XFOIL Version 6.94 Calculated polar for: agora05 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.030 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1032 0.03876 0.02804 -0.0261 0.9992 0.3611 -2.750 -0.0743 0.03663 0.02657 -0.0252 0.9992 0.4671 -2.500 -0.0592 0.03416 0.02585 -0.0197 0.9992 0.6809 -2.250 -0.0313 0.03284 0.02470 -0.0196 0.9992 1.0008 -2.000 -0.0028 0.03340 0.02467 -0.0206 0.9992 1.0008 -1.750 0.0229 0.03398 0.02491 -0.0208 0.9992 1.0008 -1.500 0.0477 0.03458 0.02529 -0.0209 0.9992 1.0008 -1.250 0.2545 0.03242 0.02292 -0.0469 0.8751 1.0008 -1.000 0.2609 0.03449 0.02492 -0.0380 0.8171 1.0008 -0.750 0.2560 0.03720 0.02753 -0.0264 0.7500 1.0008 -0.500 0.2524 0.04047 0.03068 -0.0157 0.6753 1.0008 -0.250 0.2581 0.04326 0.03333 -0.0080 0.5973 1.0008 0.000 0.3142 0.04061 0.03096 -0.0157 0.4692 1.0008 0.250 0.3313 0.03912 0.02907 -0.0088 0.4461 1.0008 0.500 0.3486 0.03808 0.02747 -0.0026 0.4262 1.0008 0.750 0.3693 0.03799 0.02684 0.0015 0.4048 1.0008 1.000 0.3934 0.03864 0.02702 0.0039 0.3856 1.0008 1.250 0.4205 0.03967 0.02777 0.0051 0.3720 1.0008 1.500 0.4474 0.04093 0.02875 0.0063 0.3599 1.0008 1.750 0.4767 0.04249 0.03019 0.0064 0.3472 1.0008 2.000 0.5044 0.04417 0.03173 0.0069 0.3372 1.0008 2.250 0.5360 0.04616 0.03377 0.0061 0.3308 1.0008 2.500 0.5680 0.04849 0.03621 0.0047 0.3244 1.0008 2.750 0.5997 0.05116 0.03902 0.0028 0.3173 1.0008 3.000 0.6317 0.05430 0.04232 0.0002 0.3119 1.0008 3.250 0.6663 0.05896 0.04739 -0.0055 0.3115 1.0008 3.500 0.6979 0.06491 0.05371 -0.0126 0.3134 1.0008