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

GOE 6K AIRFOIL (goe06k-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 6K AIRFOIL (goe06k-il)
Reynolds number: 50,000
Max Cl/Cd: 28.32 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe06k-il-50000-n5.txt
Download as CSV file: xf-goe06k-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 6K 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.4643   0.11390   0.10733  -0.0328   1.0000   0.0823
  -8.250  -0.4792   0.11266   0.10622  -0.0325   1.0000   0.0827
  -8.000  -0.4931   0.11116   0.10484  -0.0325   1.0000   0.0829
  -7.750  -0.4756   0.10471   0.09838  -0.0290   1.0000   0.0853
  -7.500  -0.4771   0.10172   0.09545  -0.0271   1.0000   0.0876
  -7.250  -0.4831   0.09923   0.09301  -0.0259   1.0000   0.0893
  -7.000  -0.4885   0.09659   0.09043  -0.0256   1.0000   0.0917
  -6.750  -0.4965   0.09417   0.08808  -0.0270   1.0000   0.0954
  -6.500  -0.5060   0.09259   0.08648  -0.0319   1.0000   0.0974
  -6.250  -0.5010   0.08762   0.08160  -0.0279   1.0000   0.0995
  -6.000  -0.4978   0.08430   0.07831  -0.0257   1.0000   0.1040
  -5.750  -0.4984   0.08229   0.07608  -0.0315   1.0000   0.1121
  -5.500  -0.4937   0.07767   0.07163  -0.0273   1.0000   0.1153
  -5.250  -0.4889   0.07472   0.06856  -0.0293   1.0000   0.1276
  -5.000  -0.4826   0.07119   0.06509  -0.0264   1.0000   0.1342
  -4.500  -0.4346   0.05988   0.05299  -0.0325   1.0000   0.0539
  -4.000  -0.3980   0.05185   0.04424  -0.0329   1.0000   0.0431
  -3.750  -0.3811   0.04839   0.04052  -0.0325   1.0000   0.0413
  -3.500  -0.3612   0.04501   0.03673  -0.0321   1.0000   0.0397
  -3.250  -0.3398   0.04187   0.03308  -0.0315   1.0000   0.0388
  -3.000  -0.3185   0.03930   0.03004  -0.0306   1.0000   0.0398
  -2.750  -0.2963   0.03697   0.02720  -0.0297   1.0000   0.0409
  -2.500  -0.2730   0.03472   0.02438  -0.0285   1.0000   0.0420
  -2.250  -0.2493   0.03272   0.02183  -0.0272   1.0000   0.0431
  -2.000  -0.2258   0.03102   0.01966  -0.0261   1.0000   0.0458
  -1.750  -0.2023   0.02953   0.01792  -0.0253   1.0000   0.0503
  -1.500  -0.1760   0.02834   0.01623  -0.0243   1.0000   0.0563
  -1.250  -0.1445   0.02718   0.01485  -0.0248   0.9982   0.0665
  -1.000  -0.1118   0.02628   0.01368  -0.0257   0.9962   0.0847
  -0.750  -0.0795   0.02540   0.01277  -0.0268   0.9942   0.1275
  -0.500  -0.0250   0.02230   0.01223  -0.0324   0.9959   1.0000
  -0.250   0.0047   0.02269   0.01204  -0.0332   0.9922   1.0000
   0.000   0.0349   0.02317   0.01208  -0.0342   0.9885   1.0000
   0.250   0.0635   0.02358   0.01216  -0.0350   0.9845   1.0000
   0.500   0.0927   0.02404   0.01232  -0.0360   0.9805   1.0000
   0.750   0.1213   0.02449   0.01255  -0.0368   0.9763   1.0000
   1.000   0.1494   0.02492   0.01278  -0.0376   0.9716   1.0000
   1.250   0.1802   0.02547   0.01318  -0.0389   0.9676   1.0000
   1.500   0.2058   0.02582   0.01344  -0.0392   0.9618   1.0000
   1.750   0.2371   0.02636   0.01390  -0.0406   0.9569   1.0000
   2.000   0.2640   0.02675   0.01426  -0.0411   0.9505   1.0000
   2.250   0.2951   0.02725   0.01475  -0.0425   0.9443   1.0000
   2.500   0.3240   0.02769   0.01523  -0.0433   0.9375   1.0000
   2.750   0.3541   0.02816   0.01575  -0.0445   0.9306   1.0000
   3.000   0.3833   0.02861   0.01630  -0.0454   0.9236   1.0000
   3.250   0.4132   0.02907   0.01691  -0.0464   0.9162   1.0000
   3.500   0.4413   0.02951   0.01751  -0.0471   0.9083   1.0000
   3.750   0.4730   0.02997   0.01821  -0.0484   0.9007   1.0000
   4.000   0.4990   0.03037   0.01883  -0.0486   0.8912   1.0000
   4.250   0.5361   0.03081   0.01958  -0.0508   0.8839   1.0000
   4.500   0.5609   0.03115   0.02026  -0.0506   0.8724   1.0000
   4.750   0.5889   0.03147   0.02094  -0.0509   0.8605   1.0000
   5.000   0.6198   0.03167   0.02161  -0.0515   0.8474   1.0000
   5.250   0.6808   0.02964   0.02037  -0.0540   0.8113   1.0000
   5.500   0.7301   0.02578   0.01330  -0.0451   0.1296   1.0000
   5.750   0.7391   0.02824   0.01523  -0.0422   0.0712   1.0000
   6.000   0.7554   0.03014   0.01715  -0.0401   0.0550   1.0000
   6.250   0.7802   0.03191   0.01913  -0.0394   0.0453   1.0000
   6.500   0.8307   0.03436   0.02184  -0.0427   0.0369   1.0000
   6.750   0.8939   0.03810   0.02616  -0.0475   0.0330   1.0000
   7.000   0.9278   0.04128   0.02977  -0.0480   0.0300   1.0000
   7.250   0.9541   0.04476   0.03373  -0.0473   0.0293   1.0000
   7.500   0.9723   0.04830   0.03775  -0.0453   0.0290   1.0000
   7.750   0.9845   0.05166   0.04157  -0.0428   0.0286   1.0000
   8.000   0.9926   0.05518   0.04544  -0.0400   0.0281   1.0000
   8.250   0.9954   0.05907   0.04961  -0.0369   0.0275   1.0000
   8.500   0.9954   0.06231   0.05327  -0.0331   0.0273   1.0000
   8.750   0.9928   0.06565   0.05697  -0.0293   0.0273   1.0000
   9.000   0.9875   0.06904   0.06066  -0.0255   0.0273   1.0000
   9.250   0.9816   0.07256   0.06441  -0.0222   0.0274   1.0000
   9.750   0.9524   0.07828   0.07072  -0.0136   0.0282   1.0000
  10.000   0.9306   0.08154   0.07425  -0.0101   0.0289   1.0000
  10.250   0.9106   0.08544   0.07831  -0.0082   0.0293   1.0000
  10.500   0.8905   0.08987   0.08288  -0.0077   0.0297   1.0000
  10.750   0.8706   0.09504   0.08815  -0.0087   0.0306   1.0000
  11.000   0.8535   0.10078   0.09394  -0.0111   0.0312   1.0000
<< Back to GOE 6K AIRFOIL (goe06k-il)

Polar data table (+)

Polar graphs


<< Back to GOE 6K AIRFOIL (goe06k-il)