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

GOE 165 (MVA MK.11) AIRFOIL (goe165-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 165 (MVA MK.11) AIRFOIL (goe165-il)
Reynolds number: 1,000,000
Max Cl/Cd: 96.57 at α=3.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe165-il-1000000.txt
Download as CSV file: xf-goe165-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 165 (MVA MK.11) 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
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.1909   0.08243   0.08097  -0.0522   0.9926   0.0106
  -9.250  -0.1871   0.07639   0.07494  -0.0554   0.9910   0.0106
  -9.000  -0.1857   0.06986   0.06841  -0.0586   0.9895   0.0107
  -8.750  -0.2111   0.05740   0.05598  -0.0637   0.9880   0.0110
  -8.500  -0.2311   0.04457   0.04308  -0.0721   0.9861   0.0109
  -8.250  -0.2276   0.03903   0.03746  -0.0786   0.9826   0.0108
  -6.000   0.0258   0.01341   0.00919  -0.1570   0.9023   0.0103
  -5.750   0.0615   0.01219   0.00758  -0.1588   0.8439   0.0102
  -5.500   0.0828   0.01195   0.00691  -0.1574   0.7652   0.0104
  -5.250   0.1072   0.01179   0.00649  -0.1566   0.7274   0.0107
  -5.000   0.1342   0.01055   0.00506  -0.1569   0.7071   0.0115
  -4.750   0.1619   0.01012   0.00450  -0.1571   0.6918   0.0121
  -4.500   0.1899   0.00978   0.00404  -0.1572   0.6791   0.0126
  -4.250   0.2182   0.00950   0.00363  -0.1572   0.6685   0.0132
  -4.000   0.2466   0.00925   0.00329  -0.1573   0.6589   0.0141
  -3.750   0.2751   0.00905   0.00299  -0.1574   0.6510   0.0149
  -3.500   0.3038   0.00879   0.00260  -0.1574   0.6425   0.0162
  -3.250   0.3323   0.00862   0.00235  -0.1574   0.6329   0.0184
  -3.000   0.3605   0.00852   0.00216  -0.1573   0.6229   0.0220
  -2.750   0.3888   0.00838   0.00204  -0.1574   0.6130   0.0483
  -2.500   0.4172   0.00831   0.00194  -0.1573   0.6045   0.0569
  -2.250   0.4453   0.00816   0.00183  -0.1574   0.5948   0.0944
  -2.000   0.4737   0.00809   0.00175  -0.1574   0.5845   0.1081
  -1.750   0.5021   0.00774   0.00158  -0.1577   0.5719   0.1977
  -1.500   0.5294   0.00765   0.00151  -0.1576   0.5457   0.2399
  -1.250   0.5535   0.00786   0.00152  -0.1569   0.4664   0.2908
  -1.000   0.5780   0.00824   0.00169  -0.1563   0.4101   0.3212
  -0.750   0.6045   0.00845   0.00180  -0.1560   0.3863   0.3325
  -0.500   0.6316   0.00862   0.00187  -0.1557   0.3698   0.3405
  -0.250   0.6588   0.00877   0.00196  -0.1555   0.3562   0.3478
   0.000   0.6860   0.00893   0.00204  -0.1552   0.3439   0.3532
   0.250   0.7134   0.00905   0.00212  -0.1550   0.3339   0.3585
   0.500   0.7408   0.00918   0.00221  -0.1548   0.3249   0.3634
   0.750   0.7679   0.00934   0.00229  -0.1546   0.3162   0.3678
   1.000   0.7955   0.00943   0.00239  -0.1544   0.3101   0.3735
   1.250   0.8227   0.00956   0.00249  -0.1542   0.3035   0.3779
   1.500   0.8500   0.00969   0.00259  -0.1540   0.2976   0.3817
   1.750   0.8773   0.00979   0.00268  -0.1538   0.2914   0.3864
   2.000   0.9043   0.00994   0.00280  -0.1536   0.2844   0.3897
   2.250   0.9315   0.01005   0.00290  -0.1534   0.2777   0.3928
   2.500   0.9583   0.01021   0.00301  -0.1531   0.2691   0.3953
   2.750   0.9848   0.01039   0.00312  -0.1528   0.2577   0.3974
   3.000   1.0114   0.01051   0.00323  -0.1525   0.2465   0.4003
   3.250   1.0372   0.01074   0.00339  -0.1521   0.2302   0.4028
   3.500   1.0612   0.01114   0.00362  -0.1514   0.2007   0.4055
   3.750   1.0768   0.01242   0.00440  -0.1493   0.1085   0.4080
   4.000   1.1011   0.01279   0.00468  -0.1487   0.0922   0.4104
   4.250   1.1235   0.01336   0.00507  -0.1477   0.0592   0.4119
   4.500   1.1491   0.01356   0.00530  -0.1473   0.0537   0.4143
   4.750   1.1745   0.01377   0.00556  -0.1468   0.0520   0.4171
   5.000   1.1997   0.01401   0.00584  -0.1463   0.0502   0.4202
   5.250   1.2243   0.01430   0.00616  -0.1457   0.0477   0.4229
   5.500   1.2487   0.01460   0.00649  -0.1450   0.0439   0.4248
   5.750   1.2741   0.01477   0.00666  -0.1446   0.0419   0.4266
   6.000   1.2982   0.01506   0.00692  -0.1439   0.0336   0.4293
   6.250   1.3201   0.01557   0.00735  -0.1429   0.0141   0.4315
   6.500   1.3431   0.01596   0.00782  -0.1420   0.0124   0.4337
   6.750   1.3654   0.01641   0.00834  -0.1409   0.0107   0.4360
   7.000   1.3873   0.01686   0.00886  -0.1398   0.0099   0.4382
   7.250   1.4092   0.01730   0.00935  -0.1388   0.0094   0.4404
   7.500   1.4303   0.01778   0.00991  -0.1375   0.0089   0.4428
   7.750   1.4504   0.01832   0.01053  -0.1362   0.0084   0.4455
   8.000   1.4694   0.01892   0.01120  -0.1347   0.0078   0.4481
   8.250   1.4849   0.01977   0.01214  -0.1326   0.0072   0.4509
   8.500   1.4934   0.02106   0.01358  -0.1293   0.0067   0.4534
   8.750   1.5101   0.02161   0.01418  -0.1274   0.0066   0.4560
   9.000   1.5217   0.02227   0.01493  -0.1246   0.0064   0.4592
   9.250   1.5326   0.02307   0.01583  -0.1218   0.0062   0.4621
   9.500   1.5417   0.02398   0.01683  -0.1188   0.0060   0.4651
   9.750   1.5495   0.02497   0.01791  -0.1157   0.0058   0.4681
  10.000   1.5553   0.02610   0.01913  -0.1125   0.0056   0.4709
  10.250   1.5599   0.02734   0.02046  -0.1094   0.0055   0.4743
  10.500   1.5637   0.02869   0.02190  -0.1063   0.0053   0.4777
  10.750   1.5661   0.03023   0.02353  -0.1034   0.0051   0.4812
  11.000   1.5666   0.03202   0.02542  -0.1007   0.0050   0.4845
  11.250   1.5665   0.03402   0.02753  -0.0983   0.0049   0.4883
  11.500   1.5631   0.03652   0.03013  -0.0962   0.0048   0.4922
  11.750   1.5567   0.03955   0.03327  -0.0944   0.0047   0.4960
  12.000   1.5475   0.04320   0.03704  -0.0931   0.0046   0.4996
  12.250   1.5345   0.04753   0.04149  -0.0922   0.0045   0.5033
  12.500   1.5170   0.05243   0.04650  -0.0912   0.0044   0.5065
  12.750   1.5099   0.05622   0.05040  -0.0906   0.0043   0.5117
  13.000   1.5116   0.05925   0.05356  -0.0908   0.0042   0.5196
  13.250   1.5113   0.06256   0.05699  -0.0909   0.0042   0.5292
  13.750   1.5045   0.06987   0.06458  -0.0911   0.0040   0.5708
  14.000   1.5013   0.07295   0.06840  -0.0911   0.0040   1.0000
  14.250   1.4970   0.07676   0.07229  -0.0911   0.0039   1.0000
  14.500   1.4932   0.08054   0.07615  -0.0912   0.0039   1.0000
  14.750   1.4897   0.08428   0.07998  -0.0912   0.0039   1.0000
<< Back to GOE 165 (MVA MK.11) AIRFOIL (goe165-il)

Polar data table (+)

Polar graphs


<< Back to GOE 165 (MVA MK.11) AIRFOIL (goe165-il)