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GOE 9K AIRFOIL (goe09k-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 9K AIRFOIL (goe09k-il)
Reynolds number: 500,000
Max Cl/Cd: 60.55 at α=1°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe09k-il-500000-n5.txt
Download as CSV file: xf-goe09k-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 9K 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.4903   0.08392   0.08177   0.0007   1.0000   0.0036
  -8.000  -0.4899   0.08007   0.07793  -0.0004   1.0000   0.0036
  -7.750  -0.4903   0.07616   0.07404  -0.0014   1.0000   0.0038
  -7.500  -0.4910   0.07238   0.07028  -0.0024   1.0000   0.0038
  -7.000  -0.5686   0.07993   0.07775  -0.0046   1.0000   0.0033
  -6.750  -0.5631   0.07609   0.07390  -0.0065   1.0000   0.0035
  -6.500  -0.5520   0.07194   0.06974  -0.0100   1.0000   0.0036
  -6.250  -0.5393   0.06761   0.06537  -0.0135   1.0000   0.0037
  -6.000  -0.5242   0.06329   0.06101  -0.0168   1.0000   0.0039
  -5.750  -0.5072   0.05884   0.05649  -0.0199   1.0000   0.0040
  -5.500  -0.4885   0.05449   0.05205  -0.0224   1.0000   0.0042
  -5.250  -0.4683   0.05012   0.04755  -0.0245   1.0000   0.0044
  -5.000  -0.4462   0.04584   0.04313  -0.0261   1.0000   0.0047
  -4.750  -0.4208   0.04157   0.03866  -0.0270   1.0000   0.0049
  -4.500  -0.3966   0.03765   0.03454  -0.0271   1.0000   0.0050
  -4.250  -0.3730   0.03394   0.03060  -0.0268   1.0000   0.0050
  -4.000  -0.3552   0.03021   0.02667  -0.0268   1.0000   0.0053
  -3.750  -0.3337   0.02727   0.02350  -0.0263   1.0000   0.0057
  -3.500  -0.3102   0.02435   0.02030  -0.0254   1.0000   0.0062
  -3.250  -0.2834   0.02172   0.01728  -0.0236   1.0000   0.0072
  -3.000  -0.2598   0.01880   0.01397  -0.0224   1.0000   0.0074
  -2.750  -0.2368   0.01646   0.01137  -0.0215   1.0000   0.0078
  -2.500  -0.2107   0.01375   0.00830  -0.0199   1.0000   0.0030
  -2.250  -0.1853   0.01186   0.00610  -0.0187   1.0000   0.0026
  -2.000  -0.1601   0.01037   0.00435  -0.0175   1.0000   0.0025
  -1.750  -0.1355   0.00925   0.00304  -0.0165   1.0000   0.0026
  -1.500  -0.1110   0.00853   0.00220  -0.0156   1.0000   0.0037
  -1.250  -0.0863   0.00800   0.00156  -0.0148   1.0000   0.0060
  -1.000  -0.0590   0.00758   0.00111  -0.0147   0.9993   0.0246
  -0.750  -0.0367   0.00562   0.00104  -0.0147   0.9983   0.6092
  -0.500   0.0061   0.00461   0.00101  -0.0178   1.0000   1.0000
  -0.250   0.0343   0.00464   0.00097  -0.0180   0.9990   1.0000
   0.000   0.0722   0.00468   0.00095  -0.0205   0.9952   1.0000
   0.250   0.1121   0.00467   0.00092  -0.0233   0.9894   1.0000
   0.500   0.1558   0.00461   0.00089  -0.0270   0.9793   1.0000
   0.750   0.1950   0.00453   0.00086  -0.0296   0.9654   1.0000
   1.000   0.2646   0.00437   0.00074  -0.0387   0.8978   1.0000
   1.250   0.2712   0.00652   0.00089  -0.0340   0.3633   1.0000
   1.500   0.2878   0.00820   0.00132  -0.0323   0.0097   1.0000
   1.750   0.3122   0.00871   0.00193  -0.0314   0.0051   1.0000
   2.000   0.3365   0.00929   0.00260  -0.0305   0.0030   1.0000
   2.250   0.3597   0.01021   0.00368  -0.0292   0.0027   1.0000
   2.500   0.3830   0.01145   0.00507  -0.0278   0.0028   1.0000
   2.750   0.4075   0.01301   0.00682  -0.0265   0.0030   1.0000
   7.250   0.7391   0.08060   0.07843  -0.0230   0.0026   1.0000
   7.750   0.7422   0.08933   0.08718  -0.0293   0.0027   1.0000
   8.000   0.7385   0.09339   0.09121  -0.0321   0.0026   1.0000
   8.250   0.7378   0.09744   0.09524  -0.0344   0.0025   1.0000
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