XFOIL Version 6.94 Calculated polar for: GOE 372 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 ------- -------- --------- --------- -------- ------- ------- 0.500 0.3894 0.01101 0.00330 -0.0817 0.6308 0.1232 1.000 0.4474 0.01083 0.00337 -0.0816 0.6367 0.1649 1.500 0.4965 0.01060 0.00308 -0.0794 0.5793 0.2184 2.000 0.5480 0.01014 0.00325 -0.0782 0.5558 0.4331 2.500 0.6160 0.00928 0.00327 -0.0797 0.4877 1.0000 3.000 0.6759 0.00943 0.00358 -0.0796 0.5195 1.0000 3.500 0.7187 0.01005 0.00357 -0.0767 0.4266 1.0000 4.000 0.7754 0.01024 0.00375 -0.0758 0.4330 1.0000 4.500 0.8288 0.01050 0.00397 -0.0747 0.4204 1.0000 5.000 0.8519 0.01331 0.00499 -0.0700 0.1420 1.0000 6.000 0.9853 0.01164 0.00516 -0.0707 0.3679 1.0000 6.500 1.0202 0.01327 0.00595 -0.0671 0.2316 1.0000 7.000 1.0581 0.01490 0.00696 -0.0644 0.1595 1.0000 8.500 1.1040 0.00116 -0.00554 -0.0454 0.1694 1.0000 9.000 1.1308 0.00159 -0.00548 -0.0415 0.1181 1.0000 9.500 1.1627 0.00196 -0.00482 -0.0378 0.1271 1.0000 10.000 1.1852 0.00258 -0.00425 -0.0337 0.1093 1.0000 10.500 1.1968 0.00361 -0.00323 -0.0288 0.0930 1.0000 11.000 1.2133 0.00502 -0.00184 -0.0256 0.0712 1.0000 11.500 1.2274 0.00675 0.00004 -0.0222 0.0647 1.0000 12.000 1.2272 0.00998 0.00330 -0.0186 0.0497 1.0000 12.500 1.2169 0.01518 0.00857 -0.0154 0.0423 1.0000 13.000 1.2074 0.02100 0.01452 -0.0127 0.0382 1.0000 13.500 1.1929 0.02864 0.02234 -0.0110 0.0360 1.0000 14.000 1.1712 0.03726 0.03119 -0.0095 0.0361 1.0000 14.500 1.1433 0.04686 0.04102 -0.0089 0.0338 1.0000 15.000 1.1101 0.05633 0.05072 -0.0086 0.0324 1.0000