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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 100,000
Max Cl/Cd: 40.21 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf25-il-100000-n5.txt
Download as CSV file: xf-raf25-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5722   0.08628   0.08139  -0.0121   1.0000   0.0219
  -8.250  -0.5764   0.08149   0.07669  -0.0149   1.0000   0.0216
  -8.000  -0.5819   0.07660   0.07187  -0.0185   1.0000   0.0210
  -7.750  -0.5839   0.07061   0.06588  -0.0235   1.0000   0.0205
  -7.500  -0.5844   0.06436   0.05957  -0.0276   1.0000   0.0201
  -7.250  -0.5835   0.05805   0.05310  -0.0305   1.0000   0.0195
  -7.000  -0.5802   0.05190   0.04668  -0.0322   1.0000   0.0190
  -6.750  -0.5739   0.04627   0.04068  -0.0327   1.0000   0.0188
  -6.500  -0.5642   0.04151   0.03551  -0.0325   1.0000   0.0188
  -6.250  -0.5513   0.03758   0.03116  -0.0318   1.0000   0.0192
  -6.000  -0.5353   0.03500   0.02827  -0.0310   1.0000   0.0209
  -5.750  -0.5182   0.03189   0.02468  -0.0298   1.0000   0.0227
  -5.500  -0.4996   0.02868   0.02084  -0.0284   1.0000   0.0237
  -5.250  -0.4791   0.02577   0.01737  -0.0270   1.0000   0.0249
  -5.000  -0.4565   0.02360   0.01466  -0.0255   1.0000   0.0272
  -4.750  -0.4353   0.02181   0.01274  -0.0246   1.0000   0.0311
  -4.500  -0.4127   0.02019   0.01090  -0.0233   1.0000   0.0350
  -4.250  -0.3908   0.01888   0.00948  -0.0222   1.0000   0.0416
  -4.000  -0.3685   0.01782   0.00826  -0.0210   1.0000   0.0488
  -3.750  -0.3471   0.01689   0.00729  -0.0199   1.0000   0.0599
  -3.500  -0.3255   0.01604   0.00636  -0.0187   1.0000   0.0681
  -3.250  -0.3035   0.01538   0.00564  -0.0177   1.0000   0.0799
  -3.000  -0.2814   0.01469   0.00493  -0.0166   1.0000   0.0942
  -2.750  -0.2597   0.01390   0.00434  -0.0156   1.0000   0.1478
  -2.500  -0.2389   0.01295   0.00397  -0.0147   1.0000   0.2843
  -2.250  -0.2176   0.01225   0.00377  -0.0138   1.0000   0.4190
  -2.000  -0.1986   0.01152   0.00374  -0.0119   1.0000   0.5987
  -1.750  -0.1412   0.01065   0.00379  -0.0160   1.0000   0.9298
  -1.500  -0.0934   0.01063   0.00354  -0.0204   1.0000   1.0000
  -1.250  -0.0752   0.01065   0.00342  -0.0188   1.0000   1.0000
  -1.000  -0.0567   0.01071   0.00332  -0.0173   1.0000   1.0000
  -0.750  -0.0381   0.01078   0.00328  -0.0159   1.0000   1.0000
  -0.500  -0.0194   0.01089   0.00330  -0.0145   1.0000   1.0000
  -0.250   0.0152   0.01101   0.00333  -0.0164   0.9936   1.0000
   0.000   0.0547   0.01112   0.00337  -0.0192   0.9844   1.0000
   0.250   0.0935   0.01121   0.00343  -0.0218   0.9747   1.0000
   0.500   0.1325   0.01129   0.00351  -0.0244   0.9648   1.0000
   0.750   0.1715   0.01133   0.00357  -0.0269   0.9542   1.0000
   1.000   0.2110   0.01134   0.00363  -0.0294   0.9419   1.0000
   1.250   0.2515   0.01132   0.00370  -0.0319   0.9278   1.0000
   1.500   0.2911   0.01131   0.00377  -0.0341   0.9127   1.0000
   1.750   0.3276   0.01132   0.00389  -0.0357   0.8971   1.0000
   2.000   0.3631   0.01130   0.00400  -0.0369   0.8739   1.0000
   2.250   0.3970   0.01125   0.00400  -0.0372   0.8353   1.0000
   2.500   0.4252   0.01127   0.00399  -0.0361   0.7804   1.0000
   2.750   0.4498   0.01142   0.00404  -0.0344   0.7113   1.0000
   3.000   0.4721   0.01174   0.00410  -0.0324   0.6164   1.0000
   3.250   0.4916   0.01238   0.00426  -0.0302   0.4941   1.0000
   3.500   0.5084   0.01344   0.00463  -0.0280   0.3301   1.0000
   3.750   0.5226   0.01532   0.00540  -0.0262   0.1197   1.0000
   4.000   0.5438   0.01631   0.00623  -0.0250   0.0815   1.0000
   4.250   0.5651   0.01728   0.00717  -0.0238   0.0604   1.0000
   4.500   0.5852   0.01845   0.00840  -0.0224   0.0459   1.0000
   4.750   0.6050   0.01985   0.00991  -0.0208   0.0381   1.0000
   5.000   0.6261   0.02113   0.01123  -0.0197   0.0316   1.0000
   5.250   0.6492   0.02266   0.01296  -0.0184   0.0277   1.0000
   5.500   0.6724   0.02429   0.01471  -0.0174   0.0247   1.0000
   5.750   0.6950   0.02626   0.01686  -0.0165   0.0221   1.0000
   6.000   0.7186   0.02871   0.01975  -0.0152   0.0207   1.0000
   6.250   0.7400   0.03167   0.02320  -0.0136   0.0200   1.0000
   6.500   0.7587   0.03493   0.02699  -0.0119   0.0196   1.0000
   6.750   0.7746   0.03843   0.03103  -0.0101   0.0192   1.0000
   7.000   0.7879   0.04199   0.03517  -0.0083   0.0179   1.0000
   7.250   0.7986   0.04568   0.03929  -0.0068   0.0171   1.0000
   7.500   0.8061   0.04972   0.04373  -0.0053   0.0167   1.0000
   7.750   0.8097   0.05425   0.04862  -0.0039   0.0168   1.0000
   8.000   0.8094   0.05898   0.05367  -0.0030   0.0170   1.0000
   8.250   0.8059   0.06367   0.05859  -0.0024   0.0172   1.0000
   8.500   0.7984   0.06839   0.06350  -0.0024   0.0174   1.0000
   8.750   0.7857   0.07304   0.06825  -0.0026   0.0176   1.0000
   9.000   0.7716   0.07794   0.07322  -0.0045   0.0179   1.0000
   9.250   0.7588   0.08454   0.07984  -0.0096   0.0182   1.0000
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