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GOE 767 AIRFOIL (goe767-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 767 AIRFOIL (goe767-il)
Reynolds number: 50,000
Max Cl/Cd: 15.49 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe767-il-50000.txt
Download as CSV file: xf-goe767-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 767 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4538   0.09580   0.08843   0.0273   1.0000   0.3853
  -7.750  -0.4445   0.09258   0.08524   0.0280   1.0000   0.4021
  -7.500  -0.6108   0.06871   0.06133  -0.0049   1.0000   0.2098
  -7.250  -0.6091   0.06315   0.05566  -0.0062   1.0000   0.2035
  -7.000  -0.6123   0.05741   0.04971  -0.0074   1.0000   0.2001
  -6.750  -0.6122   0.05233   0.04435  -0.0079   1.0000   0.2007
  -6.500  -0.6082   0.04779   0.03942  -0.0078   1.0000   0.2034
  -6.250  -0.6041   0.04350   0.03443  -0.0072   1.0000   0.2077
  -6.000  -0.5822   0.04151   0.03271  -0.0060   1.0000   0.2169
  -5.750  -0.5709   0.03849   0.02904  -0.0049   1.0000   0.2258
  -5.500  -0.5489   0.03681   0.02761  -0.0037   1.0000   0.2371
  -5.250  -0.5310   0.03468   0.02528  -0.0026   1.0000   0.2479
  -5.000  -0.5111   0.03302   0.02342  -0.0014   1.0000   0.2594
  -4.750  -0.4896   0.03141   0.02184  -0.0003   1.0000   0.2693
  -4.500  -0.4690   0.02986   0.01992   0.0007   1.0000   0.2797
  -4.250  -0.4458   0.02845   0.01868   0.0017   1.0000   0.2889
  -4.000  -0.4234   0.02712   0.01714   0.0027   1.0000   0.2992
  -3.750  -0.3999   0.02595   0.01607   0.0036   1.0000   0.3093
  -3.500  -0.3767   0.02482   0.01492   0.0045   1.0000   0.3206
  -3.250  -0.3534   0.02387   0.01396   0.0054   1.0000   0.3340
  -3.000  -0.3300   0.02291   0.01320   0.0063   1.0000   0.3480
  -2.750  -0.3080   0.02209   0.01256   0.0071   1.0000   0.3656
  -2.500  -0.2897   0.02144   0.01215   0.0081   1.0000   0.3861
  -2.250  -0.2912   0.02153   0.01248   0.0100   1.0000   0.4005
  -2.000  -0.1855   0.01968   0.01187  -0.0040   0.9389   0.5473
  -1.750  -0.0573   0.01954   0.01295  -0.0112   0.8782   0.8984
  -1.500   0.1181   0.02001   0.01231  -0.0323   0.7697   0.9936
  -1.250   0.1434   0.01999   0.01188  -0.0327   0.7271   1.0000
  -1.000   0.1595   0.02014   0.01172  -0.0313   0.6967   1.0000
  -0.750   0.1772   0.02037   0.01172  -0.0300   0.6696   1.0000
  -0.500   0.1960   0.02067   0.01182  -0.0289   0.6458   1.0000
  -0.250   0.2150   0.02106   0.01204  -0.0276   0.6251   1.0000
   0.000   0.2338   0.02152   0.01235  -0.0263   0.6062   1.0000
   0.250   0.2534   0.02204   0.01273  -0.0252   0.5886   1.0000
   0.500   0.2737   0.02262   0.01320  -0.0241   0.5726   1.0000
   0.750   0.2940   0.02326   0.01371  -0.0229   0.5589   1.0000
   1.000   0.3144   0.02392   0.01429  -0.0218   0.5457   1.0000
   1.250   0.3349   0.02471   0.01505  -0.0209   0.5324   1.0000
   1.500   0.3552   0.02548   0.01573  -0.0197   0.5213   1.0000
   1.750   0.3759   0.02633   0.01657  -0.0188   0.5102   1.0000
   2.000   0.3962   0.02732   0.01756  -0.0179   0.5000   1.0000
   2.250   0.4168   0.02819   0.01837  -0.0167   0.4898   1.0000
   2.500   0.4365   0.02936   0.01960  -0.0160   0.4797   1.0000
   2.750   0.4571   0.03026   0.02043  -0.0147   0.4702   1.0000
   3.000   0.4757   0.03161   0.02187  -0.0140   0.4594   1.0000
   3.250   0.4967   0.03238   0.02253  -0.0124   0.4499   1.0000
   3.500   0.5134   0.03401   0.02431  -0.0119   0.4385   1.0000
   3.750   0.5359   0.03459   0.02469  -0.0099   0.4296   1.0000
   4.000   0.5500   0.03662   0.02696  -0.0096   0.4175   1.0000
   4.250   0.5699   0.03773   0.02802  -0.0081   0.4084   1.0000
   4.500   0.5847   0.03963   0.03005  -0.0075   0.3972   1.0000
   4.750   0.6005   0.04148   0.03195  -0.0066   0.3879   1.0000
   5.000   0.6159   0.04333   0.03388  -0.0058   0.3778   1.0000
   5.250   0.6264   0.04602   0.03668  -0.0054   0.3687   1.0000
   5.500   0.6365   0.04875   0.03951  -0.0051   0.3600   1.0000
   5.750   0.6485   0.05125   0.04204  -0.0045   0.3523   1.0000
   6.000   0.6322   0.05754   0.04852  -0.0071   0.3466   1.0000
   6.250   0.6190   0.06328   0.05433  -0.0094   0.3443   1.0000
   6.500   0.5907   0.07052   0.06156  -0.0133   0.3488   1.0000
   6.750   0.5805   0.07542   0.06643  -0.0146   0.3521   1.0000
   7.000   0.5872   0.07969   0.07072  -0.0156   0.3562   1.0000
   7.250   0.4750   0.09269   0.08364  -0.0297   0.4724   1.0000
   7.500   0.4889   0.09635   0.08729  -0.0299   0.4653   1.0000
   7.750   0.4977   0.09867   0.08960  -0.0294   0.4515   1.0000
   8.000   0.4937   0.10104   0.09193  -0.0292   0.4415   1.0000
   8.250   0.5229   0.10510   0.09604  -0.0292   0.4311   1.0000
   8.500   0.5085   0.10674   0.09762  -0.0292   0.4221   1.0000
   8.750   0.5376   0.11097   0.10190  -0.0292   0.4123   1.0000
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