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

AG09 (ag09-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: AG09 (ag09-il)
Reynolds number: 500,000
Max Cl/Cd: 74.35 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag09-il-500000.txt
Download as CSV file: xf-ag09-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG09                                            
  
 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   0.2143   0.00454   0.00113  -0.0349   0.8921   1.0000
   0.500   0.2383   0.00468   0.00108  -0.0335   0.8465   1.0000
   0.750   0.2631   0.00487   0.00105  -0.0325   0.7998   1.0000
   1.000   0.2889   0.00508   0.00104  -0.0317   0.7514   1.0000
   1.250   0.3150   0.00532   0.00104  -0.0311   0.7008   1.0000
   1.500   0.3415   0.00556   0.00106  -0.0306   0.6486   1.0000
   1.750   0.3683   0.00583   0.00111  -0.0303   0.5963   1.0000
   2.000   0.3952   0.00611   0.00117  -0.0300   0.5435   1.0000
   2.250   0.4222   0.00640   0.00125  -0.0298   0.4932   1.0000
   2.500   0.4493   0.00669   0.00135  -0.0297   0.4471   1.0000
   2.750   0.4765   0.00697   0.00148  -0.0295   0.4053   1.0000
   3.000   0.5037   0.00725   0.00162  -0.0294   0.3673   1.0000
   3.250   0.5310   0.00753   0.00176  -0.0293   0.3325   1.0000
   3.500   0.5582   0.00780   0.00192  -0.0292   0.3000   1.0000
   3.750   0.5855   0.00809   0.00212  -0.0291   0.2700   1.0000
   4.000   0.6126   0.00838   0.00231  -0.0290   0.2414   1.0000
   4.250   0.6397   0.00868   0.00252  -0.0289   0.2143   1.0000
   4.500   0.6668   0.00899   0.00275  -0.0288   0.1882   1.0000
   4.750   0.6937   0.00933   0.00302  -0.0287   0.1617   1.0000
   5.000   0.7205   0.00970   0.00330  -0.0286   0.1348   1.0000
   5.250   0.7471   0.01012   0.00362  -0.0285   0.1064   1.0000
   5.500   0.7734   0.01064   0.00400  -0.0283   0.0760   1.0000
   5.750   0.7993   0.01128   0.00450  -0.0282   0.0464   1.0000
   6.000   0.8249   0.01209   0.00525  -0.0278   0.0263   1.0000
   6.250   0.8506   0.01292   0.00613  -0.0274   0.0189   1.0000
   6.500   0.8757   0.01388   0.00715  -0.0270   0.0153   1.0000
   6.750   0.8999   0.01510   0.00854  -0.0264   0.0138   1.0000
   7.000   0.9245   0.01611   0.00968  -0.0258   0.0128   1.0000
   7.250   0.9483   0.01734   0.01107  -0.0252   0.0121   1.0000
   7.500   0.9716   0.01873   0.01263  -0.0245   0.0117   1.0000
   7.750   0.9950   0.01994   0.01396  -0.0239   0.0110   1.0000
   8.000   1.0174   0.02127   0.01539  -0.0234   0.0103   1.0000
   8.250   1.0370   0.02367   0.01804  -0.0225   0.0099   1.0000
   8.500   1.0548   0.02667   0.02137  -0.0215   0.0098   1.0000
   8.750   1.0697   0.03026   0.02535  -0.0205   0.0097   1.0000
   9.000   1.0815   0.03416   0.02971  -0.0194   0.0097   1.0000
<< Back to AG09 (ag09-il)

Polar data table (+)

Polar graphs


<< Back to AG09 (ag09-il)