XFOIL Version 6.94 Calculated polar for: GOE 442 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 0.4758 0.00860 0.00179 -0.0908 0.6860 0.0585 0.500 0.5210 0.00875 0.00174 -0.0882 0.5965 0.1408 1.000 0.5950 0.00725 0.00194 -0.0926 0.5408 1.0000 1.500 0.6437 0.00762 0.00207 -0.0909 0.5093 1.0000 2.000 0.6930 0.00796 0.00223 -0.0894 0.4807 1.0000 2.500 0.7426 0.00828 0.00241 -0.0880 0.4539 1.0000 3.000 0.7922 0.00860 0.00259 -0.0866 0.4282 1.0000 3.500 0.8419 0.00893 0.00280 -0.0853 0.4050 1.0000 4.000 0.8874 0.00951 0.00300 -0.0834 0.3388 1.0000 4.500 0.9253 0.01092 0.00366 -0.0804 0.1935 1.0000 5.000 0.9656 0.01227 0.00452 -0.0778 0.1129 1.0000 5.500 1.0120 0.01299 0.00510 -0.0761 0.0875 1.0000 6.000 1.0478 0.01474 0.00641 -0.0727 0.0053 1.0000 6.500 1.0943 0.01537 0.00715 -0.0709 0.0058 1.0000 7.000 1.1398 0.01607 0.00798 -0.0688 0.0069 1.0000 7.500 1.1827 0.01698 0.00907 -0.0663 0.0071 1.0000 8.000 1.2207 0.01826 0.01059 -0.0629 0.0069 1.0000 8.500 1.2510 0.02006 0.01263 -0.0585 0.0066 1.0000 9.000 1.2659 0.02244 0.01520 -0.0518 0.0064 1.0000 9.500 1.2685 0.02500 0.01792 -0.0435 0.0063 1.0000 10.000 1.2658 0.02834 0.02136 -0.0357 0.0062 1.0000 10.500 1.2746 0.03136 0.02458 -0.0303 0.0060 1.0000 11.000 1.2831 0.03479 0.02826 -0.0258 0.0058 1.0000 11.500 1.2905 0.03888 0.03258 -0.0220 0.0056 1.0000 12.000 1.3011 0.04477 0.03858 -0.0194 0.0045 1.0000 12.500 1.2814 0.05040 0.04471 -0.0176 0.0038 1.0000 13.000 1.1843 0.05873 0.05376 -0.0199 0.0043 1.0000 13.500 1.1591 0.07077 0.06620 -0.0237 0.0041 1.0000 14.000 1.1322 0.08379 0.07959 -0.0291 0.0039 1.0000 14.500 1.1061 0.09675 0.09291 -0.0355 0.0039 1.0000 15.000 1.0743 0.10964 0.10616 -0.0436 0.0040 1.0000 15.500 1.0438 0.12179 0.11864 -0.0522 0.0042 1.0000 16.000 1.0230 0.13288 0.12997 -0.0595 0.0045 1.0000 16.500 0.9934 0.14845 0.14597 -0.0710 0.0057 1.0000 18.000 0.8694 0.22556 0.22375 -0.1099 0.0282 1.0000 18.500 0.8684 0.23694 0.23512 -0.1159 0.0251 1.0000 19.000 0.8737 0.24650 0.24468 -0.1202 0.0220 1.0000 19.500 0.8803 0.25557 0.25375 -0.1242 0.0188 1.0000 20.000 0.8876 0.26434 0.26254 -0.1281 0.0155 1.0000 20.500 0.8939 0.27398 0.27219 -0.1321 0.0134 1.0000 21.000 0.9009 0.28331 0.28153 -0.1359 0.0110 1.0000 21.500 0.9078 0.29281 0.29104 -0.1397 0.0091 1.0000 22.000 0.9147 0.30229 0.30053 -0.1434 0.0073 1.0000 22.500 0.9240 0.30893 0.30721 -0.1460 0.0061 1.0000 23.000 0.9283 0.32188 0.32016 -0.1506 0.0057 1.0000 23.500 0.9364 0.32903 0.32733 -0.1534 0.0046 1.0000 24.000 0.9419 0.34043 0.33875 -0.1573 0.0045 1.0000 24.500 0.9476 0.35254 0.35087 -0.1612 0.0041 1.0000 25.000 0.9539 0.36275 0.36109 -0.1646 0.0037 1.0000 25.500 0.9600 0.37069 0.36906 -0.1675 0.0032 1.0000 26.000 0.9652 0.38204 0.38042 -0.1711 0.0031 1.0000 26.500 0.9706 0.39365 0.39204 -0.1746 0.0030 1.0000 27.000 0.9759 0.40459 0.40301 -0.1779 0.0029 1.0000 27.500 0.9809 0.41524 0.41368 -0.1811 0.0029 1.0000