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

NACA 0010-66 (naca001066-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 0010-66 (naca001066-il)
Reynolds number: 200,000
Max Cl/Cd: 33.86 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca001066-il-200000.txt
Download as CSV file: xf-naca001066-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0010-66                                    
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.5771   0.09616   0.09277  -0.0248   1.0000   0.0746
 -10.000  -0.7307   0.08924   0.08567  -0.0292   1.0000   0.0663
  -9.750  -0.7484   0.08581   0.08223  -0.0272   1.0000   0.0668
  -9.500  -0.7717   0.08269   0.07908  -0.0237   1.0000   0.0672
  -9.250  -0.7979   0.08015   0.07651  -0.0184   1.0000   0.0674
  -9.000  -0.8164   0.07746   0.07377  -0.0137   1.0000   0.0681
  -8.750  -0.8304   0.07448   0.07072  -0.0096   1.0000   0.0692
  -8.500  -0.8438   0.07128   0.06742  -0.0053   1.0000   0.0707
  -8.250  -0.8585   0.06782   0.06378  -0.0006   1.0000   0.0730
  -8.000  -0.8963   0.06764   0.06282   0.0100   1.0000   0.0766
  -7.750  -0.9251   0.04906   0.04369   0.0174   1.0000   0.0500
  -7.500  -0.9283   0.04493   0.03922   0.0225   1.0000   0.0490
  -7.250  -0.9280   0.04125   0.03512   0.0274   1.0000   0.0490
  -7.000  -0.9237   0.03773   0.03114   0.0319   1.0000   0.0489
  -6.750  -0.9149   0.03455   0.02748   0.0358   1.0000   0.0489
  -6.500  -0.9014   0.03207   0.02454   0.0389   1.0000   0.0496
  -6.250  -0.8852   0.03071   0.02278   0.0416   1.0000   0.0514
  -6.000  -0.8674   0.02864   0.02031   0.0437   1.0000   0.0527
  -5.750  -0.8448   0.02640   0.01788   0.0447   1.0000   0.0542
  -5.500  -0.8231   0.02538   0.01679   0.0458   1.0000   0.0564
  -5.250  -0.8010   0.02454   0.01582   0.0469   1.0000   0.0590
  -5.000  -0.7772   0.02356   0.01467   0.0478   1.0000   0.0609
  -4.750  -0.7532   0.02273   0.01368   0.0488   1.0000   0.0627
  -4.500  -0.7283   0.02164   0.01250   0.0493   1.0000   0.0654
  -4.250  -0.7050   0.02083   0.01171   0.0501   1.0000   0.0684
  -4.000  -0.6823   0.02021   0.01108   0.0512   1.0000   0.0714
  -3.750  -0.6600   0.01967   0.01049   0.0523   1.0000   0.0747
  -3.500  -0.6387   0.01899   0.00983   0.0536   1.0000   0.0797
  -3.250  -0.6176   0.01855   0.00940   0.0549   1.0000   0.0866
  -3.000  -0.5973   0.01797   0.00889   0.0565   1.0000   0.0966
  -2.750  -0.5773   0.01737   0.00842   0.0580   1.0000   0.1213
  -2.500  -0.5588   0.01658   0.00807   0.0598   1.0000   0.1883
  -2.250  -0.4425   0.01633   0.01106   0.0420   1.0000   0.9079
  -2.000  -0.2317   0.02087   0.01521   0.0057   1.0000   0.9839
  -1.750  -0.1224   0.02110   0.01525  -0.0123   1.0000   1.0000
  -1.500  -0.1050   0.02102   0.01512  -0.0105   1.0000   1.0000
  -1.250  -0.0877   0.02096   0.01501  -0.0088   1.0000   1.0000
  -1.000  -0.0702   0.02090   0.01492  -0.0070   1.0000   1.0000
  -0.750  -0.0527   0.02086   0.01485  -0.0053   1.0000   1.0000
  -0.500  -0.0351   0.02083   0.01481  -0.0035   1.0000   1.0000
  -0.250  -0.0176   0.02082   0.01478  -0.0018   1.0000   1.0000
   0.000   0.0000   0.02081   0.01477   0.0000   1.0000   1.0000
   0.250   0.0176   0.02082   0.01477   0.0018   1.0000   1.0000
   0.500   0.0351   0.02083   0.01480   0.0035   1.0000   1.0000
   0.750   0.0527   0.02086   0.01485   0.0053   1.0000   1.0000
   1.000   0.0702   0.02090   0.01491   0.0070   1.0000   1.0000
   1.250   0.0877   0.02095   0.01500   0.0088   1.0000   1.0000
   1.500   0.1050   0.02101   0.01511   0.0105   1.0000   1.0000
   1.750   0.1224   0.02109   0.01524   0.0123   1.0000   1.0000
   2.000   0.2307   0.02087   0.01521  -0.0055   0.9841   1.0000
   2.250   0.4424   0.01632   0.01105  -0.0420   0.9079   1.0000
   2.500   0.5588   0.01658   0.00807  -0.0597   0.1886   1.0000
   2.750   0.5772   0.01736   0.00842  -0.0580   0.1213   1.0000
   3.000   0.5972   0.01797   0.00889  -0.0564   0.0966   1.0000
   3.250   0.6175   0.01854   0.00940  -0.0549   0.0866   1.0000
   3.500   0.6386   0.01898   0.00982  -0.0536   0.0797   1.0000
   3.750   0.6599   0.01967   0.01048  -0.0523   0.0747   1.0000
   4.000   0.6822   0.02021   0.01107  -0.0512   0.0714   1.0000
   4.250   0.7049   0.02082   0.01170  -0.0501   0.0684   1.0000
   4.500   0.7282   0.02164   0.01250  -0.0493   0.0653   1.0000
   4.750   0.7531   0.02273   0.01368  -0.0487   0.0627   1.0000
   5.000   0.7770   0.02355   0.01466  -0.0478   0.0609   1.0000
   5.250   0.8009   0.02453   0.01581  -0.0469   0.0590   1.0000
   5.500   0.8230   0.02538   0.01678  -0.0457   0.0564   1.0000
   5.750   0.8446   0.02639   0.01787  -0.0446   0.0542   1.0000
   6.000   0.8673   0.02863   0.02030  -0.0437   0.0527   1.0000
   6.250   0.8851   0.03073   0.02280  -0.0416   0.0514   1.0000
   6.500   0.9013   0.03206   0.02453  -0.0389   0.0497   1.0000
   6.750   0.9148   0.03454   0.02747  -0.0358   0.0489   1.0000
   7.000   0.9236   0.03773   0.03113  -0.0319   0.0489   1.0000
   7.250   0.9279   0.04124   0.03511  -0.0274   0.0490   1.0000
   7.500   0.9283   0.04493   0.03922  -0.0225   0.0490   1.0000
   7.750   0.9251   0.04906   0.04369  -0.0174   0.0500   1.0000
   8.000   0.8963   0.06763   0.06282  -0.0100   0.0766   1.0000
   8.250   0.8587   0.06781   0.06377   0.0005   0.0730   1.0000
   8.500   0.8439   0.07127   0.06741   0.0053   0.0706   1.0000
   8.750   0.8304   0.07448   0.07072   0.0096   0.0692   1.0000
   9.000   0.8165   0.07746   0.07377   0.0137   0.0681   1.0000
   9.250   0.7985   0.08014   0.07650   0.0183   0.0674   1.0000
   9.500   0.7723   0.08268   0.07907   0.0236   0.0671   1.0000
   9.750   0.7498   0.08572   0.08214   0.0272   0.0667   1.0000
  10.000   0.7321   0.08913   0.08556   0.0293   0.0662   1.0000
  10.250   0.7739   0.09381   0.09005   0.0299   0.0637   1.0000
  10.500   0.7458   0.09684   0.09314   0.0335   0.0637   1.0000
  10.750   0.7189   0.10085   0.09720   0.0347   0.0637   1.0000
<< Back to NACA 0010-66 (naca001066-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0010-66 (naca001066-il)