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

GOE 278 (DAIMLER IX) AIRFOIL (goe278-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 278 (DAIMLER IX) AIRFOIL (goe278-il)
Reynolds number: 200,000
Max Cl/Cd: 72.53 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe278-il-200000.txt
Download as CSV file: xf-goe278-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 278 (DAIMLER IX) 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
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.2735   0.08738   0.08405  -0.0287   1.0000   0.0300
  -8.250  -0.2734   0.08452   0.08123  -0.0282   1.0000   0.0305
  -8.000  -0.3674   0.09434   0.09089  -0.0272   1.0000   0.0290
  -7.750  -0.3620   0.09100   0.08757  -0.0254   1.0000   0.0294
  -7.500  -0.3623   0.08846   0.08509  -0.0245   1.0000   0.0298
  -7.250  -0.3637   0.08611   0.08279  -0.0240   1.0000   0.0302
  -7.000  -0.3669   0.08392   0.08065  -0.0233   1.0000   0.0307
  -6.750  -0.3700   0.08177   0.07853  -0.0226   1.0000   0.0313
  -6.500  -0.3718   0.07949   0.07629  -0.0223   1.0000   0.0320
  -6.250  -0.3708   0.07701   0.07383  -0.0228   1.0000   0.0330
  -6.000  -0.3652   0.07430   0.07110  -0.0247   1.0000   0.0342
  -5.750  -0.3432   0.07166   0.06828  -0.0325   1.0000   0.0354
  -5.500  -0.3258   0.06887   0.06529  -0.0357   1.0000   0.0357
  -5.250  -0.3236   0.06367   0.06018  -0.0345   1.0000   0.0362
  -5.000  -0.3153   0.06032   0.05687  -0.0337   0.9994   0.0368
  -4.750  -0.2850   0.05650   0.05301  -0.0376   0.9959   0.0383
  -4.500  -0.2459   0.05256   0.04894  -0.0435   0.9917   0.0410
  -4.250  -0.1845   0.04808   0.04386  -0.0532   0.9864   0.0457
  -4.000  -0.1531   0.04374   0.03959  -0.0570   0.9824   0.0473
  -3.750  -0.1164   0.04071   0.03645  -0.0609   0.9763   0.0504
  -3.500  -0.0597   0.03798   0.03307  -0.0668   0.9721   0.0577
  -3.250  -0.0291   0.03438   0.02959  -0.0697   0.9654   0.0606
  -3.000   0.0261   0.03517   0.02966  -0.0735   0.9597   0.0701
  -2.750   0.0630   0.02915   0.02381  -0.0784   0.9572   0.0738
  -0.500   0.3912   0.01604   0.00881  -0.0900   0.8976   0.0814
  -0.250   0.4196   0.01513   0.00769  -0.0889   0.8848   0.0713
   0.000   0.4471   0.01464   0.00714  -0.0880   0.8710   0.0699
   0.250   0.4739   0.01380   0.00628  -0.0872   0.8564   0.0709
   0.500   0.5007   0.01333   0.00577  -0.0863   0.8402   0.0701
   0.750   0.5270   0.01268   0.00512  -0.0854   0.8226   0.0699
   1.000   0.5524   0.01218   0.00464  -0.0844   0.8003   0.0704
   1.250   0.5780   0.01166   0.00412  -0.0835   0.7761   0.0723
   1.500   0.6036   0.01140   0.00378  -0.0826   0.7461   0.0758
   1.750   0.6291   0.01132   0.00355  -0.0816   0.7111   0.0833
   2.000   0.6542   0.01130   0.00333  -0.0807   0.6751   0.0957
   2.250   0.6810   0.00953   0.00328  -0.0802   0.6402   1.0000
   2.500   0.7050   0.00990   0.00332  -0.0791   0.6037   1.0000
   2.750   0.7288   0.01029   0.00344  -0.0782   0.5680   1.0000
   3.000   0.7528   0.01064   0.00359  -0.0773   0.5346   1.0000
   3.250   0.7772   0.01100   0.00375  -0.0766   0.5070   1.0000
   3.500   0.8018   0.01133   0.00394  -0.0760   0.4824   1.0000
   3.750   0.8261   0.01170   0.00413  -0.0753   0.4615   1.0000
   4.000   0.8509   0.01202   0.00437  -0.0747   0.4415   1.0000
   4.250   0.8757   0.01235   0.00461  -0.0742   0.4243   1.0000
   4.500   0.9009   0.01269   0.00488  -0.0737   0.4111   1.0000
   4.750   0.9259   0.01304   0.00517  -0.0733   0.3992   1.0000
   5.000   0.9507   0.01338   0.00547  -0.0728   0.3857   1.0000
   5.250   0.9752   0.01370   0.00576  -0.0722   0.3709   1.0000
   5.500   1.0001   0.01399   0.00608  -0.0717   0.3582   1.0000
   5.750   1.0248   0.01428   0.00642  -0.0712   0.3452   1.0000
   6.000   1.0488   0.01457   0.00673  -0.0706   0.3289   1.0000
   6.250   1.0725   0.01488   0.00704  -0.0699   0.3116   1.0000
   6.500   1.0960   0.01511   0.00731  -0.0692   0.2844   1.0000
   6.750   1.1183   0.01548   0.00759  -0.0683   0.2341   1.0000
   7.000   1.1307   0.01716   0.00848  -0.0663   0.1143   1.0000
   7.250   1.1460   0.01865   0.00963  -0.0646   0.0738   1.0000
   7.500   1.1640   0.01973   0.01071  -0.0632   0.0659   1.0000
   7.750   1.1832   0.02058   0.01168  -0.0618   0.0623   1.0000
   8.000   1.2008   0.02156   0.01278  -0.0603   0.0591   1.0000
   8.250   1.2147   0.02284   0.01410  -0.0584   0.0559   1.0000
   8.500   1.2262   0.02439   0.01569  -0.0561   0.0534   1.0000
   8.750   1.2425   0.02549   0.01689  -0.0544   0.0518   1.0000
   9.000   1.2579   0.02670   0.01821  -0.0527   0.0494   1.0000
   9.250   1.2725   0.02804   0.01960  -0.0509   0.0469   1.0000
   9.500   1.2876   0.02980   0.02135  -0.0494   0.0450   1.0000
   9.750   1.3082   0.03282   0.02438  -0.0487   0.0428   1.0000
  10.000   1.3240   0.03386   0.02564  -0.0470   0.0412   1.0000
  10.250   1.3412   0.03550   0.02748  -0.0456   0.0395   1.0000
  10.500   1.3588   0.03749   0.02965  -0.0444   0.0381   1.0000
  10.750   1.3752   0.03953   0.03181  -0.0432   0.0365   1.0000
  11.000   1.3918   0.04221   0.03454  -0.0425   0.0347   1.0000
  11.250   1.4054   0.04820   0.04083  -0.0420   0.0333   1.0000
  11.500   1.4065   0.04986   0.04282  -0.0388   0.0329   1.0000
  11.750   1.4055   0.05223   0.04552  -0.0358   0.0326   1.0000
  12.000   1.3994   0.05485   0.04842  -0.0322   0.0324   1.0000
  12.250   1.3901   0.05776   0.05161  -0.0289   0.0323   1.0000
  12.500   1.3779   0.06082   0.05494  -0.0260   0.0322   1.0000
  12.750   1.3638   0.06418   0.05856  -0.0237   0.0321   1.0000
  13.000   1.3476   0.06792   0.06255  -0.0221   0.0320   1.0000
  13.250   1.3297   0.07212   0.06698  -0.0213   0.0320   1.0000
  13.500   1.3104   0.07684   0.07192  -0.0214   0.0320   1.0000
  13.750   1.2897   0.08219   0.07749  -0.0225   0.0321   1.0000
  14.000   1.2682   0.08820   0.08369  -0.0245   0.0323   1.0000
  14.250   1.2460   0.09499   0.09065  -0.0273   0.0326   1.0000
  14.750   1.1748   0.11531   0.11162  -0.0419   0.0365   1.0000
  15.000   1.1822   0.11747   0.11374  -0.0414   0.0351   1.0000
  15.250   1.0845   0.14814   0.14470  -0.0634   0.0420   1.0000
  15.500   0.8515   0.13951   0.13628  -0.0431   0.0451   1.0000
<< Back to GOE 278 (DAIMLER IX) AIRFOIL (goe278-il)

Polar data table (+)

Polar graphs


<< Back to GOE 278 (DAIMLER IX) AIRFOIL (goe278-il)