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NACA M2 AIRFOIL (m2-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA M2 AIRFOIL (m2-il)
Reynolds number: 100,000
Max Cl/Cd: 29.37 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m2-il-100000.txt
Download as CSV file: xf-m2-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.6927   0.09590   0.09101   0.0072   1.0000   0.1207
  -8.750  -0.7231   0.08926   0.08443  -0.0033   1.0000   0.1235
  -8.500  -0.7655   0.08552   0.08030  -0.0103   1.0000   0.1248
  -8.250  -0.7233   0.07912   0.07427  -0.0056   1.0000   0.1284
  -8.000  -0.7188   0.07514   0.07025  -0.0061   1.0000   0.1329
  -7.750  -0.7421   0.06995   0.06463  -0.0111   1.0000   0.1403
  -7.500  -0.7183   0.06613   0.06102  -0.0088   1.0000   0.1465
  -7.250  -0.7183   0.06169   0.05638  -0.0100   1.0000   0.1574
  -7.000  -0.7130   0.05807   0.05254  -0.0103   1.0000   0.1708
  -6.750  -0.6962   0.04594   0.03899  -0.0117   1.0000   0.0938
  -6.500  -0.6807   0.03962   0.03175  -0.0101   1.0000   0.0809
  -6.250  -0.6609   0.03687   0.02837  -0.0088   1.0000   0.0819
  -6.000  -0.6411   0.03288   0.02410  -0.0082   1.0000   0.0834
  -5.750  -0.6180   0.03006   0.02093  -0.0074   1.0000   0.0839
  -5.500  -0.5941   0.02809   0.01884  -0.0068   1.0000   0.0869
  -5.250  -0.5695   0.02649   0.01697  -0.0060   1.0000   0.0917
  -5.000  -0.5438   0.02473   0.01480  -0.0051   1.0000   0.0943
  -4.750  -0.5182   0.02283   0.01272  -0.0045   1.0000   0.0976
  -4.500  -0.4933   0.02162   0.01152  -0.0039   1.0000   0.1050
  -4.250  -0.4678   0.02025   0.01005  -0.0031   1.0000   0.1125
  -4.000  -0.4430   0.01901   0.00882  -0.0023   1.0000   0.1229
  -3.750  -0.4200   0.01753   0.00753  -0.0013   1.0000   0.1414
  -3.500  -0.3992   0.01617   0.00646  -0.0001   1.0000   0.1777
  -3.250  -0.3781   0.01505   0.00558   0.0010   1.0000   0.2167
  -3.000  -0.3596   0.01338   0.00489   0.0022   1.0000   0.3544
  -2.750  -0.3492   0.01189   0.00509   0.0071   1.0000   0.7175
  -2.500  -0.3290   0.01200   0.00557   0.0117   1.0000   0.8506
  -2.250  -0.2946   0.01246   0.00592   0.0126   1.0000   0.9182
  -2.000  -0.2399   0.01279   0.00600   0.0085   1.0000   0.9571
  -1.750  -0.1618   0.01292   0.00586  -0.0012   1.0000   0.9784
  -1.500  -0.0824   0.01280   0.00552  -0.0118   1.0000   1.0000
  -1.250  -0.0657   0.01249   0.00517  -0.0107   1.0000   1.0000
  -1.000  -0.0499   0.01225   0.00490  -0.0093   1.0000   1.0000
  -0.750  -0.0355   0.01208   0.00470  -0.0075   1.0000   1.0000
  -0.500  -0.0228   0.01196   0.00456  -0.0052   1.0000   1.0000
  -0.250  -0.0114   0.01189   0.00448  -0.0026   1.0000   1.0000
   0.000   0.0000   0.01186   0.00446   0.0000   1.0000   1.0000
   0.250   0.0114   0.01189   0.00448   0.0026   1.0000   1.0000
   0.500   0.0228   0.01196   0.00456   0.0052   1.0000   1.0000
   0.750   0.0355   0.01208   0.00470   0.0075   1.0000   1.0000
   1.000   0.0499   0.01225   0.00490   0.0093   1.0000   1.0000
   1.250   0.0657   0.01249   0.00517   0.0107   1.0000   1.0000
   1.500   0.0825   0.01280   0.00552   0.0118   1.0000   1.0000
   1.750   0.1618   0.01292   0.00586   0.0012   0.9784   1.0000
   2.000   0.2399   0.01279   0.00600  -0.0085   0.9571   1.0000
   2.250   0.2945   0.01246   0.00592  -0.0126   0.9183   1.0000
   2.500   0.3290   0.01200   0.00557  -0.0117   0.8507   1.0000
   2.750   0.3492   0.01189   0.00509  -0.0071   0.7178   1.0000
   3.000   0.3596   0.01337   0.00489  -0.0022   0.3550   1.0000
   3.250   0.3780   0.01505   0.00557  -0.0010   0.2168   1.0000
   3.500   0.3992   0.01617   0.00646   0.0001   0.1777   1.0000
   3.750   0.4200   0.01753   0.00753   0.0013   0.1414   1.0000
   4.000   0.4430   0.01901   0.00882   0.0023   0.1229   1.0000
   4.250   0.4678   0.02024   0.01005   0.0031   0.1125   1.0000
   4.500   0.4933   0.02162   0.01152   0.0039   0.1050   1.0000
   4.750   0.5182   0.02283   0.01272   0.0045   0.0976   1.0000
   5.000   0.5438   0.02473   0.01479   0.0051   0.0943   1.0000
   5.250   0.5695   0.02649   0.01696   0.0060   0.0917   1.0000
   5.500   0.5941   0.02809   0.01885   0.0068   0.0869   1.0000
   5.750   0.6180   0.03006   0.02093   0.0074   0.0839   1.0000
   6.000   0.6412   0.03287   0.02410   0.0082   0.0834   1.0000
   6.250   0.6609   0.03687   0.02837   0.0088   0.0819   1.0000
   6.500   0.6807   0.03962   0.03175   0.0101   0.0809   1.0000
   7.250   0.7184   0.06170   0.05639   0.0100   0.1574   1.0000
   7.500   0.7184   0.06613   0.06102   0.0088   0.1465   1.0000
   7.750   0.7423   0.06995   0.06463   0.0111   0.1403   1.0000
   8.000   0.7190   0.07515   0.07026   0.0061   0.1329   1.0000
   8.250   0.7237   0.07913   0.07428   0.0056   0.1284   1.0000
   8.500   0.7655   0.08553   0.08030   0.0103   0.1248   1.0000
   8.750   0.7231   0.08930   0.08447   0.0032   0.1235   1.0000
   9.000   0.6932   0.09594   0.09106  -0.0073   0.1206   1.0000
   9.250   0.6879   0.10072   0.09579  -0.0109   0.1167   1.0000
   9.500   0.6938   0.10454   0.09961  -0.0110   0.1134   1.0000
   9.750   0.5644   0.10157   0.09678  -0.0019   0.1227   1.0000
  10.000   0.5509   0.10609   0.10126  -0.0048   0.1186   1.0000
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