GOE 517 AIRFOIL (goe517-il) Xfoil prediction polar at RE=500,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 517 AIRFOIL (goe517-il) Reynolds number: 500,000 Max Cl/Cd: 102.01 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe517-il-500000-n5.txt Download as CSV file: xf-goe517-il-500000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 517 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 -0.3142 0.09911 0.09691 -0.0296 1.0000 0.0088
-8.500 -0.3130 0.09640 0.09423 -0.0300 1.0000 0.0088
-8.250 -0.3136 0.09375 0.09163 -0.0301 1.0000 0.0089
-7.750 -0.2896 0.08597 0.08387 -0.0378 0.9870 0.0089
-7.500 -0.2757 0.08185 0.07976 -0.0426 0.9767 0.0089
-7.250 -0.2553 0.07728 0.07517 -0.0490 0.9672 0.0089
-7.000 -0.2302 0.07225 0.07012 -0.0567 0.9576 0.0089
-6.500 -0.1791 0.06181 0.05957 -0.0711 0.9331 0.0063
-6.250 -0.1529 0.05633 0.05398 -0.0785 0.9136 0.0063
-6.000 -0.1310 0.05162 0.04914 -0.0834 0.8934 0.0063
-5.750 -0.1107 0.04858 0.04597 -0.0863 0.8767 0.0070
-5.500 -0.0906 0.04483 0.04209 -0.0889 0.8629 0.0071
-5.250 -0.0695 0.04049 0.03757 -0.0913 0.8521 0.0073
-5.000 -0.0610 0.01847 0.01409 -0.0944 0.8412 0.0065
-4.750 -0.0398 0.01550 0.01054 -0.0937 0.8335 0.0067
-4.500 -0.0152 0.01430 0.00905 -0.0932 0.8267 0.0070
-4.250 0.0099 0.01326 0.00774 -0.0926 0.8197 0.0074
-4.000 0.0354 0.01234 0.00654 -0.0921 0.8130 0.0079
-3.750 0.0615 0.01169 0.00567 -0.0916 0.8065 0.0087
-3.500 0.0873 0.01103 0.00488 -0.0912 0.7994 0.0096
-3.250 0.1134 0.01057 0.00431 -0.0907 0.7924 0.0104
-3.000 0.1396 0.01012 0.00374 -0.0903 0.7846 0.0113
-2.750 0.1657 0.00972 0.00322 -0.0898 0.7763 0.0126
-2.500 0.1917 0.00946 0.00291 -0.0893 0.7657 0.0148
-2.250 0.2178 0.00920 0.00254 -0.0888 0.7542 0.0167
-2.000 0.2436 0.00893 0.00219 -0.0883 0.7422 0.0213
-1.750 0.2696 0.00873 0.00191 -0.0877 0.7291 0.0287
-1.500 0.2951 0.00854 0.00175 -0.0872 0.7130 0.0502
-1.250 0.3209 0.00849 0.00164 -0.0867 0.6947 0.0676
-1.000 0.3464 0.00848 0.00156 -0.0861 0.6732 0.0790
-0.750 0.3716 0.00853 0.00148 -0.0855 0.6468 0.0881
-0.500 0.3963 0.00859 0.00143 -0.0848 0.6189 0.0980
-0.250 0.4215 0.00865 0.00137 -0.0842 0.5957 0.1059
0.000 0.4469 0.00870 0.00134 -0.0837 0.5778 0.1134
0.250 0.4725 0.00874 0.00132 -0.0833 0.5632 0.1229
0.500 0.4984 0.00877 0.00132 -0.0829 0.5518 0.1345
1.000 0.5501 0.00880 0.00136 -0.0821 0.5312 0.1809
1.250 0.5756 0.00875 0.00142 -0.0817 0.5225 0.2387
1.750 0.6632 0.00731 0.00165 -0.0895 0.5041 1.0000
2.000 0.6880 0.00744 0.00171 -0.0888 0.4949 1.0000
2.250 0.7130 0.00754 0.00178 -0.0882 0.4861 1.0000
2.500 0.7377 0.00767 0.00186 -0.0876 0.4772 1.0000
2.750 0.7628 0.00778 0.00194 -0.0870 0.4670 1.0000
3.000 0.7875 0.00791 0.00204 -0.0864 0.4548 1.0000
3.250 0.8119 0.00806 0.00214 -0.0857 0.4400 1.0000
3.500 0.8362 0.00822 0.00225 -0.0851 0.4226 1.0000
3.750 0.8599 0.00843 0.00239 -0.0843 0.4011 1.0000
4.000 0.8830 0.00868 0.00255 -0.0834 0.3775 1.0000
4.250 0.9060 0.00896 0.00274 -0.0826 0.3563 1.0000
4.500 0.9292 0.00924 0.00296 -0.0817 0.3393 1.0000
4.750 0.9525 0.00951 0.00319 -0.0810 0.3242 1.0000
5.000 0.9760 0.00977 0.00343 -0.0802 0.3126 1.0000
5.500 1.0222 0.01035 0.00395 -0.0786 0.2841 1.0000
5.750 1.0444 0.01072 0.00423 -0.0777 0.2596 1.0000
6.000 1.0661 0.01113 0.00454 -0.0768 0.2319 1.0000
6.250 1.0868 0.01163 0.00490 -0.0757 0.1986 1.0000
6.500 1.1004 0.01276 0.00557 -0.0736 0.1181 1.0000
7.000 1.1274 0.01507 0.00731 -0.0692 0.0157 1.0000
7.250 1.1472 0.01563 0.00793 -0.0679 0.0120 1.0000
7.500 1.1667 0.01619 0.00854 -0.0666 0.0097 1.0000
7.750 1.1853 0.01681 0.00922 -0.0651 0.0082 1.0000
8.000 1.2039 0.01740 0.00990 -0.0637 0.0074 1.0000
8.250 1.2212 0.01806 0.01064 -0.0621 0.0067 1.0000
8.500 1.2358 0.01890 0.01155 -0.0601 0.0061 1.0000
8.750 1.2519 0.01956 0.01231 -0.0583 0.0055 1.0000
9.000 1.2661 0.02027 0.01309 -0.0563 0.0050 1.0000
9.250 1.2758 0.02112 0.01401 -0.0535 0.0047 1.0000
9.500 1.2807 0.02222 0.01520 -0.0501 0.0045 1.0000
9.750 1.2885 0.02319 0.01629 -0.0473 0.0043 1.0000
10.000 1.2943 0.02432 0.01755 -0.0444 0.0041 1.0000
10.250 1.2986 0.02561 0.01896 -0.0415 0.0040 1.0000
10.500 1.3023 0.02702 0.02049 -0.0388 0.0038 1.0000
10.750 1.3059 0.02853 0.02211 -0.0364 0.0037 1.0000
11.000 1.3120 0.02992 0.02359 -0.0346 0.0035 1.0000
11.250 1.3163 0.03151 0.02526 -0.0327 0.0033 1.0000
11.500 1.3160 0.03365 0.02753 -0.0307 0.0033 1.0000
11.750 1.3195 0.03552 0.02953 -0.0293 0.0030 1.0000
12.000 1.3209 0.03772 0.03188 -0.0279 0.0029 1.0000
12.250 1.3205 0.04022 0.03454 -0.0266 0.0028 1.0000
12.500 1.3184 0.04301 0.03748 -0.0254 0.0028 1.0000
12.750 1.3179 0.04574 0.04039 -0.0246 0.0027 1.0000
13.250 1.3111 0.05222 0.04726 -0.0234 0.0026 1.0000
13.500 1.3075 0.05569 0.05090 -0.0234 0.0025 1.0000
13.750 1.3022 0.05947 0.05485 -0.0236 0.0024 1.0000
14.000 1.2965 0.06349 0.05905 -0.0242 0.0024 1.0000
14.250 1.2907 0.06769 0.06339 -0.0253 0.0023 1.0000
14.500 1.2775 0.07325 0.06919 -0.0265 0.0024 1.0000
14.750 1.2713 0.07787 0.07394 -0.0283 0.0023 1.0000
15.000 1.2606 0.08344 0.07968 -0.0305 0.0023 1.0000
15.250 1.2441 0.09028 0.08675 -0.0330 0.0023 1.0000
15.500 1.2321 0.09659 0.09322 -0.0360 0.0023 1.0000
15.750 1.2182 0.10358 0.10038 -0.0394 0.0023 1.0000
16.000 1.2049 0.11072 0.10767 -0.0431 0.0022 1.0000
16.250 1.1887 0.11882 0.11594 -0.0475 0.0022 1.0000
16.500 1.1734 0.12708 0.12436 -0.0521 0.0022 1.0000
16.750 1.1535 0.13682 0.13427 -0.0575 0.0023 1.0000
17.000 1.1327 0.14733 0.14495 -0.0636 0.0023 1.0000
17.250 1.1048 0.16088 0.15869 -0.0713 0.0024 1.0000
17.500 1.0776 0.17589 0.17381 -0.0796 0.0025 1.0000
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Polar data table (+)
Polar graphs
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