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EPPLER STE 87(-3)-914 AIRFOIL (ste87391-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER STE 87(-3)-914 AIRFOIL (ste87391-il)
Reynolds number: 50,000
Max Cl/Cd: 32.23 at α=8.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ste87391-il-50000.txt
Download as CSV file: xf-ste87391-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER STE 87(-3)-914 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4601   0.11485   0.10879  -0.0142   1.0019   0.2636
  -7.500  -0.5655   0.09704   0.09123  -0.0285   1.0019   0.1414
  -7.250  -0.5964   0.08715   0.08134  -0.0352   1.0019   0.1299
  -7.000  -0.6221   0.07573   0.06964  -0.0426   1.0019   0.1204
  -6.750  -0.6311   0.06310   0.05618  -0.0509   1.0019   0.1121
  -6.500  -0.6167   0.05803   0.05077  -0.0525   1.0019   0.1109
  -6.250  -0.5969   0.05248   0.04461  -0.0553   1.0019   0.1100
  -6.000  -0.5711   0.04754   0.03887  -0.0581   1.0019   0.1114
  -5.750  -0.5478   0.04524   0.03640  -0.0585   1.0019   0.1157
  -5.500  -0.5198   0.04244   0.03303  -0.0598   1.0019   0.1210
  -5.250  -0.4942   0.04051   0.03094  -0.0600   1.0019   0.1263
  -5.000  -0.4672   0.03877   0.02895  -0.0602   1.0019   0.1350
  -4.750  -0.4418   0.03775   0.02794  -0.0602   1.0019   0.1501
  -4.500  -0.4146   0.03665   0.02688  -0.0604   1.0019   0.1764
  -4.250  -0.3839   0.03552   0.02627  -0.0613   1.0019   0.2697
  -4.000  -0.3596   0.03555   0.02643  -0.0613   1.0019   0.3433
  -3.750  -0.3423   0.03590   0.02710  -0.0596   1.0019   0.3971
  -3.500  -0.3210   0.03589   0.02710  -0.0587   1.0019   0.4377
  -3.250  -0.2973   0.03555   0.02669  -0.0584   1.0019   0.4643
  -3.000  -0.2726   0.03524   0.02630  -0.0584   1.0019   0.4875
  -2.750  -0.2479   0.03500   0.02599  -0.0583   1.0019   0.5110
  -2.500  -0.2246   0.03483   0.02587  -0.0579   1.0019   0.5365
  -2.250  -0.2028   0.03479   0.02595  -0.0570   1.0019   0.5662
  -2.000  -0.1830   0.03491   0.02622  -0.0555   1.0019   0.6050
  -1.750  -0.1718   0.03527   0.02687  -0.0516   1.0019   0.6479
  -1.500  -0.1670   0.03585   0.02765  -0.0461   1.0019   0.7011
  -1.250  -0.1653   0.03641   0.02830  -0.0397   1.0019   0.7765
  -1.000  -0.1733   0.03633   0.02824  -0.0309   1.0019   0.8598
  -0.750  -0.1856   0.03518   0.02710  -0.0217   1.0019   0.9442
  -0.500  -0.1720   0.03397   0.02577  -0.0214   1.0019   0.9981
  -0.250  -0.1460   0.03442   0.02590  -0.0238   1.0019   0.9981
   0.000  -0.1206   0.03507   0.02628  -0.0257   1.0019   0.9981
   0.250  -0.0956   0.03582   0.02682  -0.0274   1.0019   0.9981
   0.500  -0.0711   0.03666   0.02745  -0.0289   1.0019   0.9981
   0.750  -0.0471   0.03757   0.02820  -0.0304   1.0019   0.9981
   1.000  -0.0235   0.03856   0.02905  -0.0319   1.0019   0.9981
   1.250  -0.0005   0.03962   0.02999  -0.0332   1.0019   0.9981
   1.500   0.0220   0.04076   0.03100  -0.0346   1.0019   0.9981
   1.750   0.0438   0.04196   0.03212  -0.0359   1.0019   0.9981
   2.000   0.0651   0.04324   0.03333  -0.0372   1.0019   0.9981
   2.250   0.0856   0.04459   0.03463  -0.0385   1.0019   0.9981
   2.500   0.1055   0.04603   0.03602  -0.0397   1.0019   0.9981
   2.750   0.1576   0.04929   0.03919  -0.0469   0.9864   0.9981
   3.000   0.2607   0.05256   0.04232  -0.0605   0.9156   0.9981
   3.250   0.3094   0.05336   0.04306  -0.0647   0.8820   0.9981
   3.500   0.3586   0.05427   0.04394  -0.0689   0.8590   0.9981
   3.750   0.3903   0.05473   0.04440  -0.0705   0.8363   0.9981
   4.000   0.4329   0.05523   0.04491  -0.0735   0.8164   0.9981
   4.250   0.4810   0.05556   0.04526  -0.0770   0.7984   0.9981
   4.500   0.5090   0.05584   0.04559  -0.0777   0.7774   0.9981
   4.750   0.5457   0.05599   0.04578  -0.0794   0.7578   0.9981
   5.000   0.5844   0.05594   0.04581  -0.0811   0.7391   0.9981
   5.250   0.6226   0.05573   0.04566  -0.0824   0.7205   0.9981
   5.500   0.6602   0.05538   0.04540  -0.0835   0.7021   0.9981
   5.750   0.6967   0.05496   0.04507  -0.0843   0.6838   0.9981
   6.000   0.7336   0.05439   0.04461  -0.0849   0.6656   0.9981
   6.250   0.7707   0.05368   0.04402  -0.0853   0.6480   0.9981
   6.500   0.8410   0.05008   0.04060  -0.0877   0.6392   0.9981
   6.750   0.8748   0.04911   0.03974  -0.0872   0.6214   0.9981
   7.000   0.9118   0.04794   0.03870  -0.0870   0.6040   0.9981
   7.250   0.9575   0.04616   0.03707  -0.0875   0.5878   0.9981
   7.500   1.0397   0.04186   0.03291  -0.0913   0.5741   0.9981
   7.750   1.1092   0.03949   0.03065  -0.0951   0.5542   0.9981
   8.000   1.1596   0.03896   0.03020  -0.0975   0.5331   0.9981
   8.250   1.2049   0.03911   0.03041  -0.0997   0.5130   0.9981
   8.500   1.2493   0.03960   0.03097  -0.1020   0.4944   0.9981
   8.750   1.2980   0.04027   0.03169  -0.1051   0.4766   0.9981
   9.000   1.2839   0.04275   0.03441  -0.0998   0.4651   0.9981
   9.250   1.3135   0.04408   0.03586  -0.1003   0.4513   0.9981
   9.500   1.3639   0.04490   0.03676  -0.1036   0.4357   0.9981
   9.750   1.3313   0.04806   0.04021  -0.0961   0.4285   0.9981
  10.000   1.3759   0.04888   0.04112  -0.0983   0.4139   0.9981
  10.250   1.3370   0.05265   0.04515  -0.0906   0.4079   0.9981
  10.500   1.3807   0.05321   0.04583  -0.0924   0.3930   0.9981
  10.750   1.3254   0.05826   0.05112  -0.0839   0.3895   0.9981
  11.000   1.2013   0.07075   0.06364  -0.0741   0.3916   0.9981
  11.250   0.9752   0.10552   0.09790  -0.0781   0.3944   0.9981
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