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GOE 590 AIRFOIL (goe590-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 590 AIRFOIL (goe590-il)
Reynolds number: 50,000
Max Cl/Cd: 37.3 at α=5.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe590-il-50000.txt
Download as CSV file: xf-goe590-il-50000.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 =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4388   0.11294   0.10622  -0.0055   1.0000   0.1021
  -8.500  -0.4427   0.11190   0.10529  -0.0083   1.0000   0.1040
  -8.250  -0.4504   0.11144   0.10496  -0.0116   1.0000   0.1046
  -8.000  -0.4263   0.10311   0.09659  -0.0086   1.0000   0.1102
  -7.750  -0.4241   0.10040   0.09395  -0.0098   1.0000   0.1145
  -7.500  -0.4273   0.09906   0.09272  -0.0131   1.0000   0.1174
  -7.250  -0.4287   0.09894   0.09264  -0.0191   1.0000   0.1186
  -7.000  -0.4118   0.09105   0.08481  -0.0138   1.0000   0.1257
  -6.750  -0.4083   0.08933   0.08313  -0.0177   1.0000   0.1310
  -6.500  -0.4000   0.08507   0.07894  -0.0180   1.0000   0.1350
  -6.250  -0.3915   0.08207   0.07596  -0.0194   1.0000   0.1436
  -6.000  -0.3829   0.07856   0.07248  -0.0207   1.0000   0.1485
  -5.750  -0.3724   0.07710   0.07095  -0.0244   1.0000   0.1591
  -5.500  -0.3633   0.07240   0.06630  -0.0224   1.0000   0.1707
  -5.250  -0.3537   0.06915   0.06307  -0.0227   1.0000   0.1840
  -5.000  -0.3427   0.06706   0.06093  -0.0247   1.0000   0.2004
  -4.750  -0.3348   0.06285   0.05683  -0.0228   1.0000   0.2159
  -4.500  -0.3262   0.05951   0.05355  -0.0213   1.0000   0.2341
  -4.250  -0.3184   0.05668   0.05075  -0.0205   1.0000   0.2581
  -4.000  -0.3133   0.05371   0.04787  -0.0183   1.0000   0.2870
  -3.750  -0.3123   0.05103   0.04529  -0.0154   1.0000   0.3261
  -3.500  -0.3159   0.04822   0.04263  -0.0107   1.0000   0.3687
  -3.250  -0.3219   0.04556   0.04013  -0.0056   1.0000   0.4127
  -3.000  -0.3270   0.04303   0.03772  -0.0010   1.0000   0.4523
  -2.750  -0.3276   0.04093   0.03564   0.0020   1.0000   0.4880
  -2.500  -0.3271   0.03838   0.03320   0.0057   1.0000   0.5182
  -2.250  -0.3245   0.03610   0.03099   0.0090   1.0000   0.5521
  -2.000  -0.3151   0.03395   0.02878   0.0103   1.0000   0.5770
  -1.750  -0.2889   0.03211   0.02671   0.0068   0.9997   0.5839
  -1.500  -0.1037   0.03419   0.02600  -0.0278   0.9799   0.2249
  -1.250  -0.0427   0.03212   0.02306  -0.0331   0.9658   0.1849
  -1.000   0.0109   0.03000   0.02033  -0.0373   0.9525   0.1671
  -0.750   0.0643   0.02840   0.01800  -0.0414   0.9395   0.1606
  -0.500   0.1181   0.02699   0.01630  -0.0462   0.9273   0.1777
  -0.250   0.1769   0.02583   0.01472  -0.0512   0.9155   0.1878
   0.000   0.2246   0.02481   0.01364  -0.0547   0.9036   0.2114
   0.250   0.2611   0.02360   0.01287  -0.0565   0.8916   0.2956
   0.500   0.3673   0.02145   0.01141  -0.0704   0.8851   1.0000
   0.750   0.4036   0.02184   0.01156  -0.0723   0.8723   1.0000
   1.000   0.4342   0.02234   0.01190  -0.0732   0.8597   1.0000
   1.250   0.4598   0.02295   0.01240  -0.0733   0.8474   1.0000
   1.500   0.4830   0.02365   0.01301  -0.0730   0.8358   1.0000
   1.750   0.5063   0.02437   0.01367  -0.0728   0.8252   1.0000
   2.000   0.5342   0.02499   0.01426  -0.0732   0.8162   1.0000
   2.250   0.5544   0.02585   0.01511  -0.0725   0.8064   1.0000
   2.750   0.5992   0.02756   0.01693  -0.0719   0.7898   1.0000
   3.000   0.6108   0.02884   0.01825  -0.0704   0.7808   1.0000
   3.250   0.6297   0.02994   0.01942  -0.0699   0.7733   1.0000
   3.500   0.6465   0.03115   0.02072  -0.0691   0.7659   1.0000
   3.750   0.6601   0.03255   0.02221  -0.0681   0.7588   1.0000
   4.000   0.6765   0.03388   0.02374  -0.0674   0.7520   1.0000
   4.250   0.6877   0.03547   0.02546  -0.0663   0.7455   1.0000
   4.500   0.6985   0.03710   0.02721  -0.0653   0.7391   1.0000
   4.750   0.7147   0.03865   0.02894  -0.0647   0.7335   1.0000
   5.000   0.7138   0.04075   0.03113  -0.0627   0.7273   1.0000
   5.250   0.7395   0.04213   0.03281  -0.0632   0.7220   1.0000
   5.500   0.8501   0.02313   0.01422  -0.0376   0.5218   1.0000
   5.750   0.8437   0.02262   0.01226  -0.0284   0.2216   1.0000
   6.000   0.8431   0.02634   0.01484  -0.0257   0.1369   1.0000
   6.250   0.8516   0.02856   0.01701  -0.0231   0.1183   1.0000
   6.500   0.8616   0.03052   0.01896  -0.0204   0.1099   1.0000
   6.750   0.8778   0.03233   0.02081  -0.0178   0.1030   1.0000
   7.000   0.9048   0.03436   0.02283  -0.0162   0.0939   1.0000
   7.250   0.9453   0.03739   0.02577  -0.0160   0.0880   1.0000
   7.500   0.9813   0.04061   0.02921  -0.0155   0.0881   1.0000
   7.750   1.0066   0.04334   0.03260  -0.0138   0.0907   1.0000
   8.000   1.0283   0.04683   0.03668  -0.0121   0.0942   1.0000
   8.250   1.0460   0.05064   0.04095  -0.0105   0.0959   1.0000
   8.500   1.0589   0.05465   0.04540  -0.0088   0.0974   1.0000
   8.750   1.0744   0.05916   0.05023  -0.0076   0.1008   1.0000
   9.000   1.0679   0.06270   0.05467  -0.0048   0.1083   1.0000
   9.250   1.0674   0.06720   0.05966  -0.0033   0.1164   1.0000
   9.500   1.0693   0.07284   0.06562  -0.0026   0.1270   1.0000
   9.750   1.0390   0.07674   0.06992  -0.0016   0.1292   1.0000
  10.000   1.0101   0.08107   0.07440  -0.0012   0.1306   1.0000
  10.250   0.9798   0.08612   0.07951  -0.0028   0.1319   1.0000
  10.500   0.9652   0.09288   0.08634  -0.0058   0.1406   1.0000
  10.750   0.9403   0.10211   0.09556  -0.0118   0.1528   1.0000
  11.000   0.8873   0.11541   0.10867  -0.0251   0.1683   1.0000
  11.250   0.8591   0.12684   0.11994  -0.0343   0.2018   1.0000
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