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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1751 0.02529 0.01385 -0.0320 1.0000 0.0740 -2.750 -0.1485 0.02371 0.01206 -0.0306 1.0000 0.0754 -2.500 -0.1212 0.02240 0.01050 -0.0293 1.0000 0.0785 -2.250 -0.0947 0.02123 0.00922 -0.0281 1.0000 0.0828 -2.000 -0.0683 0.02019 0.00819 -0.0271 1.0000 0.0908 -1.750 -0.0413 0.01925 0.00729 -0.0263 1.0000 0.1051 -1.500 -0.0210 0.01523 0.00665 -0.0239 1.0000 1.0000 -1.250 0.0050 0.01528 0.00623 -0.0226 1.0000 1.0000 -1.000 0.0305 0.01534 0.00611 -0.0215 1.0000 1.0000 -0.750 0.0560 0.01543 0.00612 -0.0203 1.0000 1.0000 -0.500 0.0818 0.01551 0.00627 -0.0191 1.0000 1.0000 0.000 0.1831 0.01922 0.00655 -0.0238 0.2311 1.0000 0.250 0.2077 0.02016 0.00708 -0.0230 0.2087 1.0000 0.500 0.2340 0.02098 0.00772 -0.0222 0.1975 1.0000 0.750 0.2616 0.02198 0.00853 -0.0215 0.1918 1.0000 1.000 0.2914 0.02289 0.00945 -0.0211 0.1881 1.0000 1.250 0.3226 0.02392 0.01052 -0.0209 0.1861 1.0000 1.500 0.3537 0.02510 0.01179 -0.0207 0.1857 1.0000 1.750 0.3839 0.02642 0.01326 -0.0205 0.1867 1.0000 2.000 0.4134 0.02793 0.01495 -0.0203 0.1889 1.0000 2.250 0.4425 0.02971 0.01692 -0.0202 0.1918 1.0000 2.500 0.4733 0.03125 0.01893 -0.0201 0.1979 1.0000 2.750 0.5034 0.03344 0.02149 -0.0202 0.2063 1.0000 3.000 0.5333 0.03556 0.02402 -0.0204 0.2135 1.0000 3.250 0.5625 0.03794 0.02688 -0.0209 0.2173 1.0000 3.500 0.5900 0.04103 0.03019 -0.0213 0.2254 1.0000 3.750 0.6338 0.04499 0.03537 -0.0259 0.2880 1.0000 4.000 0.6249 0.05985 0.05219 -0.0892 0.7363 1.0000 4.250 0.5134 0.06058 0.05263 -0.0816 0.9014 1.0000 4.500 0.4171 0.05707 0.04880 -0.0609 1.0000 1.0000 4.750 0.4290 0.05998 0.05169 -0.0623 1.0000 1.0000 5.000 0.4407 0.06295 0.05468 -0.0637 1.0000 1.0000