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

GOE 571 AIRFOIL (goe571-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 571 AIRFOIL (goe571-il)
Reynolds number: 200,000
Max Cl/Cd: 34.03 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe571-il-200000-n5.txt
Download as CSV file: xf-goe571-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 571 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.000   0.5320   0.10959   0.10295  -0.1750   0.5856   0.0397
 -12.750   0.5375   0.10750   0.10085  -0.1758   0.5818   0.0405
 -12.500   0.5333   0.10526   0.09861  -0.1771   0.5788   0.0414
 -12.250   0.5448   0.10314   0.09644  -0.1774   0.5751   0.0416
 -12.000   0.5567   0.10127   0.09458  -0.1775   0.5716   0.0421
 -11.750   0.5644   0.09938   0.09271  -0.1778   0.5677   0.0426
 -11.500   0.5711   0.09749   0.09083  -0.1781   0.5638   0.0432
 -11.250   0.5765   0.09556   0.08889  -0.1783   0.5602   0.0438
 -11.000   0.5783   0.09350   0.08682  -0.1788   0.5570   0.0450
 -10.750   0.5643   0.09098   0.08430  -0.1803   0.5548   0.0456
 -10.500   0.5766   0.08907   0.08240  -0.1798   0.5512   0.0458
 -10.250   0.5860   0.08730   0.08066  -0.1795   0.5471   0.0461
 -10.000   0.5942   0.08560   0.07897  -0.1790   0.5427   0.0467
  -9.750   0.5993   0.08390   0.07726  -0.1786   0.5390   0.0477
  -9.500   0.6021   0.08215   0.07550  -0.1782   0.5356   0.0481
  -9.250   0.5998   0.08011   0.07347  -0.1781   0.5321   0.0494
  -9.000   0.5797   0.07702   0.07047  -0.1792   0.5301   0.0503
  -8.750   0.5838   0.07523   0.06871  -0.1783   0.5260   0.0504
  -8.500   0.5947   0.07403   0.06751  -0.1767   0.5214   0.0510
  -8.250   0.5973   0.07252   0.06598  -0.1756   0.5178   0.0516
  -8.000   0.5969   0.07099   0.06444  -0.1745   0.5147   0.0535
  -7.750   0.5617   0.06731   0.06087  -0.1755   0.5130   0.0551
  -7.500   0.5499   0.06482   0.05847  -0.1747   0.5100   0.0554
  -7.250   0.5580   0.06389   0.05755  -0.1722   0.5059   0.0558
  -7.000   0.5518   0.06229   0.05598  -0.1702   0.5026   0.0563
  -6.750   0.5339   0.06040   0.05411  -0.1680   0.4999   0.0565
  -6.500   0.5102   0.05876   0.05248  -0.1637   0.4976   0.0568
  -6.250   0.4828   0.05748   0.05123  -0.1577   0.4953   0.0569
  -5.250   0.3287   0.04318   0.03635  -0.1261   0.4880   0.0370
  -5.000   0.3237   0.04200   0.03511  -0.1213   0.4849   0.0368
  -4.750   0.3200   0.04080   0.03381  -0.1166   0.4820   0.0365
  -4.500   0.3157   0.03937   0.03224  -0.1118   0.4794   0.0362
  -4.250   0.3130   0.03801   0.03073  -0.1070   0.4766   0.0359
  -4.000   0.3081   0.03629   0.02884  -0.1017   0.4739   0.0356
  -3.750   0.3055   0.03479   0.02715  -0.0967   0.4708   0.0350
  -3.500   0.3035   0.03316   0.02527  -0.0915   0.4682   0.0347
  -3.250   0.3045   0.03181   0.02367  -0.0868   0.4655   0.0346
  -3.000   0.3075   0.03054   0.02209  -0.0823   0.4631   0.0343
  -2.750   0.3154   0.02967   0.02096  -0.0788   0.4607   0.0344
  -2.500   0.3249   0.02914   0.02029  -0.0756   0.4583   0.0348
  -2.250   0.3334   0.02866   0.01967  -0.0722   0.4558   0.0355
  -2.000   0.3425   0.02807   0.01891  -0.0688   0.4533   0.0355
  -1.750   0.3529   0.02754   0.01819  -0.0657   0.4508   0.0355
  -1.500   0.3656   0.02706   0.01752  -0.0630   0.4484   0.0355
  -1.250   0.3797   0.02665   0.01694  -0.0606   0.4461   0.0355
  -1.000   0.3960   0.02629   0.01640  -0.0586   0.4439   0.0355
  -0.750   0.4158   0.02593   0.01584  -0.0572   0.4416   0.0356
  -0.500   0.4276   0.02577   0.01562  -0.0546   0.4394   0.0358
  -0.250   0.4406   0.02569   0.01549  -0.0522   0.4370   0.0358
   0.000   0.4545   0.02562   0.01537  -0.0501   0.4346   0.0359
   0.250   0.4694   0.02556   0.01523  -0.0481   0.4325   0.0363
   0.500   0.4850   0.02554   0.01517  -0.0463   0.4303   0.0364
   0.750   0.5030   0.02550   0.01505  -0.0449   0.4285   0.0368
   1.000   0.5226   0.02548   0.01495  -0.0438   0.4266   0.0369
   1.250   0.5452   0.02538   0.01479  -0.0432   0.4249   0.0374
   1.500   0.5655   0.02540   0.01479  -0.0424   0.4232   0.0377
   1.750   0.5731   0.02570   0.01516  -0.0397   0.4209   0.0382
   2.000   0.5828   0.02601   0.01552  -0.0374   0.4186   0.0386
   2.250   0.5958   0.02632   0.01585  -0.0356   0.4167   0.0396
   2.500   0.6092   0.02666   0.01618  -0.0340   0.4146   0.0406
   2.750   0.6244   0.02698   0.01649  -0.0326   0.4128   0.0413
   3.000   0.6418   0.02722   0.01675  -0.0317   0.4111   0.0419
   3.250   0.6621   0.02744   0.01694  -0.0312   0.4094   0.0428
   3.500   0.6852   0.02763   0.01707  -0.0310   0.4078   0.0435
   3.750   0.7136   0.02771   0.01705  -0.0316   0.4061   0.0452
   4.000   0.7208   0.02848   0.01788  -0.0296   0.4042   0.0457
   4.250   0.7263   0.02938   0.01885  -0.0277   0.4021   0.0469
   4.500   0.7348   0.03026   0.01978  -0.0262   0.4000   0.0490
   5.000   0.7597   0.03187   0.02152  -0.0244   0.3962   0.0605
   5.250   0.7753   0.03259   0.02232  -0.0241   0.3944   0.0842
   5.500   1.0832   0.03238   0.02411  -0.0802   0.3917   1.0000
   5.750   1.1027   0.03282   0.02444  -0.0798   0.3901   1.0000
   6.000   1.1265   0.03310   0.02460  -0.0798   0.3886   1.0000
   6.250   1.1260   0.03459   0.02615  -0.0774   0.3869   1.0000
   6.500   1.1193   0.03653   0.02819  -0.0746   0.3849   1.0000
   6.750   1.1129   0.03859   0.03034  -0.0722   0.3825   1.0000
   7.000   1.1106   0.04049   0.03228  -0.0702   0.3800   1.0000
   7.250   1.1139   0.04211   0.03391  -0.0688   0.3780   1.0000
   7.500   1.1229   0.04331   0.03509  -0.0678   0.3760   1.0000
   7.750   1.1357   0.04429   0.03603  -0.0671   0.3744   1.0000
   8.000   1.1559   0.04472   0.03638  -0.0670   0.3730   1.0000
   8.250   1.1833   0.04464   0.03619  -0.0673   0.3717   1.0000
   8.500   1.0690   0.05633   0.04844  -0.0597   0.3642   1.0000
   8.750   1.0488   0.06059   0.05278  -0.0578   0.3603   1.0000
   9.000   1.0610   0.06177   0.05394  -0.0575   0.3586   1.0000
   9.250   1.0815   0.06215   0.05425  -0.0574   0.3574   1.0000
   9.500   1.1061   0.06211   0.05414  -0.0576   0.3563   1.0000
   9.750   1.1354   0.06164   0.05357  -0.0580   0.3555   1.0000
  11.750   0.8270   0.12052   0.11342  -0.0473   0.2940   1.0000
  12.000   0.8540   0.11979   0.11263  -0.0475   0.2949   1.0000
<< Back to GOE 571 AIRFOIL (goe571-il)

Polar data table (+)

Polar graphs


<< Back to GOE 571 AIRFOIL (goe571-il)