0279 0.01144 0.00504 -0.0557 0.3922 1.0000 5.500 1.0796 0.01173 0.00527 -0.0557 0.3742 1.0000 6.000 1.1289 0.01240 0.00569 -0.0556 0.3329 1.0000 6.500 1.1744 0.01348 0.00645 -0.0552 0.2787 1.0000 7.000 1.1672 0.01931 0.01095 -0.0499 0.0367 1.0000 7.500 1.1803 0.02126 0.01285 -0.0447 0.0043 1.0000 8.000 1.1703 0.02324 0.01496 -0.0363 0.0041 1.0000 8.500 1.1661 0.02565 0.01748 -0.0297 0.0040 1.0000 9.000 1.1691 0.02845 0.02039 -0.0252 0.0040 1.0000 9.500 1.1747 0.03163 0.02368 -0.0217 0.0040 1.0000 10.000 1.1802 0.03527 0.02745 -0.0191 0.0041 1.0000 10.500 1.1852 0.03932 0.03165 -0.0171 0.0041 1.0000 11.000 1.1884 0.04383 0.03632 -0.0156 0.0042 1.0000 11.500 1.1898 0.04885 0.04150 -0.0145 0.0044 1.0000 12.000 1.1895 0.05439 0.04722 -0.0138 0.0045 1.0000 12.500 1.1871 0.06045 0.05345 -0.0135 0.0046 1.0000 13.000 1.1818 0.06709 0.06025 -0.0134 0.0047 1.0000 13.500 1.1732 0.07437 0.06770 -0.0137 0.0048 1.0000 14.000 1.1611 0.08231 0.07580 -0.0142 0.0049 1.0000 14.500 1.1615 0.08875 0.08237 -0.0149 0.0051 1.0000 15.000 1.1642 0.09496 0.08871 -0.0156 0.0054 1.0000 15.500 1.1617 0.10198 0.09587 -0.0165 0.0057 1.0000 16.000 1.1595 0.10891 0.10293 -0.0175 0.0061 1.0000 16.500 1.1619 0.11483 0.10891 -0.0182 0.0064 1.0000 17.000 1.1745 0.11930 0.11342 -0.0187 0.0069 1.0000 17.500 1.1949 0.12220 0.11647 -0.0182 0.0081 1.0000 18.000 0.9376 0.12555 0.12060 -0.0021 0.0065 1.0000 18.500 0.9465 0.12787 0.12302 -0.0023 0.0069 1.0000 19.000 0.9927 0.12402 0.11915 0.0013 0.0086 1.0000