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

NACA 16009 (naca16009-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA 16009 (naca16009-il)
Reynolds number: 1,000,000
Max Cl/Cd: 47.07 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca16009-il-1000000-n5.txt
Download as CSV file: xf-naca16009-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 16009                                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.6648   0.07769   0.07613  -0.0369   1.0000   0.0020
 -10.000  -0.6854   0.07178   0.07017  -0.0399   1.0000   0.0020
  -9.750  -0.7091   0.06780   0.06615  -0.0383   1.0000   0.0019
  -8.750  -0.8460   0.02448   0.02069  -0.0170   0.9882   0.0020
  -8.500  -0.8290   0.02170   0.01751  -0.0157   0.9866   0.0021
  -8.250  -0.8058   0.02025   0.01583  -0.0155   0.9854   0.0023
  -8.000  -0.7804   0.01914   0.01454  -0.0156   0.9845   0.0023
  -7.750  -0.7588   0.01675   0.01179  -0.0150   0.9836   0.0026
  -7.500  -0.7315   0.01595   0.01088  -0.0155   0.9829   0.0029
  -7.250  -0.7032   0.01541   0.01026  -0.0162   0.9823   0.0030
  -7.000  -0.6762   0.01477   0.00954  -0.0165   0.9814   0.0032
  -6.750  -0.6547   0.01417   0.00886  -0.0154   0.9786   0.0035
  -6.500  -0.6304   0.01349   0.00805  -0.0150   0.9766   0.0037
  -6.250  -0.6048   0.01288   0.00735  -0.0149   0.9748   0.0041
  -6.000  -0.5779   0.01240   0.00681  -0.0150   0.9733   0.0046
  -5.750  -0.5505   0.01199   0.00635  -0.0153   0.9719   0.0050
  -5.500  -0.5232   0.01156   0.00586  -0.0155   0.9706   0.0052
  -5.250  -0.4983   0.01076   0.00494  -0.0151   0.9691   0.0059
  -5.000  -0.4755   0.01025   0.00435  -0.0142   0.9665   0.0066
  -4.750  -0.4517   0.00991   0.00394  -0.0135   0.9635   0.0072
  -4.500  -0.4261   0.00961   0.00359  -0.0132   0.9612   0.0077
  -4.250  -0.3996   0.00934   0.00329  -0.0132   0.9591   0.0084
  -4.000  -0.3728   0.00908   0.00299  -0.0131   0.9573   0.0089
  -3.750  -0.3452   0.00885   0.00271  -0.0133   0.9557   0.0095
  -3.500  -0.3204   0.00866   0.00251  -0.0127   0.9527   0.0123
  -3.250  -0.2959   0.00841   0.00230  -0.0122   0.9495   0.0257
  -3.000  -0.2704   0.00817   0.00213  -0.0119   0.9468   0.0454
  -2.750  -0.2441   0.00795   0.00198  -0.0118   0.9444   0.0663
  -2.500  -0.2175   0.00769   0.00183  -0.0118   0.9424   0.0973
  -2.250  -0.1953   0.00731   0.00169  -0.0109   0.9386   0.1666
  -2.000  -0.1787   0.00646   0.00149  -0.0090   0.9341   0.3474
  -1.750  -0.1654   0.00541   0.00122  -0.0064   0.9301   0.5672
  -1.500  -0.1482   0.00489   0.00110  -0.0042   0.9258   0.6777
  -1.250  -0.1268   0.00463   0.00104  -0.0029   0.9216   0.7366
  -1.000  -0.1046   0.00439   0.00102  -0.0016   0.9181   0.7999
  -0.750  -0.0800   0.00425   0.00106  -0.0008   0.9152   0.8531
  -0.500  -0.0548   0.00424   0.00113  -0.0002   0.9109   0.8810
  -0.250  -0.0276   0.00424   0.00117   0.0000   0.9067   0.8942
   0.000   0.0000   0.00424   0.00114   0.0000   0.9023   0.9024
   0.250   0.0276   0.00424   0.00117   0.0000   0.8942   0.9067
   0.500   0.0548   0.00424   0.00113   0.0002   0.8810   0.9109
   0.750   0.0800   0.00425   0.00106   0.0008   0.8535   0.9152
   1.000   0.1045   0.00438   0.00102   0.0016   0.7998   0.9181
   1.250   0.1267   0.00463   0.00104   0.0029   0.7361   0.9216
   1.500   0.1482   0.00489   0.00110   0.0042   0.6782   0.9257
   1.750   0.1658   0.00539   0.00121   0.0063   0.5719   0.9300
   2.000   0.1791   0.00643   0.00148   0.0089   0.3539   0.9340
   2.250   0.1953   0.00731   0.00169   0.0109   0.1666   0.9386
   2.500   0.2175   0.00770   0.00183   0.0118   0.0966   0.9424
   2.750   0.2441   0.00795   0.00198   0.0118   0.0662   0.9444
   3.000   0.2705   0.00817   0.00213   0.0119   0.0457   0.9467
   3.250   0.2959   0.00841   0.00230   0.0122   0.0256   0.9495
   3.500   0.3204   0.00867   0.00251   0.0127   0.0122   0.9527
   3.750   0.3452   0.00885   0.00271   0.0133   0.0095   0.9557
   4.000   0.3728   0.00908   0.00298   0.0131   0.0089   0.9572
   4.250   0.3996   0.00934   0.00329   0.0132   0.0084   0.9591
   4.500   0.4261   0.00961   0.00359   0.0132   0.0077   0.9611
   4.750   0.4516   0.00992   0.00394   0.0135   0.0072   0.9636
   5.000   0.4753   0.01025   0.00435   0.0142   0.0067   0.9666
   5.250   0.4983   0.01077   0.00495   0.0151   0.0059   0.9692
   5.500   0.5232   0.01155   0.00585   0.0155   0.0052   0.9706
   5.750   0.5505   0.01199   0.00635   0.0153   0.0050   0.9719
   6.000   0.5778   0.01242   0.00683   0.0150   0.0046   0.9733
   6.250   0.6048   0.01285   0.00732   0.0149   0.0041   0.9748
   6.500   0.6303   0.01349   0.00805   0.0150   0.0037   0.9766
   6.750   0.6547   0.01415   0.00884   0.0154   0.0035   0.9787
   7.000   0.6759   0.01479   0.00956   0.0166   0.0033   0.9815
   7.250   0.7034   0.01544   0.01030   0.0162   0.0031   0.9823
   7.500   0.7317   0.01596   0.01089   0.0155   0.0029   0.9830
   7.750   0.7588   0.01677   0.01181   0.0150   0.0026   0.9836
   8.000   0.7800   0.01920   0.01461   0.0157   0.0023   0.9845
   8.250   0.8060   0.02014   0.01571   0.0155   0.0023   0.9854
   8.500   0.8289   0.02172   0.01753   0.0158   0.0021   0.9866
   8.750   0.8460   0.02449   0.02070   0.0170   0.0020   0.9882
   9.250   0.7535   0.06020   0.05839   0.0308   0.0019   0.9984
   9.500   0.7358   0.06540   0.06370   0.0330   0.0019   1.0000
   9.750   0.7092   0.06781   0.06616   0.0383   0.0019   1.0000
  10.000   0.6854   0.07184   0.07023   0.0398   0.0020   1.0000
  10.250   0.6656   0.07765   0.07609   0.0368   0.0020   1.0000
<< Back to NACA 16009 (naca16009-il)

Polar data table (+)

Polar graphs


<< Back to NACA 16009 (naca16009-il)