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GOE 177 AIRFOIL (goe177-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 177 AIRFOIL (goe177-il)
Reynolds number: 500,000
Max Cl/Cd: 91.21 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe177-il-500000.txt
Download as CSV file: xf-goe177-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 177 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.3262   0.08645   0.08444  -0.0192   1.0000   0.0141
  -7.000  -0.3237   0.08387   0.08190  -0.0198   1.0000   0.0144
  -6.750  -0.3208   0.08120   0.07927  -0.0209   1.0000   0.0148
  -6.500  -0.3187   0.07858   0.07668  -0.0216   1.0000   0.0152
  -6.250  -0.3169   0.07594   0.07407  -0.0223   1.0000   0.0158
  -6.000  -0.2819   0.07034   0.06844  -0.0327   0.9959   0.0175
  -5.750  -0.2317   0.06441   0.06243  -0.0463   0.9904   0.0187
  -5.500  -0.1904   0.05824   0.05619  -0.0564   0.9856   0.0188
  -5.250  -0.1611   0.05086   0.04874  -0.0649   0.9806   0.0204
  -5.000  -0.1299   0.04849   0.04633  -0.0686   0.9727   0.0225
  -4.750  -0.0907   0.04369   0.04139  -0.0754   0.9627   0.0253
  -4.250  -0.0098   0.02252   0.01920  -0.0897   0.9380   0.0177
  -4.000   0.0198   0.01531   0.01091  -0.0912   0.9200   0.0164
  -3.750   0.0470   0.01336   0.00848  -0.0907   0.8929   0.0171
  -3.500   0.0731   0.01234   0.00708  -0.0898   0.8572   0.0181
  -3.250   0.0990   0.01174   0.00613  -0.0889   0.8188   0.0190
  -3.000   0.1251   0.01045   0.00447  -0.0884   0.7871   0.0220
  -2.500   0.1795   0.00972   0.00331  -0.0875   0.7359   0.0273
  -2.250   0.2070   0.00923   0.00267  -0.0873   0.7149   0.0358
  -2.000   0.2348   0.00907   0.00250  -0.0870   0.6943   0.0584
  -1.750   0.2624   0.00929   0.00260  -0.0868   0.6749   0.0746
  -1.500   0.2899   0.00929   0.00250  -0.0867   0.6556   0.0850
  -1.250   0.3175   0.00927   0.00236  -0.0865   0.6364   0.0924
  -1.000   0.3449   0.00926   0.00226  -0.0864   0.6174   0.0995
  -0.750   0.3726   0.00927   0.00216  -0.0863   0.5987   0.1032
  -0.500   0.4002   0.00915   0.00198  -0.0862   0.5810   0.1088
  -0.250   0.4279   0.00915   0.00190  -0.0861   0.5654   0.1144
   0.000   0.4555   0.00917   0.00180  -0.0860   0.5472   0.1184
   0.250   0.4830   0.00915   0.00173  -0.0859   0.5250   0.1266
   0.500   0.5104   0.00917   0.00168  -0.0858   0.5001   0.1440
   0.750   0.5377   0.00895   0.00174  -0.0859   0.4783   0.2756
   1.000   0.5650   0.00867   0.00181  -0.0861   0.4619   0.4388
   1.500   0.6137   0.00766   0.00187  -0.0844   0.4311   1.0000
   1.750   0.6412   0.00782   0.00191  -0.0843   0.4151   1.0000
   2.000   0.6686   0.00800   0.00196  -0.0842   0.3991   1.0000
   2.250   0.6959   0.00817   0.00204  -0.0841   0.3837   1.0000
   2.500   0.7231   0.00836   0.00213  -0.0840   0.3691   1.0000
   2.750   0.7504   0.00854   0.00222  -0.0839   0.3554   1.0000
   3.000   0.7776   0.00873   0.00233  -0.0838   0.3428   1.0000
   3.250   0.8047   0.00892   0.00245  -0.0838   0.3302   1.0000
   3.500   0.8318   0.00912   0.00259  -0.0837   0.3172   1.0000
   3.750   0.8581   0.00941   0.00275  -0.0835   0.2946   1.0000
   4.000   0.8845   0.00971   0.00291  -0.0834   0.2681   1.0000
   4.250   0.9088   0.01034   0.00319  -0.0830   0.2077   1.0000
   4.500   0.9290   0.01167   0.00398  -0.0823   0.1022   1.0000
   4.750   0.9533   0.01230   0.00442  -0.0819   0.0608   1.0000
   5.000   0.9790   0.01269   0.00479  -0.0816   0.0549   1.0000
   5.250   1.0050   0.01298   0.00509  -0.0814   0.0485   1.0000
   5.500   1.0309   0.01328   0.00538  -0.0812   0.0411   1.0000
   5.750   1.0564   0.01366   0.00565  -0.0809   0.0219   1.0000
   6.000   1.0816   0.01407   0.00604  -0.0805   0.0165   1.0000
   6.250   1.1068   0.01447   0.00653  -0.0801   0.0153   1.0000
   6.500   1.1315   0.01495   0.00711  -0.0796   0.0144   1.0000
   6.750   1.1555   0.01552   0.00780  -0.0790   0.0136   1.0000
   7.000   1.1787   0.01618   0.00860  -0.0783   0.0131   1.0000
   7.250   1.2007   0.01698   0.00951  -0.0774   0.0126   1.0000
   7.500   1.2192   0.01816   0.01082  -0.0761   0.0116   1.0000
   7.750   1.2353   0.01952   0.01230  -0.0745   0.0111   1.0000
   8.000   1.2496   0.02097   0.01385  -0.0727   0.0109   1.0000
   8.250   1.2618   0.02253   0.01552  -0.0706   0.0107   1.0000
   8.500   1.2740   0.02406   0.01715  -0.0684   0.0107   1.0000
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