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GOE 118 (MVA MK.7) AIRFOIL (goe118-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 118 (MVA MK.7) AIRFOIL (goe118-il)
Reynolds number: 1,000,000
Max Cl/Cd: 134.06 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe118-il-1000000.txt
Download as CSV file: xf-goe118-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 118 (MVA MK.7) AIRFOIL                      
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
   0.250   0.5989   0.00759   0.00119  -0.1206   0.5772   0.1121
   0.500   0.6269   0.00757   0.00117  -0.1207   0.5688   0.1183
   0.750   0.6547   0.00760   0.00117  -0.1208   0.5606   0.1244
   1.250   0.7106   0.00750   0.00122  -0.1213   0.5459   0.1929
   1.500   0.7383   0.00698   0.00137  -0.1218   0.5390   0.4906
   2.000   0.7897   0.00607   0.00154  -0.1212   0.5257   1.0000
   2.250   0.8172   0.00618   0.00161  -0.1213   0.5159   1.0000
   2.500   0.8444   0.00633   0.00169  -0.1213   0.4992   1.0000
   2.750   0.8714   0.00650   0.00177  -0.1213   0.4826   1.0000
   3.000   0.8974   0.00678   0.00190  -0.1212   0.4510   1.0000
   3.250   0.9230   0.00711   0.00206  -0.1210   0.4147   1.0000
   3.500   0.9444   0.00793   0.00246  -0.1202   0.3328   1.0000
   3.750   0.9545   0.01014   0.00362  -0.1179   0.1199   1.0000
   4.000   0.9750   0.01108   0.00425  -0.1169   0.0231   1.0000
   4.250   0.9992   0.01158   0.00484  -0.1163   0.0160   1.0000
   4.500   1.0251   0.01182   0.00507  -0.1161   0.0144   1.0000
   4.750   1.0496   0.01221   0.00553  -0.1156   0.0126   1.0000
   5.000   1.0736   0.01266   0.00602  -0.1151   0.0112   1.0000
   5.250   1.0933   0.01358   0.00705  -0.1137   0.0094   1.0000
   5.500   1.1038   0.01534   0.00899  -0.1108   0.0087   1.0000
   5.750   1.1305   0.01541   0.00906  -0.1108   0.0082   1.0000
   6.000   1.1498   0.01615   0.00986  -0.1095   0.0077   1.0000
   6.250   1.1665   0.01709   0.01086  -0.1078   0.0073   1.0000
   6.500   1.1830   0.01798   0.01183  -0.1060   0.0067   1.0000
   6.750   1.1996   0.01883   0.01272  -0.1044   0.0061   1.0000
   7.000   1.2150   0.01975   0.01368  -0.1026   0.0056   1.0000
   7.250   1.2261   0.02106   0.01503  -0.1000   0.0054   1.0000
   7.750   1.2933   0.03386   0.02804  -0.0998   0.0087   1.0000
   8.000   1.3113   0.03508   0.02944  -0.0982   0.0085   1.0000
   8.250   1.3279   0.03293   0.02737  -0.0962   0.0072   1.0000
   8.500   1.3478   0.03432   0.02892  -0.0949   0.0064   1.0000
   8.750   1.3651   0.03603   0.03076  -0.0936   0.0059   1.0000
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