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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 500,000
Max Cl/Cd: 68.18 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf25-il-500000.txt
Download as CSV file: xf-raf25-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 25 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5679   0.08017   0.07793  -0.0179   1.0000   0.0187
  -8.250  -0.5730   0.07504   0.07284  -0.0222   1.0000   0.0187
  -8.000  -0.5749   0.06867   0.06642  -0.0279   1.0000   0.0187
  -7.750  -0.5755   0.06297   0.06062  -0.0310   1.0000   0.0187
  -7.500  -0.5751   0.05784   0.05535  -0.0326   1.0000   0.0188
  -7.250  -0.5738   0.05279   0.05012  -0.0333   1.0000   0.0188
  -7.000  -0.5698   0.04816   0.04525  -0.0332   1.0000   0.0188
  -6.250  -0.5625   0.02919   0.02531  -0.0301   1.0000   0.0157
  -6.000  -0.5478   0.02563   0.02135  -0.0282   1.0000   0.0161
  -5.750  -0.5323   0.02196   0.01720  -0.0261   1.0000   0.0160
  -5.500  -0.5149   0.01858   0.01333  -0.0240   1.0000   0.0158
  -5.250  -0.4943   0.01666   0.01112  -0.0224   1.0000   0.0166
  -5.000  -0.4726   0.01538   0.00963  -0.0211   1.0000   0.0176
  -4.750  -0.4500   0.01468   0.00878  -0.0199   1.0000   0.0186
  -4.500  -0.4305   0.01271   0.00664  -0.0183   1.0000   0.0208
  -4.250  -0.4084   0.01229   0.00620  -0.0173   1.0000   0.0233
  -4.000  -0.3866   0.01168   0.00553  -0.0160   1.0000   0.0257
  -3.750  -0.3647   0.01133   0.00512  -0.0148   1.0000   0.0277
  -3.500  -0.3450   0.01034   0.00404  -0.0133   1.0000   0.0325
  -3.250  -0.3232   0.01001   0.00369  -0.0122   0.9999   0.0368
  -3.000  -0.2856   0.00972   0.00334  -0.0144   0.9969   0.0401
  -2.750  -0.2497   0.00916   0.00276  -0.0163   0.9938   0.0480
  -2.500  -0.2144   0.00885   0.00241  -0.0180   0.9898   0.0552
  -2.250  -0.1789   0.00833   0.00210  -0.0198   0.9860   0.1024
  -2.000  -0.1452   0.00734   0.00192  -0.0218   0.9830   0.3157
  -1.750  -0.1153   0.00671   0.00181  -0.0226   0.9766   0.4620
  -1.500  -0.0816   0.00622   0.00174  -0.0241   0.9721   0.5837
  -1.250  -0.0497   0.00581   0.00168  -0.0249   0.9669   0.6835
  -1.000  -0.0206   0.00532   0.00164  -0.0249   0.9604   0.7999
  -0.750   0.0247   0.00501   0.00169  -0.0282   0.9590   0.9262
  -0.500   0.0973   0.00507   0.00176  -0.0378   0.9630   0.9969
  -0.250   0.1423   0.00501   0.00167  -0.0416   0.9566   1.0000
   0.000   0.1811   0.00494   0.00157  -0.0440   0.9449   1.0000
   0.250   0.2252   0.00484   0.00141  -0.0473   0.9201   1.0000
   0.500   0.2579   0.00486   0.00128  -0.0480   0.8805   1.0000
   0.750   0.2825   0.00493   0.00125  -0.0470   0.8515   1.0000
   1.000   0.3060   0.00502   0.00126  -0.0459   0.8244   1.0000
   1.250   0.3291   0.00512   0.00126  -0.0446   0.7949   1.0000
   1.500   0.3513   0.00527   0.00127  -0.0432   0.7554   1.0000
   1.750   0.3728   0.00547   0.00128  -0.0416   0.7044   1.0000
   2.000   0.3934   0.00577   0.00131  -0.0399   0.6345   1.0000
   2.250   0.4133   0.00619   0.00141  -0.0382   0.5479   1.0000
   2.500   0.4319   0.00685   0.00157  -0.0364   0.4237   1.0000
   2.750   0.4459   0.00823   0.00193  -0.0343   0.1797   1.0000
   3.000   0.4652   0.00916   0.00232  -0.0328   0.0651   1.0000
   3.250   0.4890   0.00950   0.00267  -0.0320   0.0549   1.0000
   3.500   0.5120   0.00999   0.00319  -0.0309   0.0463   1.0000
   3.750   0.5362   0.01027   0.00351  -0.0302   0.0400   1.0000
   4.000   0.5581   0.01095   0.00422  -0.0289   0.0328   1.0000
   4.250   0.5826   0.01123   0.00450  -0.0282   0.0272   1.0000
   4.500   0.6029   0.01220   0.00552  -0.0267   0.0231   1.0000
   4.750   0.6263   0.01272   0.00613  -0.0257   0.0211   1.0000
   5.000   0.6490   0.01342   0.00689  -0.0246   0.0192   1.0000
   5.250   0.6721   0.01408   0.00757  -0.0236   0.0174   1.0000
   5.500   0.6912   0.01614   0.00977  -0.0219   0.0154   1.0000
   5.750   0.7148   0.01712   0.01088  -0.0209   0.0148   1.0000
   6.000   0.7377   0.01865   0.01262  -0.0198   0.0141   1.0000
   6.250   0.7594   0.02087   0.01513  -0.0183   0.0135   1.0000
   6.500   0.7778   0.02450   0.01921  -0.0162   0.0138   1.0000
   6.750   0.7921   0.02927   0.02431  -0.0140   0.0162   1.0000
   9.250   0.7435   0.05884   0.05660   0.0052   0.0165   1.0000
   9.500   0.6833   0.07510   0.07307  -0.0074   0.0177   1.0000
   9.750   0.6668   0.08391   0.08185  -0.0125   0.0180   1.0000
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