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EPPLER 585 AIRFOIL (e585-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 585 AIRFOIL (e585-il)
Reynolds number: 50,000
Max Cl/Cd: 8.01 at α=9.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e585-il-50000.txt
Download as CSV file: xf-e585-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 585 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  -0.4471   0.12259   0.11740  -0.0173   1.0000   0.2349
  -7.500  -0.4296   0.11865   0.11346  -0.0147   1.0000   0.2475
  -7.250  -0.4704   0.11878   0.11376  -0.0133   1.0000   0.2509
  -7.000  -0.4556   0.11522   0.11019  -0.0108   1.0000   0.2648
  -6.750  -0.4522   0.11238   0.10739  -0.0086   1.0000   0.2749
  -6.500  -0.4928   0.11222   0.10737  -0.0060   1.0000   0.2820
  -6.250  -0.4873   0.10931   0.10449  -0.0035   1.0000   0.2964
  -6.000  -0.4877   0.10681   0.10204  -0.0008   1.0000   0.3100
  -5.750  -0.4948   0.10468   0.09998   0.0021   1.0000   0.3237
  -5.500  -0.5059   0.10272   0.09809   0.0054   1.0000   0.3397
  -5.250  -0.5175   0.10092   0.09637   0.0088   1.0000   0.3575
  -5.000  -0.5519   0.09949   0.09504   0.0109   1.0000   0.3779
  -4.750  -0.5342   0.09708   0.09266   0.0162   1.0000   0.4078
  -3.500  -0.4556   0.05715   0.04975  -0.0384   1.0000   0.1276
  -3.250  -0.4249   0.05368   0.04541  -0.0396   1.0000   0.1110
  -3.000  -0.4034   0.05062   0.04224  -0.0398   1.0000   0.1075
  -2.750  -0.3781   0.04822   0.03939  -0.0403   1.0000   0.1060
  -2.500  -0.3525   0.04630   0.03699  -0.0406   1.0000   0.1068
  -2.250  -0.3273   0.04467   0.03492  -0.0405   1.0000   0.1078
  -2.000  -0.3022   0.04329   0.03312  -0.0401   1.0000   0.1094
  -1.750  -0.2793   0.04212   0.03180  -0.0396   1.0000   0.1152
  -1.500  -0.2569   0.04144   0.03091  -0.0388   1.0000   0.1253
  -1.250  -0.2363   0.04073   0.03020  -0.0373   1.0000   0.1404
  -1.000  -0.2156   0.04009   0.02971  -0.0359   1.0000   0.1683
  -0.750  -0.1822   0.03831   0.02909  -0.0373   1.0000   0.3059
  -0.500  -0.2147   0.03705   0.03046  -0.0193   1.0000   0.8811
  -0.250  -0.1961   0.03639   0.02934  -0.0169   1.0000   1.0000
   0.000  -0.1781   0.03690   0.02939  -0.0170   1.0000   1.0000
   0.250  -0.1591   0.03752   0.02962  -0.0173   1.0000   1.0000
   0.500  -0.1395   0.03823   0.03000  -0.0178   1.0000   1.0000
   0.750  -0.1195   0.03903   0.03051  -0.0185   1.0000   1.0000
   1.000  -0.0993   0.03990   0.03111  -0.0192   1.0000   1.0000
   1.250  -0.0791   0.04084   0.03179  -0.0200   1.0000   1.0000
   1.500  -0.0589   0.04184   0.03258  -0.0208   1.0000   1.0000
   1.750  -0.0388   0.04290   0.03344  -0.0216   1.0000   1.0000
   2.000  -0.0190   0.04403   0.03438  -0.0225   1.0000   1.0000
   2.250   0.0007   0.04520   0.03538  -0.0233   1.0000   1.0000
   2.500   0.0270   0.04690   0.03691  -0.0255   0.9968   1.0000
   2.750   0.0609   0.04918   0.03901  -0.0291   0.9886   1.0000
   3.000   0.0926   0.05140   0.04108  -0.0324   0.9798   1.0000
   3.250   0.1270   0.05397   0.04350  -0.0361   0.9705   1.0000
   3.500   0.1616   0.05650   0.04591  -0.0398   0.9589   1.0000
   3.750   0.1883   0.05822   0.04755  -0.0420   0.9461   1.0000
   4.000   0.2130   0.05990   0.04916  -0.0439   0.9331   1.0000
   4.250   0.2371   0.06166   0.05088  -0.0457   0.9200   1.0000
   4.500   0.2610   0.06354   0.05272  -0.0474   0.9073   1.0000
   4.750   0.2862   0.06562   0.05476  -0.0493   0.8948   1.0000
   5.000   0.3143   0.06802   0.05713  -0.0517   0.8825   1.0000
   5.250   0.3449   0.07061   0.05969  -0.0544   0.8690   1.0000
   5.500   0.3721   0.07287   0.06196  -0.0565   0.8545   1.0000
   5.750   0.3911   0.07448   0.06359  -0.0573   0.8400   1.0000
   6.000   0.4106   0.07629   0.06541  -0.0582   0.8253   1.0000
   6.250   0.4287   0.07813   0.06730  -0.0589   0.8110   1.0000
   6.500   0.4476   0.08010   0.06930  -0.0598   0.7966   1.0000
   6.750   0.4658   0.08214   0.07138  -0.0606   0.7824   1.0000
   7.000   0.4851   0.08427   0.07355  -0.0616   0.7677   1.0000
   7.250   0.5032   0.08644   0.07578  -0.0624   0.7535   1.0000
   7.500   0.5224   0.08863   0.07805  -0.0633   0.7384   1.0000
   7.750   0.5413   0.09087   0.08035  -0.0641   0.7231   1.0000
   8.000   0.6382   0.08703   0.07658  -0.0649   0.6310   1.0000
   8.250   0.6677   0.08823   0.07787  -0.0655   0.6140   1.0000
   8.500   0.6965   0.08941   0.07916  -0.0661   0.5979   1.0000
   8.750   0.7195   0.09079   0.08066  -0.0663   0.5828   1.0000
   9.000   0.7324   0.09267   0.08264  -0.0662   0.5687   1.0000
   9.250   0.7535   0.09409   0.08418  -0.0663   0.5537   1.0000
   9.500   0.7671   0.09600   0.08621  -0.0662   0.5395   1.0000
   9.750   0.7827   0.09779   0.08814  -0.0661   0.5250   1.0000
  10.000   0.7937   0.09999   0.09046  -0.0660   0.5113   1.0000
  10.250   0.8096   0.10178   0.09239  -0.0659   0.4967   1.0000
  10.500   0.8208   0.10408   0.09481  -0.0659   0.4831   1.0000
  10.750   0.8344   0.10614   0.09702  -0.0658   0.4690   1.0000
  11.000   0.8469   0.10836   0.09938  -0.0657   0.4554   1.0000
  11.250   0.8608   0.11044   0.10160  -0.0656   0.4414   1.0000
  11.500   0.8753   0.11246   0.10378  -0.0655   0.4277   1.0000
  11.750   0.8910   0.11434   0.10581  -0.0653   0.4140   1.0000
  12.000   0.9082   0.11598   0.10763  -0.0650   0.4001   1.0000
  12.250   0.9273   0.11732   0.10915  -0.0645   0.3861   1.0000
  12.500   0.9465   0.11845   0.11045  -0.0638   0.3719   1.0000
  12.750   0.9162   0.12657   0.11857  -0.0662   0.3627   1.0000
  13.000   0.9196   0.13012   0.12223  -0.0667   0.3505   1.0000
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