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=100,000 Ncrit=5


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 100,000
Max Cl/Cd: 51.19 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-raf26-il-100000-n5.txt
Download as CSV file: xf-raf26-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4844   0.09041   0.08545  -0.0229   1.0000   0.0250
  -8.000  -0.4856   0.08688   0.08199  -0.0235   1.0000   0.0241
  -7.750  -0.4889   0.08346   0.07864  -0.0244   1.0000   0.0234
  -7.500  -0.4916   0.07967   0.07492  -0.0259   1.0000   0.0227
  -7.250  -0.4915   0.07515   0.07046  -0.0289   1.0000   0.0220
  -7.000  -0.4892   0.07025   0.06559  -0.0324   1.0000   0.0213
  -6.750  -0.4849   0.06484   0.06015  -0.0361   1.0000   0.0206
  -6.500  -0.4786   0.05898   0.05421  -0.0395   1.0000   0.0199
  -6.250  -0.4705   0.05266   0.04770  -0.0424   1.0000   0.0190
  -6.000  -0.4604   0.04613   0.04084  -0.0441   1.0000   0.0182
  -5.750  -0.4478   0.04084   0.03513  -0.0445   1.0000   0.0177
  -5.500  -0.4335   0.03654   0.03038  -0.0442   1.0000   0.0177
  -5.250  -0.4179   0.03367   0.02719  -0.0437   1.0000   0.0188
  -5.000  -0.4004   0.03124   0.02443  -0.0429   1.0000   0.0205
  -4.750  -0.3813   0.02827   0.02094  -0.0418   1.0000   0.0215
  -4.500  -0.3606   0.02544   0.01756  -0.0405   1.0000   0.0220
  -4.250  -0.3388   0.02308   0.01471  -0.0392   1.0000   0.0229
  -4.000  -0.3162   0.02117   0.01239  -0.0379   1.0000   0.0247
  -3.750  -0.2940   0.01973   0.01065  -0.0367   1.0000   0.0281
  -3.500  -0.2725   0.01855   0.00938  -0.0355   1.0000   0.0313
  -3.250  -0.2506   0.01753   0.00818  -0.0342   1.0000   0.0369
  -3.000  -0.2290   0.01675   0.00732  -0.0331   1.0000   0.0470
  -2.750  -0.2073   0.01595   0.00657  -0.0321   1.0000   0.0691
  -2.500  -0.1857   0.01524   0.00597  -0.0312   1.0000   0.1110
  -2.250  -0.1590   0.01454   0.00553  -0.0314   0.9981   0.1857
  -2.000  -0.1270   0.01370   0.00540  -0.0330   0.9936   0.3740
  -1.750  -0.0951   0.01310   0.00534  -0.0341   0.9890   0.5424
  -1.500  -0.0652   0.01250   0.00525  -0.0341   0.9844   0.7140
  -1.250  -0.0078   0.01201   0.00499  -0.0399   0.9851   1.0000
  -1.000   0.0292   0.01211   0.00484  -0.0421   0.9786   1.0000
  -0.750   0.0647   0.01217   0.00471  -0.0440   0.9702   1.0000
  -0.500   0.0999   0.01222   0.00461  -0.0458   0.9615   1.0000
  -0.250   0.1367   0.01226   0.00451  -0.0478   0.9532   1.0000
   0.000   0.1727   0.01228   0.00443  -0.0496   0.9444   1.0000
   0.250   0.2072   0.01230   0.00439  -0.0511   0.9354   1.0000
   0.500   0.2462   0.01228   0.00433  -0.0534   0.9288   1.0000
   0.750   0.2804   0.01228   0.00431  -0.0548   0.9194   1.0000
   1.000   0.3198   0.01223   0.00427  -0.0571   0.9108   1.0000
   1.250   0.3621   0.01210   0.00418  -0.0598   0.8985   1.0000
   1.500   0.4006   0.01201   0.00412  -0.0616   0.8830   1.0000
   1.750   0.4362   0.01199   0.00416  -0.0629   0.8677   1.0000
   2.000   0.4701   0.01198   0.00425  -0.0639   0.8504   1.0000
   2.250   0.5038   0.01197   0.00430  -0.0646   0.8298   1.0000
   2.500   0.5342   0.01199   0.00438  -0.0646   0.8032   1.0000
   2.750   0.5618   0.01203   0.00450  -0.0640   0.7691   1.0000
   3.000   0.5902   0.01211   0.00460  -0.0635   0.7323   1.0000
   3.250   0.6167   0.01226   0.00474  -0.0625   0.6840   1.0000
   3.500   0.6414   0.01253   0.00494  -0.0613   0.6240   1.0000
   3.750   0.6613   0.01306   0.00514  -0.0591   0.5121   1.0000
   4.000   0.6766   0.01403   0.00551  -0.0564   0.3960   1.0000
   4.250   0.6899   0.01540   0.00614  -0.0539   0.2271   1.0000
   4.500   0.7062   0.01673   0.00699  -0.0520   0.1377   1.0000
   4.750   0.7227   0.01804   0.00787  -0.0504   0.0837   1.0000
   5.000   0.7426   0.01901   0.00895  -0.0489   0.0529   1.0000
   5.250   0.7621   0.02009   0.00996  -0.0472   0.0346   1.0000
   5.500   0.7814   0.02119   0.01114  -0.0455   0.0272   1.0000
   5.750   0.7993   0.02262   0.01270  -0.0436   0.0237   1.0000
   6.000   0.8187   0.02431   0.01458  -0.0418   0.0219   1.0000
   6.250   0.8401   0.02612   0.01658  -0.0404   0.0196   1.0000
   6.500   0.8606   0.02800   0.01868  -0.0392   0.0170   1.0000
   6.750   0.8809   0.03126   0.02227  -0.0379   0.0160   1.0000
   7.000   0.8998   0.03450   0.02593  -0.0362   0.0157   1.0000
   7.250   0.9158   0.03769   0.02963  -0.0342   0.0155   1.0000
   7.500   0.9279   0.04120   0.03364  -0.0319   0.0154   1.0000
   7.750   0.9359   0.04504   0.03798  -0.0293   0.0154   1.0000
   8.000   0.9406   0.04895   0.04234  -0.0266   0.0153   1.0000
   8.250   0.9423   0.05282   0.04659  -0.0241   0.0153   1.0000
   8.500   0.9403   0.05679   0.05095  -0.0216   0.0151   1.0000
   8.750   0.9343   0.06086   0.05531  -0.0192   0.0150   1.0000
   9.000   0.9247   0.06482   0.05953  -0.0169   0.0148   1.0000
   9.250   0.9098   0.06848   0.06336  -0.0146   0.0149   1.0000
   9.500   0.8938   0.07227   0.06727  -0.0134   0.0152   1.0000
   9.750   0.8764   0.07685   0.07198  -0.0141   0.0151   1.0000
  10.000   0.8601   0.08210   0.07731  -0.0165   0.0153   1.0000
  10.250   0.8455   0.08818   0.08344  -0.0202   0.0156   1.0000
<< Back to RAF 26 AIRFOIL (raf26-il)

Polar data table (+)

Polar graphs


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