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NACA 64-108 AIRFOIL (n64108-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 64-108 AIRFOIL (n64108-il)
Reynolds number: 500,000
Max Cl/Cd: 60.03 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n64108-il-500000.txt
Download as CSV file: xf-n64108-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64-108 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.6169   0.12434   0.12201   0.0106   1.0000   0.0140
 -11.000  -0.6139   0.11977   0.11746   0.0084   1.0000   0.0140
 -10.750  -0.6110   0.11530   0.11300   0.0063   1.0000   0.0140
  -5.750  -0.5354   0.02604   0.02132  -0.0225   1.0000   0.0210
  -5.500  -0.5180   0.02096   0.01577  -0.0212   1.0000   0.0194
  -5.250  -0.4962   0.01773   0.01216  -0.0198   1.0000   0.0184
  -5.000  -0.4739   0.01595   0.01014  -0.0186   1.0000   0.0188
  -4.750  -0.4523   0.01472   0.00878  -0.0173   1.0000   0.0195
  -4.500  -0.4322   0.01383   0.00780  -0.0159   1.0000   0.0204
  -4.250  -0.4022   0.01371   0.00760  -0.0165   0.9977   0.0216
  -4.000  -0.3699   0.01132   0.00511  -0.0178   0.9931   0.0236
  -3.750  -0.3348   0.01053   0.00427  -0.0196   0.9878   0.0260
  -3.500  -0.2976   0.00995   0.00364  -0.0218   0.9836   0.0291
  -3.250  -0.2629   0.00947   0.00311  -0.0234   0.9767   0.0336
  -3.000  -0.2273   0.00893   0.00253  -0.0252   0.9701   0.0442
  -2.750  -0.1970   0.00794   0.00201  -0.0263   0.9600   0.1658
  -2.500  -0.1731   0.00625   0.00167  -0.0267   0.9482   0.5279
  -2.250  -0.1468   0.00595   0.00163  -0.0263   0.9368   0.6139
  -2.000  -0.1210   0.00583   0.00160  -0.0257   0.9253   0.6576
  -1.500  -0.0705   0.00569   0.00156  -0.0241   0.9016   0.7272
  -1.250  -0.0457   0.00566   0.00160  -0.0231   0.8900   0.7648
  -1.000  -0.0202   0.00565   0.00162  -0.0224   0.8787   0.7880
  -0.750   0.0060   0.00564   0.00159  -0.0218   0.8676   0.8020
  -0.500   0.0324   0.00563   0.00157  -0.0213   0.8567   0.8137
  -0.250   0.0589   0.00563   0.00155  -0.0209   0.8455   0.8251
   0.000   0.0856   0.00562   0.00155  -0.0205   0.8343   0.8369
   0.250   0.1122   0.00563   0.00156  -0.0202   0.8232   0.8487
   0.500   0.1387   0.00565   0.00157  -0.0197   0.8123   0.8607
   0.750   0.1649   0.00567   0.00159  -0.0192   0.8017   0.8730
   1.000   0.1907   0.00568   0.00162  -0.0186   0.7894   0.8851
   1.250   0.2161   0.00569   0.00164  -0.0179   0.7756   0.8973
   1.500   0.2408   0.00570   0.00166  -0.0170   0.7545   0.9100
   1.750   0.2640   0.00575   0.00161  -0.0156   0.7186   0.9237
   2.000   0.2878   0.00580   0.00159  -0.0144   0.6839   0.9383
   2.250   0.3134   0.00585   0.00162  -0.0138   0.6576   0.9532
   2.500   0.3430   0.00595   0.00165  -0.0141   0.6248   0.9673
   2.750   0.3758   0.00626   0.00171  -0.0152   0.5453   0.9800
   3.000   0.4061   0.00760   0.00201  -0.0168   0.2832   0.9947
   3.250   0.4277   0.00918   0.00259  -0.0168   0.0575   1.0000
   3.500   0.4530   0.00966   0.00301  -0.0166   0.0396   1.0000
   3.750   0.4786   0.01020   0.00360  -0.0163   0.0315   1.0000
   4.000   0.5046   0.01068   0.00411  -0.0160   0.0276   1.0000
   4.250   0.5289   0.01152   0.00499  -0.0156   0.0245   1.0000
   4.500   0.5513   0.01288   0.00645  -0.0147   0.0228   1.0000
   4.750   0.5776   0.01336   0.00697  -0.0145   0.0217   1.0000
   5.000   0.6030   0.01413   0.00782  -0.0141   0.0204   1.0000
   5.250   0.6280   0.01518   0.00895  -0.0136   0.0195   1.0000
   5.500   0.6533   0.01636   0.01023  -0.0131   0.0185   1.0000
   5.750   0.6784   0.01793   0.01194  -0.0125   0.0179   1.0000
   6.000   0.7033   0.01912   0.01324  -0.0121   0.0168   1.0000
   6.250   0.7256   0.02166   0.01595  -0.0116   0.0155   1.0000
   6.500   0.7477   0.02449   0.01912  -0.0107   0.0155   1.0000
   6.750   0.7692   0.02719   0.02214  -0.0097   0.0156   1.0000
   7.000   0.7353   0.03007   0.02650  -0.0051   0.0230   1.0000
   7.250   0.7904   0.04121   0.03764  -0.0045   0.0208   1.0000
   7.500   0.8042   0.04472   0.04144  -0.0034   0.0193   1.0000
   7.750   0.8147   0.04828   0.04523  -0.0026   0.0182   1.0000
   8.000   0.8237   0.05137   0.04849  -0.0021   0.0173   1.0000
   8.250   0.8315   0.05394   0.05116  -0.0017   0.0166   1.0000
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