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

NACA 64-208 (naca64208-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 64-208 (naca64208-il)
Reynolds number: 50,000
Max Cl/Cd: 26.62 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca64208-il-50000.txt
Download as CSV file: xf-naca64208-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 64-208                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5358   0.10799   0.10110   0.0041   1.0000   0.2718
  -8.750  -0.5358   0.10514   0.09826   0.0039   1.0000   0.2876
  -8.500  -0.5441   0.10299   0.09620   0.0033   1.0000   0.3032
  -8.250  -0.5186   0.09796   0.09115   0.0055   1.0000   0.3271
  -8.000  -0.5150   0.09482   0.08806   0.0062   1.0000   0.3481
  -7.750  -0.5190   0.09287   0.08620   0.0073   1.0000   0.3726
  -7.500  -0.4903   0.08843   0.08171   0.0100   1.0000   0.4101
  -7.250  -0.4924   0.08660   0.07997   0.0120   1.0000   0.4416
  -7.000  -0.4669   0.08218   0.07551   0.0136   1.0000   0.4734
  -5.750  -0.5153   0.04825   0.04067  -0.0358   1.0000   0.1677
  -5.500  -0.4929   0.04354   0.03535  -0.0368   1.0000   0.1453
  -5.250  -0.4705   0.03977   0.03087  -0.0369   1.0000   0.1343
  -5.000  -0.4475   0.03619   0.02673  -0.0364   1.0000   0.1263
  -4.750  -0.4226   0.03340   0.02334  -0.0355   1.0000   0.1207
  -4.500  -0.3982   0.03069   0.02026  -0.0346   1.0000   0.1186
  -4.250  -0.3731   0.02841   0.01759  -0.0335   1.0000   0.1189
  -4.000  -0.3490   0.02646   0.01537  -0.0323   1.0000   0.1259
  -3.750  -0.3249   0.02474   0.01350  -0.0309   1.0000   0.1335
  -3.500  -0.3012   0.02309   0.01182  -0.0291   1.0000   0.1412
  -3.250  -0.2794   0.02168   0.01046  -0.0274   1.0000   0.1591
  -3.000  -0.0997   0.01827   0.00927  -0.0356   1.0000   1.0000
  -2.750  -0.0887   0.01782   0.00864  -0.0344   1.0000   1.0000
  -2.500  -0.0797   0.01742   0.00810  -0.0327   1.0000   1.0000
  -2.250  -0.0737   0.01707   0.00766  -0.0305   1.0000   1.0000
  -2.000  -0.0713   0.01678   0.00731  -0.0277   1.0000   1.0000
  -1.750  -0.0733   0.01653   0.00702  -0.0242   1.0000   1.0000
  -1.500  -0.0795   0.01632   0.00677  -0.0199   1.0000   1.0000
  -1.250  -0.0857   0.01615   0.00653  -0.0156   1.0000   1.0000
  -1.000  -0.0832   0.01606   0.00627  -0.0127   1.0000   1.0000
  -0.750  -0.0712   0.01608   0.00612  -0.0113   1.0000   1.0000
  -0.500  -0.0547   0.01618   0.00605  -0.0107   1.0000   1.0000
  -0.250  -0.0362   0.01634   0.00605  -0.0103   1.0000   1.0000
   0.000  -0.0167   0.01655   0.00613  -0.0100   1.0000   1.0000
   0.250   0.0032   0.01681   0.00626  -0.0098   1.0000   1.0000
   0.500   0.0233   0.01711   0.00646  -0.0097   1.0000   1.0000
   0.750   0.0435   0.01745   0.00672  -0.0096   1.0000   1.0000
   1.000   0.0635   0.01784   0.00705  -0.0096   1.0000   1.0000
   1.250   0.0836   0.01826   0.00744  -0.0096   1.0000   1.0000
   1.500   0.1035   0.01873   0.00789  -0.0096   1.0000   1.0000
   1.750   0.1234   0.01923   0.00839  -0.0096   1.0000   1.0000
   2.000   0.1429   0.01978   0.00895  -0.0097   1.0000   1.0000
   2.250   0.1624   0.02037   0.00957  -0.0098   1.0000   1.0000
   2.500   0.1816   0.02101   0.01026  -0.0099   1.0000   1.0000
   2.750   0.2007   0.02169   0.01104  -0.0101   1.0000   1.0000
   3.000   0.2195   0.02242   0.01186  -0.0103   1.0000   1.0000
   3.250   0.2472   0.02339   0.01296  -0.0123   0.9954   1.0000
   3.500   0.2955   0.02478   0.01460  -0.0181   0.9799   1.0000
   3.750   0.3421   0.02606   0.01624  -0.0233   0.9622   1.0000
   4.000   0.3994   0.02743   0.01804  -0.0300   0.9409   1.0000
   4.250   0.4612   0.02827   0.01946  -0.0362   0.9056   1.0000
   4.500   0.5877   0.02208   0.01108  -0.0222   0.2009   1.0000
   4.750   0.6073   0.02432   0.01295  -0.0206   0.1580   1.0000
   5.000   0.6358   0.02651   0.01490  -0.0198   0.1382   1.0000
   5.250   0.6668   0.02856   0.01707  -0.0193   0.1219   1.0000
   5.500   0.6995   0.03099   0.01970  -0.0188   0.1159   1.0000
   5.750   0.7298   0.03388   0.02278  -0.0184   0.1132   1.0000
   6.000   0.7561   0.03699   0.02628  -0.0176   0.1102   1.0000
   6.250   0.7795   0.03986   0.02975  -0.0165   0.1077   1.0000
   6.500   0.8020   0.04361   0.03395  -0.0156   0.1095   1.0000
   6.750   0.8214   0.04736   0.03850  -0.0142   0.1159   1.0000
   7.000   0.8391   0.05232   0.04390  -0.0135   0.1237   1.0000
   7.250   0.8559   0.05797   0.05006  -0.0130   0.1383   1.0000
   7.500   0.8521   0.06439   0.05768  -0.0142   0.1712   1.0000
   7.750   0.8347   0.07782   0.07181  -0.0251   0.2624   1.0000
   8.000   0.7790   0.08609   0.08000  -0.0355   0.2868   1.0000
   8.250   0.7043   0.07709   0.07132  -0.0206   0.2136   1.0000
   8.500   0.7070   0.10317   0.09667  -0.0571   0.4153   1.0000
   8.750   0.7105   0.10689   0.10042  -0.0568   0.3982   1.0000
   9.000   0.5906   0.10517   0.09905  -0.0456   0.3773   1.0000
<< Back to NACA 64-208 (naca64208-il)

Polar data table (+)

Polar graphs


<< Back to NACA 64-208 (naca64208-il)