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GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 415 AIRFOIL (goe415-il)
Reynolds number: 500,000
Max Cl/Cd: 92.2 at α=6°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe415-il-500000.txt
Download as CSV file: xf-goe415-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 415 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4185   0.09168   0.08961  -0.0084   1.0000   0.0115
  -7.750  -0.4134   0.08787   0.08583  -0.0110   1.0000   0.0115
  -7.500  -0.4161   0.08244   0.08044  -0.0130   1.0000   0.0119
  -7.250  -0.4078   0.07894   0.07696  -0.0152   1.0000   0.0121
  -7.000  -0.3977   0.07529   0.07332  -0.0182   1.0000   0.0123
  -6.750  -0.3859   0.07140   0.06944  -0.0220   1.0000   0.0126
  -6.500  -0.3727   0.06730   0.06533  -0.0262   1.0000   0.0129
  -6.250  -0.3587   0.06307   0.06109  -0.0302   1.0000   0.0132
  -6.000  -0.3453   0.05885   0.05685  -0.0336   1.0000   0.0137
  -5.750  -0.3338   0.05477   0.05271  -0.0359   1.0000   0.0142
  -5.500  -0.3129   0.04981   0.04763  -0.0398   0.9984   0.0152
  -5.250  -0.2616   0.04378   0.04119  -0.0474   0.9937   0.0167
  -5.000  -0.2260   0.03921   0.03627  -0.0518   0.9876   0.0168
  -4.750  -0.1986   0.03149   0.02833  -0.0568   0.9823   0.0181
  -4.500  -0.1684   0.02930   0.02606  -0.0590   0.9734   0.0194
  -4.250  -0.1356   0.02662   0.02312  -0.0608   0.9636   0.0209
  -4.000  -0.1017   0.02461   0.02078  -0.0617   0.9521   0.0232
  -3.750  -0.0690   0.02470   0.02054  -0.0614   0.9372   0.0246
  -3.500  -0.0423   0.02293   0.01843  -0.0609   0.9204   0.0247
  -3.250  -0.0199   0.01784   0.01285  -0.0605   0.9039   0.0264
  -3.000   0.0047   0.01620   0.01104  -0.0597   0.8848   0.0277
  -2.750   0.0319   0.01443   0.00893  -0.0582   0.8620   0.0224
  -2.500   0.0578   0.01310   0.00733  -0.0571   0.8320   0.0222
  -2.250   0.0832   0.01258   0.00652  -0.0559   0.7838   0.0235
  -2.000   0.1079   0.01162   0.00521  -0.0547   0.7319   0.0233
  -1.750   0.1337   0.01124   0.00456  -0.0538   0.6897   0.0243
  -1.500   0.1593   0.01041   0.00353  -0.0532   0.6546   0.0259
  -1.250   0.1853   0.00995   0.00294  -0.0527   0.6234   0.0301
  -1.000   0.2121   0.00970   0.00252  -0.0523   0.5966   0.0332
  -0.750   0.2392   0.00958   0.00225  -0.0520   0.5742   0.0358
  -0.500   0.2664   0.00931   0.00183  -0.0518   0.5555   0.0441
  -0.250   0.2938   0.00923   0.00163  -0.0515   0.5395   0.0550
   0.000   0.3179   0.00803   0.00163  -0.0515   0.5264   0.4691
   0.250   0.3547   0.00670   0.00165  -0.0532   0.5125   1.0000
   0.500   0.3819   0.00682   0.00163  -0.0530   0.5002   1.0000
   0.750   0.4091   0.00695   0.00163  -0.0528   0.4881   1.0000
   1.000   0.4363   0.00707   0.00164  -0.0525   0.4758   1.0000
   1.250   0.4636   0.00718   0.00166  -0.0524   0.4628   1.0000
   1.500   0.4908   0.00730   0.00168  -0.0522   0.4491   1.0000
   1.750   0.5180   0.00743   0.00172  -0.0520   0.4347   1.0000
   2.000   0.5452   0.00757   0.00176  -0.0518   0.4195   1.0000
   2.250   0.5723   0.00771   0.00181  -0.0516   0.4043   1.0000
   2.500   0.5994   0.00786   0.00188  -0.0515   0.3902   1.0000
   2.750   0.6265   0.00803   0.00198  -0.0513   0.3772   1.0000
   3.000   0.6535   0.00819   0.00208  -0.0511   0.3647   1.0000
   3.250   0.6806   0.00836   0.00220  -0.0510   0.3528   1.0000
   3.500   0.7075   0.00854   0.00232  -0.0508   0.3410   1.0000
   3.750   0.7345   0.00873   0.00248  -0.0506   0.3307   1.0000
   4.000   0.7612   0.00895   0.00264  -0.0505   0.3212   1.0000
   4.250   0.7882   0.00911   0.00279  -0.0503   0.3117   1.0000
   4.500   0.8151   0.00931   0.00297  -0.0502   0.3029   1.0000
   4.750   0.8417   0.00953   0.00317  -0.0500   0.2946   1.0000
   5.000   0.8686   0.00970   0.00334  -0.0499   0.2837   1.0000
   5.250   0.8953   0.00989   0.00353  -0.0497   0.2740   1.0000
   5.500   0.9219   0.01011   0.00371  -0.0496   0.2610   1.0000
   5.750   0.9483   0.01033   0.00393  -0.0494   0.2475   1.0000
   6.000   0.9746   0.01057   0.00415  -0.0492   0.2338   1.0000
   6.250   1.0004   0.01089   0.00439  -0.0490   0.2120   1.0000
   6.500   1.0245   0.01147   0.00473  -0.0487   0.1631   1.0000
   6.750   1.0436   0.01288   0.00564  -0.0479   0.0853   1.0000
   7.000   1.0621   0.01443   0.00683  -0.0468   0.0194   1.0000
   7.250   1.0861   0.01506   0.00758  -0.0462   0.0165   1.0000
   7.500   1.1091   0.01584   0.00852  -0.0455   0.0143   1.0000
   7.750   1.1303   0.01688   0.00974  -0.0446   0.0133   1.0000
   8.000   1.1524   0.01768   0.01066  -0.0438   0.0128   1.0000
   8.250   1.1728   0.01868   0.01180  -0.0428   0.0123   1.0000
   8.500   1.1918   0.01980   0.01305  -0.0417   0.0118   1.0000
   8.750   1.2100   0.02091   0.01426  -0.0406   0.0109   1.0000
   9.000   1.2266   0.02214   0.01559  -0.0393   0.0103   1.0000
   9.250   1.2408   0.02356   0.01712  -0.0378   0.0099   1.0000
   9.500   1.2530   0.02512   0.01877  -0.0361   0.0097   1.0000
   9.750   1.2627   0.02685   0.02060  -0.0341   0.0094   1.0000
  10.000   1.2706   0.02876   0.02259  -0.0320   0.0092   1.0000
  10.250   1.2778   0.03078   0.02468  -0.0298   0.0091   1.0000
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