XFOIL Version 6.96 Calculated polar for: GOE 16K AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -19.750 -0.6958 0.07375 0.07025 -0.1756 0.9477 0.0154 -19.500 -0.7271 0.06649 0.06285 -0.1797 0.9471 0.0153 -19.250 -0.7569 0.05950 0.05571 -0.1836 0.9466 0.0154 -19.000 -0.7774 0.05357 0.04966 -0.1873 0.9461 0.0156 -18.750 -0.8235 0.04814 0.04410 -0.1853 0.9398 0.0161 -18.500 -0.8339 0.04399 0.03984 -0.1874 0.9379 0.0163 -18.250 -0.8398 0.04030 0.03602 -0.1893 0.9363 0.0162 -18.000 -0.8467 0.03679 0.03240 -0.1908 0.9350 0.0169 -17.750 -0.8475 0.03417 0.02968 -0.1914 0.9339 0.0170 -17.500 -0.8647 0.03229 0.02770 -0.1876 0.9272 0.0170 -17.250 -0.8655 0.03048 0.02580 -0.1866 0.9245 0.0176 -17.000 -0.8604 0.02885 0.02409 -0.1860 0.9228 0.0185 -16.750 -0.8525 0.02737 0.02253 -0.1856 0.9214 0.0185 -16.500 -0.8474 0.02596 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0.9859 13.250 1.1572 0.04061 0.03430 -0.0007 0.0491 0.9861 13.500 1.1682 0.04153 0.03517 0.0008 0.0423 0.9862 13.750 1.1796 0.04238 0.03601 0.0023 0.0374 0.9863 14.000 1.1878 0.04345 0.03705 0.0041 0.0317 0.9863 14.250 1.1999 0.04428 0.03792 0.0055 0.0312 0.9864 14.500 1.2085 0.04533 0.03896 0.0072 0.0274 0.9864 14.750 1.2188 0.04629 0.03994 0.0087 0.0258 0.9865 15.000 1.2266 0.04742 0.04109 0.0105 0.0239 0.9865 15.250 1.2361 0.04845 0.04215 0.0120 0.0222 0.9865 15.500 1.2473 0.04935 0.04309 0.0133 0.0225 0.9865 15.750 1.2543 0.05059 0.04437 0.0149 0.0211 0.9866 16.000 1.2612 0.05183 0.04563 0.0166 0.0193 0.9866 16.250 1.2692 0.05299 0.04682 0.0180 0.0182 0.9866 16.500 1.2778 0.05415 0.04804 0.0194 0.0187 0.9866 16.750 1.2836 0.05555 0.04946 0.0209 0.0172 0.9866 17.000 1.2885 0.05704 0.05099 0.0224 0.0166 0.9866 17.250 1.2968 0.05829 0.05229 0.0236 0.0158 0.9866 17.500 1.3024 0.05979 0.05384 0.0249 0.0152 0.9867 17.750 1.3086 0.06146 0.05555 0.0258 0.0145 0.9868 18.000 1.3147 0.06317 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