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EPPLER 584 AIRFOIL (e584-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 584 AIRFOIL (e584-il)
Reynolds number: 50,000
Max Cl/Cd: 4.48 at α=11°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e584-il-50000.txt
Download as CSV file: xf-e584-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 584 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4685   0.13409   0.12922  -0.0169   1.0000   0.2277
  -8.000  -0.4321   0.12894   0.12402  -0.0161   0.9976   0.2387
  -7.750  -0.4754   0.12878   0.12400  -0.0152   0.9979   0.2448
  -7.500  -0.4432   0.12415   0.11933  -0.0156   0.9943   0.2585
  -7.000  -0.4679   0.12021   0.11552  -0.0117   0.9944   0.2769
  -6.750  -0.5176   0.12003   0.11548  -0.0057   1.0000   0.2790
  -6.500  -0.5744   0.11921   0.11479  -0.0033   1.0000   0.2802
  -6.250  -0.5153   0.11435   0.10989  -0.0010   1.0000   0.3029
  -6.000  -0.5496   0.11314   0.10877   0.0022   1.0000   0.3132
  -5.750  -0.5485   0.11053   0.10620   0.0051   1.0000   0.3292
  -5.500  -0.5562   0.10819   0.10391   0.0084   1.0000   0.3458
  -3.750  -0.4404   0.09477   0.09060   0.0337   1.0000   0.6492
  -3.500  -0.5996   0.07072   0.06491  -0.0102   1.0000   0.2268
  -3.250  -0.5535   0.06122   0.05387  -0.0146   1.0000   0.1230
  -3.000  -0.5295   0.05781   0.04956  -0.0143   1.0000   0.1109
  -2.750  -0.5106   0.05537   0.04680  -0.0138   1.0000   0.1097
  -2.500  -0.4903   0.05323   0.04427  -0.0133   1.0000   0.1088
  -2.250  -0.4694   0.05139   0.04205  -0.0127   1.0000   0.1079
  -2.000  -0.4477   0.04975   0.04003  -0.0120   1.0000   0.1074
  -1.750  -0.4258   0.04847   0.03837  -0.0112   1.0000   0.1084
  -1.500  -0.4038   0.04749   0.03698  -0.0103   1.0000   0.1134
  -1.250  -0.3838   0.04663   0.03614  -0.0096   1.0000   0.1208
  -1.000  -0.3624   0.04606   0.03537  -0.0085   1.0000   0.1298
  -0.750  -0.3417   0.04573   0.03506  -0.0072   1.0000   0.1461
  -0.500  -0.2122   0.04777   0.04056  -0.0152   1.0000   1.0000
  -0.250  -0.2036   0.04807   0.04048  -0.0132   1.0000   1.0000
   0.000  -0.1953   0.04837   0.04045  -0.0114   1.0000   1.0000
   0.250  -0.1871   0.04870   0.04051  -0.0096   1.0000   1.0000
   0.500  -0.1788   0.04903   0.04060  -0.0078   1.0000   1.0000
   0.750  -0.1705   0.04939   0.04074  -0.0061   1.0000   1.0000
   1.000  -0.1619   0.04978   0.04092  -0.0044   0.9999   1.0000
   1.250  -0.1340   0.05161   0.04246  -0.0067   0.9931   1.0000
   1.500  -0.1075   0.05318   0.04379  -0.0086   0.9841   1.0000
   1.750  -0.0849   0.05441   0.04481  -0.0098   0.9745   1.0000
   2.000  -0.0637   0.05569   0.04589  -0.0107   0.9648   1.0000
   2.250  -0.0402   0.05744   0.04746  -0.0120   0.9561   1.0000
   2.500  -0.0161   0.05900   0.04885  -0.0133   0.9446   1.0000
   2.750  -0.0001   0.05972   0.04944  -0.0132   0.9329   1.0000
   3.000   0.0172   0.06092   0.05051  -0.0133   0.9228   1.0000
   3.250   0.0464   0.06338   0.05281  -0.0154   0.9133   1.0000
   3.500   0.0579   0.06373   0.05308  -0.0145   0.9007   1.0000
   3.750   0.0751   0.06512   0.05437  -0.0146   0.8915   1.0000
   4.000   0.1028   0.06734   0.05647  -0.0165   0.8809   1.0000
   4.250   0.1127   0.06786   0.05693  -0.0154   0.8691   1.0000
   4.500   0.1429   0.07084   0.05979  -0.0178   0.8619   1.0000
   4.750   0.1546   0.07130   0.06022  -0.0171   0.8488   1.0000
   5.000   0.1673   0.07251   0.06138  -0.0167   0.8389   1.0000
   5.250   0.1990   0.07538   0.06418  -0.0192   0.8295   1.0000
   5.500   0.2047   0.07577   0.06456  -0.0178   0.8177   1.0000
   5.750   0.2332   0.07880   0.06753  -0.0199   0.8103   1.0000
   6.000   0.2463   0.07969   0.06842  -0.0196   0.7973   1.0000
   6.250   0.2567   0.08096   0.06968  -0.0191   0.7865   1.0000
   6.500   0.2924   0.08453   0.07323  -0.0222   0.7780   1.0000
   6.750   0.2953   0.08483   0.07354  -0.0207   0.7651   1.0000
   7.000   0.3091   0.08670   0.07542  -0.0209   0.7558   1.0000
   7.250   0.3399   0.08970   0.07842  -0.0232   0.7452   1.0000
   7.500   0.3430   0.09045   0.07921  -0.0220   0.7331   1.0000
   7.750   0.3615   0.09291   0.08169  -0.0229   0.7246   1.0000
   8.000   0.3859   0.09540   0.08420  -0.0244   0.7131   1.0000
   8.250   0.3901   0.09655   0.08539  -0.0236   0.7016   1.0000
   8.500   0.4055   0.09888   0.08776  -0.0242   0.6925   1.0000
   8.750   0.4355   0.10211   0.09106  -0.0263   0.6822   1.0000
   9.000   0.4374   0.10318   0.09217  -0.0255   0.6705   1.0000
   9.250   0.4465   0.10529   0.09434  -0.0257   0.6614   1.0000
   9.500   0.4791   0.10914   0.09825  -0.0281   0.6526   1.0000
   9.750   0.4780   0.11014   0.09932  -0.0273   0.6415   1.0000
  10.000   0.4857   0.11244   0.10170  -0.0276   0.6334   1.0000
  10.250   0.5188   0.11646   0.10580  -0.0299   0.6242   1.0000
  10.500   0.5135   0.11735   0.10676  -0.0290   0.6137   1.0000
  10.750   0.5229   0.11997   0.10945  -0.0296   0.6062   1.0000
  11.000   0.5546   0.12392   0.11352  -0.0318   0.5959   1.0000
  11.250   0.5475   0.12487   0.11454  -0.0311   0.5857   1.0000
  11.500   0.5597   0.12786   0.11762  -0.0320   0.5788   1.0000
  11.750   0.5847   0.13123   0.12109  -0.0335   0.5673   1.0000
  12.000   0.5784   0.13264   0.12257  -0.0333   0.5581   1.0000
  12.250   0.6031   0.13670   0.12676  -0.0351   0.5504   1.0000
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