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GOE 598 AIRFOIL (goe598-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 598 AIRFOIL (goe598-il)
Reynolds number: 500,000
Max Cl/Cd: 66.9 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe598-il-500000.txt
Download as CSV file: xf-goe598-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 598 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6285   0.07673   0.07466  -0.0110   1.0000   0.0159
  -8.000  -0.6276   0.06874   0.06661  -0.0190   1.0000   0.0161
  -7.750  -0.6249   0.06221   0.05998  -0.0236   1.0000   0.0164
  -7.500  -0.6180   0.05727   0.05493  -0.0262   1.0000   0.0167
  -7.250  -0.6082   0.05287   0.05041  -0.0281   1.0000   0.0170
  -7.000  -0.5961   0.04857   0.04595  -0.0296   1.0000   0.0174
  -6.750  -0.5822   0.04433   0.04152  -0.0307   1.0000   0.0180
  -6.500  -0.5664   0.04012   0.03708  -0.0314   1.0000   0.0187
  -6.250  -0.5516   0.03024   0.02642  -0.0306   1.0000   0.0145
  -6.000  -0.5356   0.02452   0.02010  -0.0298   1.0000   0.0137
  -5.750  -0.5141   0.02120   0.01629  -0.0289   1.0000   0.0140
  -5.500  -0.4907   0.01903   0.01376  -0.0281   1.0000   0.0145
  -5.250  -0.4664   0.01785   0.01225  -0.0274   1.0000   0.0159
  -5.000  -0.4440   0.01503   0.00917  -0.0265   1.0000   0.0169
  -4.750  -0.4200   0.01385   0.00790  -0.0258   1.0000   0.0182
  -4.500  -0.3957   0.01299   0.00696  -0.0251   1.0000   0.0199
  -4.250  -0.3713   0.01221   0.00609  -0.0243   1.0000   0.0219
  -4.000  -0.3463   0.01216   0.00599  -0.0237   1.0000   0.0242
  -3.750  -0.3233   0.01050   0.00419  -0.0229   1.0000   0.0280
  -3.500  -0.2990   0.00988   0.00352  -0.0223   1.0000   0.0317
  -3.250  -0.2743   0.00949   0.00308  -0.0217   1.0000   0.0353
  -3.000  -0.2494   0.00907   0.00259  -0.0211   1.0000   0.0389
  -2.750  -0.2244   0.00876   0.00227  -0.0206   1.0000   0.0448
  -2.500  -0.1994   0.00850   0.00209  -0.0202   1.0000   0.0599
  -2.250  -0.1668   0.00831   0.00194  -0.0214   0.9980   0.0844
  -2.000  -0.1285   0.00808   0.00178  -0.0239   0.9946   0.1064
  -1.750  -0.0925   0.00752   0.00163  -0.0261   0.9902   0.2139
  -1.500  -0.0561   0.00675   0.00148  -0.0286   0.9860   0.3917
  -1.250  -0.0186   0.00611   0.00144  -0.0311   0.9822   0.5621
  -1.000   0.0148   0.00573   0.00139  -0.0324   0.9727   0.6556
  -0.750   0.0461   0.00536   0.00136  -0.0330   0.9608   0.7459
  -0.500   0.0718   0.00501   0.00136  -0.0320   0.9454   0.8481
  -0.250   0.0961   0.00484   0.00136  -0.0304   0.9278   0.9295
   0.000   0.1376   0.00480   0.00131  -0.0330   0.9082   0.9884
   0.250   0.1691   0.00481   0.00122  -0.0336   0.8780   1.0000
   0.500   0.1934   0.00490   0.00117  -0.0326   0.8486   1.0000
   0.750   0.2188   0.00500   0.00116  -0.0318   0.8237   1.0000
   1.000   0.2445   0.00511   0.00117  -0.0313   0.8016   1.0000
   1.250   0.2705   0.00523   0.00119  -0.0308   0.7782   1.0000
   1.500   0.2967   0.00534   0.00124  -0.0303   0.7529   1.0000
   1.750   0.3232   0.00545   0.00129  -0.0300   0.7309   1.0000
   2.000   0.3496   0.00559   0.00134  -0.0296   0.7041   1.0000
   2.250   0.3755   0.00577   0.00140  -0.0291   0.6667   1.0000
   2.500   0.4014   0.00600   0.00150  -0.0286   0.6203   1.0000
   2.750   0.4261   0.00641   0.00158  -0.0280   0.5231   1.0000
   3.000   0.4478   0.00757   0.00188  -0.0274   0.3159   1.0000
   3.250   0.4708   0.00869   0.00228  -0.0271   0.1466   1.0000
   3.500   0.4943   0.00986   0.00292  -0.0266   0.0280   1.0000
   3.750   0.5207   0.01031   0.00342  -0.0263   0.0225   1.0000
   4.000   0.5453   0.01132   0.00456  -0.0255   0.0190   1.0000
   4.250   0.5708   0.01200   0.00532  -0.0250   0.0182   1.0000
   4.500   0.5964   0.01258   0.00595  -0.0246   0.0163   1.0000
   4.750   0.6213   0.01346   0.00690  -0.0240   0.0152   1.0000
   5.000   0.6459   0.01451   0.00803  -0.0233   0.0142   1.0000
   5.250   0.6703   0.01588   0.00951  -0.0225   0.0135   1.0000
   5.500   0.6950   0.01807   0.01193  -0.0213   0.0140   1.0000
   8.750   0.7707   0.05674   0.05457  -0.0016   0.0131   1.0000
   9.000   0.7394   0.06363   0.06160  -0.0040   0.0135   1.0000
   9.250   0.7024   0.07616   0.07418  -0.0137   0.0145   1.0000
   9.500   0.6865   0.08414   0.08214  -0.0180   0.0148   1.0000
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