Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

USA 22 AIRFOIL (usa22-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: USA 22 AIRFOIL (usa22-il)
Reynolds number: 500,000
Max Cl/Cd: 96.54 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa22-il-500000-n5.txt
Download as CSV file: xf-usa22-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 22 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3712   0.08551   0.08346  -0.0193   1.0000   0.0068
  -8.000  -0.3716   0.08133   0.07932  -0.0214   1.0000   0.0068
  -7.500  -0.3911   0.02455   0.02138  -0.0778   0.9614   0.0064
  -7.250  -0.3634   0.02021   0.01636  -0.0805   0.9488   0.0066
  -7.000  -0.3343   0.01799   0.01372  -0.0818   0.9356   0.0069
  -6.750  -0.3061   0.01654   0.01193  -0.0823   0.9210   0.0071
  -6.500  -0.2801   0.01498   0.01004  -0.0823   0.9050   0.0075
  -6.250  -0.2540   0.01412   0.00897  -0.0820   0.8884   0.0080
  -6.000  -0.2278   0.01352   0.00820  -0.0817   0.8711   0.0085
  -5.750  -0.2015   0.01302   0.00754  -0.0813   0.8527   0.0093
  -5.500  -0.1753   0.01253   0.00684  -0.0808   0.8301   0.0101
  -5.250  -0.1493   0.01207   0.00613  -0.0803   0.8031   0.0106
  -5.000  -0.1236   0.01140   0.00522  -0.0797   0.7749   0.0120
  -4.500  -0.0706   0.01084   0.00431  -0.0788   0.7288   0.0152
  -4.250  -0.0437   0.01057   0.00391  -0.0785   0.7090   0.0187
  -4.000  -0.0166   0.01043   0.00366  -0.0782   0.6908   0.0231
  -3.750   0.0107   0.01030   0.00345  -0.0780   0.6739   0.0274
  -3.500   0.0383   0.01027   0.00333  -0.0778   0.6579   0.0320
  -3.250   0.0657   0.01016   0.00311  -0.0775   0.6410   0.0349
  -3.000   0.0931   0.01008   0.00295  -0.0774   0.6234   0.0388
  -2.750   0.1205   0.01004   0.00281  -0.0772   0.6060   0.0421
  -2.500   0.1480   0.01004   0.00270  -0.0770   0.5891   0.0445
  -2.250   0.1753   0.00984   0.00238  -0.0767   0.5717   0.0481
  -2.000   0.2027   0.00975   0.00219  -0.0765   0.5541   0.0509
  -1.750   0.2302   0.00969   0.00204  -0.0764   0.5375   0.0539
  -1.500   0.2577   0.00967   0.00191  -0.0762   0.5216   0.0569
  -1.250   0.2852   0.00964   0.00179  -0.0760   0.5067   0.0622
  -1.000   0.3127   0.00958   0.00169  -0.0759   0.4934   0.0725
  -0.750   0.3397   0.00939   0.00163  -0.0758   0.4813   0.1304
  -0.500   0.3670   0.00933   0.00161  -0.0757   0.4699   0.1666
  -0.250   0.3945   0.00928   0.00162  -0.0756   0.4593   0.2027
   0.000   0.4220   0.00930   0.00164  -0.0755   0.4491   0.2336
   0.250   0.4495   0.00935   0.00168  -0.0754   0.4385   0.2591
   0.500   0.4769   0.00941   0.00172  -0.0752   0.4272   0.2796
   0.750   0.5046   0.00946   0.00176  -0.0751   0.4177   0.2928
   1.000   0.5321   0.00953   0.00179  -0.0750   0.4103   0.3052
   1.250   0.5598   0.00955   0.00185  -0.0750   0.4039   0.3231
   1.500   0.5873   0.00961   0.00189  -0.0749   0.3971   0.3352
   1.750   0.6150   0.00965   0.00194  -0.0748   0.3906   0.3457
   2.000   0.6426   0.00970   0.00200  -0.0747   0.3841   0.3598
   2.250   0.6701   0.00971   0.00208  -0.0747   0.3786   0.3822
   2.500   0.6964   0.00936   0.00217  -0.0746   0.3725   0.5647
   2.750   0.7261   0.00855   0.00229  -0.0748   0.3664   1.0000
   3.000   0.7535   0.00865   0.00238  -0.0747   0.3600   1.0000
   3.250   0.7806   0.00880   0.00247  -0.0746   0.3530   1.0000
   3.500   0.8079   0.00892   0.00259  -0.0744   0.3455   1.0000
   3.750   0.8348   0.00908   0.00270  -0.0743   0.3372   1.0000
   4.000   0.8619   0.00922   0.00283  -0.0741   0.3288   1.0000
   4.250   0.8888   0.00939   0.00297  -0.0740   0.3206   1.0000
   4.500   0.9156   0.00956   0.00313  -0.0738   0.3114   1.0000
   4.750   0.9422   0.00976   0.00330  -0.0736   0.2992   1.0000
   5.000   0.9679   0.01006   0.00349  -0.0733   0.2781   1.0000
   5.250   0.9936   0.01036   0.00373  -0.0731   0.2591   1.0000
   5.750   1.0437   0.01113   0.00430  -0.0724   0.2180   1.0000
   6.000   1.0679   0.01161   0.00465  -0.0720   0.1950   1.0000
   6.250   1.0913   0.01220   0.00509  -0.0715   0.1644   1.0000
   6.500   1.1086   0.01357   0.00594  -0.0703   0.0845   1.0000
   6.750   1.1255   0.01500   0.00708  -0.0690   0.0170   1.0000
   7.000   1.1494   0.01547   0.00760  -0.0685   0.0137   1.0000
   7.250   1.1725   0.01604   0.00826  -0.0678   0.0107   1.0000
   7.500   1.1961   0.01650   0.00881  -0.0672   0.0099   1.0000
   7.750   1.2190   0.01703   0.00943  -0.0666   0.0090   1.0000
   8.000   1.2412   0.01763   0.01010  -0.0659   0.0082   1.0000
   8.250   1.2613   0.01844   0.01098  -0.0649   0.0073   1.0000
   8.500   1.2815   0.01918   0.01181  -0.0640   0.0067   1.0000
   8.750   1.3012   0.01994   0.01267  -0.0630   0.0063   1.0000
   9.000   1.3196   0.02078   0.01360  -0.0618   0.0059   1.0000
   9.250   1.3367   0.02168   0.01462  -0.0605   0.0056   1.0000
   9.500   1.3524   0.02264   0.01566  -0.0591   0.0053   1.0000
   9.750   1.3658   0.02371   0.01682  -0.0575   0.0051   1.0000
  10.000   1.3739   0.02505   0.01825  -0.0552   0.0049   1.0000
  10.250   1.3760   0.02657   0.01989  -0.0522   0.0047   1.0000
  10.500   1.3828   0.02785   0.02128  -0.0501   0.0045   1.0000
  10.750   1.3873   0.02939   0.02293  -0.0482   0.0043   1.0000
  11.000   1.3909   0.03116   0.02482  -0.0466   0.0041   1.0000
  11.250   1.3913   0.03341   0.02720  -0.0454   0.0040   1.0000
  11.500   1.3904   0.03605   0.02996  -0.0447   0.0039   1.0000
  11.750   1.3894   0.03898   0.03301  -0.0445   0.0038   1.0000
  12.000   1.3876   0.04222   0.03637  -0.0447   0.0037   1.0000
  12.250   1.3852   0.04568   0.03995  -0.0451   0.0036   1.0000
  12.500   1.3815   0.04938   0.04379  -0.0457   0.0035   1.0000
  12.750   1.3778   0.05316   0.04769  -0.0464   0.0035   1.0000
  13.000   1.3732   0.05709   0.05172  -0.0472   0.0034   1.0000
  13.250   1.3677   0.06118   0.05591  -0.0480   0.0034   1.0000
  13.500   1.3623   0.06527   0.06011  -0.0490   0.0033   1.0000
  13.750   1.3559   0.06949   0.06443  -0.0498   0.0033   1.0000
  14.000   1.3496   0.07377   0.06880  -0.0508   0.0032   1.0000
  14.250   1.3429   0.07812   0.07324  -0.0518   0.0032   1.0000
  14.500   1.3365   0.08247   0.07768  -0.0527   0.0031   1.0000
  14.750   1.3299   0.08683   0.08213  -0.0536   0.0031   1.0000
<< Back to USA 22 AIRFOIL (usa22-il)

Polar data table (+)

Polar graphs


<< Back to USA 22 AIRFOIL (usa22-il)