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USA 45 AIRFOIL (usa45-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: USA 45 AIRFOIL (usa45-il)
Reynolds number: 50,000
Max Cl/Cd: 17.12 at α=0.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa45-il-50000.txt
Download as CSV file: xf-usa45-il-50000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2643   0.10268   0.09652  -0.0200   1.0000   0.2952
  -8.500  -0.2857   0.10234   0.09635  -0.0191   1.0000   0.3074
  -8.250  -0.2731   0.09895   0.09301  -0.0175   1.0000   0.3231
  -8.000  -0.2592   0.09552   0.08964  -0.0159   1.0000   0.3374
  -7.750  -0.2531   0.09265   0.08687  -0.0144   1.0000   0.3498
  -7.500  -0.2555   0.09043   0.08478  -0.0126   1.0000   0.3635
  -7.250  -0.2622   0.08861   0.08311  -0.0102   1.0000   0.3781
  -7.000  -0.2731   0.08715   0.08182  -0.0072   1.0000   0.3934
  -6.750  -0.3082   0.08773   0.08264  -0.0020   1.0000   0.4074
  -6.500  -0.3109   0.08594   0.08098   0.0018   1.0000   0.4239
  -6.250  -0.2851   0.08242   0.07752   0.0039   1.0000   0.4465
  -6.000  -0.2997   0.08184   0.07712   0.0088   1.0000   0.4637
  -5.750  -0.3517   0.08376   0.07931   0.0163   1.0000   0.4732
  -5.250  -0.3976   0.06422   0.05903  -0.0281   0.9557   0.2577
  -5.000  -0.3599   0.05251   0.04603  -0.0392   0.9393   0.1730
  -4.750  -0.3209   0.04821   0.04127  -0.0432   0.9173   0.1679
  -4.500  -0.2798   0.04409   0.03604  -0.0466   0.8972   0.1595
  -4.250  -0.2292   0.04075   0.03214  -0.0506   0.8783   0.1592
  -4.000  -0.1745   0.03785   0.02901  -0.0549   0.8605   0.1649
  -3.750  -0.1222   0.03541   0.02614  -0.0581   0.8428   0.1704
  -3.500  -0.0769   0.03345   0.02377  -0.0599   0.8235   0.1782
  -3.250  -0.0366   0.03201   0.02213  -0.0609   0.8035   0.1928
  -3.000   0.0092   0.03028   0.02047  -0.0626   0.7852   0.2142
  -2.750   0.0528   0.02858   0.01897  -0.0641   0.7678   0.2641
  -2.500   0.3215   0.02445   0.01660  -0.0937   0.7403   1.0000
  -2.250   0.3369   0.02456   0.01640  -0.0924   0.7227   1.0000
  -2.000   0.3533   0.02473   0.01629  -0.0911   0.7072   1.0000
  -1.750   0.3710   0.02491   0.01619  -0.0897   0.6937   1.0000
  -1.500   0.3873   0.02523   0.01628  -0.0883   0.6804   1.0000
  -1.250   0.4023   0.02572   0.01659  -0.0868   0.6677   1.0000
  -1.000   0.4186   0.02620   0.01688  -0.0853   0.6567   1.0000
  -0.750   0.4374   0.02657   0.01702  -0.0839   0.6472   1.0000
  -0.500   0.4500   0.02737   0.01774  -0.0822   0.6366   1.0000
  -0.250   0.4691   0.02782   0.01796  -0.0808   0.6284   1.0000
   0.000   0.4807   0.02875   0.01883  -0.0790   0.6188   1.0000
   0.250   0.4985   0.02938   0.01930  -0.0775   0.6116   1.0000
   0.500   0.5074   0.03057   0.02046  -0.0755   0.6034   1.0000
   0.750   0.5304   0.03098   0.02066  -0.0744   0.5973   1.0000
   1.000   0.5303   0.03277   0.02251  -0.0717   0.5891   1.0000
   1.250   0.5494   0.03343   0.02303  -0.0703   0.5828   1.0000
   1.500   0.5517   0.03523   0.02483  -0.0677   0.5764   1.0000
   1.750   0.5545   0.03701   0.02659  -0.0653   0.5704   1.0000
   2.000   0.5781   0.03760   0.02704  -0.0643   0.5654   1.0000
   2.250   0.5549   0.04104   0.03056  -0.0601   0.5593   1.0000
   2.500   0.5458   0.04367   0.03318  -0.0571   0.5539   1.0000
   2.750   0.5721   0.04432   0.03371  -0.0564   0.5495   1.0000
   3.000   0.5349   0.04898   0.03840  -0.0521   0.5464   1.0000
   3.250   0.4763   0.05478   0.04420  -0.0471   0.5456   1.0000
   3.500   0.4476   0.05894   0.04832  -0.0445   0.5455   1.0000
   3.750   0.4404   0.06221   0.05154  -0.0434   0.5466   1.0000
   4.000   0.4412   0.06525   0.05455  -0.0430   0.5489   1.0000
   4.250   0.2576   0.07701   0.06668  -0.0427   0.7157   1.0000
   4.500   0.2766   0.07956   0.06914  -0.0432   0.7094   1.0000
   4.750   0.2928   0.08134   0.07085  -0.0430   0.6971   1.0000
   5.000   0.2979   0.08274   0.07219  -0.0417   0.6855   1.0000
   5.250   0.3275   0.08604   0.07540  -0.0432   0.6778   1.0000
   5.500   0.3362   0.08735   0.07667  -0.0422   0.6646   1.0000
   5.750   0.3395   0.08885   0.07812  -0.0410   0.6541   1.0000
   6.000   0.3761   0.09271   0.08193  -0.0431   0.6463   1.0000
   6.250   0.3742   0.09341   0.08259  -0.0412   0.6328   1.0000
   6.500   0.3800   0.09518   0.08433  -0.0403   0.6217   1.0000
   6.750   0.4137   0.09895   0.08807  -0.0420   0.6140   1.0000
   7.000   0.4095   0.09976   0.08885  -0.0402   0.6014   1.0000
   7.250   0.4183   0.10196   0.09103  -0.0398   0.5916   1.0000
   7.500   0.4486   0.10540   0.09445  -0.0409   0.5818   1.0000
   7.750   0.4437   0.10644   0.09548  -0.0395   0.5699   1.0000
   8.000   0.4640   0.10979   0.09882  -0.0401   0.5630   1.0000
   8.250   0.4753   0.11171   0.10075  -0.0397   0.5506   1.0000
   8.500   0.4755   0.11342   0.10246  -0.0389   0.5393   1.0000
   8.750   0.5115   0.11797   0.10702  -0.0404   0.5325   1.0000
   9.000   0.5006   0.11847   0.10751  -0.0390   0.5203   1.0000
   9.250   0.5109   0.12131   0.11037  -0.0390   0.5127   1.0000
   9.500   0.5303   0.12412   0.11320  -0.0393   0.5016   1.0000
   9.750   0.5262   0.12579   0.11488  -0.0388   0.4920   1.0000
  10.000   0.5547   0.12981   0.11893  -0.0396   0.4847   1.0000
  10.250   0.5453   0.13093   0.12005  -0.0390   0.4747   1.0000
  10.500   0.5798   0.13576   0.12494  -0.0400   0.4678   1.0000
  10.750   0.5649   0.13609   0.12527  -0.0393   0.4572   1.0000
  11.000   0.5961   0.14091   0.13015  -0.0402   0.4518   1.0000
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