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 = 0.035 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- -3.000 -0.1605 0.02598 0.01376 -0.0321 1.0000 0.1207 -2.750 -0.1331 0.02453 0.01211 -0.0307 1.0000 0.1305 -2.500 -0.1073 0.02297 0.01071 -0.0295 1.0000 0.1515 -2.250 -0.0814 0.01997 0.00922 -0.0292 1.0000 0.3260 -1.750 -0.0348 0.01774 0.00799 -0.0240 1.0000 1.0000 -1.500 -0.0100 0.01781 0.00778 -0.0229 1.0000 1.0000 -1.250 0.0146 0.01788 0.00763 -0.0218 1.0000 1.0000 -1.000 0.0394 0.01796 0.00756 -0.0208 1.0000 1.0000 -0.750 0.0646 0.01805 0.00763 -0.0198 1.0000 1.0000 -0.500 0.0909 0.01812 0.00787 -0.0186 1.0000 1.0000 -0.250 0.1786 0.02119 0.00808 -0.0263 0.2862 1.0000 0.000 0.2026 0.02227 0.00870 -0.0251 0.2635 1.0000 0.250 0.2275 0.02339 0.00939 -0.0242 0.2435 1.0000 0.500 0.2542 0.02446 0.01013 -0.0235 0.2268 1.0000 0.750 0.2839 0.02541 0.01103 -0.0230 0.2197 1.0000 1.000 0.3153 0.02652 0.01210 -0.0228 0.2160 1.0000 1.250 0.3479 0.02773 0.01336 -0.0229 0.2142 1.0000 1.500 0.3809 0.02901 0.01481 -0.0230 0.2142 1.0000 1.750 0.4131 0.03041 0.01647 -0.0230 0.2159 1.0000 2.000 0.4447 0.03201 0.01835 -0.0231 0.2192 1.0000 2.250 0.4748 0.03388 0.02048 -0.0232 0.2234 1.0000 2.500 0.5051 0.03578 0.02271 -0.0233 0.2289 1.0000 2.750 0.5364 0.03798 0.02541 -0.0239 0.2388 1.0000 3.000 0.5680 0.04036 0.02835 -0.0248 0.2518 1.0000 3.250 0.5997 0.04317 0.03170 -0.0263 0.2698 1.0000 3.500 0.6307 0.04656 0.03551 -0.0279 0.2917 1.0000 3.750 0.6653 0.05008 0.03975 -0.0321 0.3246 1.0000 4.000 0.7060 0.05470 0.04529 -0.0423 0.3897 1.0000 4.500 0.6134 0.07093 0.06195 -0.0889 0.7510 1.0000 4.750 0.5149 0.07081 0.06156 -0.0809 0.9025 1.0000 5.000 0.4284 0.06715 0.05764 -0.0621 1.0000 1.0000