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EPPLER 598 AIRFOIL (e598-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: EPPLER 598 AIRFOIL (e598-il)
Reynolds number: 50,000
Max Cl/Cd: 9.85 at α=0°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-e598-il-50000.txt
Download as CSV file: xf-e598-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: EPPLER 598 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3190   0.11629   0.11056  -0.0065   1.0000   0.1559
  -9.000  -0.3172   0.11423   0.10863  -0.0092   1.0000   0.1617
  -8.750  -0.3318   0.11468   0.10926  -0.0144   1.0000   0.1640
  -8.500  -0.2957   0.10663   0.10124  -0.0123   1.0000   0.1732
  -8.250  -0.3005   0.10557   0.10035  -0.0166   1.0000   0.1797
  -8.000  -0.2825   0.10053   0.09541  -0.0177   1.0000   0.1856
  -7.750  -0.2759   0.09813   0.09316  -0.0206   1.0000   0.1946
  -7.500  -0.3000   0.09917   0.09445  -0.0216   1.0000   0.1966
  -7.250  -0.3298   0.10059   0.09601  -0.0197   0.9937   0.1970
  -7.000  -0.2863   0.09361   0.08899  -0.0236   0.9831   0.2095
  -6.750  -0.2604   0.08882   0.08420  -0.0308   0.9680   0.2209
  -6.500  -0.2338   0.08424   0.07959  -0.0385   0.9544   0.2364
  -6.250  -0.2044   0.07985   0.07517  -0.0435   0.9424   0.2574
  -6.000  -0.1910   0.07718   0.07245  -0.0495   0.9276   0.2804
  -5.500  -0.1541   0.07037   0.06562  -0.0508   0.9040   0.3268
  -5.250  -0.1493   0.06864   0.06387  -0.0515   0.8925   0.3647
  -5.000  -0.1286   0.06517   0.06044  -0.0469   0.8828   0.4050
  -4.250  -0.0765   0.05043   0.04412  -0.0807   0.8509   0.1785
  -4.000  -0.0556   0.04710   0.04046  -0.0815   0.8431   0.1730
  -3.750  -0.0378   0.04456   0.03749  -0.0816   0.8346   0.1737
  -3.500  -0.0144   0.04213   0.03447  -0.0819   0.8276   0.1778
  -3.250   0.0026   0.04052   0.03253  -0.0812   0.8193   0.1854
  -3.000   0.0315   0.03891   0.03024  -0.0813   0.8128   0.1978
  -2.750   0.0449   0.03819   0.02954  -0.0800   0.8045   0.2093
  -2.500   0.0749   0.03697   0.02798  -0.0801   0.7983   0.2264
  -2.250   0.0894   0.03668   0.02751  -0.0790   0.7905   0.2397
  -2.000   0.1168   0.03603   0.02662  -0.0789   0.7838   0.2600
  -1.750   0.1390   0.03584   0.02622  -0.0785   0.7772   0.2810
  -1.500   0.1604   0.03566   0.02602  -0.0781   0.7701   0.3047
  -1.250   0.1989   0.03509   0.02534  -0.0791   0.7642   0.3473
  -1.000   0.2105   0.03550   0.02586  -0.0782   0.7570   0.3818
  -0.750   0.3023   0.03233   0.02443  -0.0868   0.7498   1.0000
  -0.500   0.3116   0.03367   0.02540  -0.0854   0.7431   1.0000
  -0.250   0.3271   0.03481   0.02621  -0.0844   0.7365   1.0000
   0.000   0.3521   0.03574   0.02682  -0.0840   0.7307   1.0000
   0.250   0.3512   0.03741   0.02833  -0.0820   0.7242   1.0000
   0.500   0.3718   0.03851   0.02920  -0.0815   0.7182   1.0000
   0.750   0.3850   0.03992   0.03042  -0.0805   0.7126   1.0000
   1.000   0.3847   0.04164   0.03202  -0.0785   0.7073   1.0000
   1.250   0.4050   0.04289   0.03310  -0.0782   0.7016   1.0000
   1.500   0.4184   0.04442   0.03448  -0.0773   0.6965   1.0000
   1.750   0.4110   0.04631   0.03627  -0.0747   0.6931   1.0000
   2.000   0.4150   0.04803   0.03789  -0.0732   0.6898   1.0000
   2.250   0.4327   0.04959   0.03933  -0.0730   0.6853   1.0000
   2.500   0.4500   0.05128   0.04091  -0.0727   0.6805   1.0000
   2.750   0.4506   0.05330   0.04286  -0.0715   0.6797   1.0000
   3.000   0.4563   0.05539   0.04488  -0.0709   0.6800   1.0000
   3.250   0.4660   0.05756   0.04698  -0.0708   0.6810   1.0000
   3.750   0.3635   0.06391   0.05352  -0.0679   0.8303   1.0000
   4.000   0.3870   0.06608   0.05560  -0.0687   0.8139   1.0000
   4.250   0.4068   0.06800   0.05744  -0.0689   0.7979   1.0000
   4.500   0.4255   0.06988   0.05926  -0.0689   0.7820   1.0000
   5.000   0.4552   0.07339   0.06268  -0.0682   0.7519   1.0000
   5.250   0.4695   0.07531   0.06456  -0.0679   0.7376   1.0000
   5.500   0.4850   0.07743   0.06665  -0.0678   0.7245   1.0000
   5.750   0.5174   0.08092   0.07012  -0.0697   0.7146   1.0000
   6.000   0.5310   0.08268   0.07189  -0.0692   0.7000   1.0000
   6.250   0.5374   0.08422   0.07342  -0.0682   0.6860   1.0000
   6.500   0.5468   0.08627   0.07547  -0.0676   0.6741   1.0000
   6.750   0.5801   0.09014   0.07934  -0.0695   0.6650   1.0000
   7.000   0.5869   0.09160   0.08082  -0.0685   0.6508   1.0000
   7.250   0.5892   0.09329   0.08253  -0.0675   0.6385   1.0000
   7.500   0.6088   0.09647   0.08575  -0.0681   0.6299   1.0000
   8.000   0.6266   0.10067   0.09001  -0.0674   0.6050   1.0000
   8.250   0.6488   0.10431   0.09369  -0.0682   0.5974   1.0000
   8.500   0.6612   0.10653   0.09596  -0.0680   0.5842   1.0000
   8.750   0.6595   0.10842   0.09791  -0.0673   0.5731   1.0000
   9.000   0.6908   0.11292   0.10247  -0.0687   0.5652   1.0000
   9.250   0.6879   0.11419   0.10379  -0.0678   0.5523   1.0000
   9.500   0.6900   0.11669   0.10633  -0.0676   0.5427   1.0000
   9.750   0.7237   0.12128   0.11102  -0.0688   0.5333   1.0000
  10.000   0.7113   0.12233   0.11212  -0.0680   0.5218   1.0000
  10.250   0.7251   0.12594   0.11581  -0.0685   0.5141   1.0000
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