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

GOE 682 AIRFOIL (goe682-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 682 AIRFOIL (goe682-il)
Reynolds number: 50,000
Max Cl/Cd: 25.45 at α=8°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe682-il-50000.txt
Download as CSV file: xf-goe682-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 682 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3056   0.11101   0.10395  -0.0302   1.0000   0.2135
  -8.750  -0.3022   0.10857   0.10156  -0.0293   1.0000   0.2217
  -8.500  -0.3305   0.10982   0.10303  -0.0293   1.0000   0.2256
  -8.250  -0.3018   0.10354   0.09669  -0.0276   1.0000   0.2327
  -8.000  -0.3142   0.10272   0.09601  -0.0265   1.0000   0.2399
  -7.750  -0.3532   0.10447   0.09802  -0.0242   1.0000   0.2419
  -7.500  -0.3135   0.09756   0.09102  -0.0231   1.0000   0.2514
  -7.250  -0.3436   0.09811   0.09177  -0.0201   1.0000   0.2569
  -7.000  -0.3411   0.09497   0.08870  -0.0179   1.0000   0.2620
  -6.750  -0.3493   0.09355   0.08738  -0.0148   1.0000   0.2701
  -6.500  -0.3889   0.09448   0.08852  -0.0120   1.0000   0.2737
  -6.250  -0.3727   0.09048   0.08454  -0.0086   1.0000   0.2832
  -6.000  -0.4053   0.09069   0.08491  -0.0075   1.0000   0.2897
  -5.750  -0.3950   0.08728   0.08154  -0.0036   1.0000   0.2994
  -5.500  -0.4209   0.08656   0.08096  -0.0036   1.0000   0.3077
  -5.250  -0.4165   0.08402   0.07845   0.0000   1.0000   0.3196
  -5.000  -0.4200   0.08164   0.07613   0.0023   1.0000   0.3290
  -4.750  -0.4286   0.07974   0.07429   0.0031   1.0000   0.3420
  -4.500  -0.4328   0.07769   0.07229   0.0044   1.0000   0.3572
  -4.250  -0.4337   0.07546   0.07010   0.0065   1.0000   0.3735
  -4.000  -0.4329   0.07316   0.06786   0.0092   1.0000   0.3911
  -3.750  -0.4316   0.07101   0.06575   0.0121   1.0000   0.4110
  -3.500  -0.4331   0.06902   0.06381   0.0153   1.0000   0.4398
  -3.250  -0.4341   0.06713   0.06199   0.0203   1.0000   0.4733
  -3.000  -0.4366   0.06555   0.06048   0.0260   1.0000   0.5140
  -2.750  -0.4387   0.06365   0.05866   0.0337   1.0000   0.5571
  -2.500  -0.4432   0.06184   0.05693   0.0402   1.0000   0.6009
  -2.250  -0.2142   0.04094   0.03201  -0.0429   1.0000   0.1775
  -2.000  -0.1939   0.03931   0.03022  -0.0430   1.0000   0.1792
  -1.750  -0.1718   0.03787   0.02846  -0.0432   1.0000   0.1798
  -1.500  -0.1280   0.03683   0.02710  -0.0470   0.9912   0.1841
  -1.250  -0.0814   0.03614   0.02591  -0.0511   0.9807   0.1950
  -1.000  -0.0349   0.03565   0.02519  -0.0550   0.9699   0.2072
  -0.750   0.0032   0.03527   0.02464  -0.0575   0.9584   0.2271
  -0.500   0.0405   0.03493   0.02430  -0.0596   0.9470   0.2535
  -0.250   0.0844   0.03450   0.02397  -0.0627   0.9360   0.3119
   0.000   0.1299   0.03267   0.02356  -0.0657   0.9271   0.5348
   0.250   0.1738   0.03199   0.02335  -0.0685   0.9144   1.0000
   0.500   0.2075   0.03287   0.02381  -0.0703   0.9015   1.0000
   0.750   0.2419   0.03377   0.02441  -0.0722   0.8889   1.0000
   1.000   0.2797   0.03468   0.02505  -0.0746   0.8767   1.0000
   1.250   0.3143   0.03554   0.02570  -0.0764   0.8641   1.0000
   1.500   0.3392   0.03640   0.02641  -0.0768   0.8508   1.0000
   1.750   0.3658   0.03730   0.02718  -0.0774   0.8377   1.0000
   2.000   0.3949   0.03818   0.02795  -0.0782   0.8246   1.0000
   2.250   0.4271   0.03902   0.02868  -0.0794   0.8117   1.0000
   2.500   0.4637   0.03972   0.02929  -0.0809   0.7984   1.0000
   2.750   0.5020   0.04022   0.02973  -0.0823   0.7840   1.0000
   3.000   0.5373   0.04061   0.03007  -0.0831   0.7685   1.0000
   3.250   0.5720   0.04089   0.03033  -0.0836   0.7527   1.0000
   3.500   0.6058   0.04114   0.03057  -0.0839   0.7373   1.0000
   3.750   0.6359   0.04149   0.03091  -0.0838   0.7220   1.0000
   4.000   0.6652   0.04183   0.03128  -0.0835   0.7069   1.0000
   4.250   0.6945   0.04214   0.03161  -0.0831   0.6917   1.0000
   4.500   0.7239   0.04240   0.03191  -0.0826   0.6766   1.0000
   4.750   0.7560   0.04242   0.03199  -0.0822   0.6614   1.0000
   5.000   0.7882   0.04241   0.03204  -0.0817   0.6462   1.0000
   5.250   0.8226   0.04222   0.03192  -0.0813   0.6308   1.0000
   5.500   0.8567   0.04193   0.03172  -0.0807   0.6153   1.0000
   5.750   0.8896   0.04171   0.03158  -0.0799   0.5997   1.0000
   6.000   0.9237   0.04136   0.03133  -0.0792   0.5839   1.0000
   6.250   0.9565   0.04112   0.03116  -0.0784   0.5680   1.0000
   6.500   0.9855   0.04123   0.03135  -0.0774   0.5525   1.0000
   6.750   1.0129   0.04147   0.03169  -0.0764   0.5370   1.0000
   7.000   1.0361   0.04206   0.03236  -0.0751   0.5220   1.0000
   7.250   1.0527   0.04325   0.03363  -0.0735   0.5080   1.0000
   7.500   1.0680   0.04465   0.03512  -0.0719   0.4952   1.0000
   7.750   1.1028   0.04465   0.03523  -0.0717   0.4835   1.0000
   8.000   1.1378   0.04471   0.03538  -0.0715   0.4721   1.0000
   8.250   1.0912   0.05087   0.04166  -0.0664   0.4610   1.0000
   8.500   1.0933   0.05355   0.04442  -0.0645   0.4516   1.0000
   8.750   1.1605   0.05114   0.04221  -0.0658   0.4410   1.0000
   9.000   0.9311   0.07775   0.06833  -0.0636   0.4344   1.0000
   9.250   0.9328   0.08150   0.07216  -0.0636   0.4275   1.0000
   9.500   0.9253   0.08649   0.07718  -0.0640   0.4220   1.0000
   9.750   0.8943   0.09397   0.08465  -0.0656   0.4185   1.0000
  10.000   0.9287   0.09409   0.08496  -0.0642   0.4077   1.0000
  10.250   0.8976   0.10170   0.09253  -0.0662   0.4055   1.0000
  10.500   0.8818   0.10742   0.09828  -0.0675   0.4018   1.0000
<< Back to GOE 682 AIRFOIL (goe682-il)

Polar data table (+)

Polar graphs


<< Back to GOE 682 AIRFOIL (goe682-il)