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

GOE 275 (DAIMLER VI) AIRFOIL (goe275-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 275 (DAIMLER VI) AIRFOIL (goe275-il)
Reynolds number: 200,000
Max Cl/Cd: 72.01 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe275-il-200000.txt
Download as CSV file: xf-goe275-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 275 (DAIMLER VI) 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  -0.3346   0.10189   0.09836  -0.0266   1.0000   0.0282
  -8.250  -0.3434   0.10045   0.09700  -0.0268   1.0000   0.0282
  -8.000  -0.3534   0.09887   0.09551  -0.0267   1.0000   0.0283
  -7.750  -0.3527   0.09552   0.09221  -0.0255   1.0000   0.0284
  -7.500  -0.3426   0.09140   0.08810  -0.0215   1.0000   0.0288
  -7.250  -0.3452   0.08914   0.08589  -0.0195   1.0000   0.0291
  -7.000  -0.3488   0.08699   0.08379  -0.0181   1.0000   0.0294
  -6.750  -0.3527   0.08491   0.08176  -0.0170   1.0000   0.0297
  -6.500  -0.3557   0.08274   0.07963  -0.0164   1.0000   0.0301
  -6.250  -0.3569   0.08040   0.07733  -0.0162   1.0000   0.0306
  -6.000  -0.3558   0.07787   0.07482  -0.0168   1.0000   0.0312
  -5.750  -0.3519   0.07516   0.07212  -0.0179   1.0000   0.0318
  -5.500  -0.3438   0.07219   0.06913  -0.0201   1.0000   0.0328
  -5.250  -0.2878   0.06522   0.06192  -0.0365   0.9948   0.0344
  -5.000  -0.2650   0.06152   0.05825  -0.0383   0.9900   0.0353
  -4.750  -0.2278   0.05745   0.05410  -0.0444   0.9852   0.0367
  -4.500  -0.1849   0.05311   0.04959  -0.0517   0.9789   0.0395
  -4.250  -0.1319   0.04760   0.04370  -0.0610   0.9738   0.0419
  -4.000  -0.1032   0.04455   0.04068  -0.0636   0.9673   0.0437
  -3.500  -0.0163   0.03736   0.03291  -0.0733   0.9556   0.0517
  -3.250   0.0163   0.03483   0.03035  -0.0758   0.9497   0.0545
  -3.000   0.0644   0.03221   0.02721  -0.0801   0.9461   0.0625
  -2.750   0.0913   0.02987   0.02496  -0.0813   0.9379   0.0663
  -2.500   0.1348   0.02763   0.02236  -0.0846   0.9335   0.0766
  -2.250   0.1728   0.02624   0.02068  -0.0866   0.9271   0.0895
  -1.750   0.2506   0.02219   0.01659  -0.0924   0.9173   0.1235
  -1.500   0.2810   0.02093   0.01533  -0.0934   0.9075   0.1554
  -0.500   0.4299   0.01619   0.00954  -0.0967   0.8589   0.1497
  -0.250   0.4660   0.01458   0.00759  -0.0960   0.8405   0.0970
   0.000   0.4953   0.01406   0.00685  -0.0950   0.8159   0.0860
   0.250   0.5240   0.01334   0.00594  -0.0941   0.7880   0.0767
   0.500   0.5515   0.01310   0.00553  -0.0932   0.7529   0.0731
   0.750   0.5773   0.01260   0.00488  -0.0921   0.7117   0.0708
   1.000   0.6013   0.01241   0.00446  -0.0907   0.6673   0.0693
   1.250   0.6239   0.01240   0.00421  -0.0892   0.6248   0.0687
   1.500   0.6461   0.01247   0.00407  -0.0877   0.5855   0.0701
   1.750   0.6685   0.01260   0.00400  -0.0864   0.5499   0.0722
   2.000   0.6915   0.01274   0.00395  -0.0852   0.5199   0.0729
   2.250   0.7151   0.01294   0.00394  -0.0842   0.4963   0.0739
   2.500   0.7394   0.01309   0.00396  -0.0833   0.4765   0.0770
   2.750   0.7639   0.01327   0.00403  -0.0825   0.4599   0.0850
   3.000   0.7980   0.01178   0.00433  -0.0841   0.4450   1.0000
   3.250   0.8224   0.01208   0.00445  -0.0832   0.4318   1.0000
   3.750   0.8714   0.01266   0.00476  -0.0817   0.4075   1.0000
   4.000   0.8961   0.01295   0.00498  -0.0811   0.3976   1.0000
   4.250   0.9208   0.01330   0.00520  -0.0804   0.3890   1.0000
   4.500   0.9458   0.01354   0.00546  -0.0798   0.3803   1.0000
   4.750   0.9707   0.01392   0.00574  -0.0793   0.3728   1.0000
   5.000   0.9954   0.01416   0.00603  -0.0786   0.3642   1.0000
   5.250   1.0199   0.01453   0.00632  -0.0780   0.3560   1.0000
   5.500   1.0440   0.01478   0.00662  -0.0773   0.3465   1.0000
   5.750   1.0678   0.01510   0.00694  -0.0765   0.3369   1.0000
   6.000   1.0914   0.01544   0.00724  -0.0758   0.3273   1.0000
   6.250   1.1149   0.01569   0.00758  -0.0750   0.3176   1.0000
   6.500   1.1382   0.01604   0.00792  -0.0742   0.3083   1.0000
   6.750   1.1609   0.01630   0.00825  -0.0732   0.2975   1.0000
   7.000   1.1833   0.01654   0.00857  -0.0722   0.2855   1.0000
   7.250   1.2054   0.01680   0.00888  -0.0712   0.2732   1.0000
   7.500   1.2272   0.01706   0.00920  -0.0701   0.2601   1.0000
   7.750   1.2486   0.01734   0.00953  -0.0690   0.2464   1.0000
   8.000   1.2697   0.01766   0.00988  -0.0679   0.2308   1.0000
   8.250   1.2897   0.01806   0.01028  -0.0666   0.2135   1.0000
   8.500   1.3073   0.01865   0.01081  -0.0650   0.1954   1.0000
   8.750   1.3229   0.01949   0.01156  -0.0632   0.1776   1.0000
   9.000   1.3375   0.02046   0.01248  -0.0612   0.1609   1.0000
   9.250   1.3520   0.02142   0.01340  -0.0593   0.1474   1.0000
   9.500   1.3670   0.02226   0.01422  -0.0575   0.1370   1.0000
   9.750   1.3847   0.02285   0.01494  -0.0560   0.1279   1.0000
  10.000   1.3992   0.02361   0.01570  -0.0541   0.1209   1.0000
  10.250   1.4143   0.02423   0.01642  -0.0522   0.1147   1.0000
  10.500   1.4251   0.02511   0.01727  -0.0498   0.1100   1.0000
  10.750   1.4391   0.02578   0.01810  -0.0479   0.1045   1.0000
  11.000   1.4478   0.02674   0.01905  -0.0454   0.0995   1.0000
  11.250   1.4580   0.02770   0.02012  -0.0432   0.0949   1.0000
  11.500   1.4662   0.02875   0.02126  -0.0409   0.0904   1.0000
  11.750   1.4692   0.03031   0.02277  -0.0383   0.0861   1.0000
  12.000   1.4778   0.03136   0.02408  -0.0363   0.0819   1.0000
  12.250   1.4812   0.03283   0.02565  -0.0341   0.0772   1.0000
  12.500   1.4814   0.03469   0.02759  -0.0319   0.0722   1.0000
  12.750   1.4845   0.03619   0.02930  -0.0302   0.0650   1.0000
  13.000   1.4844   0.03818   0.03144  -0.0287   0.0568   1.0000
  13.250   1.4829   0.04046   0.03379  -0.0275   0.0487   1.0000
  13.500   1.4784   0.04326   0.03662  -0.0266   0.0429   1.0000
  13.750   1.4733   0.04640   0.03989  -0.0260   0.0391   1.0000
  14.000   1.4668   0.04990   0.04349  -0.0260   0.0366   1.0000
  14.250   1.4554   0.05424   0.04789  -0.0265   0.0350   1.0000
  14.500   1.4470   0.05848   0.05227  -0.0273   0.0337   1.0000
  14.750   1.4394   0.06281   0.05677  -0.0283   0.0323   1.0000
  15.000   1.4308   0.06742   0.06151  -0.0297   0.0312   1.0000
  15.250   1.4215   0.07232   0.06653  -0.0314   0.0301   1.0000
  15.500   1.4110   0.07752   0.07183  -0.0333   0.0293   1.0000
  15.750   1.4002   0.08277   0.07714  -0.0352   0.0286   1.0000
  16.000   1.3912   0.08779   0.08225  -0.0369   0.0278   1.0000
<< Back to GOE 275 (DAIMLER VI) AIRFOIL (goe275-il)

Polar data table (+)

Polar graphs


<< Back to GOE 275 (DAIMLER VI) AIRFOIL (goe275-il)