Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 226 (MVA H.36) AIRFOIL (goe226-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 226 (MVA H.36) AIRFOIL (goe226-il)
Reynolds number: 500,000
Max Cl/Cd: 102.68 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe226-il-500000-n5.txt
Download as CSV file: xf-goe226-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 226 (MVA H.36) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.1360   0.11503   0.11194  -0.0750   0.8489   0.0227
 -12.000  -0.1254   0.11291   0.10980  -0.0760   0.8437   0.0230
 -10.500  -0.3112   0.02787   0.02373  -0.1662   0.8126   0.0328
 -10.250  -0.2681   0.02289   0.01837  -0.1791   0.8072   0.0334
 -10.000  -0.2343   0.02095   0.01615  -0.1837   0.8018   0.0339
  -9.750  -0.2023   0.01962   0.01461  -0.1866   0.7952   0.0343
  -9.500  -0.1718   0.01855   0.01344  -0.1886   0.7889   0.0348
  -9.250  -0.1414   0.01778   0.01256  -0.1901   0.7835   0.0352
  -9.000  -0.1111   0.01712   0.01181  -0.1914   0.7770   0.0357
  -8.750  -0.0812   0.01656   0.01113  -0.1924   0.7700   0.0363
  -8.500  -0.0512   0.01602   0.01050  -0.1933   0.7624   0.0369
  -8.250  -0.0213   0.01550   0.00984  -0.1941   0.7547   0.0376
  -8.000   0.0088   0.01499   0.00920  -0.1950   0.7483   0.0382
  -7.750   0.0388   0.01453   0.00861  -0.1957   0.7413   0.0388
  -7.250   0.0986   0.01364   0.00754  -0.1971   0.7271   0.0401
  -7.000   0.1283   0.01332   0.00714  -0.1976   0.7194   0.0408
  -6.750   0.1580   0.01303   0.00677  -0.1980   0.7123   0.0416
  -6.500   0.1877   0.01277   0.00642  -0.1985   0.7046   0.0426
  -6.250   0.2173   0.01254   0.00609  -0.1988   0.6976   0.0436
  -6.000   0.2470   0.01232   0.00577  -0.1992   0.6906   0.0445
  -5.750   0.2767   0.01202   0.00541  -0.1997   0.6844   0.0456
  -5.500   0.3065   0.01180   0.00515  -0.2001   0.6784   0.0467
  -5.250   0.3361   0.01163   0.00491  -0.2004   0.6720   0.0480
  -4.750   0.3951   0.01135   0.00447  -0.2010   0.6601   0.0506
  -4.500   0.4247   0.01115   0.00424  -0.2014   0.6536   0.0525
  -4.000   0.4837   0.01092   0.00391  -0.2019   0.6438   0.0565
  -3.750   0.5133   0.01081   0.00375  -0.2022   0.6386   0.0585
  -3.500   0.5425   0.01072   0.00362  -0.2024   0.6327   0.0612
  -3.250   0.5718   0.01065   0.00350  -0.2026   0.6275   0.0643
  -3.000   0.6012   0.01056   0.00342  -0.2029   0.6223   0.0684
  -2.750   0.6303   0.01052   0.00335  -0.2030   0.6171   0.0735
  -2.500   0.6592   0.01050   0.00331  -0.2032   0.6124   0.0799
  -2.250   0.6886   0.01045   0.00328  -0.2034   0.6077   0.0868
  -2.000   0.7175   0.01044   0.00325  -0.2035   0.6020   0.0936
  -1.750   0.7461   0.01045   0.00322  -0.2036   0.5959   0.0993
  -1.500   0.7750   0.01044   0.00320  -0.2037   0.5900   0.1052
  -1.250   0.8037   0.01044   0.00318  -0.2038   0.5832   0.1100
  -1.000   0.8320   0.01046   0.00317  -0.2039   0.5765   0.1158
  -0.750   0.8608   0.01046   0.00317  -0.2040   0.5690   0.1214
  -0.500   0.8887   0.01049   0.00317  -0.2040   0.5601   0.1288
  -0.250   0.9171   0.01051   0.00319  -0.2040   0.5497   0.1379
   0.000   0.9447   0.01053   0.00324  -0.2040   0.5370   0.1597
   0.250   0.9718   0.01059   0.00332  -0.2039   0.5192   0.1992
   0.500   0.9972   0.01078   0.00346  -0.2035   0.4948   0.2290
   0.750   1.0219   0.01105   0.00363  -0.2030   0.4726   0.2485
   1.000   1.0469   0.01132   0.00381  -0.2025   0.4575   0.2620
   1.250   1.0727   0.01153   0.00398  -0.2021   0.4473   0.2743
   1.500   1.0985   0.01173   0.00415  -0.2018   0.4398   0.2857
   1.750   1.1245   0.01193   0.00432  -0.2014   0.4328   0.2972
   2.000   1.1496   0.01217   0.00453  -0.2010   0.4259   0.3112
   2.250   1.1758   0.01233   0.00471  -0.2007   0.4205   0.3240
   2.500   1.2015   0.01251   0.00490  -0.2003   0.4152   0.3364
   2.750   1.2265   0.01274   0.00511  -0.1999   0.4102   0.3493
   3.000   1.2517   0.01294   0.00533  -0.1994   0.4055   0.3646
   3.250   1.2776   0.01309   0.00552  -0.1991   0.4012   0.3791
   3.500   1.3029   0.01327   0.00573  -0.1987   0.3974   0.3949
   3.750   1.3278   0.01347   0.00596  -0.1982   0.3944   0.4131
   4.000   1.3522   0.01369   0.00621  -0.1977   0.3915   0.4343
   4.250   1.3761   0.01392   0.00648  -0.1971   0.3887   0.4577
   4.500   1.4015   0.01406   0.00670  -0.1967   0.3867   0.4799
   4.750   1.4264   0.01422   0.00693  -0.1962   0.3843   0.4983
   5.000   1.4507   0.01440   0.00719  -0.1957   0.3816   0.5194
   5.250   1.4743   0.01460   0.00746  -0.1950   0.3786   0.5449
   5.500   1.4970   0.01484   0.00777  -0.1943   0.3754   0.5773
   5.750   1.5180   0.01509   0.00812  -0.1932   0.3723   0.6254
   6.000   1.5385   0.01530   0.00843  -0.1920   0.3699   0.6760
   6.250   1.5613   0.01531   0.00873  -0.1912   0.3675   0.7805
   6.500   1.5792   0.01538   0.00897  -0.1894   0.3648   1.0000
   6.750   1.6000   0.01570   0.00931  -0.1883   0.3611   1.0000
   7.000   1.6192   0.01609   0.00970  -0.1870   0.3570   1.0000
   7.250   1.6376   0.01654   0.01013  -0.1856   0.3529   1.0000
   7.500   1.6588   0.01685   0.01049  -0.1847   0.3490   1.0000
   7.750   1.6784   0.01725   0.01092  -0.1836   0.3445   1.0000
   8.000   1.6962   0.01775   0.01142  -0.1822   0.3397   1.0000
   8.250   1.7139   0.01827   0.01195  -0.1809   0.3352   1.0000
   8.500   1.7333   0.01871   0.01244  -0.1798   0.3302   1.0000
   8.750   1.7493   0.01935   0.01308  -0.1783   0.3232   1.0000
   9.000   1.7657   0.01997   0.01371  -0.1770   0.3158   1.0000
   9.250   1.7804   0.02072   0.01445  -0.1754   0.3074   1.0000
   9.500   1.7954   0.02147   0.01522  -0.1739   0.2996   1.0000
   9.750   1.8075   0.02241   0.01615  -0.1722   0.2904   1.0000
  10.000   1.8194   0.02340   0.01712  -0.1704   0.2799   1.0000
  10.250   1.8282   0.02461   0.01830  -0.1684   0.2684   1.0000
  10.500   1.8347   0.02602   0.01966  -0.1662   0.2560   1.0000
  10.750   1.8389   0.02762   0.02121  -0.1638   0.2434   1.0000
  11.000   1.8452   0.02912   0.02268  -0.1618   0.2324   1.0000
  11.250   1.8515   0.03064   0.02420  -0.1598   0.2236   1.0000
  11.500   1.8556   0.03237   0.02591  -0.1577   0.2144   1.0000
  11.750   1.8617   0.03397   0.02753  -0.1558   0.2059   1.0000
  12.000   1.8640   0.03595   0.02949  -0.1538   0.1970   1.0000
  12.250   1.8684   0.03779   0.03134  -0.1520   0.1866   1.0000
  12.500   1.8682   0.04012   0.03364  -0.1500   0.1733   1.0000
  12.750   1.8613   0.04317   0.03661  -0.1478   0.1536   1.0000
  13.250   1.8333   0.05108   0.04432  -0.1431   0.1107   1.0000
  13.500   1.8192   0.05530   0.04849  -0.1412   0.0948   1.0000
  13.750   1.7977   0.06048   0.05359  -0.1393   0.0726   1.0000
  14.000   1.7716   0.06644   0.05946  -0.1376   0.0501   1.0000
  14.250   1.7583   0.07103   0.06407  -0.1365   0.0419   1.0000
  14.500   1.7505   0.07502   0.06811  -0.1357   0.0379   1.0000
  14.750   1.7447   0.07886   0.07203  -0.1350   0.0356   1.0000
  15.000   1.7393   0.08271   0.07596  -0.1345   0.0338   1.0000
  15.250   1.7330   0.08672   0.08005  -0.1341   0.0323   1.0000
  15.500   1.7276   0.09064   0.08406  -0.1338   0.0312   1.0000
  15.750   1.7229   0.09453   0.08805  -0.1337   0.0302   1.0000
  16.000   1.7169   0.09867   0.09228  -0.1336   0.0292   1.0000
  16.250   1.7099   0.10300   0.09670  -0.1338   0.0284   1.0000
  16.500   1.7023   0.10747   0.10126  -0.1341   0.0276   1.0000
<< Back to GOE 226 (MVA H.36) AIRFOIL (goe226-il)

Polar data table (+)

Polar graphs


<< Back to GOE 226 (MVA H.36) AIRFOIL (goe226-il)