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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 200,000
Max Cl/Cd: 58.75 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m10-il-200000.txt
Download as CSV file: xf-m10-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M10 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5775   0.10085   0.09739   0.0120   1.0000   0.0356
  -8.250  -0.5756   0.09698   0.09355   0.0086   1.0000   0.0368
  -8.000  -0.5754   0.09302   0.08964   0.0026   1.0000   0.0377
  -7.750  -0.5671   0.08829   0.08489  -0.0062   1.0000   0.0383
  -7.500  -0.5558   0.08356   0.08008  -0.0121   1.0000   0.0386
  -7.250  -0.5437   0.07883   0.07524  -0.0161   1.0000   0.0387
  -7.000  -0.5418   0.07118   0.06759  -0.0184   1.0000   0.0395
  -6.750  -0.5351   0.06807   0.06456  -0.0161   1.0000   0.0410
  -6.500  -0.5226   0.06521   0.06168  -0.0163   1.0000   0.0437
  -6.250  -0.5060   0.06101   0.05737  -0.0191   1.0000   0.0468
  -6.000  -0.4745   0.05716   0.05290  -0.0248   1.0000   0.0512
  -5.750  -0.4652   0.04995   0.04569  -0.0259   1.0000   0.0526
  -5.500  -0.4494   0.04705   0.04284  -0.0257   1.0000   0.0546
  -5.250  -0.4294   0.04418   0.03982  -0.0260   1.0000   0.0587
  -5.000  -0.4049   0.03966   0.03475  -0.0273   1.0000   0.0662
  -4.750  -0.3859   0.03754   0.03275  -0.0271   1.0000   0.0711
  -4.500  -0.3616   0.03422   0.02896  -0.0274   1.0000   0.0807
  -4.250  -0.3396   0.03193   0.02666  -0.0272   1.0000   0.0858
  -3.750  -0.2835   0.02325   0.01648  -0.0254   1.0000   0.0663
  -3.500  -0.2575   0.02059   0.01355  -0.0246   1.0000   0.0606
  -3.250  -0.2319   0.01879   0.01142  -0.0238   1.0000   0.0612
  -3.000  -0.2065   0.01761   0.00997  -0.0230   1.0000   0.0630
  -2.750  -0.1818   0.01636   0.00852  -0.0221   1.0000   0.0634
  -2.500  -0.1576   0.01546   0.00749  -0.0212   1.0000   0.0641
  -2.250  -0.1342   0.01418   0.00616  -0.0205   1.0000   0.0669
  -2.000  -0.1113   0.01350   0.00551  -0.0196   1.0000   0.0695
  -1.750  -0.0888   0.01292   0.00496  -0.0187   1.0000   0.0715
  -1.500  -0.0668   0.01248   0.00454  -0.0179   1.0000   0.0744
  -1.000   0.0023   0.01145   0.00359  -0.0214   0.9913   0.0890
  -0.750   0.0432   0.00982   0.00307  -0.0249   0.9844   0.3446
  -0.500   0.0901   0.00808   0.00310  -0.0279   0.9840   1.0000
  -0.250   0.1360   0.00803   0.00292  -0.0317   0.9712   1.0000
   0.000   0.1791   0.00799   0.00278  -0.0349   0.9558   1.0000
   0.250   0.2150   0.00796   0.00269  -0.0364   0.9350   1.0000
   0.500   0.2453   0.00795   0.00261  -0.0366   0.9136   1.0000
   0.750   0.2701   0.00800   0.00258  -0.0355   0.8899   1.0000
   1.000   0.2936   0.00807   0.00257  -0.0341   0.8680   1.0000
   1.250   0.3172   0.00817   0.00259  -0.0328   0.8462   1.0000
   1.500   0.3411   0.00829   0.00264  -0.0316   0.8258   1.0000
   1.750   0.3657   0.00841   0.00270  -0.0307   0.8050   1.0000
   2.000   0.3904   0.00855   0.00277  -0.0297   0.7859   1.0000
   2.250   0.4157   0.00869   0.00287  -0.0289   0.7643   1.0000
   2.500   0.4402   0.00882   0.00294  -0.0278   0.7383   1.0000
   2.750   0.4648   0.00895   0.00299  -0.0268   0.7084   1.0000
   3.000   0.4899   0.00909   0.00306  -0.0259   0.6791   1.0000
   3.250   0.5153   0.00926   0.00317  -0.0251   0.6496   1.0000
   3.500   0.5400   0.00944   0.00325  -0.0242   0.6063   1.0000
   3.750   0.5647   0.00968   0.00335  -0.0233   0.5550   1.0000
   4.000   0.5893   0.01003   0.00351  -0.0225   0.4915   1.0000
   4.250   0.6094   0.01119   0.00380  -0.0214   0.2968   1.0000
   4.750   0.6514   0.01465   0.00601  -0.0202   0.0626   1.0000
   5.000   0.6758   0.01551   0.00693  -0.0196   0.0548   1.0000
   5.250   0.6981   0.01691   0.00834  -0.0187   0.0493   1.0000
   5.500   0.7224   0.01787   0.00935  -0.0181   0.0438   1.0000
   5.750   0.7454   0.01948   0.01093  -0.0172   0.0405   1.0000
   6.000   0.7688   0.02224   0.01380  -0.0163   0.0385   1.0000
   6.250   0.7945   0.02305   0.01485  -0.0156   0.0357   1.0000
   6.500   0.8195   0.02505   0.01711  -0.0147   0.0344   1.0000
   6.750   0.8436   0.02769   0.02011  -0.0137   0.0344   1.0000
   7.000   0.8657   0.03125   0.02409  -0.0124   0.0359   1.0000
   7.250   0.8873   0.03588   0.02892  -0.0115   0.0397   1.0000
  14.250   0.6558   0.16055   0.15719  -0.0328   0.0386   1.0000
  14.500   0.6516   0.16341   0.16003  -0.0349   0.0383   1.0000
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