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USA 45 AIRFOIL (usa45-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: USA 45 AIRFOIL (usa45-il)
Reynolds number: 100,000
Max Cl/Cd: 32.99 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa45-il-100000.txt
Download as CSV file: xf-usa45-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2849   0.10398   0.09935  -0.0312   1.0000   0.1222
  -9.500  -0.2956   0.10144   0.09689  -0.0332   1.0000   0.1271
  -9.250  -0.3392   0.09936   0.09501  -0.0387   1.0000   0.1289
  -9.000  -0.2982   0.09477   0.09038  -0.0334   1.0000   0.1327
  -8.750  -0.2917   0.09224   0.08790  -0.0323   1.0000   0.1368
  -8.500  -0.3089   0.08978   0.08556  -0.0334   1.0000   0.1421
  -8.250  -0.3632   0.08802   0.08404  -0.0352   1.0000   0.1438
  -8.000  -0.4099   0.08685   0.08299  -0.0328   1.0000   0.1442
  -7.750  -0.3385   0.08217   0.07831  -0.0293   1.0000   0.1502
  -7.500  -0.3634   0.08090   0.07718  -0.0259   1.0000   0.1520
  -7.250  -0.3954   0.07999   0.07640  -0.0218   1.0000   0.1528
  -7.000  -0.4193   0.07455   0.07087  -0.0313   0.9880   0.1616
  -6.750  -0.3744   0.07113   0.06748  -0.0335   0.9807   0.1724
  -6.500  -0.3418   0.06631   0.06261  -0.0393   0.9693   0.1852
  -6.250  -0.3135   0.06157   0.05778  -0.0451   0.9554   0.1998
  -6.000  -0.2841   0.05727   0.05337  -0.0499   0.9410   0.2160
  -5.750  -0.2527   0.05346   0.04946  -0.0538   0.9266   0.2350
  -5.500  -0.2545   0.03562   0.02872  -0.0586   0.9085   0.0988
  -5.250  -0.2196   0.03212   0.02490  -0.0599   0.8933   0.0973
  -5.000  -0.1840   0.02937   0.02174  -0.0608   0.8765   0.0967
  -4.750  -0.1495   0.02754   0.01942  -0.0610   0.8570   0.0986
  -4.500  -0.1195   0.02568   0.01723  -0.0605   0.8324   0.1002
  -4.250  -0.0861   0.02384   0.01520  -0.0606   0.8083   0.1023
  -4.000  -0.0573   0.02265   0.01384  -0.0599   0.7792   0.1052
  -3.750  -0.0298   0.02179   0.01275  -0.0589   0.7497   0.1103
  -3.500  -0.0027   0.02092   0.01164  -0.0579   0.7218   0.1158
  -3.250   0.0231   0.02014   0.01079  -0.0569   0.6965   0.1230
  -3.000   0.0470   0.01946   0.01003  -0.0556   0.6727   0.1328
  -2.750   0.0694   0.01889   0.00947  -0.0542   0.6523   0.1522
  -2.500   0.0896   0.01806   0.00886  -0.0523   0.6351   0.2071
  -2.250   0.0972   0.01607   0.00850  -0.0486   0.6211   0.5163
  -2.000   0.1520   0.01602   0.00928  -0.0498   0.6029   0.8657
  -1.750   0.2633   0.01735   0.00997  -0.0621   0.5815   0.9500
  -1.500   0.3804   0.01757   0.00963  -0.0786   0.5606   0.9945
  -1.250   0.4143   0.01752   0.00931  -0.0805   0.5498   1.0000
  -1.000   0.4328   0.01762   0.00929  -0.0794   0.5405   1.0000
  -0.750   0.4527   0.01776   0.00920  -0.0783   0.5331   1.0000
  -0.500   0.4718   0.01792   0.00928  -0.0772   0.5250   1.0000
  -0.250   0.4922   0.01809   0.00927  -0.0762   0.5182   1.0000
   0.000   0.5119   0.01831   0.00940  -0.0751   0.5110   1.0000
   0.250   0.5320   0.01852   0.00949  -0.0740   0.5041   1.0000
   0.500   0.5532   0.01880   0.00960  -0.0730   0.4986   1.0000
   0.750   0.5722   0.01911   0.00991  -0.0717   0.4921   1.0000
   1.000   0.5930   0.01941   0.01010  -0.0706   0.4869   1.0000
   1.250   0.6140   0.01978   0.01035  -0.0696   0.4820   1.0000
   1.500   0.6325   0.02017   0.01077  -0.0682   0.4759   1.0000
   1.750   0.6534   0.02052   0.01102  -0.0671   0.4708   1.0000
   2.000   0.6750   0.02096   0.01134  -0.0661   0.4665   1.0000
   2.250   0.6924   0.02148   0.01194  -0.0646   0.4614   1.0000
   2.500   0.7123   0.02194   0.01239  -0.0633   0.4567   1.0000
   2.750   0.7344   0.02237   0.01270  -0.0624   0.4527   1.0000
   3.000   0.7524   0.02299   0.01336  -0.0609   0.4482   1.0000
   3.250   0.7699   0.02359   0.01402  -0.0594   0.4433   1.0000
   3.500   0.7904   0.02411   0.01451  -0.0582   0.4392   1.0000
   3.750   0.8133   0.02465   0.01495  -0.0574   0.4359   1.0000
   4.000   0.8284   0.02549   0.01590  -0.0556   0.4316   1.0000
   4.250   0.8444   0.02624   0.01675  -0.0539   0.4269   1.0000
   4.500   0.8651   0.02681   0.01730  -0.0528   0.4230   1.0000
   4.750   0.8888   0.02737   0.01777  -0.0522   0.4197   1.0000
   5.000   0.9012   0.02845   0.01900  -0.0501   0.4158   1.0000
   5.250   0.9135   0.02949   0.02019  -0.0480   0.4112   1.0000
   5.500   0.9337   0.03012   0.02082  -0.0469   0.4072   1.0000
   5.750   0.9584   0.03065   0.02127  -0.0464   0.4039   1.0000
   6.000   0.9693   0.03194   0.02272  -0.0443   0.4002   1.0000
   6.250   0.9738   0.03345   0.02444  -0.0414   0.3954   1.0000
   6.500   0.9916   0.03425   0.02527  -0.0401   0.3913   1.0000
   6.750   1.0203   0.03456   0.02550  -0.0401   0.3881   1.0000
   7.000   1.0265   0.03617   0.02728  -0.0376   0.3840   1.0000
   7.250   1.0182   0.03848   0.02984  -0.0337   0.3788   1.0000
   7.500   1.0370   0.03921   0.03060  -0.0326   0.3747   1.0000
   7.750   1.0761   0.03891   0.03018  -0.0336   0.3716   1.0000
   8.000   1.0487   0.04247   0.03405  -0.0281   0.3663   1.0000
   8.250   1.0182   0.04616   0.03796  -0.0229   0.3607   1.0000
   8.500   1.0706   0.04472   0.03645  -0.0246   0.3573   1.0000
   8.750   1.1249   0.04360   0.03519  -0.0268   0.3542   1.0000
   9.000   0.8185   0.07217   0.06417  -0.0155   0.3394   1.0000
   9.250   0.8944   0.06533   0.05738  -0.0122   0.3392   1.0000
   9.500   1.0950   0.05053   0.04254  -0.0165   0.3398   1.0000
   9.750   0.6867   0.09868   0.09077  -0.0223   0.3188   1.0000
  10.000   0.6738   0.10390   0.09602  -0.0237   0.3154   1.0000
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