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

GOE 499 AIRFOIL (goe499-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 499 AIRFOIL (goe499-il)
Reynolds number: 100,000
Max Cl/Cd: 69.63 at α=4.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe499-il-100000-n5.txt
Download as CSV file: xf-goe499-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 499 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.3447   0.13401   0.12915  -0.0386   1.0000   0.0335
 -10.000  -0.3479   0.13186   0.12706  -0.0375   1.0000   0.0335
  -9.750  -0.3507   0.12974   0.12498  -0.0364   1.0000   0.0335
  -9.500  -0.3384   0.12235   0.11760  -0.0338   1.0000   0.0347
  -9.250  -0.3370   0.11976   0.11504  -0.0319   1.0000   0.0357
  -9.000  -0.3378   0.11782   0.11314  -0.0301   1.0000   0.0371
  -8.750  -0.3305   0.11514   0.11046  -0.0312   0.9983   0.0401
  -8.500  -0.3186   0.11196   0.10729  -0.0353   0.9948   0.0438
  -8.250  -0.3147   0.11068   0.10605  -0.0423   0.9885   0.0458
  -8.000  -0.3069   0.10793   0.10331  -0.0466   0.9832   0.0461
  -7.750  -0.2972   0.10463   0.10002  -0.0507   0.9778   0.0463
  -7.500  -0.2866   0.10097   0.09637  -0.0546   0.9721   0.0463
  -7.250  -0.2734   0.09523   0.09062  -0.0508   0.9722   0.0483
  -7.000  -0.2583   0.09243   0.08779  -0.0512   0.9689   0.0526
  -6.750  -0.2476   0.09014   0.08554  -0.0601   0.9599   0.0591
  -6.250  -0.2156   0.08167   0.07702  -0.0671   0.9528   0.0618
  -6.000  -0.2075   0.07898   0.07433  -0.0670   0.9468   0.0632
  -5.750  -0.1898   0.07598   0.07130  -0.0699   0.9426   0.0675
  -5.250  -0.1466   0.06598   0.06117  -0.0836   0.9325   0.0541
  -5.000  -0.1070   0.05829   0.05333  -0.0972   0.9283   0.0494
  -4.750  -0.0734   0.05483   0.04975  -0.1034   0.9260   0.0532
  -4.500  -0.0293   0.04843   0.04312  -0.1151   0.9229   0.0529
  -4.250   0.0332   0.03847   0.03258  -0.1322   0.9206   0.0510
  -4.000   0.0841   0.03273   0.02614  -0.1420   0.9189   0.0524
  -3.750   0.1264   0.02946   0.02225  -0.1474   0.9171   0.0548
  -3.500   0.1668   0.02674   0.01888  -0.1514   0.9156   0.0540
  -3.250   0.2044   0.02495   0.01657  -0.1541   0.9139   0.0535
  -3.000   0.2408   0.02366   0.01489  -0.1563   0.9123   0.0534
  -2.750   0.2747   0.02272   0.01370  -0.1577   0.9098   0.0535
  -2.500   0.3006   0.02209   0.01290  -0.1575   0.9039   0.0539
  -2.250   0.3383   0.02130   0.01193  -0.1593   0.9000   0.0547
  -2.000   0.3813   0.02050   0.01097  -0.1621   0.8971   0.0563
  -1.750   0.4080   0.02008   0.01045  -0.1618   0.8887   0.0582
  -1.500   0.4478   0.01944   0.00972  -0.1639   0.8849   0.0620
  -1.250   0.4787   0.01906   0.00934  -0.1644   0.8795   0.0694
  -1.000   0.5083   0.01875   0.00912  -0.1648   0.8743   0.0907
  -0.750   0.5430   0.01832   0.00900  -0.1662   0.8712   0.1793
  -0.500   0.5728   0.01822   0.00900  -0.1666   0.8667   0.2370
  -0.250   0.5996   0.01818   0.00901  -0.1664   0.8608   0.2757
   0.000   0.6336   0.01794   0.00881  -0.1675   0.8573   0.3064
   0.250   0.6610   0.01777   0.00890  -0.1674   0.8515   0.3731
   0.500   0.6898   0.01750   0.00899  -0.1675   0.8460   0.4920
   0.750   0.7229   0.01702   0.00884  -0.1682   0.8425   0.6174
   1.000   0.7405   0.01651   0.00884  -0.1657   0.8340   1.0000
   1.250   0.7739   0.01642   0.00865  -0.1665   0.8294   1.0000
   1.500   0.7987   0.01656   0.00877  -0.1658   0.8213   1.0000
   1.750   0.8308   0.01647   0.00865  -0.1662   0.8157   1.0000
   2.000   0.8558   0.01657   0.00876  -0.1655   0.8068   1.0000
   2.250   0.8886   0.01643   0.00863  -0.1660   0.8006   1.0000
   2.500   0.9125   0.01656   0.00882  -0.1651   0.7902   1.0000
   2.750   0.9408   0.01655   0.00887  -0.1649   0.7815   1.0000
   3.000   0.9702   0.01649   0.00888  -0.1648   0.7723   1.0000
   3.250   0.9954   0.01652   0.00902  -0.1639   0.7594   1.0000
   3.500   1.0217   0.01651   0.00911  -0.1632   0.7455   1.0000
   3.750   1.0490   0.01644   0.00915  -0.1626   0.7295   1.0000
   4.000   1.0727   0.01612   0.00882  -0.1607   0.6891   1.0000
   4.250   1.1008   0.01581   0.00822  -0.1591   0.6154   1.0000
   4.500   1.1234   0.01626   0.00819  -0.1572   0.5388   1.0000
   4.750   1.1341   0.01751   0.00872  -0.1537   0.4372   1.0000
   5.000   1.1339   0.01976   0.00968  -0.1491   0.2322   1.0000
   5.250   1.1342   0.02298   0.01155  -0.1453   0.0411   1.0000
   5.500   1.1508   0.02411   0.01275  -0.1433   0.0318   1.0000
   5.750   1.1672   0.02519   0.01398  -0.1413   0.0287   1.0000
   6.000   1.1827   0.02626   0.01526  -0.1391   0.0269   1.0000
   6.250   1.1959   0.02746   0.01664  -0.1367   0.0252   1.0000
   6.500   1.2058   0.02876   0.01808  -0.1338   0.0231   1.0000
   6.750   1.2114   0.03037   0.01980  -0.1304   0.0214   1.0000
   7.000   1.2161   0.03234   0.02185  -0.1270   0.0206   1.0000
   7.250   1.2283   0.03425   0.02385  -0.1246   0.0201   1.0000
   7.500   1.2486   0.03605   0.02573  -0.1233   0.0197   1.0000
   7.750   1.2763   0.03810   0.02790  -0.1230   0.0193   1.0000
   8.000   1.3049   0.04012   0.03012  -0.1228   0.0182   1.0000
   8.250   1.3338   0.04240   0.03266  -0.1228   0.0170   1.0000
   8.500   1.3620   0.04525   0.03581  -0.1225   0.0166   1.0000
   8.750   1.3850   0.04834   0.03925  -0.1215   0.0165   1.0000
   9.000   1.4025   0.05159   0.04289  -0.1198   0.0165   1.0000
   9.250   1.4143   0.05498   0.04675  -0.1175   0.0165   1.0000
   9.500   1.4209   0.05846   0.05064  -0.1147   0.0166   1.0000
   9.750   1.4225   0.06197   0.05454  -0.1116   0.0168   1.0000
  10.000   1.4183   0.06537   0.05828  -0.1079   0.0170   1.0000
  10.250   1.4099   0.06880   0.06204  -0.1040   0.0171   1.0000
  10.500   1.3990   0.07231   0.06584  -0.1005   0.0172   1.0000
  10.750   1.3856   0.07603   0.06986  -0.0973   0.0173   1.0000
  11.000   1.3704   0.08004   0.07414  -0.0948   0.0173   1.0000
  11.250   1.3536   0.08435   0.07872  -0.0928   0.0173   1.0000
  11.500   1.3358   0.08904   0.08365  -0.0917   0.0174   1.0000
  11.750   1.3172   0.09410   0.08895  -0.0914   0.0175   1.0000
  12.000   1.2977   0.09965   0.09471  -0.0920   0.0176   1.0000
  12.250   1.2779   0.10570   0.10097  -0.0936   0.0177   1.0000
<< Back to GOE 499 AIRFOIL (goe499-il)

Polar data table (+)

Polar graphs


<< Back to GOE 499 AIRFOIL (goe499-il)