XFOIL Version 6.94 Calculated polar for: GOE 596 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 0.4158 0.01114 0.00315 -0.0796 0.6583 0.0605 1.000 0.5601 0.00852 0.00281 -0.0865 0.6200 1.0000 1.500 0.6099 0.00883 0.00290 -0.0851 0.6041 1.0000 2.000 0.6597 0.00912 0.00306 -0.0837 0.5882 1.0000 2.500 0.7096 0.00943 0.00325 -0.0823 0.5724 1.0000 3.000 0.7580 0.00969 0.00342 -0.0806 0.5504 1.0000 3.500 0.8061 0.00993 0.00359 -0.0789 0.5268 1.0000 4.000 0.8548 0.01019 0.00384 -0.0773 0.5044 1.0000 4.500 0.9026 0.01046 0.00406 -0.0755 0.4768 1.0000 5.000 0.9503 0.01076 0.00434 -0.0738 0.4469 1.0000 5.500 0.9916 0.01132 0.00460 -0.0711 0.3789 1.0000 6.000 1.0301 0.01234 0.00526 -0.0681 0.3050 1.0000 6.500 1.0508 0.01488 0.00669 -0.0630 0.1285 1.0000 7.000 1.0741 0.01721 0.00839 -0.0583 0.0374 1.0000 7.500 1.1149 0.01805 0.00928 -0.0560 0.0314 1.0000 8.000 1.1528 0.01901 0.01019 -0.0534 0.0189 1.0000 8.500 1.1827 0.02032 0.01143 -0.0495 0.0058 1.0000 9.000 1.2113 0.02150 0.01272 -0.0454 0.0046 1.0000 9.500 1.2355 0.02300 0.01438 -0.0410 0.0041 1.0000 10.000 1.2570 0.02475 0.01634 -0.0368 0.0038 1.0000 10.500 1.2769 0.02676 0.01854 -0.0331 0.0037 1.0000 11.000 1.2921 0.02933 0.02143 -0.0295 0.0036 1.0000 11.500 1.3048 0.03237 0.02472 -0.0265 0.0036 1.0000 12.000 1.3130 0.03614 0.02876 -0.0241 0.0036 1.0000 12.500 1.3175 0.04067 0.03357 -0.0224 0.0036 1.0000 13.000 1.3196 0.04584 0.03902 -0.0215 0.0036 1.0000 13.500 1.3152 0.05222 0.04570 -0.0214 0.0037 1.0000 14.000 1.3120 0.05874 0.05249 -0.0220 0.0037 1.0000 14.500 1.3042 0.06627 0.06031 -0.0233 0.0037 1.0000 15.000 1.2937 0.07463 0.06899 -0.0254 0.0038 1.0000 15.500 1.2814 0.08372 0.07837 -0.0281 0.0038 1.0000 16.000 1.2690 0.09330 0.08823 -0.0315 0.0039 1.0000 16.500 1.2545 0.10375 0.09897 -0.0357 0.0040 1.0000 17.000 1.2411 0.11446 0.10995 -0.0404 0.0040 1.0000 17.500 1.2260 0.12585 0.12163 -0.0460 0.0041 1.0000 18.000 1.2092 0.13808 0.13414 -0.0525 0.0041 1.0000 18.500 1.1907 0.15126 0.14761 -0.0600 0.0043 1.0000 19.000 1.1725 0.16510 0.16173 -0.0684 0.0043 1.0000 19.500 1.1481 0.18161 0.17850 -0.0787 0.0044 1.0000 20.000 1.1012 0.20786 0.20502 -0.0941 0.0046 1.0000 20.500 1.0635 0.23702 0.23409 -0.1087 0.0056 1.0000 21.000 1.0690 0.24807 0.24509 -0.1147 0.0062 1.0000 21.500 1.0761 0.26233 0.25930 -0.1211 0.0094 1.0000