XFOIL Version 6.94 Calculated polar for: GOA 1b 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.005 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1807 0.04585 0.01876 -0.0212 1.0000 1.0000 -2.750 -0.1558 0.04561 0.01730 -0.0208 1.0000 1.0000 -2.500 -0.1312 0.04542 0.01604 -0.0203 1.0000 1.0000 -2.250 -0.1069 0.04526 0.01495 -0.0197 1.0000 1.0000 -2.000 -0.0826 0.04513 0.01388 -0.0191 1.0000 1.0000 -1.750 -0.0586 0.04504 0.01307 -0.0185 1.0000 1.0000 -1.500 -0.0346 0.04498 0.01240 -0.0178 1.0000 1.0000 -1.250 -0.0107 0.04496 0.01186 -0.0172 1.0000 1.0000 -1.000 0.0131 0.04496 0.01144 -0.0165 1.0000 1.0000 -0.750 0.0369 0.04498 0.01115 -0.0158 1.0000 1.0000 -0.500 0.0606 0.04504 0.01087 -0.0151 1.0000 1.0000 -0.250 0.0842 0.04512 0.01081 -0.0144 1.0000 1.0000 0.000 0.1077 0.04523 0.01087 -0.0137 1.0000 1.0000 0.250 0.1313 0.04537 0.01105 -0.0130 1.0000 1.0000 0.500 0.1548 0.04553 0.01137 -0.0123 1.0000 1.0000 0.750 0.1783 0.04572 0.01177 -0.0115 1.0000 1.0000 1.000 0.2019 0.04594 0.01234 -0.0108 1.0000 1.0000 1.250 0.2257 0.04618 0.01306 -0.0101 1.0000 1.0000 1.500 0.2497 0.04644 0.01395 -0.0095 1.0000 1.0000 1.750 0.2743 0.04674 0.01504 -0.0089 1.0000 1.0000 2.000 0.2997 0.04706 0.01638 -0.0086 1.0000 1.0000 2.250 0.3266 0.04743 0.01804 -0.0087 1.0000 1.0000 2.500 0.3544 0.04800 0.02018 -0.0098 1.0000 1.0000 2.750 0.3765 0.04921 0.02294 -0.0127 1.0000 1.0000 3.000 0.3824 0.05161 0.02594 -0.0161 1.0000 1.0000 3.250 0.3812 0.05472 0.02887 -0.0191 1.0000 1.0000 3.500 0.3820 0.05783 0.03172 -0.0219 1.0000 1.0000 3.750 0.3854 0.06084 0.03452 -0.0245 1.0000 1.0000 4.000 0.3904 0.06379 0.03730 -0.0270 1.0000 1.0000 4.250 0.3965 0.06670 0.04009 -0.0293 1.0000 1.0000 4.500 0.4034 0.06959 0.04288 -0.0315 1.0000 1.0000 4.750 0.4109 0.07249 0.04569 -0.0335 1.0000 1.0000 5.000 0.4188 0.07540 0.04854 -0.0355 1.0000 1.0000