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NACA 66-206 (naca66206-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA 66-206 (naca66206-il)
Reynolds number: 500,000
Max Cl/Cd: 71.69 at α=2°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca66206-il-500000.txt
Download as CSV file: xf-naca66206-il-500000.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 =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5189   0.08578   0.08350  -0.0200   1.0000   0.0102
  -8.000  -0.5182   0.08224   0.08000  -0.0230   1.0000   0.0105
  -7.750  -0.5225   0.07841   0.07620  -0.0260   1.0000   0.0105
  -7.500  -0.5190   0.07432   0.07210  -0.0295   1.0000   0.0106
  -7.250  -0.5121   0.07053   0.06827  -0.0320   1.0000   0.0108
  -7.000  -0.5058   0.06673   0.06442  -0.0335   1.0000   0.0109
  -6.750  -0.4991   0.06293   0.06056  -0.0343   1.0000   0.0109
  -6.500  -0.4924   0.05934   0.05688  -0.0344   1.0000   0.0110
  -6.250  -0.4867   0.05591   0.05336  -0.0336   1.0000   0.0110
  -6.000  -0.4830   0.05274   0.05009  -0.0318   1.0000   0.0110
  -5.750  -0.4797   0.04976   0.04700  -0.0295   1.0000   0.0111
  -5.500  -0.4692   0.04633   0.04341  -0.0285   0.9994   0.0111
  -5.250  -0.4394   0.04133   0.03813  -0.0313   0.9967   0.0111
  -5.000  -0.4185   0.03219   0.02857  -0.0352   0.9934   0.0128
  -4.750  -0.3911   0.02986   0.02607  -0.0370   0.9904   0.0145
  -4.500  -0.3571   0.02854   0.02451  -0.0380   0.9877   0.0197
  -4.250  -0.3208   0.02780   0.02344  -0.0389   0.9854   0.0209
  -3.500  -0.2275   0.01675   0.01131  -0.0422   0.9786   0.0184
  -3.250  -0.1963   0.01354   0.00769  -0.0419   0.9761   0.0148
  -3.000  -0.1636   0.01222   0.00623  -0.0427   0.9739   0.0173
  -2.750  -0.1288   0.01193   0.00587  -0.0442   0.9721   0.0206
  -2.500  -0.0946   0.01096   0.00482  -0.0455   0.9707   0.0211
  -2.250  -0.0625   0.00934   0.00309  -0.0467   0.9690   0.0256
  -2.000  -0.0330   0.00894   0.00265  -0.0472   0.9650   0.0295
  -1.750  -0.0026   0.00853   0.00218  -0.0479   0.9611   0.0404
  -1.500   0.0197   0.00600   0.00181  -0.0483   0.9573   0.6718
  -1.250   0.0434   0.00569   0.00197  -0.0470   0.9535   0.8248
  -1.000   0.0629   0.00570   0.00213  -0.0445   0.9477   0.8815
  -0.750   0.0858   0.00575   0.00219  -0.0428   0.9435   0.9086
  -0.500   0.1055   0.00580   0.00225  -0.0405   0.9382   0.9299
  -0.250   0.1255   0.00580   0.00225  -0.0382   0.9329   0.9460
   0.000   0.1519   0.00577   0.00220  -0.0377   0.9289   0.9558
   0.250   0.1797   0.00573   0.00216  -0.0376   0.9225   0.9615
   0.500   0.2088   0.00567   0.00210  -0.0378   0.9167   0.9673
   0.750   0.2389   0.00557   0.00198  -0.0381   0.9060   0.9718
   1.000   0.2696   0.00546   0.00186  -0.0384   0.8926   0.9761
   1.250   0.3004   0.00540   0.00180  -0.0389   0.8805   0.9813
   1.500   0.3333   0.00530   0.00167  -0.0397   0.8574   0.9845
   1.750   0.3646   0.00527   0.00150  -0.0401   0.8089   0.9881
   2.000   0.3936   0.00549   0.00144  -0.0401   0.7121   0.9931
   2.250   0.4025   0.00827   0.00190  -0.0379   0.1444   1.0000
   2.500   0.4164   0.00914   0.00235  -0.0355   0.0357   1.0000
   2.750   0.4354   0.00991   0.00321  -0.0338   0.0253   1.0000
   3.000   0.4590   0.01040   0.00373  -0.0330   0.0236   1.0000
   3.250   0.4825   0.01110   0.00451  -0.0323   0.0225   1.0000
   3.500   0.5072   0.01169   0.00517  -0.0318   0.0201   1.0000
   3.750   0.5316   0.01253   0.00607  -0.0311   0.0184   1.0000
   4.000   0.5558   0.01373   0.00732  -0.0304   0.0162   1.0000
   4.250   0.5813   0.01551   0.00923  -0.0296   0.0162   1.0000
   4.500   0.6060   0.01775   0.01170  -0.0290   0.0133   1.0000
   4.750   0.6322   0.01938   0.01357  -0.0282   0.0122   1.0000
   5.000   0.6567   0.02260   0.01720  -0.0267   0.0120   1.0000
   5.500   0.6876   0.03437   0.02984  -0.0229   0.0142   1.0000
   5.750   0.7049   0.03783   0.03361  -0.0215   0.0142   1.0000
   6.000   0.7208   0.04142   0.03750  -0.0201   0.0142   1.0000
   6.500   0.7480   0.04889   0.04555  -0.0174   0.0141   1.0000
   6.750   0.7596   0.05268   0.04957  -0.0162   0.0141   1.0000
   7.000   0.7701   0.05643   0.05354  -0.0151   0.0140   1.0000
   7.250   0.7803   0.05998   0.05728  -0.0143   0.0138   1.0000
   7.500   0.7971   0.06264   0.06013  -0.0135   0.0129   1.0000
   7.750   0.8044   0.06684   0.06448  -0.0133   0.0121   1.0000
   8.000   0.8046   0.07129   0.06906  -0.0135   0.0119   1.0000
   8.250   0.8009   0.07583   0.07370  -0.0142   0.0117   1.0000
   8.500   0.7903   0.08002   0.07794  -0.0145   0.0117   1.0000
   8.750   0.7773   0.08466   0.08262  -0.0167   0.0116   1.0000
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