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

GOE 10K AIRFOIL (goe10k-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 10K AIRFOIL (goe10k-il)
Reynolds number: 100,000
Max Cl/Cd: 33.1 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe10k-il-100000-n5.txt
Download as CSV file: xf-goe10k-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 10K 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.5364   0.09961   0.09462  -0.0122   1.0000   0.0362
  -8.000  -0.5388   0.09687   0.09197  -0.0142   1.0000   0.0365
  -7.750  -0.5407   0.09380   0.08897  -0.0172   1.0000   0.0366
  -7.500  -0.5371   0.09024   0.08543  -0.0212   1.0000   0.0368
  -7.250  -0.5334   0.08573   0.08098  -0.0194   1.0000   0.0374
  -7.000  -0.5283   0.08211   0.07738  -0.0186   1.0000   0.0385
  -6.750  -0.5225   0.07847   0.07375  -0.0200   1.0000   0.0399
  -6.500  -0.5147   0.07469   0.06994  -0.0224   1.0000   0.0413
  -6.250  -0.4982   0.07089   0.06600  -0.0293   1.0000   0.0442
  -6.000  -0.4846   0.06676   0.06165  -0.0325   1.0000   0.0448
  -5.750  -0.4784   0.06192   0.05690  -0.0312   1.0000   0.0457
  -5.500  -0.4674   0.05806   0.05299  -0.0311   1.0000   0.0466
  -5.250  -0.4538   0.05422   0.04906  -0.0315   1.0000   0.0478
  -5.000  -0.4316   0.04852   0.04304  -0.0323   1.0000   0.0261
  -4.750  -0.4148   0.04455   0.03887  -0.0324   1.0000   0.0245
  -4.500  -0.3956   0.04053   0.03457  -0.0323   1.0000   0.0227
  -4.250  -0.3747   0.03648   0.03014  -0.0319   1.0000   0.0212
  -4.000  -0.3509   0.03217   0.02524  -0.0308   1.0000   0.0190
  -3.750  -0.3272   0.02964   0.02223  -0.0294   1.0000   0.0180
  -3.500  -0.3053   0.02664   0.01874  -0.0283   1.0000   0.0176
  -3.250  -0.2822   0.02383   0.01542  -0.0271   1.0000   0.0175
  -3.000  -0.2580   0.02132   0.01241  -0.0260   1.0000   0.0182
  -2.750  -0.2342   0.01928   0.01006  -0.0251   1.0000   0.0211
  -2.500  -0.2095   0.01759   0.00810  -0.0239   1.0000   0.0228
  -2.250  -0.1857   0.01633   0.00658  -0.0226   1.0000   0.0270
  -2.000  -0.1626   0.01526   0.00542  -0.0215   1.0000   0.0347
  -1.750  -0.1392   0.01446   0.00458  -0.0204   1.0000   0.0575
  -1.500  -0.1151   0.01316   0.00395  -0.0201   1.0000   0.2029
  -1.250  -0.0586   0.01076   0.00352  -0.0257   1.0000   1.0000
  -1.000  -0.0362   0.01080   0.00327  -0.0246   1.0000   1.0000
  -0.750  -0.0139   0.01086   0.00309  -0.0236   1.0000   1.0000
  -0.500   0.0085   0.01093   0.00293  -0.0227   1.0000   1.0000
  -0.250   0.0307   0.01101   0.00287  -0.0217   1.0000   1.0000
   0.000   0.0529   0.01111   0.00286  -0.0208   1.0000   1.0000
   0.250   0.0751   0.01122   0.00287  -0.0198   1.0000   1.0000
   0.500   0.0973   0.01135   0.00294  -0.0190   1.0000   1.0000
   0.750   0.1194   0.01149   0.00305  -0.0181   1.0000   1.0000
   1.000   0.1415   0.01164   0.00321  -0.0172   1.0000   1.0000
   1.250   0.1634   0.01181   0.00341  -0.0164   1.0000   1.0000
   1.500   0.1853   0.01200   0.00364  -0.0155   1.0000   1.0000
   1.750   0.2071   0.01220   0.00393  -0.0147   1.0000   1.0000
   2.000   0.2487   0.01248   0.00442  -0.0181   0.9918   1.0000
   2.250   0.3153   0.01258   0.00485  -0.0262   0.9693   1.0000
   2.500   0.3919   0.01189   0.00471  -0.0350   0.9201   1.0000
   2.750   0.4704   0.01421   0.00423  -0.0428   0.1335   1.0000
   3.000   0.4896   0.01578   0.00530  -0.0410   0.0393   1.0000
   3.250   0.5112   0.01684   0.00643  -0.0394   0.0269   1.0000
   3.500   0.5323   0.01809   0.00788  -0.0376   0.0227   1.0000
   3.750   0.5530   0.02010   0.00999  -0.0360   0.0188   1.0000
   4.000   0.5782   0.02205   0.01219  -0.0348   0.0178   1.0000
   4.250   0.6042   0.02441   0.01491  -0.0336   0.0171   1.0000
   4.500   0.6289   0.02672   0.01764  -0.0322   0.0158   1.0000
   4.750   0.6514   0.02916   0.02048  -0.0307   0.0146   1.0000
   5.000   0.6727   0.03256   0.02450  -0.0286   0.0150   1.0000
   5.250   0.6915   0.03626   0.02873  -0.0263   0.0157   1.0000
   5.500   0.7078   0.04012   0.03305  -0.0242   0.0166   1.0000
   5.750   0.7213   0.04417   0.03748  -0.0222   0.0176   1.0000
   6.250   0.7513   0.05331   0.04749  -0.0182   0.0250   1.0000
   6.500   0.7660   0.06013   0.05472  -0.0166   0.0515   1.0000
   6.750   0.7740   0.06437   0.05918  -0.0160   0.0512   1.0000
   7.000   0.7803   0.06851   0.06349  -0.0158   0.0500   1.0000
   7.250   0.7848   0.07277   0.06786  -0.0155   0.0487   1.0000
   7.500   0.7892   0.07691   0.07206  -0.0151   0.0475   1.0000
   7.750   0.7969   0.08394   0.07887  -0.0143   0.0459   1.0000
   8.000   0.7922   0.08725   0.08241  -0.0149   0.0457   1.0000
   8.250   0.7839   0.09078   0.08609  -0.0172   0.0452   1.0000
   8.500   0.7754   0.09472   0.09005  -0.0192   0.0449   1.0000
<< Back to GOE 10K AIRFOIL (goe10k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 10K AIRFOIL (goe10k-il)