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

NACA 0006 (naca0006-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: NACA 0006 (naca0006-il)
Reynolds number: 1,000,000
Max Cl/Cd: 60.29 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca0006-il-1000000-n5.txt
Download as CSV file: xf-naca0006-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 0006                                       
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.7609   0.10411   0.10261   0.0308   1.0000   0.0022
 -10.000  -0.7705   0.09706   0.09558   0.0271   1.0000   0.0022
  -9.000  -0.9288   0.02619   0.02303  -0.0075   1.0000   0.0020
  -8.750  -0.9135   0.02296   0.01936  -0.0067   1.0000   0.0021
  -8.500  -0.8941   0.02079   0.01687  -0.0060   1.0000   0.0022
  -8.250  -0.8726   0.01917   0.01498  -0.0054   1.0000   0.0022
  -8.000  -0.8500   0.01785   0.01344  -0.0049   1.0000   0.0023
  -7.750  -0.8291   0.01587   0.01114  -0.0041   1.0000   0.0024
  -7.500  -0.8054   0.01479   0.00990  -0.0036   1.0000   0.0026
  -7.250  -0.7804   0.01415   0.00918  -0.0033   1.0000   0.0027
  -7.000  -0.7550   0.01361   0.00857  -0.0031   1.0000   0.0030
  -6.750  -0.7297   0.01298   0.00785  -0.0027   1.0000   0.0032
  -6.500  -0.7045   0.01230   0.00705  -0.0024   1.0000   0.0035
  -6.250  -0.6791   0.01166   0.00628  -0.0020   1.0000   0.0037
  -6.000  -0.6533   0.01116   0.00569  -0.0017   1.0000   0.0039
  -5.750  -0.6281   0.01043   0.00485  -0.0012   1.0000   0.0044
  -5.500  -0.6022   0.00999   0.00438  -0.0009   1.0000   0.0049
  -5.250  -0.5760   0.00962   0.00396  -0.0007   1.0000   0.0055
  -5.000  -0.5497   0.00930   0.00359  -0.0004   1.0000   0.0061
  -4.750  -0.5235   0.00891   0.00315  -0.0001   1.0000   0.0076
  -4.500  -0.4973   0.00860   0.00283   0.0002   1.0000   0.0093
  -4.250  -0.4709   0.00834   0.00254   0.0005   1.0000   0.0119
  -4.000  -0.4446   0.00808   0.00232   0.0008   1.0000   0.0185
  -3.750  -0.4184   0.00786   0.00213   0.0011   1.0000   0.0267
  -3.500  -0.3921   0.00773   0.00198   0.0014   1.0000   0.0294
  -3.250  -0.3659   0.00760   0.00184   0.0017   1.0000   0.0309
  -3.000  -0.3400   0.00741   0.00165   0.0021   1.0000   0.0349
  -2.750  -0.3124   0.00725   0.00149   0.0021   0.9991   0.0380
  -2.500  -0.2787   0.00710   0.00134   0.0008   0.9953   0.0416
  -2.250  -0.2462   0.00684   0.00120  -0.0004   0.9895   0.0674
  -2.000  -0.2136   0.00655   0.00105  -0.0016   0.9818   0.1102
  -1.750  -0.1814   0.00624   0.00093  -0.0027   0.9708   0.1646
  -1.500  -0.1505   0.00598   0.00083  -0.0035   0.9547   0.2154
  -1.250  -0.1230   0.00569   0.00073  -0.0034   0.9315   0.2835
  -1.000  -0.0978   0.00548   0.00066  -0.0028   0.9011   0.3447
  -0.750  -0.0735   0.00538   0.00059  -0.0019   0.8552   0.3997
  -0.500  -0.0501   0.00537   0.00052  -0.0009   0.7729   0.4727
  -0.250  -0.0256   0.00534   0.00049  -0.0004   0.6897   0.5579
   0.000   0.0000   0.00531   0.00048   0.0000   0.6269   0.6278
   0.250   0.0258   0.00534   0.00049   0.0003   0.5580   0.6890
   0.500   0.0503   0.00538   0.00052   0.0009   0.4723   0.7716
   0.750   0.0737   0.00539   0.00059   0.0019   0.3987   0.8537
   1.000   0.0979   0.00549   0.00066   0.0028   0.3426   0.9009
   1.250   0.1232   0.00568   0.00073   0.0034   0.2847   0.9315
   1.500   0.1506   0.00598   0.00083   0.0034   0.2152   0.9547
   1.750   0.1815   0.00624   0.00093   0.0027   0.1641   0.9708
   2.000   0.2136   0.00655   0.00105   0.0016   0.1104   0.9819
   2.250   0.2462   0.00684   0.00119   0.0004   0.0681   0.9897
   2.500   0.2786   0.00710   0.00134  -0.0007   0.0416   0.9955
   2.750   0.3123   0.00725   0.00149  -0.0021   0.0380   0.9992
   3.000   0.3398   0.00741   0.00165  -0.0021   0.0350   1.0000
   3.250   0.3657   0.00760   0.00184  -0.0017   0.0309   1.0000
   3.500   0.3919   0.00773   0.00198  -0.0013   0.0294   1.0000
   3.750   0.4182   0.00786   0.00213  -0.0010   0.0267   1.0000
   4.000   0.4445   0.00808   0.00232  -0.0008   0.0186   1.0000
   4.250   0.4707   0.00834   0.00254  -0.0005   0.0118   1.0000
   4.500   0.4971   0.00860   0.00283  -0.0002   0.0093   1.0000
   4.750   0.5234   0.00891   0.00315   0.0001   0.0076   1.0000
   5.000   0.5496   0.00931   0.00360   0.0004   0.0061   1.0000
   5.250   0.5760   0.00962   0.00396   0.0007   0.0055   1.0000
   5.500   0.6022   0.00999   0.00438   0.0009   0.0049   1.0000
   5.750   0.6282   0.01042   0.00485   0.0012   0.0044   1.0000
   6.000   0.6534   0.01115   0.00568   0.0016   0.0039   1.0000
   6.250   0.6792   0.01166   0.00628   0.0019   0.0037   1.0000
   6.500   0.7047   0.01231   0.00706   0.0023   0.0035   1.0000
   6.750   0.7299   0.01299   0.00786   0.0027   0.0032   1.0000
   7.000   0.7553   0.01359   0.00855   0.0030   0.0030   1.0000
   7.250   0.7807   0.01414   0.00916   0.0033   0.0027   1.0000
   7.500   0.8057   0.01480   0.00990   0.0035   0.0026   1.0000
   7.750   0.8293   0.01588   0.01115   0.0040   0.0024   1.0000
   8.000   0.8503   0.01787   0.01346   0.0048   0.0023   1.0000
   8.250   0.8730   0.01919   0.01500   0.0053   0.0022   1.0000
   8.500   0.8945   0.02082   0.01690   0.0059   0.0022   1.0000
   8.750   0.9139   0.02301   0.01941   0.0066   0.0021   1.0000
   9.000   0.9294   0.02621   0.02305   0.0074   0.0020   1.0000
  10.000   0.7711   0.09730   0.09583  -0.0275   0.0022   1.0000
<< Back to NACA 0006 (naca0006-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0006 (naca0006-il)