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USA 45 AIRFOIL (usa45-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: USA 45 AIRFOIL (usa45-il)
Reynolds number: 50,000
Max Cl/Cd: 25.62 at α=3.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa45-il-50000-n5.txt
Download as CSV file: xf-usa45-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 45 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.3001   0.12413   0.11731  -0.0291   1.0000   0.1322
 -10.750  -0.3131   0.12193   0.11522  -0.0318   1.0000   0.1338
 -10.500  -0.3199   0.11891   0.11229  -0.0336   1.0000   0.1341
 -10.000  -0.3180   0.10246   0.09582  -0.0394   1.0000   0.0746
  -9.750  -0.3176   0.09885   0.09228  -0.0400   1.0000   0.0741
  -9.500  -0.3213   0.09508   0.08860  -0.0410   1.0000   0.0739
  -9.250  -0.3268   0.09130   0.08492  -0.0419   1.0000   0.0736
  -9.000  -0.3356   0.08741   0.08114  -0.0431   1.0000   0.0734
  -8.750  -0.3485   0.08341   0.07726  -0.0443   1.0000   0.0731
  -8.500  -0.3660   0.07985   0.07382  -0.0445   1.0000   0.0725
  -8.250  -0.3871   0.07685   0.07093  -0.0429   1.0000   0.0721
  -8.000  -0.4066   0.07379   0.06794  -0.0411   1.0000   0.0716
  -7.750  -0.4266   0.07094   0.06513  -0.0388   1.0000   0.0712
  -7.500  -0.4461   0.06817   0.06235  -0.0361   1.0000   0.0709
  -7.250  -0.4426   0.06303   0.05699  -0.0390   0.9891   0.0704
  -7.000  -0.4260   0.05723   0.05077  -0.0436   0.9731   0.0703
  -6.750  -0.4074   0.05245   0.04551  -0.0466   0.9567   0.0716
  -6.500  -0.3868   0.04821   0.04066  -0.0486   0.9403   0.0735
  -6.250  -0.3619   0.04446   0.03629  -0.0502   0.9246   0.0746
  -6.000  -0.3327   0.04120   0.03242  -0.0516   0.9096   0.0754
  -5.750  -0.2997   0.03856   0.02957  -0.0533   0.8947   0.0773
  -5.500  -0.2683   0.03673   0.02757  -0.0543   0.8774   0.0805
  -5.250  -0.2375   0.03480   0.02528  -0.0548   0.8590   0.0835
  -5.000  -0.2052   0.03289   0.02297  -0.0552   0.8401   0.0856
  -4.750  -0.1712   0.03119   0.02085  -0.0557   0.8207   0.0881
  -4.500  -0.1370   0.02971   0.01921  -0.0563   0.8008   0.0920
  -4.250  -0.1059   0.02862   0.01798  -0.0563   0.7783   0.0975
  -4.000  -0.0732   0.02756   0.01661  -0.0563   0.7556   0.1031
  -3.750  -0.0412   0.02653   0.01548  -0.0563   0.7340   0.1089
  -3.500  -0.0128   0.02580   0.01452  -0.0557   0.7105   0.1183
  -3.250   0.0141   0.02509   0.01371  -0.0551   0.6889   0.1331
  -3.000   0.0396   0.02437   0.01298  -0.0543   0.6694   0.1542
  -2.500   0.0806   0.02270   0.01189  -0.0517   0.6337   0.2921
  -2.250   0.0923   0.02139   0.01173  -0.0483   0.6193   0.5159
  -2.000   0.1549   0.02123   0.01225  -0.0509   0.6002   0.8374
  -1.750   0.2463   0.02193   0.01234  -0.0602   0.5794   0.9481
  -1.500   0.3172   0.02205   0.01192  -0.0683   0.5626   0.9847
  -1.250   0.3643   0.02206   0.01154  -0.0725   0.5496   1.0000
  -1.000   0.3842   0.02218   0.01139  -0.0715   0.5401   1.0000
  -0.750   0.4036   0.02235   0.01137  -0.0704   0.5304   1.0000
  -0.500   0.4243   0.02253   0.01128  -0.0693   0.5228   1.0000
  -0.250   0.4434   0.02278   0.01140  -0.0682   0.5138   1.0000
   0.000   0.4642   0.02300   0.01139  -0.0671   0.5071   1.0000
   0.250   0.4834   0.02332   0.01159  -0.0659   0.4994   1.0000
   0.500   0.5039   0.02360   0.01169  -0.0648   0.4928   1.0000
   0.750   0.5238   0.02394   0.01191  -0.0636   0.4862   1.0000
   1.000   0.5435   0.02431   0.01217  -0.0624   0.4798   1.0000
   1.250   0.5652   0.02464   0.01230  -0.0614   0.4747   1.0000
   1.500   0.5836   0.02512   0.01276  -0.0601   0.4685   1.0000
   1.750   0.6036   0.02555   0.01309  -0.0589   0.4628   1.0000
   2.000   0.6259   0.02592   0.01329  -0.0579   0.4583   1.0000
   2.250   0.6431   0.02651   0.01390  -0.0565   0.4524   1.0000
   2.500   0.6626   0.02704   0.01437  -0.0552   0.4473   1.0000
   2.750   0.6846   0.02749   0.01470  -0.0543   0.4432   1.0000
   3.000   0.7029   0.02813   0.01532  -0.0530   0.4384   1.0000
   3.250   0.7200   0.02881   0.01604  -0.0515   0.4333   1.0000
   3.500   0.7405   0.02935   0.01652  -0.0504   0.4289   1.0000
   3.750   0.7641   0.02982   0.01685  -0.0497   0.4254   1.0000
   4.000   0.7763   0.03081   0.01798  -0.0477   0.4203   1.0000
   4.250   0.7928   0.03161   0.01882  -0.0463   0.4158   1.0000
   4.500   0.8134   0.03223   0.01940  -0.0452   0.4120   1.0000
   4.750   0.8369   0.03276   0.01983  -0.0446   0.4088   1.0000
   5.000   0.8437   0.03408   0.02135  -0.0421   0.4036   1.0000
   5.250   0.8581   0.03505   0.02239  -0.0405   0.3992   1.0000
   5.500   0.8781   0.03576   0.02309  -0.0395   0.3957   1.0000
   5.750   0.9032   0.03627   0.02352  -0.0390   0.3928   1.0000
   6.000   0.9001   0.03820   0.02570  -0.0358   0.3874   1.0000
   6.250   0.9094   0.03950   0.02711  -0.0338   0.3830   1.0000
   6.500   0.9286   0.04028   0.02790  -0.0328   0.3794   1.0000
   6.750   0.9549   0.04075   0.02833  -0.0325   0.3767   1.0000
   7.000   0.9302   0.04399   0.03186  -0.0278   0.3708   1.0000
   7.250   0.9307   0.04584   0.03381  -0.0254   0.3662   1.0000
   7.500   0.9508   0.04656   0.03456  -0.0245   0.3629   1.0000
   7.750   0.9606   0.04795   0.03599  -0.0230   0.3593   1.0000
   8.000   0.8560   0.05685   0.04508  -0.0156   0.3498   1.0000
   8.250   0.8805   0.05718   0.04545  -0.0149   0.3470   1.0000
   8.500   0.9199   0.05635   0.04462  -0.0145   0.3450   1.0000
   9.000   0.8378   0.06996   0.05837  -0.0140   0.3291   1.0000
  10.000   0.7769   0.09044   0.07902  -0.0177   0.2999   1.0000
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