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

GOE 506 AIRFOIL (goe506-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 506 AIRFOIL (goe506-il)
Reynolds number: 200,000
Max Cl/Cd: 58.04 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe506-il-200000.txt
Download as CSV file: xf-goe506-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 506 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.2244   0.10902   0.10566  -0.0793   0.9710   0.0749
 -10.750  -0.2199   0.10681   0.10347  -0.0789   0.9653   0.0775
 -10.500  -0.2676   0.09469   0.09144  -0.0950   0.9596   0.0841
 -10.250  -0.2212   0.09623   0.09292  -0.0884   0.9594   0.0865
 -10.000  -0.1979   0.09437   0.09105  -0.0897   0.9581   0.0906
  -9.750  -0.0839   0.08008   0.07687  -0.0994   0.9428   0.1046
  -9.500  -0.2182   0.08465   0.08141  -0.0962   0.9467   0.0999
  -9.250  -0.1957   0.08268   0.07941  -0.0976   0.9449   0.1040
  -8.750  -0.4837   0.04800   0.04298  -0.0977   0.9067   0.0498
  -8.500  -0.4854   0.04454   0.03936  -0.0947   0.9018   0.0480
  -8.250  -0.4776   0.04184   0.03629  -0.0929   0.8989   0.0477
  -8.000  -0.4628   0.03909   0.03315  -0.0918   0.8970   0.0470
  -7.750  -0.4788   0.03830   0.03217  -0.0847   0.8905   0.0465
  -7.500  -0.4729   0.03662   0.03017  -0.0813   0.8866   0.0459
  -7.250  -0.4544   0.03482   0.02803  -0.0799   0.8840   0.0456
  -7.000  -0.4288   0.03323   0.02619  -0.0797   0.8822   0.0461
  -6.750  -0.3955   0.03193   0.02467  -0.0808   0.8807   0.0477
  -6.500  -0.4037   0.03185   0.02449  -0.0746   0.8733   0.0485
  -6.250  -0.3712   0.03056   0.02299  -0.0751   0.8687   0.0496
  -6.000  -0.3237   0.02912   0.02134  -0.0781   0.8660   0.0506
  -5.750  -0.2817   0.02710   0.01940  -0.0807   0.8645   0.0536
  -5.500  -0.2895   0.02715   0.01943  -0.0744   0.8552   0.0544
  -5.250  -0.2546   0.02618   0.01843  -0.0754   0.8521   0.0567
  -5.000  -0.2143   0.02515   0.01735  -0.0773   0.8500   0.0591
  -4.750  -0.2018   0.02486   0.01702  -0.0745   0.8438   0.0609
  -4.500  -0.1805   0.02404   0.01623  -0.0732   0.8381   0.0642
  -4.250  -0.1388   0.02302   0.01515  -0.0752   0.8354   0.0688
  -4.000  -0.0925   0.02183   0.01390  -0.0781   0.8339   0.0771
  -3.750  -0.0935   0.02201   0.01404  -0.0727   0.8244   0.0831
  -3.500  -0.0633   0.02072   0.01339  -0.0731   0.8214   0.2127
  -3.250  -0.0199   0.02018   0.01290  -0.0755   0.8198   0.2630
  -3.000   0.0255   0.01961   0.01231  -0.0783   0.8185   0.2909
  -2.750   0.0223   0.02009   0.01280  -0.0726   0.8087   0.2987
  -2.500   0.0635   0.01947   0.01214  -0.0747   0.8065   0.3137
  -2.250   0.1090   0.01873   0.01141  -0.0775   0.8049   0.3304
  -2.000   0.1573   0.01799   0.01067  -0.0810   0.8036   0.3467
  -1.750   0.1629   0.01807   0.01078  -0.0766   0.7951   0.3560
  -1.500   0.2009   0.01745   0.01019  -0.0781   0.7916   0.3769
  -1.250   0.2479   0.01659   0.00946  -0.0814   0.7894   0.4048
  -1.000   0.2995   0.01563   0.00865  -0.0856   0.7873   0.4445
  -0.750   0.2981   0.01549   0.00867  -0.0795   0.7771   0.4729
  -0.500   0.3418   0.01434   0.00793  -0.0820   0.7731   0.5677
  -0.250   0.3472   0.01377   0.00784  -0.0769   0.7637   0.6827
   0.000   0.5037   0.01258   0.00711  -0.1015   0.7554   0.9102
   0.250   0.5673   0.01250   0.00692  -0.1081   0.7409   0.9430
   0.500   0.6162   0.01250   0.00680  -0.1119   0.7229   0.9645
   0.750   0.6637   0.01252   0.00668  -0.1157   0.6975   0.9814
   1.000   0.7151   0.01251   0.00646  -0.1204   0.6526   0.9939
   1.250   0.7377   0.01271   0.00611  -0.1194   0.5671   1.0000
   1.500   0.7238   0.01318   0.00617  -0.1113   0.5249   1.0000
   1.750   0.7169   0.01360   0.00632  -0.1046   0.4995   1.0000
   2.000   0.7161   0.01395   0.00648  -0.0991   0.4802   1.0000
   2.250   0.7214   0.01428   0.00664  -0.0949   0.4649   1.0000
   2.500   0.7305   0.01460   0.00682  -0.0914   0.4526   1.0000
   2.750   0.7421   0.01489   0.00699  -0.0884   0.4420   1.0000
   3.000   0.7564   0.01517   0.00717  -0.0859   0.4332   1.0000
   3.250   0.7716   0.01546   0.00736  -0.0836   0.4257   1.0000
   3.500   0.7889   0.01574   0.00755  -0.0818   0.4191   1.0000
   3.750   0.8061   0.01600   0.00776  -0.0799   0.4128   1.0000
   4.000   0.8256   0.01632   0.00796  -0.0786   0.4075   1.0000
   4.250   0.8443   0.01657   0.00819  -0.0770   0.4022   1.0000
   4.500   0.8638   0.01683   0.00843  -0.0757   0.3976   1.0000
   4.750   0.8854   0.01713   0.00866  -0.0747   0.3935   1.0000
   5.000   0.9092   0.01746   0.00891  -0.0742   0.3898   1.0000
   5.250   0.9283   0.01771   0.00919  -0.0728   0.3857   1.0000
   5.500   0.9484   0.01799   0.00947  -0.0716   0.3817   1.0000
   5.750   0.9708   0.01829   0.00973  -0.0708   0.3782   1.0000
   6.000   0.9991   0.01866   0.01000  -0.0712   0.3750   1.0000
   6.250   1.0180   0.01894   0.01035  -0.0698   0.3718   1.0000
   6.500   1.0385   0.01924   0.01069  -0.0688   0.3686   1.0000
   6.750   1.0600   0.01955   0.01102  -0.0679   0.3654   1.0000
   7.000   1.0826   0.01988   0.01134  -0.0672   0.3621   1.0000
   7.250   1.1118   0.02028   0.01167  -0.0679   0.3586   1.0000
   7.500   1.1260   0.02059   0.01208  -0.0657   0.3551   1.0000
   7.750   1.1428   0.02091   0.01248  -0.0640   0.3511   1.0000
   8.000   1.1630   0.02126   0.01284  -0.0630   0.3475   1.0000
   8.250   1.1884   0.02164   0.01320  -0.0630   0.3442   1.0000
   8.500   1.2074   0.02204   0.01366  -0.0618   0.3405   1.0000
   8.750   1.2224   0.02242   0.01415  -0.0599   0.3370   1.0000
   9.000   1.2395   0.02280   0.01459  -0.0585   0.3330   1.0000
   9.250   1.2595   0.02319   0.01495  -0.0575   0.3286   1.0000
   9.500   1.2723   0.02365   0.01548  -0.0554   0.3235   1.0000
   9.750   1.2797   0.02408   0.01603  -0.0525   0.3176   1.0000
  10.000   1.2951   0.02451   0.01638  -0.0510   0.3113   1.0000
  10.250   1.2987   0.02508   0.01710  -0.0476   0.3052   1.0000
  10.500   1.3065   0.02561   0.01771  -0.0451   0.2988   1.0000
  10.750   1.3178   0.02619   0.01829  -0.0431   0.2928   1.0000
  11.000   1.3219   0.02688   0.01914  -0.0402   0.2857   1.0000
  11.250   1.3315   0.02754   0.01978  -0.0382   0.2794   1.0000
  11.500   1.3364   0.02837   0.02079  -0.0357   0.2716   1.0000
  11.750   1.3413   0.02924   0.02164  -0.0334   0.2636   1.0000
  12.000   1.3477   0.03017   0.02274  -0.0313   0.2535   1.0000
  12.250   1.3551   0.03115   0.02380  -0.0295   0.2446   1.0000
  12.500   1.3604   0.03229   0.02498  -0.0276   0.2336   1.0000
  12.750   1.3651   0.03359   0.02631  -0.0258   0.2197   1.0000
  13.000   1.3658   0.03525   0.02792  -0.0239   0.2031   1.0000
  13.250   1.3625   0.03731   0.02990  -0.0218   0.1892   1.0000
  13.500   1.3578   0.03963   0.03216  -0.0198   0.1784   1.0000
  13.750   1.3530   0.04210   0.03461  -0.0180   0.1695   1.0000
  14.000   1.3492   0.04460   0.03714  -0.0164   0.1600   1.0000
  14.250   1.3420   0.04751   0.04006  -0.0149   0.1512   1.0000
  14.500   1.3392   0.05019   0.04281  -0.0139   0.1414   1.0000
  14.750   1.3362   0.05304   0.04572  -0.0131   0.1264   1.0000
  15.000   1.3185   0.05758   0.05009  -0.0123   0.0819   1.0000
  15.250   1.2971   0.06276   0.05520  -0.0119   0.0684   1.0000
  15.750   1.2505   0.07448   0.06693  -0.0124   0.0401   1.0000
  16.000   1.2370   0.07945   0.07197  -0.0131   0.0372   1.0000
  16.250   1.2222   0.08475   0.07738  -0.0140   0.0350   1.0000
  16.500   1.2092   0.08987   0.08262  -0.0150   0.0338   1.0000
<< Back to GOE 506 AIRFOIL (goe506-il)

Polar data table (+)

Polar graphs


<< Back to GOE 506 AIRFOIL (goe506-il)