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=200,000 Ncrit=9


Details Polar file
Airfoil: AG09 (ag09-il)
Reynolds number: 200,000
Max Cl/Cd: 56.35 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag09-il-200000.txt
Download as CSV file: xf-ag09-il-200000.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 =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.6223   0.13480   0.13122   0.0354   1.0000   0.0300
 -10.000  -0.6211   0.13221   0.12866   0.0318   1.0000   0.0304
  -9.750  -0.6181   0.12910   0.12557   0.0288   1.0000   0.0306
  -9.500  -0.6143   0.12569   0.12219   0.0261   1.0000   0.0307
  -9.250  -0.6096   0.12211   0.11862   0.0238   1.0000   0.0308
  -9.000  -0.6049   0.11841   0.11495   0.0214   1.0000   0.0308
  -8.750  -0.6038   0.11140   0.10797   0.0221   1.0000   0.0315
  -8.500  -0.5946   0.10724   0.10381   0.0238   1.0000   0.0324
  -8.250  -0.5879   0.10370   0.10028   0.0234   1.0000   0.0332
  -8.000  -0.5822   0.10024   0.09684   0.0222   1.0000   0.0340
  -7.750  -0.5770   0.09675   0.09338   0.0205   1.0000   0.0349
  -7.500  -0.5719   0.09321   0.08986   0.0183   1.0000   0.0359
  -7.250  -0.5650   0.08939   0.08607   0.0150   1.0000   0.0370
  -7.000  -0.5543   0.08516   0.08184   0.0102   1.0000   0.0385
  -6.750  -0.5361   0.08025   0.07691   0.0008   1.0000   0.0408
  -6.500  -0.5058   0.07489   0.07140  -0.0121   1.0000   0.0419
  -6.250  -0.4839   0.06995   0.06632  -0.0174   1.0000   0.0421
  -6.000  -0.4816   0.06343   0.05991  -0.0168   1.0000   0.0435
  -5.750  -0.4674   0.06030   0.05676  -0.0168   1.0000   0.0450
  -5.500  -0.4478   0.05680   0.05319  -0.0188   1.0000   0.0476
  -5.250  -0.4023   0.05252   0.04826  -0.0278   1.0000   0.0545
  -5.000  -0.3929   0.04645   0.04241  -0.0280   1.0000   0.0568
  -4.750  -0.3715   0.04349   0.03939  -0.0287   1.0000   0.0592
  -4.500  -0.3424   0.03994   0.03556  -0.0309   1.0000   0.0653
  -4.250  -0.3146   0.03521   0.03045  -0.0331   1.0000   0.0699
  -4.000  -0.2791   0.02559   0.01985  -0.0342   1.0000   0.0401
  -3.750  -0.2507   0.02131   0.01500  -0.0344   1.0000   0.0380
  -3.500  -0.2221   0.01899   0.01214  -0.0342   1.0000   0.0404
  -3.250  -0.1936   0.01740   0.01011  -0.0337   1.0000   0.0422
  -3.000  -0.1665   0.01482   0.00728  -0.0335   1.0000   0.0454
  -2.750  -0.1393   0.01406   0.00643  -0.0331   1.0000   0.0522
  -2.500  -0.1124   0.01289   0.00521  -0.0327   1.0000   0.0602
  -2.250  -0.0856   0.01212   0.00440  -0.0323   1.0000   0.0724
  -2.000  -0.0590   0.01147   0.00379  -0.0319   1.0000   0.0877
  -1.750  -0.0324   0.01104   0.00337  -0.0315   1.0000   0.1084
  -1.500  -0.0062   0.01048   0.00296  -0.0312   1.0000   0.1329
  -1.250   0.0201   0.01000   0.00261  -0.0309   1.0000   0.1661
  -1.000   0.0462   0.00941   0.00237  -0.0307   1.0000   0.2312
  -0.750   0.0691   0.00695   0.00225  -0.0293   1.0000   1.0000
  -0.500   0.0950   0.00699   0.00212  -0.0289   1.0000   1.0000
  -0.250   0.1208   0.00703   0.00207  -0.0284   1.0000   1.0000
   0.000   0.1465   0.00709   0.00204  -0.0280   1.0000   1.0000
   0.250   0.1720   0.00717   0.00208  -0.0277   1.0000   1.0000
   0.500   0.1976   0.00728   0.00216  -0.0275   1.0000   1.0000
   0.750   0.2456   0.00731   0.00217  -0.0318   0.9843   1.0000
   1.000   0.2916   0.00731   0.00215  -0.0354   0.9585   1.0000
   1.250   0.3310   0.00730   0.00212  -0.0373   0.9232   1.0000
   1.500   0.3609   0.00735   0.00210  -0.0369   0.8772   1.0000
   1.750   0.3847   0.00749   0.00210  -0.0351   0.8251   1.0000
   2.000   0.4078   0.00771   0.00212  -0.0334   0.7696   1.0000
   2.250   0.4316   0.00799   0.00217  -0.0319   0.7104   1.0000
   2.500   0.4559   0.00834   0.00225  -0.0307   0.6484   1.0000
   2.750   0.4808   0.00872   0.00239  -0.0299   0.5860   1.0000
   3.000   0.5062   0.00914   0.00255  -0.0292   0.5251   1.0000
   3.250   0.5318   0.00958   0.00274  -0.0287   0.4699   1.0000
   3.500   0.5577   0.01003   0.00297  -0.0283   0.4213   1.0000
   3.750   0.5838   0.01046   0.00327  -0.0279   0.3776   1.0000
   4.000   0.6101   0.01090   0.00357  -0.0276   0.3384   1.0000
   4.250   0.6364   0.01133   0.00390  -0.0274   0.3026   1.0000
   4.500   0.6626   0.01178   0.00425  -0.0271   0.2694   1.0000
   4.750   0.6888   0.01225   0.00462  -0.0269   0.2380   1.0000
   5.000   0.7151   0.01269   0.00506  -0.0267   0.2062   1.0000
   5.250   0.7412   0.01318   0.00548  -0.0265   0.1737   1.0000
   5.500   0.7671   0.01374   0.00596  -0.0263   0.1378   1.0000
   5.750   0.7923   0.01456   0.00661  -0.0260   0.0950   1.0000
   6.000   0.8164   0.01588   0.00780  -0.0255   0.0583   1.0000
   6.250   0.8392   0.01754   0.00938  -0.0249   0.0413   1.0000
   6.500   0.8638   0.01881   0.01079  -0.0241   0.0349   1.0000
   6.750   0.8850   0.02124   0.01326  -0.0232   0.0302   1.0000
   7.000   0.9101   0.02230   0.01454  -0.0225   0.0270   1.0000
   7.250   0.9335   0.02425   0.01671  -0.0217   0.0253   1.0000
   7.500   0.9561   0.02652   0.01928  -0.0208   0.0243   1.0000
   7.750   0.9773   0.02930   0.02243  -0.0199   0.0238   1.0000
   8.000   0.9960   0.03295   0.02660  -0.0187   0.0240   1.0000
<< Back to AG09 (ag09-il)

Polar data table (+)

Polar graphs


<< Back to AG09 (ag09-il)