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NACA-M1 AIRFOIL (m1-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: NACA-M1 AIRFOIL (m1-il)
Reynolds number: 500,000
Max Cl/Cd: 50.55 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m1-il-500000-n5.txt
Download as CSV file: xf-m1-il-500000-n5.csv

 
  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA-M1 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.7796   0.08053   0.07847   0.0112   1.0000   0.0050
  -9.000  -0.7938   0.07176   0.06969   0.0017   1.0000   0.0049
  -8.750  -0.8077   0.06146   0.05921  -0.0047   1.0000   0.0048
  -8.500  -0.8768   0.03035   0.02663  -0.0067   1.0000   0.0046
  -8.250  -0.8668   0.02480   0.02033  -0.0054   1.0000   0.0049
  -8.000  -0.8480   0.02198   0.01705  -0.0045   1.0000   0.0051
  -7.750  -0.8265   0.02002   0.01469  -0.0039   1.0000   0.0054
  -7.500  -0.8033   0.01855   0.01295  -0.0033   1.0000   0.0056
  -7.250  -0.7807   0.01687   0.01102  -0.0027   1.0000   0.0061
  -7.000  -0.7558   0.01610   0.01013  -0.0025   1.0000   0.0066
  -6.750  -0.7302   0.01551   0.00946  -0.0023   1.0000   0.0074
  -6.500  -0.7049   0.01473   0.00854  -0.0019   1.0000   0.0081
  -6.250  -0.6795   0.01393   0.00760  -0.0016   1.0000   0.0087
  -6.000  -0.6545   0.01295   0.00650  -0.0011   1.0000   0.0101
  -5.750  -0.6282   0.01253   0.00605  -0.0010   1.0000   0.0117
  -5.500  -0.6015   0.01225   0.00574  -0.0008   1.0000   0.0137
  -5.250  -0.5750   0.01181   0.00530  -0.0007   1.0000   0.0170
  -5.000  -0.5480   0.01161   0.00507  -0.0006   1.0000   0.0210
  -4.750  -0.5207   0.01153   0.00494  -0.0005   1.0000   0.0243
  -4.500  -0.4933   0.01149   0.00485  -0.0004   1.0000   0.0255
  -4.250  -0.4673   0.01095   0.00425  -0.0002   1.0000   0.0276
  -4.000  -0.4409   0.01053   0.00381   0.0000   1.0000   0.0298
  -3.750  -0.4143   0.01024   0.00348   0.0001   1.0000   0.0315
  -3.500  -0.3876   0.00995   0.00315   0.0003   1.0000   0.0327
  -3.250  -0.3610   0.00968   0.00284   0.0005   1.0000   0.0336
  -3.000  -0.3343   0.00941   0.00253   0.0007   1.0000   0.0339
  -2.750  -0.3078   0.00917   0.00225   0.0010   1.0000   0.0342
  -2.500  -0.2813   0.00897   0.00201   0.0013   1.0000   0.0348
  -2.250  -0.2550   0.00880   0.00182   0.0016   1.0000   0.0358
  -2.000  -0.2290   0.00864   0.00165   0.0020   1.0000   0.0380
  -1.750  -0.2033   0.00843   0.00151   0.0024   1.0000   0.0494
  -1.500  -0.1784   0.00797   0.00138   0.0027   1.0000   0.1291
  -1.250  -0.1493   0.00745   0.00126   0.0020   0.9976   0.2367
  -1.000  -0.1161   0.00685   0.00116   0.0004   0.9922   0.3717
  -0.750  -0.0831   0.00627   0.00109  -0.0012   0.9842   0.5097
  -0.500  -0.0471   0.00572   0.00105  -0.0031   0.9616   0.6464
  -0.250  -0.0190   0.00532   0.00105  -0.0028   0.9125   0.7648
   0.000   0.0000   0.00517   0.00106   0.0000   0.8526   0.8537
   0.250   0.0189   0.00531   0.00105   0.0028   0.7647   0.9142
   0.500   0.0473   0.00572   0.00105   0.0031   0.6460   0.9624
   0.750   0.0831   0.00627   0.00109   0.0012   0.5089   0.9842
   1.000   0.1161   0.00686   0.00116  -0.0004   0.3710   0.9921
   1.250   0.1493   0.00746   0.00126  -0.0020   0.2341   0.9976
   1.500   0.1784   0.00797   0.00138  -0.0027   0.1293   1.0000
   1.750   0.2034   0.00843   0.00151  -0.0024   0.0495   1.0000
   2.000   0.2290   0.00864   0.00165  -0.0020   0.0380   1.0000
   2.250   0.2551   0.00880   0.00182  -0.0016   0.0357   1.0000
   2.500   0.2814   0.00897   0.00202  -0.0013   0.0348   1.0000
   2.750   0.3078   0.00918   0.00225  -0.0010   0.0342   1.0000
   3.000   0.3344   0.00941   0.00253  -0.0007   0.0339   1.0000
   3.250   0.3610   0.00968   0.00284  -0.0005   0.0336   1.0000
   3.500   0.3877   0.00995   0.00315  -0.0003   0.0326   1.0000
   3.750   0.4143   0.01024   0.00349  -0.0001   0.0314   1.0000
   4.000   0.4410   0.01054   0.00381   0.0000   0.0298   1.0000
   4.250   0.4673   0.01095   0.00425   0.0002   0.0276   1.0000
   4.500   0.4933   0.01150   0.00485   0.0004   0.0255   1.0000
   4.750   0.5207   0.01153   0.00494   0.0005   0.0243   1.0000
   5.000   0.5480   0.01161   0.00507   0.0005   0.0210   1.0000
   5.250   0.5751   0.01181   0.00529   0.0006   0.0170   1.0000
   5.500   0.6015   0.01225   0.00574   0.0008   0.0137   1.0000
   5.750   0.6282   0.01253   0.00605   0.0010   0.0117   1.0000
   6.000   0.6546   0.01295   0.00649   0.0011   0.0101   1.0000
   6.250   0.6795   0.01393   0.00760   0.0016   0.0087   1.0000
   6.500   0.7049   0.01472   0.00854   0.0019   0.0081   1.0000
   6.750   0.7302   0.01550   0.00945   0.0023   0.0074   1.0000
   7.000   0.7557   0.01610   0.01014   0.0025   0.0066   1.0000
   7.250   0.7807   0.01688   0.01103   0.0027   0.0061   1.0000
   7.500   0.8033   0.01856   0.01296   0.0033   0.0056   1.0000
   7.750   0.8264   0.02002   0.01470   0.0039   0.0054   1.0000
   8.000   0.8480   0.02198   0.01705   0.0045   0.0051   1.0000
   8.250   0.8667   0.02480   0.02034   0.0054   0.0049   1.0000
   8.500   0.8768   0.03033   0.02661   0.0068   0.0046   1.0000
   8.750   0.8076   0.06153   0.05928   0.0047   0.0048   1.0000
   9.000   0.7938   0.07184   0.06976  -0.0018   0.0049   1.0000
   9.250   0.7798   0.08059   0.07854  -0.0113   0.0050   1.0000
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