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GOE 711 AIRFOIL (goe711-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 711 AIRFOIL (goe711-il)
Reynolds number: 50,000
Max Cl/Cd: 20.28 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe711-il-50000.txt
Download as CSV file: xf-goe711-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 711 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.3516   0.12398   0.11854  -0.0173   1.0000   0.1725
  -7.250  -0.3806   0.12400   0.11871  -0.0151   1.0000   0.1743
  -7.000  -0.4109   0.12393   0.11877  -0.0145   1.0000   0.1752
  -6.750  -0.3854   0.11859   0.11344  -0.0105   1.0000   0.1835
  -6.500  -0.4032   0.11741   0.11237  -0.0088   1.0000   0.1883
  -6.250  -0.4294   0.11672   0.11178  -0.0107   1.0000   0.1917
  -6.000  -0.4239   0.11307   0.10821  -0.0069   1.0000   0.1977
  -5.750  -0.4323   0.11127   0.10647  -0.0064   1.0000   0.2060
  -5.500  -0.4446   0.10888   0.10414  -0.0084   1.0000   0.2115
  -5.250  -0.4420   0.10649   0.10182  -0.0048   1.0000   0.2214
  -5.000  -0.4482   0.10384   0.09922  -0.0061   1.0000   0.2298
  -4.750  -0.4516   0.10203   0.09738  -0.0103   1.0000   0.2440
  -4.500  -0.4486   0.09913   0.09459  -0.0037   1.0000   0.2519
  -4.250  -0.4488   0.09663   0.09213  -0.0041   1.0000   0.2658
  -4.000  -0.4161   0.09290   0.08836  -0.0097   0.9885   0.2984
  -3.750  -0.3849   0.08978   0.08522  -0.0117   0.9748   0.3397
  -3.250   0.0684   0.07308   0.06797  -0.0326   0.9840   1.0000
  -3.000   0.1133   0.06943   0.06427  -0.0424   0.9666   1.0000
  -2.750   0.1325   0.06711   0.06195  -0.0460   0.9489   0.9939
  -2.500   0.0720   0.06774   0.06273  -0.0322   0.9284   0.9450
  -2.250   0.0106   0.06798   0.06311  -0.0204   0.9112   0.8933
  -2.000  -0.0479   0.06777   0.06305  -0.0099   0.8931   0.8572
  -1.750   0.0359   0.05857   0.05106  -0.0887   0.8639   0.2480
  -1.500   0.1029   0.05590   0.04745  -0.0955   0.8524   0.2051
  -1.250   0.1475   0.05426   0.04522  -0.0983   0.8385   0.1885
  -1.000   0.1909   0.05318   0.04344  -0.1004   0.8245   0.1752
  -0.750   0.2292   0.05176   0.04182  -0.1021   0.8108   0.1707
  -0.500   0.2673   0.05081   0.04057  -0.1035   0.7972   0.1694
  -0.250   0.3064   0.04994   0.03940  -0.1048   0.7841   0.1712
   0.000   0.3632   0.04817   0.03732  -0.1078   0.7754   0.1723
   0.250   0.3945   0.04780   0.03674  -0.1077   0.7612   0.1745
   0.500   0.4250   0.04734   0.03624  -0.1076   0.7473   0.1808
   0.750   0.4559   0.04703   0.03583  -0.1076   0.7339   0.1921
   1.000   0.5146   0.04479   0.03365  -0.1108   0.7274   0.2169
   1.250   0.5455   0.04386   0.03348  -0.1114   0.7144   0.3249
   1.500   0.5612   0.04288   0.03368  -0.1077   0.7009   1.0000
   1.750   0.6260   0.04097   0.03099  -0.1109   0.6958   1.0000
   2.000   0.6482   0.04159   0.03135  -0.1099   0.6822   1.0000
   2.250   0.6644   0.04264   0.03221  -0.1084   0.6684   1.0000
   2.500   0.6825   0.04365   0.03305  -0.1072   0.6555   1.0000
   2.750   0.7507   0.04152   0.03060  -0.1115   0.6503   1.0000
   3.000   0.7538   0.04348   0.03249  -0.1087   0.6364   1.0000
   3.250   0.7573   0.04556   0.03449  -0.1060   0.6237   1.0000
   3.500   0.8266   0.04333   0.03202  -0.1104   0.6186   1.0000
   3.750   0.8073   0.04697   0.03568  -0.1054   0.6048   1.0000
   4.000   0.8902   0.04390   0.03235  -0.1113   0.6013   1.0000
   4.250   0.8495   0.04907   0.03762  -0.1042   0.5872   1.0000
   4.500   0.7965   0.05651   0.04516  -0.0981   0.5728   1.0000
   4.750   0.8279   0.05679   0.04534  -0.0983   0.5669   1.0000
   5.250   0.7381   0.07294   0.06167  -0.0927   0.5461   1.0000
   5.500   0.7492   0.07547   0.06417  -0.0924   0.5402   1.0000
   5.750   0.7689   0.07740   0.06605  -0.0923   0.5358   1.0000
   6.000   0.7254   0.08551   0.07425  -0.0910   0.5328   1.0000
   6.250   0.7106   0.09071   0.07949  -0.0905   0.5310   1.0000
   6.500   0.7032   0.09547   0.08428  -0.0905   0.5317   1.0000
   6.750   0.7076   0.09968   0.08849  -0.0912   0.5348   1.0000
   7.000   0.7232   0.10319   0.09200  -0.0922   0.5369   1.0000
   7.250   0.5875   0.11866   0.10792  -0.0938   0.6418   1.0000
   7.500   0.6086   0.12151   0.11072  -0.0947   0.6352   1.0000
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