Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 0010 (naca0010-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010 (naca0010-il)
Reynolds number: 100,000
Max Cl/Cd: 35.83 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca0010-il-100000.txt
Download as CSV file: xf-naca0010-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.5558   0.10754   0.10264  -0.0010   1.0000   0.1703
 -10.250  -0.5343   0.10341   0.09848   0.0011   1.0000   0.1742
 -10.000  -0.8387   0.07240   0.06713  -0.0217   1.0000   0.0863
  -9.750  -0.8376   0.06731   0.06193  -0.0216   1.0000   0.0833
  -9.500  -0.8461   0.06193   0.05638  -0.0213   1.0000   0.0816
  -9.250  -0.8578   0.05611   0.05023  -0.0205   1.0000   0.0796
  -9.000  -0.8653   0.05042   0.04406  -0.0191   1.0000   0.0779
  -8.750  -0.8624   0.04631   0.03952  -0.0175   1.0000   0.0790
  -8.500  -0.8555   0.04268   0.03542  -0.0158   1.0000   0.0810
  -8.250  -0.8453   0.03918   0.03137  -0.0140   1.0000   0.0824
  -8.000  -0.8315   0.03632   0.02792  -0.0121   1.0000   0.0842
  -7.750  -0.8149   0.03338   0.02466  -0.0110   1.0000   0.0875
  -7.500  -0.7938   0.03141   0.02257  -0.0101   1.0000   0.0909
  -7.250  -0.7723   0.02952   0.02040  -0.0090   1.0000   0.0944
  -7.000  -0.7507   0.02814   0.01856  -0.0077   1.0000   0.0987
  -6.750  -0.7269   0.02595   0.01640  -0.0072   1.0000   0.1027
  -6.500  -0.7028   0.02452   0.01487  -0.0064   1.0000   0.1073
  -6.250  -0.6789   0.02328   0.01342  -0.0054   1.0000   0.1137
  -6.000  -0.6557   0.02196   0.01224  -0.0046   1.0000   0.1219
  -5.750  -0.6326   0.02066   0.01097  -0.0036   1.0000   0.1312
  -5.500  -0.6109   0.01945   0.00988  -0.0024   1.0000   0.1459
  -5.250  -0.5910   0.01823   0.00885  -0.0009   1.0000   0.1732
  -5.000  -0.5732   0.01696   0.00799   0.0007   1.0000   0.2248
  -4.750  -0.5559   0.01583   0.00735   0.0024   1.0000   0.2991
  -4.500  -0.5374   0.01504   0.00693   0.0041   1.0000   0.3732
  -4.250  -0.5181   0.01446   0.00663   0.0058   1.0000   0.4430
  -4.000  -0.4990   0.01399   0.00647   0.0077   1.0000   0.5103
  -3.750  -0.4795   0.01362   0.00637   0.0098   1.0000   0.5737
  -3.500  -0.4591   0.01333   0.00628   0.0118   1.0000   0.6288
  -3.250  -0.4382   0.01310   0.00620   0.0139   1.0000   0.6780
  -3.000  -0.4174   0.01294   0.00619   0.0161   1.0000   0.7254
  -2.750  -0.3966   0.01288   0.00623   0.0185   1.0000   0.7710
  -2.500  -0.3752   0.01291   0.00633   0.0210   1.0000   0.8146
  -2.250  -0.3522   0.01302   0.00647   0.0232   1.0000   0.8560
  -2.000  -0.3203   0.01327   0.00669   0.0237   1.0000   0.8918
  -1.750  -0.2755   0.01358   0.00690   0.0214   1.0000   0.9233
  -1.500  -0.2169   0.01387   0.00705   0.0160   1.0000   0.9486
  -1.250  -0.1443   0.01402   0.00705   0.0073   1.0000   0.9659
  -1.000  -0.0760   0.01395   0.00688  -0.0009   1.0000   0.9830
  -0.750  -0.0075   0.01371   0.00657  -0.0096   1.0000   0.9974
  -0.500   0.0072   0.01343   0.00629  -0.0087   1.0000   1.0000
  -0.250   0.0055   0.01326   0.00614  -0.0047   1.0000   1.0000
   0.000   0.0000   0.01320   0.00609   0.0000   1.0000   1.0000
   0.250  -0.0055   0.01326   0.00614   0.0047   1.0000   1.0000
   0.500  -0.0072   0.01343   0.00629   0.0087   1.0000   1.0000
   0.750   0.0075   0.01371   0.00656   0.0096   0.9974   1.0000
   1.000   0.0760   0.01395   0.00688   0.0009   0.9830   1.0000
   1.250   0.1442   0.01401   0.00705  -0.0073   0.9659   1.0000
   1.500   0.2168   0.01387   0.00705  -0.0160   0.9486   1.0000
   1.750   0.2754   0.01358   0.00690  -0.0214   0.9233   1.0000
   2.000   0.3203   0.01327   0.00669  -0.0237   0.8918   1.0000
   2.250   0.3522   0.01302   0.00647  -0.0232   0.8560   1.0000
   2.500   0.3752   0.01291   0.00633  -0.0210   0.8146   1.0000
   2.750   0.3965   0.01287   0.00623  -0.0185   0.7710   1.0000
   3.000   0.4174   0.01294   0.00619  -0.0161   0.7254   1.0000
   3.250   0.4382   0.01310   0.00620  -0.0139   0.6780   1.0000
   3.500   0.4591   0.01333   0.00627  -0.0118   0.6288   1.0000
   3.750   0.4795   0.01362   0.00637  -0.0098   0.5737   1.0000
   4.000   0.4990   0.01399   0.00647  -0.0077   0.5103   1.0000
   4.250   0.5181   0.01446   0.00663  -0.0058   0.4430   1.0000
   4.500   0.5374   0.01504   0.00693  -0.0041   0.3732   1.0000
   4.750   0.5558   0.01583   0.00735  -0.0024   0.2991   1.0000
   5.000   0.5732   0.01696   0.00799  -0.0007   0.2248   1.0000
   5.250   0.5910   0.01823   0.00885   0.0009   0.1732   1.0000
   5.500   0.6109   0.01945   0.00988   0.0024   0.1460   1.0000
   5.750   0.6326   0.02066   0.01097   0.0036   0.1313   1.0000
   6.000   0.6557   0.02196   0.01224   0.0046   0.1219   1.0000
   6.250   0.6789   0.02328   0.01342   0.0054   0.1137   1.0000
   6.500   0.7028   0.02452   0.01487   0.0064   0.1073   1.0000
   6.750   0.7269   0.02595   0.01640   0.0072   0.1027   1.0000
   7.000   0.7507   0.02814   0.01856   0.0077   0.0987   1.0000
   7.250   0.7723   0.02953   0.02040   0.0090   0.0944   1.0000
   7.500   0.7938   0.03141   0.02257   0.0101   0.0909   1.0000
   7.750   0.8149   0.03338   0.02466   0.0110   0.0875   1.0000
   8.000   0.8316   0.03633   0.02792   0.0121   0.0842   1.0000
   8.250   0.8453   0.03919   0.03137   0.0140   0.0824   1.0000
   8.500   0.8555   0.04269   0.03542   0.0158   0.0810   1.0000
   8.750   0.8625   0.04631   0.03952   0.0175   0.0790   1.0000
   9.000   0.8654   0.05042   0.04407   0.0190   0.0779   1.0000
   9.250   0.8579   0.05613   0.05025   0.0205   0.0796   1.0000
   9.500   0.8462   0.06195   0.05640   0.0213   0.0817   1.0000
   9.750   0.8378   0.06733   0.06195   0.0216   0.0834   1.0000
  10.000   0.8396   0.07272   0.06737   0.0216   0.0846   1.0000
<< Back to NACA 0010 (naca0010-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0010 (naca0010-il)