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NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA M13 AIRFOIL (m13-il)
Reynolds number: 50,000
Max Cl/Cd: 40.57 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m13-il-50000-n5.txt
Download as CSV file: xf-m13-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4599   0.12111   0.11441   0.0016   1.0000   0.0701
  -8.750  -0.4601   0.11944   0.11282  -0.0018   1.0000   0.0712
  -8.500  -0.4612   0.11780   0.11127  -0.0056   1.0000   0.0717
  -8.250  -0.4468   0.11111   0.10459  -0.0038   1.0000   0.0735
  -8.000  -0.4383   0.10708   0.10059  -0.0041   1.0000   0.0759
  -7.750  -0.4330   0.10383   0.09735  -0.0056   1.0000   0.0784
  -7.500  -0.4294   0.10095   0.09453  -0.0081   1.0000   0.0807
  -7.250  -0.4236   0.09849   0.09212  -0.0136   1.0000   0.0828
  -7.000  -0.4141   0.09630   0.08993  -0.0209   1.0000   0.0837
  -6.750  -0.4042   0.09109   0.08479  -0.0207   1.0000   0.0848
  -6.500  -0.3943   0.08649   0.08023  -0.0186   1.0000   0.0885
  -6.250  -0.3819   0.08317   0.07690  -0.0219   1.0000   0.0936
  -5.750  -0.3547   0.07568   0.06941  -0.0279   1.0000   0.1001
  -5.500  -0.3403   0.07202   0.06574  -0.0292   1.0000   0.1035
  -5.250  -0.3146   0.06985   0.06329  -0.0366   1.0000   0.1109
  -5.000  -0.3028   0.06514   0.05867  -0.0361   1.0000   0.1130
  -4.750  -0.2810   0.06297   0.05629  -0.0395   1.0000   0.1260
  -4.500  -0.2738   0.05886   0.05233  -0.0369   1.0000   0.1364
  -4.250  -0.2595   0.05584   0.04928  -0.0373   1.0000   0.1486
  -3.750  -0.1957   0.04687   0.03927  -0.0432   1.0000   0.0713
  -3.500  -0.1843   0.04442   0.03684  -0.0423   1.0000   0.0749
  -3.250  -0.1691   0.04238   0.03462  -0.0417   1.0000   0.0783
  -3.000  -0.1332   0.03900   0.03076  -0.0447   0.9931   0.0750
  -2.750  -0.0866   0.03563   0.02658  -0.0487   0.9823   0.0714
  -2.500  -0.0450   0.03285   0.02327  -0.0520   0.9713   0.0711
  -2.250  -0.0052   0.03084   0.02103  -0.0553   0.9601   0.0768
  -2.000   0.0380   0.02888   0.01855  -0.0585   0.9499   0.0794
  -1.750   0.0791   0.02707   0.01625  -0.0609   0.9386   0.0792
  -1.500   0.1184   0.02560   0.01437  -0.0629   0.9264   0.0794
  -1.250   0.1561   0.02438   0.01279  -0.0645   0.9140   0.0803
  -1.000   0.1931   0.02339   0.01146  -0.0657   0.9015   0.0818
  -0.750   0.2291   0.02258   0.01033  -0.0667   0.8888   0.0842
  -0.500   0.2614   0.02184   0.00950  -0.0673   0.8756   0.0882
  -0.250   0.2915   0.02130   0.00881  -0.0674   0.8622   0.0948
   0.000   0.3198   0.02082   0.00828  -0.0672   0.8486   0.1059
   0.250   0.3473   0.02044   0.00795  -0.0670   0.8352   0.1387
   0.500   0.3734   0.01963   0.00774  -0.0669   0.8224   0.2551
   0.750   0.4100   0.01800   0.00741  -0.0678   0.8104   1.0000
   1.000   0.4358   0.01823   0.00737  -0.0672   0.7971   1.0000
   1.250   0.4613   0.01847   0.00741  -0.0665   0.7838   1.0000
   1.500   0.4866   0.01872   0.00751  -0.0658   0.7708   1.0000
   1.750   0.5120   0.01898   0.00767  -0.0652   0.7582   1.0000
   2.000   0.5373   0.01926   0.00785  -0.0646   0.7459   1.0000
   2.250   0.5627   0.01953   0.00806  -0.0639   0.7340   1.0000
   2.500   0.5882   0.01981   0.00831  -0.0633   0.7228   1.0000
   2.750   0.6135   0.02012   0.00861  -0.0626   0.7109   1.0000
   3.000   0.6389   0.02047   0.00899  -0.0622   0.6986   1.0000
   3.250   0.6643   0.02083   0.00939  -0.0617   0.6867   1.0000
   3.500   0.6898   0.02119   0.00984  -0.0612   0.6751   1.0000
   3.750   0.7154   0.02153   0.01024  -0.0606   0.6642   1.0000
   4.000   0.7408   0.02190   0.01071  -0.0600   0.6527   1.0000
   4.250   0.7658   0.02233   0.01129  -0.0595   0.6403   1.0000
   4.500   0.7907   0.02277   0.01193  -0.0590   0.6281   1.0000
   4.750   0.8156   0.02321   0.01254  -0.0584   0.6160   1.0000
   5.000   0.8405   0.02362   0.01316  -0.0577   0.6041   1.0000
   5.250   0.8657   0.02401   0.01375  -0.0570   0.5924   1.0000
   5.500   0.8905   0.02445   0.01450  -0.0562   0.5801   1.0000
   5.750   0.9148   0.02497   0.01535  -0.0555   0.5670   1.0000
   6.000   0.9392   0.02549   0.01624  -0.0548   0.5538   1.0000
   6.250   0.9558   0.02410   0.01478  -0.0500   0.4939   1.0000
   6.500   0.9667   0.02383   0.01421  -0.0458   0.3885   1.0000
   6.750   0.9643   0.02679   0.01532  -0.0424   0.1289   1.0000
   7.000   0.9689   0.03018   0.01802  -0.0406   0.0698   1.0000
   7.250   0.9800   0.03236   0.02030  -0.0389   0.0585   1.0000
   7.500   0.9894   0.03454   0.02260  -0.0370   0.0522   1.0000
   7.750   0.9994   0.03650   0.02474  -0.0352   0.0463   1.0000
   8.000   1.0060   0.03884   0.02710  -0.0332   0.0427   1.0000
   8.250   1.0205   0.04066   0.02919  -0.0311   0.0406   1.0000
   8.500   1.0402   0.04255   0.03130  -0.0293   0.0382   1.0000
   8.750   1.0607   0.04465   0.03361  -0.0279   0.0349   1.0000
   9.000   1.0853   0.04748   0.03653  -0.0270   0.0325   1.0000
   9.250   1.1084   0.05047   0.03993  -0.0259   0.0316   1.0000
   9.500   1.1256   0.05384   0.04374  -0.0246   0.0311   1.0000
   9.750   1.1368   0.05743   0.04777  -0.0232   0.0308   1.0000
  10.000   1.1416   0.06109   0.05186  -0.0216   0.0304   1.0000
  10.250   1.1407   0.06477   0.05595  -0.0199   0.0300   1.0000
  10.500   1.1342   0.06841   0.05994  -0.0182   0.0297   1.0000
  10.750   1.1227   0.07203   0.06386  -0.0165   0.0296   1.0000
  11.000   1.1089   0.07601   0.06811  -0.0158   0.0295   1.0000
  11.250   1.0936   0.08047   0.07280  -0.0162   0.0295   1.0000
  11.500   1.0770   0.08547   0.07802  -0.0176   0.0297   1.0000
  11.750   1.0597   0.09104   0.08377  -0.0199   0.0299   1.0000
  12.000   1.0421   0.09718   0.09008  -0.0230   0.0303   1.0000
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