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

AG11 (ag11-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: AG11 (ag11-il)
Reynolds number: 200,000
Max Cl/Cd: 59.91 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-ag11-il-200000.txt
Download as CSV file: xf-ag11-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG11                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.5202   0.10051   0.09707   0.0142   1.0000   0.0244
  -7.750  -0.5150   0.09706   0.09365   0.0123   1.0000   0.0250
  -7.500  -0.5100   0.09360   0.09022   0.0101   1.0000   0.0256
  -7.250  -0.5045   0.09007   0.08673   0.0073   1.0000   0.0262
  -7.000  -0.4948   0.08615   0.08283   0.0029   1.0000   0.0270
  -6.750  -0.4812   0.08202   0.07868  -0.0027   1.0000   0.0278
  -6.500  -0.4602   0.07771   0.07432  -0.0111   1.0000   0.0288
  -6.250  -0.4349   0.07364   0.07014  -0.0187   1.0000   0.0293
  -6.000  -0.4131   0.06980   0.06617  -0.0229   1.0000   0.0295
  -5.750  -0.3922   0.06598   0.06220  -0.0259   1.0000   0.0296
  -5.500  -0.3881   0.05892   0.05526  -0.0265   1.0000   0.0308
  -5.250  -0.3736   0.05552   0.05187  -0.0270   1.0000   0.0319
  -5.000  -0.3544   0.05219   0.04847  -0.0283   1.0000   0.0333
  -4.750  -0.3319   0.04875   0.04493  -0.0302   1.0000   0.0353
  -4.500  -0.3060   0.04533   0.04133  -0.0321   1.0000   0.0380
  -4.250  -0.2676   0.04457   0.03995  -0.0337   1.0000   0.0412
  -4.000  -0.2521   0.03802   0.03344  -0.0351   1.0000   0.0430
  -3.750  -0.2318   0.03533   0.03076  -0.0352   1.0000   0.0454
  -3.500  -0.2073   0.03301   0.02826  -0.0354   1.0000   0.0498
  -3.250  -0.1794   0.03034   0.02516  -0.0355   1.0000   0.0564
  -3.000  -0.1580   0.02806   0.02289  -0.0354   1.0000   0.0604
  -1.750  -0.0256   0.01715   0.01037  -0.0304   1.0000   0.0365
  -1.500  -0.0005   0.01592   0.00892  -0.0294   1.0000   0.0376
  -1.250   0.0254   0.01460   0.00743  -0.0286   1.0000   0.0367
  -1.000   0.0517   0.01345   0.00617  -0.0279   1.0000   0.0353
  -0.750   0.0916   0.01234   0.00504  -0.0301   0.9954   0.0363
  -0.500   0.1381   0.01133   0.00403  -0.0337   0.9864   0.0433
  -0.250   0.1759   0.00773   0.00329  -0.0353   0.9818   1.0000
   0.000   0.2221   0.00769   0.00297  -0.0388   0.9659   1.0000
   0.250   0.2638   0.00765   0.00278  -0.0414   0.9456   1.0000
   0.500   0.2985   0.00762   0.00262  -0.0422   0.9196   1.0000
   0.750   0.3267   0.00763   0.00251  -0.0416   0.8895   1.0000
   1.000   0.3515   0.00769   0.00242  -0.0402   0.8571   1.0000
   1.250   0.3755   0.00779   0.00237  -0.0387   0.8219   1.0000
   1.500   0.3997   0.00793   0.00232  -0.0373   0.7859   1.0000
   1.750   0.4244   0.00811   0.00232  -0.0362   0.7477   1.0000
   2.000   0.4494   0.00832   0.00233  -0.0352   0.7091   1.0000
   2.250   0.4747   0.00856   0.00238  -0.0343   0.6692   1.0000
   2.500   0.5002   0.00882   0.00245  -0.0335   0.6282   1.0000
   2.750   0.5258   0.00911   0.00254  -0.0329   0.5856   1.0000
   3.000   0.5516   0.00942   0.00265  -0.0323   0.5413   1.0000
   3.250   0.5773   0.00976   0.00279  -0.0318   0.4968   1.0000
   3.500   0.6031   0.01012   0.00296  -0.0314   0.4528   1.0000
   3.750   0.6289   0.01052   0.00318  -0.0310   0.4101   1.0000
   4.000   0.6548   0.01093   0.00341  -0.0307   0.3700   1.0000
   4.250   0.6806   0.01137   0.00368  -0.0304   0.3324   1.0000
   4.500   0.7065   0.01180   0.00398  -0.0302   0.2974   1.0000
   4.750   0.7322   0.01227   0.00435  -0.0299   0.2640   1.0000
   5.000   0.7578   0.01279   0.00472  -0.0297   0.2336   1.0000
   5.250   0.7833   0.01333   0.00515  -0.0295   0.2058   1.0000
   5.500   0.8087   0.01391   0.00564  -0.0292   0.1794   1.0000
   5.750   0.8339   0.01451   0.00620  -0.0289   0.1559   1.0000
   6.000   0.8585   0.01527   0.00684  -0.0286   0.1375   1.0000
   6.250   0.8829   0.01608   0.00759  -0.0281   0.1230   1.0000
   6.500   0.9077   0.01679   0.00831  -0.0278   0.1113   1.0000
   6.750   0.9326   0.01748   0.00910  -0.0273   0.1015   1.0000
   7.000   0.9565   0.01848   0.01011  -0.0268   0.0942   1.0000
   7.250   0.9810   0.01928   0.01097  -0.0264   0.0879   1.0000
   7.500   1.0043   0.02051   0.01223  -0.0258   0.0824   1.0000
   7.750   1.0287   0.02127   0.01316  -0.0253   0.0768   1.0000
   8.000   1.0515   0.02275   0.01462  -0.0248   0.0723   1.0000
   8.250   1.0751   0.02404   0.01616  -0.0242   0.0690   1.0000
   8.500   1.0983   0.02514   0.01748  -0.0236   0.0651   1.0000
   8.750   1.1206   0.02637   0.01873  -0.0231   0.0612   1.0000
   9.000   1.1409   0.02875   0.02139  -0.0224   0.0584   1.0000
   9.250   1.1617   0.03036   0.02340  -0.0215   0.0557   1.0000
   9.500   1.1818   0.03171   0.02498  -0.0207   0.0524   1.0000
   9.750   1.2010   0.03335   0.02668  -0.0202   0.0495   1.0000
  10.000   1.2119   0.03742   0.03119  -0.0190   0.0475   1.0000
  10.250   1.2247   0.03953   0.03380  -0.0176   0.0454   1.0000
  10.500   1.2357   0.04183   0.03647  -0.0164   0.0428   1.0000
  10.750   1.2462   0.04404   0.03891  -0.0153   0.0408   1.0000
  11.000   1.2570   0.04601   0.04095  -0.0144   0.0388   1.0000
  11.500   1.2315   0.05586   0.05158  -0.0113   0.0371   1.0000
  11.750   1.2125   0.05944   0.05545  -0.0097   0.0368   1.0000
  12.000   1.1897   0.06442   0.06063  -0.0105   0.0369   1.0000
  12.250   1.1643   0.07113   0.06753  -0.0139   0.0372   1.0000
  12.500   1.0049   0.11668   0.11344  -0.0482   0.0495   1.0000
  12.750   0.9228   0.15021   0.14671  -0.0666   0.0748   1.0000
  13.000   0.9348   0.15246   0.14901  -0.0651   0.0731   1.0000
<< Back to AG11 (ag11-il)

Polar data table (+)

Polar graphs


<< Back to AG11 (ag11-il)