Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il)
Reynolds number: 1,000,000
Max Cl/Cd: 80.77 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe235-il-1000000.txt
Download as CSV file: xf-goe235-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 235 (SCHTTE-LANZ) AIRFOIL                   
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4684   0.09968   0.09802   0.0039   1.0000   0.0197
  -8.250  -0.4665   0.09573   0.09409  -0.0002   1.0000   0.0198
  -8.000  -0.4623   0.09173   0.09010  -0.0045   1.0000   0.0198
  -7.750  -0.4540   0.08709   0.08547  -0.0078   1.0000   0.0199
  -7.500  -0.4430   0.08405   0.08243  -0.0083   1.0000   0.0201
  -7.250  -0.4301   0.08099   0.07938  -0.0105   1.0000   0.0202
  -7.000  -0.4157   0.07784   0.07624  -0.0135   1.0000   0.0203
  -6.750  -0.3996   0.07460   0.07300  -0.0170   1.0000   0.0205
  -6.500  -0.3819   0.07127   0.06966  -0.0209   1.0000   0.0208
  -6.250  -0.3566   0.06744   0.06580  -0.0266   0.9854   0.0212
  -6.000  -0.3312   0.06359   0.06190  -0.0319   0.9616   0.0218
  -5.750  -0.2948   0.05706   0.05516  -0.0423   0.9348   0.0233
  -5.500  -0.2721   0.05300   0.05093  -0.0453   0.9033   0.0233
  -5.250  -0.2325   0.03336   0.03113  -0.0452   0.8391   0.0236
  -5.000  -0.2135   0.03067   0.02836  -0.0462   0.8282   0.0237
  -4.750  -0.1925   0.02806   0.02567  -0.0475   0.8182   0.0239
  -4.500  -0.1700   0.02551   0.02303  -0.0490   0.8100   0.0242
  -4.250  -0.1597   0.03863   0.03590  -0.0530   0.8211   0.0244
  -4.000  -0.1329   0.03630   0.03344  -0.0543   0.8114   0.0248
  -3.750  -0.0947   0.03228   0.02907  -0.0564   0.8028   0.0272
  -3.500  -0.0653   0.02937   0.02590  -0.0573   0.7950   0.0273
  -3.250  -0.0402   0.02547   0.02182  -0.0586   0.7860   0.0277
  -3.000  -0.0139   0.02409   0.02034  -0.0590   0.7754   0.0280
  -2.750   0.0130   0.02282   0.01896  -0.0594   0.7625   0.0283
  -2.500   0.0405   0.02154   0.01754  -0.0597   0.7472   0.0287
  -2.250   0.0687   0.02021   0.01605  -0.0599   0.7288   0.0294
  -2.000   0.1006   0.01937   0.01480  -0.0593   0.7053   0.0318
  -1.750   0.1294   0.01809   0.01320  -0.0593   0.6821   0.0319
  -1.500   0.1564   0.01550   0.01043  -0.0600   0.6623   0.0328
  -1.250   0.1843   0.01485   0.00966  -0.0601   0.6446   0.0334
  -1.000   0.2127   0.01424   0.00893  -0.0602   0.6317   0.0345
  -0.750   0.2422   0.01462   0.00907  -0.0597   0.6197   0.0371
  -0.500   0.2711   0.01411   0.00836  -0.0596   0.6060   0.0372
   0.000   0.3293   0.01042   0.00421  -0.0596   0.5735   0.0321
   0.250   0.3576   0.00976   0.00342  -0.0595   0.5446   0.0323
   0.500   0.3844   0.00982   0.00302  -0.0595   0.4275   0.0328
   0.750   0.4118   0.00991   0.00285  -0.0595   0.3623   0.0333
   1.000   0.4386   0.01026   0.00279  -0.0596   0.2570   0.0338
   1.250   0.4663   0.01034   0.00270  -0.0596   0.2192   0.0344
   1.500   0.4943   0.01031   0.00261  -0.0595   0.2047   0.0353
   1.750   0.5225   0.01029   0.00256  -0.0595   0.1960   0.0365
   2.000   0.5506   0.01033   0.00255  -0.0594   0.1868   0.0373
   2.250   0.5789   0.01014   0.00234  -0.0595   0.1791   0.0389
   2.500   0.6070   0.01018   0.00235  -0.0594   0.1700   0.0404
   2.750   0.6350   0.01023   0.00237  -0.0594   0.1596   0.0422
   3.000   0.6628   0.01033   0.00238  -0.0593   0.1427   0.0444
   3.250   0.6881   0.01114   0.00276  -0.0592   0.0409   0.0481
   3.500   0.7157   0.01130   0.00291  -0.0590   0.0370   0.0516
   3.750   0.7434   0.01144   0.00306  -0.0589   0.0349   0.0593
   4.000   0.7667   0.00974   0.00339  -0.0584   0.0343   1.0000
   4.250   0.7941   0.00997   0.00360  -0.0583   0.0333   1.0000
   4.500   0.8214   0.01023   0.00384  -0.0581   0.0324   1.0000
   4.750   0.8485   0.01052   0.00412  -0.0579   0.0316   1.0000
   5.000   0.8755   0.01084   0.00443  -0.0577   0.0308   1.0000
   5.250   0.9021   0.01122   0.00482  -0.0575   0.0301   1.0000
   5.500   0.9282   0.01168   0.00531  -0.0572   0.0294   1.0000
   5.750   0.9535   0.01230   0.00596  -0.0568   0.0288   1.0000
   6.000   0.9799   0.01264   0.00632  -0.0565   0.0286   1.0000
   6.250   1.0062   0.01297   0.00666  -0.0562   0.0283   1.0000
   6.500   1.0322   0.01335   0.00705  -0.0559   0.0278   1.0000
   6.750   1.0577   0.01380   0.00752  -0.0556   0.0274   1.0000
   7.000   1.0828   0.01428   0.00803  -0.0552   0.0270   1.0000
   7.250   1.1076   0.01480   0.00856  -0.0547   0.0266   1.0000
   7.500   1.1321   0.01535   0.00913  -0.0542   0.0261   1.0000
   7.750   1.1562   0.01592   0.00974  -0.0537   0.0257   1.0000
   8.000   1.1799   0.01653   0.01036  -0.0532   0.0253   1.0000
   8.250   1.2030   0.01719   0.01103  -0.0526   0.0248   1.0000
   8.500   1.2242   0.01816   0.01200  -0.0518   0.0243   1.0000
   8.750   1.2411   0.02003   0.01390  -0.0505   0.0238   1.0000
   9.000   1.2643   0.02060   0.01452  -0.0499   0.0236   1.0000
   9.250   1.2876   0.02109   0.01506  -0.0493   0.0233   1.0000
   9.500   1.3102   0.02169   0.01572  -0.0486   0.0230   1.0000
   9.750   1.3324   0.02230   0.01639  -0.0479   0.0226   1.0000
  10.000   1.3540   0.02296   0.01711  -0.0472   0.0221   1.0000
  10.250   1.3747   0.02375   0.01794  -0.0463   0.0217   1.0000
  10.500   1.3949   0.02457   0.01881  -0.0454   0.0213   1.0000
  10.750   1.4146   0.02538   0.01967  -0.0445   0.0210   1.0000
  11.000   1.4335   0.02621   0.02055  -0.0436   0.0207   1.0000
  11.250   1.4515   0.02708   0.02146  -0.0425   0.0204   1.0000
  11.500   1.4685   0.02806   0.02247  -0.0414   0.0201   1.0000
  11.750   1.4802   0.03120   0.02567  -0.0401   0.0194   1.0000
  12.000   1.4950   0.03159   0.02618  -0.0386   0.0192   1.0000
  12.250   1.5084   0.03236   0.02706  -0.0370   0.0190   1.0000
  12.500   1.5190   0.03336   0.02819  -0.0351   0.0188   1.0000
  12.750   1.5250   0.03452   0.02946  -0.0326   0.0185   1.0000
  13.000   1.5298   0.03588   0.03094  -0.0304   0.0182   1.0000
  13.250   1.5338   0.03740   0.03257  -0.0284   0.0180   1.0000
  13.500   1.5370   0.03908   0.03435  -0.0267   0.0177   1.0000
  13.750   1.5399   0.04086   0.03624  -0.0254   0.0175   1.0000
  14.000   1.5426   0.04275   0.03822  -0.0245   0.0173   1.0000
  14.250   1.5454   0.04469   0.04024  -0.0239   0.0171   1.0000
  14.500   1.5468   0.04691   0.04255  -0.0235   0.0169   1.0000
  14.750   1.5480   0.04926   0.04499  -0.0235   0.0168   1.0000
  15.000   1.5483   0.05187   0.04768  -0.0237   0.0166   1.0000
  15.250   1.5472   0.05480   0.05070  -0.0242   0.0165   1.0000
  15.500   1.5448   0.05808   0.05408  -0.0250   0.0164   1.0000
  15.750   1.5402   0.06177   0.05786  -0.0260   0.0163   1.0000
  16.000   1.5335   0.06585   0.06203  -0.0272   0.0162   1.0000
  16.250   1.5227   0.07063   0.06693  -0.0287   0.0160   1.0000
  16.500   1.5020   0.07700   0.07347  -0.0307   0.0159   1.0000
  16.750   1.4756   0.08490   0.08157  -0.0341   0.0158   1.0000
  17.000   1.4614   0.09132   0.08815  -0.0376   0.0157   1.0000
  17.250   1.4454   0.09838   0.09536  -0.0417   0.0157   1.0000
  17.500   1.4270   0.10615   0.10329  -0.0461   0.0156   1.0000
  17.750   1.4066   0.11459   0.11188  -0.0511   0.0156   1.0000
  18.000   1.3831   0.12396   0.12142  -0.0568   0.0155   1.0000
  18.250   1.3559   0.13463   0.13226  -0.0635   0.0155   1.0000
  18.500   1.3235   0.14703   0.14484  -0.0714   0.0155   1.0000
  18.750   1.2703   0.16587   0.16392  -0.0839   0.0155   1.0000
<< Back to GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il)

Polar data table (+)

Polar graphs


<< Back to GOE 235 (SCHTTE-LANZ) AIRFOIL (goe235-il)