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GOE 507 AIRFOIL (goe507-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 507 AIRFOIL (goe507-il)
Reynolds number: 50,000
Max Cl/Cd: 37.65 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe507-il-50000.txt
Download as CSV file: xf-goe507-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 507 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4037   0.09850   0.09124  -0.0191   1.0000   0.2610
  -8.000  -0.3987   0.09499   0.08779  -0.0186   1.0000   0.2685
  -7.750  -0.4163   0.09422   0.08717  -0.0183   1.0000   0.2783
  -7.500  -0.4018   0.09023   0.08318  -0.0168   1.0000   0.2918
  -7.250  -0.3971   0.08717   0.08019  -0.0153   1.0000   0.3045
  -7.000  -0.4383   0.08773   0.08102  -0.0143   1.0000   0.3106
  -6.750  -0.4348   0.08480   0.07815  -0.0120   1.0000   0.3261
  -6.500  -0.4330   0.08200   0.07542  -0.0096   1.0000   0.3416
  -6.250  -0.4323   0.07934   0.07281  -0.0070   1.0000   0.3571
  -6.000  -0.4401   0.07735   0.07092  -0.0043   1.0000   0.3733
  -5.750  -0.4192   0.07363   0.06720  -0.0001   1.0000   0.3983
  -5.500  -0.4434   0.07292   0.06663   0.0035   1.0000   0.4208
  -5.250  -0.4403   0.07056   0.06434   0.0084   1.0000   0.4520
  -4.500  -0.3048   0.05928   0.05277   0.0192   1.0000   0.6861
  -4.250  -0.1618   0.05097   0.04401   0.0063   1.0000   0.8486
  -4.000  -0.4087   0.05021   0.04328  -0.0154   1.0000   0.3143
  -3.750  -0.3553   0.04052   0.03177  -0.0291   1.0000   0.2006
  -3.500  -0.3352   0.03787   0.02889  -0.0288   1.0000   0.1970
  -3.250  -0.3137   0.03564   0.02626  -0.0286   1.0000   0.1970
  -3.000  -0.2913   0.03374   0.02392  -0.0284   1.0000   0.1990
  -2.750  -0.2682   0.03199   0.02178  -0.0280   1.0000   0.1993
  -2.500  -0.2443   0.03049   0.01986  -0.0276   1.0000   0.2004
  -2.250  -0.2216   0.02919   0.01837  -0.0270   1.0000   0.2041
  -2.000  -0.1996   0.02827   0.01732  -0.0263   1.0000   0.2118
  -1.750  -0.1763   0.02736   0.01613  -0.0257   1.0000   0.2186
  -1.500  -0.1529   0.02650   0.01521  -0.0252   1.0000   0.2259
  -1.250  -0.1301   0.02575   0.01431  -0.0244   1.0000   0.2358
  -1.000  -0.1060   0.02512   0.01365  -0.0239   1.0000   0.2522
  -0.750  -0.0803   0.02446   0.01307  -0.0237   1.0000   0.2831
  -0.500  -0.0535   0.02319   0.01235  -0.0235   1.0000   0.3660
  -0.250  -0.0119   0.02062   0.01164  -0.0253   1.0000   1.0000
   0.000   0.0072   0.02104   0.01171  -0.0245   1.0000   1.0000
   0.250   0.0254   0.02151   0.01191  -0.0237   1.0000   1.0000
   0.500   0.0432   0.02203   0.01223  -0.0230   1.0000   1.0000
   0.750   0.0605   0.02260   0.01263  -0.0223   1.0000   1.0000
   1.000   0.0774   0.02324   0.01313  -0.0218   1.0000   1.0000
   1.250   0.0939   0.02393   0.01371  -0.0212   1.0000   1.0000
   1.500   0.1172   0.02484   0.01452  -0.0222   0.9968   1.0000
   1.750   0.1716   0.02636   0.01594  -0.0289   0.9788   1.0000
   2.000   0.2217   0.02769   0.01721  -0.0345   0.9602   1.0000
   2.250   0.2735   0.02897   0.01847  -0.0401   0.9417   1.0000
   2.500   0.3163   0.02997   0.01947  -0.0439   0.9216   1.0000
   2.750   0.3587   0.03086   0.02041  -0.0473   0.9009   1.0000
   3.000   0.4078   0.03167   0.02128  -0.0514   0.8806   1.0000
   3.250   0.4473   0.03232   0.02201  -0.0537   0.8590   1.0000
   3.500   0.4882   0.03283   0.02264  -0.0559   0.8374   1.0000
   3.750   0.5401   0.03305   0.02302  -0.0593   0.8176   1.0000
   4.000   0.5709   0.03348   0.02359  -0.0596   0.7953   1.0000
   4.250   0.6165   0.03345   0.02376  -0.0615   0.7747   1.0000
   4.500   0.6712   0.03291   0.02347  -0.0640   0.7560   1.0000
   4.750   0.6970   0.03317   0.02391  -0.0629   0.7327   1.0000
   5.000   0.7572   0.03118   0.02220  -0.0637   0.7102   1.0000
   5.250   0.8069   0.02859   0.01984  -0.0619   0.6790   1.0000
   5.500   0.8387   0.02719   0.01857  -0.0589   0.6455   1.0000
   5.750   0.8697   0.02578   0.01723  -0.0558   0.6081   1.0000
   6.000   0.8953   0.02479   0.01624  -0.0523   0.5647   1.0000
   6.250   0.9164   0.02446   0.01582  -0.0488   0.5158   1.0000
   6.500   0.9276   0.02464   0.01575  -0.0440   0.4461   1.0000
   6.750   0.9250   0.02578   0.01613  -0.0376   0.3376   1.0000
   7.000   0.9318   0.02811   0.01752  -0.0338   0.2547   1.0000
   7.250   0.9517   0.03009   0.01917  -0.0322   0.2149   1.0000
   7.500   0.9775   0.03204   0.02091  -0.0315   0.1908   1.0000
   7.750   1.0029   0.03398   0.02286  -0.0307   0.1728   1.0000
   8.000   1.0288   0.03613   0.02507  -0.0301   0.1594   1.0000
   8.250   1.0548   0.03855   0.02749  -0.0295   0.1491   1.0000
   8.500   1.0762   0.04100   0.03022  -0.0284   0.1408   1.0000
   8.750   1.0954   0.04391   0.03337  -0.0272   0.1345   1.0000
   9.000   1.1092   0.04716   0.03711  -0.0252   0.1310   1.0000
   9.250   1.1238   0.05025   0.04045  -0.0237   0.1269   1.0000
   9.500   1.1381   0.05396   0.04431  -0.0225   0.1230   1.0000
   9.750   1.1379   0.05780   0.04868  -0.0198   0.1222   1.0000
  10.000   1.1368   0.06207   0.05336  -0.0175   0.1223   1.0000
  10.250   1.1310   0.06650   0.05815  -0.0151   0.1226   1.0000
  10.500   1.1225   0.07102   0.06297  -0.0130   0.1231   1.0000
  10.750   1.1124   0.07570   0.06787  -0.0112   0.1235   1.0000
  11.000   1.0149   0.08155   0.07421  -0.0081   0.1302   1.0000
  11.250   0.9752   0.08902   0.08176  -0.0107   0.1333   1.0000
  11.500   0.9546   0.09642   0.08918  -0.0136   0.1358   1.0000
  11.750   0.8630   0.12051   0.11308  -0.0327   0.1625   1.0000
  12.000   0.8763   0.12541   0.11807  -0.0324   0.1654   1.0000
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