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RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 200,000
Max Cl/Cd: 71.28 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf26-il-200000.txt
Download as CSV file: xf-raf26-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4960   0.10001   0.09635  -0.0167   1.0000   0.0384
  -8.750  -0.4979   0.09650   0.09289  -0.0197   1.0000   0.0401
  -8.500  -0.5009   0.09313   0.08959  -0.0230   1.0000   0.0408
  -8.250  -0.4222   0.07576   0.07247  -0.0275   1.0000   0.0433
  -8.000  -0.4192   0.07249   0.06922  -0.0267   1.0000   0.0444
  -7.750  -0.4210   0.06903   0.06580  -0.0266   1.0000   0.0454
  -7.500  -0.4265   0.06560   0.06241  -0.0268   1.0000   0.0463
  -7.250  -0.5103   0.07042   0.06704  -0.0363   1.0000   0.0429
  -7.000  -0.5046   0.06792   0.06457  -0.0344   1.0000   0.0441
  -6.750  -0.4981   0.06463   0.06126  -0.0348   1.0000   0.0456
  -6.500  -0.4905   0.06063   0.05721  -0.0365   1.0000   0.0473
  -6.250  -0.4813   0.05605   0.05252  -0.0388   1.0000   0.0499
  -6.000  -0.4659   0.05252   0.04824  -0.0426   1.0000   0.0542
  -5.750  -0.4614   0.04565   0.04156  -0.0425   1.0000   0.0562
  -5.500  -0.4485   0.04306   0.03897  -0.0415   1.0000   0.0582
  -5.250  -0.4342   0.04024   0.03598  -0.0408   1.0000   0.0621
  -5.000  -0.4206   0.03645   0.03174  -0.0404   1.0000   0.0696
  -4.750  -0.3988   0.02906   0.02349  -0.0381   1.0000   0.0400
  -4.500  -0.3819   0.02465   0.01860  -0.0365   1.0000   0.0354
  -4.250  -0.3621   0.02266   0.01623  -0.0350   1.0000   0.0376
  -4.000  -0.3408   0.02020   0.01329  -0.0334   1.0000   0.0367
  -3.750  -0.3182   0.01818   0.01088  -0.0319   1.0000   0.0366
  -3.500  -0.2953   0.01677   0.00921  -0.0306   1.0000   0.0379
  -3.250  -0.2728   0.01578   0.00804  -0.0293   1.0000   0.0404
  -3.000  -0.2515   0.01463   0.00686  -0.0281   1.0000   0.0465
  -2.750  -0.2303   0.01395   0.00612  -0.0267   1.0000   0.0522
  -2.500  -0.2099   0.01324   0.00543  -0.0254   1.0000   0.0640
  -2.250  -0.1896   0.01265   0.00488  -0.0241   1.0000   0.0878
  -2.000  -0.1705   0.01160   0.00449  -0.0230   1.0000   0.2181
  -1.750  -0.1383   0.01072   0.00466  -0.0246   0.9955   0.4871
  -1.500  -0.1072   0.01008   0.00478  -0.0252   0.9895   0.6879
  -1.250  -0.0396   0.00944   0.00462  -0.0332   0.9935   1.0000
  -1.000   0.0026   0.00956   0.00456  -0.0365   0.9872   1.0000
  -0.750   0.0451   0.00970   0.00455  -0.0399   0.9818   1.0000
  -0.500   0.0843   0.00975   0.00448  -0.0425   0.9748   1.0000
  -0.250   0.1256   0.00981   0.00445  -0.0455   0.9691   1.0000
   0.000   0.1653   0.00977   0.00436  -0.0481   0.9612   1.0000
   0.250   0.2085   0.00963   0.00417  -0.0512   0.9517   1.0000
   0.500   0.2563   0.00944   0.00397  -0.0552   0.9454   1.0000
   0.750   0.2944   0.00928   0.00380  -0.0572   0.9358   1.0000
   1.000   0.3392   0.00901   0.00355  -0.0604   0.9270   1.0000
   1.250   0.3895   0.00869   0.00328  -0.0647   0.9172   1.0000
   1.500   0.4353   0.00846   0.00308  -0.0680   0.9023   1.0000
   1.750   0.4758   0.00834   0.00299  -0.0703   0.8832   1.0000
   2.000   0.5122   0.00832   0.00301  -0.0717   0.8642   1.0000
   2.250   0.5399   0.00838   0.00310  -0.0713   0.8415   1.0000
   2.500   0.5665   0.00844   0.00315  -0.0705   0.8135   1.0000
   2.750   0.5908   0.00852   0.00319  -0.0691   0.7774   1.0000
   3.000   0.6138   0.00866   0.00328  -0.0675   0.7333   1.0000
   3.250   0.6351   0.00891   0.00336  -0.0655   0.6761   1.0000
   3.500   0.6527   0.00936   0.00350  -0.0628   0.5901   1.0000
   3.750   0.6600   0.01045   0.00373  -0.0583   0.4177   1.0000
   4.000   0.6647   0.01233   0.00438  -0.0544   0.1786   1.0000
   4.250   0.6825   0.01329   0.00496  -0.0527   0.1130   1.0000
   4.500   0.7014   0.01416   0.00576  -0.0510   0.0885   1.0000
   4.750   0.7174   0.01550   0.00704  -0.0487   0.0582   1.0000
   5.000   0.7376   0.01640   0.00788  -0.0471   0.0436   1.0000
   5.250   0.7573   0.01763   0.00918  -0.0453   0.0381   1.0000
   5.500   0.7787   0.01876   0.01035  -0.0438   0.0343   1.0000
   5.750   0.8000   0.02063   0.01229  -0.0425   0.0305   1.0000
   6.000   0.8244   0.02240   0.01427  -0.0414   0.0288   1.0000
   6.250   0.8487   0.02479   0.01698  -0.0402   0.0281   1.0000
   6.500   0.8711   0.02784   0.02043  -0.0386   0.0282   1.0000
   6.750   0.8896   0.03160   0.02464  -0.0365   0.0293   1.0000
   7.000   0.9042   0.03546   0.02895  -0.0340   0.0298   1.0000
   7.250   0.9170   0.03884   0.03276  -0.0314   0.0295   1.0000
  15.750   0.6963   0.18805   0.18490  -0.0628   0.0310   1.0000
  16.000   0.6940   0.19040   0.18726  -0.0657   0.0308   1.0000
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