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GOE 113 AIRFOIL (goe113-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 113 AIRFOIL (goe113-il)
Reynolds number: 500,000
Max Cl/Cd: 77.97 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe113-il-500000-n5.txt
Download as CSV file: xf-goe113-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 113 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
  ------ -------- --------- --------- -------- -------- --------
  -7.500  -0.4274   0.08881   0.08669  -0.0133   1.0000   0.0071
  -7.250  -0.4224   0.08513   0.08304  -0.0160   1.0000   0.0071
  -7.000  -0.4198   0.08122   0.07915  -0.0168   1.0000   0.0065
  -6.750  -0.4135   0.07714   0.07508  -0.0201   1.0000   0.0056
  -6.500  -0.4056   0.07240   0.07035  -0.0240   1.0000   0.0051
  -6.250  -0.3958   0.06726   0.06520  -0.0281   1.0000   0.0048
  -6.000  -0.3839   0.06334   0.06126  -0.0308   1.0000   0.0049
  -5.750  -0.3654   0.05996   0.05783  -0.0339   0.9991   0.0055
  -5.500  -0.3311   0.05419   0.05194  -0.0413   0.9957   0.0065
  -5.250  -0.2990   0.04861   0.04622  -0.0472   0.9917   0.0062
  -5.000  -0.2643   0.04286   0.04026  -0.0527   0.9881   0.0058
  -4.750  -0.2303   0.03705   0.03419  -0.0568   0.9833   0.0055
  -4.500  -0.1948   0.03108   0.02786  -0.0601   0.9792   0.0053
  -4.250  -0.1623   0.02550   0.02184  -0.0617   0.9725   0.0054
  -4.000  -0.1293   0.01977   0.01550  -0.0626   0.9657   0.0059
  -3.750  -0.0972   0.01317   0.00793  -0.0628   0.9590   0.0071
  -3.500  -0.0662   0.01163   0.00610  -0.0633   0.9489   0.0089
  -3.250  -0.0352   0.01109   0.00544  -0.0639   0.9346   0.0104
  -3.000  -0.0050   0.01039   0.00459  -0.0642   0.9148   0.0125
  -2.750   0.0238   0.01080   0.00494  -0.0643   0.8854   0.0153
  -2.500   0.0504   0.01057   0.00448  -0.0636   0.8492   0.0199
  -2.250   0.0764   0.01092   0.00458  -0.0630   0.8097   0.0227
  -2.000   0.1019   0.01047   0.00390  -0.0625   0.7743   0.0249
  -1.500   0.1547   0.01016   0.00331  -0.0619   0.7186   0.0294
  -1.250   0.1813   0.01002   0.00299  -0.0617   0.6899   0.0306
  -1.000   0.2081   0.00980   0.00261  -0.0614   0.6639   0.0305
  -0.750   0.2350   0.00962   0.00228  -0.0612   0.6383   0.0302
  -0.500   0.2622   0.00947   0.00200  -0.0611   0.6165   0.0298
  -0.250   0.2896   0.00935   0.00177  -0.0611   0.5982   0.0296
   0.000   0.3171   0.00925   0.00158  -0.0610   0.5821   0.0295
   0.250   0.3447   0.00919   0.00143  -0.0610   0.5667   0.0295
   0.500   0.3723   0.00916   0.00131  -0.0610   0.5508   0.0297
   0.750   0.3999   0.00915   0.00122  -0.0610   0.5364   0.0302
   1.000   0.4275   0.00917   0.00116  -0.0609   0.5194   0.0314
   1.250   0.4549   0.00922   0.00113  -0.0609   0.5020   0.0334
   1.500   0.4824   0.00889   0.00117  -0.0611   0.4849   0.2220
   2.000   0.5315   0.00721   0.00133  -0.0601   0.4467   1.0000
   2.500   0.5857   0.00756   0.00150  -0.0600   0.3997   1.0000
   2.750   0.6121   0.00785   0.00161  -0.0599   0.3603   1.0000
   3.000   0.6373   0.00837   0.00182  -0.0597   0.2896   1.0000
   3.250   0.6625   0.00890   0.00209  -0.0595   0.2366   1.0000
   3.500   0.6883   0.00931   0.00233  -0.0593   0.2003   1.0000
   3.750   0.7147   0.00960   0.00253  -0.0592   0.1779   1.0000
   4.000   0.7400   0.01009   0.00278  -0.0590   0.1321   1.0000
   4.250   0.7654   0.01057   0.00311  -0.0588   0.1004   1.0000
   4.500   0.7881   0.01154   0.00371  -0.0583   0.0184   1.0000
   4.750   0.8142   0.01190   0.00414  -0.0580   0.0114   1.0000
   5.000   0.8403   0.01224   0.00458  -0.0577   0.0098   1.0000
   5.250   0.8659   0.01268   0.00514  -0.0573   0.0079   1.0000
   5.500   0.8902   0.01336   0.00595  -0.0568   0.0066   1.0000
   5.750   0.9149   0.01394   0.00663  -0.0563   0.0059   1.0000
   6.000   0.9387   0.01467   0.00745  -0.0556   0.0054   1.0000
   6.250   0.9616   0.01551   0.00839  -0.0549   0.0050   1.0000
   6.500   0.9845   0.01626   0.00922  -0.0543   0.0045   1.0000
   6.750   1.0058   0.01728   0.01033  -0.0534   0.0039   1.0000
   7.000   1.0267   0.01837   0.01155  -0.0523   0.0036   1.0000
   7.250   1.0466   0.01965   0.01295  -0.0510   0.0033   1.0000
   7.500   1.0660   0.02106   0.01448  -0.0497   0.0031   1.0000
   7.750   1.0853   0.02262   0.01617  -0.0483   0.0029   1.0000
   8.000   1.1043   0.02437   0.01806  -0.0470   0.0028   1.0000
   8.250   1.1228   0.02637   0.02025  -0.0456   0.0028   1.0000
   8.500   1.1404   0.02870   0.02282  -0.0440   0.0028   1.0000
   8.750   1.1561   0.03141   0.02586  -0.0422   0.0028   1.0000
   9.000   1.1689   0.03448   0.02927  -0.0401   0.0029   1.0000
   9.250   1.1783   0.03777   0.03290  -0.0378   0.0031   1.0000
   9.500   1.1843   0.04095   0.03640  -0.0354   0.0031   1.0000
   9.750   1.1879   0.04376   0.03948  -0.0331   0.0030   1.0000
  10.000   1.1904   0.04600   0.04191  -0.0311   0.0028   1.0000
  10.250   1.1886   0.04822   0.04430  -0.0287   0.0027   1.0000
  10.500   1.1807   0.05045   0.04667  -0.0258   0.0026   1.0000
  10.750   1.1692   0.05327   0.04965  -0.0236   0.0025   1.0000
  11.000   1.1550   0.05683   0.05339  -0.0227   0.0025   1.0000
  11.250   1.1405   0.06117   0.05792  -0.0232   0.0025   1.0000
  11.500   1.1244   0.06625   0.06317  -0.0252   0.0025   1.0000
  11.750   1.1096   0.07200   0.06909  -0.0283   0.0025   1.0000
  12.000   1.0946   0.07848   0.07573  -0.0322   0.0025   1.0000
  12.250   1.0786   0.08576   0.08317  -0.0369   0.0025   1.0000
  12.500   1.0578   0.09514   0.09274  -0.0431   0.0026   1.0000
  12.750   1.0242   0.10982   0.10765  -0.0526   0.0028   1.0000
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