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

GOE 492 AIRFOIL (goe492-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 492 AIRFOIL (goe492-il)
Reynolds number: 200,000
Max Cl/Cd: 78.61 at α=3°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe492-il-200000-n5.txt
Download as CSV file: xf-goe492-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 492 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4003   0.08680   0.08344  -0.0309   1.0000   0.0124
  -7.500  -0.4114   0.08406   0.08080  -0.0297   1.0000   0.0117
  -7.250  -0.4266   0.08064   0.07749  -0.0292   1.0000   0.0106
  -7.000  -0.4309   0.07788   0.07479  -0.0291   1.0000   0.0103
  -6.750  -0.4206   0.07466   0.07158  -0.0320   0.9985   0.0100
  -6.500  -0.3971   0.06866   0.06552  -0.0403   0.9935   0.0095
  -6.250  -0.3681   0.06158   0.05837  -0.0509   0.9881   0.0094
  -6.000  -0.3361   0.05458   0.05126  -0.0608   0.9823   0.0097
  -5.750  -0.2985   0.04562   0.04209  -0.0720   0.9779   0.0100
  -5.500  -0.2672   0.03548   0.03147  -0.0806   0.9712   0.0104
  -5.250  -0.2331   0.02778   0.02302  -0.0862   0.9677   0.0105
  -5.000  -0.2054   0.02293   0.01741  -0.0879   0.9630   0.0107
  -4.750  -0.1755   0.01944   0.01323  -0.0891   0.9595   0.0114
  -4.500  -0.1433   0.01728   0.01058  -0.0902   0.9570   0.0126
  -4.250  -0.1096   0.01617   0.00916  -0.0916   0.9551   0.0159
  -4.000  -0.0818   0.01480   0.00750  -0.0915   0.9506   0.0183
  -3.750  -0.0514   0.01400   0.00661  -0.0921   0.9469   0.0250
  -3.500  -0.0183   0.01354   0.00602  -0.0932   0.9440   0.0369
  -3.250   0.0156   0.01327   0.00560  -0.0946   0.9416   0.0484
  -3.000   0.0433   0.01282   0.00511  -0.0948   0.9365   0.0570
  -2.750   0.0738   0.01242   0.00459  -0.0954   0.9321   0.0613
  -2.500   0.1066   0.01201   0.00405  -0.0964   0.9289   0.0654
  -2.250   0.1360   0.01169   0.00369  -0.0968   0.9239   0.0721
  -2.000   0.1656   0.01139   0.00338  -0.0972   0.9186   0.0859
  -1.750   0.1972   0.01080   0.00312  -0.0983   0.9148   0.1725
  -1.500   0.2245   0.01061   0.00309  -0.0982   0.9081   0.2520
  -1.250   0.2551   0.01048   0.00298  -0.0988   0.9027   0.2893
  -1.000   0.2842   0.01033   0.00286  -0.0990   0.8963   0.3142
  -0.750   0.3127   0.01011   0.00273  -0.0992   0.8892   0.3437
  -0.500   0.3405   0.00983   0.00263  -0.0992   0.8816   0.3854
  -0.250   0.3641   0.00898   0.00266  -0.0984   0.8737   0.6494
   0.000   0.4057   0.00827   0.00257  -0.1008   0.8667   1.0000
   0.250   0.4347   0.00826   0.00249  -0.1009   0.8580   1.0000
   0.500   0.4617   0.00828   0.00245  -0.1006   0.8469   1.0000
   0.750   0.4886   0.00830   0.00242  -0.1002   0.8345   1.0000
   1.000   0.5155   0.00832   0.00242  -0.0998   0.8210   1.0000
   1.250   0.5423   0.00835   0.00241  -0.0994   0.8061   1.0000
   1.500   0.5691   0.00840   0.00242  -0.0990   0.7901   1.0000
   1.750   0.5956   0.00847   0.00250  -0.0986   0.7745   1.0000
   2.000   0.6221   0.00855   0.00259  -0.0982   0.7588   1.0000
   2.250   0.6485   0.00865   0.00270  -0.0978   0.7425   1.0000
   2.500   0.6748   0.00876   0.00283  -0.0973   0.7249   1.0000
   2.750   0.7002   0.00892   0.00299  -0.0966   0.6990   1.0000
   3.000   0.7232   0.00920   0.00309  -0.0952   0.6527   1.0000
   3.250   0.7412   0.00979   0.00327  -0.0929   0.5669   1.0000
   3.500   0.7511   0.01102   0.00361  -0.0894   0.3860   1.0000
   3.750   0.7517   0.01396   0.00482  -0.0856   0.0478   1.0000
   4.000   0.7726   0.01480   0.00561  -0.0842   0.0220   1.0000
   4.250   0.7936   0.01565   0.00659  -0.0828   0.0173   1.0000
   4.500   0.8147   0.01645   0.00755  -0.0815   0.0153   1.0000
   4.750   0.8352   0.01731   0.00851  -0.0802   0.0126   1.0000
   5.000   0.8539   0.01847   0.00977  -0.0785   0.0116   1.0000
   5.250   0.8724   0.01989   0.01130  -0.0767   0.0108   1.0000
   5.500   0.8925   0.02172   0.01321  -0.0752   0.0102   1.0000
   5.750   0.9155   0.02438   0.01606  -0.0741   0.0091   1.0000
   6.000   0.9400   0.02567   0.01757  -0.0731   0.0081   1.0000
   6.250   0.9645   0.02803   0.02022  -0.0721   0.0079   1.0000
   6.500   0.9874   0.03072   0.02326  -0.0707   0.0078   1.0000
   6.750   1.0076   0.03380   0.02674  -0.0690   0.0079   1.0000
   7.000   1.0247   0.03717   0.03051  -0.0669   0.0080   1.0000
   7.250   1.0384   0.04081   0.03462  -0.0646   0.0082   1.0000
   7.500   1.0480   0.04477   0.03897  -0.0620   0.0085   1.0000
   7.750   1.0530   0.04910   0.04366  -0.0593   0.0088   1.0000
<< Back to GOE 492 AIRFOIL (goe492-il)

Polar data table (+)

Polar graphs


<< Back to GOE 492 AIRFOIL (goe492-il)