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

RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 200,000
Max Cl/Cd: 62.14 at α=2.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf26-il-200000-n5.txt
Download as CSV file: xf-raf26-il-200000-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 =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4866   0.08436   0.08092  -0.0219   1.0000   0.0098
  -7.750  -0.4910   0.08097   0.07758  -0.0227   1.0000   0.0097
  -7.500  -0.4959   0.07740   0.07408  -0.0239   1.0000   0.0095
  -7.250  -0.4959   0.07298   0.06970  -0.0270   1.0000   0.0093
  -7.000  -0.4939   0.06805   0.06478  -0.0307   1.0000   0.0091
  -6.750  -0.4893   0.06271   0.05942  -0.0346   1.0000   0.0089
  -6.500  -0.4826   0.05700   0.05364  -0.0381   1.0000   0.0087
  -6.250  -0.4743   0.05102   0.04752  -0.0408   1.0000   0.0087
  -6.000  -0.4642   0.04532   0.04160  -0.0424   1.0000   0.0085
  -5.750  -0.4528   0.03970   0.03568  -0.0429   1.0000   0.0083
  -5.500  -0.4399   0.03450   0.03010  -0.0425   1.0000   0.0082
  -5.250  -0.4255   0.02966   0.02477  -0.0414   1.0000   0.0082
  -5.000  -0.4091   0.02540   0.01982  -0.0399   1.0000   0.0084
  -4.750  -0.3797   0.02162   0.01533  -0.0405   0.9976   0.0089
  -4.500  -0.3483   0.01881   0.01188  -0.0410   0.9951   0.0109
  -4.250  -0.3171   0.01774   0.01058  -0.0420   0.9917   0.0148
  -4.000  -0.2846   0.01688   0.00948  -0.0431   0.9883   0.0198
  -3.750  -0.2543   0.01551   0.00793  -0.0439   0.9850   0.0245
  -3.500  -0.2240   0.01458   0.00686  -0.0444   0.9806   0.0256
  -3.250  -0.1918   0.01377   0.00591  -0.0454   0.9772   0.0256
  -3.000  -0.1629   0.01315   0.00516  -0.0456   0.9711   0.0258
  -2.750  -0.1303   0.01266   0.00453  -0.0467   0.9669   0.0266
  -2.500  -0.1014   0.01227   0.00402  -0.0469   0.9600   0.0278
  -2.250  -0.0678   0.01192   0.00350  -0.0480   0.9550   0.0309
  -2.000  -0.0390   0.01138   0.00315  -0.0483   0.9470   0.0744
  -1.750  -0.0062   0.01071   0.00288  -0.0497   0.9416   0.1811
  -1.500   0.0214   0.01000   0.00276  -0.0500   0.9335   0.3393
  -1.250   0.0533   0.00940   0.00267  -0.0511   0.9287   0.4809
  -1.000   0.0782   0.00863   0.00265  -0.0503   0.9213   0.6796
  -0.750   0.1536   0.00806   0.00266  -0.0603   0.9259   1.0000
  -0.500   0.1912   0.00801   0.00251  -0.0624   0.9190   1.0000
  -0.250   0.2315   0.00795   0.00236  -0.0651   0.9114   1.0000
   0.000   0.2768   0.00784   0.00218  -0.0687   0.8982   1.0000
   0.250   0.3200   0.00777   0.00202  -0.0719   0.8793   1.0000
   0.500   0.3570   0.00778   0.00196  -0.0737   0.8615   1.0000
   0.750   0.3870   0.00783   0.00196  -0.0741   0.8422   1.0000
   1.000   0.4159   0.00790   0.00199  -0.0741   0.8196   1.0000
   1.250   0.4426   0.00799   0.00201  -0.0736   0.7942   1.0000
   1.500   0.4679   0.00810   0.00203  -0.0728   0.7637   1.0000
   1.750   0.4924   0.00825   0.00206  -0.0718   0.7274   1.0000
   2.000   0.5153   0.00847   0.00210  -0.0705   0.6836   1.0000
   2.250   0.5377   0.00871   0.00227  -0.0692   0.6433   1.0000
   2.500   0.5593   0.00900   0.00240  -0.0677   0.5966   1.0000
   2.750   0.5792   0.00938   0.00255  -0.0659   0.5252   1.0000
   3.000   0.5953   0.01003   0.00277  -0.0634   0.4342   1.0000
   3.250   0.6133   0.01068   0.00310  -0.0615   0.3584   1.0000
   3.500   0.6274   0.01180   0.00352  -0.0592   0.2050   1.0000
   3.750   0.6466   0.01258   0.00398  -0.0578   0.1279   1.0000
   4.000   0.6675   0.01319   0.00441  -0.0566   0.0960   1.0000
   4.250   0.6896   0.01369   0.00491  -0.0555   0.0843   1.0000
   4.500   0.7112   0.01424   0.00550  -0.0544   0.0756   1.0000
   4.750   0.7345   0.01461   0.00608  -0.0536   0.0614   1.0000
   5.000   0.7572   0.01507   0.00636  -0.0527   0.0257   1.0000
   5.250   0.7752   0.01623   0.00751  -0.0507   0.0137   1.0000
   5.500   0.7955   0.01713   0.00859  -0.0491   0.0115   1.0000
   5.750   0.8148   0.01824   0.00986  -0.0473   0.0099   1.0000
   6.000   0.8340   0.01935   0.01104  -0.0459   0.0075   1.0000
   6.250   0.8545   0.02040   0.01223  -0.0445   0.0060   1.0000
   6.500   0.8739   0.02195   0.01400  -0.0428   0.0055   1.0000
   6.750   0.8937   0.02383   0.01613  -0.0412   0.0052   1.0000
   7.000   0.9133   0.02601   0.01862  -0.0396   0.0050   1.0000
   7.250   0.9314   0.02869   0.02171  -0.0377   0.0049   1.0000
   7.500   0.9465   0.03186   0.02537  -0.0354   0.0049   1.0000
   7.750   0.9571   0.03567   0.02971  -0.0326   0.0048   1.0000
   8.000   0.9629   0.03999   0.03455  -0.0294   0.0048   1.0000
   8.250   0.9640   0.04468   0.03970  -0.0262   0.0049   1.0000
   8.500   0.9613   0.04959   0.04499  -0.0230   0.0050   1.0000
   8.750   0.9554   0.05422   0.04993  -0.0200   0.0051   1.0000
   9.000   0.9449   0.05878   0.05474  -0.0172   0.0051   1.0000
   9.250   0.9298   0.06266   0.05880  -0.0143   0.0051   1.0000
   9.500   0.9113   0.06674   0.06303  -0.0122   0.0052   1.0000
   9.750   0.8943   0.07099   0.06741  -0.0122   0.0053   1.0000
  10.000   0.8751   0.07674   0.07328  -0.0147   0.0053   1.0000
  10.250   0.8579   0.08416   0.08077  -0.0199   0.0053   1.0000
<< Back to RAF 26 AIRFOIL (raf26-il)

Polar data table (+)

Polar graphs


<< Back to RAF 26 AIRFOIL (raf26-il)