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NACA M13 AIRFOIL (m13-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M13 AIRFOIL (m13-il)
Reynolds number: 200,000
Max Cl/Cd: 74.13 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m13-il-200000.txt
Download as CSV file: xf-m13-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4598   0.10480   0.10143   0.0032   1.0000   0.0279
  -8.000  -0.4568   0.10270   0.09938  -0.0015   1.0000   0.0286
  -7.750  -0.4523   0.10003   0.09675  -0.0071   1.0000   0.0289
  -7.500  -0.4405   0.09648   0.09320  -0.0127   1.0000   0.0290
  -7.250  -0.4271   0.09261   0.08932  -0.0172   1.0000   0.0291
  -7.000  -0.4127   0.08851   0.08523  -0.0211   1.0000   0.0292
  -6.750  -0.4134   0.08200   0.07874  -0.0185   1.0000   0.0300
  -6.500  -0.4036   0.07829   0.07505  -0.0182   1.0000   0.0308
  -6.250  -0.3905   0.07474   0.07149  -0.0200   1.0000   0.0318
  -6.000  -0.3749   0.07112   0.06786  -0.0228   1.0000   0.0331
  -5.750  -0.3569   0.06739   0.06409  -0.0262   1.0000   0.0348
  -5.500  -0.3338   0.06350   0.06013  -0.0307   1.0000   0.0374
  -5.250  -0.3005   0.05851   0.05491  -0.0383   1.0000   0.0399
  -5.000  -0.2934   0.05555   0.05206  -0.0365   1.0000   0.0418
  -4.750  -0.2761   0.05278   0.04926  -0.0373   1.0000   0.0445
  -4.500  -0.2400   0.05026   0.04627  -0.0417   1.0000   0.0502
  -4.250  -0.2358   0.04638   0.04256  -0.0405   0.9978   0.0515
  -4.000  -0.2000   0.04317   0.03929  -0.0445   0.9902   0.0563
  -3.750  -0.1532   0.03896   0.03476  -0.0504   0.9824   0.0636
  -3.500  -0.1042   0.03724   0.03242  -0.0547   0.9707   0.0736
  -3.250  -0.0752   0.03269   0.02801  -0.0575   0.9579   0.0761
  -3.000  -0.0399   0.03040   0.02549  -0.0596   0.9431   0.0824
  -2.750  -0.0084   0.02781   0.02259  -0.0607   0.9272   0.0890
  -2.500   0.0182   0.02624   0.02082  -0.0604   0.9094   0.0952
  -2.250   0.0440   0.02436   0.01860  -0.0599   0.8922   0.1021
  -2.000   0.0700   0.02345   0.01733  -0.0590   0.8760   0.1141
  -1.750   0.0935   0.02167   0.01550  -0.0584   0.8607   0.1185
  -1.500   0.1269   0.01863   0.01166  -0.0569   0.8479   0.0772
  -1.250   0.1563   0.01655   0.00889  -0.0555   0.8351   0.0620
  -1.000   0.1831   0.01580   0.00792  -0.0548   0.8218   0.0608
  -0.750   0.2096   0.01484   0.00678  -0.0541   0.8093   0.0607
  -0.500   0.2362   0.01419   0.00596  -0.0535   0.7972   0.0611
  -0.250   0.2624   0.01335   0.00511  -0.0530   0.7850   0.0628
   0.000   0.2890   0.01298   0.00473  -0.0526   0.7726   0.0661
   0.250   0.3158   0.01275   0.00445  -0.0522   0.7606   0.0711
   0.500   0.3424   0.01237   0.00404  -0.0517   0.7491   0.0744
   0.750   0.3690   0.01204   0.00372  -0.0514   0.7380   0.0826
   1.000   0.3956   0.01171   0.00340  -0.0510   0.7272   0.1067
   1.250   0.4226   0.01149   0.00335  -0.0508   0.7152   0.1712
   1.500   0.4552   0.00959   0.00334  -0.0517   0.7037   1.0000
   1.750   0.4819   0.00973   0.00334  -0.0514   0.6924   1.0000
   2.000   0.5085   0.00987   0.00337  -0.0510   0.6814   1.0000
   2.250   0.5352   0.01002   0.00342  -0.0507   0.6704   1.0000
   2.500   0.5621   0.01016   0.00350  -0.0505   0.6583   1.0000
   2.750   0.5890   0.01032   0.00362  -0.0502   0.6466   1.0000
   3.000   0.6159   0.01047   0.00374  -0.0500   0.6351   1.0000
   3.250   0.6427   0.01064   0.00386  -0.0497   0.6238   1.0000
   3.500   0.6695   0.01081   0.00397  -0.0495   0.6127   1.0000
   3.750   0.6965   0.01096   0.00417  -0.0493   0.6007   1.0000
   4.000   0.7234   0.01114   0.00438  -0.0491   0.5891   1.0000
   4.250   0.7504   0.01133   0.00459  -0.0489   0.5781   1.0000
   4.500   0.7773   0.01153   0.00482  -0.0487   0.5676   1.0000
   4.750   0.8038   0.01165   0.00501  -0.0484   0.5522   1.0000
   5.000   0.8289   0.01160   0.00493  -0.0476   0.5232   1.0000
   5.250   0.8542   0.01165   0.00496  -0.0470   0.4913   1.0000
   5.500   0.8784   0.01185   0.00503  -0.0462   0.4441   1.0000
   5.750   0.8991   0.01263   0.00533  -0.0452   0.3169   1.0000
   6.000   0.9067   0.01636   0.00737  -0.0435   0.0503   1.0000
   6.250   0.9290   0.01739   0.00853  -0.0427   0.0411   1.0000
   6.500   0.9472   0.01901   0.01025  -0.0415   0.0367   1.0000
   6.750   0.9669   0.02030   0.01164  -0.0403   0.0350   1.0000
   7.000   0.9874   0.02148   0.01289  -0.0392   0.0324   1.0000
   7.250   1.0073   0.02279   0.01428  -0.0380   0.0299   1.0000
   7.500   1.0271   0.02444   0.01597  -0.0368   0.0288   1.0000
   7.750   1.0484   0.02635   0.01793  -0.0356   0.0282   1.0000
   8.000   1.0717   0.02851   0.02022  -0.0346   0.0281   1.0000
   8.250   1.0962   0.03103   0.02300  -0.0335   0.0289   1.0000
   8.500   1.1174   0.03351   0.02566  -0.0326   0.0279   1.0000
   8.750   1.1372   0.03655   0.02897  -0.0317   0.0273   1.0000
   9.000   1.1533   0.04277   0.03600  -0.0289   0.0345   1.0000
  15.000   0.7372   0.16010   0.15684  -0.0432   0.0497   1.0000
  15.250   0.7446   0.16394   0.16068  -0.0415   0.0479   1.0000
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