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

GOE 164 (MVA MK.10) AIRFOIL (goe164-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 164 (MVA MK.10) AIRFOIL (goe164-il)
Reynolds number: 1,000,000
Max Cl/Cd: 112.47 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe164-il-1000000.txt
Download as CSV file: xf-goe164-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 164 (MVA MK.10) 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
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.3296   0.13049   0.12885  -0.0208   1.0000   0.0101
 -10.750  -0.3277   0.12736   0.12573  -0.0214   1.0000   0.0102
 -10.500  -0.3256   0.12435   0.12273  -0.0218   1.0000   0.0102
 -10.250  -0.3231   0.12143   0.11983  -0.0220   1.0000   0.0103
 -10.000  -0.3208   0.11852   0.11693  -0.0221   1.0000   0.0103
  -9.750  -0.3142   0.11505   0.11348  -0.0237   0.9997   0.0103
  -9.500  -0.3005   0.11072   0.10914  -0.0274   0.9986   0.0103
  -9.250  -0.2891   0.10633   0.10475  -0.0302   0.9976   0.0106
  -9.000  -0.2722   0.10308   0.10150  -0.0335   0.9965   0.0108
  -8.750  -0.2556   0.09973   0.09815  -0.0370   0.9955   0.0110
  -8.500  -0.2389   0.09627   0.09469  -0.0408   0.9946   0.0114
  -8.250  -0.2251   0.09277   0.09119  -0.0440   0.9926   0.0120
  -7.000  -0.1055   0.01619   0.01285  -0.1441   0.9679   0.0109
  -6.750  -0.0717   0.01509   0.01162  -0.1459   0.9648   0.0114
  -6.500  -0.0338   0.01400   0.01037  -0.1484   0.9626   0.0119
  -6.250   0.0040   0.01295   0.00914  -0.1509   0.9608   0.0124
  -6.000   0.0408   0.01196   0.00799  -0.1529   0.9589   0.0129
  -5.750   0.0629   0.01132   0.00724  -0.1516   0.9501   0.0133
  -5.500   0.0973   0.01057   0.00635  -0.1529   0.9444   0.0137
  -5.250   0.1264   0.00954   0.00517  -0.1532   0.9304   0.0149
  -5.000   0.1638   0.00916   0.00469  -0.1550   0.8992   0.0161
  -4.750   0.1960   0.00906   0.00420  -0.1556   0.8290   0.0172
  -4.500   0.2211   0.00902   0.00391  -0.1547   0.7958   0.0178
  -4.250   0.2471   0.00859   0.00330  -0.1543   0.7734   0.0197
  -4.000   0.2736   0.00850   0.00310  -0.1538   0.7559   0.0213
  -3.500   0.3278   0.00817   0.00258  -0.1531   0.7263   0.0263
  -3.250   0.3551   0.00811   0.00244  -0.1528   0.7132   0.0290
  -3.000   0.3826   0.00790   0.00215  -0.1526   0.7004   0.0335
  -2.750   0.4099   0.00784   0.00202  -0.1523   0.6874   0.0374
  -2.500   0.4372   0.00770   0.00181  -0.1520   0.6737   0.0437
  -2.250   0.4644   0.00765   0.00168  -0.1516   0.6585   0.0490
  -2.000   0.4913   0.00737   0.00148  -0.1514   0.6405   0.0993
  -1.750   0.5177   0.00739   0.00151  -0.1510   0.6184   0.1378
  -1.500   0.5441   0.00751   0.00154  -0.1506   0.5925   0.1487
  -1.250   0.5704   0.00769   0.00158  -0.1501   0.5669   0.1554
  -1.000   0.5970   0.00780   0.00160  -0.1497   0.5455   0.1617
  -0.750   0.6238   0.00792   0.00164  -0.1493   0.5272   0.1660
  -0.500   0.6507   0.00806   0.00167  -0.1489   0.5097   0.1697
  -0.250   0.6776   0.00814   0.00168  -0.1486   0.4912   0.1738
   0.000   0.7042   0.00827   0.00173  -0.1483   0.4695   0.1784
   0.250   0.7302   0.00850   0.00182  -0.1478   0.4393   0.1835
   0.500   0.7556   0.00878   0.00190  -0.1473   0.4000   0.1858
   0.750   0.7814   0.00897   0.00196  -0.1468   0.3707   0.1900
   1.000   0.8076   0.00914   0.00205  -0.1464   0.3519   0.1935
   1.250   0.8342   0.00928   0.00213  -0.1461   0.3396   0.1972
   1.500   0.8609   0.00943   0.00221  -0.1458   0.3296   0.2000
   1.750   0.8876   0.00952   0.00228  -0.1455   0.3204   0.2049
   2.000   0.9146   0.00961   0.00236  -0.1453   0.3140   0.2094
   2.250   0.9414   0.00972   0.00245  -0.1450   0.3073   0.2135
   2.500   0.9682   0.00981   0.00254  -0.1447   0.3011   0.2189
   2.750   0.9951   0.00989   0.00264  -0.1445   0.2948   0.2253
   3.000   1.0215   0.01003   0.00275  -0.1442   0.2874   0.2330
   3.250   1.0485   0.01010   0.00286  -0.1439   0.2809   0.2436
   3.500   1.0748   0.01022   0.00299  -0.1436   0.2726   0.2623
   3.750   1.1013   0.01028   0.00316  -0.1434   0.2624   0.3241
   4.000   1.1273   0.01039   0.00334  -0.1431   0.2483   0.3956
   4.750   1.1944   0.01062   0.00424  -0.1403   0.1740   1.0000
   5.000   1.2198   0.01085   0.00444  -0.1398   0.1695   1.0000
   5.250   1.2450   0.01110   0.00467  -0.1393   0.1652   1.0000
   5.500   1.2701   0.01134   0.00490  -0.1388   0.1598   1.0000
   5.750   1.2952   0.01158   0.00512  -0.1383   0.1547   1.0000
   6.000   1.3197   0.01186   0.00537  -0.1377   0.1494   1.0000
   6.250   1.3447   0.01208   0.00561  -0.1372   0.1454   1.0000
   6.500   1.3698   0.01229   0.00582  -0.1367   0.1404   1.0000
   6.750   1.3935   0.01262   0.00609  -0.1360   0.1300   1.0000
   7.000   1.4116   0.01346   0.00666  -0.1345   0.0866   1.0000
   7.250   1.4340   0.01388   0.00707  -0.1336   0.0824   1.0000
   7.500   1.4560   0.01432   0.00751  -0.1326   0.0773   1.0000
   7.750   1.4792   0.01463   0.00786  -0.1318   0.0737   1.0000
   8.000   1.5015   0.01501   0.00827  -0.1309   0.0689   1.0000
   8.250   1.5226   0.01547   0.00871  -0.1298   0.0585   1.0000
   8.500   1.5323   0.01687   0.00982  -0.1268   0.0182   1.0000
   8.750   1.5505   0.01752   0.01050  -0.1251   0.0149   1.0000
   9.000   1.5695   0.01805   0.01107  -0.1237   0.0135   1.0000
   9.250   1.5862   0.01872   0.01179  -0.1218   0.0122   1.0000
   9.500   1.5994   0.01952   0.01268  -0.1193   0.0111   1.0000
   9.750   1.6148   0.02008   0.01329  -0.1173   0.0106   1.0000
  10.000   1.6288   0.02072   0.01399  -0.1150   0.0101   1.0000
  10.250   1.6412   0.02144   0.01477  -0.1126   0.0096   1.0000
  10.500   1.6524   0.02223   0.01562  -0.1100   0.0091   1.0000
  10.750   1.6598   0.02328   0.01675  -0.1069   0.0086   1.0000
  11.000   1.6583   0.02492   0.01853  -0.1028   0.0081   1.0000
  11.250   1.6702   0.02571   0.01938  -0.1007   0.0079   1.0000
  11.500   1.6784   0.02678   0.02053  -0.0982   0.0076   1.0000
  11.750   1.6851   0.02800   0.02184  -0.0957   0.0074   1.0000
  12.000   1.6904   0.02939   0.02331  -0.0933   0.0071   1.0000
  12.250   1.6944   0.03096   0.02497  -0.0910   0.0069   1.0000
  12.500   1.6977   0.03270   0.02679  -0.0889   0.0067   1.0000
  12.750   1.6994   0.03472   0.02891  -0.0870   0.0065   1.0000
  13.000   1.6997   0.03701   0.03129  -0.0854   0.0063   1.0000
  13.250   1.6971   0.03979   0.03417  -0.0841   0.0062   1.0000
  13.500   1.6910   0.04313   0.03762  -0.0830   0.0060   1.0000
  13.750   1.6795   0.04727   0.04190  -0.0822   0.0059   1.0000
  14.000   1.6632   0.05226   0.04703  -0.0818   0.0058   1.0000
  14.250   1.6449   0.05772   0.05265  -0.0818   0.0057   1.0000
  14.500   1.6267   0.06341   0.05848  -0.0822   0.0057   1.0000
  14.750   1.6209   0.06758   0.06276  -0.0827   0.0056   1.0000
  15.000   1.6131   0.07210   0.06739  -0.0833   0.0056   1.0000
  15.250   1.6033   0.07700   0.07240  -0.0841   0.0056   1.0000
  15.500   1.5945   0.08189   0.07740  -0.0850   0.0055   1.0000
  15.750   1.5849   0.08693   0.08255  -0.0861   0.0055   1.0000
  16.000   1.5753   0.09202   0.08774  -0.0873   0.0054   1.0000
<< Back to GOE 164 (MVA MK.10) AIRFOIL (goe164-il)

Polar data table (+)

Polar graphs


<< Back to GOE 164 (MVA MK.10) AIRFOIL (goe164-il)