XFOIL Version 6.94 Calculated polar for: GOE 496 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.080 Re = 0.450 e 6 Ncrit = 9.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------- -------- --------- --------- -------- ------- ------- 0.000 0.6251 0.00819 0.00226 -0.1454 0.7268 0.3774 0.500 0.6787 0.00820 0.00231 -0.1446 0.7044 0.4369 1.000 0.7321 0.00821 0.00239 -0.1439 0.6820 0.4975 1.500 0.7842 0.00809 0.00256 -0.1429 0.6580 0.6124 2.500 0.8875 0.00780 0.00282 -0.1404 0.6131 1.0000 3.000 0.9346 0.00823 0.00294 -0.1383 0.5564 1.0000 3.500 0.9857 0.00861 0.00321 -0.1372 0.5264 1.0000 4.000 1.0315 0.00925 0.00352 -0.1351 0.4616 1.0000 4.500 1.0698 0.01049 0.00410 -0.1319 0.3419 1.0000 5.000 1.1054 0.01214 0.00506 -0.1286 0.2254 1.0000 5.500 1.1466 0.01331 0.00586 -0.1262 0.1628 1.0000 6.000 1.1690 0.01593 0.00761 -0.1209 0.0218 1.0000 6.500 1.2098 0.01695 0.00851 -0.1183 0.0062 1.0000 7.000 1.2517 0.01778 0.00945 -0.1158 0.0042 1.0000 7.500 1.2883 0.01876 0.01056 -0.1123 0.0038 1.0000 8.000 1.3210 0.01992 0.01190 -0.1083 0.0037 1.0000 8.500 1.3507 0.02128 0.01349 -0.1040 0.0037 1.0000 9.000 1.3740 0.02311 0.01556 -0.0991 0.0036 1.0000 9.500 1.3909 0.02548 0.01817 -0.0937 0.0037 1.0000 10.000 1.4005 0.02852 0.02146 -0.0883 0.0038 1.0000 10.500 1.4004 0.03270 0.02591 -0.0828 0.0039 1.0000 11.000 1.4047 0.03693 0.03038 -0.0788 0.0041 1.0000 11.500 1.4145 0.04097 0.03461 -0.0759 0.0044 1.0000 12.000 1.4228 0.04547 0.03932 -0.0735 0.0047 1.0000 12.500 1.4288 0.05056 0.04471 -0.0715 0.0051 1.0000 13.000 1.4308 0.05673 0.05126 -0.0693 0.0061 1.0000 13.500 1.4270 0.06445 0.05942 -0.0673 0.0072 1.0000 14.000 1.4084 0.07510 0.07057 -0.0666 0.0081 1.0000