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NACA M22 AIRFOIL (m22-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: NACA M22 AIRFOIL (m22-il)
Reynolds number: 1,000,000
Max Cl/Cd: 115 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m22-il-1000000.txt
Download as CSV file: xf-m22-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4583   0.10591   0.10352   0.0353   0.6390   0.0076
  -7.500  -0.4529   0.10286   0.10046   0.0335   0.6338   0.0077
  -7.000  -0.4356   0.09661   0.09418   0.0287   0.6238   0.0080
  -6.750  -0.4232   0.09333   0.09087   0.0258   0.6185   0.0082
  -6.500  -0.4093   0.08996   0.08745   0.0229   0.6137   0.0085
  -6.250  -0.3940   0.08651   0.08397   0.0198   0.6086   0.0089
  -6.000  -0.3765   0.08284   0.08026   0.0166   0.6036   0.0099
  -5.750  -0.3512   0.07894   0.07630   0.0116   0.5991   0.0104
  -5.500  -0.3282   0.07498   0.07228   0.0076   0.5943   0.0105
  -5.250  -0.3053   0.07113   0.06836   0.0043   0.5894   0.0105
  -5.000  -0.2913   0.06649   0.06367   0.0026   0.5851   0.0109
  -4.750  -0.2703   0.06379   0.06093   0.0008   0.5800   0.0112
  -4.500  -0.2470   0.06116   0.05822  -0.0012   0.5748   0.0118
  -4.250  -0.2213   0.05829   0.05526  -0.0035   0.5702   0.0133
  -4.000  -0.1853   0.05513   0.05198  -0.0068   0.5657   0.0142
  -3.750  -0.1553   0.05162   0.04834  -0.0091   0.5612   0.0143
  -3.500  -0.1260   0.04806   0.04465  -0.0108   0.5569   0.0144
  -3.250  -0.1058   0.04368   0.04020  -0.0119   0.5528   0.0149
  -3.000  -0.0817   0.04186   0.03831  -0.0126   0.5480   0.0154
  -2.750  -0.0552   0.04008   0.03642  -0.0135   0.5433   0.0167
  -2.500  -0.0169   0.03826   0.03443  -0.0145   0.5393   0.0190
  -2.250   0.0128   0.03560   0.03163  -0.0150   0.5350   0.0191
  -1.750   0.0672   0.02863   0.02430  -0.0158   0.5275   0.0200
  -1.500   0.0925   0.02747   0.02308  -0.0162   0.5231   0.0207
  -1.250   0.1197   0.02626   0.02177  -0.0165   0.5187   0.0221
  -1.000   0.1520   0.02544   0.02079  -0.0164   0.5146   0.0254
  -0.750   0.1818   0.02404   0.01927  -0.0164   0.5109   0.0258
  -0.500   0.2110   0.02237   0.01745  -0.0162   0.5067   0.0259
  -0.250   0.2402   0.02048   0.01536  -0.0160   0.5028   0.0260
   1.000   0.3807   0.01348   0.00757  -0.0162   0.4836   0.0311
   1.250   0.4094   0.01285   0.00684  -0.0162   0.4796   0.0333
   1.500   0.4384   0.01318   0.00711  -0.0163   0.4755   0.0355
   1.750   0.4671   0.01254   0.00632  -0.0163   0.4716   0.0357
   2.500   0.5518   0.01057   0.00413  -0.0168   0.4592   0.0471
   2.750   0.5804   0.01037   0.00391  -0.0169   0.4556   0.0477
   3.250   0.6356   0.00910   0.00255  -0.0166   0.4475   0.0341
   3.500   0.6634   0.00892   0.00235  -0.0167   0.4433   0.0333
   3.750   0.6914   0.00878   0.00222  -0.0168   0.4392   0.0329
   4.000   0.7196   0.00872   0.00213  -0.0170   0.4327   0.0331
   4.250   0.7478   0.00867   0.00208  -0.0172   0.4250   0.0342
   4.500   0.7761   0.00870   0.00208  -0.0175   0.4169   0.0363
   4.750   0.8044   0.00872   0.00208  -0.0177   0.4066   0.0394
   5.000   0.8327   0.00881   0.00214  -0.0181   0.3919   0.0435
   5.250   0.8606   0.00888   0.00231  -0.0185   0.3719   0.1380
   5.500   0.8993   0.00782   0.00275  -0.0217   0.3248   0.9952
   5.750   0.9578   0.01166   0.00507  -0.0323   0.0186   1.0000
   6.000   0.9830   0.01202   0.00550  -0.0321   0.0149   1.0000
   6.250   1.0080   0.01238   0.00590  -0.0320   0.0141   1.0000
   6.500   1.0328   0.01277   0.00634  -0.0318   0.0133   1.0000
   6.750   1.0572   0.01325   0.00687  -0.0316   0.0121   1.0000
   7.000   1.0811   0.01379   0.00745  -0.0314   0.0109   1.0000
   7.250   1.1040   0.01457   0.00830  -0.0312   0.0101   1.0000
   7.500   1.1238   0.01600   0.00989  -0.0310   0.0094   1.0000
   7.750   1.1421   0.01728   0.01127  -0.0306   0.0090   1.0000
   8.000   1.1631   0.01788   0.01192  -0.0300   0.0086   1.0000
   8.250   1.1810   0.01885   0.01296  -0.0293   0.0082   1.0000
   8.500   1.1950   0.02014   0.01433  -0.0283   0.0079   1.0000
   8.750   1.2070   0.02147   0.01574  -0.0271   0.0076   1.0000
   9.000   1.2188   0.02270   0.01702  -0.0261   0.0072   1.0000
   9.250   1.2286   0.02425   0.01863  -0.0257   0.0069   1.0000
   9.500   1.2294   0.02612   0.02055  -0.0241   0.0068   1.0000
   9.750   1.2343   0.02793   0.02243  -0.0228   0.0067   1.0000
  10.000   1.2411   0.02965   0.02419  -0.0216   0.0066   1.0000
  10.250   1.2496   0.03125   0.02584  -0.0204   0.0065   1.0000
  10.500   1.2588   0.03279   0.02742  -0.0192   0.0063   1.0000
  10.750   1.2685   0.03431   0.02897  -0.0179   0.0061   1.0000
  11.000   1.2816   0.03571   0.03038  -0.0153   0.0059   1.0000
  11.500   1.3104   0.03993   0.03491  -0.0108   0.0053   1.0000
  17.000   0.8574   0.14013   0.13848  -0.0212   0.0071   1.0000
  17.250   0.8326   0.14836   0.14680  -0.0258   0.0074   1.0000
  17.500   0.8079   0.15742   0.15593  -0.0306   0.0077   1.0000
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