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NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA M2 AIRFOIL (m2-il)
Reynolds number: 50,000
Max Cl/Cd: 21.59 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m2-il-50000-n5.txt
Download as CSV file: xf-m2-il-50000-n5.csv

 
  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M2 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.7111   0.09164   0.08465  -0.0031   1.0000   0.0559
  -9.250  -0.7256   0.08452   0.07754  -0.0087   1.0000   0.0549
  -9.000  -0.7416   0.07871   0.07167  -0.0114   1.0000   0.0542
  -8.750  -0.7501   0.07370   0.06654  -0.0129   1.0000   0.0546
  -8.500  -0.7547   0.06900   0.06166  -0.0137   1.0000   0.0554
  -8.250  -0.7566   0.06431   0.05672  -0.0141   1.0000   0.0562
  -8.000  -0.7555   0.05958   0.05165  -0.0141   1.0000   0.0569
  -7.750  -0.7508   0.05493   0.04660  -0.0137   1.0000   0.0575
  -7.500  -0.7424   0.05061   0.04180  -0.0131   1.0000   0.0590
  -7.250  -0.7319   0.04644   0.03692  -0.0121   1.0000   0.0620
  -7.000  -0.7161   0.04336   0.03351  -0.0114   1.0000   0.0649
  -6.750  -0.6972   0.04065   0.03045  -0.0107   1.0000   0.0671
  -6.500  -0.6769   0.03804   0.02743  -0.0099   1.0000   0.0703
  -6.250  -0.6553   0.03560   0.02428  -0.0089   1.0000   0.0757
  -6.000  -0.6324   0.03354   0.02213  -0.0084   1.0000   0.0790
  -5.750  -0.6079   0.03156   0.01988  -0.0078   1.0000   0.0823
  -5.500  -0.5821   0.02970   0.01768  -0.0071   1.0000   0.0864
  -5.250  -0.5565   0.02811   0.01589  -0.0064   1.0000   0.0916
  -5.000  -0.5316   0.02699   0.01460  -0.0058   1.0000   0.1013
  -4.750  -0.5065   0.02564   0.01325  -0.0051   1.0000   0.1097
  -4.500  -0.4813   0.02436   0.01185  -0.0043   1.0000   0.1202
  -4.250  -0.4574   0.02316   0.01058  -0.0035   1.0000   0.1349
  -4.000  -0.4344   0.02206   0.00949  -0.0028   1.0000   0.1540
  -3.750  -0.4118   0.02101   0.00856  -0.0022   1.0000   0.1867
  -3.500  -0.3920   0.01963   0.00762  -0.0015   1.0000   0.2414
  -3.250  -0.3796   0.01762   0.00712   0.0010   1.0000   0.4732
  -3.000  -0.3638   0.01685   0.00719   0.0051   1.0000   0.6700
  -2.750  -0.3393   0.01672   0.00724   0.0078   1.0000   0.7763
  -2.500  -0.3042   0.01683   0.00726   0.0082   1.0000   0.8462
  -2.250  -0.2662   0.01695   0.00718   0.0073   1.0000   0.8949
  -2.000  -0.2200   0.01705   0.00706   0.0044   1.0000   0.9407
  -1.750  -0.1459   0.01704   0.00673  -0.0045   1.0000   0.9794
  -1.500  -0.0903   0.01677   0.00628  -0.0107   1.0000   1.0000
  -1.250  -0.0740   0.01653   0.00596  -0.0093   1.0000   1.0000
  -1.000  -0.0582   0.01635   0.00572  -0.0078   1.0000   1.0000
  -0.750  -0.0429   0.01620   0.00554  -0.0061   1.0000   1.0000
  -0.500  -0.0283   0.01611   0.00542  -0.0041   1.0000   1.0000
  -0.250  -0.0141   0.01605   0.00534  -0.0021   1.0000   1.0000
   0.000   0.0000   0.01603   0.00532   0.0000   1.0000   1.0000
   0.250   0.0141   0.01605   0.00534   0.0021   1.0000   1.0000
   0.500   0.0283   0.01611   0.00542   0.0041   1.0000   1.0000
   0.750   0.0430   0.01620   0.00554   0.0061   1.0000   1.0000
   1.000   0.0582   0.01634   0.00572   0.0078   1.0000   1.0000
   1.250   0.0740   0.01653   0.00596   0.0093   1.0000   1.0000
   1.500   0.0903   0.01677   0.00628   0.0107   1.0000   1.0000
   1.750   0.1459   0.01704   0.00673   0.0045   0.9794   1.0000
   2.000   0.2199   0.01705   0.00705  -0.0044   0.9408   1.0000
   2.250   0.2662   0.01694   0.00718  -0.0073   0.8949   1.0000
   2.500   0.3042   0.01683   0.00726  -0.0082   0.8462   1.0000
   2.750   0.3393   0.01672   0.00723  -0.0078   0.7764   1.0000
   3.000   0.3638   0.01685   0.00719  -0.0051   0.6701   1.0000
   3.250   0.3796   0.01762   0.00712  -0.0010   0.4734   1.0000
   3.500   0.3920   0.01962   0.00761   0.0015   0.2414   1.0000
   3.750   0.4118   0.02101   0.00855   0.0022   0.1866   1.0000
   4.000   0.4344   0.02206   0.00949   0.0028   0.1539   1.0000
   4.250   0.4574   0.02316   0.01058   0.0035   0.1349   1.0000
   4.500   0.4813   0.02436   0.01185   0.0043   0.1202   1.0000
   4.750   0.5065   0.02564   0.01325   0.0051   0.1097   1.0000
   5.000   0.5317   0.02699   0.01460   0.0058   0.1013   1.0000
   5.250   0.5565   0.02811   0.01588   0.0064   0.0916   1.0000
   5.500   0.5821   0.02970   0.01768   0.0071   0.0864   1.0000
   5.750   0.6079   0.03156   0.01988   0.0078   0.0823   1.0000
   6.000   0.6324   0.03354   0.02213   0.0084   0.0790   1.0000
   6.250   0.6553   0.03560   0.02429   0.0089   0.0757   1.0000
   6.500   0.6769   0.03804   0.02743   0.0099   0.0703   1.0000
   6.750   0.6972   0.04066   0.03046   0.0107   0.0671   1.0000
   7.000   0.7162   0.04337   0.03352   0.0114   0.0649   1.0000
   7.250   0.7320   0.04644   0.03693   0.0121   0.0620   1.0000
   7.500   0.7425   0.05062   0.04181   0.0130   0.0590   1.0000
   7.750   0.7509   0.05494   0.04661   0.0137   0.0575   1.0000
   8.000   0.7556   0.05959   0.05167   0.0141   0.0569   1.0000
   8.250   0.7568   0.06433   0.05674   0.0141   0.0562   1.0000
   8.500   0.7550   0.06902   0.06168   0.0137   0.0554   1.0000
   8.750   0.7505   0.07368   0.06652   0.0129   0.0545   1.0000
   9.000   0.7423   0.07868   0.07164   0.0114   0.0542   1.0000
   9.250   0.7261   0.08454   0.07757   0.0086   0.0548   1.0000
   9.500   0.7117   0.09166   0.08467   0.0030   0.0558   1.0000
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