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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 200,000
Max Cl/Cd: 69.96 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe113-il-200000.txt
Download as CSV file: xf-goe113-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4520   0.09753   0.09408  -0.0063   1.0000   0.0248
  -7.750  -0.4481   0.09471   0.09131  -0.0096   1.0000   0.0259
  -7.500  -0.4437   0.09194   0.08859  -0.0150   1.0000   0.0265
  -7.250  -0.4317   0.08837   0.08502  -0.0218   1.0000   0.0268
  -7.000  -0.4187   0.08444   0.08107  -0.0270   1.0000   0.0269
  -6.750  -0.4055   0.08043   0.07701  -0.0305   1.0000   0.0270
  -6.500  -0.3920   0.07641   0.07293  -0.0334   1.0000   0.0270
  -6.250  -0.3782   0.07245   0.06889  -0.0356   1.0000   0.0271
  -6.000  -0.3756   0.06549   0.06198  -0.0365   1.0000   0.0277
  -5.750  -0.3680   0.06147   0.05798  -0.0361   1.0000   0.0284
  -5.500  -0.3566   0.05801   0.05451  -0.0362   1.0000   0.0291
  -5.250  -0.3422   0.05460   0.05104  -0.0370   1.0000   0.0300
  -5.000  -0.3254   0.05115   0.04751  -0.0380   1.0000   0.0311
  -4.750  -0.3066   0.04769   0.04392  -0.0391   1.0000   0.0325
  -4.500  -0.2857   0.04426   0.04029  -0.0400   1.0000   0.0344
  -4.250  -0.2542   0.04238   0.03794  -0.0405   1.0000   0.0381
  -4.000  -0.2344   0.03766   0.03287  -0.0410   1.0000   0.0394
  -3.750  -0.2196   0.03417   0.02950  -0.0409   1.0000   0.0418
  -3.500  -0.1923   0.03523   0.02997  -0.0395   1.0000   0.0515
  -3.250  -0.1773   0.02956   0.02448  -0.0404   1.0000   0.0555
  -3.000  -0.1536   0.02802   0.02252  -0.0398   1.0000   0.0665
  -2.750  -0.1346   0.02597   0.02053  -0.0394   1.0000   0.0735
  -2.500  -0.1128   0.02480   0.01898  -0.0390   1.0000   0.0939
  -2.250  -0.0764   0.02244   0.01664  -0.0423   0.9964   0.1240
  -2.000  -0.0379   0.02057   0.01486  -0.0462   0.9911   0.1820
  -1.750   0.0206   0.01744   0.01032  -0.0457   0.9869   0.0633
  -1.500   0.0658   0.01592   0.00855  -0.0488   0.9820   0.0637
  -1.250   0.1075   0.01437   0.00686  -0.0513   0.9738   0.0619
  -1.000   0.1515   0.01332   0.00577  -0.0543   0.9663   0.0623
  -0.750   0.1952   0.01253   0.00498  -0.0574   0.9565   0.0657
  -0.500   0.2346   0.01172   0.00422  -0.0595   0.9411   0.0661
  -0.250   0.2705   0.01111   0.00358  -0.0608   0.9191   0.0685
   0.000   0.3046   0.01072   0.00312  -0.0616   0.8943   0.0739
   0.250   0.3354   0.01042   0.00277  -0.0617   0.8663   0.0885
   0.500   0.3616   0.00793   0.00264  -0.0607   0.8388   1.0000
   0.750   0.3876   0.00806   0.00249  -0.0598   0.8083   1.0000
   1.000   0.4131   0.00823   0.00241  -0.0589   0.7803   1.0000
   1.250   0.4387   0.00842   0.00240  -0.0582   0.7547   1.0000
   1.500   0.4643   0.00861   0.00241  -0.0575   0.7300   1.0000
   1.750   0.4901   0.00881   0.00244  -0.0568   0.7073   1.0000
   2.000   0.5162   0.00899   0.00251  -0.0564   0.6858   1.0000
   2.250   0.5422   0.00919   0.00258  -0.0559   0.6664   1.0000
   2.500   0.5686   0.00935   0.00268  -0.0555   0.6475   1.0000
   2.750   0.5951   0.00953   0.00280  -0.0551   0.6299   1.0000
   3.000   0.6215   0.00972   0.00296  -0.0548   0.6129   1.0000
   3.250   0.6479   0.00990   0.00311  -0.0544   0.5952   1.0000
   3.500   0.6743   0.01008   0.00327  -0.0541   0.5762   1.0000
   3.750   0.7003   0.01028   0.00344  -0.0537   0.5564   1.0000
   4.000   0.7255   0.01047   0.00357  -0.0531   0.5216   1.0000
   4.250   0.7498   0.01075   0.00373  -0.0523   0.4729   1.0000
   4.500   0.7745   0.01107   0.00393  -0.0517   0.4206   1.0000
   4.750   0.7959   0.01182   0.00418  -0.0508   0.3180   1.0000
   5.000   0.8174   0.01275   0.00466  -0.0502   0.2459   1.0000
   5.250   0.8414   0.01337   0.00513  -0.0497   0.2113   1.0000
   5.500   0.8658   0.01396   0.00565  -0.0493   0.1695   1.0000
   5.750   0.8850   0.01544   0.00641  -0.0483   0.0538   1.0000
   6.000   0.9071   0.01658   0.00751  -0.0473   0.0317   1.0000
   6.250   0.9306   0.01745   0.00851  -0.0464   0.0279   1.0000
   6.500   0.9526   0.01855   0.00983  -0.0454   0.0260   1.0000
   6.750   0.9717   0.02000   0.01146  -0.0439   0.0253   1.0000
   7.000   0.9917   0.02129   0.01290  -0.0426   0.0246   1.0000
   7.250   1.0119   0.02258   0.01429  -0.0414   0.0229   1.0000
   7.500   1.0315   0.02435   0.01614  -0.0399   0.0227   1.0000
   7.750   1.0525   0.02641   0.01830  -0.0385   0.0229   1.0000
   8.000   1.0745   0.02890   0.02095  -0.0373   0.0236   1.0000
   8.250   1.0961   0.03240   0.02466  -0.0362   0.0247   1.0000
   8.500   1.1209   0.03402   0.02648  -0.0351   0.0266   1.0000
   8.750   1.1512   0.04716   0.04118  -0.0307   0.0703   1.0000
  14.000   0.7140   0.15476   0.15159  -0.0489   0.0648   1.0000
  14.250   0.7139   0.15768   0.15452  -0.0498   0.0627   1.0000
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