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GOE 590 AIRFOIL (goe590-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 590 AIRFOIL (goe590-il)
Reynolds number: 100,000
Max Cl/Cd: 52.22 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe590-il-100000.txt
Download as CSV file: xf-goe590-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 590 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4103   0.09538   0.09069  -0.0107   1.0000   0.0455
  -7.500  -0.4083   0.09267   0.08804  -0.0121   1.0000   0.0469
  -7.250  -0.4058   0.09014   0.08557  -0.0144   1.0000   0.0485
  -7.000  -0.3991   0.08823   0.08368  -0.0192   1.0000   0.0499
  -6.750  -0.3874   0.08697   0.08238  -0.0252   1.0000   0.0506
  -6.500  -0.3790   0.08239   0.07784  -0.0268   1.0000   0.0514
  -6.250  -0.3755   0.07671   0.07227  -0.0218   1.0000   0.0550
  -6.000  -0.3648   0.07354   0.06911  -0.0231   1.0000   0.0575
  -5.750  -0.3514   0.07058   0.06612  -0.0253   1.0000   0.0608
  -5.500  -0.3225   0.07082   0.06602  -0.0324   1.0000   0.0639
  -5.250  -0.3185   0.06464   0.06002  -0.0307   1.0000   0.0651
  -5.000  -0.3118   0.06074   0.05617  -0.0289   1.0000   0.0676
  -4.750  -0.2991   0.05787   0.05328  -0.0287   1.0000   0.0717
  -4.500  -0.2702   0.05765   0.05263  -0.0321   1.0000   0.0777
  -4.250  -0.2693   0.05260   0.04782  -0.0296   1.0000   0.0796
  -4.000  -0.2656   0.05014   0.04543  -0.0273   1.0000   0.0829
  -3.750  -0.2584   0.04953   0.04465  -0.0257   1.0000   0.0898
  -3.500  -0.2385   0.04586   0.04093  -0.0274   0.9947   0.0940
  -3.250  -0.1936   0.04195   0.03685  -0.0330   0.9830   0.1105
  -3.000  -0.1447   0.03880   0.03333  -0.0387   0.9700   0.1339
  -2.750  -0.1028   0.03514   0.02951  -0.0428   0.9572   0.1540
  -2.500  -0.0594   0.03214   0.02624  -0.0469   0.9440   0.1891
  -1.750   0.0825   0.02296   0.01515  -0.0531   0.8996   0.1006
  -1.500   0.1161   0.02121   0.01293  -0.0531   0.8791   0.1009
  -1.250   0.1485   0.01968   0.01096  -0.0528   0.8603   0.1006
  -1.000   0.1776   0.01849   0.00944  -0.0521   0.8404   0.1014
  -0.750   0.2057   0.01763   0.00829  -0.0513   0.8219   0.1054
  -0.500   0.2321   0.01703   0.00761  -0.0506   0.8053   0.1175
  -0.250   0.2582   0.01627   0.00683  -0.0497   0.7902   0.1242
   0.000   0.2840   0.01573   0.00623  -0.0488   0.7757   0.1381
   0.250   0.3089   0.01492   0.00559  -0.0478   0.7629   0.2121
   0.500   0.3992   0.01296   0.00505  -0.0604   0.7495   1.0000
   0.750   0.4227   0.01323   0.00515  -0.0595   0.7372   1.0000
   1.000   0.4464   0.01351   0.00530  -0.0587   0.7258   1.0000
   1.250   0.4702   0.01381   0.00549  -0.0580   0.7157   1.0000
   1.500   0.4943   0.01411   0.00566  -0.0572   0.7072   1.0000
   1.750   0.5181   0.01443   0.00595  -0.0565   0.6975   1.0000
   2.000   0.5421   0.01477   0.00626  -0.0559   0.6890   1.0000
   2.250   0.5664   0.01510   0.00659  -0.0552   0.6818   1.0000
   2.500   0.5903   0.01549   0.00703  -0.0548   0.6736   1.0000
   2.750   0.6147   0.01583   0.00734  -0.0541   0.6673   1.0000
   3.000   0.6384   0.01627   0.00790  -0.0537   0.6595   1.0000
   3.250   0.6629   0.01665   0.00831  -0.0530   0.6538   1.0000
   3.500   0.6864   0.01716   0.00897  -0.0526   0.6465   1.0000
   3.750   0.7109   0.01758   0.00955  -0.0520   0.6409   1.0000
   4.000   0.7339   0.01802   0.01016  -0.0513   0.6321   1.0000
   4.250   0.7564   0.01758   0.00965  -0.0488   0.6124   1.0000
   4.500   0.7778   0.01741   0.00960  -0.0468   0.5921   1.0000
   4.750   0.7971   0.01645   0.00850  -0.0432   0.5517   1.0000
   5.000   0.8115   0.01554   0.00752  -0.0386   0.4783   1.0000
   5.250   0.8044   0.01824   0.00807  -0.0330   0.0998   1.0000
   5.500   0.8179   0.02004   0.00987  -0.0306   0.0766   1.0000
   5.750   0.8318   0.02150   0.01150  -0.0282   0.0672   1.0000
   6.000   0.8428   0.02317   0.01315  -0.0257   0.0595   1.0000
   6.250   0.8598   0.02466   0.01465  -0.0234   0.0561   1.0000
   6.500   0.8818   0.02647   0.01641  -0.0218   0.0541   1.0000
   6.750   0.9083   0.02867   0.01863  -0.0207   0.0530   1.0000
   7.000   0.9354   0.03262   0.02252  -0.0207   0.0488   1.0000
   7.250   0.9609   0.03551   0.02564  -0.0197   0.0488   1.0000
   7.500   0.9845   0.03661   0.02744  -0.0171   0.0531   1.0000
   7.750   1.0084   0.04134   0.03250  -0.0158   0.0596   1.0000
  11.250   0.7297   0.12009   0.11550  -0.0254   0.1941   1.0000
  11.500   0.7210   0.12412   0.11950  -0.0266   0.1824   1.0000
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