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

GOE 483 AIRFOIL (goe483-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 483 AIRFOIL (goe483-il)
Reynolds number: 500,000
Max Cl/Cd: 114.35 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe483-il-500000.txt
Download as CSV file: xf-goe483-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 483 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.5311   0.00784   0.00156  -0.0973   0.7162   0.1461
   0.500   0.5578   0.00782   0.00154  -0.0969   0.7006   0.1618
   0.750   0.5844   0.00784   0.00154  -0.0966   0.6830   0.1821
   1.000   0.6108   0.00789   0.00156  -0.0962   0.6634   0.2054
   1.250   0.6372   0.00795   0.00157  -0.0958   0.6410   0.2243
   1.500   0.6633   0.00804   0.00158  -0.0953   0.6179   0.2410
   1.750   0.6893   0.00815   0.00161  -0.0949   0.5951   0.2553
   2.000   0.7154   0.00825   0.00167  -0.0945   0.5751   0.2694
   2.250   0.7414   0.00836   0.00174  -0.0942   0.5571   0.2851
   2.500   0.7674   0.00846   0.00183  -0.0939   0.5407   0.3087
   2.750   0.7951   0.00728   0.00198  -0.0941   0.5259   1.0000
   3.000   0.8214   0.00746   0.00207  -0.0938   0.5127   1.0000
   3.250   0.8477   0.00763   0.00221  -0.0935   0.5001   1.0000
   3.500   0.8740   0.00780   0.00233  -0.0932   0.4880   1.0000
   3.750   0.9003   0.00797   0.00247  -0.0929   0.4762   1.0000
   4.000   0.9265   0.00815   0.00261  -0.0925   0.4638   1.0000
   4.250   0.9525   0.00833   0.00276  -0.0922   0.4488   1.0000
   4.500   0.9778   0.00856   0.00293  -0.0917   0.4245   1.0000
   4.750   1.0030   0.00881   0.00310  -0.0913   0.4002   1.0000
   5.000   1.0279   0.00910   0.00331  -0.0909   0.3740   1.0000
   5.250   1.0518   0.00950   0.00356  -0.0903   0.3376   1.0000
   5.500   1.0747   0.01004   0.00390  -0.0896   0.2918   1.0000
   5.750   1.0929   0.01121   0.00449  -0.0884   0.1907   1.0000
   6.000   1.1015   0.01377   0.00603  -0.0859   0.0232   1.0000
   6.250   1.1249   0.01431   0.00668  -0.0851   0.0190   1.0000
   6.500   1.1474   0.01495   0.00744  -0.0842   0.0168   1.0000
   6.750   1.1679   0.01587   0.00848  -0.0829   0.0150   1.0000
   7.000   1.1832   0.01738   0.01019  -0.0809   0.0136   1.0000
   7.250   1.2031   0.01819   0.01109  -0.0797   0.0127   1.0000
   7.500   1.2212   0.01919   0.01218  -0.0782   0.0122   1.0000
   7.750   1.2375   0.02036   0.01345  -0.0764   0.0117   1.0000
   8.000   1.2527   0.02166   0.01485  -0.0745   0.0110   1.0000
   8.250   1.2670   0.02315   0.01644  -0.0725   0.0107   1.0000
   8.500   1.2814   0.02492   0.01830  -0.0704   0.0106   1.0000
   8.750   1.2982   0.02716   0.02066  -0.0687   0.0110   1.0000
<< Back to GOE 483 AIRFOIL (goe483-il)

Polar data table (+)

Polar graphs


<< Back to GOE 483 AIRFOIL (goe483-il)