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EPPLER 49 AIRFOIL (e49-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 49 AIRFOIL (e49-il)
Reynolds number: 200,000
Max Cl/Cd: 82.29 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e49-il-200000.txt
Download as CSV file: xf-e49-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4131   0.12974   0.12666  -0.0114   1.0000   0.0172
  -8.250  -0.4197   0.12843   0.12539  -0.0107   1.0000   0.0173
  -8.000  -0.4259   0.12681   0.12381  -0.0094   1.0000   0.0173
  -7.750  -0.4300   0.12522   0.12225  -0.0092   0.9992   0.0174
  -7.500  -0.4186   0.12281   0.11985  -0.0137   0.9964   0.0175
  -7.000  -0.4109   0.11249   0.10956  -0.0140   0.9932   0.0184
  -6.750  -0.4022   0.10874   0.10581  -0.0139   0.9902   0.0190
  -6.500  -0.3886   0.10540   0.10247  -0.0166   0.9871   0.0195
  -6.250  -0.3729   0.10257   0.09964  -0.0204   0.9850   0.0201
  -6.000  -0.3684   0.09921   0.09629  -0.0212   0.9805   0.0205
  -5.750  -0.3475   0.09560   0.09267  -0.0264   0.9764   0.0213
  -5.500  -0.3187   0.09224   0.08928  -0.0335   0.9741   0.0224
  -5.250  -0.3105   0.08858   0.08562  -0.0351   0.9684   0.0230
  -5.000  -0.2754   0.08431   0.08130  -0.0435   0.9643   0.0243
  -4.750  -0.2252   0.08024   0.07714  -0.0555   0.9623   0.0263
  -4.500  -0.1730   0.07580   0.07256  -0.0680   0.9547   0.0281
  -4.250  -0.1050   0.07101   0.06751  -0.0821   0.9518   0.0285
  -4.000  -0.0862   0.06407   0.06061  -0.0855   0.9436   0.0304
  -3.750  -0.0456   0.06065   0.05712  -0.0913   0.9355   0.0357
  -3.500   0.0443   0.05653   0.05243  -0.1075   0.9316   0.0419
  -3.250   0.0840   0.05081   0.04669  -0.1145   0.9299   0.0445
  -3.000   0.1308   0.04833   0.04406  -0.1206   0.9283   0.0501
  -2.750   0.1795   0.04462   0.03992  -0.1268   0.9213   0.0587
  -2.500   0.2279   0.04239   0.03732  -0.1324   0.9191   0.0721
  -2.250   0.2707   0.04062   0.03541  -0.1369   0.9174   0.0875
  -1.250   0.4398   0.03392   0.02705  -0.1476   0.9056   0.0626
  -1.000   0.4708   0.03299   0.02582  -0.1477   0.8996   0.0500
  -0.750   0.5066   0.03230   0.02486  -0.1487   0.8959   0.0438
  -0.500   0.5450   0.03183   0.02431  -0.1507   0.8937   0.0426
  -0.250   0.5857   0.03113   0.02364  -0.1533   0.8923   0.0458
   0.000   0.6091   0.03076   0.02330  -0.1528   0.8837   0.0517
   0.250   0.6485   0.03043   0.02287  -0.1550   0.8812   0.0560
   0.500   0.6883   0.02852   0.02303  -0.1577   0.8802   1.0000
   0.750   0.7134   0.02835   0.02266  -0.1572   0.8713   1.0000
   1.000   0.7547   0.02795   0.02211  -0.1596   0.8683   1.0000
   1.250   0.7826   0.02764   0.02175  -0.1597   0.8601   1.0000
   1.500   0.8206   0.02717   0.02123  -0.1615   0.8563   1.0000
   1.750   0.8609   0.02668   0.02070  -0.1638   0.8542   1.0000
   2.000   0.8831   0.02651   0.02053  -0.1628   0.8445   1.0000
   2.250   0.9222   0.02593   0.01995  -0.1649   0.8419   1.0000
   2.500   0.9460   0.02571   0.01980  -0.1641   0.8326   1.0000
   2.750   0.9846   0.02497   0.01909  -0.1659   0.8293   1.0000
   3.000   1.0115   0.02444   0.01861  -0.1655   0.8196   1.0000
   3.250   1.0520   0.02329   0.01753  -0.1673   0.8155   1.0000
   3.500   1.0772   0.02280   0.01716  -0.1665   0.8049   1.0000
   3.750   1.1144   0.02176   0.01621  -0.1677   0.8010   1.0000
   4.000   1.1382   0.02134   0.01589  -0.1667   0.7891   1.0000
   4.250   1.1658   0.02078   0.01544  -0.1663   0.7770   1.0000
   4.500   1.1990   0.02001   0.01478  -0.1669   0.7640   1.0000
   4.750   1.2593   0.01759   0.01239  -0.1716   0.7270   1.0000
   5.000   1.3298   0.01616   0.01042  -0.1789   0.6243   1.0000
   5.250   1.3298   0.01739   0.01101  -0.1732   0.5182   1.0000
   5.500   1.3191   0.01921   0.01212  -0.1661   0.4120   1.0000
   5.750   1.3158   0.02095   0.01331  -0.1607   0.3228   1.0000
   6.000   1.3131   0.02301   0.01464  -0.1559   0.2129   1.0000
   6.500   1.3144   0.02778   0.01810  -0.1478   0.0242   1.0000
   6.750   1.3294   0.02886   0.01934  -0.1457   0.0213   1.0000
   7.000   1.3429   0.03008   0.02074  -0.1435   0.0196   1.0000
   7.250   1.3536   0.03151   0.02235  -0.1409   0.0187   1.0000
   7.500   1.3613   0.03321   0.02421  -0.1378   0.0182   1.0000
   7.750   1.3670   0.03516   0.02630  -0.1346   0.0177   1.0000
   8.000   1.3706   0.03789   0.02915  -0.1312   0.0167   1.0000
   8.250   1.3850   0.03926   0.03063  -0.1295   0.0159   1.0000
   8.500   1.4024   0.04128   0.03283  -0.1280   0.0156   1.0000
   8.750   1.4316   0.04373   0.03543  -0.1281   0.0156   1.0000
   9.000   1.5116   0.04997   0.04204  -0.1358   0.0168   1.0000
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