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=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA 64-206 (naca64206-il)
Reynolds number: 1,000,000
Max Cl/Cd: 68.37 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca64206-il-1000000-n5.txt
Download as CSV file: xf-naca64206-il-1000000-n5.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 =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5337   0.08375   0.08215  -0.0094   1.0000   0.0036
  -7.750  -0.5343   0.07974   0.07816  -0.0118   1.0000   0.0036
  -7.500  -0.5358   0.07559   0.07404  -0.0151   1.0000   0.0035
  -7.250  -0.5315   0.07002   0.06845  -0.0216   1.0000   0.0035
  -7.000  -0.5229   0.06508   0.06347  -0.0261   1.0000   0.0035
  -6.750  -0.5232   0.05803   0.05633  -0.0306   1.0000   0.0032
  -6.500  -0.5133   0.05260   0.05079  -0.0335   1.0000   0.0030
  -6.250  -0.5012   0.04724   0.04527  -0.0355   1.0000   0.0027
  -6.000  -0.4861   0.04228   0.04014  -0.0367   1.0000   0.0025
  -5.750  -0.4694   0.03738   0.03502  -0.0371   1.0000   0.0023
  -5.500  -0.4510   0.03246   0.02983  -0.0369   1.0000   0.0021
  -5.250  -0.4197   0.02544   0.02232  -0.0388   0.9953   0.0019
  -5.000  -0.3893   0.01167   0.00702  -0.0384   0.9893   0.0014
  -4.750  -0.3582   0.01010   0.00517  -0.0392   0.9848   0.0014
  -4.500  -0.3272   0.00907   0.00388  -0.0400   0.9787   0.0014
  -4.250  -0.2955   0.00841   0.00306  -0.0409   0.9719   0.0015
  -4.000  -0.2635   0.00795   0.00248  -0.0419   0.9642   0.0018
  -3.750  -0.2330   0.00766   0.00216  -0.0425   0.9538   0.0025
  -3.500  -0.2043   0.00744   0.00194  -0.0428   0.9412   0.0034
  -3.250  -0.1768   0.00729   0.00178  -0.0428   0.9277   0.0065
  -2.500  -0.0963   0.00716   0.00155  -0.0423   0.8878   0.0099
  -2.250  -0.0692   0.00692   0.00118  -0.0421   0.8751   0.0110
  -2.000  -0.0418   0.00677   0.00092  -0.0419   0.8625   0.0126
  -1.750  -0.0145   0.00667   0.00081  -0.0418   0.8501   0.0175
  -1.500   0.0129   0.00659   0.00072  -0.0417   0.8379   0.0255
  -1.250   0.0402   0.00650   0.00064  -0.0417   0.8256   0.0422
  -1.000   0.0676   0.00626   0.00057  -0.0418   0.8140   0.1096
  -0.750   0.0941   0.00523   0.00043  -0.0423   0.8026   0.4167
  -0.500   0.1199   0.00446   0.00045  -0.0424   0.7906   0.6735
  -0.250   0.1468   0.00433   0.00047  -0.0422   0.7787   0.7310
   0.000   0.1737   0.00428   0.00049  -0.0420   0.7662   0.7686
   0.250   0.2010   0.00430   0.00051  -0.0418   0.7497   0.7846
   0.500   0.2271   0.00444   0.00052  -0.0414   0.7042   0.7980
   0.750   0.2528   0.00466   0.00054  -0.0410   0.6435   0.8116
   1.000   0.2785   0.00493   0.00059  -0.0407   0.5736   0.8247
   1.250   0.3025   0.00557   0.00070  -0.0402   0.4219   0.8384
   1.500   0.3255   0.00653   0.00096  -0.0398   0.2170   0.8527
   1.750   0.3510   0.00694   0.00112  -0.0396   0.1336   0.8671
   2.000   0.3765   0.00727   0.00128  -0.0392   0.0728   0.8822
   2.250   0.4023   0.00745   0.00147  -0.0389   0.0483   0.8984
   2.500   0.4276   0.00767   0.00164  -0.0384   0.0243   0.9152
   2.750   0.4525   0.00783   0.00180  -0.0377   0.0160   0.9335
   3.000   0.4775   0.00809   0.00219  -0.0370   0.0129   0.9551
   3.250   0.5085   0.00830   0.00246  -0.0378   0.0126   0.9968
   3.500   0.5358   0.00852   0.00270  -0.0378   0.0124   1.0000
   3.750   0.5632   0.00870   0.00289  -0.0378   0.0118   1.0000
   4.000   0.5904   0.00892   0.00319  -0.0378   0.0106   1.0000
   4.250   0.6173   0.00923   0.00353  -0.0377   0.0086   1.0000
   4.500   0.6436   0.00975   0.00413  -0.0375   0.0077   1.0000
   4.750   0.6697   0.01025   0.00469  -0.0373   0.0069   1.0000
   5.000   0.6949   0.01100   0.00555  -0.0369   0.0060   1.0000
   5.250   0.7191   0.01200   0.00668  -0.0364   0.0050   1.0000
   5.500   0.7481   0.01150   0.00609  -0.0369   0.0044   1.0000
   5.750   0.7760   0.01135   0.00586  -0.0372   0.0031   1.0000
   6.000   0.8011   0.01204   0.00659  -0.0368   0.0018   1.0000
   6.250   0.8248   0.01322   0.00796  -0.0362   0.0016   1.0000
   6.500   0.8517   0.01329   0.00809  -0.0363   0.0015   1.0000
   6.750   0.8766   0.01403   0.00894  -0.0360   0.0013   1.0000
   7.000   0.9014   0.01480   0.00982  -0.0357   0.0011   1.0000
   7.250   0.9256   0.01579   0.01097  -0.0353   0.0010   1.0000
   7.500   0.9498   0.01671   0.01203  -0.0349   0.0008   1.0000
   7.750   0.9733   0.01792   0.01342  -0.0344   0.0007   1.0000
   8.000   0.9961   0.01936   0.01508  -0.0339   0.0006   1.0000
   8.250   0.9081   0.06088   0.05918  -0.0247   0.0006   1.0000
   8.500   0.8990   0.06786   0.06634  -0.0254   0.0006   1.0000
   8.750   0.8865   0.07419   0.07278  -0.0270   0.0006   1.0000
   9.000   0.8709   0.07855   0.07720  -0.0281   0.0007   1.0000
<< Back to NACA 64-206 (naca64206-il)

Polar data table (+)

Polar graphs


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