Airfoil Tools
Search 1638 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA 0008 (naca0008-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA 0008 (naca0008-il)
Reynolds number: 100,000
Max Cl/Cd: 34.55 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca0008-il-100000.txt
Download as CSV file: xf-naca0008-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0008                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.5657   0.09227   0.08768   0.0037   1.0000   0.1506
  -9.000  -0.5968   0.08657   0.08208  -0.0018   1.0000   0.1523
  -8.750  -0.7155   0.08997   0.08534   0.0035   1.0000   0.1405
  -8.500  -0.6980   0.08727   0.08260   0.0072   1.0000   0.1461
  -7.750  -0.7635   0.05433   0.04844  -0.0150   1.0000   0.0842
  -7.500  -0.7567   0.04621   0.03951  -0.0139   1.0000   0.0679
  -7.250  -0.7443   0.04127   0.03418  -0.0129   1.0000   0.0666
  -7.000  -0.7295   0.03756   0.02981  -0.0114   1.0000   0.0680
  -6.750  -0.7128   0.03389   0.02551  -0.0100   1.0000   0.0699
  -6.500  -0.6925   0.03043   0.02174  -0.0091   1.0000   0.0724
  -6.250  -0.6698   0.02856   0.01967  -0.0083   1.0000   0.0788
  -6.000  -0.6461   0.02650   0.01702  -0.0070   1.0000   0.0831
  -5.750  -0.6224   0.02439   0.01495  -0.0065   1.0000   0.0906
  -5.500  -0.5970   0.02293   0.01315  -0.0056   1.0000   0.0969
  -5.250  -0.5724   0.02125   0.01153  -0.0050   1.0000   0.1045
  -5.000  -0.5476   0.01991   0.01014  -0.0042   1.0000   0.1134
  -4.750  -0.5232   0.01871   0.00893  -0.0032   1.0000   0.1233
  -4.500  -0.5004   0.01744   0.00779  -0.0021   1.0000   0.1392
  -4.250  -0.4799   0.01598   0.00669  -0.0008   1.0000   0.1789
  -4.000  -0.4646   0.01393   0.00582   0.0010   1.0000   0.3548
  -3.750  -0.4471   0.01294   0.00553   0.0032   1.0000   0.5049
  -3.500  -0.4272   0.01239   0.00535   0.0054   1.0000   0.6034
  -3.250  -0.4074   0.01201   0.00528   0.0080   1.0000   0.6924
  -3.000  -0.3881   0.01184   0.00537   0.0114   1.0000   0.7806
  -2.750  -0.3652   0.01191   0.00552   0.0144   1.0000   0.8508
  -2.500  -0.3298   0.01214   0.00566   0.0147   1.0000   0.9037
  -2.250  -0.2728   0.01247   0.00575   0.0102   1.0000   0.9441
  -2.000  -0.1988   0.01262   0.00562   0.0015   1.0000   0.9717
  -1.750  -0.1253   0.01247   0.00523  -0.0080   1.0000   0.9946
  -1.500  -0.0917   0.01211   0.00478  -0.0103   1.0000   1.0000
  -1.250  -0.0747   0.01182   0.00444  -0.0093   1.0000   1.0000
  -1.000  -0.0582   0.01160   0.00418  -0.0080   1.0000   1.0000
  -0.750  -0.0424   0.01144   0.00399  -0.0063   1.0000   1.0000
  -0.500  -0.0276   0.01133   0.00387  -0.0044   1.0000   1.0000
  -0.250  -0.0137   0.01126   0.00379  -0.0022   1.0000   1.0000
   0.000   0.0000   0.01124   0.00377   0.0000   1.0000   1.0000
   0.250   0.0137   0.01126   0.00379   0.0022   1.0000   1.0000
   0.500   0.0276   0.01133   0.00387   0.0044   1.0000   1.0000
   0.750   0.0424   0.01144   0.00399   0.0063   1.0000   1.0000
   1.000   0.0582   0.01160   0.00418   0.0080   1.0000   1.0000
   1.250   0.0748   0.01182   0.00444   0.0093   1.0000   1.0000
   1.500   0.0917   0.01211   0.00478   0.0103   1.0000   1.0000
   1.750   0.1253   0.01247   0.00523   0.0080   0.9946   1.0000
   2.000   0.1988   0.01262   0.00562  -0.0015   0.9717   1.0000
   2.250   0.2728   0.01247   0.00575  -0.0102   0.9441   1.0000
   2.500   0.3298   0.01214   0.00566  -0.0147   0.9038   1.0000
   2.750   0.3652   0.01191   0.00552  -0.0144   0.8509   1.0000
   3.000   0.3881   0.01184   0.00537  -0.0114   0.7806   1.0000
   3.250   0.4074   0.01201   0.00528  -0.0080   0.6924   1.0000
   3.500   0.4273   0.01239   0.00535  -0.0054   0.6034   1.0000
   3.750   0.4471   0.01294   0.00553  -0.0032   0.5049   1.0000
   4.250   0.4799   0.01598   0.00669   0.0008   0.1789   1.0000
   4.500   0.5004   0.01744   0.00779   0.0021   0.1392   1.0000
   4.750   0.5232   0.01871   0.00893   0.0032   0.1233   1.0000
   5.000   0.5476   0.01991   0.01014   0.0042   0.1134   1.0000
   5.250   0.5724   0.02125   0.01153   0.0050   0.1045   1.0000
   5.500   0.5970   0.02294   0.01315   0.0056   0.0969   1.0000
   5.750   0.6224   0.02439   0.01495   0.0065   0.0906   1.0000
   6.000   0.6461   0.02650   0.01703   0.0070   0.0831   1.0000
   6.250   0.6698   0.02856   0.01967   0.0083   0.0788   1.0000
   6.500   0.6925   0.03043   0.02174   0.0091   0.0724   1.0000
   6.750   0.7128   0.03390   0.02552   0.0100   0.0699   1.0000
   7.000   0.7295   0.03757   0.02982   0.0114   0.0680   1.0000
   7.250   0.7443   0.04127   0.03418   0.0129   0.0666   1.0000
   7.500   0.7567   0.04621   0.03951   0.0139   0.0679   1.0000
   7.750   0.7636   0.05433   0.04843   0.0150   0.0842   1.0000
   8.500   0.6984   0.08728   0.08262  -0.0073   0.1461   1.0000
   8.750   0.7192   0.08972   0.08510  -0.0025   0.1403   1.0000
   9.000   0.7106   0.09575   0.09106  -0.0062   0.1377   1.0000
   9.250   0.5654   0.09211   0.08752  -0.0037   0.1507   1.0000
<< Back to NACA 0008 (naca0008-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0008 (naca0008-il)