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EPPLER 625 AIRFOIL (e625-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 625 AIRFOIL (e625-il)
Reynolds number: 50,000
Max Cl/Cd: 18.78 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e625-il-50000.txt
Download as CSV file: xf-e625-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 625 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4097   0.12138   0.11468  -0.0047   1.0000   0.2241
 -10.000  -0.3876   0.11641   0.10973  -0.0037   1.0000   0.2332
  -9.750  -0.4162   0.11577   0.10924  -0.0064   1.0000   0.2417
  -9.500  -0.3832   0.11021   0.10365  -0.0044   1.0000   0.2541
  -9.250  -0.3815   0.10685   0.10037  -0.0047   1.0000   0.2636
  -9.000  -0.4084   0.10585   0.09954  -0.0064   1.0000   0.2754
  -8.750  -0.3881   0.10159   0.09530  -0.0050   1.0000   0.2910
  -8.500  -0.3746   0.09789   0.09165  -0.0040   1.0000   0.3062
  -8.250  -0.3742   0.09502   0.08887  -0.0033   1.0000   0.3237
  -8.000  -0.3904   0.09298   0.08698  -0.0027   1.0000   0.3424
  -7.750  -0.3521   0.08846   0.08244  -0.0005   1.0000   0.3656
  -7.500  -0.3766   0.08730   0.08146   0.0015   1.0000   0.3924
  -7.250  -0.3259   0.08305   0.07718   0.0045   1.0000   0.4339
  -7.000  -0.3054   0.08052   0.07471   0.0078   1.0000   0.4816
  -6.750  -0.2797   0.07807   0.07230   0.0110   1.0000   0.5365
  -6.500  -0.2433   0.07496   0.06920   0.0132   1.0000   0.5971
  -5.500  -0.3499   0.06410   0.05922   0.0160   1.0000   0.5144
  -5.250  -0.4007   0.06369   0.05908   0.0234   1.0000   0.5140
  -5.000  -0.4456   0.06356   0.05914   0.0309   1.0000   0.5167
  -4.750  -0.4966   0.06263   0.05837   0.0372   1.0000   0.5140
  -4.500  -0.5680   0.05251   0.04584   0.0180   1.0000   0.1795
  -4.250  -0.5488   0.04915   0.04189   0.0185   0.9965   0.1526
  -4.000  -0.5046   0.04548   0.03741   0.0151   0.9830   0.1391
  -3.750  -0.4609   0.04231   0.03370   0.0118   0.9693   0.1338
  -3.500  -0.4159   0.03991   0.03037   0.0094   0.9553   0.1277
  -3.250  -0.3680   0.03762   0.02766   0.0061   0.9408   0.1278
  -3.000  -0.3182   0.03567   0.02564   0.0020   0.9260   0.1349
  -2.750   0.1031   0.02714   0.01962  -0.0546   0.9448   1.0000
  -2.500   0.1828   0.02640   0.01835  -0.0646   0.9259   1.0000
  -2.250   0.2420   0.02589   0.01747  -0.0706   0.9041   1.0000
  -2.000   0.2843   0.02567   0.01697  -0.0734   0.8806   1.0000
  -1.750   0.3184   0.02559   0.01665  -0.0742   0.8590   1.0000
  -1.500   0.3386   0.02588   0.01677  -0.0729   0.8359   1.0000
  -1.250   0.3596   0.02612   0.01684  -0.0715   0.8151   1.0000
  -1.000   0.3803   0.02637   0.01692  -0.0699   0.7958   1.0000
  -0.750   0.3994   0.02666   0.01706  -0.0680   0.7776   1.0000
  -0.500   0.4175   0.02700   0.01725  -0.0660   0.7601   1.0000
  -0.250   0.4348   0.02741   0.01755  -0.0640   0.7426   1.0000
   0.000   0.4526   0.02781   0.01783  -0.0621   0.7262   1.0000
   0.250   0.4706   0.02818   0.01809  -0.0601   0.7106   1.0000
   0.500   0.4882   0.02866   0.01848  -0.0583   0.6948   1.0000
   0.750   0.5058   0.02919   0.01894  -0.0566   0.6794   1.0000
   1.000   0.5231   0.02976   0.01944  -0.0549   0.6643   1.0000
   1.250   0.5404   0.03039   0.02003  -0.0532   0.6499   1.0000
   1.500   0.5578   0.03104   0.02063  -0.0516   0.6363   1.0000
   1.750   0.5763   0.03153   0.02103  -0.0498   0.6239   1.0000
   2.000   0.5958   0.03190   0.02133  -0.0480   0.6122   1.0000
   2.250   0.6108   0.03290   0.02235  -0.0466   0.5986   1.0000
   2.500   0.6259   0.03387   0.02332  -0.0450   0.5861   1.0000
   2.750   0.6452   0.03436   0.02374  -0.0432   0.5759   1.0000
   3.000   0.6611   0.03523   0.02462  -0.0415   0.5645   1.0000
   3.250   0.6719   0.03670   0.02613  -0.0400   0.5526   1.0000
   3.500   0.6906   0.03728   0.02668  -0.0382   0.5433   1.0000
   3.750   0.7025   0.03858   0.02800  -0.0364   0.5327   1.0000
   4.000   0.7083   0.04051   0.03003  -0.0346   0.5222   1.0000
   4.250   0.7331   0.04049   0.02992  -0.0328   0.5142   1.0000
   4.500   0.7243   0.04376   0.03334  -0.0306   0.5034   1.0000
   4.750   0.7491   0.04383   0.03335  -0.0289   0.4960   1.0000
   5.000   0.7298   0.04781   0.03746  -0.0261   0.4859   1.0000
   5.250   0.7398   0.04921   0.03887  -0.0239   0.4782   1.0000
   5.500   0.7010   0.05451   0.04423  -0.0204   0.4702   1.0000
   5.750   0.7282   0.05458   0.04431  -0.0187   0.4632   1.0000
   6.000   0.6205   0.06409   0.05373  -0.0125   0.4597   1.0000
   6.250   0.5811   0.06864   0.05822  -0.0092   0.4568   1.0000
   6.500   0.5697   0.07182   0.06137  -0.0074   0.4535   1.0000
   6.750   0.5848   0.07350   0.06304  -0.0061   0.4471   1.0000
   7.000   0.5600   0.07779   0.06729  -0.0047   0.4463   1.0000
   7.250   0.5489   0.08155   0.07104  -0.0040   0.4467   1.0000
   7.500   0.5526   0.08498   0.07448  -0.0038   0.4484   1.0000
   7.750   0.4396   0.09764   0.08716  -0.0093   0.5609   1.0000
   8.000   0.4443   0.09930   0.08881  -0.0081   0.5481   1.0000
   8.250   0.4402   0.10120   0.09069  -0.0067   0.5393   1.0000
   8.500   0.4645   0.10451   0.09404  -0.0071   0.5308   1.0000
   8.750   0.4541   0.10586   0.09536  -0.0055   0.5208   1.0000
   9.000   0.4832   0.10992   0.09945  -0.0062   0.5128   1.0000
   9.250   0.4741   0.11080   0.10032  -0.0046   0.5010   1.0000
   9.500   0.4839   0.11393   0.10348  -0.0045   0.4950   1.0000
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