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NACA 0006 (naca0006-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 0006 (naca0006-il)
Reynolds number: 100,000
Max Cl/Cd: 32.86 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca0006-il-100000.txt
Download as CSV file: xf-naca0006-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0006                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.5661   0.11269   0.10814   0.0140   1.0000   0.0914
  -9.750  -0.5830   0.10909   0.10461   0.0097   1.0000   0.0925
  -9.500  -0.7037   0.11447   0.10977   0.0241   1.0000   0.0814
  -9.250  -0.6934   0.11087   0.10614   0.0257   1.0000   0.0859
  -9.000  -0.6952   0.10675   0.10207   0.0227   1.0000   0.0898
  -8.750  -0.7078   0.10227   0.09768   0.0146   1.0000   0.0921
  -8.500  -0.7208   0.09688   0.09233   0.0047   1.0000   0.0927
  -8.250  -0.6967   0.09341   0.08887   0.0159   1.0000   0.0987
  -8.000  -0.6997   0.08881   0.08432   0.0114   1.0000   0.1029
  -7.750  -0.7174   0.08268   0.07796  -0.0036   1.0000   0.1066
  -7.500  -0.6995   0.07812   0.07361   0.0024   1.0000   0.1110
  -7.250  -0.7061   0.07370   0.06878  -0.0067   1.0000   0.1208
  -7.000  -0.6891   0.06863   0.06399  -0.0028   1.0000   0.1254
  -6.750  -0.6168   0.05155   0.04715  -0.0071   1.0000   0.1467
  -6.500  -0.6741   0.05985   0.05481  -0.0076   1.0000   0.1506
  -6.250  -0.6605   0.05609   0.05099  -0.0071   1.0000   0.1655
  -6.000  -0.6320   0.04340   0.03670  -0.0112   1.0000   0.0785
  -5.750  -0.6107   0.03711   0.02979  -0.0103   1.0000   0.0638
  -5.500  -0.5895   0.03283   0.02504  -0.0095   1.0000   0.0618
  -5.250  -0.5658   0.02890   0.02053  -0.0084   1.0000   0.0592
  -5.000  -0.5404   0.02585   0.01692  -0.0073   1.0000   0.0599
  -4.750  -0.5137   0.02413   0.01460  -0.0061   1.0000   0.0658
  -4.500  -0.4883   0.02113   0.01156  -0.0054   1.0000   0.0722
  -4.250  -0.4629   0.01936   0.00971  -0.0046   1.0000   0.0856
  -4.000  -0.4379   0.01784   0.00818  -0.0037   1.0000   0.1036
  -3.750  -0.4139   0.01637   0.00674  -0.0027   1.0000   0.1222
  -3.500  -0.3910   0.01504   0.00566  -0.0017   1.0000   0.1576
  -3.250  -0.3789   0.01190   0.00470   0.0009   1.0000   0.5305
  -3.000  -0.3625   0.01103   0.00463   0.0052   1.0000   0.7482
  -2.750  -0.3350   0.01089   0.00457   0.0075   1.0000   0.8707
  -2.500  -0.2163   0.01134   0.00448  -0.0080   1.0000   0.9997
  -2.250  -0.1948   0.01103   0.00396  -0.0077   1.0000   1.0000
  -2.000  -0.1733   0.01079   0.00356  -0.0072   1.0000   1.0000
  -1.750  -0.1516   0.01060   0.00324  -0.0066   1.0000   1.0000
  -1.500  -0.1297   0.01046   0.00297  -0.0059   1.0000   1.0000
  -1.250  -0.1078   0.01034   0.00275  -0.0050   1.0000   1.0000
  -1.000  -0.0859   0.01026   0.00258  -0.0042   1.0000   1.0000
  -0.750  -0.0642   0.01020   0.00246  -0.0032   1.0000   1.0000
  -0.500  -0.0427   0.01015   0.00237  -0.0022   1.0000   1.0000
  -0.250  -0.0213   0.01013   0.00231  -0.0011   1.0000   1.0000
   0.000   0.0000   0.01012   0.00230   0.0000   1.0000   1.0000
   0.250   0.0213   0.01013   0.00231   0.0011   1.0000   1.0000
   0.500   0.0427   0.01015   0.00237   0.0022   1.0000   1.0000
   0.750   0.0642   0.01020   0.00246   0.0032   1.0000   1.0000
   1.000   0.0859   0.01026   0.00258   0.0042   1.0000   1.0000
   1.250   0.1078   0.01034   0.00274   0.0050   1.0000   1.0000
   1.500   0.1297   0.01046   0.00297   0.0058   1.0000   1.0000
   1.750   0.1516   0.01060   0.00324   0.0066   1.0000   1.0000
   2.000   0.1734   0.01078   0.00356   0.0072   1.0000   1.0000
   2.250   0.1949   0.01102   0.00396   0.0077   1.0000   1.0000
   2.500   0.2162   0.01134   0.00448   0.0080   0.9999   1.0000
   2.750   0.3350   0.01089   0.00457  -0.0075   0.8708   1.0000
   3.000   0.3624   0.01103   0.00463  -0.0051   0.7484   1.0000
   3.250   0.3788   0.01190   0.00470  -0.0009   0.5312   1.0000
   3.500   0.3909   0.01504   0.00566   0.0017   0.1577   1.0000
   3.750   0.4138   0.01637   0.00674   0.0027   0.1222   1.0000
   4.000   0.4379   0.01784   0.00818   0.0037   0.1036   1.0000
   4.250   0.4628   0.01936   0.00971   0.0046   0.0856   1.0000
   4.500   0.4883   0.02113   0.01156   0.0054   0.0721   1.0000
   4.750   0.5137   0.02414   0.01461   0.0061   0.0658   1.0000
   5.000   0.5404   0.02585   0.01692   0.0073   0.0600   1.0000
   5.250   0.5658   0.02891   0.02053   0.0084   0.0592   1.0000
   5.500   0.5895   0.03284   0.02505   0.0095   0.0618   1.0000
   5.750   0.6108   0.03711   0.02980   0.0103   0.0639   1.0000
   6.750   0.6843   0.06381   0.05894   0.0064   0.1371   1.0000
   7.000   0.6893   0.06864   0.06399   0.0028   0.1253   1.0000
   7.250   0.7060   0.07370   0.06878   0.0066   0.1208   1.0000
   7.500   0.6999   0.07812   0.07361  -0.0025   0.1110   1.0000
   7.750   0.7174   0.08269   0.07798   0.0035   0.1066   1.0000
   8.000   0.7000   0.08885   0.08436  -0.0116   0.1028   1.0000
   8.250   0.6973   0.09343   0.08889  -0.0160   0.0986   1.0000
   8.500   0.7210   0.09692   0.09237  -0.0049   0.0926   1.0000
   8.750   0.7080   0.10232   0.09774  -0.0149   0.0921   1.0000
   9.000   0.6956   0.10680   0.10212  -0.0229   0.0897   1.0000
   9.250   0.6941   0.11091   0.10619  -0.0258   0.0858   1.0000
   9.500   0.5989   0.10492   0.10051  -0.0054   0.0929   1.0000
   9.750   0.5825   0.10893   0.10445  -0.0097   0.0925   1.0000
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