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USA 5 AIRFOIL (usa5-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: USA 5 AIRFOIL (usa5-il)
Reynolds number: 100,000
Max Cl/Cd: 56.29 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa5-il-100000-n5.txt
Download as CSV file: xf-usa5-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3214   0.09379   0.08935  -0.0233   1.0000   0.0356
  -7.000  -0.3276   0.09202   0.08767  -0.0219   1.0000   0.0359
  -6.750  -0.3317   0.08999   0.08572  -0.0212   1.0000   0.0366
  -6.500  -0.3343   0.08800   0.08380  -0.0212   1.0000   0.0374
  -6.250  -0.3357   0.08620   0.08207  -0.0224   1.0000   0.0383
  -6.000  -0.3338   0.08436   0.08025  -0.0246   1.0000   0.0389
  -5.750  -0.3289   0.08195   0.07785  -0.0263   1.0000   0.0391
  -5.250  -0.2880   0.07426   0.07005  -0.0347   0.9940   0.0393
  -4.750  -0.2492   0.06348   0.05923  -0.0383   0.9839   0.0271
  -4.500  -0.2167   0.05872   0.05432  -0.0440   0.9782   0.0248
  -4.250  -0.1843   0.05403   0.04949  -0.0492   0.9711   0.0235
  -4.000  -0.1475   0.04956   0.04482  -0.0545   0.9655   0.0244
  -3.750  -0.1123   0.04521   0.04024  -0.0587   0.9574   0.0252
  -3.500  -0.0753   0.04067   0.03541  -0.0625   0.9495   0.0250
  -3.250  -0.0361   0.03592   0.03027  -0.0659   0.9418   0.0248
  -3.000  -0.0023   0.03140   0.02529  -0.0675   0.9322   0.0250
  -2.750   0.0361   0.02611   0.01923  -0.0691   0.9255   0.0268
  -2.500   0.0681   0.02230   0.01449  -0.0691   0.9160   0.0300
  -2.250   0.1008   0.02050   0.01236  -0.0696   0.9063   0.0340
  -2.000   0.1377   0.01913   0.01068  -0.0709   0.8990   0.0445
  -1.750   0.1666   0.01830   0.00967  -0.0706   0.8883   0.0573
  -1.500   0.1974   0.01782   0.00898  -0.0708   0.8784   0.0742
  -1.250   0.2305   0.01736   0.00823  -0.0713   0.8693   0.0927
  -1.000   0.2615   0.01673   0.00746  -0.0716   0.8586   0.1059
  -0.750   0.2911   0.01627   0.00693  -0.0717   0.8464   0.1212
  -0.500   0.3211   0.01586   0.00647  -0.0719   0.8334   0.1369
  -0.250   0.3516   0.01547   0.00602  -0.0721   0.8196   0.1504
   0.000   0.3837   0.01513   0.00558  -0.0727   0.8049   0.1630
   0.250   0.4165   0.01481   0.00519  -0.0733   0.7894   0.1728
   0.500   0.4472   0.01457   0.00488  -0.0736   0.7723   0.1876
   0.750   0.4764   0.01436   0.00462  -0.0736   0.7526   0.2075
   1.000   0.5064   0.01405   0.00440  -0.0738   0.7317   0.2557
   1.500   0.6184   0.01264   0.00403  -0.0857   0.6734   1.0000
   1.750   0.6433   0.01283   0.00398  -0.0848   0.6446   1.0000
   2.000   0.6676   0.01305   0.00400  -0.0838   0.6164   1.0000
   2.250   0.6913   0.01330   0.00408  -0.0828   0.5881   1.0000
   2.500   0.7148   0.01356   0.00419  -0.0818   0.5613   1.0000
   2.750   0.7382   0.01384   0.00434  -0.0808   0.5368   1.0000
   3.250   0.7848   0.01444   0.00473  -0.0789   0.4942   1.0000
   3.500   0.8081   0.01477   0.00497  -0.0780   0.4765   1.0000
   3.750   0.8314   0.01510   0.00524  -0.0771   0.4613   1.0000
   4.000   0.8549   0.01544   0.00558  -0.0762   0.4480   1.0000
   4.250   0.8785   0.01578   0.00592  -0.0754   0.4365   1.0000
   4.500   0.9020   0.01615   0.00629  -0.0746   0.4260   1.0000
   4.750   0.9255   0.01651   0.00670  -0.0738   0.4155   1.0000
   5.000   0.9491   0.01686   0.00719  -0.0730   0.4053   1.0000
   5.250   0.9690   0.01727   0.00756  -0.0715   0.3808   1.0000
   5.500   0.9886   0.01768   0.00794  -0.0700   0.3551   1.0000
   5.750   1.0076   0.01808   0.00833  -0.0685   0.3220   1.0000
   6.000   1.0242   0.01866   0.00868  -0.0666   0.2554   1.0000
   6.250   1.0363   0.01995   0.00940  -0.0643   0.1698   1.0000
   6.500   1.0438   0.02210   0.01077  -0.0616   0.0426   1.0000
   6.750   1.0555   0.02369   0.01223  -0.0590   0.0227   1.0000
   7.000   1.0713   0.02477   0.01353  -0.0569   0.0202   1.0000
   7.250   1.0856   0.02598   0.01498  -0.0547   0.0188   1.0000
   7.500   1.0997   0.02710   0.01637  -0.0524   0.0176   1.0000
   7.750   1.1119   0.02833   0.01784  -0.0500   0.0164   1.0000
   8.000   1.1210   0.02970   0.01945  -0.0473   0.0156   1.0000
   8.250   1.1272   0.03119   0.02124  -0.0441   0.0151   1.0000
   8.500   1.1309   0.03270   0.02292  -0.0407   0.0148   1.0000
   8.750   1.1312   0.03432   0.02470  -0.0369   0.0145   1.0000
   9.000   1.1334   0.03595   0.02645  -0.0337   0.0144   1.0000
   9.250   1.1365   0.03779   0.02839  -0.0307   0.0141   1.0000
   9.500   1.1424   0.03975   0.03043  -0.0282   0.0137   1.0000
   9.750   1.1505   0.04187   0.03262  -0.0262   0.0129   1.0000
  10.000   1.1630   0.04424   0.03505  -0.0246   0.0123   1.0000
  10.250   1.1808   0.04726   0.03816  -0.0237   0.0116   1.0000
  10.500   1.1979   0.05039   0.04149  -0.0226   0.0115   1.0000
  10.750   1.2065   0.05302   0.04438  -0.0208   0.0116   1.0000
  11.000   1.2103   0.05581   0.04744  -0.0188   0.0116   1.0000
  11.250   1.2099   0.05861   0.05051  -0.0168   0.0117   1.0000
  11.500   1.2059   0.06151   0.05369  -0.0149   0.0118   1.0000
  11.750   1.1993   0.06459   0.05705  -0.0134   0.0119   1.0000
  12.000   1.1913   0.06780   0.06053  -0.0123   0.0121   1.0000
  12.250   1.1808   0.07161   0.06461  -0.0118   0.0122   1.0000
  12.500   1.1687   0.07578   0.06903  -0.0119   0.0123   1.0000
  12.750   1.1557   0.08022   0.07373  -0.0128   0.0125   1.0000
  13.000   1.1397   0.08531   0.07907  -0.0145   0.0126   1.0000
  13.250   1.1229   0.09098   0.08497  -0.0169   0.0127   1.0000
  13.500   1.1067   0.09694   0.09115  -0.0200   0.0129   1.0000
  13.750   1.0903   0.10342   0.09782  -0.0237   0.0130   1.0000
  14.000   1.0731   0.11061   0.10519  -0.0280   0.0131   1.0000
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