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USA 5 AIRFOIL (usa5-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: USA 5 AIRFOIL (usa5-il)
Reynolds number: 200,000
Max Cl/Cd: 70.96 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa5-il-200000-n5.txt
Download as CSV file: xf-usa5-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 5 AIRFOIL                                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3260   0.09861   0.09539  -0.0260   1.0000   0.0185
  -7.500  -0.3323   0.09681   0.09366  -0.0244   1.0000   0.0185
  -7.250  -0.3380   0.09480   0.09172  -0.0231   1.0000   0.0185
  -7.000  -0.3406   0.09253   0.08949  -0.0223   1.0000   0.0185
  -6.750  -0.3315   0.08913   0.08611  -0.0248   0.9981   0.0186
  -6.500  -0.3085   0.08454   0.08150  -0.0307   0.9936   0.0186
  -6.250  -0.2845   0.07979   0.07673  -0.0366   0.9888   0.0186
  -6.000  -0.2605   0.07515   0.07205  -0.0420   0.9838   0.0186
  -5.750  -0.2491   0.06969   0.06660  -0.0432   0.9796   0.0171
  -5.500  -0.2254   0.06499   0.06186  -0.0481   0.9734   0.0142
  -5.250  -0.1917   0.06007   0.05684  -0.0554   0.9678   0.0149
  -5.000  -0.1606   0.05549   0.05217  -0.0612   0.9599   0.0157
  -4.750  -0.1234   0.05024   0.04678  -0.0679   0.9547   0.0146
  -4.500  -0.0926   0.04541   0.04180  -0.0722   0.9458   0.0137
  -4.250  -0.0557   0.04006   0.03621  -0.0772   0.9404   0.0132
  -4.000  -0.0285   0.03539   0.03128  -0.0789   0.9307   0.0128
  -3.750   0.0005   0.03019   0.02573  -0.0803   0.9226   0.0126
  -3.500   0.0260   0.02339   0.01829  -0.0801   0.9144   0.0125
  -3.250   0.0484   0.01885   0.01297  -0.0787   0.9045   0.0131
  -3.000   0.0775   0.01691   0.01054  -0.0785   0.8955   0.0155
  -2.750   0.1071   0.01507   0.00821  -0.0783   0.8856   0.0172
  -2.500   0.1350   0.01393   0.00675  -0.0778   0.8742   0.0188
  -2.250   0.1632   0.01305   0.00572  -0.0776   0.8631   0.0258
  -2.000   0.1918   0.01232   0.00486  -0.0774   0.8516   0.0355
  -1.750   0.2205   0.01190   0.00434  -0.0773   0.8391   0.0511
  -1.500   0.2494   0.01180   0.00423  -0.0773   0.8255   0.0692
  -1.250   0.2779   0.01170   0.00401  -0.0771   0.8103   0.0832
  -1.000   0.3057   0.01155   0.00370  -0.0768   0.7935   0.0938
  -0.750   0.3330   0.01142   0.00342  -0.0765   0.7740   0.1026
  -0.500   0.3591   0.01130   0.00316  -0.0759   0.7509   0.1104
  -0.250   0.3851   0.01128   0.00299  -0.0752   0.7257   0.1209
   0.000   0.4099   0.01121   0.00280  -0.0745   0.7008   0.1301
   0.250   0.4344   0.01118   0.00264  -0.0736   0.6758   0.1361
   0.500   0.4585   0.01117   0.00252  -0.0727   0.6512   0.1416
   0.750   0.4823   0.01123   0.00242  -0.0717   0.6258   0.1480
   1.000   0.5057   0.01129   0.00238  -0.0707   0.5987   0.1575
   1.250   0.5289   0.01136   0.00236  -0.0697   0.5697   0.1722
   1.500   0.5517   0.01142   0.00236  -0.0686   0.5398   0.2010
   2.000   0.6749   0.01046   0.00259  -0.0846   0.4634   1.0000
   2.500   0.7217   0.01097   0.00284  -0.0829   0.4307   1.0000
   2.750   0.7454   0.01121   0.00300  -0.0821   0.4179   1.0000
   3.000   0.7693   0.01143   0.00319  -0.0813   0.4076   1.0000
   3.250   0.7931   0.01168   0.00338  -0.0805   0.3984   1.0000
   3.500   0.8173   0.01190   0.00358  -0.0798   0.3896   1.0000
   3.750   0.8413   0.01213   0.00381  -0.0791   0.3824   1.0000
   4.000   0.8655   0.01235   0.00407  -0.0784   0.3749   1.0000
   4.250   0.8894   0.01259   0.00433  -0.0777   0.3676   1.0000
   4.500   0.9125   0.01287   0.00459  -0.0768   0.3551   1.0000
   4.750   0.9345   0.01317   0.00483  -0.0758   0.3299   1.0000
   5.000   0.9564   0.01349   0.00512  -0.0748   0.3023   1.0000
   5.250   0.9772   0.01392   0.00540  -0.0736   0.2569   1.0000
   5.500   0.9883   0.01536   0.00611  -0.0712   0.1507   1.0000
   5.750   0.9957   0.01733   0.00731  -0.0681   0.0215   1.0000
   6.000   1.0161   0.01791   0.00803  -0.0667   0.0162   1.0000
   6.250   1.0360   0.01855   0.00887  -0.0651   0.0144   1.0000
   6.500   1.0531   0.01947   0.01001  -0.0632   0.0118   1.0000
   6.750   1.0686   0.02053   0.01130  -0.0610   0.0106   1.0000
   7.000   1.0847   0.02145   0.01240  -0.0590   0.0102   1.0000
   7.250   1.0982   0.02250   0.01362  -0.0566   0.0098   1.0000
   7.500   1.1095   0.02366   0.01490  -0.0539   0.0092   1.0000
   7.750   1.1198   0.02484   0.01618  -0.0512   0.0087   1.0000
   8.000   1.1291   0.02606   0.01745  -0.0485   0.0077   1.0000
   8.250   1.1351   0.02755   0.01905  -0.0453   0.0072   1.0000
   8.500   1.1397   0.02914   0.02068  -0.0419   0.0070   1.0000
   8.750   1.1466   0.03132   0.02284  -0.0390   0.0068   1.0000
   9.000   1.1626   0.03336   0.02491  -0.0375   0.0067   1.0000
   9.250   1.1890   0.03635   0.02795  -0.0378   0.0066   1.0000
   9.500   1.2079   0.03865   0.03042  -0.0367   0.0066   1.0000
   9.750   1.2260   0.04133   0.03328  -0.0356   0.0066   1.0000
  10.250   1.2508   0.04605   0.03843  -0.0320   0.0067   1.0000
  10.500   1.2568   0.04890   0.04155  -0.0295   0.0067   1.0000
  10.750   1.2335   0.04919   0.04219  -0.0229   0.0062   1.0000
  11.000   1.2274   0.05096   0.04424  -0.0196   0.0059   1.0000
  11.250   1.2230   0.05378   0.04729  -0.0172   0.0060   1.0000
  11.500   1.2144   0.05624   0.05002  -0.0150   0.0057   1.0000
  11.750   1.2056   0.05936   0.05338  -0.0134   0.0056   1.0000
  12.000   1.1962   0.06292   0.05715  -0.0124   0.0057   1.0000
  12.250   1.1833   0.06671   0.06118  -0.0121   0.0056   1.0000
  12.500   1.1694   0.07095   0.06565  -0.0125   0.0054   1.0000
  12.750   1.1561   0.07561   0.07050  -0.0135   0.0055   1.0000
  13.000   1.1416   0.08076   0.07584  -0.0151   0.0056   1.0000
  13.250   1.1253   0.08647   0.08175  -0.0176   0.0054   1.0000
  13.500   1.1116   0.09218   0.08761  -0.0202   0.0056   1.0000
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