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RAF 26 AIRFOIL (raf26-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: RAF 26 AIRFOIL (raf26-il)
Reynolds number: 100,000
Max Cl/Cd: 53.82 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-raf26-il-100000.txt
Download as CSV file: xf-raf26-il-100000.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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5022   0.10857   0.10343  -0.0165   1.0000   0.0790
  -9.000  -0.5133   0.10651   0.10147  -0.0211   1.0000   0.0802
  -8.750  -0.5257   0.10420   0.09928  -0.0251   1.0000   0.0805
  -8.500  -0.4967   0.09755   0.09255  -0.0188   1.0000   0.0842
  -8.250  -0.4929   0.09439   0.08943  -0.0188   1.0000   0.0876
  -8.000  -0.4077   0.07977   0.07516  -0.0234   1.0000   0.1000
  -7.750  -0.4120   0.07642   0.07186  -0.0240   1.0000   0.1041
  -7.500  -0.4278   0.07340   0.06895  -0.0257   1.0000   0.1064
  -7.250  -0.5034   0.08076   0.07609  -0.0248   1.0000   0.0980
  -7.000  -0.4987   0.07742   0.07277  -0.0256   1.0000   0.1025
  -6.750  -0.5050   0.07207   0.06726  -0.0380   1.0000   0.1084
  -6.500  -0.4922   0.06909   0.06446  -0.0312   1.0000   0.1120
  -6.250  -0.4891   0.06447   0.05959  -0.0387   1.0000   0.1224
  -6.000  -0.4770   0.06214   0.05748  -0.0330   1.0000   0.1304
  -5.750  -0.4686   0.05838   0.05366  -0.0340   1.0000   0.1407
  -5.500  -0.4584   0.05468   0.04987  -0.0353   1.0000   0.1528
  -5.250  -0.4461   0.05248   0.04724  -0.0381   1.0000   0.1767
  -5.000  -0.4342   0.04834   0.04335  -0.0349   1.0000   0.1809
  -4.750  -0.3959   0.03575   0.02892  -0.0410   1.0000   0.0797
  -4.500  -0.3748   0.03088   0.02341  -0.0396   1.0000   0.0660
  -4.250  -0.3548   0.02751   0.01964  -0.0385   1.0000   0.0634
  -4.000  -0.3329   0.02499   0.01663  -0.0371   1.0000   0.0631
  -3.750  -0.3101   0.02368   0.01472  -0.0354   1.0000   0.0672
  -3.500  -0.2871   0.02109   0.01193  -0.0344   1.0000   0.0693
  -3.250  -0.2639   0.01953   0.01025  -0.0332   1.0000   0.0738
  -3.000  -0.2411   0.01826   0.00888  -0.0320   1.0000   0.0843
  -2.750  -0.2186   0.01708   0.00772  -0.0307   1.0000   0.0981
  -2.500  -0.1973   0.01598   0.00674  -0.0294   1.0000   0.1265
  -2.250  -0.1769   0.01428   0.00580  -0.0281   1.0000   0.2423
  -2.000  -0.1182   0.01157   0.00559  -0.0327   1.0000   1.0000
  -1.750  -0.0974   0.01166   0.00534  -0.0317   1.0000   1.0000
  -1.500  -0.0769   0.01178   0.00518  -0.0306   1.0000   1.0000
  -1.250  -0.0565   0.01192   0.00506  -0.0296   1.0000   1.0000
  -1.000  -0.0362   0.01209   0.00503  -0.0285   1.0000   1.0000
  -0.750  -0.0161   0.01229   0.00506  -0.0275   1.0000   1.0000
  -0.500   0.0038   0.01251   0.00513  -0.0266   1.0000   1.0000
  -0.250   0.0236   0.01277   0.00526  -0.0256   1.0000   1.0000
   0.000   0.0433   0.01305   0.00541  -0.0247   1.0000   1.0000
   0.250   0.0627   0.01336   0.00564  -0.0239   1.0000   1.0000
   0.500   0.0820   0.01371   0.00591  -0.0230   1.0000   1.0000
   0.750   0.1012   0.01408   0.00623  -0.0223   1.0000   1.0000
   1.000   0.1402   0.01461   0.00672  -0.0255   0.9935   1.0000
   1.250   0.1851   0.01510   0.00721  -0.0297   0.9845   1.0000
   1.500   0.2276   0.01548   0.00761  -0.0333   0.9745   1.0000
   1.750   0.2721   0.01580   0.00797  -0.0372   0.9634   1.0000
   2.000   0.3226   0.01593   0.00821  -0.0419   0.9496   1.0000
   2.250   0.3780   0.01582   0.00824  -0.0471   0.9345   1.0000
   2.500   0.4363   0.01548   0.00809  -0.0525   0.9194   1.0000
   2.750   0.4941   0.01488   0.00775  -0.0573   0.9023   1.0000
   3.000   0.5554   0.01403   0.00717  -0.0621   0.8798   1.0000
   3.250   0.6015   0.01347   0.00688  -0.0640   0.8521   1.0000
   3.500   0.6358   0.01309   0.00668  -0.0635   0.8136   1.0000
   3.750   0.6643   0.01283   0.00651  -0.0618   0.7616   1.0000
   4.000   0.6878   0.01278   0.00635  -0.0588   0.6736   1.0000
   4.250   0.6921   0.01420   0.00630  -0.0524   0.3878   1.0000
   4.500   0.6947   0.01687   0.00742  -0.0482   0.1732   1.0000
   4.750   0.7106   0.01835   0.00862  -0.0459   0.1205   1.0000
   5.000   0.7254   0.02038   0.01040  -0.0433   0.0876   1.0000
   5.250   0.7479   0.02280   0.01270  -0.0419   0.0731   1.0000
   5.500   0.7730   0.02488   0.01483  -0.0411   0.0623   1.0000
   5.750   0.7999   0.02704   0.01713  -0.0403   0.0575   1.0000
   6.000   0.8263   0.03038   0.02060  -0.0398   0.0549   1.0000
   6.250   0.8490   0.03381   0.02442  -0.0384   0.0541   1.0000
   6.500   0.8679   0.03550   0.02680  -0.0358   0.0515   1.0000
   6.750   0.8852   0.03873   0.03053  -0.0337   0.0513   1.0000
   7.000   0.9001   0.04306   0.03523  -0.0316   0.0530   1.0000
   7.250   0.9108   0.04813   0.04121  -0.0278   0.0641   1.0000
   9.000   0.9004   0.08618   0.08149  -0.0183   0.1182   1.0000
   9.250   0.8614   0.08990   0.08530  -0.0192   0.1171   1.0000
   9.500   0.8335   0.09619   0.09158  -0.0253   0.1157   1.0000
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