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EPPLER 593 AIRFOIL (e593-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: EPPLER 593 AIRFOIL (e593-il)
Reynolds number: 50,000
Max Cl/Cd: 16.64 at α=3.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-e593-il-50000-n5.txt
Download as CSV file: xf-e593-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 593 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3417   0.10687   0.10006  -0.0023   0.6507   0.0564
  -8.250  -0.3344   0.10327   0.09647  -0.0039   0.6463   0.0547
  -7.750  -0.3352   0.09482   0.08807  -0.0111   0.6405   0.0487
  -7.500  -0.3322   0.09120   0.08452  -0.0138   0.6367   0.0487
  -7.250  -0.3266   0.08754   0.08088  -0.0164   0.6329   0.0485
  -7.000  -0.3204   0.08352   0.07687  -0.0198   0.6295   0.0485
  -6.750  -0.3122   0.07961   0.07293  -0.0228   0.6264   0.0482
  -6.500  -0.3029   0.07540   0.06868  -0.0262   0.6232   0.0479
  -6.250  -0.2917   0.07099   0.06424  -0.0298   0.6195   0.0476
  -6.000  -0.2792   0.06635   0.05951  -0.0334   0.6162   0.0473
  -5.750  -0.2653   0.06135   0.05432  -0.0371   0.6131   0.0473
  -5.500  -0.2499   0.05564   0.04827  -0.0408   0.6106   0.0482
  -5.250  -0.2318   0.05145   0.04379  -0.0430   0.6077   0.0502
  -5.000  -0.2105   0.04880   0.04096  -0.0443   0.6040   0.0530
  -4.750  -0.1886   0.04442   0.03607  -0.0461   0.6008   0.0555
  -4.500  -0.1656   0.04067   0.03170  -0.0471   0.5977   0.0599
  -4.250  -0.1415   0.03881   0.02951  -0.0471   0.5946   0.0653
  -4.000  -0.1160   0.03650   0.02669  -0.0474   0.5914   0.0723
  -3.750  -0.0884   0.03451   0.02412  -0.0477   0.5876   0.0815
  -3.500  -0.0624   0.03355   0.02309  -0.0477   0.5841   0.0906
  -3.250  -0.0353   0.03240   0.02165  -0.0476   0.5811   0.1011
  -3.000  -0.0080   0.03142   0.02033  -0.0474   0.5784   0.1132
  -2.750   0.0200   0.03068   0.01933  -0.0474   0.5753   0.1266
  -2.500   0.0485   0.03014   0.01864  -0.0477   0.5716   0.1411
  -2.250   0.0768   0.02968   0.01805  -0.0479   0.5681   0.1568
  -2.000   0.1049   0.02929   0.01753  -0.0478   0.5651   0.1737
  -1.750   0.1328   0.02896   0.01711  -0.0477   0.5627   0.1923
  -1.500   0.1603   0.02875   0.01686  -0.0477   0.5601   0.2145
  -1.250   0.1878   0.02869   0.01686  -0.0482   0.5565   0.2412
  -1.000   0.2143   0.02855   0.01683  -0.0484   0.5531   0.2763
  -0.750   0.2395   0.02825   0.01677  -0.0482   0.5502   0.3328
  -0.500   0.2899   0.02619   0.01657  -0.0514   0.5474   1.0000
  -0.250   0.3157   0.02666   0.01661  -0.0510   0.5454   1.0000
   0.000   0.3414   0.02749   0.01724  -0.0516   0.5417   1.0000
   0.250   0.3665   0.02824   0.01777  -0.0519   0.5382   1.0000
   0.500   0.3912   0.02888   0.01821  -0.0518   0.5350   1.0000
   0.750   0.4159   0.02947   0.01859  -0.0515   0.5324   1.0000
   1.000   0.4407   0.03004   0.01895  -0.0511   0.5303   1.0000
   1.250   0.4645   0.03101   0.01981  -0.0514   0.5272   1.0000
   1.500   0.4872   0.03215   0.02092  -0.0521   0.5230   1.0000
   1.750   0.5103   0.03306   0.02173  -0.0521   0.5196   1.0000
   2.000   0.5338   0.03384   0.02240  -0.0519   0.5168   1.0000
   2.250   0.5578   0.03453   0.02297  -0.0514   0.5146   1.0000
   2.500   0.5785   0.03584   0.02426  -0.0519   0.5109   1.0000
   2.750   0.5971   0.03743   0.02588  -0.0526   0.5060   1.0000
   3.000   0.6182   0.03852   0.02692  -0.0525   0.5025   1.0000
   3.250   0.6408   0.03934   0.02769  -0.0521   0.4998   1.0000
   3.500   0.6649   0.03997   0.02825  -0.0514   0.4977   1.0000
   3.750   0.6729   0.04279   0.03119  -0.0529   0.4904   1.0000
   4.000   0.6913   0.04411   0.03251  -0.0528   0.4864   1.0000
   4.250   0.7138   0.04490   0.03327  -0.0522   0.4835   1.0000
   4.750   0.7345   0.04925   0.03773  -0.0530   0.4720   1.0000
   5.000   0.7552   0.05019   0.03866  -0.0524   0.4684   1.0000
   5.250   0.7809   0.05062   0.03907  -0.0514   0.4660   1.0000
   5.500   0.7702   0.05474   0.04329  -0.0529   0.4560   1.0000
   5.750   0.7918   0.05556   0.04412  -0.0521   0.4522   1.0000
   6.000   0.8142   0.05635   0.04496  -0.0513   0.4490   1.0000
   6.750   0.8108   0.06438   0.05309  -0.0512   0.4260   1.0000
   7.000   0.8318   0.06522   0.05398  -0.0505   0.4215   1.0000
   7.500   0.8428   0.06935   0.05822  -0.0497   0.4077   1.0000
   7.750   0.8716   0.06951   0.05845  -0.0487   0.4046   1.0000
   8.000   0.8577   0.07324   0.06221  -0.0491   0.3939   1.0000
   8.250   0.8857   0.07336   0.06242  -0.0481   0.3903   1.0000
   8.500   0.8747   0.07698   0.06608  -0.0485   0.3799   1.0000
   8.750   0.9021   0.07707   0.06629  -0.0474   0.3761   1.0000
   9.000   0.8933   0.08056   0.06983  -0.0479   0.3655   1.0000
   9.250   0.9209   0.08054   0.06993  -0.0468   0.3617   1.0000
   9.500   0.9122   0.08411   0.07355  -0.0473   0.3508   1.0000
  10.000   0.9318   0.08759   0.07723  -0.0468   0.3361   1.0000
  10.250   0.9316   0.09049   0.08022  -0.0472   0.3268   1.0000
  10.500   0.9521   0.09095   0.08079  -0.0463   0.3212   1.0000
  10.750   0.9485   0.09435   0.08428  -0.0469   0.3110   1.0000
  11.000   0.9739   0.09404   0.08413  -0.0457   0.3063   1.0000
  11.250   0.9664   0.09811   0.08827  -0.0467   0.2955   1.0000
  11.750   0.9866   0.10151   0.09190  -0.0463   0.2802   1.0000
  12.000   0.9838   0.10516   0.09564  -0.0472   0.2703   1.0000
  12.250   1.0080   0.10461   0.09527  -0.0459   0.2650   1.0000
  12.500   1.0001   0.10923   0.09997  -0.0473   0.2544   1.0000
  12.750   1.0302   0.10744   0.09837  -0.0454   0.2500   1.0000
  13.000   1.0186   0.11286   0.10385  -0.0472   0.2387   1.0000
  13.250   1.0165   0.11680   0.10787  -0.0484   0.2295   1.0000
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