XFOIL Version 6.94 Calculated polar for: GOE 546 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 0.4527 0.00631 0.00226 -0.1004 0.7703 0.9415 0.500 0.5454 0.00634 0.00214 -0.1083 0.7450 1.0000 1.000 0.5917 0.00649 0.00210 -0.1061 0.7163 1.0000 1.500 0.6373 0.00671 0.00210 -0.1037 0.6825 1.0000 2.000 0.6828 0.00700 0.00219 -0.1013 0.6482 1.0000 2.500 0.7297 0.00733 0.00238 -0.0993 0.6195 1.0000 3.000 0.7676 0.00784 0.00252 -0.0954 0.5369 1.0000 3.500 0.8044 0.00859 0.00280 -0.0915 0.4439 1.0000 4.000 0.8219 0.01079 0.00371 -0.0848 0.2139 1.0000 4.500 0.8508 0.01253 0.00468 -0.0802 0.0731 1.0000 5.000 0.8890 0.01358 0.00548 -0.0771 0.0276 1.0000 5.500 0.9332 0.01416 0.00604 -0.0750 0.0230 1.0000 6.000 0.9767 0.01474 0.00664 -0.0728 0.0195 1.0000 6.500 1.0147 0.01563 0.00745 -0.0697 0.0035 1.0000 7.000 1.0519 0.01638 0.00828 -0.0663 0.0033 1.0000 7.500 1.0857 0.01720 0.00921 -0.0624 0.0033 1.0000 8.000 1.1175 0.01814 0.01027 -0.0582 0.0035 1.0000 8.500 1.1482 0.01918 0.01144 -0.0541 0.0037 1.0000 9.000 1.1763 0.02040 0.01281 -0.0498 0.0040 1.0000 9.500 1.2015 0.02185 0.01442 -0.0454 0.0042 1.0000 10.000 1.2239 0.02357 0.01633 -0.0410 0.0044 1.0000 10.500 1.2480 0.02526 0.01818 -0.0371 0.0048 1.0000 11.000 1.2687 0.02731 0.02042 -0.0332 0.0053 1.0000 11.500 1.2806 0.03014 0.02345 -0.0291 0.0057 1.0000 12.000 1.2874 0.03360 0.02711 -0.0252 0.0060 1.0000 12.500 1.3000 0.03684 0.03058 -0.0223 0.0066 1.0000 13.000 1.2990 0.04165 0.03565 -0.0194 0.0072 1.0000 14.000 1.2969 0.05281 0.04741 -0.0156 0.0087 1.0000 14.500 1.2851 0.06093 0.05597 -0.0139 0.0097 1.0000 15.500 1.2020 0.08852 0.08497 -0.0199 0.0112 1.0000 16.000 1.1608 0.10396 0.10082 -0.0278 0.0113 1.0000 16.500 1.1296 0.11914 0.11629 -0.0370 0.0109 1.0000 17.000 1.0997 0.13568 0.13307 -0.0475 0.0107 1.0000 17.500 1.0776 0.15151 0.14905 -0.0576 0.0104 1.0000 18.000 1.0555 0.16817 0.16581 -0.0680 0.0100 1.0000