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NACA M18 AIRFOIL (m18-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: NACA M18 AIRFOIL (m18-il)
Reynolds number: 50,000
Max Cl/Cd: 11.08 at α=0.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-m18-il-50000.txt
Download as CSV file: xf-m18-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M18 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.3725   0.12135   0.11466  -0.0115   1.0000   0.1785
  -9.500  -0.3761   0.11938   0.11276  -0.0138   1.0000   0.1851
  -9.250  -0.4018   0.11973   0.11328  -0.0189   1.0000   0.1872
  -9.000  -0.3584   0.11186   0.10533  -0.0151   1.0000   0.1973
  -8.750  -0.3769   0.11116   0.10477  -0.0191   1.0000   0.2032
  -8.500  -0.3504   0.10545   0.09907  -0.0175   1.0000   0.2106
  -8.250  -0.3605   0.10377   0.09751  -0.0202   1.0000   0.2189
  -8.000  -0.3470   0.09939   0.09319  -0.0201   1.0000   0.2250
  -7.750  -0.3507   0.09713   0.09104  -0.0216   1.0000   0.2346
  -7.500  -0.3464   0.09362   0.08764  -0.0220   1.0000   0.2413
  -7.250  -0.3546   0.09179   0.08595  -0.0217   1.0000   0.2505
  -7.000  -0.3950   0.09253   0.08692  -0.0179   1.0000   0.2523
  -6.750  -0.4353   0.09343   0.08796  -0.0135   1.0000   0.2531
  -6.500  -0.4238   0.08976   0.08437  -0.0097   1.0000   0.2599
  -6.250  -0.4468   0.08917   0.08388  -0.0062   1.0000   0.2653
  -6.000  -0.4898   0.09002   0.08476  -0.0049   1.0000   0.2700
  -5.750  -0.4767   0.08608   0.08092  -0.0008   1.0000   0.2783
  -5.500  -0.5035   0.08529   0.08015  -0.0001   1.0000   0.2878
  -5.250  -0.4849   0.08134   0.07621  -0.0006   0.9945   0.3063
  -5.000  -0.4641   0.07777   0.07259  -0.0038   0.9844   0.3380
  -4.750   0.0721   0.05352   0.04768  -0.0332   1.0000   1.0000
  -4.500   0.0606   0.05402   0.04826  -0.0308   1.0000   1.0000
  -4.250   0.0494   0.05440   0.04871  -0.0285   1.0000   1.0000
  -4.000  -0.0186   0.05752   0.05201  -0.0149   0.9831   0.9349
  -3.750  -0.1190   0.05896   0.05362  -0.0034   0.9600   0.8040
  -3.500  -0.1946   0.05830   0.05312   0.0017   0.9418   0.7330
  -3.250  -0.2388   0.05633   0.05125   0.0039   0.9256   0.6932
  -3.000  -0.2899   0.05388   0.04883   0.0033   0.9093   0.6513
  -2.750  -0.1607   0.04854   0.04053  -0.0482   0.8922   0.2286
  -2.500  -0.1308   0.04697   0.03796  -0.0485   0.8811   0.1876
  -2.250  -0.0960   0.04467   0.03546  -0.0503   0.8722   0.1780
  -2.000  -0.0593   0.04324   0.03338  -0.0515   0.8628   0.1684
  -1.750  -0.0336   0.04221   0.03208  -0.0515   0.8530   0.1665
  -1.500   0.0115   0.04100   0.03042  -0.0542   0.8449   0.1689
  -1.250   0.0295   0.04079   0.02991  -0.0529   0.8345   0.1706
  -1.000   0.0814   0.03931   0.02829  -0.0567   0.8276   0.1755
  -0.750   0.0956   0.03944   0.02832  -0.0551   0.8174   0.1799
  -0.500   0.1607   0.03842   0.02707  -0.0605   0.8108   0.2020
  -0.250   0.1733   0.03878   0.02746  -0.0590   0.8006   0.2153
   0.000   0.2190   0.03761   0.02659  -0.0612   0.7947   0.2754
   0.250   0.2165   0.03778   0.02726  -0.0575   0.7860   0.3428
   0.500   0.3967   0.03580   0.02598  -0.0822   0.7790   1.0000
   0.750   0.3912   0.03743   0.02750  -0.0784   0.7697   1.0000
   1.000   0.4161   0.03820   0.02809  -0.0782   0.7630   1.0000
   1.250   0.4095   0.03993   0.02972  -0.0743   0.7553   1.0000
   1.500   0.4229   0.04107   0.03073  -0.0728   0.7483   1.0000
   1.750   0.4341   0.04237   0.03191  -0.0710   0.7421   1.0000
   2.000   0.4176   0.04425   0.03372  -0.0659   0.7360   1.0000
   2.250   0.4453   0.04522   0.03457  -0.0663   0.7299   1.0000
   2.500   0.4349   0.04701   0.03629  -0.0620   0.7248   1.0000
   2.750   0.4255   0.04879   0.03800  -0.0581   0.7202   1.0000
   3.000   0.4431   0.05010   0.03922  -0.0574   0.7147   1.0000
   3.250   0.4536   0.05164   0.04070  -0.0560   0.7095   1.0000
   3.500   0.4452   0.05359   0.04260  -0.0529   0.7062   1.0000
   3.750   0.4503   0.05529   0.04425  -0.0512   0.7018   1.0000
   4.000   0.4835   0.05655   0.04545  -0.0523   0.6946   1.0000
   4.250   0.4740   0.05868   0.04755  -0.0495   0.6921   1.0000
   4.500   0.4729   0.06070   0.04956  -0.0477   0.6895   1.0000
   4.750   0.4760   0.06277   0.05161  -0.0466   0.6879   1.0000
   5.000   0.4803   0.06490   0.05373  -0.0456   0.6868   1.0000
   5.250   0.4832   0.06740   0.05623  -0.0451   0.6906   1.0000
   5.500   0.4986   0.07009   0.05892  -0.0459   0.6945   1.0000
   6.250   0.4467   0.07848   0.06741  -0.0421   0.7652   1.0000
   6.500   0.4575   0.08043   0.06937  -0.0415   0.7539   1.0000
   6.750   0.4888   0.08374   0.07270  -0.0437   0.7439   1.0000
   7.250   0.5030   0.08689   0.07589  -0.0415   0.7178   1.0000
   7.500   0.5229   0.08979   0.07882  -0.0424   0.7095   1.0000
   8.000   0.5434   0.09367   0.08278  -0.0413   0.6844   1.0000
   8.250   0.5646   0.09693   0.08609  -0.0424   0.6774   1.0000
   8.500   0.5760   0.09890   0.08811  -0.0420   0.6648   1.0000
   8.750   0.5770   0.10069   0.08994  -0.0408   0.6537   1.0000
   9.000   0.6115   0.10498   0.09433  -0.0431   0.6459   1.0000
   9.250   0.6061   0.10590   0.09530  -0.0411   0.6331   1.0000
   9.500   0.6101   0.10811   0.09755  -0.0404   0.6224   1.0000
   9.750   0.6423   0.11233   0.10188  -0.0423   0.6128   1.0000
  10.000   0.6455   0.11388   0.10351  -0.0412   0.5988   1.0000
  10.250   0.6441   0.11574   0.10543  -0.0402   0.5865   1.0000
  10.500   0.6539   0.11851   0.10829  -0.0402   0.5751   1.0000
  10.750   0.6807   0.12251   0.11240  -0.0413   0.5636   1.0000
  11.000   0.6897   0.12483   0.11483  -0.0410   0.5495   1.0000
  11.250   0.6835   0.12667   0.11672  -0.0401   0.5375   1.0000
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