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

GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 415 AIRFOIL (goe415-il)
Reynolds number: 50,000
Max Cl/Cd: 37.23 at α=7.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe415-il-50000-n5.txt
Download as CSV file: xf-goe415-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 415 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4295   0.13307   0.12629   0.0002   1.0000   0.0700
 -10.000  -0.4317   0.13156   0.12486  -0.0031   1.0000   0.0716
  -9.750  -0.4340   0.12993   0.12331  -0.0065   1.0000   0.0721
  -9.500  -0.4124   0.12264   0.11602  -0.0030   1.0000   0.0770
  -9.250  -0.4068   0.11954   0.11296  -0.0041   1.0000   0.0814
  -9.000  -0.4084   0.11761   0.11112  -0.0073   1.0000   0.0851
  -8.750  -0.4137   0.11622   0.10985  -0.0115   1.0000   0.0861
  -8.500  -0.3938   0.10995   0.10357  -0.0082   1.0000   0.0909
  -8.250  -0.3888   0.10696   0.10064  -0.0094   1.0000   0.0958
  -8.000  -0.3959   0.10562   0.09944  -0.0138   1.0000   0.0996
  -7.750  -0.3882   0.10131   0.09521  -0.0138   1.0000   0.1022
  -7.500  -0.3770   0.09762   0.09155  -0.0129   1.0000   0.1080
  -7.250  -0.3826   0.09607   0.09013  -0.0197   1.0000   0.1140
  -7.000  -0.3711   0.09151   0.08559  -0.0170   1.0000   0.1183
  -6.750  -0.3663   0.08884   0.08299  -0.0206   1.0000   0.1274
  -6.250  -0.3544   0.08271   0.07698  -0.0290   1.0000   0.1439
  -6.000  -0.3431   0.07864   0.07299  -0.0233   1.0000   0.1494
  -5.500  -0.3287   0.07255   0.06701  -0.0283   1.0000   0.1742
  -5.000  -0.2766   0.05916   0.05315  -0.0417   1.0000   0.0697
  -4.750  -0.2570   0.05446   0.04825  -0.0436   1.0000   0.0554
  -4.500  -0.2406   0.05078   0.04445  -0.0445   1.0000   0.0525
  -4.250  -0.2208   0.04698   0.04040  -0.0459   1.0000   0.0502
  -4.000  -0.1988   0.04330   0.03635  -0.0472   1.0000   0.0483
  -3.750  -0.1736   0.04000   0.03246  -0.0481   1.0000   0.0460
  -3.500  -0.1496   0.03796   0.02986  -0.0479   1.0000   0.0445
  -3.250  -0.1278   0.03511   0.02672  -0.0479   1.0000   0.0438
  -3.000  -0.1045   0.03275   0.02401  -0.0478   1.0000   0.0433
  -2.750  -0.0574   0.03000   0.02065  -0.0515   0.9891   0.0436
  -2.500  -0.0110   0.02740   0.01761  -0.0553   0.9773   0.0475
  -2.250   0.0338   0.02545   0.01534  -0.0581   0.9632   0.0507
  -2.000   0.0788   0.02384   0.01337  -0.0607   0.9484   0.0543
  -1.750   0.1244   0.02242   0.01170  -0.0635   0.9328   0.0642
  -1.500   0.1687   0.02134   0.01031  -0.0659   0.9150   0.0729
  -1.250   0.2098   0.02018   0.00919  -0.0681   0.8947   0.1054
  -1.000   0.2467   0.01671   0.00860  -0.0687   0.8741   1.0000
  -0.750   0.2803   0.01675   0.00803  -0.0689   0.8465   1.0000
  -0.500   0.3113   0.01679   0.00759  -0.0688   0.8175   1.0000
  -0.250   0.3411   0.01684   0.00723  -0.0685   0.7887   1.0000
   0.000   0.3699   0.01691   0.00692  -0.0679   0.7612   1.0000
   0.250   0.3978   0.01703   0.00666  -0.0672   0.7351   1.0000
   0.500   0.4248   0.01720   0.00648  -0.0664   0.7105   1.0000
   0.750   0.4511   0.01743   0.00640  -0.0656   0.6869   1.0000
   1.000   0.4770   0.01771   0.00639  -0.0648   0.6639   1.0000
   1.250   0.5028   0.01800   0.00642  -0.0641   0.6430   1.0000
   1.500   0.5284   0.01833   0.00654  -0.0634   0.6226   1.0000
   1.750   0.5541   0.01866   0.00668  -0.0628   0.6035   1.0000
   2.000   0.5798   0.01901   0.00685  -0.0622   0.5859   1.0000
   2.250   0.6055   0.01938   0.00705  -0.0616   0.5695   1.0000
   2.500   0.6316   0.01978   0.00734  -0.0612   0.5534   1.0000
   2.750   0.6578   0.02020   0.00766  -0.0608   0.5378   1.0000
   3.000   0.6839   0.02063   0.00802  -0.0605   0.5231   1.0000
   3.250   0.7100   0.02108   0.00845  -0.0601   0.5094   1.0000
   3.500   0.7361   0.02155   0.00887  -0.0598   0.4964   1.0000
   3.750   0.7618   0.02204   0.00937  -0.0594   0.4833   1.0000
   4.000   0.7874   0.02255   0.00998  -0.0591   0.4708   1.0000
   4.250   0.8129   0.02309   0.01057  -0.0588   0.4589   1.0000
   4.500   0.8384   0.02363   0.01117  -0.0584   0.4478   1.0000
   4.750   0.8641   0.02416   0.01173  -0.0580   0.4379   1.0000
   5.000   0.8891   0.02477   0.01253  -0.0577   0.4270   1.0000
   5.250   0.9139   0.02541   0.01332  -0.0573   0.4166   1.0000
   5.500   0.9390   0.02601   0.01402  -0.0569   0.4072   1.0000
   5.750   0.9636   0.02663   0.01485  -0.0564   0.3972   1.0000
   6.000   0.9871   0.02725   0.01565  -0.0558   0.3852   1.0000
   6.250   1.0103   0.02780   0.01637  -0.0551   0.3726   1.0000
   6.500   1.0333   0.02834   0.01709  -0.0544   0.3601   1.0000
   6.750   1.0564   0.02892   0.01785  -0.0537   0.3486   1.0000
   7.000   1.0800   0.02955   0.01872  -0.0530   0.3386   1.0000
   7.250   1.1020   0.03037   0.01989  -0.0522   0.3276   1.0000
   7.500   1.1205   0.03046   0.02009  -0.0507   0.3060   1.0000
   7.750   1.1345   0.03047   0.02018  -0.0489   0.2743   1.0000
   8.000   1.1513   0.03095   0.02080  -0.0475   0.2520   1.0000
   8.250   1.1677   0.03167   0.02177  -0.0463   0.2262   1.0000
   8.500   1.1821   0.03258   0.02281  -0.0449   0.1892   1.0000
   8.750   1.1933   0.03399   0.02420  -0.0434   0.1381   1.0000
   9.250   1.1831   0.04069   0.02993  -0.0397   0.0397   1.0000
   9.500   1.1784   0.04368   0.03302  -0.0379   0.0352   1.0000
   9.750   1.1705   0.04700   0.03650  -0.0366   0.0329   1.0000
  10.000   1.1609   0.05086   0.04051  -0.0363   0.0315   1.0000
  10.250   1.1527   0.05499   0.04490  -0.0370   0.0305   1.0000
  10.500   1.1438   0.05967   0.04982  -0.0384   0.0299   1.0000
  10.750   1.1343   0.06477   0.05514  -0.0404   0.0295   1.0000
  11.000   1.1246   0.07005   0.06063  -0.0424   0.0291   1.0000
  11.250   1.1150   0.07535   0.06612  -0.0444   0.0288   1.0000
  11.500   1.1056   0.08060   0.07154  -0.0462   0.0284   1.0000
  11.750   1.0970   0.08572   0.07681  -0.0479   0.0281   1.0000
  12.000   1.0893   0.09068   0.08190  -0.0494   0.0278   1.0000
  12.250   1.0827   0.09542   0.08677  -0.0508   0.0274   1.0000
<< Back to GOE 415 AIRFOIL (goe415-il)

Polar data table (+)

Polar graphs


<< Back to GOE 415 AIRFOIL (goe415-il)