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NACA 16-006 (naca16006-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA 16-006 (naca16006-il)
Reynolds number: 1,000,000
Max Cl/Cd: 40.83 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-naca16006-il-1000000.txt
Download as CSV file: xf-naca16006-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA 16-006                                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5253   0.08273   0.08117  -0.0118   1.0000   0.0062
  -8.750  -0.5299   0.07788   0.07633  -0.0137   1.0000   0.0062
  -8.500  -0.5364   0.07280   0.07126  -0.0162   1.0000   0.0062
  -8.250  -0.5474   0.06728   0.06575  -0.0210   1.0000   0.0062
  -8.000  -0.5605   0.06277   0.06121  -0.0230   1.0000   0.0062
  -7.750  -0.5714   0.05860   0.05700  -0.0228   1.0000   0.0062
  -7.500  -0.5772   0.05412   0.05245  -0.0224   1.0000   0.0062
  -7.250  -0.5819   0.04977   0.04803  -0.0212   1.0000   0.0062
  -7.000  -0.5850   0.04559   0.04377  -0.0194   1.0000   0.0062
  -6.750  -0.5868   0.04165   0.03973  -0.0170   1.0000   0.0062
  -6.500  -0.5879   0.03779   0.03575  -0.0142   1.0000   0.0062
  -6.250  -0.5878   0.03409   0.03193  -0.0111   1.0000   0.0063
  -6.000  -0.5861   0.03060   0.02829  -0.0079   1.0000   0.0063
  -5.750  -0.5830   0.02720   0.02472  -0.0046   1.0000   0.0063
  -5.500  -0.5785   0.02385   0.02118  -0.0013   1.0000   0.0063
  -5.000  -0.5839   0.02435   0.02060   0.0063   0.9991   0.0055
  -4.750  -0.5590   0.01954   0.01526   0.0077   0.9979   0.0057
  -4.500  -0.5338   0.01576   0.01095   0.0086   0.9969   0.0070
  -4.250  -0.5055   0.01402   0.00900   0.0085   0.9962   0.0075
  -4.000  -0.4759   0.01292   0.00776   0.0080   0.9954   0.0081
  -3.750  -0.4448   0.01245   0.00723   0.0070   0.9947   0.0093
  -3.500  -0.4171   0.01063   0.00523   0.0073   0.9942   0.0089
  -3.250  -0.3902   0.00956   0.00399   0.0076   0.9930   0.0094
  -3.000  -0.3613   0.00907   0.00345   0.0073   0.9913   0.0116
  -2.750  -0.3307   0.00885   0.00320   0.0065   0.9897   0.0127
  -2.500  -0.3013   0.00817   0.00242   0.0061   0.9881   0.0188
  -2.250  -0.2698   0.00793   0.00218   0.0051   0.9867   0.0299
  -2.000  -0.2375   0.00775   0.00199   0.0039   0.9854   0.0378
  -1.750  -0.2071   0.00713   0.00178   0.0029   0.9843   0.1579
  -1.500  -0.1831   0.00548   0.00157   0.0025   0.9833   0.5665
  -1.250  -0.1588   0.00448   0.00154   0.0031   0.9822   0.8299
  -1.000  -0.1320   0.00435   0.00158   0.0035   0.9797   0.8848
  -0.750  -0.1037   0.00432   0.00161   0.0035   0.9767   0.9142
  -0.500  -0.0727   0.00434   0.00167   0.0030   0.9745   0.9383
  -0.250  -0.0399   0.00449   0.00186   0.0023   0.9729   0.9612
   0.000   0.0000   0.00463   0.00200   0.0000   0.9714   0.9714
   0.250   0.0398   0.00449   0.00186  -0.0023   0.9612   0.9729
   0.500   0.0727   0.00434   0.00167  -0.0029   0.9384   0.9745
   0.750   0.1036   0.00432   0.00161  -0.0034   0.9143   0.9768
   1.000   0.1318   0.00435   0.00158  -0.0034   0.8857   0.9798
   1.250   0.1588   0.00448   0.00154  -0.0031   0.8293   0.9822
   1.500   0.1831   0.00548   0.00156  -0.0025   0.5672   0.9833
   1.750   0.2071   0.00713   0.00178  -0.0029   0.1581   0.9843
   2.000   0.2376   0.00775   0.00199  -0.0040   0.0376   0.9855
   2.250   0.2698   0.00792   0.00218  -0.0051   0.0299   0.9868
   2.500   0.3013   0.00817   0.00241  -0.0061   0.0186   0.9882
   2.750   0.3308   0.00885   0.00320  -0.0066   0.0127   0.9898
   3.000   0.3613   0.00907   0.00345  -0.0073   0.0116   0.9914
   3.250   0.3903   0.00956   0.00399  -0.0076   0.0094   0.9931
   3.500   0.4173   0.01064   0.00523  -0.0074   0.0089   0.9943
   3.750   0.4451   0.01243   0.00721  -0.0071   0.0092   0.9947
   4.000   0.4761   0.01291   0.00775  -0.0080   0.0081   0.9955
   4.250   0.5057   0.01400   0.00898  -0.0085   0.0075   0.9963
   4.500   0.5340   0.01576   0.01095  -0.0087   0.0070   0.9971
   4.750   0.5593   0.01954   0.01526  -0.0078   0.0057   0.9980
   5.000   0.5843   0.02442   0.02068  -0.0063   0.0055   0.9992
   5.250   0.5897   0.03244   0.02930  -0.0027   0.0063   1.0000
   5.500   0.5985   0.03587   0.03296   0.0006   0.0063   1.0000
   5.750   0.6061   0.03932   0.03661   0.0039   0.0063   1.0000
   6.000   0.6126   0.04279   0.04026   0.0071   0.0063   1.0000
   6.250   0.6180   0.04630   0.04393   0.0101   0.0062   1.0000
   6.500   0.6226   0.04990   0.04767   0.0129   0.0062   1.0000
   6.750   0.6267   0.05361   0.05151   0.0152   0.0062   1.0000
   7.000   0.6304   0.05747   0.05548   0.0170   0.0062   1.0000
   7.250   0.6335   0.06148   0.05962   0.0182   0.0062   1.0000
   7.500   0.6362   0.06570   0.06393   0.0186   0.0062   1.0000
   7.750   0.6386   0.07014   0.06844   0.0180   0.0062   1.0000
   8.000   0.6400   0.07476   0.07313   0.0165   0.0062   1.0000
   8.250   0.6392   0.07942   0.07783   0.0143   0.0062   1.0000
   8.500   0.6341   0.08415   0.08258   0.0107   0.0062   1.0000
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