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

NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-206 (naca66206-il)
Reynolds number: 200,000
Max Cl/Cd: 52.92 at α=2.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66206-il-200000.txt
Download as CSV file: xf-naca66206-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-206                                     
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5224   0.08643   0.08289  -0.0303   1.0000   0.0211
  -7.750  -0.5264   0.08032   0.07681  -0.0324   1.0000   0.0215
  -7.500  -0.5286   0.07492   0.07142  -0.0334   1.0000   0.0220
  -7.250  -0.5271   0.07094   0.06746  -0.0328   1.0000   0.0228
  -7.000  -0.5229   0.06736   0.06386  -0.0333   1.0000   0.0235
  -6.750  -0.5181   0.06377   0.06023  -0.0337   1.0000   0.0244
  -6.500  -0.5121   0.06021   0.05661  -0.0341   1.0000   0.0253
  -6.250  -0.5051   0.05669   0.05300  -0.0342   1.0000   0.0269
  -6.000  -0.4950   0.05344   0.04960  -0.0341   1.0000   0.0307
  -5.750  -0.4746   0.05318   0.04881  -0.0330   1.0000   0.0337
  -5.500  -0.4713   0.04785   0.04328  -0.0319   1.0000   0.0344
  -5.250  -0.4665   0.04300   0.03853  -0.0309   1.0000   0.0361
  -5.000  -0.4563   0.04027   0.03572  -0.0295   1.0000   0.0381
  -4.750  -0.4429   0.03772   0.03299  -0.0282   1.0000   0.0410
  -4.500  -0.4204   0.03846   0.03305  -0.0257   1.0000   0.0468
  -4.250  -0.4093   0.03232   0.02688  -0.0255   1.0000   0.0493
  -4.000  -0.3926   0.02997   0.02445  -0.0246   1.0000   0.0534
  -3.750  -0.3727   0.02785   0.02186  -0.0235   1.0000   0.0628
  -3.500  -0.3530   0.02607   0.01982  -0.0227   1.0000   0.0757
  -3.250  -0.3330   0.02409   0.01773  -0.0223   1.0000   0.0903
  -3.000  -0.2989   0.02214   0.01498  -0.0191   1.0000   0.0465
  -2.750  -0.2712   0.01842   0.01096  -0.0185   1.0000   0.0390
  -2.500  -0.2459   0.01708   0.00941  -0.0177   1.0000   0.0410
  -2.250  -0.2198   0.01558   0.00768  -0.0167   1.0000   0.0381
  -2.000  -0.1955   0.01447   0.00649  -0.0158   1.0000   0.0388
  -1.750  -0.1719   0.01364   0.00564  -0.0151   1.0000   0.0423
  -1.500  -0.1486   0.01296   0.00489  -0.0145   1.0000   0.0507
  -1.250  -0.1184   0.01245   0.00431  -0.0153   0.9983   0.0589
  -1.000  -0.1034   0.00943   0.00424  -0.0130   0.9972   0.8060
  -0.750  -0.0664   0.00931   0.00426  -0.0133   0.9955   1.0000
  -0.500  -0.0316   0.00948   0.00426  -0.0153   0.9919   1.0000
  -0.250   0.0024   0.00963   0.00426  -0.0172   0.9878   1.0000
   0.000   0.0390   0.00984   0.00437  -0.0196   0.9840   1.0000
   0.250   0.0763   0.01005   0.00452  -0.0221   0.9802   1.0000
   0.500   0.1097   0.01018   0.00461  -0.0237   0.9746   1.0000
   0.750   0.1478   0.01036   0.00479  -0.0263   0.9705   1.0000
   1.000   0.1834   0.01053   0.00499  -0.0284   0.9660   1.0000
   1.250   0.2173   0.01066   0.00515  -0.0301   0.9600   1.0000
   1.500   0.2599   0.01079   0.00536  -0.0334   0.9559   1.0000
   1.750   0.3097   0.01062   0.00532  -0.0376   0.9448   1.0000
   2.000   0.4086   0.00900   0.00407  -0.0486   0.9105   1.0000
   2.250   0.4297   0.00812   0.00312  -0.0439   0.8375   1.0000
   2.500   0.4270   0.00923   0.00258  -0.0355   0.4469   1.0000
   2.750   0.4298   0.01209   0.00359  -0.0319   0.0602   1.0000
   3.000   0.4516   0.01295   0.00450  -0.0307   0.0466   1.0000
   3.250   0.4735   0.01389   0.00545  -0.0296   0.0413   1.0000
   3.500   0.4948   0.01543   0.00696  -0.0283   0.0388   1.0000
   3.750   0.5197   0.01714   0.00876  -0.0275   0.0385   1.0000
   4.000   0.5462   0.01811   0.00992  -0.0269   0.0339   1.0000
   4.250   0.5737   0.02015   0.01220  -0.0260   0.0337   1.0000
   4.500   0.6007   0.02332   0.01565  -0.0251   0.0365   1.0000
   5.750   0.7172   0.03826   0.03274  -0.0174   0.0435   1.0000
   6.000   0.7341   0.04143   0.03602  -0.0167   0.0402   1.0000
   6.250   0.7400   0.05006   0.04466  -0.0167   0.0379   1.0000
   6.500   0.7590   0.04965   0.04505  -0.0139   0.0355   1.0000
   6.750   0.7728   0.05314   0.04889  -0.0128   0.0329   1.0000
   7.000   0.7826   0.05710   0.05309  -0.0120   0.0310   1.0000
   7.250   0.7901   0.06107   0.05726  -0.0115   0.0298   1.0000
   7.500   0.7952   0.06507   0.06141  -0.0111   0.0288   1.0000
   7.750   0.7982   0.06906   0.06550  -0.0109   0.0278   1.0000
   8.000   0.7995   0.07306   0.06960  -0.0108   0.0273   1.0000
   8.250   0.7993   0.07719   0.07378  -0.0107   0.0267   1.0000
   8.500   0.7984   0.08208   0.07868  -0.0107   0.0260   1.0000
   8.750   0.7866   0.08810   0.08474  -0.0113   0.0256   1.0000
   9.000   0.7746   0.09305   0.08973  -0.0118   0.0254   1.0000
<< Back to NACA 66-206 (naca66206-il)

Polar data table (+)

Polar graphs


<< Back to NACA 66-206 (naca66206-il)