Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

EPPLER 904 AIRFOIL (e904-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 904 AIRFOIL (e904-il)
Reynolds number: 100,000
Max Cl/Cd: 41.95 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e904-il-100000-n5.txt
Download as CSV file: xf-e904-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 904 AIRFOIL                              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.750  -0.3731   0.11083   0.10595  -0.0363   1.0000   0.0166
 -10.500  -0.3746   0.10715   0.10231  -0.0369   1.0000   0.0158
 -10.250  -0.4626   0.11046   0.10527  -0.0345   1.0000   0.0176
 -10.000  -0.4632   0.10659   0.10146  -0.0352   1.0000   0.0165
  -9.750  -0.4639   0.10294   0.09786  -0.0361   1.0000   0.0156
  -9.500  -0.4663   0.09895   0.09394  -0.0371   1.0000   0.0149
  -9.250  -0.4690   0.09519   0.09023  -0.0381   1.0000   0.0143
  -9.000  -0.4728   0.09141   0.08652  -0.0392   1.0000   0.0138
  -8.750  -0.4795   0.08724   0.08243  -0.0405   1.0000   0.0133
  -8.500  -0.4883   0.08315   0.07843  -0.0419   1.0000   0.0128
  -8.250  -0.4999   0.07960   0.07496  -0.0431   1.0000   0.0127
  -8.000  -0.5185   0.07657   0.07200  -0.0420   1.0000   0.0125
  -7.750  -0.5386   0.07431   0.06979  -0.0389   1.0000   0.0125
  -7.500  -0.5554   0.07164   0.06712  -0.0361   1.0000   0.0123
  -7.250  -0.5705   0.06871   0.06416  -0.0331   1.0000   0.0119
  -7.000  -0.5826   0.06533   0.06070  -0.0303   1.0000   0.0115
  -6.750  -0.5874   0.06148   0.05669  -0.0279   0.9997   0.0108
  -6.500  -0.5687   0.05751   0.05245  -0.0290   0.9956   0.0102
  -6.250  -0.5530   0.05421   0.04885  -0.0294   0.9919   0.0097
  -6.000  -0.5447   0.04963   0.04398  -0.0292   0.9875   0.0095
  -5.750  -0.5323   0.04532   0.03931  -0.0290   0.9838   0.0092
  -5.500  -0.5171   0.04146   0.03503  -0.0285   0.9804   0.0090
  -5.250  -0.5025   0.03803   0.03116  -0.0271   0.9764   0.0088
  -5.000  -0.4843   0.03485   0.02749  -0.0259   0.9733   0.0087
  -4.750  -0.4621   0.03184   0.02395  -0.0250   0.9709   0.0086
  -4.500  -0.4365   0.02923   0.02068  -0.0246   0.9690   0.0085
  -4.250  -0.4149   0.02713   0.01816  -0.0231   0.9663   0.0086
  -4.000  -0.3912   0.02535   0.01605  -0.0221   0.9639   0.0086
  -3.750  -0.3657   0.02374   0.01416  -0.0214   0.9616   0.0089
  -3.500  -0.3396   0.02253   0.01273  -0.0209   0.9595   0.0097
  -3.250  -0.3151   0.02109   0.01112  -0.0205   0.9576   0.0130
  -3.000  -0.2884   0.02043   0.01028  -0.0204   0.9556   0.0202
  -2.750  -0.2687   0.01946   0.00916  -0.0186   0.9522   0.0281
  -2.500  -0.2617   0.01626   0.00868  -0.0155   0.9497   0.5053
  -2.250  -0.2304   0.01635   0.00982  -0.0126   0.9500   0.8910
  -2.000  -0.1862   0.01677   0.00974  -0.0152   0.9496   0.9266
  -1.750  -0.1000   0.01741   0.00995  -0.0264   0.9550   0.9598
  -1.500  -0.0074   0.01759   0.00981  -0.0397   0.9608   0.9910
  -1.250   0.0419   0.01754   0.00958  -0.0449   0.9610   1.0000
  -1.000   0.0698   0.01755   0.00941  -0.0457   0.9578   1.0000
  -0.750   0.0939   0.01756   0.00933  -0.0457   0.9538   1.0000
  -0.500   0.1161   0.01758   0.00928  -0.0453   0.9490   1.0000
  -0.250   0.1442   0.01761   0.00926  -0.0460   0.9455   1.0000
   0.000   0.1713   0.01766   0.00927  -0.0465   0.9418   1.0000
   0.250   0.1917   0.01771   0.00930  -0.0456   0.9361   1.0000
   0.500   0.2218   0.01775   0.00935  -0.0466   0.9324   1.0000
   0.750   0.2470   0.01781   0.00942  -0.0467   0.9275   1.0000
   1.000   0.2711   0.01786   0.00950  -0.0464   0.9218   1.0000
   1.250   0.3059   0.01788   0.00958  -0.0483   0.9181   1.0000
   1.500   0.3263   0.01793   0.00970  -0.0472   0.9105   1.0000
   1.750   0.3638   0.01788   0.00976  -0.0494   0.9058   1.0000
   2.000   0.3888   0.01787   0.01002  -0.0490   0.8974   1.0000
   2.250   0.4370   0.01761   0.00996  -0.0530   0.8920   1.0000
   2.500   0.4685   0.01741   0.00995  -0.0536   0.8815   1.0000
   2.750   0.5019   0.01718   0.00994  -0.0544   0.8713   1.0000
   3.000   0.5388   0.01691   0.00996  -0.0559   0.8624   1.0000
   3.250   0.5733   0.01665   0.01003  -0.0567   0.8525   1.0000
   3.500   0.6141   0.01464   0.00678  -0.0522   0.4676   1.0000
   3.750   0.5973   0.01771   0.00751  -0.0446   0.0747   1.0000
   4.000   0.6060   0.01912   0.00856  -0.0409   0.0249   1.0000
   4.250   0.6189   0.01996   0.00959  -0.0378   0.0192   1.0000
   4.500   0.6292   0.02096   0.01077  -0.0343   0.0172   1.0000
   4.750   0.6367   0.02230   0.01226  -0.0304   0.0159   1.0000
   5.000   0.6492   0.02373   0.01376  -0.0273   0.0152   1.0000
   5.250   0.6688   0.02535   0.01549  -0.0254   0.0146   1.0000
   5.500   0.6934   0.02737   0.01771  -0.0242   0.0142   1.0000
   5.750   0.7179   0.02966   0.02029  -0.0229   0.0136   1.0000
   6.000   0.7368   0.03142   0.02239  -0.0207   0.0116   1.0000
   6.250   0.7521   0.03319   0.02446  -0.0182   0.0100   1.0000
   6.500   0.7600   0.03469   0.02614  -0.0154   0.0076   1.0000
   6.750   0.7684   0.03737   0.02916  -0.0123   0.0072   1.0000
   7.000   0.7759   0.04007   0.03222  -0.0088   0.0071   1.0000
   7.250   0.7810   0.04289   0.03540  -0.0051   0.0071   1.0000
   7.500   0.7861   0.04546   0.03845  -0.0016   0.0072   1.0000
   7.750   0.7900   0.04812   0.04144   0.0019   0.0074   1.0000
   8.000   0.7902   0.05103   0.04467   0.0055   0.0074   1.0000
   8.250   0.7891   0.05404   0.04797   0.0091   0.0077   1.0000
   8.500   0.7834   0.05732   0.05152   0.0127   0.0078   1.0000
   8.750   0.7742   0.06056   0.05499   0.0164   0.0081   1.0000
   9.000   0.7606   0.06367   0.05828   0.0201   0.0083   1.0000
   9.250   0.7463   0.06708   0.06186   0.0227   0.0085   1.0000
   9.500   0.7301   0.07085   0.06576   0.0241   0.0086   1.0000
   9.750   0.7147   0.07502   0.07004   0.0242   0.0088   1.0000
  10.000   0.6994   0.07993   0.07504   0.0227   0.0090   1.0000
  10.250   0.6857   0.08572   0.08089   0.0195   0.0091   1.0000
  10.500   0.6763   0.09280   0.08799   0.0145   0.0094   1.0000
<< Back to EPPLER 904 AIRFOIL (e904-il)

Polar data table (+)

Polar graphs


<< Back to EPPLER 904 AIRFOIL (e904-il)