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=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 281 (DAIMLER XII) AIRFOIL (goe281-il)
Reynolds number: 50,000
Max Cl/Cd: 42.09 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe281-il-50000-n5.txt
Download as CSV file: xf-goe281-il-50000-n5.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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3449   0.11659   0.10977  -0.0262   1.0000   0.0715
  -8.750  -0.3500   0.11554   0.10883  -0.0285   1.0000   0.0722
  -8.500  -0.3567   0.11445   0.10788  -0.0303   1.0000   0.0725
  -8.250  -0.3309   0.10633   0.09970  -0.0268   1.0000   0.0769
  -8.000  -0.3291   0.10377   0.09721  -0.0268   1.0000   0.0809
  -7.750  -0.3354   0.10230   0.09587  -0.0273   1.0000   0.0842
  -7.500  -0.3433   0.10149   0.09521  -0.0306   1.0000   0.0859
  -7.250  -0.3396   0.09737   0.09118  -0.0292   1.0000   0.0882
  -7.000  -0.3334   0.09398   0.08784  -0.0269   1.0000   0.0937
  -6.750  -0.3349   0.09204   0.08599  -0.0287   1.0000   0.0985
  -6.500  -0.3355   0.09143   0.08540  -0.0365   1.0000   0.1009
  -6.250  -0.3318   0.08627   0.08034  -0.0293   1.0000   0.1044
  -6.000  -0.3284   0.08363   0.07775  -0.0287   1.0000   0.1098
  -5.750  -0.3215   0.08185   0.07593  -0.0361   1.0000   0.1161
  -5.250  -0.3064   0.07630   0.07037  -0.0378   1.0000   0.1307
  -5.000  -0.3072   0.07256   0.06680  -0.0318   1.0000   0.1357
  -4.500  -0.2879   0.06718   0.06139  -0.0345   1.0000   0.1625
  -4.250  -0.2737   0.06422   0.05842  -0.0355   0.9986   0.1781
  -4.000  -0.2409   0.06053   0.05461  -0.0406   0.9917   0.2069
  -3.750  -0.2133   0.05709   0.05113  -0.0428   0.9852   0.2393
  -3.250  -0.0772   0.04540   0.03780  -0.0662   0.9725   0.0984
  -3.000  -0.0322   0.04164   0.03350  -0.0706   0.9663   0.0828
  -2.750   0.0075   0.03891   0.03034  -0.0741   0.9591   0.0814
  -2.500   0.0510   0.03642   0.02734  -0.0778   0.9523   0.0805
  -2.250   0.0919   0.03422   0.02458  -0.0804   0.9430   0.0766
  -2.000   0.1387   0.03241   0.02208  -0.0836   0.9348   0.0736
  -1.750   0.1804   0.03104   0.02022  -0.0859   0.9240   0.0725
  -1.500   0.2184   0.02946   0.01840  -0.0879   0.9128   0.0740
  -1.250   0.2571   0.02823   0.01702  -0.0901   0.9032   0.0774
  -1.000   0.2953   0.02723   0.01578  -0.0917   0.8946   0.0779
  -0.750   0.3285   0.02648   0.01479  -0.0922   0.8839   0.0781
  -0.500   0.3638   0.02579   0.01390  -0.0931   0.8744   0.0786
  -0.250   0.4002   0.02512   0.01307  -0.0943   0.8656   0.0796
   0.000   0.4311   0.02468   0.01251  -0.0946   0.8540   0.0810
   0.250   0.4627   0.02430   0.01201  -0.0950   0.8427   0.0830
   0.500   0.4948   0.02397   0.01156  -0.0954   0.8318   0.0858
   0.750   0.5287   0.02352   0.01111  -0.0962   0.8217   0.0947
   1.000   0.5588   0.02323   0.01086  -0.0963   0.8092   0.1095
   1.250   0.5847   0.02097   0.01081  -0.0958   0.7968   1.0000
   1.500   0.6139   0.02109   0.01063  -0.0954   0.7836   1.0000
   1.750   0.6427   0.02121   0.01052  -0.0951   0.7702   1.0000
   2.000   0.6711   0.02133   0.01046  -0.0947   0.7564   1.0000
   2.250   0.6991   0.02144   0.01044  -0.0943   0.7417   1.0000
   2.500   0.7265   0.02152   0.01039  -0.0936   0.7252   1.0000
   2.750   0.7539   0.02155   0.01029  -0.0927   0.7073   1.0000
   3.000   0.7815   0.02156   0.01020  -0.0919   0.6894   1.0000
   3.250   0.8062   0.02176   0.01032  -0.0909   0.6692   1.0000
   3.500   0.8321   0.02195   0.01044  -0.0901   0.6508   1.0000
   3.750   0.8581   0.02218   0.01065  -0.0893   0.6334   1.0000
   4.000   0.8840   0.02245   0.01088  -0.0886   0.6163   1.0000
   4.250   0.9088   0.02278   0.01121  -0.0878   0.5983   1.0000
   4.500   0.9327   0.02317   0.01168  -0.0869   0.5795   1.0000
   4.750   0.9566   0.02354   0.01209  -0.0861   0.5607   1.0000
   5.000   0.9808   0.02388   0.01246  -0.0852   0.5421   1.0000
   5.250   1.0037   0.02429   0.01293  -0.0842   0.5222   1.0000
   5.500   1.0264   0.02464   0.01335  -0.0830   0.5006   1.0000
   5.750   1.0484   0.02501   0.01371  -0.0818   0.4778   1.0000
   6.000   1.0703   0.02543   0.01408  -0.0804   0.4548   1.0000
   6.250   1.0915   0.02601   0.01467  -0.0792   0.4318   1.0000
   6.500   1.1130   0.02666   0.01536  -0.0780   0.4109   1.0000
   6.750   1.1341   0.02739   0.01616  -0.0769   0.3910   1.0000
   7.000   1.1545   0.02818   0.01710  -0.0758   0.3715   1.0000
   7.250   1.1719   0.02890   0.01786  -0.0742   0.3481   1.0000
   7.500   1.1854   0.02963   0.01862  -0.0722   0.3199   1.0000
   7.750   1.1999   0.03044   0.01957  -0.0704   0.2950   1.0000
   8.000   1.2139   0.03133   0.02049  -0.0685   0.2734   1.0000
   8.250   1.2235   0.03240   0.02145  -0.0663   0.2469   1.0000
   8.500   1.2363   0.03348   0.02264  -0.0645   0.2275   1.0000
   8.750   1.2476   0.03464   0.02391  -0.0626   0.2085   1.0000
   9.000   1.2553   0.03598   0.02529  -0.0605   0.1815   1.0000
   9.250   1.2643   0.03725   0.02676  -0.0583   0.1621   1.0000
   9.500   1.2680   0.03892   0.02839  -0.0559   0.1329   1.0000
   9.750   1.2699   0.04097   0.03038  -0.0537   0.1032   1.0000
  10.000   1.2605   0.04427   0.03322  -0.0513   0.0516   1.0000
  10.250   1.2540   0.04757   0.03640  -0.0495   0.0433   1.0000
  10.500   1.2492   0.05084   0.03976  -0.0482   0.0398   1.0000
  10.750   1.2432   0.05441   0.04348  -0.0472   0.0378   1.0000
  11.000   1.2378   0.05810   0.04735  -0.0468   0.0366   1.0000
  11.250   1.2321   0.06199   0.05148  -0.0467   0.0357   1.0000
  11.500   1.2253   0.06622   0.05594  -0.0470   0.0350   1.0000
  11.750   1.2175   0.07073   0.06068  -0.0477   0.0344   1.0000
  12.000   1.2092   0.07548   0.06564  -0.0486   0.0338   1.0000
  12.250   1.2005   0.08041   0.07077  -0.0497   0.0333   1.0000
  12.500   1.1918   0.08544   0.07599  -0.0510   0.0328   1.0000
  12.750   1.1834   0.09051   0.08123  -0.0523   0.0322   1.0000
  13.000   1.1755   0.09556   0.08644  -0.0537   0.0317   1.0000
  13.250   1.1684   0.10052   0.09158  -0.0551   0.0311   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)