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GOE 511 AIRFOIL (goe511-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 511 AIRFOIL (goe511-il)
Reynolds number: 50,000
Max Cl/Cd: 25.23 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe511-il-50000-n5.txt
Download as CSV file: xf-goe511-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 511 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.0598   0.12073   0.11443  -0.0669   0.9211   0.1355
  -8.750  -0.0689   0.11931   0.11305  -0.0708   0.9103   0.1381
  -8.500  -0.0696   0.11658   0.11038  -0.0719   0.8983   0.1390
  -8.250  -0.0295   0.11121   0.10497  -0.0726   0.8933   0.1437
  -8.000  -0.0269   0.10886   0.10266  -0.0729   0.8808   0.1481
  -7.750  -0.0446   0.10773   0.10158  -0.0745   0.8664   0.1520
  -7.500  -0.0683   0.10663   0.10052  -0.0747   0.8509   0.1527
  -7.250  -0.0432   0.10174   0.09564  -0.0754   0.8427   0.1542
  -6.750  -0.0349   0.09607   0.09000  -0.0737   0.8173   0.1565
  -6.500  -0.0601   0.08847   0.08215  -0.0778   0.8030   0.0968
  -6.000  -0.0649   0.08035   0.07382  -0.0784   0.7779   0.0872
  -5.750  -0.0532   0.07699   0.07036  -0.0785   0.7689   0.0867
  -5.500  -0.0444   0.07375   0.06701  -0.0782   0.7586   0.0863
  -5.250  -0.0356   0.07067   0.06381  -0.0776   0.7483   0.0858
  -5.000  -0.0206   0.06716   0.06014  -0.0777   0.7399   0.0848
  -4.750  -0.0101   0.06400   0.05681  -0.0770   0.7299   0.0837
  -4.500   0.0053   0.06035   0.05291  -0.0769   0.7214   0.0825
  -4.250   0.0214   0.05658   0.04882  -0.0767   0.7135   0.0814
  -4.000   0.0332   0.05352   0.04546  -0.0753   0.7031   0.0812
  -3.750   0.0613   0.04994   0.04148  -0.0760   0.6978   0.0825
  -3.500   0.0667   0.04820   0.03946  -0.0729   0.6849   0.0832
  -3.250   0.0946   0.04485   0.03559  -0.0730   0.6789   0.0840
  -3.000   0.1046   0.04324   0.03361  -0.0702   0.6673   0.0842
  -2.750   0.1337   0.04072   0.03055  -0.0700   0.6604   0.0849
  -2.500   0.1524   0.03929   0.02869  -0.0683   0.6505   0.0858
  -2.250   0.1806   0.03772   0.02664  -0.0678   0.6423   0.0880
  -2.000   0.2100   0.03676   0.02553  -0.0678   0.6342   0.0909
  -1.750   0.2327   0.03609   0.02463  -0.0668   0.6245   0.0936
  -1.500   0.2823   0.03447   0.02252  -0.0697   0.6186   0.0970
  -1.250   0.2989   0.03444   0.02234  -0.0679   0.6074   0.0993
  -1.000   0.3591   0.03336   0.02100  -0.0731   0.6008   0.1080
  -0.750   0.3863   0.03336   0.02092  -0.0732   0.5908   0.1144
  -0.500   0.4349   0.03265   0.02007  -0.0765   0.5834   0.1264
  -0.250   0.4596   0.03267   0.02020  -0.0759   0.5752   0.1419
   0.000   0.4865   0.03265   0.02032  -0.0756   0.5672   0.1705
   0.250   0.5249   0.03247   0.02004  -0.0769   0.5611   0.2355
   0.500   0.5291   0.03317   0.02074  -0.0730   0.5521   0.2650
   0.750   0.5584   0.03316   0.02065  -0.0731   0.5457   0.3134
   1.000   0.5808   0.03330   0.02078  -0.0723   0.5391   0.3562
   1.250   0.5922   0.03370   0.02125  -0.0698   0.5317   0.3840
   1.500   0.6257   0.03343   0.02092  -0.0708   0.5259   0.4152
   1.750   0.6511   0.03348   0.02093  -0.0705   0.5202   0.4370
   2.000   0.6527   0.03422   0.02178  -0.0665   0.5129   0.4522
   2.500   0.8564   0.03459   0.02275  -0.0954   0.4977   1.0000
   2.750   0.8576   0.03549   0.02358  -0.0912   0.4920   1.0000
   3.000   0.8779   0.03603   0.02393  -0.0900   0.4872   1.0000
   3.250   0.9149   0.03626   0.02387  -0.0915   0.4833   1.0000
   3.500   0.8914   0.03778   0.02550  -0.0836   0.4776   1.0000
   3.750   0.8839   0.03908   0.02680  -0.0784   0.4720   1.0000
   4.000   0.9013   0.03971   0.02729  -0.0769   0.4676   1.0000
   4.250   0.9339   0.03991   0.02728  -0.0776   0.4643   1.0000
   4.500   0.9041   0.04222   0.02972  -0.0698   0.4588   1.0000
   4.750   0.8711   0.04518   0.03281  -0.0625   0.4526   1.0000
   5.000   0.8784   0.04647   0.03403  -0.0602   0.4487   1.0000
   5.250   0.9051   0.04677   0.03420  -0.0600   0.4460   1.0000
   5.750   0.7964   0.05816   0.04595  -0.0465   0.4294   1.0000
   6.000   0.8204   0.05844   0.04611  -0.0458   0.4272   1.0000
   6.250   0.8504   0.05828   0.04581  -0.0456   0.4256   1.0000
   6.750   0.7519   0.07384   0.06169  -0.0400   0.4080   1.0000
   7.750   0.6842   0.09393   0.08194  -0.0369   0.3849   1.0000
   8.000   0.6766   0.09780   0.08583  -0.0364   0.3797   1.0000
   8.250   0.6850   0.09996   0.08795  -0.0359   0.3766   1.0000
   8.500   0.7016   0.10121   0.08914  -0.0354   0.3742   1.0000
   8.750   0.7247   0.10168   0.08955  -0.0347   0.3722   1.0000
   9.000   0.6879   0.10876   0.09675  -0.0347   0.3636   1.0000
   9.250   0.6965   0.11080   0.09877  -0.0342   0.3596   1.0000
   9.500   0.7158   0.11162   0.09953  -0.0336   0.3565   1.0000
   9.750   0.7414   0.11176   0.09961  -0.0329   0.3544   1.0000
  10.000   0.7089   0.11823   0.10618  -0.0332   0.3448   1.0000
  10.250   0.7237   0.11945   0.10737  -0.0327   0.3406   1.0000
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