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RAF 27 AIRFOIL (raf27-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAF 27 AIRFOIL (raf27-il)
Reynolds number: 100,000
Max Cl/Cd: 39.5 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf27-il-100000.txt
Download as CSV file: xf-raf27-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: RAF 27 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6443   0.09158   0.08662  -0.0105   1.0000   0.1673
  -9.250  -0.7829   0.06509   0.05965  -0.0280   1.0000   0.0811
  -9.000  -0.8126   0.05769   0.05160  -0.0252   1.0000   0.0686
  -8.750  -0.8123   0.05318   0.04688  -0.0235   1.0000   0.0678
  -8.500  -0.8119   0.04915   0.04252  -0.0212   1.0000   0.0675
  -8.250  -0.8079   0.04512   0.03816  -0.0190   1.0000   0.0667
  -8.000  -0.8016   0.04129   0.03392  -0.0165   1.0000   0.0657
  -7.750  -0.7922   0.03786   0.03005  -0.0140   1.0000   0.0652
  -7.500  -0.7797   0.03487   0.02660  -0.0116   1.0000   0.0657
  -7.250  -0.7652   0.03284   0.02404  -0.0091   1.0000   0.0683
  -7.000  -0.7480   0.03003   0.02093  -0.0074   1.0000   0.0711
  -6.750  -0.7270   0.02786   0.01862  -0.0061   1.0000   0.0742
  -6.500  -0.7056   0.02625   0.01681  -0.0047   1.0000   0.0792
  -6.250  -0.6836   0.02440   0.01483  -0.0035   1.0000   0.0852
  -6.000  -0.6613   0.02290   0.01332  -0.0023   1.0000   0.0919
  -5.750  -0.6406   0.02145   0.01191  -0.0009   1.0000   0.1013
  -5.500  -0.6211   0.02011   0.01059   0.0009   1.0000   0.1126
  -5.250  -0.6043   0.01887   0.00952   0.0029   1.0000   0.1303
  -5.000  -0.5901   0.01745   0.00842   0.0054   1.0000   0.1667
  -4.750  -0.5823   0.01561   0.00749   0.0088   1.0000   0.2855
  -4.500  -0.5716   0.01460   0.00711   0.0120   1.0000   0.4046
  -4.250  -0.5565   0.01409   0.00683   0.0146   1.0000   0.4813
  -4.000  -0.5405   0.01368   0.00667   0.0173   1.0000   0.5467
  -3.750  -0.5249   0.01342   0.00667   0.0204   1.0000   0.6120
  -3.500  -0.5087   0.01328   0.00669   0.0235   1.0000   0.6695
  -3.250  -0.4908   0.01316   0.00664   0.0262   1.0000   0.7153
  -3.000  -0.4716   0.01309   0.00664   0.0287   1.0000   0.7573
  -2.750  -0.4518   0.01314   0.00677   0.0315   1.0000   0.8019
  -2.500  -0.4274   0.01332   0.00697   0.0335   1.0000   0.8439
  -2.250  -0.3935   0.01357   0.00714   0.0334   1.0000   0.8779
  -2.000  -0.3467   0.01392   0.00736   0.0304   1.0000   0.9047
  -1.750  -0.2943   0.01426   0.00755   0.0261   1.0000   0.9276
  -1.500  -0.2281   0.01459   0.00770   0.0189   1.0000   0.9434
  -1.250  -0.1625   0.01476   0.00775   0.0116   1.0000   0.9585
  -1.000  -0.0975   0.01477   0.00764   0.0040   1.0000   0.9737
  -0.750  -0.0317   0.01461   0.00741  -0.0040   1.0000   0.9885
  -0.500   0.0245   0.01432   0.00708  -0.0106   1.0000   1.0000
  -0.250   0.0170   0.01415   0.00695  -0.0061   1.0000   1.0000
   0.000   0.0000   0.01410   0.00692   0.0000   1.0000   1.0000
   0.250  -0.0169   0.01415   0.00695   0.0061   1.0000   1.0000
   0.500  -0.0244   0.01432   0.00708   0.0106   1.0000   1.0000
   0.750   0.0319   0.01460   0.00741   0.0040   0.9884   1.0000
   1.000   0.0975   0.01477   0.00764  -0.0040   0.9737   1.0000
   1.250   0.1625   0.01476   0.00774  -0.0116   0.9585   1.0000
   1.500   0.2283   0.01458   0.00770  -0.0190   0.9434   1.0000
   1.750   0.2943   0.01426   0.00755  -0.0261   0.9277   1.0000
   2.000   0.3466   0.01392   0.00736  -0.0304   0.9047   1.0000
   2.250   0.3935   0.01358   0.00714  -0.0334   0.8782   1.0000
   2.500   0.4275   0.01332   0.00697  -0.0335   0.8439   1.0000
   2.750   0.4518   0.01314   0.00676  -0.0315   0.8017   1.0000
   3.000   0.4717   0.01309   0.00664  -0.0287   0.7572   1.0000
   3.250   0.4909   0.01316   0.00666  -0.0262   0.7154   1.0000
   3.500   0.5089   0.01328   0.00670  -0.0235   0.6697   1.0000
   3.750   0.5250   0.01342   0.00667  -0.0204   0.6121   1.0000
   4.000   0.5405   0.01369   0.00667  -0.0173   0.5463   1.0000
   4.250   0.5566   0.01409   0.00683  -0.0147   0.4814   1.0000
   4.500   0.5717   0.01460   0.00711  -0.0120   0.4043   1.0000
   4.750   0.5824   0.01561   0.00749  -0.0088   0.2857   1.0000
   5.000   0.5900   0.01747   0.00843  -0.0054   0.1663   1.0000
   5.250   0.6043   0.01886   0.00952  -0.0029   0.1302   1.0000
   5.500   0.6211   0.02011   0.01060  -0.0009   0.1125   1.0000
   5.750   0.6407   0.02145   0.01192   0.0009   0.1015   1.0000
   6.000   0.6614   0.02291   0.01333   0.0022   0.0920   1.0000
   6.250   0.6837   0.02440   0.01483   0.0035   0.0852   1.0000
   6.500   0.7056   0.02624   0.01681   0.0047   0.0791   1.0000
   6.750   0.7270   0.02786   0.01862   0.0061   0.0741   1.0000
   7.000   0.7480   0.03004   0.02094   0.0074   0.0711   1.0000
   7.250   0.7653   0.03289   0.02409   0.0091   0.0684   1.0000
   7.500   0.7797   0.03487   0.02659   0.0116   0.0657   1.0000
   7.750   0.7923   0.03786   0.03005   0.0140   0.0652   1.0000
   8.000   0.8015   0.04129   0.03393   0.0165   0.0657   1.0000
   8.250   0.8078   0.04513   0.03817   0.0190   0.0667   1.0000
   8.500   0.8115   0.04913   0.04251   0.0213   0.0674   1.0000
   8.750   0.8125   0.05321   0.04690   0.0235   0.0679   1.0000
   9.000   0.8141   0.05792   0.05179   0.0250   0.0688   1.0000
   9.250   0.8022   0.06649   0.06081   0.0267   0.0822   1.0000
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