XFOIL Version 6.94 Calculated polar for: GOE 547 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 0.5382 0.00833 0.00362 -0.1286 0.8180 0.9873 0.500 0.5997 0.00839 0.00352 -0.1298 0.8026 1.0000 1.000 0.6471 0.00843 0.00340 -0.1277 0.7821 1.0000 1.500 0.6942 0.00847 0.00331 -0.1254 0.7597 1.0000 2.000 0.7376 0.00852 0.00322 -0.1223 0.7300 1.0000 2.500 0.7749 0.00862 0.00307 -0.1176 0.6776 1.0000 3.000 0.8182 0.00890 0.00318 -0.1147 0.6434 1.0000 3.500 0.8500 0.00944 0.00332 -0.1093 0.5660 1.0000 4.000 0.8855 0.01001 0.00366 -0.1050 0.4961 1.0000 4.500 0.9118 0.01103 0.00420 -0.0992 0.3948 1.0000 5.000 0.9115 0.01363 0.00550 -0.0892 0.1799 1.0000 5.500 0.9284 0.01541 0.00660 -0.0824 0.0617 1.0000 6.000 0.9542 0.01681 0.00770 -0.0771 0.0099 1.0000 6.500 0.9905 0.01765 0.00864 -0.0736 0.0088 1.0000 7.000 1.0261 0.01857 0.00972 -0.0701 0.0082 1.0000 7.500 1.0599 0.01963 0.01105 -0.0664 0.0081 1.0000 8.000 1.0905 0.02095 0.01257 -0.0623 0.0082 1.0000 8.500 1.1181 0.02251 0.01433 -0.0582 0.0083 1.0000 9.000 1.1393 0.02458 0.01665 -0.0534 0.0085 1.0000 9.500 1.1543 0.02721 0.01951 -0.0482 0.0087 1.0000 10.000 1.1655 0.03042 0.02295 -0.0431 0.0091 1.0000 10.500 1.1790 0.03396 0.02678 -0.0386 0.0094 1.0000 11.000 1.1983 0.03803 0.03112 -0.0348 0.0100 1.0000 11.500 1.2265 0.04368 0.03719 -0.0318 0.0106 1.0000 12.000 1.2436 0.04686 0.04061 -0.0286 0.0112 1.0000 12.500 1.2399 0.05394 0.04824 -0.0249 0.0109 1.0000 13.000 1.2389 0.05886 0.05352 -0.0222 0.0114 1.0000 13.500 1.2038 0.07015 0.06569 -0.0197 0.0136 1.0000 14.000 1.1673 0.08070 0.07668 -0.0207 0.0137 1.0000 14.500 1.1124 0.09640 0.09285 -0.0256 0.0160 1.0000 15.000 1.0801 0.10957 0.10634 -0.0332 0.0145 1.0000 15.500 1.0346 0.12917 0.12626 -0.0458 0.0148 1.0000 16.000 0.9781 0.15894 0.15621 -0.0644 0.0153 1.0000 16.500 0.9728 0.17149 0.16875 -0.0704 0.0181 1.0000