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NACA 64-208 (naca64208-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 64-208 (naca64208-il)
Reynolds number: 500,000
Max Cl/Cd: 75.77 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca64208-il-500000.txt
Download as CSV file: xf-naca64208-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64-208                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5528   0.08568   0.08343  -0.0165   1.0000   0.0150
  -8.750  -0.5539   0.08123   0.07901  -0.0192   1.0000   0.0151
  -8.500  -0.5573   0.07645   0.07426  -0.0227   1.0000   0.0154
  -8.250  -0.5695   0.07018   0.06803  -0.0292   1.0000   0.0152
  -8.000  -0.5747   0.06502   0.06281  -0.0333   1.0000   0.0153
  -7.750  -0.5753   0.06020   0.05790  -0.0359   1.0000   0.0156
  -7.500  -0.5722   0.05570   0.05328  -0.0376   1.0000   0.0159
  -7.250  -0.5659   0.05145   0.04889  -0.0386   1.0000   0.0164
  -7.000  -0.5568   0.04727   0.04453  -0.0390   1.0000   0.0173
  -6.750  -0.5332   0.04534   0.04229  -0.0381   1.0000   0.0199
  -6.500  -0.5210   0.04207   0.03874  -0.0374   1.0000   0.0200
  -6.000  -0.5098   0.03012   0.02604  -0.0353   1.0000   0.0213
  -5.750  -0.4956   0.02786   0.02367  -0.0341   1.0000   0.0222
  -5.000  -0.4210   0.01836   0.01304  -0.0330   0.9958   0.0183
  -4.750  -0.3863   0.01607   0.01043  -0.0344   0.9929   0.0187
  -4.500  -0.3519   0.01465   0.00881  -0.0357   0.9892   0.0194
  -4.250  -0.3175   0.01295   0.00695  -0.0373   0.9858   0.0212
  -4.000  -0.2821   0.01156   0.00551  -0.0391   0.9830   0.0230
  -3.750  -0.2481   0.01081   0.00469  -0.0406   0.9781   0.0248
  -3.500  -0.2132   0.01016   0.00400  -0.0422   0.9730   0.0273
  -3.250  -0.1770   0.00981   0.00362  -0.0441   0.9687   0.0304
  -3.000  -0.1463   0.00899   0.00273  -0.0448   0.9597   0.0377
  -2.750  -0.1144   0.00855   0.00226  -0.0457   0.9517   0.0511
  -2.500  -0.0874   0.00718   0.00179  -0.0465   0.9419   0.2923
  -2.250  -0.0634   0.00611   0.00164  -0.0463   0.9304   0.5613
  -2.000  -0.0375   0.00589   0.00161  -0.0458   0.9195   0.6336
  -1.750  -0.0112   0.00580   0.00156  -0.0453   0.9088   0.6707
  -1.500   0.0149   0.00574   0.00153  -0.0447   0.8982   0.7068
  -1.250   0.0400   0.00568   0.00158  -0.0438   0.8870   0.7506
  -1.000   0.0655   0.00566   0.00161  -0.0430   0.8760   0.7797
  -0.750   0.0913   0.00565   0.00162  -0.0423   0.8651   0.8005
  -0.500   0.1178   0.00565   0.00159  -0.0419   0.8545   0.8138
  -0.250   0.1444   0.00566   0.00157  -0.0414   0.8440   0.8256
   0.000   0.1710   0.00565   0.00157  -0.0411   0.8329   0.8378
   0.250   0.1976   0.00566   0.00158  -0.0407   0.8221   0.8499
   0.500   0.2240   0.00568   0.00160  -0.0402   0.8115   0.8622
   0.750   0.2501   0.00570   0.00162  -0.0397   0.8012   0.8750
   1.000   0.2760   0.00572   0.00165  -0.0391   0.7896   0.8880
   1.250   0.3012   0.00573   0.00168  -0.0384   0.7767   0.9010
   1.500   0.3257   0.00575   0.00170  -0.0374   0.7612   0.9147
   1.750   0.3490   0.00576   0.00168  -0.0361   0.7363   0.9296
   2.000   0.3713   0.00580   0.00162  -0.0346   0.6997   0.9459
   2.250   0.3980   0.00584   0.00161  -0.0342   0.6701   0.9632
   2.500   0.4314   0.00593   0.00167  -0.0353   0.6429   0.9803
   2.750   0.4645   0.00613   0.00171  -0.0366   0.5932   1.0000
   3.000   0.4890   0.00656   0.00182  -0.0362   0.5085   1.0000
   3.250   0.5078   0.00796   0.00221  -0.0354   0.2725   1.0000
   3.500   0.5272   0.00960   0.00287  -0.0348   0.0585   1.0000
   3.750   0.5533   0.01013   0.00335  -0.0347   0.0387   1.0000
   4.000   0.5793   0.01071   0.00396  -0.0344   0.0312   1.0000
   4.250   0.6053   0.01122   0.00451  -0.0342   0.0272   1.0000
   4.500   0.6288   0.01222   0.00556  -0.0335   0.0241   1.0000
   4.750   0.6521   0.01330   0.00673  -0.0328   0.0226   1.0000
   5.000   0.6781   0.01380   0.00730  -0.0325   0.0211   1.0000
   5.250   0.7027   0.01473   0.00830  -0.0320   0.0201   1.0000
   5.500   0.7275   0.01578   0.00943  -0.0314   0.0192   1.0000
   5.750   0.7524   0.01697   0.01071  -0.0309   0.0182   1.0000
   6.000   0.7774   0.01837   0.01222  -0.0304   0.0175   1.0000
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