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RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 500,000
Max Cl/Cd: 69.76 at α=1.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf26-il-500000-n5.txt
Download as CSV file: xf-raf26-il-500000-n5.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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5008   0.09019   0.08795  -0.0183   1.0000   0.0051
  -8.500  -0.5010   0.08695   0.08474  -0.0193   1.0000   0.0049
  -8.250  -0.5035   0.08331   0.08112  -0.0204   1.0000   0.0048
  -8.000  -0.5093   0.07964   0.07750  -0.0213   1.0000   0.0048
  -7.750  -0.5173   0.07629   0.07420  -0.0218   1.0000   0.0047
  -7.500  -0.5179   0.07179   0.06972  -0.0251   1.0000   0.0046
  -7.250  -0.5174   0.06624   0.06418  -0.0296   1.0000   0.0045
  -7.000  -0.5137   0.06023   0.05813  -0.0343   1.0000   0.0044
  -6.750  -0.5083   0.05344   0.05125  -0.0385   1.0000   0.0043
  -6.500  -0.4984   0.04548   0.04309  -0.0425   0.9995   0.0042
  -6.250  -0.4738   0.03244   0.02942  -0.0494   0.9957   0.0039
  -6.000  -0.4535   0.02190   0.01783  -0.0509   0.9915   0.0039
  -5.750  -0.4258   0.01872   0.01411  -0.0517   0.9888   0.0041
  -5.500  -0.3972   0.01649   0.01146  -0.0522   0.9865   0.0046
  -5.250  -0.3702   0.01495   0.00957  -0.0522   0.9827   0.0050
  -5.000  -0.3407   0.01374   0.00812  -0.0528   0.9797   0.0057
  -4.750  -0.3128   0.01271   0.00694  -0.0530   0.9759   0.0062
  -4.500  -0.2853   0.01189   0.00600  -0.0529   0.9707   0.0067
  -4.250  -0.2570   0.01114   0.00513  -0.0531   0.9659   0.0074
  -3.750  -0.2040   0.00988   0.00363  -0.0524   0.9490   0.0100
  -3.500  -0.1725   0.00935   0.00301  -0.0531   0.9427   0.0130
  -3.250  -0.1406   0.00897   0.00256  -0.0540   0.9355   0.0176
  -3.000  -0.1049   0.00857   0.00217  -0.0557   0.9300   0.0339
  -2.750  -0.0680   0.00822   0.00189  -0.0579   0.9247   0.0584
  -2.500  -0.0286   0.00788   0.00166  -0.0606   0.9193   0.0988
  -2.250   0.0139   0.00749   0.00141  -0.0641   0.9111   0.1544
  -2.000   0.0558   0.00717   0.00121  -0.0675   0.8967   0.2194
  -1.750   0.0888   0.00690   0.00109  -0.0688   0.8806   0.2929
  -1.500   0.1166   0.00663   0.00104  -0.0689   0.8671   0.3853
  -1.250   0.1433   0.00645   0.00099  -0.0687   0.8537   0.4460
  -1.000   0.1693   0.00633   0.00096  -0.0683   0.8391   0.4947
  -0.750   0.1947   0.00622   0.00092  -0.0678   0.8241   0.5418
  -0.500   0.2183   0.00603   0.00092  -0.0668   0.8069   0.6178
  -0.250   0.2382   0.00571   0.00093  -0.0649   0.7861   0.7383
   0.000   0.2634   0.00535   0.00095  -0.0640   0.7653   0.8815
   0.500   0.3768   0.00573   0.00098  -0.0771   0.6712   1.0000
   0.750   0.3989   0.00592   0.00101  -0.0758   0.6368   1.0000
   1.000   0.4212   0.00612   0.00106  -0.0745   0.6039   1.0000
   1.250   0.4430   0.00635   0.00112  -0.0732   0.5572   1.0000
   1.500   0.4609   0.00686   0.00122  -0.0712   0.4577   1.0000
   1.750   0.4824   0.00722   0.00138  -0.0699   0.4113   1.0000
   2.000   0.5051   0.00751   0.00153  -0.0689   0.3738   1.0000
   2.250   0.5235   0.00818   0.00172  -0.0672   0.2549   1.0000
   2.500   0.5434   0.00875   0.00200  -0.0657   0.1785   1.0000
   2.750   0.5643   0.00926   0.00224  -0.0644   0.1076   1.0000
   3.000   0.5873   0.00958   0.00246  -0.0635   0.0889   1.0000
   3.250   0.6112   0.00982   0.00270  -0.0627   0.0810   1.0000
   3.500   0.6350   0.01008   0.00299  -0.0619   0.0734   1.0000
   3.750   0.6592   0.01030   0.00323  -0.0611   0.0673   1.0000
   4.000   0.6833   0.01052   0.00344  -0.0604   0.0545   1.0000
   4.250   0.7050   0.01101   0.00372  -0.0593   0.0200   1.0000
   4.500   0.7275   0.01145   0.00417  -0.0581   0.0122   1.0000
   4.750   0.7492   0.01200   0.00483  -0.0568   0.0086   1.0000
   5.000   0.7717   0.01247   0.00539  -0.0557   0.0077   1.0000
   5.250   0.7935   0.01303   0.00605  -0.0544   0.0069   1.0000
   5.500   0.8151   0.01360   0.00669  -0.0532   0.0061   1.0000
   5.750   0.8345   0.01446   0.00760  -0.0516   0.0051   1.0000
   6.000   0.8555   0.01516   0.00840  -0.0503   0.0049   1.0000
   6.250   0.8758   0.01602   0.00941  -0.0488   0.0045   1.0000
   6.500   0.8958   0.01701   0.01053  -0.0473   0.0040   1.0000
   6.750   0.9157   0.01815   0.01181  -0.0457   0.0039   1.0000
   7.000   0.9356   0.01936   0.01318  -0.0443   0.0037   1.0000
   7.250   0.9549   0.02075   0.01476  -0.0428   0.0035   1.0000
   7.500   0.9734   0.02245   0.01669  -0.0411   0.0035   1.0000
   7.750   0.9905   0.02431   0.01880  -0.0394   0.0033   1.0000
   8.000   1.0052   0.02654   0.02133  -0.0373   0.0033   1.0000
   8.250   1.0140   0.02964   0.02479  -0.0347   0.0031   1.0000
   8.500   1.0251   0.03233   0.02793  -0.0320   0.0029   1.0000
   8.750   1.0268   0.03673   0.03286  -0.0284   0.0027   1.0000
   9.000   1.0194   0.04225   0.03886  -0.0241   0.0026   1.0000
   9.250   1.0032   0.04927   0.04629  -0.0197   0.0026   1.0000
   9.500   0.9874   0.05463   0.05190  -0.0160   0.0026   1.0000
   9.750   0.9670   0.05871   0.05615  -0.0121   0.0026   1.0000
  10.000   0.9504   0.06200   0.05957  -0.0101   0.0027   1.0000
  10.250   0.9288   0.06702   0.06472  -0.0101   0.0027   1.0000
  10.500   0.9064   0.07351   0.07133  -0.0130   0.0027   1.0000
  10.750   0.8897   0.08087   0.07878  -0.0186   0.0028   1.0000
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