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NACA 16009 (naca16009-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 16009 (naca16009-il)
Reynolds number: 50,000
Max Cl/Cd: 17.99 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca16009-il-50000.txt
Download as CSV file: xf-naca16009-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.5917   0.12785   0.12063  -0.0051   1.0000   0.2285
 -10.500  -0.5973   0.12466   0.11751  -0.0058   1.0000   0.2385
 -10.250  -0.6118   0.12260   0.11555  -0.0069   1.0000   0.2494
 -10.000  -0.6090   0.11884   0.11182  -0.0063   1.0000   0.2630
  -9.750  -0.6110   0.11544   0.10848  -0.0059   1.0000   0.2768
  -9.500  -0.5875   0.11025   0.10325  -0.0036   1.0000   0.2950
  -9.250  -0.5854   0.10693   0.09997  -0.0025   1.0000   0.3137
  -9.000  -0.6041   0.10483   0.09799  -0.0019   1.0000   0.3313
  -8.750  -0.5819   0.10028   0.09341   0.0005   1.0000   0.3556
  -8.500  -0.5741   0.09668   0.08983   0.0025   1.0000   0.3794
  -8.250  -0.5690   0.09339   0.08658   0.0047   1.0000   0.4059
  -8.000  -0.5719   0.09100   0.08426   0.0078   1.0000   0.4397
  -6.500  -0.7102   0.06376   0.05736   0.0003   1.0000   0.2964
  -6.250  -0.7116   0.05477   0.04683  -0.0040   1.0000   0.1504
  -6.000  -0.7017   0.05027   0.04160  -0.0009   1.0000   0.1243
  -5.750  -0.6894   0.04645   0.03716   0.0023   1.0000   0.1114
  -5.500  -0.6753   0.04275   0.03292   0.0052   1.0000   0.1041
  -5.250  -0.6577   0.03991   0.02937   0.0081   1.0000   0.0991
  -5.000  -0.6383   0.03710   0.02606   0.0104   1.0000   0.0993
  -4.750  -0.6168   0.03428   0.02305   0.0118   1.0000   0.1050
  -4.500  -0.2346   0.02518   0.01557  -0.0367   1.0000   1.0000
  -4.250  -0.2215   0.02455   0.01462  -0.0353   1.0000   1.0000
  -4.000  -0.2081   0.02401   0.01371  -0.0339   1.0000   1.0000
  -3.750  -0.1947   0.02354   0.01298  -0.0323   1.0000   1.0000
  -3.500  -0.1813   0.02312   0.01233  -0.0306   1.0000   1.0000
  -3.250  -0.1679   0.02276   0.01176  -0.0288   1.0000   1.0000
  -3.000  -0.1545   0.02244   0.01126  -0.0269   1.0000   1.0000
  -2.750  -0.1410   0.02216   0.01081  -0.0250   1.0000   1.0000
  -2.500  -0.1277   0.02191   0.01042  -0.0230   1.0000   1.0000
  -2.250  -0.1144   0.02170   0.01002  -0.0209   1.0000   1.0000
  -2.000  -0.1013   0.02151   0.00972  -0.0187   1.0000   1.0000
  -1.750  -0.0883   0.02135   0.00946  -0.0165   1.0000   1.0000
  -1.500  -0.0754   0.02122   0.00924  -0.0143   1.0000   1.0000
  -1.250  -0.0627   0.02110   0.00906  -0.0120   1.0000   1.0000
  -1.000  -0.0500   0.02101   0.00892  -0.0096   1.0000   1.0000
  -0.750  -0.0374   0.02095   0.00879  -0.0072   1.0000   1.0000
  -0.500  -0.0249   0.02090   0.00871  -0.0048   1.0000   1.0000
  -0.250  -0.0124   0.02087   0.00866  -0.0024   1.0000   1.0000
   0.000   0.0000   0.02086   0.00865   0.0000   1.0000   1.0000
   0.250   0.0125   0.02087   0.00866   0.0024   1.0000   1.0000
   0.500   0.0249   0.02090   0.00871   0.0048   1.0000   1.0000
   0.750   0.0374   0.02094   0.00879   0.0072   1.0000   1.0000
   1.000   0.0500   0.02101   0.00892   0.0096   1.0000   1.0000
   1.250   0.0627   0.02110   0.00906   0.0120   1.0000   1.0000
   1.500   0.0754   0.02121   0.00924   0.0143   1.0000   1.0000
   1.750   0.0883   0.02134   0.00945   0.0165   1.0000   1.0000
   2.000   0.1013   0.02150   0.00971   0.0187   1.0000   1.0000
   2.250   0.1144   0.02169   0.01001   0.0209   1.0000   1.0000
   2.500   0.1277   0.02190   0.01041   0.0230   1.0000   1.0000
   2.750   0.1411   0.02215   0.01080   0.0250   1.0000   1.0000
   3.000   0.1545   0.02243   0.01124   0.0269   1.0000   1.0000
   3.250   0.1680   0.02275   0.01175   0.0288   1.0000   1.0000
   3.500   0.1814   0.02311   0.01232   0.0306   1.0000   1.0000
   3.750   0.1948   0.02352   0.01296   0.0323   1.0000   1.0000
   4.000   0.2083   0.02399   0.01369   0.0338   1.0000   1.0000
   4.250   0.2216   0.02453   0.01460   0.0353   1.0000   1.0000
   4.500   0.2348   0.02516   0.01555   0.0366   1.0000   1.0000
   4.750   0.6167   0.03428   0.02304  -0.0118   0.1051   1.0000
   5.000   0.6382   0.03710   0.02605  -0.0104   0.0993   1.0000
   5.250   0.6576   0.03990   0.02936  -0.0081   0.0991   1.0000
   5.500   0.6752   0.04274   0.03291  -0.0051   0.1040   1.0000
   5.750   0.6894   0.04644   0.03714  -0.0022   0.1114   1.0000
   6.000   0.7016   0.05025   0.04159   0.0010   0.1243   1.0000
   6.250   0.7115   0.05476   0.04682   0.0041   0.1504   1.0000
   6.500   0.7100   0.06419   0.05784  -0.0013   0.3081   1.0000
   6.750   0.6567   0.07123   0.06507  -0.0089   0.4034   1.0000
   8.250   0.5697   0.09340   0.08659  -0.0048   0.4057   1.0000
   8.500   0.5748   0.09668   0.08984  -0.0026   0.3793   1.0000
   8.750   0.5820   0.10026   0.09338  -0.0006   0.3554   1.0000
   9.000   0.6042   0.10480   0.09796   0.0018   0.3313   1.0000
   9.250   0.5858   0.10691   0.09995   0.0024   0.3135   1.0000
   9.500   0.5888   0.11029   0.10329   0.0035   0.2947   1.0000
   9.750   0.6110   0.11542   0.10845   0.0057   0.2767   1.0000
  10.000   0.6091   0.11882   0.11179   0.0061   0.2629   1.0000
  10.250   0.4672   0.11365   0.10694   0.0132   0.2745   1.0000
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