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

AG19 (ag19-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: AG19 (ag19-il)
Reynolds number: 100,000
Max Cl/Cd: 48.61 at α=4°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-ag19-il-100000-n5.txt
Download as CSV file: xf-ag19-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: AG19                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.5320   0.10933   0.10441   0.0143   1.0000   0.0369
  -8.250  -0.5300   0.10654   0.10169   0.0097   1.0000   0.0384
  -8.000  -0.5267   0.10350   0.09871   0.0052   1.0000   0.0388
  -7.750  -0.5203   0.09994   0.09521   0.0005   1.0000   0.0391
  -7.500  -0.5093   0.09571   0.09100  -0.0055   1.0000   0.0392
  -7.250  -0.5072   0.09098   0.08632   0.0031   1.0000   0.0421
  -7.000  -0.4975   0.08741   0.08278   0.0012   1.0000   0.0439
  -6.750  -0.4864   0.08339   0.07879  -0.0026   1.0000   0.0451
  -6.500  -0.4732   0.07905   0.07446  -0.0073   1.0000   0.0458
  -6.250  -0.4576   0.07441   0.06982  -0.0126   1.0000   0.0462
  -6.000  -0.4394   0.06938   0.06477  -0.0183   1.0000   0.0455
  -5.750  -0.4080   0.05969   0.05487  -0.0295   1.0000   0.0257
  -5.500  -0.3871   0.05459   0.04967  -0.0336   1.0000   0.0245
  -5.250  -0.3619   0.04900   0.04390  -0.0384   1.0000   0.0234
  -5.000  -0.3341   0.04318   0.03781  -0.0429   1.0000   0.0224
  -4.750  -0.3049   0.03750   0.03172  -0.0466   1.0000   0.0216
  -4.500  -0.2753   0.03241   0.02605  -0.0491   1.0000   0.0210
  -4.250  -0.2457   0.02815   0.02119  -0.0506   1.0000   0.0207
  -4.000  -0.2166   0.02482   0.01727  -0.0513   1.0000   0.0209
  -3.750  -0.1883   0.02230   0.01424  -0.0514   1.0000   0.0219
  -3.500  -0.1605   0.02064   0.01217  -0.0512   1.0000   0.0253
  -3.250  -0.1329   0.01906   0.01022  -0.0508   1.0000   0.0274
  -3.000  -0.1063   0.01716   0.00818  -0.0505   1.0000   0.0299
  -2.750  -0.0798   0.01601   0.00694  -0.0501   1.0000   0.0342
  -2.500  -0.0534   0.01511   0.00594  -0.0498   1.0000   0.0444
  -2.250  -0.0266   0.01421   0.00500  -0.0494   1.0000   0.0595
  -2.000   0.0001   0.01337   0.00432  -0.0493   1.0000   0.0997
  -1.750   0.0264   0.01257   0.00394  -0.0493   1.0000   0.1972
  -1.500   0.0523   0.01184   0.00375  -0.0493   1.0000   0.3492
  -1.250   0.0777   0.01106   0.00361  -0.0489   1.0000   0.5490
  -1.000   0.0960   0.00991   0.00341  -0.0459   1.0000   1.0000
  -0.750   0.1221   0.00998   0.00327  -0.0457   1.0000   1.0000
  -0.500   0.1478   0.01007   0.00320  -0.0454   1.0000   1.0000
  -0.250   0.1861   0.01017   0.00314  -0.0478   0.9887   1.0000
   0.000   0.2292   0.01025   0.00305  -0.0509   0.9713   1.0000
   0.250   0.2690   0.01031   0.00300  -0.0532   0.9506   1.0000
   0.500   0.3077   0.01035   0.00297  -0.0552   0.9282   1.0000
   0.750   0.3430   0.01040   0.00294  -0.0563   0.9021   1.0000
   1.000   0.3756   0.01047   0.00292  -0.0567   0.8733   1.0000
   1.250   0.4048   0.01055   0.00293  -0.0564   0.8408   1.0000
   1.500   0.4322   0.01067   0.00294  -0.0556   0.8064   1.0000
   1.750   0.4583   0.01083   0.00302  -0.0546   0.7692   1.0000
   2.000   0.4839   0.01101   0.00307  -0.0535   0.7298   1.0000
   2.250   0.5094   0.01124   0.00316  -0.0525   0.6873   1.0000
   2.500   0.5348   0.01151   0.00327  -0.0515   0.6415   1.0000
   2.750   0.5600   0.01183   0.00341  -0.0506   0.5918   1.0000
   3.000   0.5850   0.01220   0.00364  -0.0497   0.5400   1.0000
   3.250   0.6099   0.01263   0.00386  -0.0489   0.4885   1.0000
   3.500   0.6350   0.01309   0.00413  -0.0482   0.4392   1.0000
   3.750   0.6601   0.01359   0.00445  -0.0476   0.3936   1.0000
   4.000   0.6854   0.01410   0.00482  -0.0472   0.3521   1.0000
   4.250   0.7106   0.01464   0.00530  -0.0467   0.3137   1.0000
   4.500   0.7360   0.01518   0.00576  -0.0463   0.2784   1.0000
   4.750   0.7612   0.01576   0.00626  -0.0459   0.2460   1.0000
   5.000   0.7864   0.01635   0.00682  -0.0455   0.2156   1.0000
   5.250   0.8115   0.01699   0.00742  -0.0452   0.1877   1.0000
   5.500   0.8363   0.01768   0.00815  -0.0448   0.1620   1.0000
   5.750   0.8610   0.01842   0.00891  -0.0444   0.1398   1.0000
   6.000   0.8855   0.01921   0.00972  -0.0439   0.1188   1.0000
   6.250   0.9096   0.02010   0.01067  -0.0434   0.1022   1.0000
   6.500   0.9332   0.02106   0.01164  -0.0430   0.0857   1.0000
   6.750   0.9564   0.02217   0.01282  -0.0424   0.0728   1.0000
   7.000   0.9791   0.02333   0.01407  -0.0418   0.0598   1.0000
   7.250   1.0018   0.02458   0.01558  -0.0411   0.0496   1.0000
   7.500   1.0237   0.02602   0.01721  -0.0404   0.0404   1.0000
   7.750   1.0438   0.02784   0.01919  -0.0395   0.0341   1.0000
   8.000   1.0642   0.02966   0.02128  -0.0386   0.0288   1.0000
   8.250   1.0823   0.03177   0.02358  -0.0378   0.0244   1.0000
   8.500   1.1001   0.03451   0.02676  -0.0365   0.0219   1.0000
   8.750   1.1155   0.03755   0.03022  -0.0354   0.0201   1.0000
   9.000   1.1294   0.04010   0.03305  -0.0345   0.0182   1.0000
   9.500   1.1423   0.04745   0.04123  -0.0325   0.0155   1.0000
   9.750   1.1413   0.05189   0.04619  -0.0316   0.0149   1.0000
  10.000   1.1330   0.05672   0.05145  -0.0310   0.0147   1.0000
  10.250   1.1177   0.06143   0.05648  -0.0308   0.0146   1.0000
  10.500   1.0994   0.06679   0.06207  -0.0329   0.0147   1.0000
  10.750   1.0808   0.07368   0.06916  -0.0382   0.0149   1.0000
  11.000   1.0619   0.08258   0.07820  -0.0458   0.0152   1.0000
  11.250   1.0431   0.09250   0.08819  -0.0534   0.0158   1.0000
  11.500   1.0249   0.10206   0.09776  -0.0593   0.0163   1.0000
<< Back to AG19 (ag19-il)

Polar data table (+)

Polar graphs


<< Back to AG19 (ag19-il)