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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 500,000
Max Cl/Cd: 86.79 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe113-il-500000.txt
Download as CSV file: xf-goe113-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 113 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4586   0.09905   0.09684  -0.0077   1.0000   0.0119
  -8.000  -0.4540   0.09550   0.09332  -0.0099   1.0000   0.0119
  -7.750  -0.4502   0.09195   0.08979  -0.0121   1.0000   0.0119
  -7.500  -0.4433   0.08798   0.08584  -0.0158   1.0000   0.0120
  -7.250  -0.4334   0.08367   0.08153  -0.0201   1.0000   0.0120
  -7.000  -0.4287   0.07759   0.07547  -0.0245   1.0000   0.0122
  -6.750  -0.4251   0.07355   0.07144  -0.0247   1.0000   0.0125
  -6.500  -0.4161   0.07036   0.06825  -0.0257   1.0000   0.0129
  -6.250  -0.4053   0.06697   0.06484  -0.0277   1.0000   0.0134
  -6.000  -0.3927   0.06337   0.06122  -0.0300   1.0000   0.0141
  -5.750  -0.3783   0.05963   0.05743  -0.0322   1.0000   0.0149
  -5.500  -0.3604   0.05583   0.05356  -0.0346   1.0000   0.0164
  -5.250  -0.3283   0.05290   0.05043  -0.0374   1.0000   0.0178
  -5.000  -0.3101   0.04946   0.04686  -0.0379   1.0000   0.0179
  -4.750  -0.2923   0.04599   0.04324  -0.0380   1.0000   0.0180
  -4.500  -0.2742   0.04257   0.03967  -0.0379   1.0000   0.0180
  -4.250  -0.2558   0.03921   0.03611  -0.0375   1.0000   0.0180
  -4.000  -0.2216   0.03495   0.03157  -0.0402   0.9981   0.0181
  -2.750  -0.0166   0.00509   0.00020  -0.0555   0.9723   0.0281
  -1.750   0.1065   0.01216   0.00640  -0.0594   0.9496   0.0392
  -1.500   0.1389   0.01110   0.00522  -0.0600   0.9302   0.0391
  -1.250   0.1677   0.00990   0.00392  -0.0598   0.9027   0.0399
  -1.000   0.1937   0.00930   0.00318  -0.0590   0.8624   0.0404
  -0.750   0.2184   0.00903   0.00269  -0.0580   0.8150   0.0408
  -0.500   0.2437   0.00884   0.00230  -0.0573   0.7751   0.0423
  -0.250   0.2700   0.00873   0.00201  -0.0568   0.7426   0.0426
   0.000   0.2969   0.00865   0.00179  -0.0565   0.7155   0.0425
   0.250   0.3240   0.00858   0.00160  -0.0563   0.6930   0.0434
   0.500   0.3512   0.00855   0.00145  -0.0561   0.6702   0.0458
   0.750   0.3785   0.00855   0.00134  -0.0559   0.6480   0.0498
   1.000   0.4058   0.00857   0.00129  -0.0557   0.6280   0.0599
   1.250   0.4284   0.00626   0.00138  -0.0551   0.6115   1.0000
   1.500   0.4557   0.00638   0.00138  -0.0549   0.5937   1.0000
   1.750   0.4830   0.00650   0.00140  -0.0548   0.5774   1.0000
   2.000   0.5104   0.00663   0.00145  -0.0547   0.5629   1.0000
   2.250   0.5377   0.00675   0.00150  -0.0546   0.5482   1.0000
   2.500   0.5651   0.00688   0.00157  -0.0546   0.5333   1.0000
   2.750   0.5923   0.00702   0.00165  -0.0545   0.5177   1.0000
   3.000   0.6195   0.00717   0.00176  -0.0544   0.4980   1.0000
   3.250   0.6457   0.00744   0.00184  -0.0541   0.4525   1.0000
   3.500   0.6713   0.00780   0.00195  -0.0539   0.3927   1.0000
   3.750   0.6960   0.00838   0.00217  -0.0536   0.3109   1.0000
   4.000   0.7204   0.00905   0.00249  -0.0533   0.2438   1.0000
   4.250   0.7460   0.00950   0.00279  -0.0531   0.2079   1.0000
   4.500   0.7716   0.00993   0.00304  -0.0529   0.1686   1.0000
   4.750   0.7960   0.01060   0.00337  -0.0526   0.1121   1.0000
   5.000   0.8176   0.01180   0.00408  -0.0519   0.0225   1.0000
   5.250   0.8437   0.01220   0.00457  -0.0515   0.0179   1.0000
   5.500   0.8695   0.01267   0.00523  -0.0510   0.0162   1.0000
   5.750   0.8946   0.01325   0.00594  -0.0505   0.0156   1.0000
   6.000   0.9190   0.01394   0.00675  -0.0499   0.0148   1.0000
   6.250   0.9427   0.01471   0.00760  -0.0492   0.0134   1.0000
   6.500   0.9653   0.01564   0.00865  -0.0483   0.0128   1.0000
   6.750   0.9867   0.01674   0.00989  -0.0473   0.0122   1.0000
   7.000   1.0070   0.01805   0.01130  -0.0460   0.0118   1.0000
   7.250   1.0271   0.01956   0.01294  -0.0445   0.0120   1.0000
   7.500   1.0474   0.02169   0.01525  -0.0427   0.0137   1.0000
   9.000   1.1590   0.03633   0.03130  -0.0341   0.0144   1.0000
   9.250   1.1716   0.03854   0.03377  -0.0324   0.0132   1.0000
   9.500   1.1804   0.04111   0.03659  -0.0306   0.0123   1.0000
   9.750   1.1862   0.04389   0.03959  -0.0287   0.0117   1.0000
  10.000   1.1895   0.04683   0.04272  -0.0268   0.0113   1.0000
  10.250   1.1912   0.05004   0.04607  -0.0249   0.0109   1.0000
  10.500   1.1853   0.05352   0.04975  -0.0227   0.0106   1.0000
  10.750   1.1723   0.05685   0.05326  -0.0197   0.0105   1.0000
  11.000   1.1558   0.06034   0.05691  -0.0174   0.0104   1.0000
  11.250   1.1389   0.06417   0.06092  -0.0167   0.0103   1.0000
  11.500   1.1213   0.06871   0.06562  -0.0176   0.0103   1.0000
  11.750   1.1042   0.07387   0.07094  -0.0200   0.0103   1.0000
  12.000   1.0885   0.07947   0.07670  -0.0236   0.0104   1.0000
  12.250   1.0713   0.08619   0.08357  -0.0284   0.0104   1.0000
  12.500   1.0537   0.09386   0.09141  -0.0343   0.0106   1.0000
  12.750   1.0338   0.10333   0.10103  -0.0415   0.0108   1.0000
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