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NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: NACA M16 AIRFOIL (m16-il)
Reynolds number: 50,000
Max Cl/Cd: 38.13 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m16-il-50000-n5.txt
Download as CSV file: xf-m16-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4683   0.11194   0.10584   0.0119   1.0000   0.0617
  -8.000  -0.4639   0.10910   0.10306   0.0091   1.0000   0.0649
  -7.750  -0.4621   0.10724   0.10129   0.0041   1.0000   0.0669
  -7.500  -0.4547   0.10559   0.09966  -0.0032   1.0000   0.0678
  -7.250  -0.4452   0.09937   0.09353  -0.0019   1.0000   0.0694
  -7.000  -0.4347   0.09470   0.08883  -0.0012   1.0000   0.0727
  -6.750  -0.4234   0.09107   0.08522  -0.0039   1.0000   0.0762
  -6.500  -0.4088   0.08826   0.08239  -0.0098   1.0000   0.0807
  -6.000  -0.3807   0.07994   0.07410  -0.0149   1.0000   0.0853
  -5.750  -0.3631   0.07670   0.07081  -0.0181   1.0000   0.0941
  -5.500  -0.3462   0.07264   0.06675  -0.0209   1.0000   0.0994
  -5.250  -0.3190   0.07075   0.06462  -0.0270   1.0000   0.1104
  -4.750  -0.2782   0.06363   0.05736  -0.0316   1.0000   0.1250
  -4.500  -0.2686   0.05873   0.05261  -0.0303   1.0000   0.1297
  -4.250  -0.2497   0.05576   0.04958  -0.0320   0.9896   0.1461
  -3.750  -0.1804   0.04841   0.04180  -0.0406   0.9469   0.1898
  -3.500  -0.1187   0.04330   0.03573  -0.0471   0.9313   0.0925
  -3.250  -0.0840   0.04028   0.03229  -0.0493   0.9149   0.0913
  -3.000  -0.0465   0.03737   0.02864  -0.0509   0.8998   0.0813
  -2.750  -0.0177   0.03497   0.02587  -0.0514   0.8849   0.0794
  -2.500   0.0111   0.03295   0.02343  -0.0516   0.8707   0.0780
  -2.250   0.0396   0.03119   0.02123  -0.0515   0.8573   0.0773
  -2.000   0.0679   0.02979   0.01935  -0.0512   0.8444   0.0797
  -1.750   0.0943   0.02864   0.01788  -0.0510   0.8319   0.0857
  -1.500   0.1221   0.02747   0.01636  -0.0506   0.8201   0.0863
  -1.250   0.1502   0.02641   0.01490  -0.0501   0.8091   0.0855
  -1.000   0.1778   0.02550   0.01365  -0.0494   0.7987   0.0852
  -0.750   0.2069   0.02476   0.01258  -0.0492   0.7874   0.0853
  -0.500   0.2363   0.02413   0.01169  -0.0490   0.7767   0.0858
  -0.250   0.2640   0.02359   0.01096  -0.0484   0.7669   0.0868
   0.000   0.2904   0.02315   0.01036  -0.0477   0.7572   0.0884
   0.250   0.3164   0.02282   0.00988  -0.0472   0.7468   0.0908
   0.500   0.3417   0.02256   0.00948  -0.0465   0.7374   0.0945
   0.750   0.3667   0.02232   0.00917  -0.0457   0.7283   0.1000
   1.000   0.3925   0.02217   0.00901  -0.0454   0.7182   0.1116
   1.250   0.4177   0.02187   0.00890  -0.0448   0.7092   0.1534
   1.500   0.4580   0.02005   0.00873  -0.0470   0.7001   1.0000
   1.750   0.4836   0.02037   0.00886  -0.0467   0.6900   1.0000
   2.000   0.5086   0.02066   0.00902  -0.0460   0.6812   1.0000
   2.250   0.5337   0.02097   0.00923  -0.0455   0.6721   1.0000
   2.500   0.5591   0.02134   0.00956  -0.0452   0.6623   1.0000
   2.750   0.5842   0.02164   0.00983  -0.0445   0.6541   1.0000
   3.000   0.6095   0.02203   0.01023  -0.0442   0.6444   1.0000
   3.250   0.6349   0.02245   0.01069  -0.0440   0.6349   1.0000
   3.500   0.6598   0.02275   0.01099  -0.0432   0.6273   1.0000
   3.750   0.6852   0.02326   0.01166  -0.0431   0.6169   1.0000
   4.000   0.7104   0.02371   0.01222  -0.0428   0.6078   1.0000
   4.250   0.7355   0.02408   0.01269  -0.0422   0.5995   1.0000
   4.500   0.7605   0.02466   0.01347  -0.0421   0.5893   1.0000
   4.750   0.7855   0.02514   0.01416  -0.0416   0.5805   1.0000
   5.000   0.8105   0.02560   0.01483  -0.0411   0.5715   1.0000
   5.250   0.8350   0.02626   0.01577  -0.0410   0.5612   1.0000
   5.500   0.8599   0.02677   0.01655  -0.0404   0.5525   1.0000
   5.750   0.8847   0.02719   0.01733  -0.0397   0.5425   1.0000
   6.000   0.9056   0.02577   0.01604  -0.0362   0.5036   1.0000
   6.250   0.9254   0.02451   0.01481  -0.0328   0.4458   1.0000
   6.500   0.9392   0.02463   0.01427  -0.0298   0.2918   1.0000
   6.750   0.9351   0.02931   0.01696  -0.0286   0.0698   1.0000
   7.000   0.9466   0.03173   0.01939  -0.0276   0.0556   1.0000
   7.250   0.9574   0.03392   0.02179  -0.0266   0.0481   1.0000
   7.500   0.9655   0.03620   0.02422  -0.0254   0.0432   1.0000
   7.750   0.9711   0.03857   0.02675  -0.0240   0.0411   1.0000
   8.000   0.9785   0.04063   0.02903  -0.0224   0.0395   1.0000
   8.250   0.9864   0.04263   0.03122  -0.0207   0.0377   1.0000
   8.500   0.9950   0.04454   0.03326  -0.0188   0.0354   1.0000
   8.750   1.0115   0.04653   0.03518  -0.0165   0.0326   1.0000
   9.000   1.0439   0.04816   0.03714  -0.0147   0.0312   1.0000
   9.250   1.0753   0.05065   0.04005  -0.0133   0.0304   1.0000
   9.500   1.0981   0.05373   0.04349  -0.0121   0.0300   1.0000
   9.750   1.1119   0.05701   0.04717  -0.0110   0.0293   1.0000
  10.000   1.1189   0.06038   0.05095  -0.0097   0.0286   1.0000
  10.250   1.1205   0.06386   0.05479  -0.0085   0.0279   1.0000
  10.500   1.1163   0.06732   0.05857  -0.0072   0.0275   1.0000
  10.750   1.1079   0.07093   0.06247  -0.0063   0.0273   1.0000
  11.000   1.0970   0.07491   0.06672  -0.0062   0.0273   1.0000
  11.250   1.0841   0.07933   0.07139  -0.0069   0.0273   1.0000
  11.500   1.0695   0.08420   0.07650  -0.0084   0.0275   1.0000
  11.750   1.0537   0.08959   0.08209  -0.0106   0.0277   1.0000
  12.000   1.0368   0.09552   0.08821  -0.0135   0.0280   1.0000
  12.250   1.0195   0.10200   0.09485  -0.0172   0.0283   1.0000
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