XFOIL Version 6.94 Calculated polar for: EPPLER 604 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.5115 0.01155 0.00593 -0.1134 0.6753 0.6929 0.500 0.5687 0.01154 0.00588 -0.1141 0.6691 0.6963 1.000 0.6283 0.01167 0.00591 -0.1155 0.6629 0.7005 1.500 0.6829 0.01171 0.00598 -0.1159 0.6570 0.7047 2.000 0.7389 0.01169 0.00595 -0.1165 0.6501 0.7081 2.500 0.7957 0.01167 0.00592 -0.1172 0.6434 0.7115 3.000 0.8495 0.01180 0.00611 -0.1172 0.6361 0.7151 3.500 0.9017 0.01180 0.00617 -0.1169 0.6274 0.7195 4.000 0.9602 0.01191 0.00622 -0.1180 0.6185 0.7238 4.500 1.0076 0.01192 0.00633 -0.1169 0.6090 0.7277 5.000 1.0595 0.01188 0.00631 -0.1166 0.5990 0.7318 5.500 1.1042 0.01196 0.00653 -0.1148 0.5878 0.7360 6.000 1.1523 0.01206 0.00663 -0.1137 0.5752 0.7412 6.500 1.1888 0.01215 0.00685 -0.1104 0.5610 0.7465 7.000 1.2208 0.01225 0.00702 -0.1061 0.5454 0.7513 7.500 1.2467 0.01253 0.00736 -0.1008 0.5270 0.7565 8.000 1.2709 0.01302 0.00788 -0.0954 0.5054 0.7625 8.500 1.2911 0.01373 0.00860 -0.0897 0.4801 0.7686 9.000 1.3033 0.01469 0.00958 -0.0829 0.4510 0.7755 9.500 1.3092 0.01615 0.01098 -0.0758 0.4187 0.7829 10.000 1.3099 0.01814 0.01289 -0.0687 0.3816 0.7902 10.500 1.3079 0.02053 0.01524 -0.0619 0.3445 0.7982 11.000 1.3038 0.02349 0.01809 -0.0558 0.3075 0.8069 11.500 1.2995 0.02678 0.02130 -0.0504 0.2724 0.8161 12.000 1.3017 0.03005 0.02454 -0.0462 0.2402 0.8270 12.500 1.3011 0.03374 0.02820 -0.0424 0.2110 0.8385 13.000 1.3058 0.03737 0.03182 -0.0395 0.1832 0.8521 14.000 1.3091 0.04517 0.03973 -0.0337 0.1366 0.9098