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

GOE 590 AIRFOIL (goe590-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 590 AIRFOIL (goe590-il)
Reynolds number: 100,000
Max Cl/Cd: 53.72 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe590-il-100000-n5.txt
Download as CSV file: xf-goe590-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 590 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4084   0.10833   0.10358  -0.0127   1.0000   0.0281
  -8.250  -0.4061   0.10585   0.10116  -0.0147   1.0000   0.0282
  -8.000  -0.4038   0.10321   0.09859  -0.0166   1.0000   0.0282
  -7.750  -0.3999   0.10049   0.09592  -0.0190   1.0000   0.0283
  -7.500  -0.3926   0.09736   0.09282  -0.0215   1.0000   0.0283
  -7.250  -0.3840   0.09415   0.08963  -0.0242   1.0000   0.0284
  -7.000  -0.3740   0.09082   0.08631  -0.0263   1.0000   0.0284
  -6.750  -0.3632   0.08736   0.08285  -0.0283   1.0000   0.0284
  -6.500  -0.3570   0.08231   0.07788  -0.0281   1.0000   0.0287
  -6.250  -0.3506   0.07787   0.07348  -0.0271   1.0000   0.0292
  -6.000  -0.3412   0.07435   0.06999  -0.0276   1.0000   0.0296
  -5.750  -0.3311   0.07091   0.06652  -0.0283   1.0000   0.0301
  -5.500  -0.3198   0.06763   0.06324  -0.0293   1.0000   0.0306
  -5.250  -0.3078   0.06445   0.06006  -0.0301   1.0000   0.0311
  -5.000  -0.2953   0.06140   0.05700  -0.0308   1.0000   0.0316
  -4.750  -0.2799   0.05821   0.05376  -0.0319   0.9979   0.0316
  -4.500  -0.2282   0.05382   0.04897  -0.0396   0.9795   0.0282
  -4.250  -0.2053   0.04899   0.04412  -0.0422   0.9619   0.0262
  -4.000  -0.1662   0.04476   0.03960  -0.0464   0.9432   0.0271
  -3.750  -0.1320   0.04067   0.03525  -0.0494   0.9244   0.0262
  -3.500  -0.0977   0.03680   0.03106  -0.0517   0.9049   0.0253
  -3.000  -0.0334   0.02972   0.02310  -0.0536   0.8647   0.0255
  -2.750  -0.0015   0.02631   0.01894  -0.0530   0.8454   0.0283
  -2.500   0.0236   0.02396   0.01625  -0.0527   0.8247   0.0307
  -2.250   0.0510   0.02185   0.01362  -0.0519   0.8057   0.0336
  -2.000   0.0786   0.01998   0.01115  -0.0511   0.7881   0.0401
  -1.750   0.1055   0.01881   0.00957  -0.0505   0.7706   0.0465
  -1.500   0.1323   0.01771   0.00814  -0.0499   0.7543   0.0534
  -1.250   0.1592   0.01679   0.00695  -0.0493   0.7390   0.0583
  -1.000   0.1860   0.01616   0.00608  -0.0488   0.7242   0.0647
  -0.750   0.2120   0.01573   0.00551  -0.0483   0.7103   0.0716
  -0.500   0.2379   0.01526   0.00487  -0.0477   0.6971   0.0747
  -0.250   0.2641   0.01490   0.00437  -0.0472   0.6846   0.0796
   0.000   0.2907   0.01464   0.00399  -0.0468   0.6727   0.0915
   0.250   0.3168   0.01438   0.00372  -0.0464   0.6617   0.1280
   0.500   0.3412   0.01394   0.00361  -0.0460   0.6507   0.2553
   1.000   0.4453   0.01249   0.00341  -0.0562   0.6291   1.0000
   1.250   0.4699   0.01264   0.00341  -0.0555   0.6204   1.0000
   1.500   0.4946   0.01280   0.00349  -0.0550   0.6118   1.0000
   1.750   0.5192   0.01297   0.00357  -0.0543   0.6048   1.0000
   2.000   0.5439   0.01314   0.00371  -0.0538   0.5966   1.0000
   2.250   0.5686   0.01333   0.00384  -0.0532   0.5902   1.0000
   2.500   0.5934   0.01354   0.00406  -0.0527   0.5831   1.0000
   2.750   0.6182   0.01375   0.00429  -0.0521   0.5775   1.0000
   3.000   0.6430   0.01398   0.00459  -0.0517   0.5710   1.0000
   3.250   0.6678   0.01422   0.00487  -0.0511   0.5652   1.0000
   3.500   0.6927   0.01448   0.00521  -0.0506   0.5598   1.0000
   3.750   0.7175   0.01476   0.00567  -0.0501   0.5542   1.0000
   4.000   0.7426   0.01504   0.00607  -0.0496   0.5494   1.0000
   4.250   0.7671   0.01534   0.00657  -0.0491   0.5427   1.0000
   4.500   0.7902   0.01550   0.00685  -0.0480   0.5273   1.0000
   4.750   0.8125   0.01560   0.00710  -0.0466   0.5068   1.0000
   5.000   0.8289   0.01543   0.00691  -0.0439   0.4401   1.0000
   5.250   0.8260   0.01818   0.00774  -0.0397   0.0882   1.0000
   5.500   0.8384   0.02008   0.00942  -0.0375   0.0396   1.0000
   5.750   0.8538   0.02142   0.01093  -0.0356   0.0296   1.0000
   6.000   0.8673   0.02282   0.01255  -0.0335   0.0257   1.0000
   6.250   0.8802   0.02420   0.01414  -0.0311   0.0232   1.0000
   6.500   0.8932   0.02559   0.01566  -0.0289   0.0198   1.0000
   6.750   0.9046   0.02770   0.01771  -0.0265   0.0179   1.0000
   7.000   0.9245   0.02957   0.01969  -0.0250   0.0173   1.0000
   7.250   0.9484   0.03194   0.02224  -0.0237   0.0166   1.0000
   7.500   0.9723   0.03442   0.02498  -0.0226   0.0158   1.0000
   7.750   0.9934   0.03689   0.02780  -0.0212   0.0144   1.0000
   8.000   1.0117   0.03968   0.03096  -0.0197   0.0136   1.0000
   8.250   1.0272   0.04291   0.03461  -0.0178   0.0136   1.0000
   8.500   1.0386   0.04640   0.03852  -0.0158   0.0137   1.0000
   8.750   1.0465   0.04990   0.04251  -0.0136   0.0139   1.0000
   9.000   1.0505   0.05353   0.04652  -0.0114   0.0141   1.0000
   9.250   1.0510   0.05706   0.05040  -0.0092   0.0144   1.0000
   9.500   1.0473   0.06063   0.05426  -0.0070   0.0146   1.0000
   9.750   1.0390   0.06415   0.05805  -0.0048   0.0148   1.0000
  10.000   1.0265   0.06731   0.06142  -0.0025   0.0150   1.0000
  10.250   1.0119   0.07074   0.06502  -0.0012   0.0152   1.0000
  10.500   0.9957   0.07469   0.06915  -0.0012   0.0153   1.0000
  10.750   0.9800   0.07898   0.07359  -0.0022   0.0154   1.0000
  11.000   0.9640   0.08381   0.07856  -0.0042   0.0155   1.0000
  11.250   0.9481   0.08917   0.08405  -0.0070   0.0156   1.0000
  11.500   0.9320   0.09524   0.09021  -0.0106   0.0157   1.0000
<< Back to GOE 590 AIRFOIL (goe590-il)

Polar data table (+)

Polar graphs


<< Back to GOE 590 AIRFOIL (goe590-il)