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

GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 55 AIRFOIL (goe55-il)
Reynolds number: 1,000,000
Max Cl/Cd: 77.62 at α=7°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe55-il-1000000-n5.txt
Download as CSV file: xf-goe55-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 55 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -1.0591   0.02837   0.02545  -0.0299   1.0000   0.0174
 -11.000  -1.0338   0.02770   0.02474  -0.0299   1.0000   0.0177
 -10.750  -1.0069   0.02789   0.02497  -0.0295   1.0000   0.0179
 -10.500  -0.9806   0.02767   0.02474  -0.0293   1.0000   0.0181
 -10.250  -0.9553   0.02685   0.02385  -0.0294   1.0000   0.0184
 -10.000  -0.9297   0.02604   0.02296  -0.0294   1.0000   0.0186
  -9.750  -0.9040   0.02515   0.02199  -0.0295   1.0000   0.0190
  -9.500  -0.8787   0.02385   0.02053  -0.0298   1.0000   0.0194
  -9.250  -0.8537   0.02212   0.01858  -0.0302   1.0000   0.0198
  -9.000  -0.8280   0.02060   0.01683  -0.0306   1.0000   0.0203
  -8.750  -0.8014   0.01959   0.01565  -0.0307   1.0000   0.0206
  -8.500  -0.7745   0.01873   0.01465  -0.0308   1.0000   0.0209
  -8.250  -0.7473   0.01797   0.01375  -0.0308   1.0000   0.0210
  -8.000  -0.7205   0.01662   0.01221  -0.0311   1.0000   0.0215
  -7.750  -0.6930   0.01606   0.01158  -0.0312   1.0000   0.0218
  -7.500  -0.6652   0.01555   0.01102  -0.0311   1.0000   0.0221
  -7.250  -0.6374   0.01505   0.01045  -0.0311   1.0000   0.0223
  -7.000  -0.6094   0.01454   0.00988  -0.0311   1.0000   0.0226
  -6.750  -0.5814   0.01415   0.00943  -0.0311   1.0000   0.0229
  -6.500  -0.5532   0.01370   0.00893  -0.0311   1.0000   0.0232
  -6.250  -0.5250   0.01315   0.00830  -0.0311   1.0000   0.0235
  -6.000  -0.4967   0.01261   0.00767  -0.0311   1.0000   0.0238
  -5.750  -0.4682   0.01211   0.00710  -0.0311   1.0000   0.0241
  -5.500  -0.4398   0.01164   0.00657  -0.0310   1.0000   0.0244
  -5.250  -0.4112   0.01122   0.00609  -0.0310   1.0000   0.0247
  -5.000  -0.3826   0.01083   0.00565  -0.0310   1.0000   0.0250
  -4.750  -0.3610   0.01123   0.00541  -0.0291   0.7577   0.0252
  -4.500  -0.3331   0.01138   0.00513  -0.0290   0.6377   0.0255
  -4.250  -0.3045   0.01162   0.00483  -0.0292   0.4719   0.0257
  -4.000  -0.2758   0.01128   0.00428  -0.0293   0.4205   0.0261
  -3.750  -0.2470   0.01100   0.00387  -0.0294   0.3910   0.0265
  -3.500  -0.2183   0.01079   0.00357  -0.0294   0.3721   0.0269
  -3.250  -0.1895   0.01061   0.00333  -0.0293   0.3584   0.0273
  -3.000  -0.1608   0.01046   0.00312  -0.0293   0.3477   0.0277
  -2.500  -0.1032   0.01018   0.00277  -0.0292   0.3331   0.0286
  -2.250  -0.0744   0.01006   0.00261  -0.0292   0.3276   0.0291
  -1.750  -0.0168   0.00984   0.00233  -0.0290   0.3188   0.0302
  -1.500   0.0120   0.00975   0.00223  -0.0290   0.3151   0.0308
  -1.250   0.0408   0.00969   0.00214  -0.0289   0.3099   0.0312
  -1.000   0.0697   0.00959   0.00201  -0.0289   0.3048   0.0323
  -0.750   0.0986   0.00950   0.00192  -0.0288   0.3015   0.0336
  -0.500   0.1274   0.00944   0.00186  -0.0287   0.2968   0.0349
  -0.250   0.1562   0.00941   0.00181  -0.0287   0.2890   0.0364
   0.000   0.1849   0.00937   0.00176  -0.0286   0.2812   0.0381
   0.250   0.2137   0.00934   0.00172  -0.0286   0.2752   0.0424
   0.500   0.2425   0.00930   0.00170  -0.0285   0.2699   0.0483
   0.750   0.2713   0.00927   0.00168  -0.0285   0.2604   0.0621
   1.000   0.3001   0.00922   0.00172  -0.0285   0.2531   0.0875
   1.250   0.3288   0.00925   0.00172  -0.0284   0.2437   0.0961
   1.500   0.3575   0.00927   0.00174  -0.0284   0.2318   0.1034
   1.750   0.3861   0.00936   0.00176  -0.0283   0.2100   0.1096
   2.000   0.4145   0.00966   0.00187  -0.0284   0.1573   0.1186
   2.250   0.4434   0.00970   0.00204  -0.0287   0.1271   0.2265
   2.750   0.4948   0.00813   0.00227  -0.0280   0.0967   1.0000
   3.000   0.5233   0.00842   0.00243  -0.0280   0.0736   1.0000
   3.250   0.5517   0.00889   0.00270  -0.0281   0.0360   1.0000
   3.500   0.5802   0.00907   0.00284  -0.0280   0.0338   1.0000
   3.750   0.6087   0.00926   0.00301  -0.0279   0.0322   1.0000
   4.000   0.6371   0.00946   0.00319  -0.0278   0.0309   1.0000
   4.250   0.6655   0.00966   0.00339  -0.0278   0.0299   1.0000
   4.500   0.6938   0.00989   0.00362  -0.0277   0.0292   1.0000
   4.750   0.7221   0.01009   0.00382  -0.0276   0.0289   1.0000
   5.000   0.7503   0.01030   0.00403  -0.0275   0.0286   1.0000
   5.250   0.7784   0.01052   0.00427  -0.0275   0.0282   1.0000
   5.500   0.8065   0.01076   0.00452  -0.0274   0.0277   1.0000
   5.750   0.8345   0.01102   0.00478  -0.0273   0.0273   1.0000
   6.000   0.8625   0.01129   0.00508  -0.0272   0.0269   1.0000
   6.250   0.8903   0.01158   0.00538  -0.0271   0.0265   1.0000
   6.500   0.9181   0.01188   0.00571  -0.0270   0.0262   1.0000
   6.750   0.9457   0.01220   0.00606  -0.0270   0.0258   1.0000
   7.000   0.9733   0.01254   0.00643  -0.0269   0.0255   1.0000
   7.250   1.0007   0.01291   0.00682  -0.0268   0.0251   1.0000
   7.500   1.0279   0.01331   0.00726  -0.0267   0.0248   1.0000
   7.750   1.0549   0.01374   0.00774  -0.0265   0.0244   1.0000
   8.000   1.0816   0.01429   0.00833  -0.0264   0.0239   1.0000
   8.250   1.1077   0.01499   0.00910  -0.0263   0.0234   1.0000
   8.500   1.1340   0.01554   0.00972  -0.0261   0.0232   1.0000
   8.750   1.1605   0.01595   0.01017  -0.0260   0.0230   1.0000
   9.000   1.1870   0.01631   0.01058  -0.0258   0.0228   1.0000
   9.250   1.2133   0.01672   0.01103  -0.0256   0.0226   1.0000
   9.500   1.2393   0.01717   0.01154  -0.0254   0.0223   1.0000
   9.750   1.2653   0.01760   0.01203  -0.0252   0.0219   1.0000
  10.000   1.2910   0.01806   0.01254  -0.0250   0.0215   1.0000
  10.250   1.3165   0.01851   0.01304  -0.0248   0.0210   1.0000
  10.500   1.3419   0.01897   0.01356  -0.0245   0.0206   1.0000
  10.750   1.3671   0.01944   0.01407  -0.0243   0.0202   1.0000
  11.000   1.3923   0.01985   0.01451  -0.0241   0.0198   1.0000
  11.250   1.4171   0.02033   0.01503  -0.0238   0.0194   1.0000
  11.500   1.4410   0.02098   0.01572  -0.0235   0.0190   1.0000
  11.750   1.4635   0.02185   0.01667  -0.0231   0.0185   1.0000
  12.000   1.4879   0.02229   0.01719  -0.0228   0.0183   1.0000
  12.250   1.5117   0.02281   0.01777  -0.0225   0.0179   1.0000
  12.500   1.5349   0.02339   0.01843  -0.0221   0.0175   1.0000
  12.750   1.5578   0.02400   0.01911  -0.0217   0.0170   1.0000
  13.000   1.5805   0.02458   0.01976  -0.0213   0.0166   1.0000
  13.250   1.6033   0.02511   0.02033  -0.0209   0.0160   1.0000
  13.500   1.6260   0.02562   0.02088  -0.0205   0.0155   1.0000
  13.750   1.6479   0.02623   0.02151  -0.0200   0.0149   1.0000
  14.000   1.6675   0.02711   0.02248  -0.0194   0.0145   1.0000
  14.250   1.6862   0.02806   0.02353  -0.0188   0.0140   1.0000
  14.500   1.7042   0.02902   0.02458  -0.0181   0.0135   1.0000
  14.750   1.7215   0.03000   0.02563  -0.0174   0.0129   1.0000
  15.000   1.7376   0.03105   0.02673  -0.0166   0.0123   1.0000
  15.250   1.7512   0.03232   0.02807  -0.0158   0.0117   1.0000
  15.500   1.7600   0.03396   0.02983  -0.0147   0.0114   1.0000
  15.750   1.7631   0.03589   0.03190  -0.0133   0.0110   1.0000
  16.000   1.7537   0.03863   0.03478  -0.0116   0.0108   1.0000
  16.250   1.7322   0.04577   0.04216  -0.0171   0.0107   1.0000
  16.500   1.6887   0.06110   0.05781  -0.0304   0.0110   1.0000
  16.750   1.6272   0.07509   0.07198  -0.0384   0.0114   1.0000
  17.000   1.5735   0.08731   0.08435  -0.0449   0.0116   1.0000
<< Back to GOE 55 AIRFOIL (goe55-il)

Polar data table (+)

Polar graphs


<< Back to GOE 55 AIRFOIL (goe55-il)