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EPPLER STE 871-514 AIRFOIL (ste87151-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER STE 871-514 AIRFOIL (ste87151-il)
Reynolds number: 50,000
Max Cl/Cd: 28.03 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ste87151-il-50000.txt
Download as CSV file: xf-ste87151-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER STE 871-514 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4812   0.11776   0.11138  -0.0098   1.0020   0.3424
  -8.250  -0.4902   0.11555   0.10924  -0.0075   1.0020   0.3606
  -7.750  -0.4877   0.11041   0.10420  -0.0024   1.0020   0.4029
  -7.500  -0.4753   0.10736   0.10116  -0.0001   1.0020   0.4264
  -7.250  -0.4648   0.10452   0.09833   0.0023   1.0020   0.4509
  -7.000  -0.4586   0.10201   0.09585   0.0049   1.0020   0.4765
  -6.750  -0.4672   0.10042   0.09435   0.0091   1.0020   0.5079
  -5.250  -0.6251   0.05576   0.04747  -0.0230   1.0020   0.1472
  -5.000  -0.6012   0.05133   0.04251  -0.0230   1.0020   0.1299
  -4.750  -0.5756   0.04776   0.03824  -0.0228   1.0020   0.1184
  -4.500  -0.5488   0.04486   0.03459  -0.0223   1.0020   0.1109
  -4.250  -0.5237   0.04235   0.03173  -0.0217   1.0020   0.1079
  -4.000  -0.4980   0.04027   0.02925  -0.0210   1.0020   0.1070
  -3.750  -0.4729   0.03864   0.02727  -0.0201   1.0020   0.1097
  -3.500  -0.4474   0.03726   0.02552  -0.0191   1.0020   0.1128
  -3.250  -0.4238   0.03582   0.02413  -0.0177   1.0020   0.1162
  -3.000  -0.4004   0.03491   0.02312  -0.0158   1.0020   0.1217
  -2.750  -0.1756   0.03660   0.02789  -0.0281   1.0020   0.9980
  -2.500  -0.1669   0.03630   0.02730  -0.0258   1.0020   0.9980
  -2.250  -0.1581   0.03604   0.02679  -0.0236   1.0020   0.9980
  -2.000  -0.1493   0.03582   0.02635  -0.0214   1.0020   0.9980
  -1.750  -0.1404   0.03563   0.02597  -0.0192   1.0020   0.9980
  -1.500  -0.1315   0.03548   0.02563  -0.0169   1.0020   0.9980
  -1.250  -0.1227   0.03535   0.02534  -0.0147   1.0020   0.9980
  -1.000  -0.1139   0.03526   0.02511  -0.0124   1.0020   0.9980
  -0.750  -0.1052   0.03519   0.02491  -0.0102   1.0020   0.9980
  -0.500  -0.0966   0.03514   0.02475  -0.0079   1.0020   0.9980
  -0.250  -0.0882   0.03511   0.02462  -0.0057   1.0020   0.9980
   0.000  -0.0799   0.03511   0.02451  -0.0034   1.0020   0.9980
   0.250  -0.0718   0.03512   0.02444  -0.0011   1.0020   0.9980
   0.500  -0.0638   0.03515   0.02440   0.0011   1.0020   0.9980
   0.750  -0.0559   0.03520   0.02439   0.0033   1.0020   0.9980
   1.000  -0.0482   0.03527   0.02440   0.0055   1.0020   0.9980
   1.250  -0.0405   0.03535   0.02444   0.0077   1.0020   0.9980
   1.500  -0.0328   0.03546   0.02450   0.0098   1.0020   0.9980
   1.750  -0.0251   0.03558   0.02460   0.0119   1.0020   0.9980
   2.000  -0.0170   0.03575   0.02474   0.0138   1.0020   0.9980
   2.250  -0.0077   0.03599   0.02497   0.0154   1.0020   0.9980
   2.500   0.0031   0.03632   0.02529   0.0166   1.0020   0.9980
   2.750   0.0154   0.03676   0.02572   0.0175   1.0020   0.9980
   3.000   0.0291   0.03731   0.02627   0.0181   1.0020   0.9980
   3.250   0.0439   0.03796   0.02693   0.0183   1.0020   0.9980
   3.500   0.0595   0.03873   0.02771   0.0183   1.0020   0.9980
   3.750   0.0757   0.03959   0.02861   0.0180   1.0020   0.9980
   4.000   0.0923   0.04056   0.02961   0.0176   1.0020   0.9980
   4.250   0.1090   0.04162   0.03073   0.0170   1.0020   0.9980
   4.500   0.1257   0.04279   0.03196   0.0163   1.0020   0.9980
   4.750   0.1423   0.04406   0.03330   0.0155   1.0020   0.9980
   5.000   0.1635   0.04571   0.03503   0.0136   0.9997   0.9980
   5.250   0.3037   0.05119   0.04073  -0.0072   0.9036   0.9980
   5.500   0.3481   0.05197   0.04163  -0.0108   0.8671   0.9980
   5.750   0.3962   0.05273   0.04255  -0.0147   0.8385   0.9980
   6.000   0.4445   0.05303   0.04303  -0.0183   0.8118   0.9980
   6.250   0.4834   0.05282   0.04306  -0.0201   0.7839   0.9980
   6.500   0.5301   0.05210   0.04256  -0.0222   0.7554   0.9980
   6.750   0.5856   0.05027   0.04099  -0.0242   0.7259   0.9980
   7.000   0.6584   0.04616   0.03729  -0.0259   0.6960   0.9980
   7.250   0.7156   0.04164   0.03313  -0.0249   0.6628   0.9980
   7.500   0.7841   0.03502   0.02696  -0.0229   0.6233   0.9980
   7.750   0.8392   0.02994   0.02158  -0.0189   0.4986   0.9980
   8.000   0.8569   0.03111   0.02130  -0.0151   0.3591   0.9980
   8.250   0.8878   0.03308   0.02251  -0.0152   0.2883   0.9980
   8.500   0.9428   0.03512   0.02426  -0.0192   0.2438   0.9980
   8.750   1.0042   0.03754   0.02654  -0.0246   0.2168   0.9980
   9.000   1.0527   0.04014   0.02928  -0.0284   0.2008   0.9980
   9.250   1.0851   0.04270   0.03214  -0.0297   0.1911   0.9980
   9.500   1.1260   0.04570   0.03514  -0.0325   0.1821   0.9980
   9.750   1.1361   0.04810   0.03810  -0.0301   0.1781   0.9980
  10.000   1.1498   0.05074   0.04113  -0.0286   0.1737   0.9980
  10.250   1.1799   0.05402   0.04446  -0.0299   0.1685   0.9980
  10.500   1.1824   0.05717   0.04805  -0.0271   0.1669   0.9980
  10.750   1.1766   0.06024   0.05160  -0.0233   0.1658   0.9980
  11.000   1.1662   0.06347   0.05525  -0.0193   0.1649   0.9980
  11.250   1.1504   0.06677   0.05891  -0.0150   0.1646   0.9980
  11.500   1.1311   0.07039   0.06284  -0.0109   0.1648   0.9980
  11.750   1.1091   0.07438   0.06709  -0.0072   0.1654   0.9980
  12.000   1.0866   0.07881   0.07175  -0.0043   0.1662   0.9980
  12.250   1.0669   0.08373   0.07685  -0.0024   0.1672   0.9980
  12.500   0.6617   0.11641   0.10967  -0.0019   0.2112   0.9980
  12.750   0.6749   0.12077   0.11412  -0.0023   0.2142   0.9980
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