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RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 500,000
Max Cl/Cd: 88.76 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf26-il-500000.txt
Download as CSV file: xf-raf26-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 26 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4104   0.08428   0.08210  -0.0218   1.0000   0.0169
  -8.750  -0.4132   0.07987   0.07772  -0.0232   1.0000   0.0181
  -8.500  -0.4156   0.07568   0.07354  -0.0244   1.0000   0.0184
  -8.250  -0.4163   0.07194   0.06983  -0.0262   1.0000   0.0192
  -8.000  -0.4205   0.06789   0.06580  -0.0272   1.0000   0.0193
  -7.750  -0.4275   0.06389   0.06184  -0.0282   1.0000   0.0194
  -7.500  -0.4392   0.06021   0.05820  -0.0288   1.0000   0.0194
  -7.250  -0.4460   0.05537   0.05338  -0.0321   1.0000   0.0194
  -7.000  -0.4491   0.04955   0.04751  -0.0361   1.0000   0.0194
  -6.750  -0.4501   0.04390   0.04176  -0.0385   1.0000   0.0195
  -6.500  -0.4496   0.03884   0.03656  -0.0395   1.0000   0.0195
  -4.750  -0.3886   0.01735   0.01220  -0.0383   0.9956   0.0149
  -4.500  -0.3571   0.01529   0.00979  -0.0390   0.9938   0.0155
  -4.250  -0.3257   0.01397   0.00827  -0.0396   0.9912   0.0166
  -4.000  -0.2931   0.01293   0.00705  -0.0406   0.9885   0.0178
  -3.750  -0.2564   0.01278   0.00682  -0.0424   0.9858   0.0196
  -3.500  -0.2246   0.01086   0.00473  -0.0433   0.9838   0.0226
  -3.250  -0.1937   0.01029   0.00411  -0.0439   0.9790   0.0260
  -3.000  -0.1601   0.00973   0.00348  -0.0451   0.9750   0.0330
  -2.750  -0.1248   0.00918   0.00293  -0.0467   0.9722   0.0498
  -2.500  -0.0949   0.00866   0.00261  -0.0473   0.9668   0.0992
  -2.250  -0.0640   0.00806   0.00238  -0.0482   0.9621   0.1972
  -2.000  -0.0311   0.00738   0.00222  -0.0497   0.9589   0.3482
  -1.750  -0.0032   0.00696   0.00206  -0.0498   0.9514   0.4422
  -1.500   0.0303   0.00647   0.00187  -0.0510   0.9462   0.5329
  -1.250   0.0591   0.00598   0.00174  -0.0511   0.9376   0.6443
  -1.000   0.0922   0.00528   0.00164  -0.0519   0.9315   0.8161
  -0.750   0.1770   0.00519   0.00183  -0.0646   0.9365   1.0000
  -0.500   0.2256   0.00508   0.00165  -0.0692   0.9281   1.0000
  -0.250   0.2776   0.00497   0.00147  -0.0745   0.9164   1.0000
   0.000   0.3264   0.00493   0.00131  -0.0791   0.8951   1.0000
   0.250   0.3593   0.00499   0.00126  -0.0801   0.8731   1.0000
   0.500   0.3860   0.00508   0.00126  -0.0797   0.8520   1.0000
   0.750   0.4106   0.00516   0.00127  -0.0788   0.8309   1.0000
   1.000   0.4346   0.00525   0.00128  -0.0778   0.8081   1.0000
   1.250   0.4577   0.00535   0.00130  -0.0765   0.7815   1.0000
   1.500   0.4802   0.00549   0.00132  -0.0752   0.7507   1.0000
   1.750   0.5021   0.00566   0.00138  -0.0737   0.7155   1.0000
   2.000   0.5228   0.00589   0.00144  -0.0720   0.6719   1.0000
   2.250   0.5428   0.00619   0.00153  -0.0702   0.6199   1.0000
   2.500   0.5615   0.00658   0.00164  -0.0681   0.5454   1.0000
   2.750   0.5757   0.00735   0.00186  -0.0654   0.4212   1.0000
   3.000   0.5933   0.00804   0.00212  -0.0634   0.3131   1.0000
   3.250   0.6068   0.00920   0.00256  -0.0609   0.1491   1.0000
   3.500   0.6270   0.00980   0.00287  -0.0595   0.0932   1.0000
   3.750   0.6498   0.01015   0.00320  -0.0585   0.0838   1.0000
   4.000   0.6734   0.01043   0.00352  -0.0576   0.0755   1.0000
   4.250   0.6968   0.01073   0.00382  -0.0567   0.0651   1.0000
   4.500   0.7195   0.01113   0.00406  -0.0556   0.0318   1.0000
   4.750   0.7401   0.01181   0.00475  -0.0540   0.0222   1.0000
   5.000   0.7619   0.01235   0.00536  -0.0527   0.0195   1.0000
   5.250   0.7831   0.01297   0.00607  -0.0512   0.0176   1.0000
   5.500   0.8034   0.01371   0.00687  -0.0497   0.0160   1.0000
   5.750   0.8213   0.01486   0.00812  -0.0476   0.0147   1.0000
   6.000   0.8364   0.01716   0.01064  -0.0450   0.0138   1.0000
   6.250   0.8586   0.01812   0.01173  -0.0438   0.0132   1.0000
   6.500   0.8812   0.01901   0.01275  -0.0427   0.0124   1.0000
   6.750   0.9024   0.02064   0.01458  -0.0413   0.0120   1.0000
   7.000   0.9226   0.02266   0.01686  -0.0396   0.0114   1.0000
   7.250   0.9397   0.02564   0.02022  -0.0374   0.0114   1.0000
   7.500   0.9495   0.03033   0.02546  -0.0340   0.0121   1.0000
   7.750   0.9228   0.04616   0.04227  -0.0264   0.0213   1.0000
   8.000   0.9000   0.03561   0.03228  -0.0203   0.0198   1.0000
   8.250   0.9021   0.03949   0.03643  -0.0175   0.0187   1.0000
   8.500   0.8978   0.04383   0.04099  -0.0148   0.0181   1.0000
   8.750   0.8884   0.04822   0.04557  -0.0120   0.0176   1.0000
   9.000   0.8729   0.05216   0.04966  -0.0088   0.0174   1.0000
   9.250   0.8503   0.05578   0.05341  -0.0055   0.0174   1.0000
   9.500   0.8243   0.06013   0.05787  -0.0043   0.0176   1.0000
   9.750   0.7958   0.06575   0.06360  -0.0056   0.0180   1.0000
  10.000   0.7668   0.07314   0.07107  -0.0096   0.0187   1.0000
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