Airfoil Tools
Search 1638 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=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 16009 (naca16009-il)
Reynolds number: 100,000
Max Cl/Cd: 25.11 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca16009-il-100000.txt
Download as CSV file: xf-naca16009-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5022   0.10256   0.09781  -0.0295   1.0000   0.0996
 -10.000  -0.5261   0.09799   0.09332  -0.0337   1.0000   0.1002
  -9.750  -0.6052   0.10114   0.09616  -0.0237   1.0000   0.0928
  -9.500  -0.6111   0.09688   0.09195  -0.0258   1.0000   0.0960
  -9.250  -0.6274   0.09192   0.08711  -0.0302   1.0000   0.0980
  -9.000  -0.6517   0.08821   0.08341  -0.0315   1.0000   0.0989
  -8.750  -0.6803   0.08581   0.08100  -0.0293   1.0000   0.0996
  -8.500  -0.7106   0.08416   0.07919  -0.0263   1.0000   0.1003
  -8.250  -0.6874   0.07777   0.07301  -0.0256   1.0000   0.1045
  -8.000  -0.6929   0.07478   0.06993  -0.0232   1.0000   0.1084
  -7.750  -0.7229   0.07365   0.06854  -0.0192   1.0000   0.1134
  -7.500  -0.7185   0.06849   0.06342  -0.0174   1.0000   0.1167
  -7.250  -0.7158   0.06554   0.06045  -0.0148   1.0000   0.1228
  -7.000  -0.7256   0.06251   0.05724  -0.0113   1.0000   0.1302
  -6.750  -0.7363   0.06144   0.05579  -0.0068   1.0000   0.1415
  -6.500  -0.7212   0.05660   0.05119  -0.0054   1.0000   0.1475
  -6.250  -0.7204   0.05379   0.04827  -0.0022   1.0000   0.1600
  -6.000  -0.7170   0.05112   0.04551   0.0010   1.0000   0.1749
  -5.500  -0.6814   0.03890   0.03082   0.0104   1.0000   0.0601
  -5.250  -0.6663   0.03516   0.02682   0.0127   1.0000   0.0579
  -5.000  -0.6491   0.03258   0.02379   0.0153   1.0000   0.0583
  -4.750  -0.6283   0.02976   0.02056   0.0174   1.0000   0.0568
  -4.500  -0.6039   0.02713   0.01751   0.0190   1.0000   0.0553
  -4.250  -0.5772   0.02496   0.01500   0.0202   1.0000   0.0557
  -4.000  -0.5501   0.02318   0.01297   0.0212   1.0000   0.0584
  -3.750  -0.5259   0.02155   0.01130   0.0221   1.0000   0.0672
  -3.500  -0.5044   0.02008   0.00988   0.0237   1.0000   0.0792
  -3.250  -0.1524   0.01878   0.01104  -0.0313   1.0000   1.0000
  -3.000  -0.1394   0.01845   0.01058  -0.0294   1.0000   1.0000
  -2.750  -0.1266   0.01816   0.01018  -0.0274   1.0000   1.0000
  -2.500  -0.1140   0.01791   0.00983  -0.0253   1.0000   1.0000
  -2.250  -0.1017   0.01769   0.00948  -0.0231   1.0000   1.0000
  -2.000  -0.0896   0.01750   0.00922  -0.0208   1.0000   1.0000
  -1.750  -0.0778   0.01734   0.00900  -0.0184   1.0000   1.0000
  -1.500  -0.0663   0.01721   0.00881  -0.0159   1.0000   1.0000
  -1.250  -0.0549   0.01710   0.00866  -0.0133   1.0000   1.0000
  -1.000  -0.0437   0.01701   0.00853  -0.0107   1.0000   1.0000
  -0.750  -0.0327   0.01695   0.00842  -0.0081   1.0000   1.0000
  -0.500  -0.0217   0.01690   0.00836  -0.0054   1.0000   1.0000
  -0.250  -0.0108   0.01687   0.00831  -0.0027   1.0000   1.0000
   0.000   0.0000   0.01686   0.00830   0.0000   1.0000   1.0000
   0.250   0.0108   0.01687   0.00831   0.0027   1.0000   1.0000
   0.500   0.0217   0.01690   0.00835   0.0054   1.0000   1.0000
   0.750   0.0327   0.01694   0.00842   0.0081   1.0000   1.0000
   1.000   0.0437   0.01701   0.00853   0.0107   1.0000   1.0000
   1.250   0.0549   0.01710   0.00865   0.0134   1.0000   1.0000
   1.500   0.0663   0.01720   0.00880   0.0159   1.0000   1.0000
   1.750   0.0779   0.01734   0.00899   0.0184   1.0000   1.0000
   2.000   0.0897   0.01749   0.00921   0.0208   1.0000   1.0000
   2.250   0.1017   0.01768   0.00947   0.0231   1.0000   1.0000
   2.500   0.1141   0.01790   0.00982   0.0253   1.0000   1.0000
   2.750   0.1266   0.01815   0.01017   0.0274   1.0000   1.0000
   3.000   0.1394   0.01844   0.01057   0.0294   1.0000   1.0000
   3.250   0.1524   0.01877   0.01103   0.0313   1.0000   1.0000
   3.500   0.5043   0.02008   0.00987  -0.0237   0.0792   1.0000
   3.750   0.5258   0.02155   0.01130  -0.0221   0.0673   1.0000
   4.000   0.5500   0.02317   0.01297  -0.0212   0.0585   1.0000
   4.250   0.5771   0.02495   0.01498  -0.0202   0.0558   1.0000
   4.500   0.6037   0.02712   0.01749  -0.0190   0.0553   1.0000
   4.750   0.6282   0.02975   0.02054  -0.0174   0.0568   1.0000
   5.000   0.6490   0.03258   0.02379  -0.0153   0.0583   1.0000
   5.250   0.6662   0.03515   0.02681  -0.0127   0.0579   1.0000
   5.500   0.6814   0.03888   0.03080  -0.0104   0.0601   1.0000
   6.750   0.7362   0.06143   0.05577   0.0068   0.1415   1.0000
   7.000   0.7255   0.06250   0.05723   0.0113   0.1302   1.0000
   7.250   0.7158   0.06554   0.06045   0.0148   0.1228   1.0000
   7.500   0.7185   0.06848   0.06342   0.0174   0.1167   1.0000
   7.750   0.7229   0.07364   0.06853   0.0192   0.1134   1.0000
   8.000   0.6930   0.07477   0.06992   0.0232   0.1084   1.0000
   8.250   0.6876   0.07777   0.07300   0.0256   0.1045   1.0000
   8.500   0.7107   0.08415   0.07919   0.0263   0.1003   1.0000
   8.750   0.6803   0.08580   0.08099   0.0293   0.0996   1.0000
   9.000   0.6518   0.08821   0.08342   0.0314   0.0989   1.0000
   9.250   0.6277   0.09194   0.08713   0.0301   0.0979   1.0000
   9.500   0.6116   0.09690   0.09197   0.0257   0.0960   1.0000
<< Back to NACA 16009 (naca16009-il)

Polar data table (+)

Polar graphs


<< Back to NACA 16009 (naca16009-il)