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=100,000 Ncrit=5


Details Polar file
Airfoil: NACA 16009 (naca16009-il)
Reynolds number: 100,000
Max Cl/Cd: 28.82 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca16009-il-100000-n5.txt
Download as CSV file: xf-naca16009-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.4783   0.11250   0.10752  -0.0262   1.0000   0.0268
 -11.000  -0.4820   0.10825   0.10330  -0.0271   1.0000   0.0257
 -10.750  -0.5956   0.11125   0.10605  -0.0219   1.0000   0.0284
 -10.500  -0.5962   0.10703   0.10185  -0.0224   1.0000   0.0267
 -10.250  -0.5992   0.10224   0.09711  -0.0243   1.0000   0.0255
 -10.000  -0.6041   0.09710   0.09202  -0.0267   1.0000   0.0245
  -9.750  -0.6110   0.09150   0.08647  -0.0300   1.0000   0.0237
  -9.500  -0.6218   0.08544   0.08045  -0.0342   1.0000   0.0231
  -9.250  -0.6368   0.08031   0.07531  -0.0361   1.0000   0.0225
  -9.000  -0.6551   0.07615   0.07111  -0.0353   1.0000   0.0220
  -8.750  -0.6757   0.07258   0.06750  -0.0323   1.0000   0.0215
  -8.500  -0.6950   0.06785   0.06270  -0.0290   1.0000   0.0204
  -8.250  -0.7041   0.06408   0.05867  -0.0252   1.0000   0.0194
  -7.750  -0.7105   0.05837   0.05248  -0.0182   1.0000   0.0187
  -7.500  -0.7160   0.05483   0.04874  -0.0144   1.0000   0.0185
  -7.250  -0.7191   0.05141   0.04509  -0.0105   1.0000   0.0184
  -7.000  -0.7198   0.04808   0.04147  -0.0066   1.0000   0.0182
  -6.750  -0.7180   0.04483   0.03789  -0.0028   1.0000   0.0181
  -6.500  -0.7134   0.04168   0.03437   0.0010   1.0000   0.0181
  -6.250  -0.7060   0.03868   0.03095   0.0045   1.0000   0.0181
  -6.000  -0.6955   0.03585   0.02767   0.0077   1.0000   0.0183
  -5.750  -0.6819   0.03326   0.02462   0.0105   1.0000   0.0185
  -5.500  -0.6657   0.03081   0.02172   0.0130   1.0000   0.0192
  -5.250  -0.6499   0.02852   0.01922   0.0146   1.0000   0.0218
  -5.000  -0.6300   0.02701   0.01747   0.0162   1.0000   0.0243
  -4.750  -0.6066   0.02509   0.01524   0.0175   1.0000   0.0252
  -4.500  -0.5835   0.02346   0.01340   0.0187   1.0000   0.0264
  -4.250  -0.5622   0.02211   0.01188   0.0202   1.0000   0.0279
  -4.000  -0.5426   0.02103   0.01055   0.0220   1.0000   0.0297
  -3.750  -0.5251   0.01986   0.00932   0.0239   1.0000   0.0336
  -3.500  -0.5059   0.01912   0.00853   0.0255   1.0000   0.0445
  -3.250  -0.4867   0.01827   0.00771   0.0272   1.0000   0.0684
  -3.000  -0.4689   0.01702   0.00705   0.0287   1.0000   0.1627
  -2.750  -0.4256   0.01477   0.00785   0.0273   1.0000   0.8316
  -2.500  -0.3773   0.01581   0.00858   0.0250   1.0000   0.9139
  -2.250  -0.2313   0.01759   0.00977   0.0029   1.0000   0.9700
  -2.000  -0.1905   0.01749   0.00948  -0.0004   1.0000   0.9769
  -1.750  -0.1550   0.01739   0.00918  -0.0027   1.0000   0.9835
  -1.500  -0.1191   0.01725   0.00892  -0.0051   1.0000   0.9888
  -1.250  -0.0853   0.01715   0.00873  -0.0071   1.0000   0.9943
  -1.000  -0.0505   0.01702   0.00853  -0.0093   1.0000   0.9987
  -0.750  -0.0327   0.01695   0.00841  -0.0081   1.0000   1.0000
  -0.500  -0.0218   0.01690   0.00834  -0.0054   1.0000   1.0000
  -0.250  -0.0109   0.01687   0.00830  -0.0027   1.0000   1.0000
   0.000   0.0000   0.01686   0.00829   0.0000   1.0000   1.0000
   0.250   0.0109   0.01687   0.00830   0.0027   1.0000   1.0000
   0.500   0.0218   0.01690   0.00834   0.0054   1.0000   1.0000
   0.750   0.0327   0.01695   0.00841   0.0081   1.0000   1.0000
   1.000   0.0505   0.01702   0.00853   0.0093   0.9987   1.0000
   1.250   0.0852   0.01715   0.00873   0.0071   0.9943   1.0000
   1.500   0.1190   0.01724   0.00892   0.0051   0.9888   1.0000
   1.750   0.1550   0.01739   0.00918   0.0027   0.9835   1.0000
   2.000   0.1905   0.01749   0.00947   0.0004   0.9769   1.0000
   2.250   0.2313   0.01759   0.00976  -0.0029   0.9700   1.0000
   2.500   0.3773   0.01580   0.00858  -0.0250   0.9139   1.0000
   2.750   0.4256   0.01477   0.00785  -0.0273   0.8316   1.0000
   3.000   0.4688   0.01702   0.00705  -0.0287   0.1631   1.0000
   3.250   0.4866   0.01826   0.00771  -0.0272   0.0684   1.0000
   3.500   0.5058   0.01912   0.00853  -0.0255   0.0443   1.0000
   3.750   0.5250   0.01986   0.00932  -0.0239   0.0336   1.0000
   4.000   0.5425   0.02102   0.01054  -0.0220   0.0297   1.0000
   4.250   0.5621   0.02210   0.01188  -0.0202   0.0279   1.0000
   4.500   0.5834   0.02345   0.01339  -0.0187   0.0264   1.0000
   4.750   0.6065   0.02508   0.01524  -0.0174   0.0252   1.0000
   5.000   0.6299   0.02701   0.01746  -0.0161   0.0243   1.0000
   5.250   0.6498   0.02850   0.01921  -0.0146   0.0218   1.0000
   5.500   0.6656   0.03080   0.02172  -0.0129   0.0191   1.0000
   5.750   0.6818   0.03325   0.02461  -0.0105   0.0185   1.0000
   6.000   0.6954   0.03584   0.02766  -0.0077   0.0183   1.0000
   6.250   0.7059   0.03867   0.03094  -0.0045   0.0181   1.0000
   6.500   0.7134   0.04167   0.03436  -0.0010   0.0181   1.0000
   6.750   0.7179   0.04482   0.03787   0.0028   0.0181   1.0000
   7.000   0.7197   0.04807   0.04146   0.0066   0.0182   1.0000
   7.250   0.7191   0.05140   0.04508   0.0105   0.0184   1.0000
   7.500   0.7160   0.05482   0.04873   0.0144   0.0185   1.0000
   7.750   0.7105   0.05836   0.05247   0.0182   0.0187   1.0000
   8.250   0.7041   0.06407   0.05866   0.0252   0.0194   1.0000
   8.500   0.6951   0.06787   0.06271   0.0290   0.0204   1.0000
   8.750   0.6758   0.07259   0.06751   0.0322   0.0215   1.0000
   9.000   0.6553   0.07616   0.07113   0.0353   0.0220   1.0000
   9.250   0.6371   0.08035   0.07535   0.0361   0.0226   1.0000
   9.500   0.6222   0.08549   0.08050   0.0341   0.0231   1.0000
   9.750   0.6115   0.09155   0.08652   0.0298   0.0237   1.0000
<< Back to NACA 16009 (naca16009-il)

Polar data table (+)

Polar graphs


<< Back to NACA 16009 (naca16009-il)