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

GOE 427 AIRFOIL (goe427-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 427 AIRFOIL (goe427-il)
Reynolds number: 500,000
Max Cl/Cd: 94.74 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe427-il-500000-n5.txt
Download as CSV file: xf-goe427-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 427 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.3272   0.11441   0.11214  -0.0248   1.0000   0.0053
  -9.250  -0.3254   0.11187   0.10962  -0.0244   1.0000   0.0053
  -9.000  -0.3249   0.10915   0.10690  -0.0238   1.0000   0.0052
  -8.750  -0.3189   0.10486   0.10262  -0.0247   0.9993   0.0048
  -8.500  -0.3065   0.10087   0.09864  -0.0278   0.9976   0.0046
  -8.250  -0.2918   0.09724   0.09502  -0.0314   0.9953   0.0043
  -8.000  -0.2783   0.09372   0.09149  -0.0346   0.9925   0.0041
  -7.750  -0.2639   0.08992   0.08771  -0.0384   0.9892   0.0040
  -7.500  -0.2475   0.08627   0.08407  -0.0427   0.9861   0.0039
  -7.250  -0.2332   0.08258   0.08039  -0.0465   0.9807   0.0040
  -7.000  -0.2105   0.07831   0.07611  -0.0527   0.9776   0.0041
  -6.750  -0.1907   0.07420   0.07199  -0.0579   0.9721   0.0043
  -6.500  -0.1588   0.06920   0.06698  -0.0668   0.9682   0.0051
  -6.250  -0.1335   0.06526   0.06303  -0.0728   0.9661   0.0063
  -6.000  -0.1153   0.06455   0.06232  -0.0737   0.9605   0.0101
  -5.750  -0.0841   0.06025   0.05799  -0.0806   0.9563   0.0095
  -5.500  -0.0519   0.05569   0.05337  -0.0877   0.9514   0.0084
  -5.250  -0.0216   0.05149   0.04912  -0.0935   0.9444   0.0079
  -5.000   0.0145   0.04707   0.04461  -0.1003   0.9394   0.0078
  -4.750   0.0454   0.04333   0.04080  -0.1049   0.9308   0.0082
  -4.500   0.0790   0.03968   0.03703  -0.1095   0.9234   0.0091
  -4.250   0.1129   0.03625   0.03343  -0.1135   0.9154   0.0106
  -4.000   0.1458   0.03294   0.02996  -0.1166   0.9060   0.0117
  -3.750   0.1790   0.02927   0.02609  -0.1196   0.8962   0.0115
  -3.500   0.2122   0.02548   0.02204  -0.1219   0.8856   0.0114
  -3.250   0.2441   0.02150   0.01774  -0.1234   0.8748   0.0117
  -3.000   0.2744   0.01794   0.01381  -0.1239   0.8644   0.0128
  -2.750   0.3021   0.01521   0.01063  -0.1243   0.8534   0.0154
  -2.500   0.3287   0.01476   0.01005  -0.1243   0.8416   0.0176
  -2.250   0.3558   0.01363   0.00863  -0.1241   0.8297   0.0193
  -2.000   0.3828   0.01279   0.00745  -0.1238   0.8169   0.0235
  -1.750   0.4092   0.01207   0.00647  -0.1233   0.8008   0.0231
  -1.500   0.4352   0.01128   0.00544  -0.1228   0.7835   0.0213
  -1.250   0.4610   0.01032   0.00420  -0.1222   0.7675   0.0204
  -1.000   0.4871   0.00963   0.00332  -0.1216   0.7543   0.0196
  -0.750   0.5130   0.00912   0.00266  -0.1211   0.7416   0.0189
  -0.500   0.5388   0.00874   0.00217  -0.1205   0.7287   0.0183
  -0.250   0.5646   0.00849   0.00182  -0.1199   0.7149   0.0178
   0.000   0.5903   0.00833   0.00155  -0.1194   0.7004   0.0174
   0.250   0.6158   0.00828   0.00138  -0.1188   0.6833   0.0172
   0.500   0.6390   0.00838   0.00127  -0.1177   0.6394   0.0171
   0.750   0.6597   0.00870   0.00122  -0.1162   0.5731   0.0173
   1.000   0.6815   0.00905   0.00125  -0.1150   0.5239   0.0178
   1.250   0.7054   0.00926   0.00129  -0.1142   0.4937   0.0198
   1.500   0.7304   0.00933   0.00136  -0.1137   0.4747   0.0531
   1.750   0.7550   0.00941   0.00150  -0.1131   0.4492   0.1106
   2.000   0.7790   0.00958   0.00163  -0.1125   0.4134   0.1500
   2.250   0.8062   0.00851   0.00200  -0.1132   0.3573   1.0000
   2.500   0.8274   0.00909   0.00227  -0.1122   0.2964   1.0000
   2.750   0.8431   0.01032   0.00279  -0.1103   0.1528   1.0000
   3.000   0.8630   0.01114   0.00322  -0.1090   0.0710   1.0000
   3.250   0.8860   0.01160   0.00352  -0.1082   0.0356   1.0000
   3.500   0.9106   0.01185   0.00375  -0.1077   0.0332   1.0000
   3.750   0.9353   0.01210   0.00401  -0.1071   0.0317   1.0000
   4.000   0.9598   0.01236   0.00434  -0.1066   0.0308   1.0000
   4.250   0.9842   0.01263   0.00465  -0.1060   0.0303   1.0000
   4.500   1.0083   0.01292   0.00498  -0.1054   0.0299   1.0000
   4.750   1.0321   0.01324   0.00534  -0.1047   0.0292   1.0000
   5.000   1.0563   0.01345   0.00546  -0.1043   0.0209   1.0000
   5.250   1.0806   0.01366   0.00557  -0.1039   0.0082   1.0000
   5.500   1.1032   0.01412   0.00599  -0.1030   0.0042   1.0000
   5.750   1.1260   0.01452   0.00648  -0.1021   0.0037   1.0000
   6.000   1.1484   0.01497   0.00701  -0.1012   0.0035   1.0000
   6.250   1.1701   0.01548   0.00772  -0.1001   0.0034   1.0000
   6.500   1.1910   0.01607   0.00843  -0.0989   0.0034   1.0000
   6.750   1.2109   0.01674   0.00922  -0.0975   0.0033   1.0000
   7.000   1.2294   0.01752   0.01014  -0.0959   0.0033   1.0000
   7.250   1.2464   0.01842   0.01118  -0.0941   0.0033   1.0000
   7.500   1.2629   0.01934   0.01222  -0.0923   0.0031   1.0000
   7.750   1.2790   0.02025   0.01324  -0.0904   0.0029   1.0000
   8.000   1.2948   0.02117   0.01426  -0.0886   0.0025   1.0000
   8.250   1.3089   0.02225   0.01547  -0.0864   0.0024   1.0000
   8.500   1.3211   0.02354   0.01690  -0.0840   0.0023   1.0000
   8.750   1.3321   0.02484   0.01834  -0.0814   0.0021   1.0000
   9.000   1.3422   0.02628   0.01995  -0.0787   0.0020   1.0000
   9.250   1.3519   0.02777   0.02160  -0.0761   0.0019   1.0000
   9.500   1.3608   0.02942   0.02343  -0.0734   0.0018   1.0000
   9.750   1.3684   0.03120   0.02540  -0.0708   0.0017   1.0000
  10.000   1.3742   0.03310   0.02762  -0.0680   0.0016   1.0000
  10.250   1.3780   0.03528   0.03004  -0.0652   0.0015   1.0000
  10.500   1.3791   0.03757   0.03256  -0.0623   0.0015   1.0000
  10.750   1.3765   0.04014   0.03536  -0.0593   0.0014   1.0000
  11.000   1.3721   0.04303   0.03850  -0.0565   0.0014   1.0000
  11.250   1.3652   0.04628   0.04202  -0.0540   0.0014   1.0000
  11.500   1.3557   0.04980   0.04577  -0.0519   0.0014   1.0000
  11.750   1.3430   0.05375   0.04995  -0.0504   0.0013   1.0000
  12.000   1.3288   0.05820   0.05463  -0.0496   0.0013   1.0000
  12.250   1.3142   0.06311   0.05977  -0.0496   0.0013   1.0000
  12.500   1.2980   0.06863   0.06551  -0.0505   0.0014   1.0000
  12.750   1.2804   0.07481   0.07188  -0.0523   0.0013   1.0000
  13.000   1.2624   0.08172   0.07900  -0.0551   0.0014   1.0000
  13.250   1.2432   0.08949   0.08695  -0.0590   0.0014   1.0000
  13.500   1.2240   0.09798   0.09560  -0.0637   0.0014   1.0000
  13.750   1.2046   0.10726   0.10504  -0.0691   0.0014   1.0000
  14.000   1.1819   0.11854   0.11648  -0.0758   0.0014   1.0000
  14.250   1.1561   0.13255   0.13063  -0.0840   0.0015   1.0000
<< Back to GOE 427 AIRFOIL (goe427-il)

Polar data table (+)

Polar graphs


<< Back to GOE 427 AIRFOIL (goe427-il)