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GOE 559 AIRFOIL (goe559-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 559 AIRFOIL (goe559-il)
Reynolds number: 100,000
Max Cl/Cd: 40.4 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe559-il-100000.txt
Download as CSV file: xf-goe559-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 559 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3386   0.09445   0.09027  -0.0421   1.0000   0.0388
  -8.000  -0.3432   0.09173   0.08766  -0.0437   1.0000   0.0393
  -7.750  -0.3477   0.08916   0.08518  -0.0445   1.0000   0.0394
  -7.500  -0.3591   0.08659   0.08273  -0.0455   1.0000   0.0396
  -7.250  -0.3656   0.08394   0.08014  -0.0463   1.0000   0.0397
  -7.000  -0.3728   0.08160   0.07784  -0.0462   1.0000   0.0398
  -6.750  -0.3831   0.07969   0.07596  -0.0452   1.0000   0.0400
  -6.500  -0.3932   0.07818   0.07447  -0.0429   1.0000   0.0400
  -6.250  -0.4113   0.07744   0.07376  -0.0396   0.9999   0.0402
  -6.000  -0.3897   0.06975   0.06624  -0.0426   0.9932   0.0438
  -5.750  -0.3517   0.06411   0.06034  -0.0520   0.9809   0.0503
  -5.500  -0.3202   0.05802   0.05403  -0.0588   0.9674   0.0557
  -5.250  -0.2833   0.05445   0.04982  -0.0653   0.9515   0.0658
  -5.000  -0.2529   0.04840   0.04372  -0.0698   0.9403   0.0807
  -4.750  -0.2221   0.04381   0.03887  -0.0739   0.9275   0.1088
  -4.500  -0.1990   0.03981   0.03500  -0.0760   0.9129   0.1694
  -3.250  -0.0571   0.02566   0.01944  -0.0775   0.8222   0.2896
  -3.000  -0.0063   0.02409   0.01630  -0.0776   0.8039   0.1757
  -2.750   0.0331   0.02262   0.01420  -0.0766   0.7872   0.1247
  -2.500   0.0686   0.02223   0.01315  -0.0752   0.7712   0.0933
  -2.250   0.1008   0.02131   0.01194  -0.0743   0.7564   0.0771
  -2.000   0.1309   0.02045   0.01078  -0.0734   0.7425   0.0657
  -1.750   0.1572   0.01920   0.00958  -0.0725   0.7296   0.0586
  -1.500   0.1824   0.01874   0.00894  -0.0711   0.7176   0.0516
  -1.250   0.2049   0.01812   0.00811  -0.0698   0.7071   0.0481
  -1.000   0.2302   0.01779   0.00730  -0.0689   0.6977   0.0453
  -0.750   0.2551   0.01761   0.00684  -0.0681   0.6876   0.0456
  -0.500   0.3693   0.01463   0.00629  -0.0843   0.6762   1.0000
  -0.250   0.3924   0.01485   0.00609  -0.0836   0.6687   1.0000
   0.000   0.4134   0.01509   0.00616  -0.0828   0.6598   1.0000
   0.250   0.4364   0.01534   0.00613  -0.0822   0.6533   1.0000
   0.500   0.4582   0.01563   0.00634  -0.0815   0.6458   1.0000
   0.750   0.4824   0.01590   0.00645  -0.0811   0.6406   1.0000
   1.000   0.5034   0.01624   0.00682  -0.0804   0.6335   1.0000
   1.250   0.5274   0.01652   0.00709  -0.0800   0.6281   1.0000
   1.500   0.5495   0.01688   0.00751  -0.0794   0.6222   1.0000
   1.750   0.5722   0.01723   0.00796  -0.0789   0.6166   1.0000
   2.000   0.5901   0.01619   0.00665  -0.0753   0.5798   1.0000
   2.250   0.5969   0.01517   0.00518  -0.0697   0.5145   1.0000
   2.500   0.6100   0.01510   0.00496  -0.0666   0.4667   1.0000
   2.750   0.6200   0.01550   0.00481  -0.0632   0.3570   1.0000
   3.000   0.6186   0.01805   0.00617  -0.0589   0.0453   1.0000
   3.250   0.6398   0.01850   0.00679  -0.0575   0.0419   1.0000
   3.500   0.6602   0.01906   0.00763  -0.0559   0.0419   1.0000
   3.750   0.6788   0.01976   0.00861  -0.0540   0.0431   1.0000
   4.000   0.6961   0.02056   0.00968  -0.0520   0.0442   1.0000
   4.250   0.7130   0.02134   0.01076  -0.0498   0.0456   1.0000
   4.500   0.7278   0.02224   0.01192  -0.0473   0.0482   1.0000
   4.750   0.7384   0.02331   0.01322  -0.0443   0.0523   1.0000
   5.000   0.7388   0.02482   0.01486  -0.0400   0.0551   1.0000
   5.250   0.7600   0.02518   0.01541  -0.0379   0.0642   1.0000
   5.500   0.7787   0.02568   0.01608  -0.0353   0.0745   1.0000
   5.750   0.8001   0.02618   0.01669  -0.0327   0.0883   1.0000
   6.000   0.8317   0.02678   0.01728  -0.0313   0.1046   1.0000
   6.250   0.8703   0.02658   0.01740  -0.0301   0.1230   1.0000
   6.500   0.9223   0.02747   0.01829  -0.0310   0.1373   1.0000
   6.750   0.9804   0.02935   0.02016  -0.0328   0.1474   1.0000
   7.000   1.0293   0.03344   0.02416  -0.0346   0.1434   1.0000
   7.250   1.0501   0.03525   0.02633  -0.0323   0.1368   1.0000
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