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

GOE 115 (MVA MK.2) AIRFOIL (goe115-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 115 (MVA MK.2) AIRFOIL (goe115-il)
Reynolds number: 50,000
Max Cl/Cd: 37.86 at α=6.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe115-il-50000-n5.txt
Download as CSV file: xf-goe115-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 115 (MVA MK.2) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3802   0.11404   0.10702  -0.0337   1.0000   0.0799
  -9.750  -0.3801   0.10907   0.10209  -0.0359   1.0000   0.0668
  -9.250  -0.3918   0.09948   0.09275  -0.0421   1.0000   0.0578
  -9.000  -0.3860   0.09607   0.08934  -0.0421   1.0000   0.0561
  -8.750  -0.3840   0.09258   0.08592  -0.0424   1.0000   0.0550
  -8.500  -0.3854   0.08898   0.08241  -0.0431   1.0000   0.0540
  -8.250  -0.3895   0.08548   0.07902  -0.0437   1.0000   0.0531
  -8.000  -0.3973   0.08195   0.07564  -0.0442   1.0000   0.0522
  -7.750  -0.4054   0.07811   0.07194  -0.0453   1.0000   0.0514
  -7.500  -0.4135   0.07351   0.06747  -0.0477   1.0000   0.0504
  -7.250  -0.4235   0.06752   0.06156  -0.0516   1.0000   0.0491
  -7.000  -0.4361   0.06058   0.05452  -0.0564   1.0000   0.0476
  -6.750  -0.4458   0.05443   0.04795  -0.0600   1.0000   0.0463
  -6.500  -0.4444   0.05047   0.04366  -0.0605   1.0000   0.0461
  -6.250  -0.4391   0.04728   0.04032  -0.0600   1.0000   0.0472
  -6.000  -0.4310   0.04473   0.03764  -0.0594   1.0000   0.0490
  -5.750  -0.3995   0.04045   0.03275  -0.0638   0.9918   0.0516
  -5.500  -0.3643   0.03644   0.02798  -0.0675   0.9829   0.0531
  -5.250  -0.3280   0.03353   0.02433  -0.0704   0.9742   0.0560
  -5.000  -0.2936   0.03107   0.02154  -0.0727   0.9655   0.0605
  -4.750  -0.2600   0.02920   0.01931  -0.0742   0.9560   0.0635
  -4.500  -0.2247   0.02779   0.01749  -0.0759   0.9471   0.0692
  -4.250  -0.1895   0.02632   0.01588  -0.0778   0.9385   0.0753
  -4.000  -0.1558   0.02527   0.01458  -0.0791   0.9285   0.0799
  -3.750  -0.1176   0.02425   0.01330  -0.0814   0.9206   0.0865
  -3.500  -0.0815   0.02321   0.01226  -0.0835   0.9117   0.1024
  -3.250  -0.0474   0.02242   0.01218  -0.0853   0.9025   0.2255
  -3.000  -0.0096   0.02267   0.01205  -0.0870   0.8945   0.2998
  -2.750   0.0194   0.02283   0.01203  -0.0873   0.8835   0.3387
  -2.500   0.0508   0.02278   0.01177  -0.0880   0.8741   0.3614
  -2.000   0.1149   0.02255   0.01111  -0.0896   0.8565   0.3946
  -1.750   0.1448   0.02245   0.01098  -0.0900   0.8481   0.4244
  -1.500   0.1744   0.02232   0.01083  -0.0903   0.8398   0.4534
  -1.250   0.2016   0.02221   0.01074  -0.0903   0.8308   0.4760
  -1.000   0.2341   0.02202   0.01048  -0.0911   0.8241   0.4932
  -0.750   0.2602   0.02200   0.01045  -0.0910   0.8150   0.5096
  -0.500   0.2922   0.02181   0.01026  -0.0917   0.8091   0.5304
  -0.250   0.3166   0.02180   0.01036  -0.0914   0.8000   0.5567
   0.000   0.3469   0.02149   0.01026  -0.0916   0.7944   0.6038
   0.250   0.3691   0.02108   0.01046  -0.0904   0.7860   0.7325
   0.500   0.4164   0.02093   0.01031  -0.0942   0.7809   1.0000
   0.750   0.4406   0.02137   0.01059  -0.0939   0.7728   1.0000
   1.000   0.4697   0.02163   0.01071  -0.0943   0.7668   1.0000
   1.250   0.4933   0.02210   0.01108  -0.0939   0.7594   1.0000
   1.500   0.5204   0.02245   0.01135  -0.0939   0.7534   1.0000
   1.750   0.5452   0.02289   0.01176  -0.0937   0.7469   1.0000
   2.000   0.5698   0.02334   0.01219  -0.0934   0.7402   1.0000
   2.250   0.5983   0.02365   0.01248  -0.0935   0.7352   1.0000
   2.500   0.6183   0.02430   0.01319  -0.0927   0.7271   1.0000
   2.750   0.6477   0.02456   0.01347  -0.0928   0.7220   1.0000
   3.000   0.6675   0.02515   0.01413  -0.0918   0.7126   1.0000
   3.250   0.6957   0.02531   0.01437  -0.0915   0.7050   1.0000
   3.500   0.7194   0.02566   0.01480  -0.0907   0.6954   1.0000
   3.750   0.7421   0.02610   0.01535  -0.0898   0.6862   1.0000
   4.000   0.7703   0.02632   0.01570  -0.0895   0.6794   1.0000
   4.250   0.7900   0.02699   0.01657  -0.0885   0.6702   1.0000
   4.500   0.8198   0.02716   0.01691  -0.0883   0.6641   1.0000
   4.750   0.8377   0.02795   0.01790  -0.0871   0.6542   1.0000
   5.000   0.8639   0.02831   0.01852  -0.0866   0.6466   1.0000
   5.250   0.8863   0.02888   0.01936  -0.0857   0.6376   1.0000
   5.500   0.9068   0.02955   0.02031  -0.0846   0.6279   1.0000
   5.750   0.9410   0.02809   0.01906  -0.0825   0.6038   1.0000
   6.000   0.9608   0.02638   0.01742  -0.0777   0.5562   1.0000
   6.250   0.9693   0.02560   0.01671  -0.0724   0.4920   1.0000
   6.500   0.9679   0.02579   0.01595  -0.0659   0.3221   1.0000
   6.750   0.9469   0.02927   0.01782  -0.0604   0.1255   1.0000
   7.000   0.9385   0.03208   0.02010  -0.0566   0.0841   1.0000
   7.250   0.9368   0.03430   0.02225  -0.0535   0.0708   1.0000
   7.500   0.9369   0.03648   0.02446  -0.0509   0.0628   1.0000
   7.750   0.9386   0.03860   0.02672  -0.0486   0.0578   1.0000
   8.000   0.9388   0.04098   0.02919  -0.0466   0.0546   1.0000
   8.250   0.9428   0.04314   0.03147  -0.0447   0.0514   1.0000
   8.500   0.9515   0.04501   0.03354  -0.0429   0.0472   1.0000
   8.750   0.9629   0.04681   0.03541  -0.0412   0.0436   1.0000
   9.000   1.0095   0.04790   0.03659  -0.0394   0.0402   1.0000
   9.250   1.0564   0.04962   0.03866  -0.0394   0.0364   1.0000
   9.500   1.0977   0.05239   0.04165  -0.0402   0.0335   1.0000
   9.750   1.1318   0.05600   0.04555  -0.0408   0.0324   1.0000
  10.000   1.1522   0.05974   0.04960  -0.0403   0.0318   1.0000
  10.250   1.1657   0.06386   0.05401  -0.0394   0.0313   1.0000
  10.500   1.1682   0.06756   0.05799  -0.0376   0.0309   1.0000
  10.750   1.1608   0.07039   0.06119  -0.0347   0.0306   1.0000
  11.000   1.1511   0.07342   0.06454  -0.0322   0.0304   1.0000
  11.250   1.1390   0.07671   0.06814  -0.0303   0.0303   1.0000
  11.500   1.1252   0.08032   0.07202  -0.0290   0.0301   1.0000
  11.750   1.1103   0.08431   0.07628  -0.0285   0.0301   1.0000
  12.000   1.0957   0.08863   0.08082  -0.0286   0.0302   1.0000
  12.250   1.0787   0.09340   0.08581  -0.0294   0.0302   1.0000
  12.500   1.0615   0.09857   0.09117  -0.0309   0.0302   1.0000
  12.750   1.0460   0.10401   0.09677  -0.0329   0.0304   1.0000
  13.000   1.0299   0.10990   0.10280  -0.0355   0.0306   1.0000
  13.250   1.0148   0.11608   0.10910  -0.0385   0.0307   1.0000
  13.500   1.0022   0.12240   0.11552  -0.0418   0.0309   1.0000
<< Back to GOE 115 (MVA MK.2) AIRFOIL (goe115-il)

Polar data table (+)

Polar graphs


<< Back to GOE 115 (MVA MK.2) AIRFOIL (goe115-il)