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

GOE 207 (AVIATIK V8) AIRFOIL (goe207-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 207 (AVIATIK V8) AIRFOIL (goe207-il)
Reynolds number: 200,000
Max Cl/Cd: 70.43 at α=9.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe207-il-200000.txt
Download as CSV file: xf-goe207-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 207 (AVIATIK V8) 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.2312   0.09830   0.09526  -0.0148   1.0000   0.0302
  -7.500  -0.2241   0.09577   0.09279  -0.0162   1.0000   0.0308
  -7.250  -0.2187   0.09346   0.09056  -0.0176   1.0000   0.0313
  -7.000  -0.1941   0.09017   0.08728  -0.0252   0.9687   0.0320
  -6.750  -0.1555   0.08753   0.08452  -0.0404   0.9264   0.0324
  -6.500  -0.1381   0.08378   0.08069  -0.0431   0.8943   0.0326
  -6.250  -0.1315   0.08003   0.07687  -0.0401   0.8656   0.0328
  -6.000  -0.1210   0.07704   0.07378  -0.0396   0.8388   0.0331
  -5.750  -0.1071   0.07428   0.07091  -0.0406   0.8125   0.0337
  -5.250  -0.0700   0.06891   0.06526  -0.0455   0.7657   0.0354
  -5.000  -0.0471   0.06618   0.06238  -0.0489   0.7444   0.0363
  -4.750  -0.0139   0.06384   0.05982  -0.0549   0.7244   0.0375
  -4.500   0.0328   0.06198   0.05761  -0.0642   0.7060   0.0381
  -4.250   0.0434   0.05807   0.05367  -0.0629   0.6893   0.0384
  -4.000   0.0611   0.05510   0.05062  -0.0633   0.6730   0.0389
  -3.750   0.0846   0.05256   0.04796  -0.0650   0.6574   0.0397
  -3.500   0.1120   0.05020   0.04544  -0.0674   0.6425   0.0409
  -3.250   0.1441   0.04807   0.04312  -0.0704   0.6284   0.0431
  -3.000   0.1959   0.04722   0.04181  -0.0762   0.6148   0.0447
  -2.750   0.2135   0.04373   0.03828  -0.0766   0.6030   0.0452
  -2.500   0.2372   0.04132   0.03579  -0.0776   0.5902   0.0461
  -2.250   0.2655   0.03938   0.03373  -0.0791   0.5781   0.0476
  -2.000   0.2969   0.03771   0.03184  -0.0808   0.5670   0.0498
  -1.750   0.3431   0.03765   0.03126  -0.0832   0.5557   0.0526
  -1.500   0.3654   0.03445   0.02811  -0.0843   0.5448   0.0535
  -1.250   0.3920   0.03268   0.02620  -0.0853   0.5347   0.0551
  -1.000   0.4226   0.03140   0.02476  -0.0863   0.5236   0.0585
  -0.750   0.4596   0.03062   0.02363  -0.0873   0.5129   0.0624
  -0.500   0.4860   0.02879   0.02168  -0.0882   0.5030   0.0640
  -0.250   0.5158   0.02759   0.02039  -0.0890   0.4918   0.0671
   0.000   0.5516   0.02760   0.01997  -0.0892   0.4817   0.0725
   0.250   0.5785   0.02560   0.01793  -0.0902   0.4720   0.0745
   0.500   0.6080   0.02466   0.01690  -0.0908   0.4615   0.0790
   0.750   0.6395   0.02400   0.01592  -0.0911   0.4526   0.0853
   1.000   0.6683   0.02301   0.01489  -0.0917   0.4423   0.0908
   1.250   0.6984   0.02232   0.01398  -0.0920   0.4333   0.0998
   1.500   0.7274   0.02167   0.01320  -0.0924   0.4241   0.1140
   1.750   0.7559   0.02105   0.01248  -0.0928   0.4152   0.1312
   2.000   0.7843   0.02052   0.01180  -0.0931   0.4069   0.1582
   2.250   0.8123   0.01983   0.01108  -0.0936   0.3984   0.1908
   2.500   0.8406   0.01922   0.01035  -0.0938   0.3910   0.2219
   2.750   0.8701   0.01912   0.01010  -0.0937   0.3837   0.2310
   3.000   0.9047   0.01869   0.00916  -0.0915   0.3779   0.0763
   3.250   0.9326   0.01853   0.00883  -0.0913   0.3723   0.0763
   3.500   0.9609   0.01830   0.00857  -0.0911   0.3660   0.0744
   3.750   0.9886   0.01817   0.00838  -0.0910   0.3607   0.0740
   4.000   1.0162   0.01819   0.00835  -0.0910   0.3560   0.0759
   4.250   1.0440   0.01818   0.00843  -0.0910   0.3509   0.0807
   4.500   1.0715   0.01819   0.00844  -0.0910   0.3462   0.0911
   4.750   1.0992   0.01837   0.00848  -0.0910   0.3421   0.1120
   5.000   1.1202   0.01721   0.00866  -0.0895   0.3382   1.0000
   5.250   1.1472   0.01754   0.00894  -0.0894   0.3338   1.0000
   5.500   1.1740   0.01789   0.00922  -0.0894   0.3300   1.0000
   5.750   1.2009   0.01836   0.00957  -0.0893   0.3267   1.0000
   6.000   1.2274   0.01880   0.01004  -0.0893   0.3235   1.0000
   6.250   1.2536   0.01913   0.01043  -0.0892   0.3194   1.0000
   6.500   1.2796   0.01937   0.01063  -0.0890   0.3143   1.0000
   6.750   1.3055   0.01977   0.01094  -0.0889   0.3092   1.0000
   7.000   1.3306   0.01999   0.01129  -0.0886   0.3043   1.0000
   7.250   1.3559   0.02020   0.01151  -0.0884   0.2990   1.0000
   7.500   1.3811   0.02058   0.01178  -0.0882   0.2938   1.0000
   7.750   1.4054   0.02076   0.01215  -0.0878   0.2885   1.0000
   8.000   1.4299   0.02097   0.01239  -0.0875   0.2832   1.0000
   8.250   1.4539   0.02130   0.01269  -0.0871   0.2778   1.0000
   8.500   1.4772   0.02142   0.01298  -0.0866   0.2707   1.0000
   8.750   1.4996   0.02161   0.01316  -0.0860   0.2624   1.0000
   9.000   1.5221   0.02172   0.01341  -0.0855   0.2527   1.0000
   9.250   1.5438   0.02192   0.01368  -0.0848   0.2404   1.0000
   9.500   1.5648   0.02224   0.01402  -0.0841   0.2258   1.0000
   9.750   1.5841   0.02277   0.01452  -0.0832   0.2068   1.0000
  10.000   1.6010   0.02358   0.01522  -0.0821   0.1720   1.0000
  10.250   1.5902   0.02671   0.01778  -0.0783   0.1048   1.0000
  10.750   1.5688   0.03193   0.02279  -0.0708   0.0448   1.0000
  11.000   1.5645   0.03444   0.02542  -0.0689   0.0415   1.0000
  11.250   1.5627   0.03705   0.02818  -0.0678   0.0397   1.0000
  11.500   1.5604   0.04001   0.03134  -0.0673   0.0387   1.0000
  11.750   1.5556   0.04351   0.03504  -0.0673   0.0380   1.0000
  12.000   1.5480   0.04762   0.03934  -0.0677   0.0373   1.0000
  12.250   1.5375   0.05230   0.04422  -0.0685   0.0369   1.0000
  12.500   1.5245   0.05753   0.04963  -0.0697   0.0365   1.0000
  12.750   1.5088   0.06320   0.05549  -0.0709   0.0362   1.0000
  13.000   1.4914   0.06920   0.06167  -0.0723   0.0360   1.0000
  13.250   1.4727   0.07542   0.06806  -0.0738   0.0358   1.0000
  13.500   1.4535   0.08180   0.07460  -0.0753   0.0357   1.0000
  13.750   1.4346   0.08834   0.08130  -0.0770   0.0355   1.0000
  14.000   1.4168   0.09487   0.08798  -0.0788   0.0354   1.0000
<< Back to GOE 207 (AVIATIK V8) AIRFOIL (goe207-il)

Polar data table (+)

Polar graphs


<< Back to GOE 207 (AVIATIK V8) AIRFOIL (goe207-il)