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

GOE 506 AIRFOIL (goe506-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 506 AIRFOIL (goe506-il)
Reynolds number: 50,000
Max Cl/Cd: 27.98 at α=7.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe506-il-50000-n5.txt
Download as CSV file: xf-goe506-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 506 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3957   0.10724   0.10110  -0.0509   0.9875   0.0583
  -9.750  -0.4056   0.10056   0.09446  -0.0551   0.9840   0.0580
  -9.500  -0.4256   0.09260   0.08655  -0.0598   0.9791   0.0571
  -9.250  -0.4597   0.08239   0.07627  -0.0681   0.9738   0.0557
  -9.000  -0.5108   0.07584   0.06958  -0.0698   0.9662   0.0542
  -8.750  -0.5425   0.07022   0.06363  -0.0704   0.9606   0.0534
  -8.500  -0.5565   0.06665   0.05983  -0.0689   0.9550   0.0538
  -8.250  -0.5616   0.06288   0.05572  -0.0683   0.9505   0.0541
  -8.000  -0.5622   0.05961   0.05210  -0.0672   0.9469   0.0547
  -7.750  -0.5638   0.05684   0.04900  -0.0648   0.9429   0.0551
  -7.500  -0.5575   0.05424   0.04604  -0.0634   0.9392   0.0560
  -7.250  -0.5452   0.05188   0.04328  -0.0626   0.9359   0.0577
  -7.000  -0.5287   0.04969   0.04056  -0.0621   0.9332   0.0599
  -6.750  -0.5207   0.04807   0.03849  -0.0595   0.9298   0.0612
  -6.500  -0.5060   0.04637   0.03673  -0.0581   0.9266   0.0627
  -6.250  -0.4876   0.04501   0.03524  -0.0572   0.9233   0.0646
  -6.000  -0.4664   0.04383   0.03388  -0.0566   0.9205   0.0677
  -5.750  -0.4421   0.04279   0.03254  -0.0564   0.9180   0.0714
  -5.500  -0.4272   0.04186   0.03143  -0.0544   0.9140   0.0734
  -5.250  -0.4106   0.04085   0.03039  -0.0530   0.9098   0.0761
  -5.000  -0.3868   0.04001   0.02944  -0.0527   0.9054   0.0813
  -4.750  -0.3581   0.03924   0.02847  -0.0531   0.9001   0.0872
  -4.500  -0.3391   0.03836   0.02757  -0.0518   0.8923   0.0923
  -4.250  -0.3063   0.03754   0.02671  -0.0530   0.8869   0.1035
  -4.000  -0.2867   0.03672   0.02615  -0.0521   0.8798   0.1252
  -3.750  -0.2589   0.03637   0.02614  -0.0524   0.8739   0.2053
  -3.500  -0.2313   0.03649   0.02620  -0.0525   0.8681   0.2613
  -3.250  -0.2123   0.03659   0.02622  -0.0511   0.8601   0.2949
  -3.000  -0.1803   0.03674   0.02629  -0.0519   0.8538   0.3319
  -2.750  -0.1597   0.03666   0.02620  -0.0506   0.8433   0.3565
  -2.500  -0.1319   0.03648   0.02590  -0.0504   0.8326   0.3814
  -2.250  -0.0895   0.03604   0.02537  -0.0524   0.8235   0.4055
  -2.000  -0.0596   0.03558   0.02477  -0.0525   0.8113   0.4195
  -1.750  -0.0304   0.03514   0.02422  -0.0525   0.7995   0.4319
  -1.500   0.0049   0.03467   0.02369  -0.0535   0.7909   0.4486
  -1.250   0.0350   0.03426   0.02328  -0.0537   0.7814   0.4693
  -1.000   0.0591   0.03390   0.02300  -0.0529   0.7706   0.4929
  -0.750   0.0971   0.03320   0.02244  -0.0541   0.7638   0.5328
  -0.500   0.1166   0.03280   0.02228  -0.0524   0.7515   0.5832
  -0.250   0.1437   0.03215   0.02213  -0.0515   0.7407   0.6851
   0.000   0.2790   0.03170   0.02201  -0.0710   0.7349   1.0000
   0.250   0.2974   0.03154   0.02163  -0.0693   0.7214   1.0000
   0.500   0.3183   0.03132   0.02121  -0.0679   0.7083   1.0000
   0.750   0.3519   0.03071   0.02038  -0.0681   0.6997   1.0000
   1.000   0.3698   0.03055   0.02007  -0.0662   0.6846   1.0000
   1.250   0.3879   0.03041   0.01978  -0.0642   0.6688   1.0000
   1.750   0.4175   0.03050   0.01962  -0.0597   0.6306   1.0000
   2.000   0.4390   0.03037   0.01935  -0.0584   0.6136   1.0000
   2.250   0.4646   0.03012   0.01895  -0.0575   0.5981   1.0000
   2.500   0.4938   0.02976   0.01843  -0.0571   0.5840   1.0000
   2.750   0.5266   0.02933   0.01780  -0.0573   0.5718   1.0000
   3.000   0.5631   0.02880   0.01704  -0.0578   0.5605   1.0000
   3.250   0.5942   0.02863   0.01667  -0.0579   0.5474   1.0000
   3.500   0.6244   0.02860   0.01644  -0.0580   0.5352   1.0000
   3.750   0.6564   0.02859   0.01622  -0.0584   0.5244   1.0000
   4.000   0.6877   0.02871   0.01615  -0.0588   0.5142   1.0000
   4.250   0.7129   0.02910   0.01643  -0.0585   0.5042   1.0000
   4.500   0.7475   0.02925   0.01638  -0.0594   0.4958   1.0000
   4.750   0.7667   0.02988   0.01699  -0.0584   0.4868   1.0000
   5.000   0.8006   0.03013   0.01710  -0.0594   0.4793   1.0000
   5.250   0.8171   0.03084   0.01782  -0.0580   0.4712   1.0000
   5.500   0.8464   0.03124   0.01815  -0.0584   0.4644   1.0000
   5.750   0.8655   0.03189   0.01880  -0.0574   0.4569   1.0000
   6.000   0.8877   0.03246   0.01935  -0.0568   0.4499   1.0000
   6.250   0.9139   0.03295   0.01981  -0.0568   0.4436   1.0000
   6.500   0.9277   0.03376   0.02070  -0.0551   0.4373   1.0000
   6.750   0.9553   0.03422   0.02113  -0.0553   0.4317   1.0000
   7.000   0.9720   0.03499   0.02197  -0.0541   0.4260   1.0000
   7.250   0.9858   0.03584   0.02290  -0.0525   0.4203   1.0000
   7.500   1.0143   0.03625   0.02328  -0.0528   0.4148   1.0000
   7.750   1.0229   0.03729   0.02443  -0.0505   0.4094   1.0000
   8.000   1.0347   0.03824   0.02549  -0.0487   0.4039   1.0000
   8.250   1.0627   0.03868   0.02594  -0.0490   0.3991   1.0000
   8.500   1.0678   0.03990   0.02729  -0.0464   0.3938   1.0000
   8.750   1.0744   0.04105   0.02855  -0.0441   0.3879   1.0000
   9.000   1.1031   0.04140   0.02892  -0.0443   0.3829   1.0000
   9.250   1.1013   0.04300   0.03069  -0.0413   0.3777   1.0000
   9.500   1.1027   0.04458   0.03242  -0.0387   0.3725   1.0000
   9.750   1.1257   0.04519   0.03310  -0.0384   0.3680   1.0000
  10.000   1.1350   0.04649   0.03450  -0.0367   0.3633   1.0000
  10.250   1.1146   0.04936   0.03761  -0.0326   0.3575   1.0000
  10.500   1.1297   0.05036   0.03870  -0.0316   0.3525   1.0000
  10.750   1.1536   0.05088   0.03927  -0.0313   0.3478   1.0000
  11.000   1.0961   0.05667   0.04533  -0.0256   0.3404   1.0000
  11.250   1.1137   0.05747   0.04621  -0.0248   0.3352   1.0000
  11.500   1.1160   0.05951   0.04835  -0.0233   0.3297   1.0000
  11.750   1.0447   0.06829   0.05732  -0.0200   0.3192   1.0000
  12.000   1.0877   0.06659   0.05568  -0.0197   0.3162   1.0000
  12.250   0.9886   0.08027   0.06948  -0.0187   0.3009   1.0000
  12.500   1.0202   0.07934   0.06864  -0.0179   0.2988   1.0000
<< Back to GOE 506 AIRFOIL (goe506-il)

Polar data table (+)

Polar graphs


<< Back to GOE 506 AIRFOIL (goe506-il)