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

GOE 380 AIRFOIL (goe380-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 380 AIRFOIL (goe380-il)
Reynolds number: 50,000
Max Cl/Cd: 31.12 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe380-il-50000.txt
Download as CSV file: xf-goe380-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 380 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.3122   0.10812   0.10168  -0.0235   1.0000   0.0939
  -7.750  -0.3194   0.10759   0.10129  -0.0240   1.0000   0.0945
  -7.500  -0.3244   0.10733   0.10116  -0.0264   1.0000   0.0949
  -7.250  -0.3088   0.09979   0.09365  -0.0223   1.0000   0.0977
  -7.000  -0.3051   0.09685   0.09079  -0.0217   1.0000   0.1004
  -6.750  -0.3042   0.09466   0.08864  -0.0218   1.0000   0.1031
  -6.500  -0.3047   0.09312   0.08720  -0.0229   1.0000   0.1056
  -6.250  -0.3050   0.09275   0.08691  -0.0261   1.0000   0.1071
  -6.000  -0.3017   0.08888   0.08314  -0.0248   1.0000   0.1087
  -5.750  -0.2981   0.08520   0.07953  -0.0222   1.0000   0.1118
  -5.500  -0.2951   0.08288   0.07727  -0.0219   1.0000   0.1154
  -5.250  -0.2887   0.08221   0.07658  -0.0257   1.0000   0.1196
  -5.000  -0.2854   0.07875   0.07321  -0.0245   1.0000   0.1217
  -4.750  -0.2826   0.07549   0.07002  -0.0221   1.0000   0.1259
  -4.500  -0.2689   0.07475   0.06917  -0.0263   1.0000   0.1330
  -4.250  -0.2668   0.07076   0.06530  -0.0239   1.0000   0.1362
  -4.000  -0.2595   0.06824   0.06279  -0.0233   1.0000   0.1434
  -3.750  -0.2483   0.06573   0.06026  -0.0245   1.0000   0.1503
  -3.500  -0.2317   0.06427   0.05867  -0.0270   1.0000   0.1620
  -3.250  -0.2285   0.06077   0.05529  -0.0243   1.0000   0.1695
  -3.000  -0.2170   0.05844   0.05290  -0.0247   1.0000   0.1833
  -2.500  -0.1918   0.05440   0.04876  -0.0260   1.0000   0.2322
  -2.250  -0.1849   0.05180   0.04622  -0.0247   1.0000   0.2616
  -2.000  -0.1794   0.04939   0.04391  -0.0228   1.0000   0.3050
  -1.250  -0.1415   0.04167   0.03642  -0.0187   0.9902   0.4747
  -1.000  -0.0941   0.03871   0.03337  -0.0233   0.9751   0.5234
  -0.750  -0.0259   0.03627   0.03062  -0.0329   0.9589   0.5352
  -0.250   0.1858   0.03626   0.02763  -0.0626   0.9227   0.1861
   0.000   0.2390   0.03471   0.02544  -0.0659   0.9074   0.1653
   0.250   0.2875   0.03330   0.02358  -0.0688   0.8924   0.1683
   0.500   0.3414   0.03193   0.02169  -0.0723   0.8776   0.1866
   0.750   0.4001   0.03024   0.01976  -0.0767   0.8637   0.2341
   1.000   0.4489   0.02904   0.01858  -0.0797   0.8490   0.2868
   1.250   0.5225   0.02660   0.01754  -0.0870   0.8352   1.0000
   1.500   0.5667   0.02664   0.01702  -0.0886   0.8191   1.0000
   1.750   0.6092   0.02658   0.01659  -0.0901   0.8031   1.0000
   2.000   0.6448   0.02665   0.01643  -0.0906   0.7862   1.0000
   2.250   0.6715   0.02702   0.01664  -0.0899   0.7675   1.0000
   2.500   0.7031   0.02717   0.01666  -0.0896   0.7504   1.0000
   2.750   0.7349   0.02729   0.01668  -0.0894   0.7341   1.0000
   3.000   0.7655   0.02744   0.01674  -0.0889   0.7183   1.0000
   3.250   0.7912   0.02784   0.01707  -0.0879   0.7017   1.0000
   3.500   0.8130   0.02850   0.01770  -0.0866   0.6847   1.0000
   3.750   0.8371   0.02903   0.01826  -0.0855   0.6687   1.0000
   4.000   0.8614   0.02957   0.01877  -0.0844   0.6534   1.0000
   4.250   0.8862   0.03004   0.01923  -0.0832   0.6379   1.0000
   4.500   0.9114   0.03043   0.01961  -0.0820   0.6225   1.0000
   4.750   0.9370   0.03077   0.01994  -0.0807   0.6070   1.0000
   5.000   0.9622   0.03116   0.02041  -0.0794   0.5917   1.0000
   5.250   0.9864   0.03170   0.02098  -0.0782   0.5768   1.0000
   5.500   1.0063   0.03260   0.02196  -0.0767   0.5614   1.0000
   5.750   1.0258   0.03356   0.02301  -0.0752   0.5462   1.0000
   6.000   1.0453   0.03456   0.02413  -0.0738   0.5312   1.0000
   6.250   1.0644   0.03561   0.02537  -0.0723   0.5160   1.0000
   6.500   1.0830   0.03672   0.02663  -0.0707   0.5008   1.0000
   6.750   1.1013   0.03788   0.02794  -0.0692   0.4855   1.0000
   7.000   1.1191   0.03911   0.02934  -0.0676   0.4706   1.0000
   7.250   1.1360   0.04048   0.03088  -0.0660   0.4562   1.0000
   7.500   1.1520   0.04193   0.03252  -0.0644   0.4421   1.0000
   7.750   1.1666   0.04358   0.03445  -0.0628   0.4291   1.0000
   8.000   1.1824   0.04528   0.03635  -0.0613   0.4175   1.0000
   8.250   1.2063   0.04630   0.03757  -0.0602   0.4066   1.0000
   8.500   1.2031   0.04964   0.04119  -0.0580   0.3962   1.0000
   8.750   1.1890   0.05417   0.04595  -0.0558   0.3879   1.0000
   9.000   1.2234   0.05437   0.04641  -0.0550   0.3789   1.0000
   9.250   1.1225   0.06725   0.05908  -0.0523   0.3761   1.0000
   9.500   1.0129   0.08452   0.07575  -0.0574   0.3753   1.0000
   9.750   0.9824   0.09258   0.08373  -0.0599   0.3741   1.0000
  10.000   0.9640   0.09924   0.09038  -0.0619   0.3740   1.0000
  10.250   0.9521   0.10533   0.09650  -0.0637   0.3757   1.0000
  10.500   0.9603   0.11022   0.10152  -0.0649   0.3792   1.0000
  10.750   1.1285   0.08260   0.07539  -0.0447   0.3109   1.0000
  11.000   1.2393   0.04913   0.04051  -0.0231   0.1118   1.0000
  11.250   1.2289   0.05293   0.04423  -0.0221   0.1014   1.0000
  11.500   1.2199   0.05675   0.04794  -0.0215   0.0944   1.0000
  11.750   1.2152   0.06040   0.05167  -0.0210   0.0872   1.0000
  12.000   1.2109   0.06401   0.05520  -0.0206   0.0822   1.0000
  12.250   1.2101   0.06751   0.05889  -0.0201   0.0767   1.0000
  12.500   1.2136   0.07039   0.06171  -0.0188   0.0719   1.0000
  12.750   1.2275   0.07294   0.06456  -0.0167   0.0692   1.0000
  13.000   1.2370   0.07636   0.06834  -0.0154   0.0674   1.0000
  13.250   1.2371   0.08085   0.07315  -0.0153   0.0666   1.0000
  13.500   1.2282   0.08623   0.07884  -0.0164   0.0666   1.0000
  13.750   1.2123   0.09257   0.08547  -0.0187   0.0669   1.0000
  14.000   1.1919   0.09976   0.09290  -0.0221   0.0675   1.0000
  14.250   1.1690   0.10781   0.10116  -0.0264   0.0684   1.0000
  14.500   1.1450   0.11670   0.11020  -0.0315   0.0693   1.0000
  14.750   1.1226   0.12602   0.11962  -0.0369   0.0704   1.0000
  15.000   1.1051   0.13497   0.12862  -0.0418   0.0713   1.0000
<< Back to GOE 380 AIRFOIL (goe380-il)

Polar data table (+)

Polar graphs


<< Back to GOE 380 AIRFOIL (goe380-il)