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NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: NACA 66-206 (naca66206-il)
Reynolds number: 200,000
Max Cl/Cd: 50.68 at α=2°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-naca66206-il-200000-n5.txt
Download as CSV file: xf-naca66206-il-200000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5140   0.08558   0.08206  -0.0244   1.0000   0.0107
  -7.750  -0.5142   0.08151   0.07800  -0.0275   1.0000   0.0107
  -7.000  -0.5149   0.06750   0.06396  -0.0323   1.0000   0.0097
  -6.750  -0.5123   0.06334   0.05975  -0.0333   1.0000   0.0092
  -6.500  -0.5073   0.05914   0.05548  -0.0340   1.0000   0.0088
  -6.250  -0.5007   0.05521   0.05145  -0.0342   1.0000   0.0084
  -6.000  -0.4926   0.05129   0.04739  -0.0340   1.0000   0.0081
  -5.750  -0.4837   0.04754   0.04348  -0.0332   1.0000   0.0079
  -5.500  -0.4737   0.04385   0.03959  -0.0320   1.0000   0.0077
  -5.250  -0.4622   0.04033   0.03585  -0.0305   1.0000   0.0076
  -5.000  -0.4397   0.03653   0.03172  -0.0307   0.9982   0.0084
  -4.250  -0.3435   0.02703   0.02097  -0.0337   0.9881   0.0107
  -4.000  -0.3135   0.02372   0.01722  -0.0343   0.9848   0.0108
  -3.750  -0.2836   0.02028   0.01324  -0.0348   0.9816   0.0110
  -3.500  -0.2531   0.01719   0.00972  -0.0354   0.9794   0.0124
  -3.250  -0.2214   0.01637   0.00869  -0.0365   0.9767   0.0172
  -3.000  -0.1921   0.01492   0.00704  -0.0365   0.9732   0.0190
  -2.750  -0.1610   0.01392   0.00590  -0.0371   0.9703   0.0215
  -2.500  -0.1296   0.01287   0.00482  -0.0383   0.9680   0.0268
  -2.250  -0.0998   0.01219   0.00404  -0.0389   0.9642   0.0294
  -2.000  -0.0694   0.01174   0.00344  -0.0396   0.9605   0.0345
  -1.750  -0.0376   0.01128   0.00291  -0.0407   0.9575   0.0513
  -1.500  -0.0237   0.00855   0.00294  -0.0388   0.9538   0.7640
  -1.250  -0.0077   0.00852   0.00315  -0.0350   0.9478   0.8692
  -1.000   0.0136   0.00858   0.00323  -0.0324   0.9440   0.9243
  -0.750   0.0486   0.00861   0.00319  -0.0333   0.9428   0.9565
  -0.500   0.0846   0.00859   0.00307  -0.0352   0.9400   0.9625
  -0.250   0.1178   0.00858   0.00300  -0.0367   0.9361   0.9681
   0.000   0.1542   0.00856   0.00295  -0.0387   0.9332   0.9720
   0.250   0.1900   0.00854   0.00292  -0.0407   0.9305   0.9767
   0.500   0.2262   0.00853   0.00293  -0.0428   0.9276   0.9806
   0.750   0.2600   0.00853   0.00297  -0.0444   0.9229   0.9860
   1.000   0.2964   0.00849   0.00298  -0.0463   0.9180   0.9898
   1.250   0.3303   0.00840   0.00296  -0.0475   0.9076   0.9946
   1.500   0.3613   0.00816   0.00280  -0.0475   0.8830   0.9997
   1.750   0.3828   0.00804   0.00271  -0.0457   0.8574   1.0000
   2.000   0.4019   0.00793   0.00251  -0.0430   0.8077   1.0000
   2.250   0.4087   0.00863   0.00209  -0.0374   0.5437   1.0000
   2.500   0.4100   0.01110   0.00267  -0.0335   0.1231   1.0000
   2.750   0.4307   0.01196   0.00317  -0.0325   0.0450   1.0000
   3.000   0.4544   0.01250   0.00379  -0.0318   0.0362   1.0000
   3.250   0.4781   0.01308   0.00452  -0.0312   0.0299   1.0000
   3.500   0.5009   0.01385   0.00537  -0.0304   0.0253   1.0000
   3.750   0.5241   0.01470   0.00633  -0.0296   0.0228   1.0000
   4.000   0.5471   0.01585   0.00758  -0.0286   0.0207   1.0000
   4.250   0.5713   0.01692   0.00873  -0.0279   0.0177   1.0000
   4.500   0.5947   0.01915   0.01108  -0.0272   0.0144   1.0000
   4.750   0.6209   0.02076   0.01296  -0.0264   0.0132   1.0000
   5.000   0.6469   0.02280   0.01537  -0.0255   0.0114   1.0000
   5.250   0.6715   0.02412   0.01703  -0.0249   0.0086   1.0000
   5.500   0.6944   0.02665   0.01995  -0.0237   0.0079   1.0000
   5.750   0.7154   0.02974   0.02347  -0.0224   0.0076   1.0000
   6.000   0.7342   0.03334   0.02752  -0.0208   0.0075   1.0000
   6.250   0.7511   0.03743   0.03205  -0.0190   0.0075   1.0000
   6.500   0.7658   0.04228   0.03735  -0.0170   0.0079   1.0000
   6.750   0.7778   0.04755   0.04303  -0.0152   0.0084   1.0000
   7.000   0.7871   0.05273   0.04855  -0.0138   0.0089   1.0000
   7.250   0.7938   0.05781   0.05390  -0.0128   0.0095   1.0000
   7.500   0.7975   0.06283   0.05914  -0.0123   0.0100   1.0000
   7.750   0.7982   0.06775   0.06424  -0.0122   0.0106   1.0000
   8.000   0.7961   0.07247   0.06909  -0.0126   0.0111   1.0000
   8.250   0.7907   0.07716   0.07387  -0.0134   0.0115   1.0000
   8.500   0.7799   0.08157   0.07833  -0.0142   0.0119   1.0000
   8.750   0.7702   0.08642   0.08321  -0.0170   0.0120   1.0000
   9.000   0.7624   0.09226   0.08904  -0.0220   0.0123   1.0000
   9.250   0.7575   0.09782   0.09457  -0.0257   0.0128   1.0000
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