XFOIL Version 6.94 Calculated polar for: nabac 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.018 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.3381 0.04824 0.03793 0.0043 0.9999 0.2062 -2.750 -0.3010 0.04577 0.03527 0.0038 0.9999 0.2024 -2.500 -0.2618 0.04338 0.03281 0.0032 0.9999 0.2003 -2.250 -0.2200 0.04123 0.03066 0.0023 0.9999 0.2003 -2.000 -0.1761 0.04171 0.02887 0.0030 0.4335 0.2030 -1.750 -0.1414 0.04141 0.02775 0.0027 0.3960 0.2085 -1.500 -0.1064 0.04095 0.02685 0.0025 0.3772 0.2177 -1.250 -0.0731 0.04056 0.02615 0.0026 0.3645 0.2304 -1.000 -0.0392 0.03988 0.02560 0.0027 0.3556 0.2514 -0.750 -0.0067 0.03897 0.02519 0.0028 0.3486 0.2962 -0.500 0.0695 0.03747 0.02516 -0.0038 0.3402 1.0001 -0.250 0.1014 0.03829 0.02543 -0.0031 0.3376 1.0001 0.000 0.1329 0.03921 0.02588 -0.0026 0.3360 1.0001 0.250 0.1645 0.04021 0.02657 -0.0022 0.3354 1.0001 0.500 0.1961 0.04130 0.02740 -0.0021 0.3360 1.0001 0.750 0.2277 0.04251 0.02837 -0.0021 0.3373 1.0001 1.000 0.2590 0.04385 0.02948 -0.0023 0.3390 1.0001 1.250 0.2898 0.04536 0.03075 -0.0027 0.3410 1.0001 1.500 0.3231 0.04662 0.03209 -0.0040 0.3449 1.0001 1.750 0.3561 0.04823 0.03374 -0.0057 0.3500 1.0001 2.000 0.3879 0.05015 0.03565 -0.0075 0.3554 1.0001 2.250 0.4182 0.05229 0.03768 -0.0091 0.3607 1.0001 2.500 0.4482 0.05454 0.03993 -0.0110 0.3663 1.0001 2.750 0.4808 0.05720 0.04289 -0.0157 0.3772 1.0001 3.000 0.5082 0.06008 0.04575 -0.0180 0.3853 1.0001 3.250 0.5370 0.06358 0.04956 -0.0243 0.3995 1.0001 3.500 0.5618 0.06704 0.05297 -0.0269 0.4101 1.0001 3.750 0.5798 0.07167 0.05786 -0.0346 0.4304 1.0001 4.000 0.5890 0.07650 0.06282 -0.0412 0.4520 1.0001 4.250 0.5915 0.08131 0.06764 -0.0466 0.4756 1.0001 4.500 0.5847 0.08616 0.07246 -0.0512 0.5031 1.0001 4.750 0.5837 0.09078 0.07698 -0.0549 0.5325 1.0001 5.000 0.5487 0.09515 0.08124 -0.0564 0.5709 1.0001