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NACA M3 AIRFOIL (m3-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA M3 AIRFOIL (m3-il)
Reynolds number: 50,000
Max Cl/Cd: 27.33 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m3-il-50000.txt
Download as CSV file: xf-m3-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M3 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.7223   0.08756   0.08003  -0.0204   1.0000   0.1683
  -9.750  -0.7340   0.08121   0.07365  -0.0224   1.0000   0.1633
  -9.500  -0.8167   0.07118   0.06331  -0.0239   1.0000   0.1508
  -9.250  -0.8235   0.06643   0.05839  -0.0229   1.0000   0.1496
  -9.000  -0.8316   0.06182   0.05353  -0.0215   1.0000   0.1487
  -8.750  -0.8370   0.05748   0.04886  -0.0197   1.0000   0.1482
  -8.500  -0.8382   0.05342   0.04441  -0.0177   1.0000   0.1481
  -8.250  -0.8346   0.04967   0.04024  -0.0156   1.0000   0.1483
  -8.000  -0.8266   0.04621   0.03634  -0.0136   1.0000   0.1487
  -7.750  -0.8160   0.04309   0.03274  -0.0115   1.0000   0.1501
  -7.500  -0.8017   0.04031   0.02961  -0.0098   1.0000   0.1535
  -7.250  -0.7812   0.03818   0.02744  -0.0087   1.0000   0.1594
  -7.000  -0.7632   0.03591   0.02477  -0.0071   1.0000   0.1650
  -6.750  -0.7409   0.03366   0.02242  -0.0060   1.0000   0.1714
  -6.500  -0.7201   0.03188   0.02037  -0.0046   1.0000   0.1824
  -6.250  -0.6973   0.03010   0.01862  -0.0035   1.0000   0.1980
  -6.000  -0.6734   0.02828   0.01708  -0.0025   1.0000   0.2198
  -5.750  -0.6539   0.02660   0.01563  -0.0007   1.0000   0.2573
  -5.500  -0.6382   0.02490   0.01447   0.0018   1.0000   0.3157
  -5.250  -0.6272   0.02349   0.01369   0.0053   1.0000   0.3967
  -5.000  -0.6178   0.02260   0.01341   0.0096   1.0000   0.4863
  -4.750  -0.6054   0.02227   0.01355   0.0143   1.0000   0.5676
  -4.500  -0.5887   0.02238   0.01393   0.0188   1.0000   0.6345
  -4.250  -0.5687   0.02280   0.01445   0.0232   1.0000   0.6905
  -4.000  -0.5447   0.02351   0.01516   0.0273   1.0000   0.7398
  -3.750  -0.5109   0.02454   0.01609   0.0300   1.0000   0.7842
  -3.500  -0.4689   0.02549   0.01682   0.0306   1.0000   0.8260
  -3.250  -0.3880   0.02679   0.01773   0.0244   1.0000   0.8632
  -3.000  -0.2913   0.02731   0.01782   0.0133   1.0000   0.8979
  -2.750  -0.2139   0.02701   0.01721   0.0041   1.0000   0.9292
  -2.500  -0.1425   0.02628   0.01624  -0.0050   1.0000   0.9579
  -2.250  -0.0683   0.02519   0.01498  -0.0153   1.0000   0.9841
  -2.000  -0.0118   0.02404   0.01371  -0.0229   1.0000   1.0000
  -1.750  -0.0031   0.02344   0.01310  -0.0217   1.0000   1.0000
  -1.500   0.0039   0.02293   0.01260  -0.0199   1.0000   1.0000
  -1.250   0.0087   0.02252   0.01221  -0.0177   1.0000   1.0000
  -1.000   0.0112   0.02220   0.01191  -0.0150   1.0000   1.0000
  -0.750   0.0115   0.02197   0.01169  -0.0118   1.0000   1.0000
  -0.500   0.0094   0.02181   0.01154  -0.0082   1.0000   1.0000
  -0.250   0.0052   0.02172   0.01146  -0.0042   1.0000   1.0000
   0.000   0.0000   0.02169   0.01144   0.0000   1.0000   1.0000
   0.250  -0.0052   0.02172   0.01146   0.0042   1.0000   1.0000
   0.500  -0.0094   0.02181   0.01154   0.0082   1.0000   1.0000
   0.750  -0.0115   0.02196   0.01169   0.0118   1.0000   1.0000
   1.000  -0.0112   0.02220   0.01190   0.0150   1.0000   1.0000
   1.250  -0.0087   0.02252   0.01220   0.0177   1.0000   1.0000
   1.500  -0.0039   0.02293   0.01259   0.0199   1.0000   1.0000
   1.750   0.0031   0.02343   0.01309   0.0217   1.0000   1.0000
   2.000   0.0119   0.02404   0.01370   0.0229   1.0000   1.0000
   2.250   0.0682   0.02518   0.01496   0.0153   0.9841   1.0000
   2.500   0.1424   0.02627   0.01623   0.0050   0.9580   1.0000
   2.750   0.2138   0.02700   0.01720  -0.0041   0.9292   1.0000
   3.000   0.2912   0.02731   0.01781  -0.0133   0.8980   1.0000
   3.250   0.3878   0.02678   0.01773  -0.0243   0.8632   1.0000
   3.500   0.4690   0.02548   0.01681  -0.0307   0.8261   1.0000
   3.750   0.5108   0.02454   0.01609  -0.0300   0.7843   1.0000
   4.000   0.5446   0.02351   0.01516  -0.0273   0.7399   1.0000
   4.250   0.5687   0.02280   0.01445  -0.0232   0.6906   1.0000
   4.500   0.5886   0.02238   0.01393  -0.0188   0.6346   1.0000
   4.750   0.6053   0.02227   0.01355  -0.0143   0.5677   1.0000
   5.000   0.6177   0.02260   0.01341  -0.0096   0.4864   1.0000
   5.250   0.6272   0.02348   0.01369  -0.0053   0.3968   1.0000
   5.500   0.6381   0.02490   0.01447  -0.0018   0.3159   1.0000
   5.750   0.6538   0.02659   0.01563   0.0007   0.2574   1.0000
   6.000   0.6734   0.02827   0.01708   0.0025   0.2198   1.0000
   6.250   0.6972   0.03010   0.01862   0.0035   0.1980   1.0000
   6.500   0.7200   0.03187   0.02036   0.0046   0.1824   1.0000
   6.750   0.7409   0.03366   0.02242   0.0060   0.1714   1.0000
   7.000   0.7632   0.03591   0.02477   0.0071   0.1650   1.0000
   7.250   0.7812   0.03817   0.02744   0.0087   0.1594   1.0000
   7.500   0.8017   0.04031   0.02960   0.0098   0.1535   1.0000
   7.750   0.8160   0.04309   0.03274   0.0115   0.1501   1.0000
   8.000   0.8266   0.04621   0.03634   0.0136   0.1487   1.0000
   8.250   0.8346   0.04967   0.04024   0.0156   0.1483   1.0000
   8.500   0.8382   0.05342   0.04442   0.0176   0.1481   1.0000
   8.750   0.8371   0.05748   0.04886   0.0196   0.1482   1.0000
   9.000   0.8317   0.06183   0.05354   0.0214   0.1487   1.0000
   9.250   0.8236   0.06644   0.05841   0.0229   0.1496   1.0000
   9.500   0.8169   0.07120   0.06332   0.0238   0.1508   1.0000
   9.750   0.7344   0.08125   0.07369   0.0223   0.1633   1.0000
  10.000   0.7228   0.08761   0.08008   0.0203   0.1683   1.0000
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