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NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M16 AIRFOIL (m16-il)
Reynolds number: 200,000
Max Cl/Cd: 70.26 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m16-il-200000.txt
Download as CSV file: xf-m16-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -6.750  -0.4324   0.08696   0.08394  -0.0074   1.0000   0.0276
  -6.500  -0.4277   0.08115   0.07818  -0.0060   1.0000   0.0283
  -6.250  -0.4156   0.07704   0.07409  -0.0065   1.0000   0.0291
  -6.000  -0.3992   0.07321   0.07026  -0.0089   1.0000   0.0301
  -5.750  -0.3791   0.06941   0.06645  -0.0123   1.0000   0.0316
  -5.500  -0.3539   0.06560   0.06258  -0.0169   1.0000   0.0344
  -5.250  -0.3021   0.06294   0.05959  -0.0264   0.9385   0.0382
  -5.000  -0.2769   0.06048   0.05686  -0.0279   0.9068   0.0385
  -4.750  -0.2639   0.05480   0.05100  -0.0284   0.8859   0.0392
  -4.500  -0.2553   0.05076   0.04693  -0.0273   0.8672   0.0405
  -4.250  -0.2392   0.04807   0.04413  -0.0269   0.8508   0.0424
  -4.000  -0.2171   0.04555   0.04143  -0.0273   0.8366   0.0463
  -3.750  -0.1726   0.04543   0.04062  -0.0288   0.8238   0.0520
  -3.500  -0.1612   0.03938   0.03461  -0.0292   0.8130   0.0543
  -3.250  -0.1399   0.03717   0.03228  -0.0291   0.8019   0.0576
  -3.000  -0.1041   0.03539   0.02990  -0.0297   0.7912   0.0666
  -2.750  -0.0856   0.03253   0.02713  -0.0298   0.7810   0.0704
  -2.500  -0.0544   0.03108   0.02514  -0.0298   0.7724   0.0811
  -2.250  -0.0315   0.02869   0.02278  -0.0301   0.7623   0.0847
  -2.000  -0.0026   0.02715   0.02088  -0.0301   0.7535   0.0963
  -1.750   0.0245   0.02601   0.01944  -0.0300   0.7453   0.1095
  -1.500   0.0513   0.02464   0.01795  -0.0302   0.7360   0.1237
  -1.250   0.0780   0.02338   0.01651  -0.0302   0.7281   0.1379
  -1.000   0.1046   0.02202   0.01506  -0.0303   0.7194   0.1532
  -0.750   0.1442   0.01928   0.01127  -0.0287   0.7127   0.0806
  -0.500   0.1724   0.01755   0.00942  -0.0285   0.7047   0.0741
  -0.250   0.2011   0.01670   0.00838  -0.0283   0.6966   0.0743
   0.000   0.2296   0.01572   0.00712  -0.0278   0.6893   0.0693
   0.250   0.2582   0.01517   0.00644  -0.0277   0.6809   0.0672
   0.500   0.2855   0.01469   0.00584  -0.0273   0.6739   0.0667
   0.750   0.3133   0.01410   0.00528  -0.0272   0.6652   0.0674
   1.000   0.3403   0.01367   0.00485  -0.0269   0.6581   0.0697
   1.250   0.3680   0.01338   0.00460  -0.0269   0.6496   0.0736
   1.500   0.3955   0.01321   0.00441  -0.0267   0.6420   0.0788
   1.750   0.4231   0.01303   0.00426  -0.0266   0.6340   0.0931
   2.000   0.4607   0.01080   0.00412  -0.0285   0.6257   1.0000
   2.250   0.4872   0.01093   0.00407  -0.0281   0.6182   1.0000
   2.500   0.5144   0.01105   0.00416  -0.0280   0.6095   1.0000
   2.750   0.5408   0.01121   0.00419  -0.0277   0.6025   1.0000
   3.000   0.5682   0.01133   0.00433  -0.0277   0.5934   1.0000
   3.250   0.5952   0.01148   0.00444  -0.0275   0.5857   1.0000
   3.500   0.6223   0.01162   0.00459  -0.0274   0.5772   1.0000
   3.750   0.6495   0.01177   0.00477  -0.0273   0.5688   1.0000
   4.000   0.6765   0.01193   0.00489  -0.0271   0.5612   1.0000
   4.250   0.7040   0.01209   0.00513  -0.0271   0.5519   1.0000
   4.500   0.7311   0.01222   0.00533  -0.0270   0.5429   1.0000
   4.750   0.7576   0.01207   0.00518  -0.0265   0.5237   1.0000
   5.000   0.7840   0.01196   0.00505  -0.0261   0.5029   1.0000
   5.250   0.8108   0.01193   0.00508  -0.0258   0.4800   1.0000
   5.500   0.8368   0.01191   0.00498  -0.0254   0.4303   1.0000
   5.750   0.8529   0.01490   0.00613  -0.0257   0.0929   1.0000
   6.000   0.8756   0.01652   0.00752  -0.0255   0.0425   1.0000
   6.250   0.8994   0.01758   0.00871  -0.0253   0.0356   1.0000
   6.500   0.9204   0.01910   0.01041  -0.0248   0.0328   1.0000
   6.750   0.9416   0.02037   0.01179  -0.0242   0.0317   1.0000
   7.000   0.9615   0.02182   0.01335  -0.0232   0.0309   1.0000
   7.250   0.9813   0.02335   0.01494  -0.0221   0.0300   1.0000
   7.500   1.0019   0.02475   0.01638  -0.0212   0.0278   1.0000
   7.750   1.0230   0.02647   0.01816  -0.0201   0.0270   1.0000
   8.000   1.0459   0.02856   0.02038  -0.0190   0.0274   1.0000
   8.250   1.0703   0.03123   0.02326  -0.0178   0.0290   1.0000
   8.500   1.0939   0.03529   0.02755  -0.0167   0.0319   1.0000
  14.750   0.9391   0.17785   0.17431  -0.0617   0.0422   1.0000
  15.000   0.9402   0.18353   0.17997  -0.0652   0.0399   1.0000
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