XFOIL Version 6.96 Calculated polar for: GOE 403 AIRFOIL 1 1 Reynolds number fixed Mach number fixed xtrf = 1.000 (top) 1.000 (bottom) Mach = 0.000 Re = 0.500 e 6 Ncrit = 5.000 alpha CL CD CDp CM Top_Xtr Bot_Xtr ------ -------- --------- --------- -------- -------- -------- -8.750 -0.3537 0.09699 0.09479 -0.0244 1.0000 0.0041 -8.500 -0.3532 0.09338 0.09122 -0.0253 1.0000 0.0042 -8.000 -0.3502 0.08731 0.08521 -0.0263 1.0000 0.0041 -7.750 -0.3434 0.08667 0.08459 -0.0261 1.0000 0.0053 -7.500 -0.3462 0.08414 0.08211 -0.0257 1.0000 0.0056 -7.000 -0.3226 0.07540 0.07340 -0.0352 0.9862 0.0055 -6.750 -0.3008 0.06991 0.06791 -0.0431 0.9744 0.0054 -6.500 -0.2708 0.06275 0.06070 -0.0544 0.9613 0.0054 -6.000 -0.1996 0.01762 0.01371 -0.0971 0.9121 0.0048 -5.750 -0.1720 0.01543 0.01103 -0.0977 0.8923 0.0051 -5.500 -0.1455 0.01393 0.00913 -0.0978 0.8760 0.0055 -5.250 -0.1194 0.01282 0.00767 -0.0975 0.8619 0.0061 -5.000 -0.0932 0.01200 0.00655 -0.0972 0.8497 0.0066 -4.750 -0.0672 0.01111 0.00540 -0.0969 0.8392 0.0083 -4.500 -0.0408 0.01061 0.00472 -0.0965 0.8295 0.0097 -4.250 -0.0142 0.01008 0.00399 -0.0962 0.8203 0.0121 -4.000 0.0127 0.00982 0.00369 -0.0959 0.8118 0.0198 -3.750 0.0397 0.00974 0.00354 -0.0956 0.8032 0.0251 -3.500 0.0670 0.00968 0.00339 -0.0954 0.7948 0.0295 -3.250 0.0942 0.00975 0.00333 -0.0952 0.7864 0.0320 -3.000 0.1208 0.00942 0.00296 -0.0950 0.7774 0.0359 -2.750 0.1477 0.00925 0.00270 -0.0947 0.7682 0.0381 -2.500 0.1745 0.00908 0.00243 -0.0944 0.7583 0.0396 -2.250 0.2013 0.00892 0.00216 -0.0942 0.7473 0.0407 -2.000 0.2281 0.00878 0.00194 -0.0939 0.7353 0.0415 -1.750 0.2549 0.00866 0.00173 -0.0936 0.7216 0.0422 -1.500 0.2814 0.00854 0.00151 -0.0932 0.7057 0.0439 -1.250 0.3077 0.00845 0.00135 -0.0928 0.6871 0.0484 -1.000 0.3340 0.00840 0.00122 -0.0924 0.6650 0.0538 -0.750 0.3599 0.00837 0.00113 -0.0920 0.6408 0.0701 -0.500 0.3856 0.00839 0.00110 -0.0916 0.6159 0.0930 -0.250 0.4116 0.00844 0.00107 -0.0912 0.5927 0.1068 0.000 0.4376 0.00850 0.00107 -0.0908 0.5728 0.1184 0.250 0.4639 0.00855 0.00107 -0.0905 0.5559 0.1318 0.500 0.4902 0.00856 0.00110 -0.0903 0.5420 0.1571 1.000 0.5426 0.00849 0.00122 -0.0898 0.5184 0.2666 1.250 0.5674 0.00807 0.00134 -0.0895 0.5090 0.4882 1.500 0.6026 0.00705 0.00144 -0.0912 0.4996 1.0000 1.750 0.6289 0.00718 0.00151 -0.0909 0.4904 1.0000 2.000 0.6554 0.00730 0.00159 -0.0906 0.4810 1.0000 2.250 0.6814 0.00745 0.00167 -0.0903 0.4669 1.0000 2.500 0.7064 0.00768 0.00177 -0.0898 0.4386 1.0000 2.750 0.7319 0.00788 0.00189 -0.0894 0.4182 1.0000 3.000 0.7566 0.00817 0.00202 -0.0888 0.3825 1.0000 3.250 0.7798 0.00863 0.00221 -0.0881 0.3275 1.0000 3.500 0.7948 0.01012 0.00285 -0.0864 0.1600 1.0000 3.750 0.8120 0.01149 0.00364 -0.0848 0.0163 1.0000 4.000 0.8372 0.01181 0.00404 -0.0842 0.0116 1.0000 4.250 0.8623 0.01211 0.00441 -0.0837 0.0103 1.0000 4.500 0.8867 0.01251 0.00490 -0.0831 0.0090 1.0000 4.750 0.9104 0.01300 0.00547 -0.0823 0.0076 1.0000 5.000 0.9320 0.01375 0.00631 -0.0813 0.0065 1.0000 5.250 0.9544 0.01438 0.00702 -0.0803 0.0061 1.0000 5.500 0.9753 0.01516 0.00788 -0.0791 0.0056 1.0000 5.750 0.9950 0.01606 0.00886 -0.0777 0.0053 1.0000 6.000 1.0140 0.01699 0.00989 -0.0762 0.0050 1.0000 6.250 1.0341 0.01771 0.01063 -0.0751 0.0044 1.0000 6.500 1.0513 0.01884 0.01184 -0.0735 0.0039 1.0000 6.750 1.0688 0.02008 0.01318 -0.0717 0.0037 1.0000 7.000 1.0865 0.02154 0.01474 -0.0699 0.0034 1.0000 7.250 1.1057 0.02316 0.01645 -0.0683 0.0033 1.0000 7.500 1.1265 0.02498 0.01839 -0.0671 0.0031 1.0000 7.750 1.1483 0.02708 0.02063 -0.0659 0.0031 1.0000 8.000 1.1698 0.02942 0.02321 -0.0647 0.0031 1.0000 8.250 1.1894 0.03196 0.02598 -0.0633 0.0032 1.0000 8.500 1.2063 0.03475 0.02903 -0.0616 0.0034 1.0000 8.750 1.2199 0.03767 0.03222 -0.0596 0.0035 1.0000 9.000 1.2303 0.04066 0.03546 -0.0575 0.0037 1.0000 9.250 1.2365 0.04390 0.03894 -0.0552 0.0038 1.0000 16.250 0.8242 0.19424 0.19248 -0.0982 0.0058 1.0000 16.500 0.8268 0.19841 0.19666 -0.1002 0.0049 1.0000