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GOE 590 AIRFOIL (goe590-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 590 AIRFOIL (goe590-il)
Reynolds number: 200,000
Max Cl/Cd: 67.61 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe590-il-200000.txt
Download as CSV file: xf-goe590-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4155   0.10070   0.09728  -0.0104   1.0000   0.0246
  -8.000  -0.4137   0.09840   0.09503  -0.0125   1.0000   0.0248
  -7.750  -0.4131   0.09609   0.09277  -0.0146   1.0000   0.0249
  -7.500  -0.4069   0.09339   0.09009  -0.0177   1.0000   0.0250
  -7.250  -0.3980   0.09045   0.08715  -0.0205   1.0000   0.0251
  -7.000  -0.3875   0.08727   0.08397  -0.0228   1.0000   0.0252
  -6.750  -0.3759   0.08402   0.08071  -0.0249   1.0000   0.0253
  -6.500  -0.3633   0.08066   0.07733  -0.0269   1.0000   0.0253
  -6.250  -0.3498   0.07718   0.07382  -0.0286   1.0000   0.0254
  -6.000  -0.3447   0.07122   0.06789  -0.0298   1.0000   0.0259
  -5.750  -0.3404   0.06649   0.06319  -0.0289   1.0000   0.0266
  -5.500  -0.3317   0.06305   0.05977  -0.0283   1.0000   0.0274
  -5.250  -0.3204   0.05997   0.05668  -0.0283   1.0000   0.0283
  -5.000  -0.3077   0.05700   0.05367  -0.0285   1.0000   0.0294
  -4.750  -0.2945   0.05417   0.05080  -0.0286   1.0000   0.0311
  -4.500  -0.2822   0.05171   0.04830  -0.0282   1.0000   0.0341
  -4.250  -0.2300   0.05008   0.04625  -0.0335   0.9932   0.0378
  -4.000  -0.2004   0.04284   0.03893  -0.0383   0.9818   0.0393
  -3.750  -0.1664   0.03920   0.03523  -0.0420   0.9687   0.0422
  -3.500  -0.1268   0.03598   0.03177  -0.0454   0.9514   0.0472
  -3.250  -0.0855   0.03218   0.02755  -0.0484   0.9328   0.0529
  -3.000  -0.0521   0.02978   0.02496  -0.0504   0.9106   0.0597
  -2.750  -0.0226   0.02757   0.02248  -0.0510   0.8850   0.0720
  -2.500   0.0032   0.02549   0.02002  -0.0503   0.8590   0.0815
  -2.250   0.0274   0.02391   0.01806  -0.0491   0.8331   0.0937
  -2.000   0.0568   0.02034   0.01363  -0.0467   0.8121   0.0654
  -1.750   0.0819   0.01830   0.01124  -0.0454   0.7909   0.0612
  -1.500   0.1078   0.01633   0.00882  -0.0441   0.7707   0.0584
  -1.250   0.1342   0.01506   0.00713  -0.0431   0.7521   0.0592
  -1.000   0.1603   0.01476   0.00646  -0.0422   0.7346   0.0641
  -0.750   0.1860   0.01355   0.00511  -0.0415   0.7181   0.0663
  -0.500   0.2114   0.01301   0.00450  -0.0408   0.7029   0.0700
  -0.250   0.2366   0.01265   0.00404  -0.0401   0.6894   0.0743
   0.000   0.2614   0.01232   0.00367  -0.0395   0.6771   0.0819
   0.250   0.2861   0.01205   0.00332  -0.0387   0.6660   0.0890
   0.500   0.3102   0.01164   0.00300  -0.0378   0.6550   0.1202
   0.750   0.4079   0.00948   0.00283  -0.0531   0.6423   1.0000
   1.000   0.4324   0.00960   0.00284  -0.0524   0.6332   1.0000
   1.250   0.4569   0.00974   0.00283  -0.0517   0.6254   1.0000
   1.500   0.4816   0.00986   0.00289  -0.0511   0.6166   1.0000
   1.750   0.5063   0.01001   0.00297  -0.0505   0.6098   1.0000
   2.000   0.5310   0.01015   0.00308  -0.0500   0.6025   1.0000
   2.250   0.5558   0.01033   0.00323  -0.0494   0.5962   1.0000
   2.500   0.5807   0.01049   0.00339  -0.0488   0.5894   1.0000
   2.750   0.6057   0.01070   0.00356  -0.0483   0.5842   1.0000
   3.000   0.6306   0.01088   0.00381  -0.0478   0.5779   1.0000
   3.250   0.6554   0.01109   0.00397  -0.0472   0.5711   1.0000
   3.500   0.6792   0.01120   0.00416  -0.0464   0.5592   1.0000
   3.750   0.7030   0.01132   0.00433  -0.0455   0.5472   1.0000
   4.000   0.7269   0.01146   0.00452  -0.0447   0.5356   1.0000
   4.250   0.7495   0.01152   0.00454  -0.0434   0.5159   1.0000
   4.500   0.7704   0.01141   0.00447  -0.0418   0.4821   1.0000
   4.750   0.7850   0.01161   0.00419  -0.0389   0.3268   1.0000
   5.000   0.7866   0.01479   0.00600  -0.0355   0.0564   1.0000
   5.250   0.8046   0.01583   0.00713  -0.0337   0.0433   1.0000
   5.500   0.8193   0.01718   0.00857  -0.0316   0.0378   1.0000
   5.750   0.8324   0.01878   0.01024  -0.0291   0.0356   1.0000
   6.000   0.8499   0.02003   0.01156  -0.0271   0.0339   1.0000
   6.250   0.8688   0.02114   0.01272  -0.0256   0.0296   1.0000
   6.500   0.8884   0.02279   0.01441  -0.0239   0.0286   1.0000
   6.750   0.9111   0.02483   0.01655  -0.0226   0.0284   1.0000
   7.000   0.9372   0.02758   0.01960  -0.0212   0.0304   1.0000
   7.250   0.9682   0.03119   0.02336  -0.0202   0.0388   1.0000
  11.000   0.8679   0.07730   0.07375   0.0061   0.0400   1.0000
  11.250   0.8311   0.08538   0.08198   0.0015   0.0416   1.0000
  11.500   0.8021   0.09385   0.09054  -0.0034   0.0431   1.0000
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