Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

RAF 25 AIRFOIL (raf25-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: RAF 25 AIRFOIL (raf25-il)
Reynolds number: 1,000,000
Max Cl/Cd: 72.23 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf25-il-1000000.txt
Download as CSV file: xf-raf25-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4869   0.07482   0.07328  -0.0155   1.0000   0.0122
  -8.750  -0.4920   0.06972   0.06819  -0.0173   1.0000   0.0122
  -8.500  -0.4992   0.06429   0.06279  -0.0194   1.0000   0.0123
  -8.250  -0.5083   0.05884   0.05735  -0.0221   1.0000   0.0123
  -8.000  -0.5416   0.05235   0.05090  -0.0268   1.0000   0.0113
  -7.750  -0.5548   0.04538   0.04385  -0.0316   1.0000   0.0107
  -7.500  -0.6418   0.03850   0.03628  -0.0340   1.0000   0.0071
  -7.250  -0.6681   0.02227   0.01882  -0.0294   1.0000   0.0071
  -7.000  -0.6531   0.01954   0.01572  -0.0276   1.0000   0.0075
  -6.750  -0.6344   0.01785   0.01376  -0.0261   1.0000   0.0077
  -6.500  -0.6142   0.01659   0.01230  -0.0248   1.0000   0.0079
  -6.250  -0.5963   0.01448   0.00984  -0.0231   1.0000   0.0081
  -6.000  -0.5793   0.01222   0.00726  -0.0212   1.0000   0.0089
  -5.750  -0.5566   0.01181   0.00683  -0.0202   1.0000   0.0097
  -5.500  -0.5338   0.01143   0.00640  -0.0193   1.0000   0.0104
  -5.250  -0.5116   0.01089   0.00578  -0.0181   1.0000   0.0113
  -5.000  -0.4892   0.01048   0.00531  -0.0170   1.0000   0.0120
  -4.750  -0.4674   0.00997   0.00472  -0.0158   1.0000   0.0126
  -4.500  -0.4367   0.00910   0.00375  -0.0166   0.9987   0.0150
  -4.250  -0.4005   0.00900   0.00367  -0.0185   0.9967   0.0175
  -4.000  -0.3645   0.00874   0.00338  -0.0204   0.9947   0.0195
  -3.750  -0.3317   0.00824   0.00286  -0.0216   0.9921   0.0243
  -3.500  -0.2979   0.00809   0.00269  -0.0229   0.9892   0.0273
  -3.250  -0.2637   0.00781   0.00238  -0.0244   0.9864   0.0302
  -3.000  -0.2290   0.00747   0.00201  -0.0260   0.9841   0.0348
  -2.750  -0.1964   0.00726   0.00178  -0.0270   0.9804   0.0388
  -2.500  -0.1643   0.00705   0.00155  -0.0280   0.9754   0.0425
  -2.250  -0.1301   0.00677   0.00135  -0.0294   0.9718   0.0588
  -2.000  -0.0991   0.00617   0.00114  -0.0303   0.9657   0.1660
  -1.750  -0.0669   0.00560   0.00099  -0.0316   0.9584   0.2902
  -1.500  -0.0337   0.00515   0.00087  -0.0330   0.9472   0.3923
  -1.250   0.0004   0.00476   0.00078  -0.0345   0.9305   0.4979
  -1.000   0.0290   0.00448   0.00072  -0.0347   0.9106   0.5860
  -0.750   0.0553   0.00433   0.00069  -0.0342   0.8938   0.6452
  -0.500   0.0801   0.00418   0.00068  -0.0334   0.8764   0.7068
  -0.250   0.1023   0.00395   0.00068  -0.0320   0.8577   0.7922
   0.000   0.1251   0.00368   0.00071  -0.0305   0.8348   0.9074
   0.250   0.1877   0.00379   0.00078  -0.0380   0.7981   0.9739
   0.500   0.2219   0.00391   0.00080  -0.0394   0.7736   0.9856
   0.750   0.2587   0.00405   0.00083  -0.0413   0.7411   0.9936
   1.000   0.2976   0.00426   0.00083  -0.0438   0.6859   0.9991
   1.250   0.3236   0.00448   0.00084  -0.0434   0.6274   1.0000
   1.500   0.3450   0.00479   0.00087  -0.0421   0.5553   1.0000
   1.750   0.3666   0.00516   0.00094  -0.0408   0.4759   1.0000
   2.000   0.3867   0.00572   0.00106  -0.0394   0.3594   1.0000
   2.250   0.4058   0.00649   0.00128  -0.0379   0.2125   1.0000
   2.500   0.4243   0.00740   0.00158  -0.0363   0.0603   1.0000
   2.750   0.4484   0.00765   0.00178  -0.0355   0.0494   1.0000
   3.000   0.4732   0.00781   0.00195  -0.0349   0.0451   1.0000
   3.250   0.4978   0.00799   0.00215  -0.0342   0.0407   1.0000
   3.500   0.5216   0.00833   0.00250  -0.0333   0.0341   1.0000
   3.750   0.5471   0.00841   0.00258  -0.0328   0.0311   1.0000
   4.000   0.5711   0.00871   0.00284  -0.0320   0.0232   1.0000
   4.250   0.5959   0.00891   0.00301  -0.0313   0.0190   1.0000
   4.500   0.6183   0.00949   0.00368  -0.0301   0.0155   1.0000
   4.750   0.6428   0.00975   0.00396  -0.0294   0.0142   1.0000
   5.000   0.6669   0.01005   0.00428  -0.0286   0.0127   1.0000
   5.250   0.6898   0.01056   0.00481  -0.0276   0.0111   1.0000
   5.500   0.7080   0.01183   0.00626  -0.0256   0.0100   1.0000
   5.750   0.7318   0.01225   0.00672  -0.0248   0.0096   1.0000
   6.000   0.7547   0.01285   0.00742  -0.0238   0.0091   1.0000
   6.250   0.7772   0.01357   0.00822  -0.0228   0.0084   1.0000
   6.500   0.7993   0.01445   0.00920  -0.0216   0.0079   1.0000
   6.750   0.8213   0.01542   0.01028  -0.0205   0.0075   1.0000
   7.000   0.8436   0.01621   0.01114  -0.0197   0.0071   1.0000
   7.250   0.8620   0.01817   0.01332  -0.0181   0.0066   1.0000
   7.500   0.8755   0.02169   0.01730  -0.0158   0.0064   1.0000
   7.750   0.8890   0.02481   0.02080  -0.0136   0.0064   1.0000
   8.000   0.8942   0.02955   0.02606  -0.0107   0.0064   1.0000
   8.250   0.9016   0.03330   0.03017  -0.0083   0.0064   1.0000
   8.500   0.8877   0.04126   0.03869  -0.0043   0.0064   1.0000
   8.750   0.9022   0.04306   0.04061  -0.0031   0.0065   1.0000
<< Back to RAF 25 AIRFOIL (raf25-il)

Polar data table (+)

Polar graphs


<< Back to RAF 25 AIRFOIL (raf25-il)