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

NACA 65-206 (naca65206-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA 65-206 (naca65206-il)
Reynolds number: 50,000
Max Cl/Cd: 29.4 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca65206-il-50000-n5.txt
Download as CSV file: xf-naca65206-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 65-206                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5608   0.09441   0.08774  -0.0112   1.0000   0.1053
  -7.750  -0.5583   0.09086   0.08425  -0.0127   1.0000   0.1092
  -7.500  -0.5731   0.08792   0.08130  -0.0238   1.0000   0.1151
  -6.500  -0.5211   0.06675   0.05975  -0.0317   1.0000   0.0591
  -6.250  -0.5099   0.06232   0.05526  -0.0326   1.0000   0.0566
  -6.000  -0.4912   0.05797   0.05030  -0.0352   1.0000   0.0471
  -5.750  -0.4769   0.05363   0.04583  -0.0355   1.0000   0.0463
  -5.500  -0.4611   0.04958   0.04159  -0.0358   1.0000   0.0450
  -5.250  -0.4431   0.04567   0.03739  -0.0361   1.0000   0.0431
  -5.000  -0.4229   0.04196   0.03329  -0.0361   1.0000   0.0414
  -4.750  -0.4009   0.03850   0.02936  -0.0359   1.0000   0.0401
  -4.500  -0.3777   0.03532   0.02567  -0.0354   1.0000   0.0395
  -4.250  -0.3537   0.03246   0.02233  -0.0347   1.0000   0.0396
  -4.000  -0.3293   0.02995   0.01937  -0.0339   1.0000   0.0408
  -3.750  -0.3038   0.02830   0.01708  -0.0329   1.0000   0.0472
  -3.500  -0.2801   0.02582   0.01444  -0.0320   1.0000   0.0503
  -3.250  -0.2559   0.02397   0.01226  -0.0305   1.0000   0.0526
  -3.000  -0.2326   0.02248   0.01057  -0.0289   1.0000   0.0558
  -2.750  -0.2096   0.02131   0.00911  -0.0274   1.0000   0.0598
  -2.500  -0.1875   0.02022   0.00790  -0.0265   1.0000   0.0732
  -2.250  -0.1648   0.01914   0.00675  -0.0258   1.0000   0.0899
  -2.000  -0.1409   0.01785   0.00564  -0.0254   1.0000   0.1310
  -1.750  -0.1078   0.01461   0.00543  -0.0223   1.0000   1.0000
  -1.500  -0.0913   0.01454   0.00488  -0.0207   1.0000   1.0000
  -1.250  -0.0729   0.01452   0.00453  -0.0195   1.0000   1.0000
  -1.000  -0.0533   0.01453   0.00427  -0.0186   1.0000   1.0000
  -0.750  -0.0328   0.01457   0.00407  -0.0178   1.0000   1.0000
  -0.500  -0.0118   0.01465   0.00395  -0.0172   1.0000   1.0000
  -0.250   0.0095   0.01475   0.00389  -0.0166   1.0000   1.0000
   0.000   0.0311   0.01488   0.00385  -0.0161   1.0000   1.0000
   0.250   0.0527   0.01504   0.00390  -0.0156   1.0000   1.0000
   0.500   0.0743   0.01522   0.00401  -0.0151   1.0000   1.0000
   0.750   0.0958   0.01544   0.00417  -0.0147   1.0000   1.0000
   1.000   0.1173   0.01568   0.00440  -0.0143   1.0000   1.0000
   1.250   0.1387   0.01595   0.00470  -0.0140   1.0000   1.0000
   1.500   0.1642   0.01629   0.00507  -0.0145   0.9973   1.0000
   1.750   0.2022   0.01673   0.00562  -0.0174   0.9879   1.0000
   2.000   0.2403   0.01717   0.00621  -0.0203   0.9783   1.0000
   2.250   0.2779   0.01760   0.00685  -0.0231   0.9680   1.0000
   2.500   0.3151   0.01802   0.00763  -0.0256   0.9570   1.0000
   2.750   0.3523   0.01844   0.00837  -0.0281   0.9452   1.0000
   3.000   0.3909   0.01884   0.00921  -0.0306   0.9325   1.0000
   3.250   0.4830   0.01643   0.00684  -0.0298   0.5529   1.0000
   3.500   0.4840   0.02040   0.00778  -0.0259   0.0993   1.0000
   3.750   0.5062   0.02187   0.00921  -0.0251   0.0743   1.0000
   4.000   0.5288   0.02330   0.01079  -0.0242   0.0657   1.0000
   4.250   0.5534   0.02480   0.01258  -0.0231   0.0609   1.0000
   4.500   0.5813   0.02658   0.01460  -0.0221   0.0567   1.0000
   4.750   0.6089   0.02860   0.01680  -0.0216   0.0477   1.0000
   5.000   0.6370   0.03086   0.01935  -0.0209   0.0429   1.0000
   5.250   0.6641   0.03372   0.02272  -0.0202   0.0410   1.0000
   5.500   0.6890   0.03706   0.02645  -0.0195   0.0399   1.0000
   5.750   0.7114   0.04078   0.03064  -0.0185   0.0394   1.0000
   6.000   0.7309   0.04438   0.03481  -0.0174   0.0380   1.0000
   6.250   0.7493   0.04757   0.03873  -0.0161   0.0355   1.0000
   6.500   0.7644   0.05136   0.04308  -0.0152   0.0337   1.0000
   6.750   0.7769   0.05560   0.04775  -0.0144   0.0339   1.0000
   7.000   0.7868   0.05999   0.05252  -0.0140   0.0345   1.0000
   7.250   0.7938   0.06456   0.05740  -0.0137   0.0351   1.0000
   7.500   0.7984   0.06913   0.06221  -0.0138   0.0359   1.0000
   7.750   0.8011   0.07373   0.06698  -0.0140   0.0367   1.0000
   8.000   0.8024   0.07856   0.07191  -0.0144   0.0376   1.0000
   8.250   0.8010   0.08298   0.07659  -0.0156   0.0410   1.0000
   8.500   0.7895   0.08813   0.08185  -0.0182   0.0431   1.0000
   8.750   0.7807   0.09328   0.08700  -0.0213   0.0447   1.0000
   9.000   0.7754   0.09891   0.09260  -0.0254   0.0466   1.0000
<< Back to NACA 65-206 (naca65206-il)

Polar data table (+)

Polar graphs


<< Back to NACA 65-206 (naca65206-il)