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

GOE 419 AIRFOIL (goe419-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 419 AIRFOIL (goe419-il)
Reynolds number: 1,000,000
Max Cl/Cd: 124.97 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe419-il-1000000.txt
Download as CSV file: xf-goe419-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 419 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3824   0.08975   0.08825  -0.0188   1.0000   0.0075
  -7.500  -0.3875   0.08776   0.08630  -0.0176   1.0000   0.0076
  -7.250  -0.3879   0.08546   0.08403  -0.0178   1.0000   0.0077
  -7.000  -0.3699   0.08159   0.08016  -0.0233   0.9988   0.0079
  -4.750  -0.0895   0.03854   0.03651  -0.0807   0.9468   0.0082
  -4.500  -0.0461   0.03316   0.03093  -0.0868   0.9395   0.0082
  -4.250   0.0011   0.02704   0.02452  -0.0932   0.9321   0.0081
  -4.000   0.0387   0.01403   0.01038  -0.0964   0.9179   0.0074
  -3.750   0.0728   0.01341   0.00955  -0.0978   0.9038   0.0082
  -3.500   0.0967   0.00976   0.00514  -0.0969   0.8899   0.0088
  -3.000   0.1476   0.00797   0.00292  -0.0957   0.8684   0.0112
  -2.750   0.1732   0.00747   0.00227  -0.0950   0.8596   0.0135
  -2.500   0.1999   0.00740   0.00214  -0.0947   0.8510   0.0155
  -2.250   0.2248   0.00692   0.00157  -0.0940   0.8421   0.0232
  -2.000   0.2521   0.00710   0.00176  -0.0939   0.8336   0.0264
  -1.500   0.3033   0.00668   0.00121  -0.0930   0.8154   0.0378
  -1.250   0.3294   0.00661   0.00107  -0.0926   0.8058   0.0390
  -1.000   0.3553   0.00650   0.00089  -0.0921   0.7954   0.0405
  -0.750   0.3809   0.00638   0.00073  -0.0916   0.7837   0.0487
  -0.500   0.4060   0.00618   0.00068  -0.0910   0.7707   0.1102
  -0.250   0.4310   0.00603   0.00067  -0.0905   0.7554   0.1815
   0.000   0.4556   0.00589   0.00068  -0.0899   0.7365   0.2666
   0.250   0.5044   0.00432   0.00082  -0.0954   0.7143   0.9869
   0.500   0.5451   0.00445   0.00082  -0.0984   0.6892   1.0000
   0.750   0.5680   0.00459   0.00084  -0.0974   0.6674   1.0000
   1.000   0.5911   0.00473   0.00088  -0.0964   0.6480   1.0000
   1.500   0.6361   0.00510   0.00102  -0.0941   0.5939   1.0000
   1.750   0.6589   0.00528   0.00109  -0.0931   0.5632   1.0000
   2.000   0.6767   0.00578   0.00119  -0.0911   0.4702   1.0000
   2.250   0.6919   0.00658   0.00144  -0.0887   0.3549   1.0000
   2.500   0.6928   0.00880   0.00226  -0.0839   0.0272   1.0000
   2.750   0.7170   0.00901   0.00250  -0.0831   0.0179   1.0000
   3.000   0.7406   0.00932   0.00290  -0.0822   0.0139   1.0000
   3.250   0.7619   0.00988   0.00359  -0.0807   0.0110   1.0000
   3.500   0.7830   0.01047   0.00426  -0.0793   0.0102   1.0000
   3.750   0.8072   0.01071   0.00451  -0.0786   0.0092   1.0000
   4.000   0.8283   0.01130   0.00517  -0.0772   0.0079   1.0000
   4.250   0.8463   0.01223   0.00619  -0.0751   0.0077   1.0000
   4.500   0.8625   0.01347   0.00750  -0.0726   0.0080   1.0000
   4.750   0.8792   0.01538   0.00946  -0.0701   0.0092   1.0000
   6.250   1.0138   0.02791   0.02261  -0.0638   0.0047   1.0000
   6.500   1.0311   0.03067   0.02563  -0.0617   0.0047   1.0000
   6.750   1.0461   0.03351   0.02873  -0.0594   0.0046   1.0000
   7.000   1.0589   0.03646   0.03195  -0.0569   0.0046   1.0000
   7.250   1.0704   0.03934   0.03508  -0.0544   0.0046   1.0000
   7.500   1.0794   0.04243   0.03846  -0.0516   0.0046   1.0000
   7.750   1.0860   0.04558   0.04186  -0.0488   0.0046   1.0000
   8.000   1.0902   0.04876   0.04528  -0.0459   0.0046   1.0000
   8.250   1.0919   0.05197   0.04871  -0.0429   0.0046   1.0000
   8.500   1.0908   0.05523   0.05219  -0.0398   0.0046   1.0000
   8.750   1.0868   0.05841   0.05555  -0.0368   0.0046   1.0000
   9.000   1.0793   0.06152   0.05883  -0.0336   0.0046   1.0000
   9.250   1.0661   0.06416   0.06162  -0.0296   0.0046   1.0000
   9.500   1.0490   0.06674   0.06431  -0.0258   0.0046   1.0000
   9.750   1.0313   0.06974   0.06743  -0.0234   0.0046   1.0000
  10.000   1.0137   0.07321   0.07102  -0.0224   0.0046   1.0000
  10.250   0.9958   0.07730   0.07522  -0.0227   0.0046   1.0000
  10.500   0.9783   0.08201   0.08003  -0.0243   0.0046   1.0000
<< Back to GOE 419 AIRFOIL (goe419-il)

Polar data table (+)

Polar graphs


<< Back to GOE 419 AIRFOIL (goe419-il)