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

GOE 510 AIRFOIL (goe510-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 510 AIRFOIL (goe510-il)
Reynolds number: 50,000
Max Cl/Cd: 29.53 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe510-il-50000-n5.txt
Download as CSV file: xf-goe510-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 510 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
  ------ -------- --------- --------- -------- -------- --------
  -7.000  -0.2442   0.10573   0.09936  -0.0422   0.9601   0.1620
  -6.750  -0.2643   0.10454   0.09819  -0.0464   0.9467   0.1653
  -6.500  -0.2680   0.10205   0.09566  -0.0504   0.9347   0.1660
  -6.000  -0.2509   0.08705   0.08033  -0.0593   0.9172   0.1034
  -5.750  -0.2297   0.08374   0.07701  -0.0603   0.9113   0.1014
  -5.500  -0.2207   0.08046   0.07368  -0.0612   0.9004   0.1005
  -5.250  -0.2095   0.07707   0.07019  -0.0625   0.8895   0.0997
  -4.750  -0.1776   0.06954   0.06236  -0.0662   0.8692   0.0967
  -4.250  -0.1459   0.05850   0.05038  -0.0708   0.8506   0.0904
  -4.000  -0.1342   0.05577   0.04743  -0.0696   0.8392   0.0903
  -3.750  -0.1049   0.05336   0.04492  -0.0710   0.8335   0.0918
  -3.500  -0.0928   0.05186   0.04332  -0.0692   0.8218   0.0930
  -3.250  -0.0616   0.04907   0.04022  -0.0707   0.8166   0.0943
  -3.000  -0.0502   0.04699   0.03783  -0.0686   0.8046   0.0943
  -2.750  -0.0173   0.04402   0.03442  -0.0699   0.7997   0.0946
  -2.500  -0.0037   0.04249   0.03258  -0.0677   0.7876   0.0950
  -2.250   0.0324   0.04034   0.03004  -0.0690   0.7828   0.0973
  -2.000   0.0479   0.03931   0.02873  -0.0670   0.7705   0.0997
  -1.750   0.0868   0.03726   0.02619  -0.0683   0.7659   0.1023
  -1.500   0.1053   0.03643   0.02504  -0.0666   0.7536   0.1034
  -1.250   0.1471   0.03484   0.02327  -0.0684   0.7488   0.1056
  -1.000   0.1666   0.03441   0.02278  -0.0668   0.7362   0.1082
  -0.750   0.2030   0.03356   0.02175  -0.0679   0.7283   0.1134
  -0.500   0.2423   0.03279   0.02073  -0.0694   0.7190   0.1188
  -0.250   0.2791   0.03214   0.02005  -0.0706   0.7098   0.1237
   0.000   0.3189   0.03139   0.01918  -0.0720   0.7013   0.1317
   0.250   0.3473   0.03104   0.01887  -0.0718   0.6904   0.1420
   0.500   0.3884   0.03022   0.01808  -0.0734   0.6831   0.1615
   0.750   0.4108   0.03002   0.01801  -0.0721   0.6710   0.1859
   1.000   0.4519   0.02934   0.01741  -0.0736   0.6640   0.2528
   1.250   0.4705   0.02944   0.01750  -0.0718   0.6515   0.2953
   1.500   0.4952   0.02928   0.01736  -0.0711   0.6410   0.3415
   1.750   0.5270   0.02861   0.01681  -0.0714   0.6322   0.4015
   2.000   0.5406   0.02818   0.01677  -0.0690   0.6209   0.4650
   2.500   0.7305   0.02755   0.01656  -0.0938   0.5986   1.0000
   2.750   0.7477   0.02794   0.01679  -0.0919   0.5879   1.0000
   3.000   0.7780   0.02802   0.01665  -0.0920   0.5791   1.0000
   3.250   0.7874   0.02861   0.01715  -0.0888   0.5683   1.0000
   3.500   0.8190   0.02867   0.01700  -0.0892   0.5604   1.0000
   3.750   0.8249   0.02937   0.01766  -0.0855   0.5503   1.0000
   4.000   0.8541   0.02951   0.01762  -0.0855   0.5429   1.0000
   4.250   0.8599   0.03021   0.01830  -0.0819   0.5333   1.0000
   4.500   0.8904   0.03034   0.01827  -0.0821   0.5264   1.0000
   4.750   0.8921   0.03116   0.01910  -0.0778   0.5175   1.0000
   5.000   0.9238   0.03128   0.01907  -0.0783   0.5108   1.0000
   5.250   0.9231   0.03216   0.01999  -0.0737   0.5025   1.0000
   5.500   0.9489   0.03244   0.02017  -0.0733   0.4958   1.0000
   5.750   0.9565   0.03318   0.02090  -0.0700   0.4885   1.0000
   6.000   0.9715   0.03373   0.02142  -0.0680   0.4814   1.0000
   6.250   0.9987   0.03403   0.02164  -0.0678   0.4754   1.0000
   6.500   0.9950   0.03516   0.02283  -0.0631   0.4680   1.0000
   6.750   1.0235   0.03541   0.02301  -0.0631   0.4620   1.0000
   7.000   1.0267   0.03644   0.02409  -0.0595   0.4553   1.0000
   7.250   1.0355   0.03733   0.02500  -0.0568   0.4485   1.0000
   7.500   1.0732   0.03735   0.02493  -0.0581   0.4434   1.0000
   7.750   1.0546   0.03923   0.02695  -0.0520   0.4362   1.0000
   8.000   1.0699   0.03999   0.02773  -0.0503   0.4301   1.0000
   8.250   1.1063   0.04001   0.02769  -0.0513   0.4251   1.0000
   8.500   1.0732   0.04282   0.03069  -0.0442   0.4176   1.0000
   8.750   1.0926   0.04345   0.03134  -0.0431   0.4120   1.0000
   9.000   1.1221   0.04370   0.03156  -0.0432   0.4072   1.0000
   9.250   1.0717   0.04803   0.03610  -0.0357   0.3987   1.0000
   9.500   1.0978   0.04825   0.03635  -0.0352   0.3939   1.0000
   9.750   1.0821   0.05121   0.03941  -0.0317   0.3873   1.0000
  10.000   1.0426   0.05630   0.04463  -0.0276   0.3783   1.0000
  10.250   1.0839   0.05515   0.04348  -0.0277   0.3752   1.0000
  10.500   0.9853   0.06684   0.05534  -0.0237   0.3605   1.0000
  10.750   1.0209   0.06560   0.05413  -0.0228   0.3581   1.0000
  11.250   0.9757   0.07642   0.06510  -0.0214   0.3408   1.0000
  11.750   0.9376   0.08769   0.07650  -0.0216   0.3230   1.0000
  12.250   0.9252   0.09574   0.08468  -0.0217   0.3084   1.0000
  12.750   0.9059   0.10495   0.09400  -0.0225   0.2933   1.0000
  13.250   0.8926   0.11337   0.10253  -0.0236   0.2789   1.0000
  13.500   0.9176   0.11288   0.10213  -0.0224   0.2765   1.0000
  13.750   0.8841   0.12125   0.11053  -0.0248   0.2654   1.0000
  14.000   0.9026   0.12176   0.11112  -0.0241   0.2621   1.0000
<< Back to GOE 510 AIRFOIL (goe510-il)

Polar data table (+)

Polar graphs


<< Back to GOE 510 AIRFOIL (goe510-il)