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RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 1,000,000
Max Cl/Cd: 94.08 at α=1.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf26-il-1000000.txt
Download as CSV file: xf-raf26-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.4219   0.09515   0.09358  -0.0190   1.0000   0.0079
  -9.750  -0.4253   0.09045   0.08890  -0.0199   1.0000   0.0080
  -9.500  -0.4283   0.08600   0.08446  -0.0208   1.0000   0.0082
  -9.250  -0.4303   0.08191   0.08037  -0.0215   1.0000   0.0083
  -9.000  -0.4315   0.07818   0.07665  -0.0221   1.0000   0.0084
  -8.750  -0.4311   0.07507   0.07355  -0.0225   1.0000   0.0088
  -8.500  -0.4361   0.07058   0.06909  -0.0235   1.0000   0.0087
  -8.250  -0.4401   0.06678   0.06530  -0.0241   1.0000   0.0090
  -8.000  -0.4494   0.06222   0.06076  -0.0251   1.0000   0.0089
  -7.750  -0.4620   0.05833   0.05690  -0.0254   1.0000   0.0090
  -5.500  -0.4269   0.01503   0.01067  -0.0459   0.9923   0.0081
  -5.250  -0.3992   0.01269   0.00800  -0.0461   0.9900   0.0084
  -5.000  -0.3683   0.01142   0.00657  -0.0470   0.9878   0.0090
  -4.750  -0.3348   0.01075   0.00582  -0.0483   0.9859   0.0099
  -4.500  -0.3004   0.01020   0.00520  -0.0498   0.9844   0.0110
  -4.250  -0.2695   0.00966   0.00458  -0.0505   0.9815   0.0118
  -4.000  -0.2401   0.00919   0.00404  -0.0508   0.9771   0.0123
  -3.750  -0.2091   0.00829   0.00302  -0.0515   0.9738   0.0153
  -3.500  -0.1753   0.00798   0.00268  -0.0528   0.9716   0.0179
  -3.250  -0.1493   0.00772   0.00239  -0.0523   0.9639   0.0195
  -3.000  -0.1181   0.00726   0.00189  -0.0529   0.9591   0.0289
  -2.750  -0.0910   0.00694   0.00160  -0.0527   0.9497   0.0430
  -2.500  -0.0571   0.00656   0.00136  -0.0541   0.9430   0.0815
  -2.250  -0.0153   0.00615   0.00113  -0.0573   0.9341   0.1351
  -2.000   0.0277   0.00562   0.00096  -0.0611   0.9271   0.2509
  -1.750   0.0702   0.00519   0.00085  -0.0648   0.9169   0.3671
  -1.500   0.1069   0.00500   0.00077  -0.0669   0.9012   0.4320
  -1.250   0.1360   0.00487   0.00073  -0.0672   0.8844   0.4895
  -1.000   0.1616   0.00472   0.00072  -0.0667   0.8694   0.5585
  -0.750   0.1847   0.00447   0.00073  -0.0656   0.8556   0.6570
  -0.500   0.2065   0.00429   0.00073  -0.0641   0.8374   0.7399
  -0.250   0.2270   0.00408   0.00074  -0.0623   0.8183   0.8318
   0.000   0.3133   0.00395   0.00082  -0.0755   0.7875   0.9807
   0.250   0.3545   0.00414   0.00085  -0.0785   0.7541   0.9977
   0.500   0.3855   0.00429   0.00085  -0.0792   0.7185   1.0000
   0.750   0.4073   0.00444   0.00086  -0.0778   0.6844   1.0000
   1.000   0.4295   0.00461   0.00088  -0.0765   0.6497   1.0000
   1.250   0.4516   0.00480   0.00093  -0.0752   0.6132   1.0000
   1.500   0.4729   0.00505   0.00098  -0.0738   0.5578   1.0000
   1.750   0.4889   0.00572   0.00112  -0.0714   0.4278   1.0000
   2.000   0.5105   0.00606   0.00126  -0.0701   0.3769   1.0000
   2.250   0.5271   0.00682   0.00147  -0.0680   0.2343   1.0000
   2.500   0.5461   0.00743   0.00171  -0.0663   0.1448   1.0000
   2.750   0.5673   0.00788   0.00191  -0.0650   0.0887   1.0000
   3.000   0.5911   0.00810   0.00210  -0.0641   0.0799   1.0000
   3.250   0.6156   0.00826   0.00229  -0.0634   0.0757   1.0000
   3.500   0.6393   0.00850   0.00251  -0.0625   0.0685   1.0000
   3.750   0.6643   0.00862   0.00265  -0.0619   0.0650   1.0000
   4.000   0.6888   0.00879   0.00279  -0.0612   0.0559   1.0000
   4.250   0.7108   0.00920   0.00302  -0.0600   0.0230   1.0000
   4.500   0.7334   0.00959   0.00339  -0.0589   0.0160   1.0000
   4.750   0.7554   0.01006   0.00394  -0.0575   0.0133   1.0000
   5.000   0.7784   0.01041   0.00434  -0.0565   0.0123   1.0000
   5.250   0.8007   0.01084   0.00483  -0.0553   0.0113   1.0000
   5.500   0.8226   0.01131   0.00534  -0.0540   0.0100   1.0000
   5.750   0.8402   0.01229   0.00645  -0.0519   0.0087   1.0000
   6.000   0.8580   0.01335   0.00764  -0.0498   0.0082   1.0000
   6.250   0.8791   0.01404   0.00840  -0.0485   0.0080   1.0000
   6.500   0.9008   0.01467   0.00912  -0.0472   0.0076   1.0000
   6.750   0.9219   0.01546   0.01000  -0.0459   0.0072   1.0000
   7.000   0.9432   0.01623   0.01085  -0.0447   0.0066   1.0000
   7.250   0.9637   0.01721   0.01194  -0.0433   0.0063   1.0000
   7.500   0.9841   0.01817   0.01303  -0.0420   0.0059   1.0000
   7.750   1.0030   0.01946   0.01446  -0.0404   0.0056   1.0000
   8.000   1.0194   0.02146   0.01670  -0.0384   0.0055   1.0000
   8.250   1.0040   0.03203   0.02848  -0.0303   0.0069   1.0000
   8.500   1.0479   0.02572   0.02146  -0.0340   0.0054   1.0000
   8.750   1.0566   0.02857   0.02464  -0.0312   0.0053   1.0000
  14.750   0.6848   0.16904   0.16761  -0.0623   0.0079   1.0000
  15.000   0.6874   0.17319   0.17176  -0.0647   0.0070   1.0000
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