GOE 419 AIRFOIL (goe419-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5
| Details | Polar file |
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Airfoil: GOE 419 AIRFOIL (goe419-il) Reynolds number: 1,000,000 Max Cl/Cd: 119.34 at α=1.5° Description: Mach=0 Ncrit=5 Source: Xfoil prediction Download polar: xf-goe419-il-1000000-n5.txt Download as CSV file: xf-goe419-il-1000000-n5.csv |
XFOIL Version 6.96
Calculated polar for: GOE 419 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
------ -------- --------- --------- -------- -------- --------
-9.750 -0.3907 0.11335 0.11173 -0.0174 1.0000 0.0019
-9.500 -0.3862 0.11045 0.10884 -0.0180 1.0000 0.0019
-9.000 -0.3784 0.10448 0.10289 -0.0192 1.0000 0.0019
-6.500 -0.2568 0.06772 0.06617 -0.0538 0.9509 0.0036
-6.250 -0.2172 0.06197 0.06036 -0.0641 0.9405 0.0036
-6.000 -0.1690 0.05579 0.05409 -0.0762 0.9330 0.0036
-5.500 -0.0759 0.02432 0.02226 -0.0875 0.8523 0.0038
-5.000 -0.0278 0.03649 0.03405 -0.1022 0.8637 0.0031
-4.750 -0.0042 0.03277 0.03015 -0.1035 0.8522 0.0029
-4.500 0.0202 0.02841 0.02556 -0.1043 0.8427 0.0026
-4.250 0.0440 0.02170 0.01841 -0.1040 0.8339 0.0022
-4.000 0.0631 0.01295 0.00878 -0.1016 0.8262 0.0019
-3.750 0.0863 0.01050 0.00573 -0.1004 0.8175 0.0019
-3.500 0.1104 0.00901 0.00383 -0.0994 0.8081 0.0020
-3.250 0.1355 0.00831 0.00294 -0.0988 0.7982 0.0023
-3.000 0.1609 0.00791 0.00240 -0.0982 0.7875 0.0027
-2.750 0.1866 0.00761 0.00198 -0.0977 0.7774 0.0035
-2.500 0.2121 0.00730 0.00156 -0.0972 0.7664 0.0050
-2.250 0.2378 0.00708 0.00122 -0.0967 0.7536 0.0078
-2.000 0.2630 0.00690 0.00104 -0.0961 0.7371 0.0214
-1.750 0.2884 0.00689 0.00094 -0.0956 0.7151 0.0299
-1.500 0.3134 0.00690 0.00083 -0.0951 0.6908 0.0348
-1.250 0.3385 0.00691 0.00074 -0.0945 0.6691 0.0377
-1.000 0.3639 0.00694 0.00066 -0.0941 0.6514 0.0393
-0.750 0.3896 0.00695 0.00059 -0.0937 0.6372 0.0447
-0.500 0.4154 0.00692 0.00057 -0.0933 0.6252 0.0661
-0.250 0.4415 0.00690 0.00056 -0.0931 0.6146 0.0850
0.000 0.4675 0.00688 0.00057 -0.0928 0.6049 0.1084
0.250 0.4934 0.00685 0.00059 -0.0925 0.5954 0.1495
0.500 0.5194 0.00678 0.00063 -0.0923 0.5860 0.2038
0.750 0.5452 0.00674 0.00068 -0.0920 0.5762 0.2544
1.000 0.5707 0.00665 0.00074 -0.0917 0.5657 0.3324
1.500 0.6480 0.00543 0.00100 -0.0975 0.5307 1.0000
1.750 0.6684 0.00581 0.00110 -0.0961 0.4678 1.0000
2.000 0.6889 0.00622 0.00128 -0.0947 0.4122 1.0000
2.250 0.7092 0.00669 0.00148 -0.0933 0.3494 1.0000
2.500 0.7179 0.00820 0.00204 -0.0900 0.1372 1.0000
2.750 0.7350 0.00911 0.00253 -0.0880 0.0117 1.0000
3.000 0.7590 0.00939 0.00285 -0.0873 0.0068 1.0000
3.250 0.7830 0.00965 0.00312 -0.0866 0.0046 1.0000
3.750 0.8292 0.01041 0.00401 -0.0847 0.0029 1.0000
4.000 0.8517 0.01085 0.00448 -0.0838 0.0022 1.0000
4.250 0.8734 0.01140 0.00514 -0.0825 0.0018 1.0000
4.500 0.8932 0.01214 0.00599 -0.0809 0.0015 1.0000
4.750 0.9126 0.01292 0.00687 -0.0793 0.0013 1.0000
5.000 0.9292 0.01407 0.00813 -0.0770 0.0012 1.0000
5.250 0.9446 0.01578 0.00996 -0.0744 0.0012 1.0000
5.500 0.9669 0.01886 0.01314 -0.0723 0.0013 1.0000
5.750 0.9946 0.02213 0.01661 -0.0709 0.0016 1.0000
6.000 1.0168 0.02476 0.01948 -0.0693 0.0020 1.0000
7.000 1.0772 0.03674 0.03244 -0.0606 0.0023 1.0000
7.250 1.0884 0.03974 0.03570 -0.0579 0.0023 1.0000
7.500 1.0975 0.04275 0.03896 -0.0552 0.0023 1.0000
7.750 1.1043 0.04586 0.04231 -0.0523 0.0023 1.0000
8.000 1.1089 0.04904 0.04571 -0.0494 0.0023 1.0000
8.250 1.1104 0.05233 0.04922 -0.0464 0.0023 1.0000
8.500 1.1092 0.05554 0.05263 -0.0433 0.0023 1.0000
8.750 1.1043 0.05882 0.05610 -0.0401 0.0023 1.0000
9.000 1.0963 0.06193 0.05936 -0.0369 0.0023 1.0000
9.250 1.0812 0.06455 0.06212 -0.0326 0.0023 1.0000
9.500 1.0646 0.06731 0.06500 -0.0293 0.0023 1.0000
9.750 1.0468 0.07055 0.06836 -0.0273 0.0023 1.0000
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Polar data table (+)
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