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

AG10 (ag10-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG10 (ag10-il)
Reynolds number: 100,000
Max Cl/Cd: 41.67 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag10-il-100000-n5.txt
Download as CSV file: xf-ag10-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG10                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5971   0.10308   0.09814   0.0241   1.0000   0.0431
  -7.750  -0.5918   0.09964   0.09474   0.0225   1.0000   0.0447
  -7.500  -0.5871   0.09615   0.09127   0.0201   1.0000   0.0464
  -7.250  -0.5793   0.09218   0.08732   0.0139   1.0000   0.0491
  -7.000  -0.5624   0.08727   0.08238   0.0010   1.0000   0.0505
  -6.750  -0.5445   0.08227   0.07726  -0.0062   1.0000   0.0507
  -6.500  -0.5346   0.07681   0.07184  -0.0072   1.0000   0.0513
  -6.000  -0.5014   0.06352   0.05834  -0.0126   1.0000   0.0309
  -5.750  -0.4839   0.05933   0.05409  -0.0148   1.0000   0.0299
  -5.500  -0.4617   0.05425   0.04885  -0.0183   1.0000   0.0289
  -5.250  -0.4363   0.04862   0.04298  -0.0222   1.0000   0.0280
  -5.000  -0.4089   0.04274   0.03676  -0.0256   1.0000   0.0273
  -4.750  -0.3807   0.03741   0.03100  -0.0281   1.0000   0.0277
  -4.500  -0.3516   0.03259   0.02561  -0.0298   1.0000   0.0295
  -4.250  -0.3221   0.02813   0.02041  -0.0309   1.0000   0.0305
  -4.000  -0.2927   0.02462   0.01617  -0.0313   1.0000   0.0312
  -3.750  -0.2654   0.02246   0.01373  -0.0315   1.0000   0.0330
  -3.500  -0.2380   0.02121   0.01227  -0.0315   1.0000   0.0366
  -3.250  -0.2095   0.01940   0.01000  -0.0312   1.0000   0.0393
  -3.000  -0.1822   0.01805   0.00856  -0.0310   1.0000   0.0430
  -2.750  -0.1549   0.01710   0.00747  -0.0307   1.0000   0.0487
  -2.500  -0.1279   0.01606   0.00640  -0.0304   1.0000   0.0543
  -2.250  -0.1007   0.01533   0.00556  -0.0300   1.0000   0.0630
  -2.000  -0.0740   0.01462   0.00492  -0.0297   1.0000   0.0750
  -1.750  -0.0472   0.01400   0.00435  -0.0294   1.0000   0.0910
  -1.500  -0.0207   0.01343   0.00390  -0.0291   1.0000   0.1140
  -1.250   0.0059   0.01293   0.00352  -0.0287   1.0000   0.1492
  -1.000   0.0321   0.01238   0.00327  -0.0285   1.0000   0.2120
  -0.750   0.0582   0.01151   0.00308  -0.0284   1.0000   0.3882
  -0.500   0.0811   0.00979   0.00289  -0.0263   1.0000   1.0000
  -0.250   0.1073   0.00982   0.00276  -0.0259   1.0000   1.0000
   0.000   0.1333   0.00985   0.00269  -0.0255   1.0000   1.0000
   0.250   0.1592   0.00990   0.00268  -0.0250   1.0000   1.0000
   0.500   0.1848   0.00997   0.00270  -0.0246   1.0000   1.0000
   0.750   0.2151   0.01005   0.00276  -0.0252   0.9927   1.0000
   1.000   0.2597   0.01012   0.00280  -0.0285   0.9530   1.0000
   1.250   0.3017   0.01020   0.00283  -0.0310   0.8951   1.0000
   1.500   0.3350   0.01036   0.00284  -0.0312   0.8173   1.0000
   1.750   0.3600   0.01066   0.00284  -0.0295   0.7319   1.0000
   2.000   0.3836   0.01106   0.00290  -0.0278   0.6476   1.0000
   2.250   0.4081   0.01152   0.00301  -0.0267   0.5721   1.0000
   2.500   0.4334   0.01198   0.00320  -0.0259   0.5089   1.0000
   2.750   0.4593   0.01243   0.00341  -0.0253   0.4566   1.0000
   3.000   0.4855   0.01287   0.00366  -0.0249   0.4127   1.0000
   3.250   0.5118   0.01330   0.00394  -0.0245   0.3745   1.0000
   3.500   0.5383   0.01372   0.00429  -0.0241   0.3393   1.0000
   3.750   0.5648   0.01415   0.00463  -0.0238   0.3061   1.0000
   4.000   0.5912   0.01460   0.00500  -0.0236   0.2746   1.0000
   4.250   0.6176   0.01507   0.00544  -0.0233   0.2444   1.0000
   4.500   0.6439   0.01557   0.00589  -0.0230   0.2149   1.0000
   4.750   0.6702   0.01612   0.00638  -0.0228   0.1862   1.0000
   5.000   0.6963   0.01671   0.00693  -0.0226   0.1584   1.0000
   5.250   0.7222   0.01737   0.00759  -0.0224   0.1317   1.0000
   5.500   0.7480   0.01811   0.00831  -0.0221   0.1066   1.0000
   5.750   0.7733   0.01900   0.00915  -0.0219   0.0846   1.0000
   6.000   0.7984   0.01998   0.01015  -0.0216   0.0664   1.0000
   6.250   0.8230   0.02113   0.01134  -0.0212   0.0534   1.0000
   6.500   0.8472   0.02237   0.01269  -0.0207   0.0444   1.0000
   6.750   0.8709   0.02378   0.01424  -0.0202   0.0381   1.0000
   7.000   0.8939   0.02525   0.01580  -0.0197   0.0340   1.0000
   7.250   0.9169   0.02693   0.01775  -0.0191   0.0302   1.0000
   7.500   0.9394   0.02862   0.01964  -0.0185   0.0277   1.0000
   7.750   0.9602   0.03076   0.02190  -0.0180   0.0261   1.0000
   8.000   0.9806   0.03346   0.02504  -0.0172   0.0250   1.0000
   8.250   0.9995   0.03628   0.02840  -0.0165   0.0233   1.0000
   8.500   1.0169   0.03880   0.03128  -0.0160   0.0218   1.0000
   8.750   1.0325   0.04139   0.03415  -0.0157   0.0208   1.0000
   9.000   1.0439   0.04492   0.03802  -0.0152   0.0203   1.0000
   9.250   1.0492   0.04948   0.04306  -0.0150   0.0201   1.0000
   9.500   1.0473   0.05475   0.04887  -0.0152   0.0200   1.0000
   9.750   1.0387   0.06035   0.05489  -0.0163   0.0199   1.0000
  10.000   1.0238   0.06598   0.06081  -0.0183   0.0200   1.0000
  10.250   1.0058   0.07243   0.06744  -0.0232   0.0201   1.0000
<< Back to AG10 (ag10-il)

Polar data table (+)

Polar graphs


<< Back to AG10 (ag10-il)