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NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA M13 AIRFOIL (m13-il)
Reynolds number: 100,000
Max Cl/Cd: 56.73 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m13-il-100000.txt
Download as CSV file: xf-m13-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M13 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4794   0.12001   0.11522   0.0031   1.0000   0.0496
  -8.750  -0.4802   0.11842   0.11370  -0.0012   1.0000   0.0499
  -8.500  -0.4803   0.11647   0.11181  -0.0060   1.0000   0.0500
  -8.250  -0.4579   0.10701   0.10230   0.0013   1.0000   0.0524
  -8.000  -0.4510   0.10356   0.09887   0.0004   1.0000   0.0543
  -7.750  -0.4455   0.10037   0.09570  -0.0013   1.0000   0.0564
  -7.500  -0.4414   0.09746   0.09284  -0.0039   1.0000   0.0585
  -7.250  -0.4341   0.09529   0.09071  -0.0115   1.0000   0.0603
  -7.000  -0.4182   0.09364   0.08898  -0.0229   1.0000   0.0610
  -6.750  -0.4145   0.08674   0.08220  -0.0168   1.0000   0.0626
  -6.500  -0.4048   0.08289   0.07838  -0.0155   1.0000   0.0659
  -6.250  -0.3905   0.07946   0.07493  -0.0192   1.0000   0.0699
  -6.000  -0.3566   0.07862   0.07377  -0.0338   1.0000   0.0732
  -5.750  -0.3531   0.07192   0.06727  -0.0300   1.0000   0.0745
  -5.500  -0.3423   0.06800   0.06341  -0.0289   1.0000   0.0771
  -5.250  -0.3239   0.06465   0.06001  -0.0312   1.0000   0.0816
  -5.000  -0.2947   0.06143   0.05654  -0.0378   1.0000   0.0866
  -4.750  -0.2853   0.05751   0.05271  -0.0361   1.0000   0.0895
  -4.500  -0.2553   0.05573   0.05057  -0.0407   1.0000   0.0995
  -4.250  -0.2508   0.05162   0.04668  -0.0382   1.0000   0.1025
  -4.000  -0.2393   0.04938   0.04441  -0.0374   1.0000   0.1083
  -3.750  -0.2265   0.04741   0.04224  -0.0376   1.0000   0.1146
  -3.500  -0.2234   0.04532   0.04021  -0.0353   1.0000   0.1178
  -3.250  -0.2083   0.04426   0.03885  -0.0355   1.0000   0.1279
  -3.000  -0.1671   0.04047   0.03492  -0.0404   0.9930   0.1430
  -2.750  -0.1218   0.03762   0.03180  -0.0457   0.9845   0.1703
  -2.500  -0.0803   0.03461   0.02865  -0.0500   0.9755   0.1987
  -2.250  -0.0386   0.03159   0.02559  -0.0543   0.9684   0.2422
  -2.000  -0.0036   0.02882   0.02287  -0.0569   0.9588   0.3053
  -1.750   0.0318   0.02606   0.02016  -0.0590   0.9499   0.3771
  -1.500   0.1244   0.02356   0.01541  -0.0657   0.9430   0.1106
  -1.250   0.1667   0.02187   0.01315  -0.0673   0.9319   0.0985
  -1.000   0.2046   0.02027   0.01136  -0.0689   0.9205   0.0966
  -0.750   0.2397   0.01940   0.01021  -0.0696   0.9081   0.0992
  -0.500   0.2703   0.01834   0.00917  -0.0698   0.8949   0.1039
  -0.250   0.2987   0.01765   0.00841  -0.0693   0.8814   0.1062
   0.000   0.3252   0.01709   0.00780  -0.0683   0.8677   0.1106
   0.250   0.3498   0.01652   0.00730  -0.0670   0.8543   0.1190
   0.500   0.3748   0.01608   0.00684  -0.0658   0.8412   0.1430
   0.750   0.4115   0.01357   0.00639  -0.0670   0.8285   1.0000
   1.000   0.4358   0.01377   0.00631  -0.0658   0.8148   1.0000
   1.250   0.4602   0.01398   0.00634  -0.0647   0.8012   1.0000
   1.500   0.4847   0.01420   0.00643  -0.0637   0.7879   1.0000
   1.750   0.5094   0.01442   0.00654  -0.0629   0.7748   1.0000
   2.000   0.5342   0.01466   0.00668  -0.0620   0.7620   1.0000
   2.250   0.5592   0.01491   0.00685  -0.0612   0.7495   1.0000
   2.500   0.5843   0.01516   0.00706  -0.0604   0.7373   1.0000
   2.750   0.6095   0.01540   0.00725  -0.0596   0.7257   1.0000
   3.000   0.6349   0.01566   0.00748  -0.0589   0.7135   1.0000
   3.250   0.6604   0.01597   0.00779  -0.0583   0.7007   1.0000
   3.500   0.6860   0.01628   0.00812  -0.0578   0.6881   1.0000
   3.750   0.7115   0.01660   0.00851  -0.0573   0.6757   1.0000
   4.000   0.7371   0.01691   0.00886  -0.0567   0.6637   1.0000
   4.250   0.7628   0.01721   0.00921  -0.0561   0.6521   1.0000
   4.500   0.7887   0.01747   0.00951  -0.0553   0.6413   1.0000
   4.750   0.8144   0.01783   0.01004  -0.0549   0.6287   1.0000
   5.000   0.8401   0.01821   0.01056  -0.0545   0.6162   1.0000
   5.250   0.8655   0.01846   0.01096  -0.0537   0.6024   1.0000
   5.500   0.8876   0.01758   0.00995  -0.0507   0.5686   1.0000
   5.750   0.9094   0.01687   0.00922  -0.0482   0.5275   1.0000
   6.000   0.9293   0.01638   0.00867  -0.0456   0.4584   1.0000
   6.250   0.9301   0.01950   0.00963  -0.0421   0.1126   1.0000
   6.500   0.9453   0.02179   0.01165  -0.0404   0.0756   1.0000
   6.750   0.9626   0.02339   0.01333  -0.0389   0.0651   1.0000
   7.000   0.9773   0.02522   0.01515  -0.0371   0.0587   1.0000
   7.250   0.9959   0.02682   0.01682  -0.0354   0.0556   1.0000
   7.500   1.0170   0.02867   0.01870  -0.0340   0.0533   1.0000
   7.750   1.0409   0.03080   0.02083  -0.0329   0.0513   1.0000
   8.000   1.0667   0.03470   0.02459  -0.0326   0.0480   1.0000
   8.250   1.0904   0.03693   0.02715  -0.0316   0.0472   1.0000
   8.500   1.1139   0.04060   0.03110  -0.0307   0.0476   1.0000
   8.750   1.1363   0.04329   0.03416  -0.0294   0.0488   1.0000
   9.000   1.1518   0.04617   0.03799  -0.0266   0.0534   1.0000
   9.250   1.1654   0.05137   0.04367  -0.0250   0.0579   1.0000
   9.500   1.1743   0.05685   0.04996  -0.0224   0.0695   1.0000
   9.750   1.1155   0.05572   0.04957  -0.0147   0.0850   1.0000
  10.250   1.0391   0.06543   0.06058  -0.0071   0.1046   1.0000
  10.500   1.0103   0.06983   0.06511  -0.0060   0.1039   1.0000
  10.750   0.9831   0.07483   0.07021  -0.0062   0.1027   1.0000
  11.000   0.9582   0.07889   0.07427  -0.0069   0.0926   1.0000
  11.250   0.9303   0.08487   0.08032  -0.0091   0.0902   1.0000
  11.500   0.9039   0.09162   0.08712  -0.0118   0.0890   1.0000
  11.750   0.8757   0.09931   0.09484  -0.0156   0.0883   1.0000
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