Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: NACA M16 AIRFOIL (m16-il)
Reynolds number: 100,000
Max Cl/Cd: 53.56 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m16-il-100000.txt
Download as CSV file: xf-m16-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4820   0.10608   0.10174   0.0163   1.0000   0.0464
  -7.750  -0.4765   0.10282   0.09853   0.0138   1.0000   0.0481
  -7.500  -0.4715   0.10000   0.09577   0.0098   1.0000   0.0499
  -7.250  -0.4591   0.09833   0.09409   0.0006   1.0000   0.0513
  -7.000  -0.4398   0.09693   0.09259  -0.0089   1.0000   0.0519
  -6.750  -0.4364   0.08898   0.08478  -0.0055   1.0000   0.0532
  -6.500  -0.4271   0.08420   0.08005  -0.0039   1.0000   0.0551
  -6.250  -0.4131   0.08038   0.07623  -0.0059   1.0000   0.0576
  -6.000  -0.3955   0.07668   0.07252  -0.0095   1.0000   0.0607
  -5.750  -0.3577   0.07534   0.07090  -0.0204   1.0000   0.0654
  -5.500  -0.3411   0.06985   0.06540  -0.0228   1.0000   0.0667
  -5.250  -0.3306   0.06495   0.06063  -0.0218   1.0000   0.0692
  -5.000  -0.3084   0.06161   0.05725  -0.0244   1.0000   0.0756
  -4.750  -0.2783   0.05828   0.05374  -0.0300   1.0000   0.0816
  -4.500  -0.2766   0.05584   0.05140  -0.0286   1.0000   0.0843
  -4.250  -0.2243   0.05228   0.04744  -0.0373   0.9857   0.0951
  -4.000  -0.1891   0.04776   0.04289  -0.0413   0.9718   0.1036
  -3.750  -0.1493   0.04424   0.03913  -0.0460   0.9568   0.1169
  -3.500  -0.1100   0.04222   0.03667  -0.0498   0.9402   0.1368
  -3.250  -0.0840   0.03968   0.03396  -0.0506   0.9229   0.1510
  -3.000  -0.0621   0.03720   0.03136  -0.0504   0.9069   0.1659
  -2.750  -0.0417   0.03498   0.02905  -0.0497   0.8920   0.1827
  -2.500  -0.0216   0.03317   0.02711  -0.0488   0.8782   0.2121
  -2.250  -0.0022   0.03140   0.02522  -0.0479   0.8654   0.2516
  -2.000   0.0172   0.03006   0.02374  -0.0466   0.8535   0.3034
  -1.750   0.0352   0.02764   0.02134  -0.0449   0.8429   0.3427
  -1.500   0.0575   0.02567   0.01930  -0.0440   0.8314   0.3804
  -1.250   0.0855   0.02427   0.01768  -0.0439   0.8205   0.3974
  -1.000   0.1546   0.02419   0.01577  -0.0458   0.8112   0.1326
  -0.750   0.1828   0.02282   0.01400  -0.0446   0.8018   0.1162
  -0.500   0.2124   0.02196   0.01271  -0.0440   0.7912   0.1068
  -0.250   0.2393   0.02099   0.01162  -0.0434   0.7814   0.1043
   0.000   0.2651   0.02026   0.01071  -0.0422   0.7729   0.1029
   0.250   0.2929   0.01966   0.01003  -0.0420   0.7623   0.1032
   0.500   0.3198   0.01920   0.00949  -0.0414   0.7525   0.1054
   0.750   0.3445   0.01870   0.00897  -0.0402   0.7446   0.1107
   1.000   0.3715   0.01845   0.00875  -0.0400   0.7339   0.1234
   1.250   0.3979   0.01822   0.00851  -0.0394   0.7245   0.1363
   1.500   0.4229   0.01780   0.00823  -0.0384   0.7166   0.1861
   1.750   0.4657   0.01608   0.00802  -0.0410   0.7059   1.0000
   2.000   0.4913   0.01637   0.00817  -0.0405   0.6965   1.0000
   2.250   0.5159   0.01658   0.00824  -0.0395   0.6884   1.0000
   2.500   0.5423   0.01691   0.00856  -0.0395   0.6779   1.0000
   2.750   0.5680   0.01722   0.00883  -0.0391   0.6687   1.0000
   3.000   0.5931   0.01746   0.00900  -0.0383   0.6605   1.0000
   3.250   0.6195   0.01785   0.00942  -0.0383   0.6501   1.0000
   3.500   0.6453   0.01818   0.00979  -0.0379   0.6412   1.0000
   3.750   0.6709   0.01846   0.01007  -0.0373   0.6327   1.0000
   4.000   0.6972   0.01889   0.01060  -0.0373   0.6224   1.0000
   4.250   0.7230   0.01924   0.01099  -0.0368   0.6140   1.0000
   4.500   0.7488   0.01957   0.01144  -0.0364   0.6048   1.0000
   4.750   0.7751   0.02006   0.01209  -0.0364   0.5947   1.0000
   5.000   0.8008   0.02038   0.01251  -0.0359   0.5861   1.0000
   5.250   0.8250   0.02003   0.01222  -0.0342   0.5680   1.0000
   5.500   0.8476   0.01873   0.01084  -0.0312   0.5394   1.0000
   5.750   0.8714   0.01777   0.00992  -0.0290   0.5086   1.0000
   6.000   0.8945   0.01670   0.00882  -0.0267   0.4460   1.0000
   6.250   0.9031   0.01978   0.00972  -0.0255   0.0970   1.0000
   6.500   0.9219   0.02191   0.01165  -0.0248   0.0670   1.0000
   6.750   0.9421   0.02340   0.01333  -0.0240   0.0599   1.0000
   7.000   0.9585   0.02520   0.01527  -0.0229   0.0547   1.0000
   7.250   0.9761   0.02668   0.01686  -0.0217   0.0507   1.0000
   7.500   0.9937   0.02836   0.01853  -0.0201   0.0490   1.0000
   7.750   1.0144   0.03022   0.02035  -0.0186   0.0479   1.0000
   8.000   1.0391   0.03235   0.02250  -0.0174   0.0476   1.0000
   8.250   1.0653   0.03478   0.02503  -0.0164   0.0474   1.0000
   8.500   1.0898   0.03807   0.02837  -0.0158   0.0457   1.0000
   8.750   1.1128   0.04102   0.03166  -0.0148   0.0457   1.0000
   9.000   1.1330   0.04374   0.03514  -0.0129   0.0496   1.0000
   9.250   1.1509   0.04835   0.04020  -0.0116   0.0539   1.0000
   9.500   1.1657   0.05257   0.04521  -0.0094   0.0633   1.0000
   9.750   1.1741   0.05946   0.05316  -0.0070   0.0882   1.0000
  10.750   1.0149   0.07074   0.06603   0.0042   0.1099   1.0000
  11.000   0.9878   0.07619   0.07158   0.0034   0.1093   1.0000
  11.250   0.9616   0.08224   0.07770   0.0018   0.1087   1.0000
  11.500   0.9345   0.08897   0.08448  -0.0004   0.1082   1.0000
  11.750   0.9068   0.09624   0.09174  -0.0032   0.1079   1.0000
  12.000   0.8768   0.10422   0.09974  -0.0070   0.1076   1.0000
<< Back to NACA M16 AIRFOIL (m16-il)

Polar data table (+)

Polar graphs


<< Back to NACA M16 AIRFOIL (m16-il)