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

GOE 92 AIRFOIL (goe92-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 92 AIRFOIL (goe92-il)
Reynolds number: 500,000
Max Cl/Cd: 97.82 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe92-il-500000-n5.txt
Download as CSV file: xf-goe92-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 92 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.1466   0.08954   0.08725  -0.0360   0.8708   0.0078
  -8.750  -0.1424   0.08637   0.08407  -0.0367   0.8636   0.0081
  -8.500  -0.1381   0.08314   0.08082  -0.0374   0.8565   0.0086
  -8.250  -0.1337   0.07992   0.07759  -0.0386   0.8496   0.0091
  -8.000  -0.1306   0.07692   0.07457  -0.0403   0.8418   0.0093
  -7.750  -0.1273   0.07377   0.07141  -0.0418   0.8336   0.0094
  -7.500  -0.1235   0.07057   0.06819  -0.0431   0.8247   0.0094
  -7.250  -0.1191   0.06726   0.06487  -0.0445   0.8148   0.0094
  -7.000  -0.1143   0.06389   0.06149  -0.0461   0.8053   0.0094
  -6.750  -0.1086   0.06039   0.05797  -0.0483   0.7957   0.0094
  -6.500  -0.0987   0.05657   0.05412  -0.0515   0.7850   0.0094
  -6.250  -0.0869   0.05258   0.05008  -0.0549   0.7728   0.0094
  -6.000  -0.0734   0.04854   0.04598  -0.0583   0.7581   0.0094
  -5.750  -0.0584   0.04448   0.04182  -0.0616   0.7390   0.0094
  -5.250  -0.0562   0.05552   0.05265  -0.0694   0.7509   0.0086
  -5.000  -0.0308   0.05147   0.04842  -0.0737   0.7237   0.0076
  -4.750   0.0025   0.04493   0.04159  -0.0800   0.6913   0.0067
  -4.500   0.0286   0.04177   0.03813  -0.0825   0.6526   0.0068
  -4.250   0.0553   0.03899   0.03509  -0.0846   0.6244   0.0070
  -4.000   0.0834   0.03601   0.03187  -0.0865   0.6042   0.0072
  -3.750   0.1120   0.03335   0.02898  -0.0880   0.5878   0.0081
  -3.500   0.1427   0.02958   0.02489  -0.0894   0.5744   0.0087
  -3.250   0.1731   0.02566   0.02059  -0.0901   0.5621   0.0089
  -3.000   0.2035   0.02121   0.01563  -0.0904   0.5509   0.0096
  -2.750   0.2306   0.02013   0.01434  -0.0904   0.5374   0.0103
  -2.500   0.2583   0.01896   0.01292  -0.0904   0.5244   0.0118
  -2.250   0.2877   0.01540   0.00872  -0.0898   0.5134   0.0141
  -2.000   0.3147   0.01540   0.00860  -0.0898   0.4982   0.0150
  -1.750   0.3419   0.01506   0.00809  -0.0897   0.4824   0.0172
  -1.500   0.3701   0.01410   0.00672  -0.0892   0.4667   0.0199
  -1.250   0.3968   0.01356   0.00605  -0.0892   0.4498   0.0215
  -1.000   0.4237   0.01339   0.00576  -0.0891   0.4326   0.0230
  -0.750   0.4508   0.01303   0.00522  -0.0889   0.4156   0.0244
  -0.500   0.4778   0.01284   0.00487  -0.0888   0.3993   0.0265
  -0.250   0.5049   0.01267   0.00454  -0.0886   0.3851   0.0276
   0.000   0.5318   0.01258   0.00431  -0.0884   0.3715   0.0282
   0.500   0.5853   0.01202   0.00356  -0.0881   0.3502   0.0297
   0.750   0.6120   0.01184   0.00332  -0.0880   0.3425   0.0302
   1.000   0.6392   0.01172   0.00316  -0.0879   0.3369   0.0300
   1.250   0.6662   0.01165   0.00304  -0.0878   0.3315   0.0298
   1.500   0.6932   0.01160   0.00294  -0.0877   0.3269   0.0299
   1.750   0.7205   0.01152   0.00285  -0.0877   0.3233   0.0301
   2.000   0.7477   0.01149   0.00279  -0.0876   0.3193   0.0308
   2.250   0.7747   0.01151   0.00277  -0.0876   0.3154   0.0320
   2.500   0.8018   0.01156   0.00279  -0.0875   0.3119   0.0338
   2.750   0.8290   0.01160   0.00283  -0.0875   0.3080   0.0357
   3.000   0.8560   0.01169   0.00288  -0.0874   0.3037   0.0372
   3.250   0.8828   0.01181   0.00297  -0.0873   0.3003   0.0384
   3.500   0.9096   0.01192   0.00309  -0.0873   0.2977   0.0432
   3.750   0.9367   0.01199   0.00320  -0.0872   0.2952   0.0494
   4.250   0.9882   0.01061   0.00367  -0.0871   0.2894   1.0000
   4.500   1.0146   0.01079   0.00382  -0.0869   0.2864   1.0000
   4.750   1.0408   0.01098   0.00399  -0.0868   0.2832   1.0000
   5.000   1.0676   0.01112   0.00416  -0.0868   0.2805   1.0000
   5.250   1.0939   0.01128   0.00432  -0.0867   0.2727   1.0000
   5.500   1.1199   0.01148   0.00448  -0.0865   0.2625   1.0000
   5.750   1.1455   0.01171   0.00466  -0.0864   0.2485   1.0000
   6.000   1.1696   0.01211   0.00491  -0.0860   0.2189   1.0000
   6.250   1.1835   0.01370   0.00588  -0.0845   0.1266   1.0000
   6.500   1.2041   0.01449   0.00651  -0.0837   0.1005   1.0000
   6.750   1.2173   0.01612   0.00773  -0.0819   0.0195   1.0000
   7.000   1.2406   0.01654   0.00823  -0.0813   0.0131   1.0000
   7.250   1.2636   0.01698   0.00877  -0.0807   0.0113   1.0000
   7.500   1.2851   0.01755   0.00945  -0.0800   0.0092   1.0000
   7.750   1.3063   0.01813   0.01013  -0.0792   0.0080   1.0000
   8.000   1.3271   0.01872   0.01081  -0.0783   0.0072   1.0000
   8.250   1.3466   0.01939   0.01159  -0.0773   0.0065   1.0000
   8.500   1.3636   0.02024   0.01254  -0.0760   0.0059   1.0000
   8.750   1.3804   0.02105   0.01344  -0.0747   0.0053   1.0000
   9.000   1.3957   0.02193   0.01442  -0.0731   0.0049   1.0000
   9.250   1.4081   0.02289   0.01549  -0.0712   0.0046   1.0000
   9.500   1.4171   0.02392   0.01662  -0.0688   0.0043   1.0000
   9.750   1.4243   0.02510   0.01789  -0.0664   0.0041   1.0000
  10.000   1.4285   0.02654   0.01946  -0.0639   0.0039   1.0000
  10.250   1.4275   0.02849   0.02154  -0.0614   0.0038   1.0000
  10.500   1.4232   0.03089   0.02408  -0.0591   0.0037   1.0000
  10.750   1.4219   0.03324   0.02657  -0.0575   0.0036   1.0000
  11.000   1.4195   0.03588   0.02935  -0.0562   0.0036   1.0000
  11.250   1.4167   0.03877   0.03238  -0.0554   0.0034   1.0000
  11.500   1.4127   0.04203   0.03577  -0.0549   0.0033   1.0000
  11.750   1.4084   0.04559   0.03947  -0.0550   0.0031   1.0000
  12.000   1.4036   0.04944   0.04344  -0.0554   0.0030   1.0000
  12.250   1.3978   0.05355   0.04767  -0.0561   0.0029   1.0000
  12.500   1.3912   0.05784   0.05209  -0.0568   0.0028   1.0000
  12.750   1.3833   0.06230   0.05666  -0.0575   0.0027   1.0000
  13.000   1.3750   0.06680   0.06127  -0.0583   0.0027   1.0000
  13.250   1.3681   0.07120   0.06577  -0.0591   0.0026   1.0000
  13.500   1.3608   0.07562   0.07028  -0.0598   0.0026   1.0000
  13.750   1.3546   0.07992   0.07467  -0.0606   0.0026   1.0000
  14.000   1.3492   0.08414   0.07898  -0.0614   0.0025   1.0000
  14.250   1.3444   0.08829   0.08321  -0.0622   0.0025   1.0000
<< Back to GOE 92 AIRFOIL (goe92-il)

Polar data table (+)

Polar graphs


<< Back to GOE 92 AIRFOIL (goe92-il)