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NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: NACA M16 AIRFOIL (m16-il)
Reynolds number: 500,000
Max Cl/Cd: 92.65 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m16-il-500000.txt
Download as CSV file: xf-m16-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M16 AIRFOIL                                
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5229   0.11456   0.11261   0.0262   1.0000   0.0139
  -8.500  -0.5192   0.11112   0.10918   0.0233   1.0000   0.0139
  -8.250  -0.5152   0.10757   0.10565   0.0204   1.0000   0.0140
  -8.000  -0.5108   0.10389   0.10200   0.0173   1.0000   0.0140
  -7.750  -0.5031   0.09985   0.09797   0.0135   1.0000   0.0140
  -7.500  -0.5007   0.09431   0.09246   0.0159   1.0000   0.0145
  -7.250  -0.4910   0.09081   0.08897   0.0138   1.0000   0.0148
  -7.000  -0.4794   0.08734   0.08551   0.0109   1.0000   0.0153
  -6.750  -0.4656   0.08389   0.08201   0.0075   0.9312   0.0161
  -6.500  -0.4569   0.08084   0.07880   0.0060   0.8826   0.0171
  -6.250  -0.4415   0.07724   0.07505   0.0027   0.8536   0.0181
  -6.000  -0.4135   0.07371   0.07133  -0.0038   0.8308   0.0190
  -5.750  -0.3903   0.06982   0.06728  -0.0076   0.8123   0.0192
  -5.500  -0.3681   0.06579   0.06310  -0.0104   0.7962   0.0193
  -5.250  -0.3451   0.06156   0.05872  -0.0129   0.7820   0.0194
  -5.000  -0.3212   0.05727   0.05423  -0.0152   0.7693   0.0194
  -4.500  -0.2897   0.04861   0.04535  -0.0174   0.7467   0.0211
  -4.250  -0.2660   0.04632   0.04295  -0.0184   0.7358   0.0229
  -4.000  -0.2284   0.04421   0.04057  -0.0201   0.7264   0.0267
  -3.750  -0.1985   0.04128   0.03741  -0.0209   0.7176   0.0271
  -3.250  -0.1477   0.03074   0.02625  -0.0226   0.7028   0.0286
  -3.000  -0.1241   0.02936   0.02479  -0.0231   0.6941   0.0300
  -2.750  -0.0979   0.02782   0.02308  -0.0233   0.6868   0.0324
  -2.500  -0.0646   0.02734   0.02234  -0.0229   0.6789   0.0382
  -2.250  -0.0370   0.02263   0.01707  -0.0230   0.6731   0.0398
  -2.000  -0.0112   0.02093   0.01532  -0.0234   0.6656   0.0419
  -1.750   0.0160   0.01992   0.01415  -0.0236   0.6591   0.0445
  -1.500   0.0456   0.01937   0.01339  -0.0234   0.6519   0.0506
  -1.250   0.0735   0.01730   0.01097  -0.0235   0.6459   0.0548
  -1.000   0.1012   0.01655   0.01021  -0.0238   0.6389   0.0581
  -0.500   0.1606   0.01362   0.00667  -0.0231   0.6264   0.0486
  -0.250   0.1893   0.01247   0.00533  -0.0230   0.6200   0.0457
   0.000   0.2177   0.01171   0.00444  -0.0228   0.6138   0.0439
   0.250   0.2459   0.01127   0.00396  -0.0228   0.6070   0.0447
   0.500   0.2739   0.01082   0.00344  -0.0228   0.6008   0.0440
   0.750   0.3020   0.01041   0.00300  -0.0227   0.5939   0.0432
   1.000   0.3300   0.01014   0.00268  -0.0227   0.5876   0.0432
   1.250   0.3583   0.00988   0.00243  -0.0228   0.5803   0.0440
   1.500   0.3864   0.00974   0.00225  -0.0228   0.5739   0.0461
   1.750   0.4149   0.00959   0.00212  -0.0230   0.5664   0.0490
   2.000   0.4431   0.00955   0.00203  -0.0231   0.5597   0.0516
   2.250   0.4717   0.00945   0.00197  -0.0232   0.5519   0.0587
   2.500   0.4942   0.00769   0.00204  -0.0229   0.5455   0.7989
   2.750   0.5305   0.00730   0.00205  -0.0242   0.5370   1.0000
   3.000   0.5579   0.00739   0.00208  -0.0242   0.5295   1.0000
   3.250   0.5854   0.00747   0.00213  -0.0242   0.5214   1.0000
   3.500   0.6130   0.00755   0.00222  -0.0242   0.5133   1.0000
   3.750   0.6406   0.00766   0.00229  -0.0242   0.5048   1.0000
   4.000   0.6682   0.00774   0.00234  -0.0242   0.4898   1.0000
   4.250   0.6959   0.00783   0.00241  -0.0243   0.4723   1.0000
   4.500   0.7236   0.00795   0.00254  -0.0244   0.4573   1.0000
   4.750   0.7510   0.00813   0.00264  -0.0245   0.4297   1.0000
   5.000   0.7783   0.00840   0.00280  -0.0246   0.3952   1.0000
   5.250   0.8018   0.01026   0.00357  -0.0253   0.1771   1.0000
   5.500   0.8249   0.01233   0.00491  -0.0258   0.0248   1.0000
   5.750   0.8515   0.01288   0.00558  -0.0258   0.0218   1.0000
   6.000   0.8777   0.01349   0.00628  -0.0258   0.0200   1.0000
   6.250   0.9031   0.01424   0.00713  -0.0257   0.0184   1.0000
   6.500   0.9273   0.01522   0.00818  -0.0255   0.0167   1.0000
   6.750   0.9463   0.01723   0.01034  -0.0249   0.0152   1.0000
   7.000   0.9687   0.01831   0.01151  -0.0244   0.0147   1.0000
   7.250   0.9911   0.01936   0.01263  -0.0238   0.0143   1.0000
   7.500   1.0123   0.02073   0.01407  -0.0230   0.0139   1.0000
   7.750   1.0330   0.02248   0.01591  -0.0219   0.0138   1.0000
   8.000   1.0540   0.02501   0.01853  -0.0208   0.0143   1.0000
   8.250   1.0754   0.02902   0.02268  -0.0198   0.0149   1.0000
   8.500   1.1012   0.02904   0.02277  -0.0193   0.0160   1.0000
   8.750   1.1242   0.03300   0.02731  -0.0172   0.0196   1.0000
   9.000   1.1426   0.03509   0.02959  -0.0164   0.0182   1.0000
   9.250   1.1590   0.03746   0.03211  -0.0157   0.0172   1.0000
   9.500   1.1750   0.04036   0.03510  -0.0151   0.0164   1.0000
   9.750   1.1643   0.05149   0.04670  -0.0138   0.0150   1.0000
  10.000   1.1678   0.05484   0.05034  -0.0123   0.0150   1.0000
  10.250   1.1674   0.05806   0.05384  -0.0108   0.0150   1.0000
  10.500   1.1628   0.06112   0.05715  -0.0092   0.0149   1.0000
  10.750   1.1541   0.06395   0.06021  -0.0075   0.0149   1.0000
  11.000   1.1389   0.06653   0.06296  -0.0055   0.0148   1.0000
  11.250   1.1229   0.06969   0.06628  -0.0052   0.0148   1.0000
  11.500   1.1068   0.07350   0.07025  -0.0060   0.0147   1.0000
  11.750   1.0902   0.07803   0.07493  -0.0078   0.0147   1.0000
  12.000   1.0725   0.08336   0.08040  -0.0103   0.0147   1.0000
  12.250   1.0543   0.08937   0.08654  -0.0134   0.0147   1.0000
  12.500   1.0359   0.09588   0.09316  -0.0170   0.0148   1.0000
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