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USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 200,000
Max Cl/Cd: 65.11 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa50-il-200000.txt
Download as CSV file: xf-usa50-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 50 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4933   0.08713   0.08370  -0.0197   1.0000   0.0365
  -7.750  -0.5011   0.08393   0.08057  -0.0218   1.0000   0.0372
  -7.500  -0.5034   0.07973   0.07639  -0.0268   1.0000   0.0379
  -7.250  -0.5014   0.07552   0.07211  -0.0309   1.0000   0.0383
  -7.000  -0.4981   0.07176   0.06823  -0.0329   1.0000   0.0385
  -6.750  -0.4941   0.06814   0.06446  -0.0335   1.0000   0.0387
  -6.500  -0.4969   0.06129   0.05760  -0.0337   1.0000   0.0395
  -6.250  -0.4911   0.05765   0.05405  -0.0322   1.0000   0.0407
  -6.000  -0.4832   0.05473   0.05111  -0.0311   1.0000   0.0422
  -5.750  -0.4737   0.05153   0.04782  -0.0305   1.0000   0.0446
  -5.500  -0.4562   0.04966   0.04536  -0.0303   1.0000   0.0508
  -5.250  -0.4504   0.04400   0.03952  -0.0295   1.0000   0.0524
  -5.000  -0.4388   0.04118   0.03685  -0.0284   1.0000   0.0558
  -4.750  -0.4239   0.03851   0.03397  -0.0271   1.0000   0.0598
  -4.500  -0.4087   0.03554   0.03039  -0.0253   1.0000   0.0659
  -4.250  -0.3930   0.03244   0.02727  -0.0241   1.0000   0.0678
  -4.000  -0.3703   0.02620   0.02010  -0.0208   1.0000   0.0410
  -3.750  -0.3490   0.02268   0.01617  -0.0188   1.0000   0.0333
  -3.500  -0.3260   0.01946   0.01234  -0.0168   1.0000   0.0308
  -3.250  -0.3026   0.01764   0.01021  -0.0153   1.0000   0.0311
  -3.000  -0.2798   0.01683   0.00925  -0.0142   1.0000   0.0353
  -2.750  -0.2563   0.01583   0.00805  -0.0129   1.0000   0.0378
  -2.500  -0.2341   0.01437   0.00651  -0.0114   1.0000   0.0409
  -2.250  -0.2126   0.01371   0.00581  -0.0101   1.0000   0.0488
  -2.000  -0.1917   0.01302   0.00520  -0.0087   1.0000   0.0738
  -1.750  -0.1714   0.01244   0.00465  -0.0074   1.0000   0.1049
  -1.500  -0.1513   0.01207   0.00444  -0.0063   1.0000   0.1387
  -1.250  -0.0621   0.00913   0.00438  -0.0189   1.0000   1.0000
  -1.000  -0.0436   0.00925   0.00434  -0.0173   1.0000   1.0000
  -0.750  -0.0249   0.00940   0.00436  -0.0158   0.9999   1.0000
  -0.500   0.0239   0.00959   0.00440  -0.0206   0.9936   1.0000
  -0.250   0.0709   0.00969   0.00436  -0.0249   0.9860   1.0000
   0.000   0.1170   0.00973   0.00432  -0.0290   0.9779   1.0000
   0.250   0.1686   0.00968   0.00422  -0.0341   0.9694   1.0000
   0.500   0.2185   0.00947   0.00399  -0.0386   0.9576   1.0000
   0.750   0.2649   0.00921   0.00372  -0.0423   0.9446   1.0000
   1.000   0.3086   0.00894   0.00346  -0.0454   0.9314   1.0000
   1.250   0.3551   0.00869   0.00323  -0.0491   0.9152   1.0000
   1.500   0.4051   0.00844   0.00304  -0.0535   0.8945   1.0000
   1.750   0.4560   0.00825   0.00285  -0.0581   0.8653   1.0000
   2.000   0.4981   0.00820   0.00275  -0.0607   0.8231   1.0000
   2.250   0.5297   0.00833   0.00276  -0.0611   0.7745   1.0000
   2.500   0.5551   0.00858   0.00282  -0.0602   0.7293   1.0000
   2.750   0.5775   0.00887   0.00296  -0.0587   0.6898   1.0000
   3.000   0.5969   0.00919   0.00309  -0.0566   0.6415   1.0000
   3.250   0.6158   0.00953   0.00323  -0.0545   0.5939   1.0000
   3.500   0.6361   0.00985   0.00343  -0.0528   0.5574   1.0000
   3.750   0.6542   0.01026   0.00369  -0.0506   0.5052   1.0000
   4.000   0.6716   0.01066   0.00388  -0.0483   0.4285   1.0000
   4.250   0.6827   0.01167   0.00417  -0.0451   0.2824   1.0000
   4.500   0.6884   0.01381   0.00511  -0.0415   0.0634   1.0000
   4.750   0.7072   0.01472   0.00590  -0.0395   0.0355   1.0000
   5.000   0.7257   0.01564   0.00692  -0.0375   0.0300   1.0000
   5.250   0.7430   0.01668   0.00809  -0.0352   0.0285   1.0000
   5.500   0.7602   0.01783   0.00939  -0.0330   0.0276   1.0000
   5.750   0.7781   0.01923   0.01088  -0.0308   0.0272   1.0000
   6.000   0.7985   0.02082   0.01257  -0.0291   0.0267   1.0000
   6.250   0.8202   0.02215   0.01401  -0.0277   0.0244   1.0000
   6.500   0.8431   0.02428   0.01635  -0.0263   0.0241   1.0000
   6.750   0.8659   0.02725   0.01966  -0.0246   0.0255   1.0000
   7.000   0.8848   0.03103   0.02382  -0.0226   0.0278   1.0000
   7.250   0.9040   0.03644   0.02988  -0.0194   0.0385   1.0000
  12.250   0.6735   0.13006   0.12678  -0.0180   0.0520   1.0000
  12.500   0.6652   0.13503   0.13174  -0.0216   0.0518   1.0000
  12.750   0.6526   0.13958   0.13627  -0.0260   0.0513   1.0000
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