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USA 50 AIRFOIL (usa50-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: USA 50 AIRFOIL (usa50-il)
Reynolds number: 1,000,000
Max Cl/Cd: 92.27 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa50-il-1000000.txt
Download as CSV file: xf-usa50-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5155   0.08110   0.07957  -0.0167   1.0000   0.0074
  -7.750  -0.5264   0.07810   0.07661  -0.0169   1.0000   0.0075
  -7.500  -0.5297   0.07375   0.07226  -0.0197   1.0000   0.0075
  -7.250  -0.5303   0.06916   0.06766  -0.0225   1.0000   0.0075
  -7.000  -0.5285   0.06460   0.06307  -0.0248   1.0000   0.0077
  -6.750  -0.5253   0.05987   0.05828  -0.0265   1.0000   0.0079
  -6.500  -0.5203   0.05535   0.05369  -0.0273   1.0000   0.0081
  -6.250  -0.5142   0.05083   0.04906  -0.0274   1.0000   0.0083
  -6.000  -0.4956   0.04557   0.04362  -0.0293   0.9991   0.0090
  -5.750  -0.4625   0.04215   0.04006  -0.0314   0.9975   0.0099
  -5.500  -0.4577   0.02014   0.01658  -0.0319   0.9930   0.0062
  -5.250  -0.4332   0.01650   0.01241  -0.0315   0.9904   0.0063
  -5.000  -0.4030   0.01551   0.01128  -0.0324   0.9883   0.0070
  -4.750  -0.3714   0.01461   0.01024  -0.0335   0.9866   0.0078
  -4.500  -0.3393   0.01342   0.00886  -0.0344   0.9851   0.0085
  -4.250  -0.3069   0.01226   0.00750  -0.0354   0.9840   0.0089
  -4.000  -0.2736   0.01144   0.00656  -0.0365   0.9830   0.0094
  -3.750  -0.2489   0.01061   0.00562  -0.0357   0.9788   0.0098
  -3.500  -0.2192   0.00985   0.00476  -0.0360   0.9760   0.0100
  -3.250  -0.1907   0.00852   0.00321  -0.0360   0.9735   0.0118
  -3.000  -0.1576   0.00807   0.00271  -0.0371   0.9718   0.0140
  -2.750  -0.1231   0.00775   0.00232  -0.0386   0.9704   0.0163
  -2.500  -0.0989   0.00738   0.00195  -0.0376   0.9640   0.0249
  -2.250  -0.0673   0.00706   0.00177  -0.0384   0.9605   0.0583
  -2.000  -0.0330   0.00686   0.00159  -0.0398   0.9578   0.0690
  -1.750  -0.0071   0.00668   0.00141  -0.0393   0.9489   0.0786
  -1.500   0.0278   0.00638   0.00119  -0.0409   0.9409   0.1015
  -1.250   0.0661   0.00541   0.00099  -0.0438   0.9312   0.3475
  -1.000   0.1071   0.00476   0.00085  -0.0471   0.9143   0.5233
  -0.750   0.1398   0.00430   0.00076  -0.0484   0.8881   0.6766
  -0.500   0.1599   0.00393   0.00074  -0.0464   0.8552   0.8136
  -0.250   0.2304   0.00385   0.00086  -0.0558   0.8139   0.9549
   0.000   0.2675   0.00421   0.00096  -0.0577   0.7547   0.9755
   0.250   0.2988   0.00449   0.00102  -0.0584   0.7073   0.9841
   0.500   0.3386   0.00473   0.00108  -0.0612   0.6647   0.9901
   0.750   0.3745   0.00497   0.00113  -0.0631   0.6219   0.9948
   1.000   0.4148   0.00519   0.00115  -0.0661   0.5748   0.9986
   1.250   0.4459   0.00533   0.00116  -0.0669   0.5493   1.0000
   1.500   0.4689   0.00542   0.00119  -0.0659   0.5332   1.0000
   1.750   0.4911   0.00557   0.00123  -0.0648   0.5082   1.0000
   2.000   0.5144   0.00567   0.00127  -0.0638   0.4866   1.0000
   2.250   0.5370   0.00582   0.00132  -0.0627   0.4537   1.0000
   2.500   0.5578   0.00612   0.00137  -0.0613   0.3846   1.0000
   2.750   0.5760   0.00668   0.00153  -0.0595   0.2940   1.0000
   3.000   0.5954   0.00718   0.00172  -0.0579   0.2198   1.0000
   3.250   0.6092   0.00824   0.00212  -0.0554   0.0697   1.0000
   3.500   0.6290   0.00879   0.00245  -0.0538   0.0175   1.0000
   3.750   0.6518   0.00906   0.00277  -0.0526   0.0140   1.0000
   4.000   0.6731   0.00951   0.00336  -0.0511   0.0112   1.0000
   4.250   0.6948   0.00991   0.00383  -0.0497   0.0108   1.0000
   4.500   0.7169   0.01026   0.00422  -0.0484   0.0102   1.0000
   4.750   0.7385   0.01065   0.00466  -0.0471   0.0094   1.0000
   5.000   0.7590   0.01118   0.00528  -0.0455   0.0088   1.0000
   5.250   0.7791   0.01174   0.00591  -0.0439   0.0082   1.0000
   5.500   0.7989   0.01234   0.00656  -0.0422   0.0075   1.0000
   5.750   0.8152   0.01333   0.00759  -0.0400   0.0066   1.0000
   6.000   0.8296   0.01484   0.00920  -0.0372   0.0063   1.0000
   6.250   0.8484   0.01602   0.01048  -0.0353   0.0064   1.0000
   6.500   0.8689   0.01711   0.01169  -0.0336   0.0067   1.0000
   7.250   0.9216   0.02460   0.02008  -0.0262   0.0085   1.0000
   7.500   0.9377   0.02612   0.02178  -0.0242   0.0079   1.0000
   7.750   0.9524   0.02766   0.02347  -0.0222   0.0074   1.0000
   8.000   0.9655   0.02947   0.02545  -0.0200   0.0071   1.0000
   8.250   0.9786   0.03104   0.02707  -0.0183   0.0067   1.0000
   8.500   0.9846   0.03393   0.03018  -0.0155   0.0064   1.0000
   8.750   0.9508   0.04354   0.04039  -0.0086   0.0060   1.0000
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