XFOIL Version 6.94 Calculated polar for: GOA 3 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.008 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1522 0.03763 0.01800 -0.0252 0.9999 1.0001 -2.750 -0.1277 0.03750 0.01692 -0.0246 0.9999 1.0001 -2.500 -0.1037 0.03741 0.01602 -0.0239 0.9999 1.0001 -2.250 -0.0800 0.03734 0.01527 -0.0231 0.9999 1.0001 -2.000 -0.0564 0.03731 0.01466 -0.0223 0.9999 1.0001 -1.750 -0.0329 0.03730 0.01409 -0.0214 0.9999 1.0001 -1.500 -0.0095 0.03732 0.01371 -0.0206 0.9999 1.0001 -1.250 0.0139 0.03737 0.01343 -0.0197 0.9999 1.0001 -1.000 0.0373 0.03744 0.01327 -0.0188 0.9999 1.0001 -0.750 0.0608 0.03754 0.01323 -0.0179 0.9999 1.0001 -0.500 0.0844 0.03765 0.01329 -0.0170 0.9999 1.0001 -0.250 0.1081 0.03779 0.01342 -0.0161 0.9999 1.0001 0.000 0.1322 0.03794 0.01373 -0.0152 0.9999 1.0001 0.250 0.1568 0.03811 0.01420 -0.0143 0.9999 1.0001 0.500 0.1824 0.03828 0.01484 -0.0135 0.9999 1.0001 0.750 0.2100 0.03843 0.01570 -0.0128 0.9999 1.0001 1.000 0.2414 0.03855 0.01685 -0.0127 0.9999 1.0001 1.250 0.2760 0.03879 0.01837 -0.0142 0.9999 1.0001 1.500 0.3020 0.03983 0.02061 -0.0173 0.9999 1.0001 1.750 0.3046 0.04248 0.02360 -0.0205 0.9999 1.0001 2.000 0.3005 0.04572 0.02659 -0.0232 0.9999 1.0001 2.250 0.4829 0.05315 0.03424 -0.0572 0.7633 1.0001 2.500 0.5230 0.05592 0.03694 -0.0606 0.7200 1.0001 2.750 0.5564 0.05899 0.03992 -0.0636 0.6973 1.0001 3.000 0.5703 0.06235 0.04319 -0.0654 0.6887 1.0001 3.250 0.5915 0.06586 0.04663 -0.0678 0.6816 1.0001 3.500 0.5984 0.06945 0.05010 -0.0692 0.6804 1.0001 3.750 0.6053 0.07308 0.05363 -0.0707 0.6809 1.0001 4.000 0.6050 0.07657 0.05700 -0.0713 0.6840 1.0001 4.250 0.6024 0.08000 0.06028 -0.0718 0.6889 1.0001 4.500 0.6061 0.08363 0.06379 -0.0730 0.6944 1.0001 4.750 0.6086 0.08722 0.06732 -0.0742 0.7012 1.0001 5.000 0.6071 0.09064 0.07061 -0.0749 0.7102 1.0001