XFOIL Version 6.94 Calculated polar for: GOA 1 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 -0.1915 0.03393 0.01844 -0.0317 1.0000 0.2394 -2.750 -0.1614 0.03174 0.01614 -0.0309 1.0000 0.2641 -2.500 -0.1339 0.02927 0.01405 -0.0296 1.0000 0.3109 -2.250 -0.1174 0.02506 0.01235 -0.0260 1.0000 0.5158 -2.000 -0.0765 0.02376 0.00983 -0.0239 1.0000 1.0000 -1.750 -0.0508 0.02374 0.00906 -0.0227 1.0000 1.0000 -1.500 -0.0257 0.02375 0.00857 -0.0216 1.0000 1.0000 -1.250 -0.0010 0.02377 0.00821 -0.0205 1.0000 1.0000 -1.000 0.0235 0.02381 0.00796 -0.0196 1.0000 1.0000 -0.750 0.0479 0.02387 0.00775 -0.0186 1.0000 1.0000 -0.500 0.0723 0.02394 0.00764 -0.0177 1.0000 1.0000 -0.250 0.0967 0.02403 0.00762 -0.0167 1.0000 1.0000 0.000 0.1215 0.02412 0.00753 -0.0157 1.0000 1.0000 0.250 0.1462 0.02424 0.00782 -0.0148 1.0000 1.0000 0.500 0.1719 0.02435 0.00833 -0.0137 1.0000 1.0000 0.750 0.2010 0.02441 0.00915 -0.0128 1.0000 1.0000 1.000 0.2411 0.02428 0.01037 -0.0134 1.0000 1.0000 1.250 0.3257 0.02898 0.01211 -0.0197 0.4389 1.0000 1.500 0.3549 0.03084 0.01362 -0.0193 0.4122 1.0000 1.750 0.3868 0.03261 0.01531 -0.0194 0.3954 1.0000 2.000 0.4200 0.03438 0.01720 -0.0200 0.3819 1.0000 2.250 0.4518 0.03626 0.01933 -0.0205 0.3673 1.0000 2.500 0.4829 0.03828 0.02153 -0.0210 0.3543 1.0000 2.750 0.5149 0.04039 0.02398 -0.0219 0.3478 1.0000 3.000 0.5481 0.04269 0.02678 -0.0235 0.3491 1.0000 3.250 0.5806 0.04526 0.02988 -0.0254 0.3528 1.0000 3.500 0.6116 0.04816 0.03321 -0.0273 0.3580 1.0000 3.750 0.6421 0.05124 0.03692 -0.0307 0.3670 1.0000 4.000 0.6708 0.05484 0.04087 -0.0333 0.3766 1.0000 4.250 0.6986 0.05881 0.04537 -0.0384 0.3921 1.0000 4.500 0.7236 0.06320 0.05015 -0.0437 0.4102 1.0000 4.750 0.7387 0.06796 0.05532 -0.0517 0.4353 1.0000 5.000 0.7555 0.07303 0.06059 -0.0581 0.4621 1.0000