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GOE 5K AIRFOIL (goe05k-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 5K AIRFOIL (goe05k-il)
Reynolds number: 200,000
Max Cl/Cd: 33.45 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe05k-il-200000.txt
Download as CSV file: xf-goe05k-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 5K AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.5200   0.09207   0.08857  -0.0115   1.0000   0.0159
  -7.500  -0.5208   0.08928   0.08582  -0.0122   1.0000   0.0161
  -7.250  -0.5193   0.08618   0.08276  -0.0139   1.0000   0.0165
  -7.000  -0.5132   0.08289   0.07950  -0.0171   1.0000   0.0170
  -6.750  -0.5035   0.07947   0.07608  -0.0212   1.0000   0.0174
  -6.500  -0.4897   0.07610   0.07267  -0.0254   1.0000   0.0177
  -6.250  -0.4744   0.07276   0.06925  -0.0287   1.0000   0.0179
  -6.000  -0.4593   0.06938   0.06578  -0.0307   1.0000   0.0181
  -5.750  -0.4527   0.06383   0.06020  -0.0319   1.0000   0.0185
  -5.500  -0.4521   0.05909   0.05551  -0.0306   1.0000   0.0196
  -5.250  -0.4416   0.05548   0.05184  -0.0307   1.0000   0.0209
  -5.000  -0.4266   0.05189   0.04815  -0.0314   1.0000   0.0223
  -4.750  -0.4091   0.04835   0.04445  -0.0321   1.0000   0.0240
  -4.500  -0.3889   0.04482   0.04072  -0.0327   1.0000   0.0262
  -4.250  -0.3581   0.04314   0.03849  -0.0327   1.0000   0.0283
  -4.000  -0.3461   0.03795   0.03330  -0.0325   1.0000   0.0295
  -3.750  -0.3301   0.03498   0.03024  -0.0320   1.0000   0.0331
  -3.500  -0.3068   0.03215   0.02700  -0.0312   1.0000   0.0381
  -3.250  -0.2858   0.02981   0.02435  -0.0303   1.0000   0.0463
  -3.000  -0.2640   0.02812   0.02225  -0.0293   1.0000   0.0571
  -2.750  -0.2436   0.02527   0.01923  -0.0287   1.0000   0.0703
  -2.500  -0.2228   0.02277   0.01664  -0.0281   1.0000   0.0871
  -1.750  -0.1363   0.01653   0.00899  -0.0232   1.0000   0.0350
  -1.500  -0.1085   0.01486   0.00705  -0.0219   1.0000   0.0269
  -1.250  -0.0831   0.01364   0.00568  -0.0209   1.0000   0.0274
  -1.000  -0.0585   0.01240   0.00438  -0.0198   1.0000   0.0277
  -0.750  -0.0349   0.01155   0.00349  -0.0188   1.0000   0.0369
  -0.500   0.0155   0.00819   0.00272  -0.0235   1.0000   1.0000
  -0.250   0.0379   0.00829   0.00261  -0.0226   1.0000   1.0000
   0.000   0.0601   0.00841   0.00258  -0.0217   1.0000   1.0000
   0.250   0.0824   0.00854   0.00261  -0.0209   1.0000   1.0000
   0.500   0.1045   0.00868   0.00267  -0.0201   1.0000   1.0000
   0.750   0.1267   0.00883   0.00278  -0.0193   1.0000   1.0000
   1.000   0.1486   0.00901   0.00295  -0.0185   1.0000   1.0000
   1.250   0.1705   0.00920   0.00316  -0.0178   1.0000   1.0000
   1.500   0.1920   0.00941   0.00341  -0.0170   1.0000   1.0000
   1.750   0.2137   0.00964   0.00371  -0.0163   1.0000   1.0000
   2.000   0.2350   0.00989   0.00407  -0.0155   1.0000   1.0000
   2.250   0.2981   0.01025   0.00479  -0.0234   0.9882   1.0000
   2.500   0.4499   0.01345   0.00490  -0.0441   0.0222   1.0000
   2.750   0.4716   0.01488   0.00641  -0.0422   0.0232   1.0000
   3.000   0.4974   0.01688   0.00851  -0.0406   0.0270   1.0000
   3.250   0.5266   0.01909   0.01096  -0.0393   0.0347   1.0000
   5.250   0.7032   0.04163   0.03644  -0.0244   0.0350   1.0000
   5.500   0.7162   0.04537   0.04021  -0.0235   0.0312   1.0000
   5.750   0.7280   0.04972   0.04491  -0.0217   0.0298   1.0000
   6.000   0.7460   0.05270   0.04834  -0.0199   0.0266   1.0000
   6.250   0.7576   0.05663   0.05249  -0.0190   0.0244   1.0000
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