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

NACA M21 AIRFOIL (m21-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M21 AIRFOIL (m21-il)
Reynolds number: 200,000
Max Cl/Cd: 70.14 at α=10.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m21-il-200000.txt
Download as CSV file: xf-m21-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M21 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.2386   0.09048   0.08660  -0.0301   0.7585   0.0524
  -7.500  -0.2374   0.08771   0.08381  -0.0343   0.7505   0.0539
  -7.250  -0.2324   0.08567   0.08157  -0.0459   0.7445   0.0550
  -7.000  -0.2224   0.08121   0.07706  -0.0478   0.7383   0.0553
  -6.750  -0.2116   0.07718   0.07305  -0.0460   0.7316   0.0558
  -6.500  -0.2000   0.07409   0.06988  -0.0451   0.7262   0.0564
  -6.250  -0.1862   0.07125   0.06704  -0.0456   0.7192   0.0573
  -6.000  -0.1715   0.06857   0.06427  -0.0466   0.7131   0.0588
  -5.750  -0.1551   0.06584   0.06142  -0.0484   0.7081   0.0609
  -5.500  -0.1210   0.06460   0.05972  -0.0558   0.7017   0.0643
  -5.250  -0.1080   0.05984   0.05489  -0.0561   0.6964   0.0649
  -5.000  -0.0942   0.05651   0.05154  -0.0555   0.6917   0.0655
  -4.750  -0.0771   0.05387   0.04892  -0.0554   0.6852   0.0663
  -4.500  -0.0584   0.05161   0.04657  -0.0555   0.6798   0.0678
  -4.250  -0.0376   0.04952   0.04431  -0.0557   0.6754   0.0702
  -4.000   0.0099   0.03191   0.02651  -0.0573   0.6600   0.0758
  -3.750   0.0249   0.02862   0.02328  -0.0572   0.6552   0.0766
  -3.500   0.0427   0.02634   0.02099  -0.0569   0.6499   0.0778
  -3.250   0.0623   0.02440   0.01893  -0.0565   0.6453   0.0794
  -3.000   0.0839   0.02264   0.01703  -0.0563   0.6410   0.0823
  -2.750   0.1121   0.02099   0.01490  -0.0563   0.6354   0.0888
  -2.500   0.1313   0.01891   0.01284  -0.0561   0.6309   0.0902
  -2.250   0.1527   0.01745   0.01126  -0.0557   0.6270   0.0924
  -2.000   0.1832   0.01767   0.01101  -0.0550   0.6211   0.1014
  -1.750   0.2028   0.01498   0.00830  -0.0550   0.6167   0.1032
  -1.500   0.2254   0.01373   0.00700  -0.0548   0.6126   0.1057
  -1.250   0.2505   0.01295   0.00604  -0.0544   0.6087   0.1119
  -1.000   0.2747   0.01198   0.00497  -0.0542   0.6032   0.1208
  -0.750   0.3002   0.01130   0.00407  -0.0539   0.5983   0.1339
  -0.500   0.3251   0.01053   0.00322  -0.0537   0.5946   0.1400
  -0.250   0.3724   0.02189   0.01312  -0.0539   0.5964   0.0891
   0.000   0.3995   0.02107   0.01229  -0.0540   0.5925   0.0927
   0.250   0.4280   0.02021   0.01132  -0.0539   0.5868   0.0929
   0.500   0.4564   0.01957   0.01054  -0.0538   0.5819   0.0952
   0.750   0.4849   0.01927   0.01002  -0.0536   0.5780   0.1002
   1.000   0.5130   0.01865   0.00936  -0.0537   0.5732   0.1036
   1.250   0.5406   0.01825   0.00901  -0.0538   0.5679   0.1086
   1.500   0.5686   0.01801   0.00867  -0.0537   0.5636   0.1142
   1.750   0.5962   0.01764   0.00824  -0.0535   0.5597   0.1184
   2.000   0.6225   0.01744   0.00818  -0.0535   0.5542   0.1249
   2.250   0.6495   0.01728   0.00801  -0.0533   0.5495   0.1298
   2.500   0.6759   0.01700   0.00772  -0.0530   0.5457   0.1362
   2.750   0.7018   0.01697   0.00777  -0.0528   0.5408   0.1440
   3.000   0.7273   0.01688   0.00775  -0.0524   0.5357   0.1516
   3.250   0.7538   0.01683   0.00768  -0.0522   0.5316   0.1645
   3.500   0.7804   0.01679   0.00765  -0.0519   0.5281   0.1948
   3.750   0.9000   0.01561   0.00819  -0.0716   0.5194   1.0000
   4.000   0.9253   0.01574   0.00822  -0.0711   0.5153   1.0000
   4.250   0.9508   0.01595   0.00831  -0.0707   0.5117   1.0000
   4.500   0.9742   0.01620   0.00865  -0.0702   0.5061   1.0000
   4.750   0.9987   0.01636   0.00879  -0.0698   0.5014   1.0000
   5.000   1.0242   0.01651   0.00884  -0.0694   0.4976   1.0000
   5.250   1.0478   0.01681   0.00919  -0.0689   0.4931   1.0000
   5.500   1.0712   0.01708   0.00951  -0.0683   0.4880   1.0000
   5.750   1.0960   0.01724   0.00964  -0.0678   0.4838   1.0000
   6.000   1.1217   0.01745   0.00975  -0.0675   0.4803   1.0000
   6.250   1.1430   0.01781   0.01027  -0.0668   0.4748   1.0000
   6.500   1.1666   0.01806   0.01056  -0.0662   0.4703   1.0000
   6.750   1.1918   0.01825   0.01070  -0.0658   0.4665   1.0000
   7.000   1.2149   0.01859   0.01110  -0.0652   0.4622   1.0000
   7.250   1.2363   0.01894   0.01156  -0.0644   0.4569   1.0000
   7.500   1.2603   0.01918   0.01182  -0.0640   0.4528   1.0000
   7.750   1.2862   0.01941   0.01198  -0.0637   0.4493   1.0000
   8.000   1.3056   0.01988   0.01264  -0.0627   0.4441   1.0000
   8.250   1.3276   0.02013   0.01296  -0.0619   0.4388   1.0000
   8.500   1.3541   0.02012   0.01284  -0.0617   0.4335   1.0000
   8.750   1.3710   0.02039   0.01332  -0.0601   0.4258   1.0000
   9.000   1.3953   0.02030   0.01316  -0.0595   0.4190   1.0000
   9.250   1.4128   0.02059   0.01358  -0.0581   0.4116   1.0000
   9.500   1.4343   0.02063   0.01364  -0.0571   0.4045   1.0000
   9.750   1.4514   0.02087   0.01399  -0.0556   0.3966   1.0000
  10.000   1.4704   0.02097   0.01411  -0.0543   0.3886   1.0000
  10.250   1.4847   0.02127   0.01453  -0.0524   0.3799   1.0000
  10.500   1.5017   0.02141   0.01463  -0.0508   0.3707   1.0000
  10.750   1.5098   0.02179   0.01519  -0.0480   0.3597   1.0000
  11.000   1.5166   0.02224   0.01572  -0.0451   0.3475   1.0000
  11.250   1.5183   0.02282   0.01634  -0.0414   0.3347   1.0000
  11.500   1.5137   0.02373   0.01729  -0.0373   0.3201   1.0000
  11.750   1.5065   0.02521   0.01877  -0.0339   0.3006   1.0000
  12.000   1.4944   0.02749   0.02099  -0.0310   0.2738   1.0000
  12.250   1.4758   0.03078   0.02413  -0.0285   0.2403   1.0000
  12.500   1.4518   0.03500   0.02815  -0.0265   0.2046   1.0000
  12.750   1.4249   0.03984   0.03280  -0.0250   0.1712   1.0000
  13.000   1.3975   0.04498   0.03779  -0.0238   0.1412   1.0000
  13.250   1.3691   0.05048   0.04313  -0.0229   0.1110   1.0000
  13.500   1.3415   0.05625   0.04874  -0.0224   0.0841   1.0000
  13.750   1.3174   0.06189   0.05428  -0.0223   0.0650   1.0000
  14.000   1.2991   0.06708   0.05942  -0.0223   0.0541   1.0000
  14.250   1.2844   0.07198   0.06434  -0.0225   0.0484   1.0000
  14.500   1.2753   0.07631   0.06876  -0.0228   0.0448   1.0000
  14.750   1.2654   0.08083   0.07334  -0.0232   0.0423   1.0000
  15.000   1.2540   0.08562   0.07820  -0.0237   0.0406   1.0000
  15.250   1.2464   0.08997   0.08266  -0.0243   0.0392   1.0000
<< Back to NACA M21 AIRFOIL (m21-il)

Polar data table (+)

Polar graphs


<< Back to NACA M21 AIRFOIL (m21-il)