GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 113 AIRFOIL (goe113-il) Reynolds number: 1,000,000 Max Cl/Cd: 83.29 at α=3.75° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe113-il-1000000-n5.txt Download as CSV file: xf-goe113-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 113 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 1.000 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-8.750 -0.3435 0.08737 0.08589 -0.0150 1.0000 0.0024
-8.500 -0.3425 0.08356 0.08210 -0.0158 1.0000 0.0024
-8.250 -0.4483 0.09628 0.09472 -0.0063 1.0000 0.0024
-8.000 -0.4440 0.09285 0.09131 -0.0080 1.0000 0.0023
-7.750 -0.4406 0.08933 0.08780 -0.0098 1.0000 0.0023
-7.500 -0.4385 0.08587 0.08437 -0.0116 1.0000 0.0023
-7.250 -0.4340 0.08208 0.08060 -0.0146 1.0000 0.0022
-7.000 -0.4268 0.07785 0.07638 -0.0186 1.0000 0.0022
-6.750 -0.4194 0.07354 0.07207 -0.0223 1.0000 0.0022
-6.500 -0.4022 0.06792 0.06643 -0.0288 0.9980 0.0023
-6.250 -0.3742 0.06100 0.05945 -0.0377 0.9945 0.0025
-6.000 -0.3454 0.05411 0.05247 -0.0458 0.9899 0.0026
-5.750 -0.3160 0.04836 0.04659 -0.0519 0.9854 0.0027
-5.500 -0.2863 0.04407 0.04218 -0.0562 0.9800 0.0029
-5.250 -0.2574 0.04038 0.03836 -0.0594 0.9730 0.0032
-5.000 -0.2271 0.03524 0.03301 -0.0624 0.9644 0.0034
-4.750 -0.1967 0.02764 0.02497 -0.0647 0.9533 0.0033
-4.500 -0.1703 0.01133 0.00691 -0.0642 0.9417 0.0037
-4.250 -0.1438 0.01002 0.00531 -0.0639 0.9212 0.0041
-4.000 -0.1183 0.00942 0.00445 -0.0632 0.8888 0.0044
-3.750 -0.0931 0.00886 0.00362 -0.0625 0.8495 0.0051
-3.500 -0.0674 0.00854 0.00308 -0.0620 0.8136 0.0061
-3.250 -0.0411 0.00835 0.00264 -0.0616 0.7774 0.0071
-3.000 -0.0142 0.00820 0.00236 -0.0614 0.7453 0.0098
-2.750 0.0132 0.00836 0.00247 -0.0612 0.7146 0.0127
-2.500 0.0408 0.00875 0.00277 -0.0611 0.6816 0.0171
-2.250 0.0684 0.00906 0.00294 -0.0611 0.6496 0.0206
-2.000 0.0961 0.00931 0.00308 -0.0611 0.6267 0.0222
-1.750 0.1241 0.00963 0.00331 -0.0611 0.6081 0.0230
-1.500 0.1514 0.00925 0.00281 -0.0611 0.5898 0.0239
-1.250 0.1787 0.00883 0.00232 -0.0611 0.5722 0.0255
-1.000 0.2063 0.00867 0.00207 -0.0611 0.5563 0.0262
-0.750 0.2341 0.00851 0.00181 -0.0612 0.5408 0.0256
-0.500 0.2619 0.00840 0.00161 -0.0612 0.5257 0.0252
-0.250 0.2898 0.00831 0.00145 -0.0612 0.5130 0.0248
0.000 0.3178 0.00824 0.00132 -0.0613 0.5001 0.0246
0.250 0.3457 0.00818 0.00121 -0.0614 0.4868 0.0244
0.500 0.3735 0.00816 0.00111 -0.0614 0.4709 0.0242
0.750 0.4012 0.00816 0.00103 -0.0614 0.4504 0.0242
1.000 0.4289 0.00819 0.00096 -0.0615 0.4278 0.0245
1.250 0.4564 0.00827 0.00091 -0.0615 0.3989 0.0257
1.500 0.4835 0.00843 0.00093 -0.0615 0.3611 0.0304
1.750 0.5102 0.00868 0.00101 -0.0614 0.3153 0.0339
2.000 0.5368 0.00896 0.00113 -0.0614 0.2677 0.0431
2.500 0.5849 0.00728 0.00152 -0.0607 0.2142 1.0000
2.750 0.6122 0.00745 0.00163 -0.0607 0.2009 1.0000
3.000 0.6389 0.00775 0.00177 -0.0607 0.1687 1.0000
3.250 0.6656 0.00805 0.00193 -0.0606 0.1392 1.0000
3.500 0.6921 0.00836 0.00212 -0.0606 0.1153 1.0000
3.750 0.7188 0.00863 0.00230 -0.0605 0.0990 1.0000
4.000 0.7438 0.00924 0.00267 -0.0602 0.0408 1.0000
4.250 0.7698 0.00965 0.00299 -0.0600 0.0125 1.0000
4.500 0.7967 0.00986 0.00325 -0.0599 0.0100 1.0000
4.750 0.8232 0.01017 0.00359 -0.0598 0.0074 1.0000
5.000 0.8496 0.01046 0.00391 -0.0596 0.0060 1.0000
5.250 0.8759 0.01077 0.00427 -0.0594 0.0051 1.0000
5.500 0.9015 0.01120 0.00475 -0.0591 0.0043 1.0000
5.750 0.9269 0.01168 0.00531 -0.0588 0.0039 1.0000
6.000 0.9524 0.01211 0.00579 -0.0585 0.0034 1.0000
6.250 0.9779 0.01248 0.00620 -0.0582 0.0029 1.0000
6.500 1.0028 0.01296 0.00673 -0.0579 0.0027 1.0000
6.750 1.0254 0.01386 0.00774 -0.0571 0.0024 1.0000
7.000 1.0492 0.01450 0.00846 -0.0565 0.0022 1.0000
7.250 1.0717 0.01533 0.00941 -0.0558 0.0020 1.0000
7.500 1.0926 0.01639 0.01059 -0.0547 0.0018 1.0000
7.750 1.1121 0.01766 0.01199 -0.0534 0.0017 1.0000
8.000 1.1303 0.01918 0.01366 -0.0519 0.0016 1.0000
8.250 1.1471 0.02112 0.01577 -0.0501 0.0016 1.0000
8.500 1.1630 0.02368 0.01857 -0.0481 0.0016 1.0000
8.750 1.1795 0.02614 0.02125 -0.0463 0.0015 1.0000
9.000 1.1968 0.02773 0.02301 -0.0451 0.0015 1.0000
9.250 1.2132 0.02917 0.02461 -0.0438 0.0014 1.0000
9.500 1.2280 0.03057 0.02617 -0.0425 0.0013 1.0000
9.750 1.2360 0.03312 0.02897 -0.0404 0.0012 1.0000
10.000 1.2359 0.03706 0.03327 -0.0373 0.0012 1.0000
10.250 1.2365 0.04021 0.03668 -0.0345 0.0012 1.0000
10.500 1.2322 0.04333 0.04004 -0.0313 0.0011 1.0000
10.750 1.2208 0.04622 0.04313 -0.0275 0.0011 1.0000
11.000 1.2071 0.04949 0.04659 -0.0248 0.0011 1.0000
11.250 1.1914 0.05336 0.05066 -0.0236 0.0011 1.0000
11.500 1.1765 0.05784 0.05531 -0.0241 0.0011 1.0000
11.750 1.1598 0.06324 0.06088 -0.0261 0.0011 1.0000
12.000 1.1431 0.06939 0.06719 -0.0295 0.0011 1.0000
12.250 1.1263 0.07627 0.07422 -0.0339 0.0011 1.0000
12.500 1.1091 0.08383 0.08190 -0.0390 0.0011 1.0000
12.750 1.0916 0.09214 0.09033 -0.0445 0.0011 1.0000
13.000 1.0733 0.10132 0.09961 -0.0507 0.0012 1.0000
13.250 1.0545 0.11139 0.10977 -0.0570 0.0012 1.0000
13.750 0.8844 0.10785 0.10624 -0.0369 0.0013 1.0000
14.000 0.8540 0.11475 0.11323 -0.0395 0.0013 1.0000
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Polar data table (+)
Polar graphs
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