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RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 50,000
Max Cl/Cd: 28.3 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf25-il-50000.txt
Download as CSV file: xf-raf25-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5492   0.09987   0.09291   0.0074   1.0000   0.2792
  -8.000  -0.5489   0.09702   0.09011   0.0077   1.0000   0.2964
  -7.750  -0.5539   0.09461   0.08780   0.0078   1.0000   0.3123
  -7.500  -0.5384   0.09018   0.08338   0.0094   1.0000   0.3321
  -7.250  -0.5353   0.08713   0.08038   0.0104   1.0000   0.3516
  -7.000  -0.5440   0.08489   0.07827   0.0114   1.0000   0.3700
  -6.750  -0.5251   0.08071   0.07406   0.0132   1.0000   0.3928
  -6.500  -0.5193   0.07744   0.07085   0.0149   1.0000   0.4156
  -6.250  -0.5152   0.07449   0.06796   0.0171   1.0000   0.4420
  -6.000  -0.5022   0.07132   0.06475   0.0198   1.0000   0.4764
  -5.500  -0.5114   0.04569   0.03763  -0.0300   1.0000   0.1425
  -5.250  -0.4916   0.04079   0.03203  -0.0302   1.0000   0.1296
  -5.000  -0.4710   0.03677   0.02743  -0.0296   1.0000   0.1242
  -4.750  -0.4497   0.03387   0.02406  -0.0287   1.0000   0.1284
  -4.500  -0.4264   0.03109   0.02067  -0.0276   1.0000   0.1311
  -4.250  -0.4013   0.02853   0.01751  -0.0264   1.0000   0.1329
  -4.000  -0.3774   0.02625   0.01512  -0.0254   1.0000   0.1428
  -3.750  -0.3525   0.02434   0.01303  -0.0244   1.0000   0.1559
  -3.500  -0.3260   0.02247   0.01106  -0.0232   1.0000   0.1690
  -3.250  -0.2995   0.02083   0.00948  -0.0224   1.0000   0.1990
  -3.000  -0.2754   0.01890   0.00802  -0.0212   1.0000   0.2665
  -2.750  -0.1714   0.01545   0.00684  -0.0289   1.0000   1.0000
  -2.500  -0.1598   0.01512   0.00618  -0.0270   1.0000   1.0000
  -2.250  -0.1446   0.01493   0.00566  -0.0254   1.0000   1.0000
  -2.000  -0.1279   0.01481   0.00526  -0.0238   1.0000   1.0000
  -1.750  -0.1100   0.01474   0.00493  -0.0223   1.0000   1.0000
  -1.500  -0.0915   0.01470   0.00462  -0.0209   1.0000   1.0000
  -1.250  -0.0724   0.01470   0.00442  -0.0196   1.0000   1.0000
  -1.000  -0.0531   0.01472   0.00427  -0.0183   1.0000   1.0000
  -0.750  -0.0337   0.01477   0.00418  -0.0170   1.0000   1.0000
  -0.500  -0.0141   0.01484   0.00413  -0.0158   1.0000   1.0000
  -0.250   0.0055   0.01495   0.00411  -0.0146   1.0000   1.0000
   0.000   0.0249   0.01507   0.00416  -0.0135   1.0000   1.0000
   0.250   0.0444   0.01523   0.00425  -0.0124   1.0000   1.0000
   0.500   0.0637   0.01542   0.00440  -0.0113   1.0000   1.0000
   0.750   0.0828   0.01563   0.00459  -0.0102   1.0000   1.0000
   1.000   0.1018   0.01588   0.00484  -0.0092   1.0000   1.0000
   1.250   0.1206   0.01617   0.00514  -0.0082   1.0000   1.0000
   1.500   0.1392   0.01649   0.00549  -0.0073   1.0000   1.0000
   1.750   0.1575   0.01686   0.00591  -0.0065   1.0000   1.0000
   2.000   0.1755   0.01728   0.00639  -0.0057   1.0000   1.0000
   2.250   0.1931   0.01775   0.00698  -0.0050   1.0000   1.0000
   2.500   0.2104   0.01828   0.00761  -0.0044   1.0000   1.0000
   2.750   0.2270   0.01889   0.00833  -0.0039   1.0000   1.0000
   3.000   0.2432   0.01957   0.00914  -0.0035   1.0000   1.0000
   3.250   0.2588   0.02034   0.01006  -0.0032   1.0000   1.0000
   3.500   0.2737   0.02122   0.01111  -0.0031   1.0000   1.0000
   3.750   0.3772   0.02285   0.01356  -0.0191   0.9556   1.0000
   4.000   0.5372   0.01898   0.00875  -0.0206   0.3109   1.0000
   4.250   0.5521   0.02207   0.01056  -0.0184   0.1942   1.0000
   4.500   0.5772   0.02409   0.01239  -0.0173   0.1568   1.0000
   4.750   0.6077   0.02641   0.01460  -0.0166   0.1380   1.0000
   5.000   0.6366   0.02870   0.01705  -0.0159   0.1233   1.0000
   5.250   0.6658   0.03125   0.01997  -0.0149   0.1183   1.0000
   5.500   0.6919   0.03450   0.02326  -0.0143   0.1122   1.0000
   5.750   0.7147   0.03702   0.02649  -0.0126   0.1086   1.0000
   6.000   0.7370   0.04054   0.03050  -0.0113   0.1100   1.0000
   6.250   0.7578   0.04424   0.03473  -0.0099   0.1138   1.0000
   6.500   0.7734   0.04876   0.04006  -0.0084   0.1232   1.0000
   6.750   0.7873   0.05387   0.04584  -0.0075   0.1370   1.0000
   7.000   0.7953   0.06044   0.05317  -0.0081   0.1640   1.0000
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