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NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NACA M22 AIRFOIL (m22-il)
Reynolds number: 200,000
Max Cl/Cd: 74.9 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m22-il-200000.txt
Download as CSV file: xf-m22-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M22 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
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.3035   0.02488   0.01759  -0.0243   0.5873   0.0948
   0.500   0.3331   0.02400   0.01653  -0.0245   0.5820   0.1052
   0.750   0.3621   0.02332   0.01567  -0.0245   0.5768   0.1179
   1.000   0.3895   0.02253   0.01469  -0.0244   0.5729   0.1315
   1.250   0.4172   0.02151   0.01366  -0.0248   0.5676   0.1463
   2.000   0.5095   0.01869   0.01004  -0.0236   0.5538   0.0871
   2.250   0.5392   0.01774   0.00893  -0.0234   0.5489   0.0735
   2.500   0.5670   0.01701   0.00810  -0.0232   0.5450   0.0699
   2.750   0.5951   0.01656   0.00761  -0.0231   0.5400   0.0673
   3.000   0.6225   0.01619   0.00725  -0.0230   0.5348   0.0664
   3.250   0.6492   0.01592   0.00694  -0.0227   0.5308   0.0672
   3.500   0.6764   0.01580   0.00690  -0.0227   0.5262   0.0698
   3.750   0.7034   0.01572   0.00684  -0.0226   0.5209   0.0744
   4.000   0.7300   0.01563   0.00672  -0.0223   0.5168   0.0885
   4.250   0.8034   0.01432   0.00707  -0.0325   0.5101   1.0000
   4.500   0.8295   0.01450   0.00724  -0.0323   0.5054   1.0000
   4.750   0.8551   0.01468   0.00733  -0.0318   0.5017   1.0000
   5.000   0.8815   0.01495   0.00769  -0.0319   0.4963   1.0000
   5.250   0.9066   0.01478   0.00741  -0.0312   0.4877   1.0000
   5.500   0.9317   0.01435   0.00691  -0.0304   0.4725   1.0000
   5.750   0.9578   0.01431   0.00699  -0.0303   0.4619   1.0000
   6.000   0.9837   0.01427   0.00701  -0.0300   0.4510   1.0000
   6.250   1.0093   0.01388   0.00660  -0.0295   0.4244   1.0000
   6.500   1.0351   0.01382   0.00651  -0.0293   0.3906   1.0000
   6.750   1.0596   0.01442   0.00681  -0.0294   0.3118   1.0000
   7.000   1.0675   0.01928   0.01000  -0.0311   0.0396   1.0000
   7.250   1.0877   0.02040   0.01123  -0.0307   0.0323   1.0000
   7.500   1.1075   0.02142   0.01244  -0.0302   0.0307   1.0000
   7.750   1.1247   0.02268   0.01391  -0.0295   0.0295   1.0000
   8.000   1.1380   0.02423   0.01566  -0.0288   0.0288   1.0000
   8.250   1.1466   0.02609   0.01771  -0.0279   0.0282   1.0000
   8.500   1.1507   0.02830   0.02004  -0.0273   0.0271   1.0000
   8.750   1.1480   0.03071   0.02252  -0.0262   0.0261   1.0000
   9.000   1.1456   0.03318   0.02505  -0.0249   0.0253   1.0000
   9.250   1.1461   0.03551   0.02743  -0.0236   0.0250   1.0000
   9.500   1.1496   0.03755   0.02951  -0.0220   0.0249   1.0000
   9.750   1.1576   0.03916   0.03115  -0.0200   0.0250   1.0000
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