GOE 429 AIRFOIL (goe429-il) Xfoil prediction polar at RE=50,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 429 AIRFOIL (goe429-il) Reynolds number: 50,000 Max Cl/Cd: 26.9 at α=7° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe429-il-50000-n5.txt Download as CSV file: xf-goe429-il-50000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 429 AIRFOIL
1 1 Reynolds number fixed Mach number fixed
xtrf = 1.000 (top) 1.000 (bottom)
Mach = 0.000 Re = 0.050 e 6 Ncrit = 5.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.8610 0.10084 0.09378 0.0158 1.0000 0.0759
-12.000 -0.9023 0.08763 0.08049 0.0067 1.0000 0.0752
-11.750 -0.9393 0.07703 0.06975 -0.0010 1.0000 0.0745
-11.500 -0.9702 0.06893 0.06142 -0.0065 1.0000 0.0743
-11.250 -0.9957 0.06275 0.05494 -0.0098 1.0000 0.0749
-11.000 -1.0165 0.05770 0.04945 -0.0107 1.0000 0.0762
-10.750 -1.0057 0.05575 0.04759 -0.0102 1.0000 0.0801
-10.500 -1.0033 0.05278 0.04439 -0.0102 1.0000 0.0844
-10.250 -1.0059 0.04888 0.03985 -0.0102 1.0000 0.0890
-10.000 -0.9890 0.04728 0.03838 -0.0098 1.0000 0.0947
-9.750 -0.9797 0.04456 0.03522 -0.0096 1.0000 0.1014
-9.500 -0.9628 0.04262 0.03326 -0.0092 1.0000 0.1077
-9.250 -0.9465 0.04049 0.03084 -0.0088 1.0000 0.1147
-9.000 -0.9275 0.03860 0.02884 -0.0085 1.0000 0.1212
-8.750 -0.9074 0.03691 0.02694 -0.0081 1.0000 0.1283
-8.500 -0.8863 0.03525 0.02515 -0.0078 1.0000 0.1352
-8.250 -0.8644 0.03385 0.02362 -0.0074 1.0000 0.1429
-8.000 -0.8417 0.03249 0.02219 -0.0070 1.0000 0.1502
-7.750 -0.8186 0.03121 0.02068 -0.0067 1.0000 0.1596
-7.500 -0.7953 0.03011 0.01962 -0.0064 1.0000 0.1685
-7.250 -0.7715 0.02897 0.01841 -0.0061 1.0000 0.1790
-7.000 -0.7476 0.02789 0.01727 -0.0058 1.0000 0.1922
-6.750 -0.7236 0.02683 0.01616 -0.0055 1.0000 0.2072
-6.500 -0.6997 0.02582 0.01519 -0.0052 1.0000 0.2250
-6.250 -0.6758 0.02487 0.01430 -0.0049 1.0000 0.2467
-6.000 -0.6515 0.02400 0.01348 -0.0046 1.0000 0.2713
-5.750 -0.6270 0.02328 0.01282 -0.0043 1.0000 0.2976
-5.500 -0.6021 0.02267 0.01229 -0.0040 1.0000 0.3242
-5.250 -0.5767 0.02214 0.01178 -0.0037 1.0000 0.3509
-5.000 -0.5510 0.02167 0.01132 -0.0034 1.0000 0.3775
-4.750 -0.5253 0.02126 0.01096 -0.0030 1.0000 0.4025
-4.500 -0.4994 0.02089 0.01063 -0.0027 1.0000 0.4302
-4.250 -0.4740 0.02060 0.01040 -0.0022 1.0000 0.4609
-4.000 -0.4489 0.02037 0.01029 -0.0014 1.0000 0.4894
-3.750 -0.4235 0.02013 0.01014 -0.0007 1.0000 0.5142
-3.500 -0.3978 0.01988 0.00993 -0.0002 1.0000 0.5357
-3.250 -0.3721 0.01963 0.00970 0.0003 1.0000 0.5565
-3.000 -0.3469 0.01944 0.00956 0.0008 1.0000 0.5759
-2.750 -0.3224 0.01929 0.00945 0.0015 1.0000 0.5948
-2.500 -0.2989 0.01915 0.00935 0.0022 1.0000 0.6139
-2.250 -0.2769 0.01902 0.00926 0.0030 1.0000 0.6333
-2.000 -0.2566 0.01893 0.00917 0.0039 1.0000 0.6531
-1.750 -0.2375 0.01887 0.00916 0.0051 1.0000 0.6716
-1.500 -0.2052 0.01886 0.00918 0.0039 0.9885 0.6921
-1.250 -0.1616 0.01885 0.00918 0.0008 0.9696 0.7160
-1.000 -0.1213 0.01884 0.00924 -0.0012 0.9499 0.7384
-0.750 -0.0828 0.01882 0.00927 -0.0027 0.9304 0.7616
-0.500 -0.0487 0.01879 0.00929 -0.0032 0.9095 0.7835
-0.250 -0.0184 0.01875 0.00930 -0.0028 0.8873 0.8034
0.000 0.0101 0.01871 0.00928 -0.0018 0.8651 0.8229
0.250 0.0356 0.01865 0.00926 -0.0004 0.8406 0.8429
0.500 0.0615 0.01857 0.00920 0.0011 0.8171 0.8632
0.750 0.0892 0.01849 0.00914 0.0022 0.7912 0.8837
1.000 0.1195 0.01840 0.00905 0.0028 0.7662 0.9058
1.250 0.1530 0.01833 0.00898 0.0026 0.7386 0.9293
1.500 0.1904 0.01827 0.00889 0.0014 0.7084 0.9537
1.750 0.2291 0.01821 0.00880 -0.0003 0.6751 0.9811
2.000 0.2597 0.01818 0.00869 -0.0009 0.6408 1.0000
2.250 0.2799 0.01822 0.00863 0.0003 0.6068 1.0000
2.500 0.3025 0.01832 0.00860 0.0012 0.5685 1.0000
2.750 0.3262 0.01847 0.00858 0.0021 0.5276 1.0000
3.000 0.3505 0.01873 0.00863 0.0028 0.4866 1.0000
3.250 0.3753 0.01910 0.00874 0.0035 0.4485 1.0000
3.500 0.4006 0.01957 0.00899 0.0039 0.4157 1.0000
3.750 0.4264 0.02012 0.00938 0.0041 0.3880 1.0000
4.000 0.4526 0.02071 0.00982 0.0042 0.3652 1.0000
4.250 0.4793 0.02132 0.01034 0.0043 0.3465 1.0000
4.500 0.5060 0.02196 0.01093 0.0044 0.3313 1.0000
4.750 0.5329 0.02262 0.01160 0.0044 0.3174 1.0000
5.000 0.5599 0.02329 0.01232 0.0045 0.3049 1.0000
5.250 0.5864 0.02397 0.01303 0.0046 0.2923 1.0000
5.500 0.6123 0.02459 0.01366 0.0047 0.2772 1.0000
5.750 0.6377 0.02516 0.01425 0.0049 0.2605 1.0000
6.000 0.6628 0.02573 0.01488 0.0050 0.2443 1.0000
6.250 0.6877 0.02632 0.01553 0.0051 0.2283 1.0000
6.500 0.7124 0.02695 0.01624 0.0052 0.2127 1.0000
6.750 0.7364 0.02757 0.01697 0.0054 0.1943 1.0000
7.000 0.7598 0.02825 0.01773 0.0055 0.1755 1.0000
7.250 0.7826 0.02910 0.01862 0.0057 0.1574 1.0000
7.500 0.8051 0.03023 0.01990 0.0060 0.1378 1.0000
7.750 0.8265 0.03153 0.02127 0.0063 0.1219 1.0000
8.000 0.8473 0.03289 0.02264 0.0067 0.1106 1.0000
8.250 0.8677 0.03428 0.02407 0.0070 0.1019 1.0000
8.500 0.8879 0.03595 0.02590 0.0075 0.0959 1.0000
8.750 0.9073 0.03766 0.02776 0.0080 0.0915 1.0000
9.000 0.9259 0.03927 0.02941 0.0084 0.0885 1.0000
9.250 0.9430 0.04150 0.03192 0.0089 0.0854 1.0000
9.500 0.9580 0.04378 0.03447 0.0093 0.0823 1.0000
9.750 0.9721 0.04577 0.03659 0.0096 0.0795 1.0000
10.000 0.9867 0.04751 0.03833 0.0101 0.0773 1.0000
10.250 0.9931 0.05084 0.04202 0.0103 0.0759 1.0000
10.500 0.9933 0.05469 0.04627 0.0103 0.0749 1.0000
10.750 0.9851 0.05907 0.05099 0.0098 0.0742 1.0000
11.000 0.9656 0.06440 0.05661 0.0081 0.0740 1.0000
11.250 0.9343 0.07230 0.06475 0.0030 0.0742 1.0000
11.500 0.8841 0.08545 0.07811 -0.0070 0.0751 1.0000
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Polar data table (+)
Polar graphs
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