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

NACA 66-209 (naca66209-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-209 (naca66209-il)
Reynolds number: 50,000
Max Cl/Cd: 22.09 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66209-il-50000.txt
Download as CSV file: xf-naca66209-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 66-209                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.5087   0.12186   0.11466  -0.0156   1.0000   0.2205
 -10.250  -0.5244   0.12067   0.11358  -0.0176   1.0000   0.2293
 -10.000  -0.5176   0.11679   0.10973  -0.0172   1.0000   0.2431
  -9.750  -0.5118   0.11323   0.10621  -0.0169   1.0000   0.2569
  -9.500  -0.5108   0.11012   0.10315  -0.0167   1.0000   0.2712
  -9.250  -0.4974   0.10575   0.09879  -0.0155   1.0000   0.2884
  -9.000  -0.4884   0.10244   0.09550  -0.0142   1.0000   0.3103
  -8.750  -0.4800   0.09910   0.09215  -0.0127   1.0000   0.3345
  -8.500  -0.4728   0.09615   0.08924  -0.0110   1.0000   0.3622
  -8.250  -0.4651   0.09325   0.08638  -0.0088   1.0000   0.3940
  -8.000  -0.4623   0.09082   0.08400  -0.0066   1.0000   0.4248
  -7.750  -0.4510   0.08754   0.08075  -0.0044   1.0000   0.4568
  -7.500  -0.4305   0.08403   0.07723  -0.0025   1.0000   0.4912
  -7.000  -0.4036   0.07793   0.07116   0.0013   1.0000   0.5578
  -6.250  -0.3773   0.05797   0.05212  -0.0024   1.0000   0.4885
  -6.000  -0.4466   0.05526   0.04970  -0.0005   1.0000   0.4304
  -5.750  -0.4917   0.05244   0.04707   0.0020   1.0000   0.4149
  -5.250  -0.5742   0.04760   0.04004  -0.0204   1.0000   0.1941
  -5.000  -0.5514   0.04333   0.03481  -0.0204   1.0000   0.1551
  -4.750  -0.5302   0.03981   0.03064  -0.0192   1.0000   0.1377
  -4.500  -0.5079   0.03701   0.02720  -0.0178   1.0000   0.1273
  -4.250  -0.4846   0.03439   0.02394  -0.0165   1.0000   0.1204
  -4.000  -0.4612   0.03225   0.02134  -0.0152   1.0000   0.1200
  -3.750  -0.4383   0.03031   0.01921  -0.0141   1.0000   0.1244
  -3.500  -0.4132   0.02859   0.01717  -0.0129   1.0000   0.1257
  -3.250  -0.3869   0.02705   0.01534  -0.0116   1.0000   0.1282
  -3.000  -0.1380   0.02162   0.01272  -0.0352   1.0000   1.0000
  -2.750  -0.1339   0.02148   0.01224  -0.0320   1.0000   1.0000
  -2.500  -0.1299   0.02135   0.01188  -0.0288   1.0000   1.0000
  -2.250  -0.1259   0.02122   0.01154  -0.0256   1.0000   1.0000
  -2.000  -0.1217   0.02110   0.01124  -0.0224   1.0000   1.0000
  -1.750  -0.1174   0.02098   0.01096  -0.0192   1.0000   1.0000
  -1.500  -0.1129   0.02087   0.01069  -0.0161   1.0000   1.0000
  -1.250  -0.1083   0.02075   0.01043  -0.0129   1.0000   1.0000
  -1.000  -0.1034   0.02063   0.01016  -0.0098   1.0000   1.0000
  -0.750  -0.0981   0.02052   0.00992  -0.0067   1.0000   1.0000
  -0.500  -0.0922   0.02042   0.00970  -0.0038   1.0000   1.0000
  -0.250  -0.0851   0.02033   0.00951  -0.0010   1.0000   1.0000
   0.000  -0.0765   0.02028   0.00935   0.0016   1.0000   1.0000
   0.250  -0.0654   0.02029   0.00925   0.0036   1.0000   1.0000
   0.500  -0.0517   0.02035   0.00921   0.0052   1.0000   1.0000
   0.750  -0.0360   0.02047   0.00922   0.0065   1.0000   1.0000
   1.000  -0.0188   0.02064   0.00932   0.0074   1.0000   1.0000
   1.250  -0.0007   0.02086   0.00948   0.0082   1.0000   1.0000
   1.500   0.0181   0.02112   0.00970   0.0088   1.0000   1.0000
   1.750   0.0373   0.02142   0.00998   0.0093   1.0000   1.0000
   2.000   0.0568   0.02175   0.01031   0.0098   1.0000   1.0000
   2.250   0.0765   0.02213   0.01070   0.0102   1.0000   1.0000
   2.500   0.0961   0.02255   0.01115   0.0105   1.0000   1.0000
   2.750   0.1157   0.02300   0.01166   0.0108   1.0000   1.0000
   3.000   0.1352   0.02350   0.01223   0.0110   1.0000   1.0000
   3.250   0.1546   0.02405   0.01290   0.0112   1.0000   1.0000
   3.500   0.1736   0.02465   0.01361   0.0113   1.0000   1.0000
   3.750   0.1923   0.02530   0.01439   0.0114   1.0000   1.0000
   4.000   0.2107   0.02602   0.01526   0.0115   1.0000   1.0000
   4.250   0.2284   0.02681   0.01622   0.0115   1.0000   1.0000
   4.500   0.2458   0.02768   0.01733   0.0115   1.0000   1.0000
   4.750   0.2624   0.02864   0.01850   0.0114   1.0000   1.0000
   5.000   0.3230   0.03097   0.02137   0.0027   0.9747   1.0000
   5.250   0.6248   0.02828   0.01650  -0.0123   0.1274   1.0000
   5.500   0.6663   0.03062   0.01893  -0.0136   0.1162   1.0000
   5.750   0.7016   0.03318   0.02179  -0.0138   0.1136   1.0000
   6.000   0.7311   0.03594   0.02497  -0.0131   0.1139   1.0000
   6.250   0.7561   0.03887   0.02838  -0.0120   0.1149   1.0000
   6.500   0.7765   0.04181   0.03176  -0.0104   0.1139   1.0000
   6.750   0.7938   0.04509   0.03562  -0.0086   0.1157   1.0000
   7.000   0.8132   0.04936   0.04013  -0.0075   0.1210   1.0000
   7.250   0.8241   0.05329   0.04486  -0.0051   0.1333   1.0000
   7.500   0.8305   0.05785   0.05023  -0.0030   0.1540   1.0000
   8.000   0.8059   0.07205   0.06572  -0.0061   0.2492   1.0000
   8.250   0.7575   0.07964   0.07339  -0.0114   0.2909   1.0000
<< Back to NACA 66-209 (naca66209-il)

Polar data table (+)

Polar graphs


<< Back to NACA 66-209 (naca66209-il)