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USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 1,000,000
Max Cl/Cd: 97.25 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa46-il-1000000.txt
Download as CSV file: xf-usa46-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 46 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.7109   0.02778   0.02533  -0.0583   1.0000   0.0065
  -8.750  -0.6949   0.02441   0.02160  -0.0591   1.0000   0.0067
  -8.500  -0.6762   0.02192   0.01879  -0.0592   1.0000   0.0069
  -8.250  -0.6548   0.02043   0.01708  -0.0589   1.0000   0.0071
  -8.000  -0.6323   0.01932   0.01580  -0.0585   1.0000   0.0072
  -7.750  -0.6139   0.01622   0.01223  -0.0582   1.0000   0.0075
  -7.500  -0.5933   0.01404   0.00972  -0.0576   1.0000   0.0079
  -7.250  -0.5700   0.01310   0.00865  -0.0570   1.0000   0.0084
  -7.000  -0.5461   0.01246   0.00792  -0.0564   1.0000   0.0090
  -6.750  -0.5219   0.01198   0.00737  -0.0558   1.0000   0.0097
  -6.500  -0.4977   0.01149   0.00679  -0.0551   1.0000   0.0104
  -6.250  -0.4734   0.01107   0.00630  -0.0544   1.0000   0.0109
  -6.000  -0.4500   0.01005   0.00510  -0.0536   1.0000   0.0118
  -5.750  -0.4261   0.00946   0.00444  -0.0528   1.0000   0.0133
  -5.500  -0.4021   0.00913   0.00406  -0.0520   1.0000   0.0147
  -5.250  -0.3730   0.00885   0.00373  -0.0523   0.9993   0.0161
  -5.000  -0.3389   0.00834   0.00321  -0.0537   0.9975   0.0213
  -4.750  -0.3043   0.00802   0.00290  -0.0552   0.9955   0.0282
  -4.500  -0.2696   0.00792   0.00279  -0.0567   0.9935   0.0325
  -4.250  -0.2367   0.00759   0.00247  -0.0578   0.9905   0.0392
  -4.000  -0.2027   0.00746   0.00233  -0.0591   0.9875   0.0433
  -3.750  -0.1676   0.00719   0.00205  -0.0606   0.9847   0.0492
  -3.500  -0.1332   0.00699   0.00185  -0.0620   0.9805   0.0551
  -3.250  -0.0993   0.00673   0.00163  -0.0632   0.9739   0.0663
  -3.000  -0.0639   0.00648   0.00143  -0.0648   0.9672   0.0821
  -2.750  -0.0292   0.00621   0.00129  -0.0663   0.9575   0.1133
  -2.500   0.0056   0.00608   0.00118  -0.0677   0.9439   0.1333
  -2.250   0.0379   0.00599   0.00108  -0.0685   0.9249   0.1449
  -2.000   0.0666   0.00597   0.00100  -0.0684   0.9046   0.1543
  -1.750   0.0938   0.00596   0.00092  -0.0681   0.8850   0.1626
  -1.500   0.1205   0.00594   0.00086  -0.0676   0.8649   0.1703
  -1.250   0.1470   0.00596   0.00080  -0.0671   0.8435   0.1778
  -1.000   0.1736   0.00593   0.00075  -0.0667   0.8228   0.1898
  -0.750   0.2003   0.00590   0.00071  -0.0663   0.8032   0.2056
  -0.500   0.2267   0.00581   0.00067  -0.0659   0.7818   0.2427
  -0.250   0.2521   0.00548   0.00065  -0.0655   0.7580   0.3777
   0.000   0.2767   0.00505   0.00068  -0.0650   0.7328   0.5632
   0.250   0.3012   0.00485   0.00074  -0.0642   0.7055   0.6844
   0.500   0.3265   0.00488   0.00079  -0.0635   0.6738   0.7423
   0.750   0.3524   0.00497   0.00083  -0.0629   0.6429   0.7726
   1.000   0.3783   0.00505   0.00087  -0.0623   0.6138   0.7955
   1.250   0.4041   0.00512   0.00091  -0.0617   0.5860   0.8184
   1.500   0.4292   0.00513   0.00095  -0.0610   0.5620   0.8507
   1.750   0.4668   0.00502   0.00099  -0.0631   0.5299   1.0000
   2.000   0.4936   0.00518   0.00106  -0.0627   0.5062   1.0000
   2.250   0.5203   0.00535   0.00114  -0.0624   0.4783   1.0000
   2.500   0.5459   0.00566   0.00123  -0.0620   0.4190   1.0000
   2.750   0.5705   0.00612   0.00137  -0.0614   0.3465   1.0000
   3.000   0.5963   0.00645   0.00153  -0.0610   0.3079   1.0000
   3.250   0.6225   0.00672   0.00169  -0.0607   0.2765   1.0000
   3.500   0.6486   0.00699   0.00184  -0.0604   0.2441   1.0000
   3.750   0.6740   0.00739   0.00203  -0.0599   0.1967   1.0000
   4.000   0.6953   0.00845   0.00251  -0.0591   0.0727   1.0000
   4.250   0.7192   0.00918   0.00298  -0.0583   0.0198   1.0000
   4.500   0.7451   0.00954   0.00334  -0.0579   0.0161   1.0000
   4.750   0.7705   0.01004   0.00394  -0.0572   0.0138   1.0000
   5.000   0.7965   0.01037   0.00431  -0.0568   0.0132   1.0000
   5.250   0.8221   0.01075   0.00472  -0.0563   0.0119   1.0000
   5.500   0.8474   0.01119   0.00520  -0.0558   0.0109   1.0000
   5.750   0.8717   0.01180   0.00586  -0.0551   0.0101   1.0000
   6.000   0.8931   0.01295   0.00714  -0.0539   0.0093   1.0000
   6.250   0.9152   0.01396   0.00824  -0.0528   0.0089   1.0000
   6.500   0.9398   0.01447   0.00881  -0.0522   0.0085   1.0000
   6.750   0.9633   0.01525   0.00966  -0.0514   0.0081   1.0000
   7.000   0.9856   0.01631   0.01083  -0.0504   0.0077   1.0000
   7.250   1.0081   0.01733   0.01193  -0.0494   0.0073   1.0000
   7.500   1.0310   0.01819   0.01286  -0.0487   0.0070   1.0000
   7.750   1.0537   0.01903   0.01378  -0.0479   0.0066   1.0000
   8.000   1.0751   0.02026   0.01512  -0.0470   0.0065   1.0000
   8.250   1.0956   0.02169   0.01667  -0.0459   0.0063   1.0000
   8.500   1.1145   0.02353   0.01870  -0.0447   0.0062   1.0000
   8.750   1.1311   0.02584   0.02122  -0.0433   0.0061   1.0000
   9.000   1.1435   0.02902   0.02469  -0.0415   0.0060   1.0000
   9.250   1.1489   0.03332   0.02941  -0.0392   0.0059   1.0000
   9.500   1.1433   0.03902   0.03559  -0.0361   0.0058   1.0000
   9.750   1.1403   0.04310   0.03999  -0.0335   0.0058   1.0000
  10.000   1.1394   0.04606   0.04316  -0.0313   0.0058   1.0000
  10.250   1.1303   0.04923   0.04655  -0.0285   0.0059   1.0000
  10.500   1.1227   0.05107   0.04851  -0.0254   0.0059   1.0000
  10.750   0.9061   0.07024   0.06848  -0.0305   0.0090   1.0000
  11.000   0.8853   0.07842   0.07676  -0.0355   0.0090   1.0000
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