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GOE 598 AIRFOIL (goe598-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 598 AIRFOIL (goe598-il)
Reynolds number: 100,000
Max Cl/Cd: 43.07 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe598-il-100000.txt
Download as CSV file: xf-goe598-il-100000.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 =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6071   0.09943   0.09471   0.0047   1.0000   0.0957
  -8.250  -0.6189   0.09551   0.09089  -0.0037   1.0000   0.0984
  -8.000  -0.6307   0.09039   0.08567  -0.0179   1.0000   0.0992
  -7.750  -0.6066   0.08653   0.08192  -0.0039   1.0000   0.1041
  -7.500  -0.6025   0.08199   0.07738  -0.0086   1.0000   0.1093
  -7.250  -0.6051   0.07551   0.07084  -0.0196   1.0000   0.1141
  -7.000  -0.5910   0.07212   0.06747  -0.0161   1.0000   0.1181
  -6.750  -0.5871   0.06656   0.06165  -0.0243   1.0000   0.1276
  -6.500  -0.5721   0.06278   0.05797  -0.0221   1.0000   0.1314
  -6.250  -0.5610   0.05812   0.05310  -0.0256   1.0000   0.1424
  -6.000  -0.5469   0.05426   0.04905  -0.0272   1.0000   0.1553
  -5.750  -0.5315   0.05083   0.04560  -0.0268   1.0000   0.1701
  -5.500  -0.5168   0.04807   0.04278  -0.0265   1.0000   0.1969
  -5.250  -0.4717   0.03385   0.02624  -0.0310   1.0000   0.0711
  -5.000  -0.4491   0.02966   0.02171  -0.0305   1.0000   0.0677
  -4.750  -0.4241   0.02630   0.01779  -0.0296   1.0000   0.0655
  -4.500  -0.3983   0.02375   0.01477  -0.0287   1.0000   0.0664
  -4.250  -0.3722   0.02224   0.01278  -0.0277   1.0000   0.0728
  -4.000  -0.3472   0.02013   0.01059  -0.0270   1.0000   0.0774
  -3.750  -0.3214   0.01878   0.00904  -0.0261   1.0000   0.0844
  -3.500  -0.2971   0.01744   0.00774  -0.0253   1.0000   0.0973
  -3.250  -0.2734   0.01603   0.00644  -0.0243   1.0000   0.1110
  -3.000  -0.2506   0.01455   0.00527  -0.0234   1.0000   0.1534
  -2.750  -0.2331   0.01185   0.00457  -0.0221   1.0000   0.5261
  -2.500  -0.2180   0.01096   0.00447  -0.0179   1.0000   0.7576
  -2.250  -0.1795   0.01068   0.00435  -0.0179   1.0000   0.9355
  -2.000  -0.1227   0.01058   0.00392  -0.0239   1.0000   1.0000
  -1.750  -0.1043   0.01052   0.00367  -0.0226   1.0000   1.0000
  -1.500  -0.0845   0.01050   0.00345  -0.0214   1.0000   1.0000
  -1.250  -0.0638   0.01051   0.00332  -0.0204   1.0000   1.0000
  -1.000  -0.0427   0.01056   0.00324  -0.0195   1.0000   1.0000
  -0.750  -0.0212   0.01063   0.00321  -0.0187   1.0000   1.0000
  -0.500   0.0003   0.01074   0.00323  -0.0180   1.0000   1.0000
  -0.250   0.0219   0.01089   0.00329  -0.0174   1.0000   1.0000
   0.000   0.0434   0.01107   0.00342  -0.0168   1.0000   1.0000
   0.250   0.0649   0.01131   0.00361  -0.0164   1.0000   1.0000
   0.500   0.0864   0.01158   0.00386  -0.0160   1.0000   1.0000
   0.750   0.1079   0.01190   0.00417  -0.0158   1.0000   1.0000
   1.000   0.1294   0.01226   0.00454  -0.0157   1.0000   1.0000
   1.250   0.1509   0.01267   0.00495  -0.0157   1.0000   1.0000
   1.500   0.1916   0.01307   0.00541  -0.0194   0.9931   1.0000
   1.750   0.2485   0.01334   0.00581  -0.0258   0.9776   1.0000
   2.000   0.2995   0.01355   0.00621  -0.0308   0.9638   1.0000
   2.250   0.3552   0.01356   0.00644  -0.0363   0.9465   1.0000
   2.500   0.4124   0.01332   0.00651  -0.0413   0.9276   1.0000
   2.750   0.4514   0.01314   0.00657  -0.0425   0.9039   1.0000
   3.000   0.4858   0.01274   0.00638  -0.0416   0.8729   1.0000
   3.250   0.5083   0.01241   0.00617  -0.0380   0.8281   1.0000
   3.500   0.5211   0.01210   0.00562  -0.0314   0.7268   1.0000
   3.750   0.5335   0.01262   0.00544  -0.0259   0.5031   1.0000
   4.000   0.5385   0.01676   0.00687  -0.0224   0.0871   1.0000
   4.250   0.5609   0.01815   0.00821  -0.0211   0.0704   1.0000
   4.500   0.5832   0.01988   0.00989  -0.0197   0.0643   1.0000
   4.750   0.6079   0.02140   0.01143  -0.0188   0.0573   1.0000
   5.000   0.6331   0.02404   0.01401  -0.0180   0.0533   1.0000
   5.250   0.6604   0.02629   0.01654  -0.0171   0.0528   1.0000
   5.500   0.6866   0.02935   0.01995  -0.0162   0.0533   1.0000
   5.750   0.7140   0.03131   0.02241  -0.0149   0.0559   1.0000
   6.000   0.7381   0.03547   0.02743  -0.0130   0.0628   1.0000
   6.250   0.7588   0.03969   0.03198  -0.0121   0.0652   1.0000
   7.500   0.8373   0.07168   0.06661  -0.0106   0.1460   1.0000
   7.750   0.8246   0.07485   0.07023  -0.0134   0.1348   1.0000
   8.000   0.8167   0.08025   0.07574  -0.0160   0.1288   1.0000
   9.000   0.7375   0.09042   0.08611  -0.0062   0.1185   1.0000
   9.250   0.7090   0.09559   0.09130  -0.0100   0.1184   1.0000
   9.500   0.6796   0.10185   0.09751  -0.0166   0.1181   1.0000
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