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=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 682 AIRFOIL (goe682-il)
Reynolds number: 100,000
Max Cl/Cd: 57.45 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe682-il-100000-n5.txt
Download as CSV file: xf-goe682-il-100000-n5.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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.3371   0.09155   0.08667  -0.0384   1.0000   0.0547
  -8.250  -0.3492   0.08910   0.08431  -0.0371   1.0000   0.0537
  -8.000  -0.3682   0.08686   0.08218  -0.0350   1.0000   0.0525
  -7.750  -0.3746   0.08364   0.07902  -0.0365   0.9961   0.0525
  -7.500  -0.3574   0.07842   0.07378  -0.0441   0.9870   0.0535
  -7.250  -0.3398   0.07235   0.06766  -0.0529   0.9774   0.0539
  -7.000  -0.3231   0.06511   0.06031  -0.0626   0.9673   0.0532
  -6.750  -0.3058   0.05688   0.05187  -0.0723   0.9570   0.0526
  -6.500  -0.2862   0.04938   0.04404  -0.0798   0.9480   0.0530
  -6.250  -0.2618   0.04318   0.03740  -0.0855   0.9410   0.0547
  -6.000  -0.2420   0.03805   0.03173  -0.0883   0.9318   0.0553
  -5.750  -0.2123   0.03368   0.02673  -0.0916   0.9270   0.0560
  -5.500  -0.1898   0.03092   0.02348  -0.0920   0.9179   0.0567
  -5.250  -0.1573   0.02857   0.02055  -0.0938   0.9133   0.0585
  -5.000  -0.1321   0.02699   0.01854  -0.0938   0.9047   0.0600
  -4.750  -0.1003   0.02539   0.01678  -0.0950   0.8994   0.0615
  -4.500  -0.0719   0.02420   0.01541  -0.0954   0.8922   0.0627
  -4.250  -0.0416   0.02312   0.01415  -0.0959   0.8854   0.0643
  -4.000  -0.0072   0.02205   0.01289  -0.0971   0.8805   0.0662
  -3.750   0.0187   0.02131   0.01199  -0.0966   0.8710   0.0686
  -3.500   0.0538   0.02056   0.01102  -0.0977   0.8654   0.0722
  -3.250   0.0791   0.01973   0.01020  -0.0971   0.8548   0.0753
  -3.000   0.1108   0.01901   0.00944  -0.0976   0.8471   0.0787
  -2.750   0.1384   0.01845   0.00880  -0.0973   0.8371   0.0830
  -2.500   0.1669   0.01791   0.00822  -0.0972   0.8285   0.0886
  -2.250   0.1955   0.01746   0.00776  -0.0972   0.8200   0.0979
  -2.000   0.2226   0.01704   0.00739  -0.0968   0.8110   0.1087
  -1.750   0.2517   0.01663   0.00704  -0.0969   0.8031   0.1302
  -1.500   0.2779   0.01630   0.00690  -0.0964   0.7937   0.1700
  -1.250   0.3076   0.01600   0.00668  -0.0965   0.7861   0.2217
  -1.000   0.3333   0.01581   0.00660  -0.0960   0.7761   0.2635
  -0.750   0.3630   0.01554   0.00641  -0.0962   0.7685   0.3114
  -0.500   0.3883   0.01526   0.00635  -0.0957   0.7585   0.3778
  -0.250   0.4103   0.01436   0.00636  -0.0942   0.7502   0.6249
   0.000   0.4755   0.01366   0.00618  -0.1008   0.7419   1.0000
   0.250   0.5011   0.01376   0.00611  -0.1002   0.7324   1.0000
   0.500   0.5280   0.01383   0.00602  -0.0997   0.7234   1.0000
   0.750   0.5524   0.01397   0.00604  -0.0989   0.7129   1.0000
   1.000   0.5789   0.01407   0.00600  -0.0983   0.7038   1.0000
   1.250   0.6042   0.01419   0.00600  -0.0976   0.6930   1.0000
   1.500   0.6287   0.01431   0.00602  -0.0968   0.6806   1.0000
   1.750   0.6536   0.01442   0.00601  -0.0959   0.6672   1.0000
   2.000   0.6782   0.01452   0.00600  -0.0949   0.6525   1.0000
   2.250   0.7025   0.01464   0.00600  -0.0940   0.6374   1.0000
   2.500   0.7271   0.01478   0.00604  -0.0931   0.6236   1.0000
   2.750   0.7520   0.01494   0.00613  -0.0923   0.6114   1.0000
   3.000   0.7759   0.01513   0.00628  -0.0914   0.5985   1.0000
   3.250   0.8000   0.01532   0.00643  -0.0906   0.5857   1.0000
   3.500   0.8241   0.01552   0.00659  -0.0897   0.5731   1.0000
   3.750   0.8482   0.01572   0.00674  -0.0888   0.5599   1.0000
   4.000   0.8719   0.01593   0.00691  -0.0879   0.5464   1.0000
   4.250   0.8952   0.01616   0.00712  -0.0869   0.5322   1.0000
   4.500   0.9182   0.01640   0.00734  -0.0859   0.5179   1.0000
   4.750   0.9411   0.01666   0.00758  -0.0849   0.5033   1.0000
   5.000   0.9637   0.01693   0.00783  -0.0839   0.4884   1.0000
   5.250   0.9859   0.01723   0.00810  -0.0828   0.4731   1.0000
   5.500   1.0078   0.01756   0.00839  -0.0816   0.4577   1.0000
   5.750   1.0295   0.01792   0.00873  -0.0805   0.4426   1.0000
   6.000   1.0509   0.01830   0.00910  -0.0793   0.4278   1.0000
   6.250   1.0717   0.01871   0.00949  -0.0780   0.4126   1.0000
   6.500   1.0921   0.01915   0.00992  -0.0767   0.3971   1.0000
   6.750   1.1119   0.01961   0.01039  -0.0754   0.3815   1.0000
   7.000   1.1315   0.02009   0.01088  -0.0740   0.3667   1.0000
   7.250   1.1509   0.02059   0.01139  -0.0726   0.3528   1.0000
   7.500   1.1699   0.02112   0.01192  -0.0712   0.3398   1.0000
   7.750   1.1885   0.02168   0.01249  -0.0697   0.3276   1.0000
   8.000   1.2068   0.02224   0.01309  -0.0683   0.3154   1.0000
   8.250   1.2246   0.02284   0.01372  -0.0667   0.3041   1.0000
   8.500   1.2414   0.02346   0.01435  -0.0651   0.2933   1.0000
   8.750   1.2571   0.02411   0.01499  -0.0633   0.2830   1.0000
   9.000   1.2730   0.02476   0.01574  -0.0616   0.2727   1.0000
   9.250   1.2881   0.02546   0.01645  -0.0598   0.2648   1.0000
   9.500   1.3027   0.02615   0.01724  -0.0579   0.2564   1.0000
   9.750   1.3153   0.02690   0.01804  -0.0557   0.2476   1.0000
  10.000   1.3250   0.02770   0.01887  -0.0533   0.2373   1.0000
  10.250   1.3334   0.02854   0.01983  -0.0509   0.2258   1.0000
  10.500   1.3395   0.02949   0.02083  -0.0483   0.2135   1.0000
  10.750   1.3437   0.03056   0.02194  -0.0457   0.2002   1.0000
  11.000   1.3477   0.03173   0.02317  -0.0433   0.1867   1.0000
  11.250   1.3522   0.03300   0.02452  -0.0412   0.1731   1.0000
  11.500   1.3571   0.03435   0.02597  -0.0393   0.1592   1.0000
  11.750   1.3614   0.03587   0.02759  -0.0376   0.1424   1.0000
  12.000   1.3610   0.03781   0.02955  -0.0359   0.1219   1.0000
  12.250   1.3557   0.04034   0.03196  -0.0342   0.1001   1.0000
  12.500   1.3472   0.04336   0.03488  -0.0328   0.0844   1.0000
  12.750   1.3384   0.04666   0.03814  -0.0316   0.0711   1.0000
  13.000   1.3299   0.05012   0.04164  -0.0309   0.0593   1.0000
  13.250   1.3211   0.05380   0.04538  -0.0305   0.0515   1.0000
  13.500   1.3118   0.05772   0.04939  -0.0304   0.0466   1.0000
  13.750   1.3009   0.06206   0.05381  -0.0307   0.0430   1.0000
  14.000   1.2882   0.06685   0.05870  -0.0315   0.0405   1.0000
  14.250   1.2777   0.07161   0.06361  -0.0325   0.0384   1.0000
  14.500   1.2660   0.07676   0.06891  -0.0338   0.0368   1.0000
  14.750   1.2525   0.08242   0.07470  -0.0357   0.0355   1.0000
  15.000   1.2387   0.08837   0.08078  -0.0378   0.0346   1.0000
<< Back to GOE 682 AIRFOIL (goe682-il)

Polar data table (+)

Polar graphs


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