GOE 409 AIRFOIL (goe409-il) Xfoil prediction polar at RE=50,000 Ncrit=9
| Details | Polar file |
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Airfoil: GOE 409 AIRFOIL (goe409-il) Reynolds number: 50,000 Max Cl/Cd: 28.72 at α=4.75° Description: Mach=0 Ncrit=9 Source: Xfoil prediction Download polar: xf-goe409-il-50000.txt Download as CSV file: xf-goe409-il-50000.csv |
XFOIL Version 6.96
Calculated polar for: GOE 409 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 = 9.000
alpha CL CD CDp CM Top_Xtr Bot_Xtr
------ -------- --------- --------- -------- -------- --------
-10.500 -0.6213 0.09370 0.08586 -0.0256 1.0000 0.2044
-10.250 -0.6221 0.08816 0.08033 -0.0271 1.0000 0.2000
-10.000 -0.8112 0.07067 0.06285 -0.0325 1.0000 0.1744
-9.750 -0.8189 0.06646 0.05854 -0.0307 1.0000 0.1734
-9.500 -0.8292 0.06249 0.05443 -0.0282 1.0000 0.1722
-9.250 -0.8406 0.05874 0.05048 -0.0252 1.0000 0.1710
-9.000 -0.8509 0.05530 0.04677 -0.0217 1.0000 0.1703
-8.750 -0.8586 0.05220 0.04334 -0.0180 1.0000 0.1708
-8.500 -0.8625 0.04936 0.04014 -0.0144 1.0000 0.1718
-8.250 -0.8622 0.04668 0.03707 -0.0110 1.0000 0.1729
-8.000 -0.8577 0.04413 0.03413 -0.0079 1.0000 0.1737
-7.750 -0.8497 0.04177 0.03135 -0.0050 1.0000 0.1745
-7.500 -0.8322 0.03932 0.02881 -0.0036 1.0000 0.1768
-7.250 -0.8140 0.03748 0.02690 -0.0021 1.0000 0.1809
-7.000 -0.7982 0.03577 0.02493 -0.0001 1.0000 0.1859
-6.750 -0.7822 0.03414 0.02289 0.0020 1.0000 0.1899
-6.500 -0.7570 0.03233 0.02117 0.0025 1.0000 0.1953
-6.250 -0.7336 0.03087 0.01958 0.0035 1.0000 0.2022
-6.000 -0.7081 0.02942 0.01816 0.0042 1.0000 0.2118
-5.750 -0.6842 0.02812 0.01693 0.0051 1.0000 0.2265
-5.500 -0.6608 0.02676 0.01574 0.0061 1.0000 0.2487
-5.250 -0.6417 0.02501 0.01448 0.0079 1.0000 0.2926
-5.000 -0.6410 0.02308 0.01370 0.0129 1.0000 0.3986
-4.750 -0.6420 0.02235 0.01364 0.0193 1.0000 0.5201
-4.500 -0.6323 0.02233 0.01395 0.0245 1.0000 0.5994
-4.250 -0.6175 0.02275 0.01453 0.0292 1.0000 0.6596
-4.000 -0.5968 0.02354 0.01535 0.0333 1.0000 0.7087
-3.750 -0.5700 0.02443 0.01617 0.0361 1.0000 0.7472
-3.500 -0.5367 0.02553 0.01713 0.0379 1.0000 0.7816
-3.250 -0.4853 0.02726 0.01864 0.0374 1.0000 0.8166
-3.000 -0.3495 0.03056 0.02145 0.0244 1.0000 0.8611
-2.750 -0.2662 0.03104 0.02163 0.0158 1.0000 0.8904
-2.500 -0.1891 0.03091 0.02128 0.0072 1.0000 0.9211
-2.250 -0.1198 0.03021 0.02041 -0.0012 1.0000 0.9483
-2.000 -0.0680 0.02936 0.01946 -0.0073 1.0000 0.9699
-1.750 -0.0057 0.02811 0.01812 -0.0158 1.0000 0.9885
-1.500 0.0346 0.02717 0.01714 -0.0206 1.0000 1.0000
-1.250 0.0335 0.02682 0.01682 -0.0178 1.0000 1.0000
-1.000 0.0302 0.02652 0.01654 -0.0147 1.0000 1.0000
-0.750 0.0248 0.02626 0.01631 -0.0113 1.0000 1.0000
-0.500 0.0176 0.02607 0.01613 -0.0077 1.0000 1.0000
-0.250 0.0091 0.02595 0.01602 -0.0039 1.0000 1.0000
0.000 0.0000 0.02591 0.01599 0.0000 1.0000 1.0000
0.250 -0.0091 0.02595 0.01602 0.0039 1.0000 1.0000
0.500 -0.0175 0.02606 0.01613 0.0077 1.0000 1.0000
0.750 -0.0248 0.02626 0.01630 0.0113 1.0000 1.0000
1.000 -0.0301 0.02651 0.01653 0.0147 1.0000 1.0000
1.250 -0.0335 0.02681 0.01681 0.0178 1.0000 1.0000
1.500 -0.0345 0.02716 0.01713 0.0206 1.0000 1.0000
1.750 0.0056 0.02809 0.01810 0.0158 0.9885 1.0000
2.000 0.0680 0.02935 0.01946 0.0073 0.9699 1.0000
2.250 0.1199 0.03019 0.02040 0.0012 0.9482 1.0000
2.500 0.1894 0.03089 0.02127 -0.0072 0.9209 1.0000
2.750 0.2664 0.03102 0.02161 -0.0158 0.8903 1.0000
3.000 0.3505 0.03057 0.02146 -0.0246 0.8618 1.0000
3.250 0.4862 0.02722 0.01860 -0.0375 0.8165 1.0000
3.500 0.5366 0.02552 0.01712 -0.0379 0.7816 1.0000
3.750 0.5699 0.02443 0.01616 -0.0361 0.7472 1.0000
4.000 0.5966 0.02354 0.01534 -0.0333 0.7087 1.0000
4.250 0.6174 0.02276 0.01455 -0.0293 0.6602 1.0000
4.500 0.6323 0.02233 0.01395 -0.0245 0.5997 1.0000
4.750 0.6419 0.02235 0.01364 -0.0193 0.5202 1.0000
5.000 0.6409 0.02307 0.01369 -0.0129 0.3987 1.0000
5.250 0.6416 0.02501 0.01447 -0.0078 0.2926 1.0000
5.500 0.6607 0.02676 0.01574 -0.0061 0.2488 1.0000
5.750 0.6840 0.02810 0.01691 -0.0051 0.2267 1.0000
6.000 0.7080 0.02941 0.01815 -0.0041 0.2118 1.0000
6.250 0.7335 0.03086 0.01957 -0.0035 0.2022 1.0000
6.500 0.7568 0.03233 0.02116 -0.0025 0.1952 1.0000
6.750 0.7821 0.03413 0.02288 -0.0020 0.1900 1.0000
7.000 0.7980 0.03577 0.02492 0.0001 0.1858 1.0000
7.250 0.8139 0.03747 0.02689 0.0021 0.1810 1.0000
7.500 0.8321 0.03931 0.02880 0.0036 0.1768 1.0000
7.750 0.8497 0.04176 0.03134 0.0050 0.1745 1.0000
8.000 0.8576 0.04413 0.03412 0.0079 0.1737 1.0000
8.250 0.8622 0.04668 0.03706 0.0110 0.1729 1.0000
8.500 0.8626 0.04935 0.04013 0.0144 0.1718 1.0000
8.750 0.8586 0.05220 0.04333 0.0180 0.1708 1.0000
9.000 0.8510 0.05531 0.04677 0.0217 0.1704 1.0000
9.250 0.8407 0.05874 0.05048 0.0251 0.1710 1.0000
9.500 0.8290 0.06250 0.05444 0.0282 0.1721 1.0000
9.750 0.8190 0.06648 0.05856 0.0306 0.1734 1.0000
10.000 0.8112 0.07069 0.06287 0.0325 0.1744 1.0000
10.250 0.6224 0.08820 0.08037 0.0269 0.1997 1.0000
10.500 0.5944 0.09846 0.09054 0.0210 0.2194 1.0000
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Polar data table (+)
Polar graphs
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