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=500,000 Ncrit=5


Details Polar file
Airfoil: NACA 0006 (naca0006-il)
Reynolds number: 500,000
Max Cl/Cd: 47.61 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca0006-il-500000-n5.txt
Download as CSV file: xf-naca0006-il-500000-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 =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.7553   0.08620   0.08416   0.0189   1.0000   0.0038
  -9.000  -0.7635   0.07934   0.07734   0.0133   1.0000   0.0038
  -8.750  -0.7723   0.07106   0.06905   0.0035   1.0000   0.0037
  -8.500  -0.7784   0.06178   0.05964  -0.0037   1.0000   0.0037
  -8.250  -0.7851   0.05186   0.04946  -0.0078   1.0000   0.0037
  -8.000  -0.8037   0.03568   0.03252  -0.0089   1.0000   0.0036
  -7.750  -0.7984   0.02789   0.02390  -0.0078   1.0000   0.0038
  -7.500  -0.7839   0.02371   0.01914  -0.0068   1.0000   0.0041
  -7.250  -0.7629   0.02183   0.01701  -0.0063   1.0000   0.0043
  -7.000  -0.7412   0.01988   0.01474  -0.0056   1.0000   0.0045
  -6.750  -0.7184   0.01823   0.01283  -0.0050   1.0000   0.0047
  -6.500  -0.6948   0.01680   0.01115  -0.0043   1.0000   0.0050
  -6.250  -0.6708   0.01555   0.00965  -0.0037   1.0000   0.0054
  -6.000  -0.6464   0.01447   0.00839  -0.0031   1.0000   0.0058
  -5.750  -0.6215   0.01362   0.00739  -0.0026   1.0000   0.0062
  -5.500  -0.5972   0.01261   0.00626  -0.0021   1.0000   0.0072
  -5.250  -0.5715   0.01215   0.00575  -0.0018   1.0000   0.0083
  -5.000  -0.5460   0.01154   0.00505  -0.0013   1.0000   0.0094
  -4.750  -0.5202   0.01105   0.00445  -0.0009   1.0000   0.0103
  -4.500  -0.4948   0.01039   0.00373  -0.0004   1.0000   0.0150
  -4.250  -0.4687   0.01007   0.00339  -0.0001   1.0000   0.0197
  -4.000  -0.4425   0.00977   0.00310   0.0002   1.0000   0.0277
  -3.750  -0.4162   0.00955   0.00282   0.0005   1.0000   0.0323
  -3.500  -0.3902   0.00928   0.00254   0.0008   1.0000   0.0373
  -3.250  -0.3641   0.00906   0.00229   0.0011   1.0000   0.0420
  -3.000  -0.3382   0.00881   0.00205   0.0015   1.0000   0.0477
  -2.750  -0.3124   0.00855   0.00182   0.0019   1.0000   0.0605
  -2.250  -0.2620   0.00777   0.00145   0.0026   1.0000   0.1646
  -2.000  -0.2368   0.00741   0.00131   0.0030   1.0000   0.2270
  -1.750  -0.2116   0.00703   0.00118   0.0033   1.0000   0.2997
  -1.250  -0.1501   0.00622   0.00102   0.0014   0.9927   0.4780
  -1.000  -0.1164   0.00586   0.00096  -0.0001   0.9846   0.5625
  -0.750  -0.0847   0.00541   0.00094  -0.0010   0.9737   0.6727
  -0.500  -0.0539   0.00509   0.00094  -0.0016   0.9591   0.7544
  -0.250  -0.0256   0.00481   0.00097  -0.0012   0.9394   0.8430
   0.000   0.0000   0.00470   0.00101   0.0000   0.9073   0.9077
   0.250   0.0257   0.00480   0.00097   0.0012   0.8434   0.9394
   0.500   0.0539   0.00509   0.00094   0.0016   0.7544   0.9591
   0.750   0.0847   0.00541   0.00094   0.0010   0.6726   0.9738
   1.000   0.1164   0.00586   0.00096   0.0001   0.5627   0.9846
   1.250   0.1500   0.00622   0.00102  -0.0014   0.4786   0.9928
   1.500   0.1841   0.00667   0.00110  -0.0031   0.3761   0.9987
   1.750   0.2115   0.00703   0.00118  -0.0033   0.2996   1.0000
   2.000   0.2367   0.00741   0.00131  -0.0029   0.2272   1.0000
   2.250   0.2618   0.00777   0.00145  -0.0026   0.1651   1.0000
   2.750   0.3122   0.00855   0.00182  -0.0018   0.0607   1.0000
   3.000   0.3380   0.00880   0.00205  -0.0015   0.0478   1.0000
   3.250   0.3640   0.00906   0.00229  -0.0011   0.0420   1.0000
   3.500   0.3901   0.00928   0.00254  -0.0008   0.0373   1.0000
   3.750   0.4161   0.00955   0.00282  -0.0004   0.0323   1.0000
   4.000   0.4424   0.00977   0.00309  -0.0002   0.0277   1.0000
   4.250   0.4686   0.01007   0.00339   0.0001   0.0196   1.0000
   4.500   0.4947   0.01039   0.00373   0.0005   0.0150   1.0000
   4.750   0.5202   0.01105   0.00445   0.0009   0.0103   1.0000
   5.000   0.5460   0.01154   0.00505   0.0013   0.0094   1.0000
   5.250   0.5715   0.01214   0.00575   0.0018   0.0083   1.0000
   5.500   0.5973   0.01260   0.00624   0.0021   0.0072   1.0000
   5.750   0.6216   0.01361   0.00738   0.0026   0.0062   1.0000
   6.000   0.6465   0.01447   0.00839   0.0031   0.0058   1.0000
   6.250   0.6709   0.01554   0.00964   0.0037   0.0054   1.0000
   6.500   0.6949   0.01681   0.01117   0.0043   0.0050   1.0000
   6.750   0.7185   0.01824   0.01284   0.0049   0.0047   1.0000
   7.000   0.7414   0.01988   0.01474   0.0056   0.0045   1.0000
   7.250   0.7631   0.02184   0.01701   0.0062   0.0043   1.0000
   7.500   0.7841   0.02372   0.01916   0.0068   0.0041   1.0000
   7.750   0.7986   0.02793   0.02395   0.0077   0.0038   1.0000
   8.000   0.8037   0.03582   0.03267   0.0089   0.0036   1.0000
   8.250   0.7853   0.05197   0.04957   0.0077   0.0037   1.0000
   8.500   0.7787   0.06187   0.05972   0.0036   0.0037   1.0000
   8.750   0.7726   0.07120   0.06919  -0.0038   0.0037   1.0000
   9.000   0.7641   0.07942   0.07742  -0.0135   0.0038   1.0000
   9.250   0.7557   0.08640   0.08435  -0.0192   0.0038   1.0000
<< Back to NACA 0006 (naca0006-il)

Polar data table (+)

Polar graphs


<< Back to NACA 0006 (naca0006-il)