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EPPLER 472 AIRFOIL (e472-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 472 AIRFOIL (e472-il)
Reynolds number: 50,000
Max Cl/Cd: 17.46 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e472-il-50000.txt
Download as CSV file: xf-e472-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 472 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
  ------ -------- --------- --------- -------- -------- --------
 -12.000  -0.5697   0.14406   0.13544   0.0217   1.0000   0.2965
 -11.750  -0.5466   0.13932   0.13064   0.0220   1.0000   0.3042
 -11.500  -0.5781   0.14035   0.13177   0.0205   1.0000   0.3119
 -11.250  -0.5351   0.13315   0.12445   0.0214   1.0000   0.3206
 -11.000  -0.5581   0.13279   0.12418   0.0204   1.0000   0.3293
 -10.750  -0.5234   0.12707   0.11837   0.0210   1.0000   0.3382
 -10.500  -0.5414   0.12588   0.11725   0.0204   1.0000   0.3467
 -10.250  -0.5125   0.12113   0.11243   0.0208   1.0000   0.3568
 -10.000  -0.5130   0.11836   0.10968   0.0207   1.0000   0.3664
  -9.750  -0.5218   0.11712   0.10848   0.0210   1.0000   0.3792
  -9.500  -0.4931   0.11212   0.10342   0.0211   1.0000   0.3873
  -9.000  -0.4915   0.10740   0.09875   0.0218   1.0000   0.4122
  -8.750  -0.4715   0.10358   0.09490   0.0222   1.0000   0.4238
  -8.500  -0.4679   0.10084   0.09219   0.0226   1.0000   0.4360
  -8.000  -0.4678   0.09660   0.08801   0.0243   1.0000   0.4645
  -7.750  -0.4438   0.09289   0.08427   0.0246   1.0000   0.4786
  -7.500  -0.4319   0.09007   0.08147   0.0253   1.0000   0.4945
  -7.250  -0.6389   0.06912   0.06074   0.0031   1.0000   0.3123
  -7.000  -0.6386   0.06460   0.05616   0.0028   1.0000   0.3099
  -6.750  -0.6434   0.05984   0.05133   0.0025   1.0000   0.3078
  -6.500  -0.6516   0.05481   0.04616   0.0023   1.0000   0.3057
  -6.250  -0.6568   0.05042   0.04155   0.0028   1.0000   0.3070
  -6.000  -0.6597   0.04641   0.03721   0.0037   1.0000   0.3102
  -5.750  -0.6492   0.04370   0.03437   0.0049   1.0000   0.3143
  -5.500  -0.6332   0.04169   0.03230   0.0062   1.0000   0.3189
  -5.250  -0.6219   0.03932   0.02969   0.0076   1.0000   0.3245
  -5.000  -0.6099   0.03705   0.02716   0.0090   1.0000   0.3306
  -4.750  -0.5902   0.03566   0.02578   0.0103   1.0000   0.3371
  -4.500  -0.5753   0.03374   0.02358   0.0116   1.0000   0.3434
  -4.250  -0.5560   0.03232   0.02214   0.0130   1.0000   0.3504
  -4.000  -0.5373   0.03102   0.02072   0.0143   1.0000   0.3590
  -3.750  -0.5180   0.02967   0.01928   0.0156   1.0000   0.3666
  -3.500  -0.4973   0.02854   0.01811   0.0168   1.0000   0.3750
  -3.250  -0.4770   0.02744   0.01697   0.0180   1.0000   0.3846
  -3.000  -0.4564   0.02647   0.01589   0.0192   1.0000   0.3954
  -2.750  -0.4344   0.02554   0.01506   0.0204   1.0000   0.4053
  -2.500  -0.4129   0.02465   0.01413   0.0215   1.0000   0.4170
  -2.250  -0.3915   0.02389   0.01335   0.0226   1.0000   0.4310
  -2.000  -0.3691   0.02316   0.01276   0.0237   1.0000   0.4452
  -1.750  -0.3465   0.02246   0.01218   0.0247   1.0000   0.4610
  -1.500  -0.3242   0.02183   0.01165   0.0257   1.0000   0.4805
  -1.250  -0.3010   0.02122   0.01125   0.0266   1.0000   0.5031
  -1.000  -0.2785   0.02067   0.01092   0.0276   1.0000   0.5308
  -0.750  -0.2569   0.02013   0.01067   0.0288   1.0000   0.5669
  -0.500  -0.2369   0.01956   0.01057   0.0303   1.0000   0.6171
  -0.250  -0.2185   0.01888   0.01065   0.0326   1.0000   0.7008
   0.000   0.0000   0.01901   0.01158   0.0000   1.0000   1.0000
   0.250   0.2185   0.01888   0.01065  -0.0326   0.7006   1.0000
   0.500   0.2370   0.01956   0.01057  -0.0303   0.6170   1.0000
   0.750   0.2568   0.02013   0.01067  -0.0288   0.5666   1.0000
   1.000   0.2785   0.02067   0.01092  -0.0276   0.5308   1.0000
   1.250   0.3010   0.02122   0.01125  -0.0266   0.5032   1.0000
   1.500   0.3242   0.02183   0.01165  -0.0257   0.4805   1.0000
   1.750   0.3465   0.02246   0.01218  -0.0247   0.4611   1.0000
   2.000   0.3690   0.02316   0.01276  -0.0237   0.4453   1.0000
   2.250   0.3915   0.02389   0.01334  -0.0226   0.4311   1.0000
   2.500   0.4129   0.02465   0.01413  -0.0215   0.4170   1.0000
   2.750   0.4344   0.02554   0.01506  -0.0204   0.4053   1.0000
   3.000   0.4564   0.02647   0.01589  -0.0192   0.3953   1.0000
   3.250   0.4770   0.02744   0.01695  -0.0180   0.3845   1.0000
   3.500   0.4973   0.02854   0.01810  -0.0168   0.3750   1.0000
   3.750   0.5179   0.02967   0.01928  -0.0156   0.3666   1.0000
   4.000   0.5373   0.03102   0.02071  -0.0143   0.3590   1.0000
   4.250   0.5560   0.03232   0.02214  -0.0130   0.3505   1.0000
   4.500   0.5753   0.03374   0.02358  -0.0116   0.3433   1.0000
   4.750   0.5901   0.03565   0.02578  -0.0103   0.3371   1.0000
   5.000   0.6098   0.03705   0.02716  -0.0090   0.3306   1.0000
   5.250   0.6219   0.03932   0.02969  -0.0076   0.3245   1.0000
   5.500   0.6332   0.04169   0.03230  -0.0062   0.3189   1.0000
   5.750   0.6491   0.04371   0.03438  -0.0050   0.3144   1.0000
   6.000   0.6599   0.04639   0.03718  -0.0037   0.3103   1.0000
   6.250   0.6569   0.05042   0.04155  -0.0028   0.3071   1.0000
   6.500   0.6516   0.05483   0.04617  -0.0023   0.3057   1.0000
   6.750   0.6435   0.05984   0.05132  -0.0025   0.3077   1.0000
   7.000   0.6394   0.06456   0.05612  -0.0028   0.3100   1.0000
   7.250   0.6399   0.06905   0.06068  -0.0031   0.3122   1.0000
   7.500   0.4328   0.09008   0.08148  -0.0254   0.4948   1.0000
   7.750   0.4464   0.09302   0.08442  -0.0248   0.4794   1.0000
   8.000   0.4679   0.09655   0.08796  -0.0244   0.4644   1.0000
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