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

GOE 114 (MVA MK.1) AIRFOIL (goe114-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 114 (MVA MK.1) AIRFOIL (goe114-il)
Reynolds number: 50,000
Max Cl/Cd: 32.44 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe114-il-50000-n5.txt
Download as CSV file: xf-goe114-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 114 (MVA MK.1) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3843   0.09878   0.09233  -0.0442   1.0000   0.0566
  -8.500  -0.3911   0.09469   0.08839  -0.0459   1.0000   0.0529
  -8.250  -0.3963   0.09099   0.08479  -0.0472   1.0000   0.0487
  -8.000  -0.3988   0.08799   0.08190  -0.0464   1.0000   0.0470
  -7.750  -0.4065   0.08508   0.07912  -0.0459   1.0000   0.0457
  -7.500  -0.4180   0.08212   0.07631  -0.0454   1.0000   0.0446
  -7.250  -0.4278   0.07864   0.07297  -0.0460   1.0000   0.0431
  -6.750  -0.4498   0.06873   0.06304  -0.0528   1.0000   0.0386
  -6.500  -0.4527   0.06552   0.05984  -0.0521   1.0000   0.0378
  -6.250  -0.4553   0.06180   0.05604  -0.0526   1.0000   0.0371
  -6.000  -0.4550   0.05774   0.05181  -0.0535   1.0000   0.0363
  -5.750  -0.4505   0.05364   0.04744  -0.0547   1.0000   0.0353
  -5.500  -0.4415   0.04955   0.04295  -0.0558   1.0000   0.0342
  -5.250  -0.4282   0.04580   0.03870  -0.0565   1.0000   0.0332
  -5.000  -0.3962   0.04183   0.03379  -0.0597   0.9954   0.0315
  -4.750  -0.3626   0.03904   0.03025  -0.0618   0.9899   0.0304
  -4.500  -0.3310   0.03596   0.02673  -0.0633   0.9851   0.0295
  -4.250  -0.2981   0.03356   0.02393  -0.0645   0.9802   0.0289
  -4.000  -0.2665   0.03165   0.02168  -0.0651   0.9749   0.0286
  -3.750  -0.2337   0.03007   0.01980  -0.0659   0.9696   0.0287
  -3.500  -0.2030   0.02892   0.01833  -0.0665   0.9634   0.0300
  -3.250  -0.1698   0.02745   0.01661  -0.0681   0.9578   0.0333
  -3.000  -0.1365   0.02635   0.01509  -0.0695   0.9514   0.0363
  -2.750  -0.1008   0.02543   0.01373  -0.0713   0.9453   0.0404
  -2.500  -0.0621   0.02342   0.01272  -0.0745   0.9411   0.2382
  -2.250  -0.0326   0.02404   0.01317  -0.0746   0.9323   0.3606
  -2.000  -0.0077   0.02443   0.01354  -0.0740   0.9242   0.4344
  -1.750   0.0241   0.02441   0.01330  -0.0748   0.9175   0.4729
  -1.500   0.0551   0.02429   0.01288  -0.0757   0.9098   0.4825
  -1.250   0.0921   0.02416   0.01242  -0.0778   0.9036   0.4925
  -1.000   0.1224   0.02412   0.01214  -0.0787   0.8954   0.5020
  -0.750   0.1597   0.02399   0.01189  -0.0808   0.8894   0.5124
  -0.500   0.1888   0.02396   0.01174  -0.0814   0.8809   0.5230
  -0.250   0.2271   0.02383   0.01157  -0.0836   0.8752   0.5358
   0.000   0.2544   0.02382   0.01157  -0.0839   0.8661   0.5493
   0.250   0.2911   0.02364   0.01149  -0.0857   0.8603   0.5685
   0.500   0.3175   0.02358   0.01158  -0.0857   0.8510   0.5924
   0.750   0.3459   0.02335   0.01167  -0.0859   0.8431   0.6338
   1.000   0.3889   0.02259   0.01169  -0.0886   0.8364   1.0000
   1.250   0.4169   0.02287   0.01193  -0.0890   0.8269   1.0000
   1.500   0.4531   0.02296   0.01201  -0.0905   0.8205   1.0000
   1.750   0.4771   0.02329   0.01237  -0.0901   0.8099   1.0000
   2.000   0.5046   0.02357   0.01273  -0.0902   0.8008   1.0000
   2.250   0.5373   0.02370   0.01309  -0.0910   0.7932   1.0000
   2.500   0.5607   0.02408   0.01365  -0.0903   0.7825   1.0000
   2.750   0.6034   0.02225   0.01203  -0.0891   0.7481   1.0000
   3.000   0.6281   0.02045   0.01028  -0.0841   0.6799   1.0000
   3.250   0.6458   0.01991   0.00955  -0.0797   0.5907   1.0000
   3.500   0.6419   0.02190   0.00915  -0.0731   0.2317   1.0000
   3.750   0.6459   0.02453   0.01075  -0.0701   0.0442   1.0000
   4.000   0.6638   0.02563   0.01200  -0.0684   0.0382   1.0000
   4.250   0.6820   0.02662   0.01325  -0.0667   0.0350   1.0000
   4.500   0.6964   0.02795   0.01485  -0.0647   0.0312   1.0000
   4.750   0.7095   0.02931   0.01643  -0.0625   0.0292   1.0000
   5.000   0.7217   0.03067   0.01804  -0.0602   0.0285   1.0000
   5.250   0.7325   0.03210   0.01968  -0.0576   0.0282   1.0000
   5.500   0.7429   0.03358   0.02135  -0.0549   0.0281   1.0000
   5.750   0.7588   0.03504   0.02296  -0.0528   0.0282   1.0000
   6.000   0.7978   0.03677   0.02492  -0.0528   0.0282   1.0000
   6.250   0.8550   0.03912   0.02741  -0.0548   0.0275   1.0000
   6.500   0.8990   0.04186   0.03039  -0.0558   0.0271   1.0000
   6.750   0.9342   0.04500   0.03381  -0.0558   0.0283   1.0000
   7.000   0.9655   0.04890   0.03787  -0.0559   0.0298   1.0000
   7.250   0.9863   0.05078   0.04039  -0.0534   0.0324   1.0000
   7.500   1.0055   0.05411   0.04417  -0.0515   0.0354   1.0000
   7.750   1.0319   0.05997   0.05003  -0.0517   0.0387   1.0000
   8.000   1.0359   0.06047   0.05147  -0.0468   0.0434   1.0000
   8.250   1.0600   0.06652   0.05764  -0.0467   0.0495   1.0000
   8.500   1.0551   0.06740   0.05940  -0.0417   0.0553   1.0000
   8.750   1.0658   0.07130   0.06369  -0.0395   0.0644   1.0000
   9.000   1.0616   0.07522   0.06801  -0.0368   0.0700   1.0000
   9.250   1.0820   0.08228   0.07522  -0.0365   0.0859   1.0000
   9.750   1.0770   0.09261   0.08642  -0.0334   0.1267   1.0000
  10.250   1.0147   0.09782   0.09182  -0.0291   0.1125   1.0000
  10.500   0.9915   0.10232   0.09640  -0.0296   0.1125   1.0000
<< Back to GOE 114 (MVA MK.1) AIRFOIL (goe114-il)

Polar data table (+)

Polar graphs


<< Back to GOE 114 (MVA MK.1) AIRFOIL (goe114-il)