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 = 0.045 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1709 0.02392 0.01251 -0.0308 1.0000 0.0830 -2.750 -0.1441 0.02240 0.01087 -0.0296 1.0000 0.0855 -2.500 -0.1175 0.02123 0.00960 -0.0284 1.0000 0.0929 -2.250 -0.0911 0.02011 0.00845 -0.0275 1.0000 0.1064 -2.000 -0.0639 0.01880 0.00740 -0.0268 1.0000 0.1488 -1.750 -0.0430 0.01519 0.00664 -0.0240 1.0000 1.0000 -1.500 -0.0176 0.01523 0.00629 -0.0228 1.0000 1.0000 -1.250 0.0074 0.01530 0.00612 -0.0217 1.0000 1.0000 -1.000 0.0323 0.01537 0.00598 -0.0207 1.0000 1.0000 -0.750 0.0573 0.01544 0.00589 -0.0197 1.0000 1.0000 -0.500 0.0827 0.01551 0.00599 -0.0187 1.0000 1.0000 -0.250 0.1588 0.01817 0.00607 -0.0247 0.2663 1.0000 0.000 0.1829 0.01921 0.00657 -0.0238 0.2316 1.0000 0.250 0.2075 0.02016 0.00710 -0.0230 0.2088 1.0000 0.500 0.2338 0.02097 0.00773 -0.0222 0.1976 1.0000 0.750 0.2614 0.02198 0.00853 -0.0215 0.1917 1.0000 1.000 0.2912 0.02289 0.00945 -0.0211 0.1881 1.0000 1.250 0.3224 0.02392 0.01052 -0.0209 0.1860 1.0000 1.500 0.3535 0.02509 0.01177 -0.0207 0.1856 1.0000 1.750 0.3838 0.02641 0.01323 -0.0205 0.1866 1.0000 2.000 0.4133 0.02792 0.01492 -0.0203 0.1888 1.0000 2.250 0.4425 0.02969 0.01685 -0.0202 0.1917 1.0000 2.500 0.4734 0.03124 0.01888 -0.0201 0.1978 1.0000 2.750 0.5035 0.03343 0.02149 -0.0202 0.2062 1.0000 3.000 0.5336 0.03555 0.02404 -0.0204 0.2135 1.0000 3.250 0.5627 0.03793 0.02692 -0.0208 0.2173 1.0000 3.500 0.5903 0.04103 0.03030 -0.0212 0.2254 1.0000 4.000 0.6256 0.05985 0.05220 -0.0892 0.7360 1.0000 4.500 0.4171 0.05704 0.04876 -0.0608 1.0000 1.0000 4.750 0.4290 0.05995 0.05164 -0.0623 1.0000 1.0000 5.000 0.4407 0.06292 0.05460 -0.0637 1.0000 1.0000