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

AG04 (ag04-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: AG04 (ag04-il)
Reynolds number: 100,000
Max Cl/Cd: 45.42 at α=4.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag04-il-100000.txt
Download as CSV file: xf-ag04-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG04                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.5721   0.11485   0.11004   0.0197   1.0000   0.0753
  -8.750  -0.5753   0.11205   0.10731   0.0150   1.0000   0.0779
  -8.500  -0.5830   0.10963   0.10498   0.0079   1.0000   0.0787
  -8.250  -0.5769   0.10407   0.09947   0.0078   1.0000   0.0799
  -8.000  -0.5595   0.09965   0.09501   0.0121   1.0000   0.0831
  -7.750  -0.5541   0.09612   0.09152   0.0108   1.0000   0.0865
  -7.500  -0.5540   0.09258   0.08805   0.0058   1.0000   0.0903
  -7.250  -0.5503   0.08696   0.08242  -0.0118   1.0000   0.0928
  -7.000  -0.5398   0.08331   0.07881  -0.0023   1.0000   0.0953
  -6.750  -0.5283   0.07963   0.07514  -0.0033   1.0000   0.0994
  -6.500  -0.5143   0.07305   0.06842  -0.0190   1.0000   0.1067
  -6.250  -0.5028   0.06978   0.06525  -0.0143   1.0000   0.1095
  -6.000  -0.4834   0.06421   0.05943  -0.0248   1.0000   0.1204
  -5.750  -0.4697   0.06072   0.05606  -0.0219   1.0000   0.1233
  -5.500  -0.4492   0.05594   0.05108  -0.0272   1.0000   0.1351
  -5.250  -0.4312   0.05268   0.04783  -0.0267   1.0000   0.1418
  -5.000  -0.3885   0.03652   0.03026  -0.0374   1.0000   0.0618
  -4.750  -0.3632   0.03187   0.02505  -0.0381   1.0000   0.0598
  -4.500  -0.3367   0.02800   0.02058  -0.0383   1.0000   0.0596
  -4.250  -0.3090   0.02445   0.01637  -0.0379   1.0000   0.0585
  -4.000  -0.2813   0.02186   0.01328  -0.0373   1.0000   0.0608
  -3.750  -0.2543   0.01985   0.01095  -0.0367   1.0000   0.0691
  -3.500  -0.2273   0.01811   0.00898  -0.0359   1.0000   0.0806
  -3.250  -0.2011   0.01681   0.00769  -0.0353   1.0000   0.1049
  -3.000  -0.1756   0.01547   0.00654  -0.0348   1.0000   0.1447
  -2.750  -0.1504   0.01435   0.00568  -0.0342   1.0000   0.2024
  -2.500  -0.1258   0.01328   0.00506  -0.0337   1.0000   0.2823
  -2.250  -0.1021   0.01214   0.00460  -0.0330   1.0000   0.4211
  -2.000  -0.0830   0.01080   0.00434  -0.0305   1.0000   0.6698
  -1.750  -0.0448   0.01001   0.00392  -0.0311   1.0000   1.0000
  -1.500  -0.0197   0.01002   0.00366  -0.0307   1.0000   1.0000
  -1.250   0.0051   0.01005   0.00347  -0.0302   1.0000   1.0000
  -1.000   0.0297   0.01010   0.00332  -0.0297   1.0000   1.0000
  -0.750   0.0542   0.01016   0.00325  -0.0293   1.0000   1.0000
  -0.500   0.0784   0.01026   0.00323  -0.0289   1.0000   1.0000
  -0.250   0.1023   0.01038   0.00326  -0.0285   1.0000   1.0000
   0.000   0.1262   0.01055   0.00334  -0.0282   1.0000   1.0000
   0.250   0.1498   0.01076   0.00351  -0.0280   1.0000   1.0000
   0.500   0.1729   0.01104   0.00375  -0.0279   1.0000   1.0000
   0.750   0.1977   0.01139   0.00409  -0.0284   0.9988   1.0000
   1.000   0.2600   0.01157   0.00428  -0.0358   0.9771   1.0000
   1.250   0.3215   0.01160   0.00436  -0.0424   0.9556   1.0000
   1.500   0.3707   0.01155   0.00437  -0.0460   0.9277   1.0000
   1.750   0.4075   0.01152   0.00439  -0.0468   0.8936   1.0000
   2.000   0.4368   0.01152   0.00438  -0.0458   0.8579   1.0000
   2.250   0.4610   0.01160   0.00440  -0.0439   0.8184   1.0000
   2.500   0.4839   0.01174   0.00446  -0.0417   0.7775   1.0000
   2.750   0.5068   0.01195   0.00458  -0.0397   0.7353   1.0000
   3.000   0.5299   0.01221   0.00470  -0.0378   0.6920   1.0000
   3.250   0.5535   0.01253   0.00487  -0.0362   0.6466   1.0000
   3.500   0.5774   0.01290   0.00509  -0.0349   0.6000   1.0000
   3.750   0.6014   0.01332   0.00537  -0.0336   0.5539   1.0000
   4.000   0.6256   0.01379   0.00567  -0.0325   0.5075   1.0000
   4.250   0.6499   0.01431   0.00604  -0.0316   0.4619   1.0000
   4.500   0.6742   0.01489   0.00646  -0.0307   0.4182   1.0000
   4.750   0.6986   0.01551   0.00693  -0.0299   0.3762   1.0000
   5.000   0.7230   0.01618   0.00753  -0.0292   0.3362   1.0000
   5.250   0.7472   0.01690   0.00814  -0.0286   0.2978   1.0000
   5.500   0.7711   0.01770   0.00878  -0.0279   0.2613   1.0000
   5.750   0.7950   0.01854   0.00958  -0.0273   0.2242   1.0000
   6.000   0.8185   0.01955   0.01050  -0.0266   0.1895   1.0000
   6.250   0.8416   0.02063   0.01156  -0.0259   0.1565   1.0000
   6.500   0.8647   0.02201   0.01288  -0.0252   0.1293   1.0000
   6.750   0.8875   0.02344   0.01434  -0.0244   0.1060   1.0000
   7.000   0.9106   0.02524   0.01614  -0.0235   0.0898   1.0000
   7.250   0.9330   0.02705   0.01801  -0.0228   0.0760   1.0000
   7.500   0.9559   0.02966   0.02095  -0.0217   0.0671   1.0000
   7.750   0.9766   0.03223   0.02371  -0.0209   0.0589   1.0000
   8.000   0.9968   0.03548   0.02755  -0.0196   0.0546   1.0000
   8.250   1.0143   0.03891   0.03107  -0.0190   0.0499   1.0000
   8.500   1.0259   0.04274   0.03570  -0.0175   0.0470   1.0000
   8.750   1.0332   0.04771   0.04133  -0.0163   0.0462   1.0000
   9.000   1.0345   0.05322   0.04742  -0.0154   0.0463   1.0000
   9.250   1.0303   0.05894   0.05359  -0.0150   0.0468   1.0000
   9.500   1.0218   0.06463   0.05961  -0.0151   0.0475   1.0000
   9.750   1.0099   0.07020   0.06540  -0.0156   0.0481   1.0000
  10.000   0.9961   0.07555   0.07087  -0.0163   0.0488   1.0000
  10.250   0.9302   0.09280   0.08843  -0.0350   0.0547   1.0000
<< Back to AG04 (ag04-il)

Polar data table (+)

Polar graphs


<< Back to AG04 (ag04-il)