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EPPLER 49 AIRFOIL (e49-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 49 AIRFOIL (e49-il)
Reynolds number: 500,000
Max Cl/Cd: 119.1 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e49-il-500000.txt
Download as CSV file: xf-e49-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 49 AIRFOIL                               
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -15.250  -0.3898   0.18771   0.18536  -0.0336   0.9856   0.0060
 -15.000  -0.3817   0.18528   0.18292  -0.0351   0.9853   0.0062
  -8.750  -0.2880   0.11661   0.11444  -0.0429   0.9587   0.0084
  -8.500  -0.2794   0.11344   0.11127  -0.0447   0.9547   0.0087
  -8.250  -0.2665   0.11024   0.10807  -0.0478   0.9527   0.0092
  -8.000  -0.2511   0.10722   0.10504  -0.0517   0.9514   0.0097
  -7.750  -0.2611   0.10540   0.10325  -0.0486   0.9428   0.0103
  -7.500  -0.2524   0.10299   0.10083  -0.0528   0.9337   0.0107
  -5.750  -0.1210   0.07278   0.07054  -0.0857   0.9085   0.0114
  -5.500  -0.0940   0.06881   0.06654  -0.0905   0.9070   0.0116
  -5.250  -0.0627   0.06504   0.06273  -0.0960   0.9059   0.0121
  -5.000  -0.0220   0.06075   0.05836  -0.1041   0.9052   0.0127
  -4.750   0.0006   0.05751   0.05507  -0.1075   0.8994   0.0133
  -4.500   0.0461   0.05313   0.05059  -0.1160   0.8977   0.0147
  -4.250   0.1093   0.04932   0.04658  -0.1260   0.8968   0.0173
  -4.000   0.1609   0.04519   0.04223  -0.1337   0.8962   0.0175
  -3.750   0.2215   0.03748   0.03411  -0.1454   0.8969   0.0185
  -3.500   0.2612   0.03479   0.03127  -0.1501   0.8962   0.0196
  -1.250   0.5930   0.02241   0.01676  -0.1693   0.8842   0.0297
  -1.000   0.6284   0.02152   0.01583  -0.1708   0.8831   0.0259
  -0.750   0.6645   0.02080   0.01507  -0.1725   0.8821   0.0251
   0.000   0.7854   0.01865   0.01278  -0.1798   0.8795   0.0350
   0.500   0.8363   0.01661   0.01276  -0.1795   0.8699   1.0000
   0.750   0.8716   0.01624   0.01235  -0.1811   0.8687   1.0000
   1.000   0.9082   0.01581   0.01189  -0.1829   0.8676   1.0000
   1.250   0.9456   0.01531   0.01137  -0.1849   0.8667   1.0000
   1.500   0.9604   0.01528   0.01135  -0.1822   0.8571   1.0000
   1.750   0.9989   0.01450   0.01057  -0.1842   0.8550   1.0000
   2.250   1.0504   0.01373   0.00985  -0.1832   0.8419   1.0000
   2.500   1.0683   0.01371   0.00985  -0.1812   0.8304   1.0000
   3.250   1.1587   0.01249   0.00871  -0.1825   0.7925   1.0000
   3.500   1.2422   0.01043   0.00616  -0.1936   0.7121   1.0000
   3.750   1.2584   0.01101   0.00642  -0.1913   0.6592   1.0000
   4.000   1.2568   0.01208   0.00709  -0.1853   0.5836   1.0000
   4.250   1.2508   0.01351   0.00799  -0.1788   0.4861   1.0000
   4.500   1.2507   0.01499   0.00888  -0.1738   0.3838   1.0000
   4.750   1.2581   0.01632   0.00966  -0.1705   0.2900   1.0000
   5.000   1.2648   0.01793   0.01056  -0.1673   0.1753   1.0000
   5.250   1.2669   0.02015   0.01191  -0.1633   0.0360   1.0000
   5.500   1.2828   0.02113   0.01276  -0.1614   0.0113   1.0000
   5.750   1.3019   0.02176   0.01351  -0.1600   0.0101   1.0000
   6.000   1.3203   0.02244   0.01431  -0.1585   0.0099   1.0000
   6.250   1.3375   0.02322   0.01523  -0.1568   0.0094   1.0000
   6.500   1.3539   0.02407   0.01618  -0.1551   0.0087   1.0000
   6.750   1.3689   0.02505   0.01726  -0.1532   0.0083   1.0000
   7.000   1.3823   0.02617   0.01850  -0.1509   0.0080   1.0000
   7.250   1.3939   0.02746   0.01990  -0.1484   0.0076   1.0000
   7.500   1.4034   0.02894   0.02149  -0.1456   0.0074   1.0000
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