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GOE 625 AIRFOIL (goe625-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 625 AIRFOIL (goe625-il)
Reynolds number: 100,000
Max Cl/Cd: 33.81 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe625-il-100000.txt
Download as CSV file: xf-goe625-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 625 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.1389   0.13423   0.12933  -0.0615   0.9692   0.1437
 -10.750  -0.1346   0.13113   0.12621  -0.0677   0.9657   0.1498
 -10.500  -0.1588   0.12849   0.12363  -0.0723   0.9571   0.1518
 -10.250  -0.1059   0.12287   0.11795  -0.0732   0.9544   0.1543
 -10.000  -0.0755   0.11896   0.11400  -0.0770   0.9514   0.1592
  -9.750  -0.1036   0.11769   0.11277  -0.0816   0.9415   0.1657
  -9.500  -0.0870   0.11228   0.10735  -0.0855   0.9373   0.1678
  -9.250  -0.0380   0.10773   0.10273  -0.0879   0.9354   0.1714
  -9.000   0.0327   0.09831   0.09353  -0.1011   0.9026   0.1841
  -8.750   0.0934   0.09292   0.08807  -0.1032   0.9007   0.1881
  -8.500   0.1144   0.08895   0.08406  -0.1066   0.8977   0.1963
  -8.250   0.0224   0.08896   0.08416  -0.1077   0.8808   0.2010
  -8.000   0.1131   0.08230   0.07741  -0.1081   0.8808   0.2051
  -7.750   0.1291   0.07999   0.07510  -0.1072   0.8715   0.2103
  -7.500   0.0487   0.07799   0.07316  -0.1102   0.8594   0.2198
  -7.250   0.0044   0.08311   0.07810  -0.1059   0.8722   0.2184
  -7.000   0.0047   0.07887   0.07386  -0.1072   0.8635   0.2227
  -6.750   0.0621   0.07591   0.07083  -0.1080   0.8614   0.2298
  -6.500  -0.0121   0.07316   0.06803  -0.1101   0.8426   0.2411
  -6.250   0.0460   0.07006   0.06495  -0.1092   0.8408   0.2451
  -6.000  -0.0658   0.05102   0.04456  -0.1177   0.8216   0.1678
  -5.750  -0.0652   0.04864   0.04182  -0.1149   0.8075   0.1656
  -5.500  -0.0340   0.04498   0.03766  -0.1170   0.8032   0.1663
  -5.250  -0.0225   0.04293   0.03538  -0.1150   0.7906   0.1668
  -5.000   0.0151   0.04025   0.03268  -0.1169   0.7854   0.1691
  -4.750   0.0410   0.03882   0.03118  -0.1166   0.7757   0.1716
  -4.500   0.0720   0.03686   0.02897  -0.1171   0.7673   0.1738
  -4.250   0.1132   0.03452   0.02626  -0.1191   0.7630   0.1763
  -4.000   0.1306   0.03350   0.02501  -0.1173   0.7485   0.1789
  -3.750   0.1723   0.03172   0.02274  -0.1191   0.7435   0.1825
  -3.500   0.1914   0.03063   0.02154  -0.1176   0.7290   0.1849
  -3.250   0.2209   0.02957   0.02041  -0.1176   0.7186   0.1885
  -3.000   0.2532   0.02862   0.01931  -0.1180   0.7087   0.1939
  -2.750   0.2794   0.02802   0.01847  -0.1174   0.6968   0.1993
  -2.500   0.3143   0.02680   0.01712  -0.1183   0.6884   0.2053
  -2.250   0.3376   0.02640   0.01668  -0.1173   0.6761   0.2119
  -2.000   0.3734   0.02550   0.01556  -0.1183   0.6685   0.2226
  -1.750   0.3942   0.02528   0.01542  -0.1171   0.6570   0.2329
  -1.500   0.4280   0.02455   0.01460  -0.1179   0.6501   0.2477
  -1.250   0.4495   0.02448   0.01453  -0.1168   0.6405   0.2623
  -1.000   0.4791   0.02414   0.01416  -0.1170   0.6334   0.2808
  -0.750   0.5070   0.02390   0.01399  -0.1170   0.6269   0.2998
  -0.500   0.5284   0.02391   0.01413  -0.1160   0.6188   0.3182
  -0.250   0.5600   0.02360   0.01384  -0.1165   0.6132   0.3447
   0.000   0.5826   0.02360   0.01406  -0.1157   0.6068   0.3775
   0.250   0.6035   0.02349   0.01435  -0.1147   0.6004   0.4438
   0.500   0.6435   0.02231   0.01472  -0.1144   0.5957   0.9258
   0.750   0.7305   0.02268   0.01472  -0.1258   0.5899   1.0000
   1.000   0.7403   0.02330   0.01528  -0.1230   0.5841   1.0000
   1.250   0.7601   0.02365   0.01548  -0.1218   0.5791   1.0000
   1.500   0.7883   0.02385   0.01545  -0.1219   0.5752   1.0000
   1.750   0.8061   0.02445   0.01595  -0.1205   0.5711   1.0000
   2.000   0.8142   0.02531   0.01683  -0.1176   0.5662   1.0000
   2.250   0.8326   0.02588   0.01732  -0.1162   0.5617   1.0000
   2.500   0.8598   0.02619   0.01748  -0.1162   0.5578   1.0000
   2.750   0.8944   0.02645   0.01752  -0.1174   0.5546   1.0000
   3.000   0.8931   0.02776   0.01897  -0.1132   0.5497   1.0000
   3.250   0.9044   0.02870   0.01992  -0.1109   0.5449   1.0000
   3.500   0.9295   0.02912   0.02025  -0.1106   0.5407   1.0000
   3.750   0.9670   0.02916   0.02009  -0.1121   0.5370   1.0000
   4.000   0.9710   0.03041   0.02141  -0.1089   0.5318   1.0000
   4.250   0.9768   0.03151   0.02257  -0.1059   0.5259   1.0000
   4.500   1.0071   0.03168   0.02262  -0.1063   0.5213   1.0000
   4.750   1.0507   0.03152   0.02226  -0.1087   0.5177   1.0000
   5.000   1.0388   0.03349   0.02440  -0.1033   0.5124   1.0000
   5.250   1.0318   0.03528   0.02630  -0.0989   0.5068   1.0000
   5.500   1.0621   0.03551   0.02645  -0.0993   0.5027   1.0000
   5.750   1.1109   0.03510   0.02587  -0.1023   0.4994   1.0000
   6.000   1.0815   0.03800   0.02897  -0.0951   0.4937   1.0000
   6.250   0.9812   0.04500   0.03628  -0.0810   0.4841   1.0000
   6.500   1.0738   0.04139   0.03245  -0.0875   0.4830   1.0000
   6.750   1.1533   0.03959   0.03045  -0.0941   0.4805   1.0000
   7.000   0.8964   0.06048   0.05198  -0.0710   0.4588   1.0000
   7.250   0.9603   0.05669   0.04808  -0.0721   0.4589   1.0000
   7.500   1.0351   0.05251   0.04377  -0.0745   0.4589   1.0000
   7.750   1.1402   0.04715   0.03821  -0.0805   0.4591   1.0000
  10.000   0.7071   0.12423   0.11622  -0.0669   0.3607   1.0000
  10.250   0.6972   0.12975   0.12179  -0.0680   0.3610   1.0000
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