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

GOE 286 AIRFOIL (goe286-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 286 AIRFOIL (goe286-il)
Reynolds number: 500,000
Max Cl/Cd: 78.82 at α=6.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe286-il-500000-n5.txt
Download as CSV file: xf-goe286-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 286 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4055   0.08747   0.08532  -0.0171   1.0000   0.0141
  -8.250  -0.4042   0.08326   0.08114  -0.0200   1.0000   0.0147
  -8.000  -0.4072   0.07371   0.07159  -0.0315   0.9849   0.0159
  -7.500  -0.3716   0.06726   0.06508  -0.0400   0.9594   0.0164
  -7.250  -0.3574   0.06352   0.06129  -0.0437   0.9439   0.0167
  -7.000  -0.3425   0.05907   0.05675  -0.0480   0.9266   0.0172
  -6.750  -0.3277   0.04219   0.03940  -0.0618   0.9137   0.0197
  -6.500  -0.3094   0.03948   0.03652  -0.0624   0.8895   0.0200
  -6.250  -0.2905   0.03765   0.03448  -0.0622   0.8595   0.0203
  -6.000  -0.2696   0.03573   0.03230  -0.0623   0.8303   0.0207
  -5.750  -0.2470   0.03312   0.02941  -0.0628   0.8056   0.0215
  -5.500  -0.2257   0.02456   0.02006  -0.0643   0.7891   0.0230
  -5.250  -0.2013   0.02094   0.01570  -0.0643   0.7641   0.0241
  -5.000  -0.1757   0.01933   0.01362  -0.0642   0.7271   0.0245
  -4.500  -0.1236   0.01679   0.01038  -0.0641   0.6596   0.0252
  -4.250  -0.0963   0.01605   0.00941  -0.0641   0.6419   0.0255
  -4.000  -0.0686   0.01538   0.00855  -0.0641   0.6294   0.0258
  -3.750  -0.0407   0.01481   0.00783  -0.0642   0.6186   0.0262
  -3.500  -0.0127   0.01437   0.00724  -0.0642   0.6082   0.0268
  -3.250   0.0154   0.01385   0.00659  -0.0642   0.5963   0.0272
  -3.000   0.0436   0.01334   0.00594  -0.0642   0.5852   0.0275
  -2.750   0.0719   0.01290   0.00539  -0.0642   0.5746   0.0279
  -2.500   0.1002   0.01251   0.00489  -0.0642   0.5628   0.0283
  -2.250   0.1285   0.01216   0.00446  -0.0642   0.5507   0.0287
  -2.000   0.1568   0.01186   0.00408  -0.0642   0.5354   0.0290
  -1.750   0.1850   0.01163   0.00373  -0.0642   0.5132   0.0293
  -1.500   0.2130   0.01149   0.00343  -0.0642   0.4782   0.0296
  -1.250   0.2408   0.01143   0.00319  -0.0642   0.4439   0.0299
  -1.000   0.2687   0.01141   0.00303  -0.0643   0.4179   0.0303
  -0.750   0.2967   0.01132   0.00282  -0.0643   0.3966   0.0309
  -0.500   0.3247   0.01119   0.00260  -0.0644   0.3783   0.0315
  -0.250   0.3529   0.01112   0.00246  -0.0644   0.3632   0.0321
   0.000   0.3812   0.01107   0.00235  -0.0645   0.3509   0.0329
   0.250   0.4095   0.01107   0.00228  -0.0645   0.3383   0.0338
   0.500   0.4377   0.01109   0.00223  -0.0646   0.3255   0.0348
   1.000   0.4939   0.01123   0.00220  -0.0647   0.2935   0.0372
   1.500   0.5492   0.01147   0.00229  -0.0648   0.2477   0.0558
   1.750   0.5765   0.01162   0.00241  -0.0648   0.2171   0.1066
   2.000   0.6014   0.01206   0.00277  -0.0649   0.1335   0.2244
   2.500   0.6533   0.01061   0.00311  -0.0644   0.1197   1.0000
   2.750   0.6809   0.01079   0.00324  -0.0644   0.1175   1.0000
   3.000   0.7084   0.01098   0.00338  -0.0644   0.1150   1.0000
   3.250   0.7358   0.01118   0.00353  -0.0643   0.1125   1.0000
   3.500   0.7632   0.01138   0.00370  -0.0643   0.1104   1.0000
   3.750   0.7906   0.01157   0.00387  -0.0642   0.1088   1.0000
   4.000   0.8179   0.01175   0.00403  -0.0642   0.1081   1.0000
   4.250   0.8453   0.01192   0.00420  -0.0641   0.1073   1.0000
   4.500   0.8725   0.01210   0.00438  -0.0641   0.1063   1.0000
   4.750   0.8996   0.01229   0.00457  -0.0640   0.1051   1.0000
   5.000   0.9266   0.01250   0.00478  -0.0639   0.1037   1.0000
   5.250   0.9534   0.01272   0.00500  -0.0639   0.1022   1.0000
   5.500   0.9801   0.01295   0.00523  -0.0638   0.1004   1.0000
   5.750   1.0065   0.01321   0.00549  -0.0636   0.0986   1.0000
   6.000   1.0328   0.01348   0.00577  -0.0635   0.0966   1.0000
   6.250   1.0591   0.01373   0.00603  -0.0634   0.0952   1.0000
   6.500   1.0858   0.01391   0.00624  -0.0633   0.0936   1.0000
   6.750   1.1122   0.01411   0.00647  -0.0631   0.0886   1.0000
   7.000   1.1359   0.01468   0.00682  -0.0628   0.0592   1.0000
   7.250   1.1603   0.01513   0.00727  -0.0624   0.0560   1.0000
   7.500   1.1846   0.01556   0.00773  -0.0620   0.0535   1.0000
   7.750   1.2089   0.01598   0.00818  -0.0616   0.0513   1.0000
   8.000   1.2334   0.01634   0.00859  -0.0613   0.0500   1.0000
   8.250   1.2577   0.01671   0.00901  -0.0609   0.0485   1.0000
   8.500   1.2813   0.01713   0.00949  -0.0605   0.0462   1.0000
   8.750   1.3041   0.01762   0.01002  -0.0599   0.0441   1.0000
   9.000   1.3263   0.01814   0.01058  -0.0593   0.0418   1.0000
   9.250   1.3500   0.01846   0.01096  -0.0589   0.0390   1.0000
   9.500   1.3722   0.01893   0.01145  -0.0583   0.0337   1.0000
   9.750   1.3889   0.01989   0.01225  -0.0570   0.0164   1.0000
  10.000   1.4074   0.02063   0.01304  -0.0559   0.0141   1.0000
  10.250   1.4249   0.02141   0.01388  -0.0547   0.0128   1.0000
  10.500   1.4402   0.02229   0.01484  -0.0532   0.0117   1.0000
  10.750   1.4552   0.02312   0.01577  -0.0517   0.0109   1.0000
  11.000   1.4690   0.02392   0.01665  -0.0500   0.0104   1.0000
  11.250   1.4789   0.02481   0.01763  -0.0477   0.0099   1.0000
  11.500   1.4876   0.02585   0.01876  -0.0456   0.0095   1.0000
  11.750   1.4952   0.02706   0.02006  -0.0436   0.0091   1.0000
  12.000   1.5011   0.02848   0.02158  -0.0417   0.0088   1.0000
  12.250   1.5044   0.03023   0.02345  -0.0399   0.0085   1.0000
  12.500   1.5049   0.03240   0.02573  -0.0385   0.0082   1.0000
  12.750   1.5096   0.03434   0.02778  -0.0376   0.0080   1.0000
  13.000   1.5129   0.03656   0.03012  -0.0369   0.0078   1.0000
  13.250   1.5146   0.03909   0.03276  -0.0366   0.0076   1.0000
  13.500   1.5150   0.04193   0.03573  -0.0366   0.0074   1.0000
  13.750   1.5138   0.04516   0.03907  -0.0369   0.0072   1.0000
  14.000   1.5111   0.04872   0.04275  -0.0375   0.0070   1.0000
  14.250   1.5065   0.05259   0.04675  -0.0383   0.0069   1.0000
  14.500   1.5000   0.05674   0.05103  -0.0391   0.0068   1.0000
  14.750   1.4917   0.06119   0.05560  -0.0401   0.0067   1.0000
  15.000   1.4818   0.06595   0.06048  -0.0413   0.0066   1.0000
  15.250   1.4711   0.07095   0.06561  -0.0427   0.0065   1.0000
  15.500   1.4596   0.07625   0.07103  -0.0443   0.0064   1.0000
  15.750   1.4476   0.08172   0.07661  -0.0460   0.0064   1.0000
  16.000   1.4352   0.08732   0.08233  -0.0478   0.0063   1.0000
  16.250   1.4228   0.09305   0.08817  -0.0497   0.0063   1.0000
  16.500   1.4103   0.09884   0.09406  -0.0518   0.0062   1.0000
<< Back to GOE 286 AIRFOIL (goe286-il)

Polar data table (+)

Polar graphs


<< Back to GOE 286 AIRFOIL (goe286-il)