XFOIL Version 6.96 Calculated polar for: EPPLER E1212 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.100 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -5.000 -0.1878 0.02931 0.02149 -0.0704 0.7578 0.2880 -4.500 -0.1402 0.02818 0.02032 -0.0686 0.7278 0.2970 -4.000 -0.0937 0.02688 0.01843 -0.0673 0.7012 0.3087 -3.500 -0.0433 0.02594 0.01743 -0.0658 0.6766 0.3186 -3.000 0.0072 0.02501 0.01604 -0.0648 0.6537 0.3303 -2.500 0.0591 0.02440 0.01532 -0.0636 0.6317 0.3415 -2.000 0.1113 0.02379 0.01445 -0.0626 0.6108 0.3538 -1.500 0.1640 0.02343 0.01392 -0.0616 0.5905 0.3672 -1.000 0.2165 0.02307 0.01352 -0.0606 0.5710 0.3813 -0.500 0.2692 0.02284 0.01320 -0.0597 0.5522 0.3973 0.000 0.3217 0.02277 0.01311 -0.0587 0.5344 0.4167 0.500 0.3740 0.02274 0.01313 -0.0577 0.5175 0.4395 1.000 0.4263 0.02276 0.01323 -0.0566 0.5022 0.4705 1.500 0.4804 0.02266 0.01318 -0.0556 0.4882 0.5157 2.000 0.5271 0.02268 0.01372 -0.0536 0.4726 0.5916 2.500 0.5677 0.02251 0.01442 -0.0494 0.4587 0.7774 3.000 0.6652 0.02290 0.01482 -0.0565 0.4428 1.0000 3.500 0.7147 0.02355 0.01512 -0.0553 0.4310 1.0000 4.000 0.7558 0.02483 0.01640 -0.0532 0.4179 1.0000 4.500 0.8043 0.02583 0.01722 -0.0519 0.4070 1.0000 5.000 0.8499 0.02694 0.01827 -0.0504 0.3956 1.0000 5.500 0.8905 0.02858 0.01995 -0.0485 0.3847 1.0000 6.000 0.9423 0.02942 0.02062 -0.0476 0.3751 1.0000 6.500 0.9717 0.03187 0.02331 -0.0449 0.3647 1.0000 7.000 1.0260 0.03269 0.02394 -0.0443 0.3564 1.0000 7.500 1.0401 0.03609 0.02770 -0.0404 0.3462 1.0000 8.000 1.0918 0.03719 0.02869 -0.0397 0.3391 1.0000 8.500 1.0597 0.04361 0.03561 -0.0331 0.3296 1.0000 9.000 1.1006 0.04518 0.03716 -0.0316 0.3226 1.0000