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E230 (9.96%) (e230-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: E230 (9.96%) (e230-il)
Reynolds number: 50,000
Max Cl/Cd: 23.03 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e230-il-50000.txt
Download as CSV file: xf-e230-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: E230  (9.96%)                                   
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.7498   0.11143   0.10501   0.0219   1.0000   0.1377
 -10.250  -0.7918   0.09498   0.08862   0.0078   1.0000   0.1091
 -10.000  -0.7902   0.08960   0.08326   0.0065   1.0000   0.1074
  -9.750  -0.8034   0.08326   0.07692   0.0036   1.0000   0.1046
  -9.500  -0.8877   0.07535   0.06854   0.0007   1.0000   0.0961
  -9.250  -0.8817   0.07015   0.06330   0.0009   1.0000   0.0948
  -9.000  -0.8864   0.06531   0.05826   0.0016   1.0000   0.0936
  -8.750  -0.8915   0.06055   0.05318   0.0027   1.0000   0.0928
  -8.500  -0.8932   0.05600   0.04822   0.0041   1.0000   0.0925
  -8.250  -0.8906   0.05174   0.04347   0.0058   1.0000   0.0932
  -8.000  -0.8838   0.04775   0.03891   0.0077   1.0000   0.0943
  -7.750  -0.8735   0.04414   0.03460   0.0097   1.0000   0.0960
  -7.500  -0.8534   0.04036   0.03072   0.0105   1.0000   0.0996
  -7.250  -0.8318   0.03723   0.02726   0.0117   1.0000   0.1052
  -7.000  -0.8098   0.03440   0.02424   0.0128   1.0000   0.1158
  -6.750  -0.7870   0.03176   0.02140   0.0140   1.0000   0.1348
  -6.500  -0.7683   0.02933   0.01927   0.0157   1.0000   0.1728
  -6.250  -0.7568   0.02763   0.01795   0.0184   1.0000   0.2322
  -6.000  -0.7432   0.02604   0.01669   0.0213   1.0000   0.2962
  -5.750  -0.7271   0.02478   0.01579   0.0240   1.0000   0.3577
  -5.500  -0.7099   0.02376   0.01505   0.0267   1.0000   0.4180
  -5.250  -0.6913   0.02290   0.01440   0.0294   1.0000   0.4765
  -5.000  -0.6716   0.02221   0.01389   0.0322   1.0000   0.5337
  -4.750  -0.6509   0.02167   0.01350   0.0350   1.0000   0.5899
  -4.500  -0.6258   0.02142   0.01339   0.0376   1.0000   0.6438
  -4.250  -0.5992   0.02134   0.01331   0.0402   1.0000   0.6962
  -4.000  -0.5632   0.02169   0.01360   0.0417   1.0000   0.7449
  -3.750  -0.5169   0.02237   0.01408   0.0417   1.0000   0.7898
  -3.500  -0.4599   0.02316   0.01454   0.0395   1.0000   0.8303
  -3.250  -0.3837   0.02394   0.01490   0.0333   1.0000   0.8662
  -3.000  -0.3197   0.02403   0.01465   0.0276   1.0000   0.9018
  -2.750  -0.2456   0.02364   0.01397   0.0191   1.0000   0.9340
  -2.500  -0.1744   0.02275   0.01282   0.0101   1.0000   0.9654
  -2.250  -0.1002   0.02134   0.01121  -0.0004   1.0000   0.9956
  -2.000  -0.0748   0.02042   0.01023  -0.0025   1.0000   1.0000
  -1.750  -0.0583   0.01975   0.00954  -0.0027   1.0000   1.0000
  -1.500  -0.0413   0.01919   0.00898  -0.0027   1.0000   1.0000
  -1.250  -0.0239   0.01871   0.00851  -0.0027   1.0000   1.0000
  -1.000  -0.0062   0.01833   0.00814  -0.0026   1.0000   1.0000
  -0.750   0.0113   0.01803   0.00788  -0.0023   1.0000   1.0000
  -0.500   0.0276   0.01781   0.00772  -0.0018   1.0000   1.0000
  -0.250   0.0410   0.01771   0.00769  -0.0009   1.0000   1.0000
   0.000   0.0466   0.01777   0.00781   0.0013   1.0000   1.0000
   0.250   0.0408   0.01802   0.00808   0.0052   1.0000   1.0000
   0.500   0.0309   0.01838   0.00843   0.0097   1.0000   1.0000
   0.750   0.0222   0.01879   0.00882   0.0138   1.0000   1.0000
   1.000   0.0167   0.01925   0.00924   0.0173   1.0000   1.0000
   1.250   0.0143   0.01976   0.00973   0.0203   1.0000   1.0000
   1.500   0.0965   0.02062   0.01083   0.0078   0.9748   1.0000
   1.750   0.2080   0.02116   0.01175  -0.0088   0.9410   1.0000
   2.000   0.3090   0.02114   0.01216  -0.0217   0.9029   1.0000
   2.250   0.3504   0.02125   0.01246  -0.0230   0.8630   1.0000
   2.500   0.3744   0.02140   0.01274  -0.0207   0.8267   1.0000
   2.750   0.3899   0.02161   0.01302  -0.0170   0.7908   1.0000
   3.000   0.4043   0.02173   0.01319  -0.0126   0.7570   1.0000
   3.250   0.4177   0.02179   0.01327  -0.0080   0.7240   1.0000
   3.500   0.4319   0.02195   0.01350  -0.0040   0.6870   1.0000
   3.750   0.4463   0.02198   0.01351   0.0005   0.6516   1.0000
   4.000   0.4620   0.02209   0.01362   0.0044   0.6117   1.0000
   4.250   0.4783   0.02219   0.01366   0.0082   0.5704   1.0000
   4.500   0.4954   0.02239   0.01377   0.0118   0.5250   1.0000
   4.750   0.5132   0.02269   0.01396   0.0150   0.4758   1.0000
   5.000   0.5313   0.02318   0.01425   0.0180   0.4228   1.0000
   5.250   0.5496   0.02386   0.01463   0.0207   0.3680   1.0000
   5.500   0.5681   0.02488   0.01544   0.0228   0.3111   1.0000
   5.750   0.5864   0.02626   0.01663   0.0249   0.2558   1.0000
   6.000   0.6029   0.02771   0.01778   0.0268   0.2009   1.0000
   6.250   0.6200   0.02957   0.01942   0.0285   0.1581   1.0000
   6.500   0.6395   0.03188   0.02162   0.0299   0.1342   1.0000
   6.750   0.6595   0.03438   0.02426   0.0312   0.1205   1.0000
   7.000   0.6784   0.03740   0.02779   0.0326   0.1138   1.0000
   7.250   0.6972   0.04045   0.03082   0.0336   0.1084   1.0000
   7.500   0.7084   0.04401   0.03512   0.0350   0.1057   1.0000
   7.750   0.7161   0.04789   0.03954   0.0361   0.1035   1.0000
   8.000   0.7195   0.05213   0.04422   0.0371   0.1027   1.0000
   8.250   0.7176   0.05683   0.04930   0.0378   0.1039   1.0000
   8.500   0.7119   0.06172   0.05445   0.0381   0.1054   1.0000
   8.750   0.7077   0.06647   0.05933   0.0383   0.1070   1.0000
   9.000   0.7137   0.07115   0.06405   0.0387   0.1086   1.0000
   9.250   0.6311   0.08111   0.07413   0.0306   0.1180   1.0000
   9.500   0.6192   0.08822   0.08120   0.0264   0.1242   1.0000
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