XFOIL Version 6.94 Calculated polar for: GOE 509 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.300 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.4526 0.01067 0.00324 -0.0732 0.6449 0.2328 0.500 0.4901 0.01078 0.00327 -0.0690 0.5901 0.3122 1.000 0.5255 0.01098 0.00330 -0.0646 0.5372 0.3597 1.500 0.5480 0.01041 0.00328 -0.0577 0.5055 0.5947 2.500 0.8601 0.01169 0.00475 -0.1018 0.4516 1.0000 3.000 0.9012 0.01215 0.00504 -0.0990 0.4392 1.0000 3.500 0.9429 0.01242 0.00532 -0.0963 0.4295 1.0000 4.000 0.9842 0.01295 0.00568 -0.0936 0.4162 1.0000 5.000 1.0459 0.01347 0.00614 -0.0837 0.3593 1.0000 5.500 1.0692 0.01399 0.00633 -0.0775 0.2978 1.0000 6.000 1.0877 0.01491 0.00690 -0.0706 0.2371 1.0000 6.500 1.0897 0.01648 0.00796 -0.0609 0.1668 1.0000 7.000 1.1110 0.01745 0.00880 -0.0549 0.1431 1.0000 7.500 1.1207 0.01910 0.01009 -0.0474 0.0805 1.0000 8.000 1.1237 0.02128 0.01205 -0.0394 0.0260 1.0000 8.500 1.1494 0.02239 0.01324 -0.0350 0.0220 1.0000 9.000 1.1710 0.02381 0.01468 -0.0305 0.0052 1.0000 9.500 1.1917 0.02542 0.01639 -0.0262 0.0039 1.0000 10.000 1.2113 0.02723 0.01838 -0.0222 0.0037 1.0000 10.500 1.2289 0.02937 0.02072 -0.0185 0.0038 1.0000 11.000 1.2440 0.03188 0.02343 -0.0152 0.0040 1.0000 11.500 1.2566 0.03483 0.02659 -0.0123 0.0041 1.0000 12.000 1.2650 0.03837 0.03036 -0.0097 0.0043 1.0000 12.500 1.2688 0.04265 0.03489 -0.0076 0.0045 1.0000 13.000 1.2666 0.04792 0.04041 -0.0061 0.0047 1.0000 13.500 1.2577 0.05443 0.04719 -0.0055 0.0048 1.0000 14.000 1.2476 0.06164 0.05467 -0.0058 0.0050 1.0000 14.500 1.2410 0.06895 0.06223 -0.0068 0.0054 1.0000 15.000 1.2238 0.07811 0.07165 -0.0087 0.0055 1.0000 15.500 1.2062 0.08752 0.08131 -0.0109 0.0057 1.0000 16.000 1.1849 0.09769 0.09171 -0.0136 0.0060 1.0000 16.500 1.1672 0.10749 0.10171 -0.0164 0.0061 1.0000 17.000 1.1529 0.11688 0.11127 -0.0192 0.0063 1.0000 17.500 1.1475 0.12463 0.11914 -0.0214 0.0066 1.0000 18.000 1.1559 0.12990 0.12459 -0.0222 0.0075 1.0000 18.500 1.1808 0.13055 0.12524 -0.0193 0.0085 1.0000