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NACA M10 AIRFOIL (m10-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA M10 AIRFOIL (m10-il)
Reynolds number: 100,000
Max Cl/Cd: 46.5 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m10-il-100000.txt
Download as CSV file: xf-m10-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5714   0.10068   0.09588   0.0092   1.0000   0.0737
  -8.000  -0.5728   0.09725   0.09250   0.0047   1.0000   0.0761
  -7.750  -0.5734   0.09376   0.08900  -0.0068   1.0000   0.0777
  -7.500  -0.5687   0.08895   0.08412  -0.0133   1.0000   0.0786
  -7.250  -0.5580   0.08390   0.07920  -0.0055   1.0000   0.0814
  -7.000  -0.5481   0.08012   0.07541  -0.0064   1.0000   0.0852
  -6.750  -0.5346   0.07619   0.07125  -0.0166   1.0000   0.0916
  -6.500  -0.5249   0.07068   0.06575  -0.0178   1.0000   0.0938
  -6.250  -0.5128   0.06714   0.06227  -0.0158   1.0000   0.0984
  -6.000  -0.4958   0.06272   0.05757  -0.0214   1.0000   0.1079
  -5.750  -0.4812   0.05948   0.05434  -0.0203   1.0000   0.1169
  -5.500  -0.4647   0.05533   0.05008  -0.0219   1.0000   0.1249
  -5.250  -0.4454   0.05169   0.04623  -0.0238   1.0000   0.1372
  -5.000  -0.4262   0.04840   0.04278  -0.0246   1.0000   0.1512
  -4.750  -0.4074   0.04522   0.03952  -0.0246   1.0000   0.1663
  -4.500  -0.3889   0.04262   0.03681  -0.0246   1.0000   0.1937
  -4.250  -0.3709   0.03996   0.03412  -0.0239   1.0000   0.2223
  -4.000  -0.3530   0.03730   0.03146  -0.0229   1.0000   0.2525
  -3.750  -0.3363   0.03491   0.02910  -0.0214   1.0000   0.2943
  -3.500  -0.2649   0.02685   0.01846  -0.0262   1.0000   0.1050
  -3.250  -0.2389   0.02406   0.01540  -0.0256   1.0000   0.0997
  -3.000  -0.2119   0.02195   0.01283  -0.0248   1.0000   0.0963
  -2.750  -0.1867   0.02063   0.01132  -0.0241   1.0000   0.0999
  -2.500  -0.1613   0.01941   0.00989  -0.0232   1.0000   0.1026
  -2.250  -0.1362   0.01824   0.00853  -0.0224   1.0000   0.1036
  -2.000  -0.1115   0.01732   0.00748  -0.0215   1.0000   0.1058
  -1.750  -0.0878   0.01622   0.00649  -0.0207   1.0000   0.1101
  -1.500  -0.0646   0.01554   0.00587  -0.0198   1.0000   0.1179
  -1.250  -0.0422   0.01489   0.00536  -0.0190   1.0000   0.1320
  -1.000  -0.0196   0.01420   0.00480  -0.0182   1.0000   0.1530
  -0.750   0.0270   0.01098   0.00413  -0.0207   1.0000   1.0000
  -0.500   0.0453   0.01116   0.00412  -0.0193   1.0000   1.0000
  -0.250   0.0634   0.01141   0.00423  -0.0182   1.0000   1.0000
   0.000   0.0817   0.01172   0.00441  -0.0173   1.0000   1.0000
   0.250   0.1001   0.01210   0.00470  -0.0167   1.0000   1.0000
   0.500   0.1432   0.01238   0.00490  -0.0208   0.9909   1.0000
   0.750   0.1975   0.01255   0.00503  -0.0267   0.9765   1.0000
   1.000   0.2521   0.01263   0.00511  -0.0326   0.9620   1.0000
   1.250   0.3046   0.01265   0.00516  -0.0377   0.9468   1.0000
   1.500   0.3518   0.01264   0.00522  -0.0415   0.9299   1.0000
   1.750   0.3914   0.01265   0.00528  -0.0435   0.9108   1.0000
   2.000   0.4229   0.01272   0.00540  -0.0437   0.8892   1.0000
   2.250   0.4504   0.01282   0.00554  -0.0429   0.8681   1.0000
   2.500   0.4747   0.01296   0.00576  -0.0414   0.8458   1.0000
   2.750   0.4979   0.01313   0.00596  -0.0397   0.8236   1.0000
   3.000   0.5206   0.01329   0.00615  -0.0378   0.8012   1.0000
   3.250   0.5414   0.01336   0.00622  -0.0351   0.7724   1.0000
   3.500   0.5610   0.01333   0.00619  -0.0320   0.7343   1.0000
   3.750   0.5813   0.01331   0.00611  -0.0292   0.6926   1.0000
   4.000   0.6025   0.01333   0.00606  -0.0266   0.6459   1.0000
   4.250   0.6236   0.01341   0.00603  -0.0243   0.5803   1.0000
   4.500   0.6426   0.01384   0.00604  -0.0216   0.4503   1.0000
   4.750   0.6525   0.01723   0.00732  -0.0193   0.1204   1.0000
   5.000   0.6743   0.01870   0.00867  -0.0182   0.0975   1.0000
   5.250   0.6959   0.02015   0.01008  -0.0170   0.0848   1.0000
   5.500   0.7176   0.02205   0.01178  -0.0160   0.0761   1.0000
   5.750   0.7433   0.02364   0.01351  -0.0149   0.0715   1.0000
   6.000   0.7678   0.02569   0.01549  -0.0143   0.0648   1.0000
   6.250   0.7938   0.02827   0.01832  -0.0135   0.0624   1.0000
   6.500   0.8195   0.03085   0.02127  -0.0125   0.0618   1.0000
   6.750   0.8436   0.03386   0.02472  -0.0115   0.0619   1.0000
   7.000   0.8654   0.03662   0.02802  -0.0102   0.0604   1.0000
   7.250   0.8850   0.04056   0.03241  -0.0091   0.0617   1.0000
   7.500   0.9031   0.04570   0.03847  -0.0071   0.0732   1.0000
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