XFOIL Version 6.94 Calculated polar for: GOA 2 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.026 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1630 0.02866 0.01519 -0.0324 1.0000 0.1833 -2.750 -0.1363 0.02665 0.01342 -0.0312 1.0000 0.2200 -2.500 -0.1119 0.02324 0.01147 -0.0299 1.0000 0.3546 -2.250 -0.0859 0.02037 0.00973 -0.0257 1.0000 1.0000 -2.000 -0.0609 0.02038 0.00909 -0.0245 1.0000 1.0000 -1.750 -0.0363 0.02041 0.00870 -0.0233 1.0000 1.0000 -1.500 -0.0119 0.02046 0.00839 -0.0223 1.0000 1.0000 -1.250 0.0124 0.02052 0.00816 -0.0213 1.0000 1.0000 -1.000 0.0368 0.02059 0.00802 -0.0204 1.0000 1.0000 -0.750 0.0613 0.02067 0.00800 -0.0194 1.0000 1.0000 -0.500 0.0862 0.02077 0.00810 -0.0184 1.0000 1.0000 -0.250 0.1121 0.02086 0.00843 -0.0173 1.0000 1.0000 0.000 0.1414 0.02088 0.00902 -0.0162 1.0000 1.0000 0.250 0.2291 0.02457 0.00981 -0.0236 0.3414 1.0000 0.500 0.2551 0.02585 0.01070 -0.0226 0.3194 1.0000 0.750 0.2832 0.02713 0.01167 -0.0219 0.3034 1.0000 1.000 0.3124 0.02846 0.01277 -0.0215 0.2865 1.0000 1.250 0.3427 0.02984 0.01401 -0.0214 0.2716 1.0000 1.500 0.3752 0.03122 0.01546 -0.0215 0.2642 1.0000 1.750 0.4088 0.03273 0.01714 -0.0218 0.2625 1.0000 2.000 0.4425 0.03431 0.01904 -0.0223 0.2629 1.0000 2.250 0.4758 0.03601 0.02119 -0.0229 0.2653 1.0000 2.500 0.5078 0.03795 0.02354 -0.0234 0.2695 1.0000 2.750 0.5389 0.04023 0.02613 -0.0241 0.2743 1.0000 3.000 0.5712 0.04246 0.02902 -0.0254 0.2833 1.0000 3.250 0.6008 0.04534 0.03217 -0.0264 0.2918 1.0000 3.500 0.6324 0.04823 0.03572 -0.0287 0.3070 1.0000 3.750 0.6629 0.05162 0.03959 -0.0314 0.3252 1.0000 4.000 0.6942 0.05542 0.04410 -0.0370 0.3553 1.0000 4.250 0.7242 0.06000 0.04917 -0.0436 0.3925 1.0000 4.500 0.7433 0.06575 0.05550 -0.0575 0.4540 1.0000 4.750 0.7382 0.07258 0.06250 -0.0734 0.5336 1.0000 5.000 0.7049 0.07834 0.06809 -0.0838 0.6158 1.0000