Xfoil polar. Reynolds number fixed. Mach number fixed Polar key,xf-ma409sm-il-500000 Airfoil,ma409sm-il Reynolds number,500000 Ncrit,9 Mach,0 Max Cl/Cd,121.638 Max Cl/Cd alpha,2.5 Url,http://airfoiltools.com/polar/csv?polar=xf-ma409sm-il-500000 Alpha,Cl,Cd,Cdp,Cm,Top_Xtr,Bot_Xtr 0.250,0.4700,0.00528,0.00176,-0.1085,0.9080,1.0000 0.500,0.4983,0.00526,0.00169,-0.1083,0.8969,1.0000 0.750,0.5263,0.00524,0.00163,-0.1081,0.8843,1.0000 1.000,0.5539,0.00525,0.00159,-0.1078,0.8699,1.0000 1.250,0.5813,0.00528,0.00156,-0.1074,0.8535,1.0000 1.500,0.6084,0.00534,0.00155,-0.1070,0.8354,1.0000 1.750,0.6351,0.00543,0.00159,-0.1065,0.8144,1.0000 2.000,0.6614,0.00554,0.00162,-0.1059,0.7914,1.0000 2.500,0.7128,0.00586,0.00175,-0.1046,0.7358,1.0000 2.750,0.7380,0.00607,0.00185,-0.1039,0.7033,1.0000 3.000,0.7623,0.00634,0.00199,-0.1030,0.6640,1.0000 3.250,0.7839,0.00680,0.00214,-0.1015,0.5958,1.0000 3.500,0.8054,0.00735,0.00235,-0.1002,0.5230,1.0000 3.750,0.8277,0.00788,0.00260,-0.0992,0.4598,1.0000 4.000,0.8503,0.00844,0.00287,-0.0982,0.3937,1.0000 4.250,0.8727,0.00908,0.00320,-0.0973,0.3212,1.0000 4.500,0.8937,0.00992,0.00361,-0.0963,0.2307,1.0000 4.750,0.9131,0.01105,0.00418,-0.0951,0.1217,1.0000 5.000,0.9315,0.01242,0.00503,-0.0937,0.0347,1.0000 5.250,0.9551,0.01308,0.00573,-0.0927,0.0255,1.0000 5.500,0.9754,0.01427,0.00704,-0.0912,0.0207,1.0000 5.750,0.9992,0.01483,0.00768,-0.0903,0.0197,1.0000 6.000,1.0225,0.01545,0.00837,-0.0895,0.0182,1.0000 6.250,1.0443,0.01628,0.00929,-0.0883,0.0171,1.0000 6.500,1.0658,0.01719,0.01027,-0.0871,0.0162,1.0000 6.750,1.0867,0.01820,0.01135,-0.0858,0.0153,1.0000 7.000,1.1057,0.01966,0.01289,-0.0843,0.0142,1.0000 7.250,1.1227,0.02237,0.01578,-0.0824,0.0129,1.0000 7.500,1.1453,0.02330,0.01688,-0.0814,0.0123,1.0000 7.750,1.1666,0.02489,0.01864,-0.0801,0.0117,1.0000 8.000,1.1869,0.02684,0.02081,-0.0787,0.0112,1.0000