GOE 429 AIRFOIL (goe429-il) Xfoil prediction polar at RE=100,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 429 AIRFOIL (goe429-il) Reynolds number: 100,000 Max Cl/Cd: 35.54 at α=6.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe429-il-100000.txt Download as CSV file: xf-goe429-il-100000.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 = 0.100 e 6 Ncrit = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-12.250 -0.9074 0.09431 0.08937 0.0117 1.0000 0.0738
-12.000 -0.9377 0.08407 0.07907 0.0049 1.0000 0.0728
-11.750 -0.9732 0.07448 0.06935 -0.0021 1.0000 0.0714
-11.500 -1.0089 0.06658 0.06123 -0.0074 1.0000 0.0701
-11.250 -1.0393 0.06087 0.05524 -0.0092 1.0000 0.0696
-11.000 -1.0563 0.05600 0.04999 -0.0095 1.0000 0.0702
-10.750 -1.0653 0.05144 0.04496 -0.0093 1.0000 0.0716
-10.500 -1.0687 0.04716 0.04004 -0.0086 1.0000 0.0735
-10.250 -1.0594 0.04336 0.03597 -0.0082 1.0000 0.0763
-10.000 -1.0405 0.04167 0.03426 -0.0078 1.0000 0.0812
-9.750 -1.0301 0.03912 0.03097 -0.0069 1.0000 0.0864
-9.500 -1.0090 0.03653 0.02858 -0.0067 1.0000 0.0926
-9.250 -0.9919 0.03437 0.02590 -0.0061 1.0000 0.1002
-9.000 -0.9694 0.03255 0.02420 -0.0059 1.0000 0.1083
-8.750 -0.9480 0.03059 0.02205 -0.0055 1.0000 0.1167
-8.500 -0.9254 0.02934 0.02058 -0.0052 1.0000 0.1261
-8.250 -0.9013 0.02790 0.01926 -0.0050 1.0000 0.1349
-8.000 -0.8772 0.02654 0.01772 -0.0047 1.0000 0.1440
-7.750 -0.8520 0.02550 0.01662 -0.0045 1.0000 0.1531
-7.500 -0.8264 0.02437 0.01554 -0.0043 1.0000 0.1618
-7.250 -0.8005 0.02330 0.01434 -0.0040 1.0000 0.1717
-7.000 -0.7744 0.02242 0.01343 -0.0038 1.0000 0.1827
-6.750 -0.7483 0.02144 0.01254 -0.0036 1.0000 0.1945
-6.500 -0.7222 0.02051 0.01164 -0.0034 1.0000 0.2092
-6.250 -0.6961 0.01957 0.01074 -0.0033 1.0000 0.2283
-6.000 -0.6705 0.01858 0.00995 -0.0030 1.0000 0.2532
-5.750 -0.6452 0.01770 0.00928 -0.0028 1.0000 0.2882
-5.500 -0.6192 0.01707 0.00879 -0.0026 1.0000 0.3260
-5.250 -0.5928 0.01664 0.00847 -0.0023 1.0000 0.3586
-5.000 -0.5659 0.01630 0.00819 -0.0021 1.0000 0.3877
-4.750 -0.5389 0.01597 0.00798 -0.0019 1.0000 0.4131
-4.500 -0.5115 0.01566 0.00771 -0.0017 1.0000 0.4392
-4.250 -0.4843 0.01536 0.00752 -0.0015 1.0000 0.4641
-4.000 -0.4570 0.01508 0.00733 -0.0013 1.0000 0.4894
-3.750 -0.4298 0.01484 0.00719 -0.0010 1.0000 0.5157
-3.500 -0.4032 0.01469 0.00719 -0.0005 1.0000 0.5413
-3.250 -0.3767 0.01462 0.00723 -0.0001 1.0000 0.5685
-3.000 -0.3512 0.01462 0.00737 0.0006 1.0000 0.5923
-2.750 -0.3268 0.01460 0.00742 0.0012 1.0000 0.6158
-2.500 -0.3062 0.01463 0.00755 0.0024 1.0000 0.6350
-2.250 -0.2875 0.01466 0.00764 0.0036 1.0000 0.6534
-2.000 -0.2685 0.01469 0.00772 0.0046 1.0000 0.6717
-1.750 -0.2486 0.01473 0.00780 0.0054 1.0000 0.6900
-1.500 -0.2062 0.01474 0.00781 0.0022 0.9894 0.7122
-1.250 -0.1562 0.01470 0.00787 -0.0018 0.9751 0.7324
-1.000 -0.1076 0.01463 0.00787 -0.0054 0.9595 0.7529
-0.750 -0.0648 0.01456 0.00784 -0.0075 0.9411 0.7732
-0.500 -0.0334 0.01451 0.00783 -0.0073 0.9197 0.7927
-0.250 -0.0093 0.01447 0.00785 -0.0053 0.8984 0.8113
0.000 0.0107 0.01444 0.00787 -0.0022 0.8785 0.8316
0.250 0.0289 0.01439 0.00786 0.0011 0.8579 0.8542
0.500 0.0471 0.01430 0.00782 0.0044 0.8369 0.8768
0.750 0.0674 0.01419 0.00772 0.0074 0.8161 0.8988
1.000 0.0915 0.01405 0.00757 0.0095 0.7947 0.9211
1.250 0.1238 0.01393 0.00744 0.0098 0.7690 0.9419
1.500 0.1629 0.01382 0.00727 0.0086 0.7415 0.9615
1.750 0.2066 0.01372 0.00711 0.0063 0.7096 0.9801
2.000 0.2550 0.01361 0.00691 0.0027 0.6704 0.9971
2.250 0.2770 0.01345 0.00665 0.0033 0.6346 1.0000
2.500 0.2946 0.01336 0.00643 0.0047 0.5938 1.0000
2.750 0.3164 0.01341 0.00625 0.0056 0.5441 1.0000
3.000 0.3411 0.01368 0.00618 0.0061 0.4899 1.0000
3.250 0.3670 0.01419 0.00632 0.0063 0.4416 1.0000
3.500 0.3936 0.01481 0.00661 0.0064 0.4050 1.0000
3.750 0.4210 0.01542 0.00701 0.0063 0.3761 1.0000
4.000 0.4486 0.01604 0.00747 0.0062 0.3538 1.0000
4.250 0.4763 0.01658 0.00791 0.0061 0.3335 1.0000
4.500 0.5037 0.01706 0.00828 0.0061 0.3144 1.0000
4.750 0.5311 0.01753 0.00869 0.0060 0.2978 1.0000
5.000 0.5587 0.01804 0.00918 0.0060 0.2846 1.0000
5.250 0.5861 0.01853 0.00965 0.0059 0.2720 1.0000
5.500 0.6133 0.01902 0.01008 0.0059 0.2600 1.0000
5.750 0.6404 0.01939 0.01049 0.0059 0.2464 1.0000
6.000 0.6672 0.01971 0.01085 0.0060 0.2306 1.0000
6.250 0.6936 0.02006 0.01124 0.0060 0.2137 1.0000
6.500 0.7199 0.02034 0.01169 0.0060 0.1888 1.0000
6.750 0.7446 0.02095 0.01219 0.0061 0.1536 1.0000
7.000 0.7692 0.02197 0.01311 0.0064 0.1269 1.0000
7.250 0.7939 0.02312 0.01421 0.0068 0.1142 1.0000
7.500 0.8185 0.02438 0.01546 0.0072 0.1068 1.0000
7.750 0.8432 0.02551 0.01660 0.0076 0.1008 1.0000
8.000 0.8674 0.02688 0.01798 0.0079 0.0961 1.0000
8.250 0.8919 0.02827 0.01955 0.0084 0.0920 1.0000
8.500 0.9159 0.02975 0.02110 0.0088 0.0891 1.0000
8.750 0.9397 0.03146 0.02278 0.0091 0.0868 1.0000
9.000 0.9617 0.03366 0.02525 0.0096 0.0848 1.0000
9.250 0.9815 0.03592 0.02790 0.0100 0.0825 1.0000
9.500 0.9995 0.03847 0.03080 0.0105 0.0808 1.0000
9.750 1.0145 0.04154 0.03424 0.0110 0.0798 1.0000
10.000 1.0261 0.04485 0.03791 0.0114 0.0789 1.0000
10.250 1.0372 0.04790 0.04122 0.0118 0.0777 1.0000
10.500 1.0484 0.05079 0.04426 0.0121 0.0763 1.0000
10.750 1.0517 0.05485 0.04856 0.0123 0.0755 1.0000
11.000 1.0337 0.06031 0.05453 0.0120 0.0760 1.0000
11.250 0.8384 0.09555 0.09058 -0.0151 0.0873 1.0000
11.500 0.8469 0.09777 0.09278 -0.0145 0.0865 1.0000
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Polar data table (+)
Polar graphs
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