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

NACA 64-206 (naca64206-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 64-206 (naca64206-il)
Reynolds number: 200,000
Max Cl/Cd: 56.17 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca64206-il-200000.txt
Download as CSV file: xf-naca64206-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64-206                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5472   0.09001   0.08653  -0.0076   1.0000   0.0268
  -8.000  -0.5465   0.08618   0.08276  -0.0104   1.0000   0.0275
  -7.750  -0.5483   0.08214   0.07879  -0.0163   1.0000   0.0287
  -7.500  -0.5424   0.07733   0.07393  -0.0234   1.0000   0.0291
  -6.750  -0.5271   0.06145   0.05778  -0.0330   1.0000   0.0310
  -6.500  -0.5188   0.05799   0.05432  -0.0326   1.0000   0.0328
  -6.250  -0.5065   0.05428   0.05052  -0.0337   1.0000   0.0345
  -6.000  -0.4917   0.05043   0.04651  -0.0351   1.0000   0.0366
  -5.750  -0.4733   0.04657   0.04239  -0.0363   1.0000   0.0397
  -5.500  -0.4460   0.04623   0.04131  -0.0364   1.0000   0.0427
  -5.250  -0.4359   0.03859   0.03376  -0.0376   1.0000   0.0450
  -5.000  -0.4176   0.03572   0.03082  -0.0375   1.0000   0.0480
  -4.500  -0.3749   0.03010   0.02462  -0.0371   1.0000   0.0603
  -4.250  -0.3525   0.02784   0.02200  -0.0368   1.0000   0.0718
  -4.000  -0.3307   0.02569   0.01966  -0.0364   1.0000   0.0853
  -3.750  -0.2966   0.02131   0.01452  -0.0337   1.0000   0.0395
  -3.500  -0.2715   0.01913   0.01194  -0.0327   1.0000   0.0396
  -3.250  -0.2465   0.01717   0.00971  -0.0316   1.0000   0.0389
  -3.000  -0.2217   0.01539   0.00771  -0.0304   1.0000   0.0371
  -2.750  -0.1977   0.01407   0.00629  -0.0294   1.0000   0.0378
  -2.500  -0.1740   0.01305   0.00521  -0.0284   1.0000   0.0406
  -2.250  -0.1506   0.01206   0.00424  -0.0278   1.0000   0.0481
  -2.000  -0.1265   0.01138   0.00352  -0.0272   1.0000   0.0544
  -1.750  -0.1004   0.01010   0.00259  -0.0273   1.0000   0.1555
  -1.500  -0.0929   0.00783   0.00289  -0.0230   1.0000   0.8167
  -1.250  -0.0877   0.00772   0.00299  -0.0172   1.0000   0.9207
  -1.000  -0.0499   0.00761   0.00277  -0.0195   1.0000   1.0000
  -0.750  -0.0261   0.00771   0.00274  -0.0196   1.0000   1.0000
  -0.500   0.0072   0.00785   0.00276  -0.0214   0.9968   1.0000
  -0.250   0.0501   0.00799   0.00279  -0.0251   0.9898   1.0000
   0.000   0.0940   0.00813   0.00286  -0.0288   0.9831   1.0000
   0.250   0.1351   0.00821   0.00291  -0.0320   0.9752   1.0000
   0.500   0.1799   0.00828   0.00298  -0.0358   0.9692   1.0000
   0.750   0.2204   0.00831   0.00304  -0.0387   0.9607   1.0000
   1.000   0.2663   0.00832   0.00310  -0.0426   0.9553   1.0000
   1.250   0.3042   0.00833   0.00317  -0.0448   0.9453   1.0000
   1.500   0.3421   0.00833   0.00325  -0.0469   0.9350   1.0000
   1.750   0.3774   0.00826   0.00334  -0.0480   0.9207   1.0000
   2.000   0.4047   0.00802   0.00312  -0.0463   0.8895   1.0000
   2.250   0.4223   0.00783   0.00282  -0.0422   0.8315   1.0000
   2.500   0.4432   0.00789   0.00275  -0.0397   0.7713   1.0000
   2.750   0.4621   0.00825   0.00267  -0.0368   0.6518   1.0000
   3.000   0.4654   0.01185   0.00348  -0.0331   0.0734   1.0000
   3.250   0.4900   0.01285   0.00444  -0.0325   0.0487   1.0000
   3.500   0.5140   0.01396   0.00566  -0.0318   0.0422   1.0000
   3.750   0.5385   0.01512   0.00686  -0.0310   0.0383   1.0000
   4.000   0.5635   0.01657   0.00834  -0.0302   0.0366   1.0000
   4.250   0.5889   0.01865   0.01045  -0.0296   0.0342   1.0000
   4.500   0.6156   0.02064   0.01271  -0.0289   0.0315   1.0000
   4.750   0.6423   0.02350   0.01587  -0.0281   0.0320   1.0000
   5.750   0.7448   0.03778   0.03229  -0.0221   0.0542   1.0000
   6.000   0.7641   0.04062   0.03525  -0.0217   0.0480   1.0000
   6.250   0.7739   0.04887   0.04355  -0.0222   0.0455   1.0000
   6.500   0.7952   0.04850   0.04400  -0.0201   0.0391   1.0000
   6.750   0.8089   0.05214   0.04785  -0.0198   0.0366   1.0000
   7.000   0.8208   0.05601   0.05178  -0.0197   0.0349   1.0000
   7.250   0.8254   0.06350   0.05925  -0.0204   0.0334   1.0000
   7.500   0.8254   0.07004   0.06604  -0.0207   0.0330   1.0000
   7.750   0.8286   0.07380   0.07007  -0.0207   0.0328   1.0000
   8.000   0.8292   0.07725   0.07377  -0.0211   0.0325   1.0000
   8.250   0.8241   0.08122   0.07794  -0.0226   0.0320   1.0000
   8.500   0.8145   0.08555   0.08238  -0.0244   0.0317   1.0000
   8.750   0.8037   0.09055   0.08741  -0.0277   0.0317   1.0000
<< Back to NACA 64-206 (naca64206-il)

Polar data table (+)

Polar graphs


<< Back to NACA 64-206 (naca64206-il)