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RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 50,000
Max Cl/Cd: 37.08 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf26-il-50000-n5.txt
Download as CSV file: xf-raf26-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.4918   0.11608   0.10886  -0.0148   1.0000   0.0961
  -9.250  -0.4911   0.11283   0.10568  -0.0162   1.0000   0.0980
  -9.000  -0.5051   0.11143   0.10444  -0.0204   1.0000   0.1013
  -8.500  -0.4849   0.10004   0.09304  -0.0210   1.0000   0.0660
  -8.250  -0.4848   0.09530   0.08835  -0.0229   1.0000   0.0519
  -8.000  -0.4943   0.09099   0.08417  -0.0274   1.0000   0.0462
  -7.750  -0.4909   0.08736   0.08059  -0.0269   1.0000   0.0447
  -7.500  -0.4892   0.08339   0.07668  -0.0284   1.0000   0.0432
  -7.250  -0.4878   0.07898   0.07234  -0.0309   1.0000   0.0420
  -7.000  -0.4851   0.07422   0.06760  -0.0339   1.0000   0.0405
  -6.750  -0.4813   0.06878   0.06214  -0.0375   1.0000   0.0390
  -6.500  -0.4756   0.06173   0.05474  -0.0430   1.0000   0.0364
  -6.250  -0.4670   0.05755   0.05045  -0.0435   1.0000   0.0361
  -6.000  -0.4568   0.05327   0.04597  -0.0442   1.0000   0.0358
  -5.750  -0.4447   0.04895   0.04136  -0.0448   1.0000   0.0356
  -5.500  -0.4302   0.04469   0.03664  -0.0451   1.0000   0.0360
  -5.250  -0.4147   0.04087   0.03243  -0.0450   1.0000   0.0374
  -5.000  -0.3978   0.03830   0.02968  -0.0445   1.0000   0.0401
  -4.750  -0.3783   0.03509   0.02598  -0.0438   1.0000   0.0414
  -4.500  -0.3571   0.03205   0.02238  -0.0429   1.0000   0.0425
  -4.250  -0.3345   0.02940   0.01917  -0.0418   1.0000   0.0447
  -4.000  -0.3124   0.02728   0.01670  -0.0409   1.0000   0.0500
  -3.750  -0.2895   0.02552   0.01459  -0.0397   1.0000   0.0562
  -3.500  -0.2664   0.02378   0.01258  -0.0383   1.0000   0.0630
  -3.250  -0.2430   0.02250   0.01119  -0.0375   1.0000   0.0806
  -3.000  -0.2192   0.02118   0.00973  -0.0365   1.0000   0.1050
  -2.750  -0.1947   0.01984   0.00848  -0.0361   1.0000   0.1508
  -2.500  -0.1707   0.01834   0.00767  -0.0359   1.0000   0.2842
  -2.250  -0.1518   0.01713   0.00739  -0.0342   1.0000   0.5124
  -1.750  -0.0848   0.01565   0.00650  -0.0349   1.0000   1.0000
  -1.500  -0.0633   0.01573   0.00616  -0.0341   1.0000   1.0000
  -1.250  -0.0420   0.01585   0.00594  -0.0332   1.0000   1.0000
  -1.000  -0.0208   0.01599   0.00579  -0.0323   1.0000   1.0000
  -0.750   0.0002   0.01615   0.00568  -0.0315   1.0000   1.0000
  -0.500   0.0210   0.01634   0.00567  -0.0307   1.0000   1.0000
  -0.250   0.0416   0.01656   0.00570  -0.0298   1.0000   1.0000
   0.000   0.0621   0.01681   0.00580  -0.0290   1.0000   1.0000
   0.250   0.0823   0.01709   0.00593  -0.0282   1.0000   1.0000
   0.500   0.1024   0.01740   0.00613  -0.0275   1.0000   1.0000
   0.750   0.1333   0.01780   0.00645  -0.0289   0.9947   1.0000
   1.000   0.1746   0.01821   0.00682  -0.0324   0.9834   1.0000
   1.250   0.2168   0.01856   0.00716  -0.0359   0.9707   1.0000
   1.500   0.2592   0.01884   0.00747  -0.0392   0.9570   1.0000
   1.750   0.2982   0.01908   0.00780  -0.0419   0.9440   1.0000
   2.000   0.3343   0.01934   0.00814  -0.0439   0.9322   1.0000
   2.250   0.3706   0.01960   0.00853  -0.0459   0.9207   1.0000
   2.500   0.4080   0.01983   0.00898  -0.0480   0.9095   1.0000
   2.750   0.4479   0.02000   0.00937  -0.0504   0.8975   1.0000
   3.000   0.4900   0.01997   0.00961  -0.0528   0.8800   1.0000
   3.250   0.5387   0.01967   0.00970  -0.0556   0.8570   1.0000
   3.500   0.5782   0.01941   0.00977  -0.0565   0.8300   1.0000
   3.750   0.6108   0.01915   0.00981  -0.0559   0.7944   1.0000
   4.000   0.6439   0.01863   0.00954  -0.0543   0.7333   1.0000
   4.250   0.6716   0.01849   0.00948  -0.0521   0.6571   1.0000
   4.500   0.6948   0.01874   0.00969  -0.0496   0.5555   1.0000
   4.750   0.7069   0.02015   0.01002  -0.0454   0.3447   1.0000
   5.000   0.7151   0.02270   0.01144  -0.0426   0.1828   1.0000
   5.250   0.7313   0.02467   0.01301  -0.0411   0.1166   1.0000
   5.500   0.7474   0.02695   0.01501  -0.0395   0.0781   1.0000
   5.750   0.7688   0.02916   0.01737  -0.0380   0.0595   1.0000
   6.000   0.7936   0.03109   0.01950  -0.0369   0.0479   1.0000
   6.250   0.8196   0.03346   0.02198  -0.0362   0.0423   1.0000
   6.500   0.8469   0.03610   0.02503  -0.0352   0.0380   1.0000
   6.750   0.8673   0.03839   0.02751  -0.0344   0.0342   1.0000
   7.000   0.8889   0.04180   0.03158  -0.0327   0.0324   1.0000
   7.250   0.9057   0.04549   0.03584  -0.0309   0.0317   1.0000
   7.500   0.9180   0.04932   0.04020  -0.0288   0.0313   1.0000
   7.750   0.9261   0.05332   0.04470  -0.0265   0.0311   1.0000
   8.000   0.9302   0.05736   0.04920  -0.0243   0.0307   1.0000
   8.250   0.9310   0.06136   0.05359  -0.0223   0.0304   1.0000
   8.500   0.9275   0.06549   0.05807  -0.0204   0.0302   1.0000
   8.750   0.9203   0.06962   0.06248  -0.0187   0.0300   1.0000
   9.000   0.9095   0.07375   0.06683  -0.0173   0.0301   1.0000
   9.250   0.8943   0.07773   0.07095  -0.0159   0.0303   1.0000
   9.500   0.8786   0.08200   0.07532  -0.0158   0.0308   1.0000
   9.750   0.8623   0.08698   0.08037  -0.0175   0.0311   1.0000
  10.000   0.8486   0.09258   0.08600  -0.0204   0.0317   1.0000
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