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GOE 117 (MVA MK.4) AIRFOIL (goe117-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 117 (MVA MK.4) AIRFOIL (goe117-il)
Reynolds number: 500,000
Max Cl/Cd: 107.84 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe117-il-500000.txt
Download as CSV file: xf-goe117-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 117 (MVA MK.4) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.4826   0.00861   0.00177  -0.1018   0.6031   0.3428
   0.500   0.5092   0.00855   0.00181  -0.1016   0.5977   0.3713
   0.750   0.5353   0.00845   0.00187  -0.1014   0.5924   0.4243
   1.000   0.5522   0.00759   0.00200  -0.0993   0.5877   0.7843
   1.250   0.6103   0.00733   0.00207  -0.1058   0.5813   1.0000
   1.500   0.6360   0.00744   0.00213  -0.1054   0.5762   1.0000
   1.750   0.6620   0.00758   0.00222  -0.1050   0.5719   1.0000
   2.000   0.6881   0.00766   0.00231  -0.1046   0.5675   1.0000
   2.250   0.7141   0.00776   0.00239  -0.1043   0.5632   1.0000
   2.500   0.7401   0.00793   0.00250  -0.1039   0.5584   1.0000
   2.750   0.7656   0.00797   0.00259  -0.1034   0.5511   1.0000
   3.000   0.7914   0.00811   0.00269  -0.1030   0.5454   1.0000
   3.250   0.8167   0.00816   0.00278  -0.1025   0.5365   1.0000
   3.500   0.8416   0.00827   0.00287  -0.1019   0.5269   1.0000
   3.750   0.8663   0.00837   0.00296  -0.1012   0.5154   1.0000
   4.000   0.8907   0.00844   0.00304  -0.1005   0.4999   1.0000
   4.250   0.9151   0.00853   0.00313  -0.0998   0.4808   1.0000
   4.500   0.9382   0.00870   0.00323  -0.0989   0.4501   1.0000
   4.750   0.9514   0.00959   0.00357  -0.0963   0.3376   1.0000
   5.000   0.9441   0.01249   0.00514  -0.0909   0.0700   1.0000
   5.250   0.9623   0.01324   0.00576  -0.0892   0.0416   1.0000
   5.500   0.9828   0.01378   0.00634  -0.0879   0.0344   1.0000
   5.750   0.9989   0.01464   0.00726  -0.0859   0.0280   1.0000
   6.000   1.0198   0.01508   0.00775  -0.0847   0.0257   1.0000
   6.250   1.0382   0.01568   0.00841  -0.0830   0.0233   1.0000
   6.500   1.0538   0.01643   0.00920  -0.0810   0.0215   1.0000
   6.750   1.0559   0.01799   0.01089  -0.0767   0.0199   1.0000
   7.000   1.0741   0.01846   0.01141  -0.0751   0.0188   1.0000
   7.250   1.0870   0.01913   0.01213  -0.0726   0.0178   1.0000
   7.500   1.0959   0.02002   0.01309  -0.0695   0.0170   1.0000
   7.750   1.1059   0.02092   0.01405  -0.0667   0.0162   1.0000
   8.000   1.1159   0.02191   0.01509  -0.0640   0.0156   1.0000
   8.250   1.1262   0.02300   0.01623  -0.0615   0.0150   1.0000
   8.500   1.1368   0.02456   0.01782  -0.0592   0.0144   1.0000
   8.750   1.1679   0.02783   0.02120  -0.0600   0.0136   1.0000
   9.000   1.1823   0.02853   0.02203  -0.0580   0.0131   1.0000
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