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

GOE 281 (DAIMLER XII) AIRFOIL (goe281-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 281 (DAIMLER XII) AIRFOIL (goe281-il)
Reynolds number: 100,000
Max Cl/Cd: 58.98 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe281-il-100000.txt
Download as CSV file: xf-goe281-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 281 (DAIMLER XII) 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3337   0.09715   0.09249  -0.0267   1.0000   0.0544
  -7.500  -0.3402   0.09582   0.09126  -0.0281   1.0000   0.0553
  -7.250  -0.3427   0.09457   0.09009  -0.0326   1.0000   0.0560
  -7.000  -0.3402   0.09297   0.08850  -0.0377   1.0000   0.0564
  -6.750  -0.3397   0.08801   0.08363  -0.0352   1.0000   0.0572
  -6.500  -0.3384   0.08405   0.07974  -0.0292   1.0000   0.0589
  -6.250  -0.3365   0.08140   0.07712  -0.0276   1.0000   0.0607
  -6.000  -0.3342   0.07888   0.07465  -0.0274   1.0000   0.0629
  -5.750  -0.3302   0.07633   0.07212  -0.0285   1.0000   0.0656
  -5.500  -0.3080   0.07475   0.07030  -0.0389   1.0000   0.0699
  -5.250  -0.3050   0.07038   0.06599  -0.0382   1.0000   0.0711
  -5.000  -0.3075   0.06748   0.06324  -0.0335   1.0000   0.0733
  -4.750  -0.3012   0.06517   0.06094  -0.0327   1.0000   0.0777
  -4.500  -0.2725   0.06219   0.05764  -0.0407   1.0000   0.0851
  -4.250  -0.2712   0.05919   0.05480  -0.0377   1.0000   0.0874
  -4.000  -0.2603   0.05696   0.05255  -0.0375   1.0000   0.0927
  -3.750  -0.2168   0.05274   0.04811  -0.0449   0.9950   0.1038
  -3.500  -0.1706   0.04893   0.04411  -0.0518   0.9878   0.1195
  -3.250  -0.1225   0.04546   0.04040  -0.0589   0.9811   0.1460
  -3.000  -0.0800   0.04244   0.03723  -0.0641   0.9732   0.1752
  -2.750  -0.0408   0.03979   0.03454  -0.0679   0.9656   0.2100
  -1.500   0.1966   0.02762   0.02070  -0.0901   0.9306   0.2248
  -1.250   0.2617   0.02555   0.01735  -0.0935   0.9270   0.1256
  -1.000   0.3001   0.02476   0.01610  -0.0945   0.9167   0.1133
  -0.750   0.3495   0.02279   0.01402  -0.0981   0.9127   0.1050
  -0.500   0.3843   0.02227   0.01324  -0.0986   0.9017   0.0991
  -0.250   0.4211   0.02128   0.01222  -0.0998   0.8925   0.0972
   0.000   0.4612   0.02024   0.01120  -0.1014   0.8852   0.0963
   0.250   0.4924   0.01961   0.01061  -0.1014   0.8737   0.0972
   0.500   0.5240   0.01894   0.00999  -0.1016   0.8629   0.1027
   0.750   0.5576   0.01838   0.00938  -0.1018   0.8528   0.1074
   1.250   0.6200   0.01739   0.00833  -0.1010   0.8270   0.1339
   1.500   0.6474   0.01519   0.00798  -0.0997   0.8116   1.0000
   1.750   0.6728   0.01513   0.00770  -0.0981   0.7925   1.0000
   2.000   0.6992   0.01510   0.00746  -0.0969   0.7747   1.0000
   2.250   0.7257   0.01512   0.00733  -0.0959   0.7577   1.0000
   2.500   0.7522   0.01518   0.00724  -0.0949   0.7406   1.0000
   2.750   0.7778   0.01531   0.00725  -0.0940   0.7218   1.0000
   3.000   0.8034   0.01547   0.00731  -0.0930   0.7024   1.0000
   3.250   0.8295   0.01564   0.00736  -0.0922   0.6840   1.0000
   3.500   0.8555   0.01586   0.00749  -0.0914   0.6657   1.0000
   3.750   0.8805   0.01614   0.00772  -0.0906   0.6453   1.0000
   4.000   0.9060   0.01638   0.00787  -0.0897   0.6256   1.0000
   4.250   0.9307   0.01663   0.00804  -0.0888   0.6043   1.0000
   4.500   0.9551   0.01683   0.00821  -0.0878   0.5816   1.0000
   4.750   0.9791   0.01703   0.00839  -0.0868   0.5584   1.0000
   5.000   1.0035   0.01725   0.00856  -0.0859   0.5367   1.0000
   5.250   1.0275   0.01752   0.00883  -0.0850   0.5142   1.0000
   5.500   1.0517   0.01783   0.00909  -0.0841   0.4924   1.0000
   5.750   1.0749   0.01825   0.00948  -0.0832   0.4680   1.0000
   6.000   1.0979   0.01880   0.00993  -0.0822   0.4433   1.0000
   6.250   1.1194   0.01938   0.01039  -0.0810   0.4154   1.0000
   6.500   1.1398   0.01991   0.01089  -0.0797   0.3865   1.0000
   6.750   1.1597   0.02037   0.01134  -0.0784   0.3582   1.0000
   7.000   1.1793   0.02079   0.01176  -0.0771   0.3306   1.0000
   7.250   1.1996   0.02129   0.01231  -0.0759   0.3066   1.0000
   7.500   1.2186   0.02184   0.01285  -0.0746   0.2808   1.0000
   7.750   1.2362   0.02249   0.01347  -0.0732   0.2531   1.0000
   8.000   1.2546   0.02323   0.01420  -0.0718   0.2335   1.0000
   8.250   1.2726   0.02395   0.01494  -0.0705   0.2078   1.0000
   8.500   1.2901   0.02473   0.01569  -0.0691   0.1773   1.0000
   8.750   1.3064   0.02570   0.01655  -0.0675   0.1377   1.0000
   9.000   1.3165   0.02733   0.01785  -0.0652   0.0863   1.0000
   9.250   1.3214   0.02941   0.01960  -0.0623   0.0541   1.0000
   9.500   1.3292   0.03109   0.02133  -0.0596   0.0487   1.0000
   9.750   1.3347   0.03275   0.02313  -0.0567   0.0457   1.0000
  10.000   1.3347   0.03463   0.02513  -0.0532   0.0434   1.0000
  10.250   1.3321   0.03677   0.02741  -0.0499   0.0417   1.0000
  10.500   1.3331   0.03875   0.02960  -0.0473   0.0404   1.0000
  10.750   1.3327   0.04097   0.03200  -0.0449   0.0394   1.0000
  11.000   1.3319   0.04335   0.03455  -0.0427   0.0389   1.0000
  11.250   1.3318   0.04584   0.03719  -0.0408   0.0384   1.0000
  11.500   1.3334   0.04834   0.03985  -0.0391   0.0380   1.0000
  11.750   1.3379   0.05076   0.04239  -0.0374   0.0377   1.0000
  12.000   1.3464   0.05307   0.04481  -0.0357   0.0375   1.0000
  12.250   1.3588   0.05536   0.04725  -0.0340   0.0374   1.0000
  12.500   1.3727   0.05791   0.04999  -0.0324   0.0375   1.0000
  12.750   1.3811   0.06094   0.05327  -0.0310   0.0376   1.0000
  13.000   1.3835   0.06434   0.05694  -0.0298   0.0374   1.0000
  13.250   1.3807   0.06807   0.06093  -0.0289   0.0372   1.0000
  13.500   1.3740   0.07217   0.06534  -0.0285   0.0370   1.0000
  13.750   1.3632   0.07670   0.07016  -0.0285   0.0370   1.0000
  14.000   1.3488   0.08179   0.07553  -0.0292   0.0372   1.0000
  14.250   1.3314   0.08743   0.08146  -0.0306   0.0375   1.0000
  14.500   1.3115   0.09369   0.08799  -0.0329   0.0378   1.0000
  14.750   1.2897   0.10059   0.09515  -0.0361   0.0382   1.0000
  15.000   1.2665   0.10826   0.10302  -0.0402   0.0386   1.0000
  15.250   1.2424   0.11684   0.11182  -0.0454   0.0392   1.0000
  15.500   1.2176   0.12633   0.12150  -0.0515   0.0398   1.0000
  15.750   1.1930   0.13655   0.13187  -0.0582   0.0406   1.0000
  16.000   1.1709   0.14695   0.14235  -0.0648   0.0414   1.0000
<< Back to GOE 281 (DAIMLER XII) AIRFOIL (goe281-il)

Polar data table (+)

Polar graphs


<< Back to GOE 281 (DAIMLER XII) AIRFOIL (goe281-il)