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GOE 9K AIRFOIL (goe09k-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 9K AIRFOIL (goe09k-il)
Reynolds number: 1,000,000
Max Cl/Cd: 77.65 at α=1°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe09k-il-1000000.txt
Download as CSV file: xf-goe09k-il-1000000.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 =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.000  -0.5682   0.07949   0.07793  -0.0048   1.0000   0.0019
  -6.750  -0.5652   0.07539   0.07383  -0.0064   1.0000   0.0021
  -6.500  -0.5523   0.07098   0.06940  -0.0104   1.0000   0.0021
  -6.250  -0.5403   0.06677   0.06516  -0.0138   1.0000   0.0023
  -6.000  -0.5250   0.06246   0.06080  -0.0171   1.0000   0.0025
  -5.750  -0.5078   0.05797   0.05626  -0.0200   1.0000   0.0026
  -5.500  -0.4881   0.05357   0.05177  -0.0225   1.0000   0.0031
  -5.250  -0.4654   0.04904   0.04713  -0.0246   1.0000   0.0032
  -5.000  -0.4423   0.04469   0.04263  -0.0259   1.0000   0.0032
  -4.750  -0.4197   0.04062   0.03842  -0.0264   1.0000   0.0033
  -4.500  -0.3979   0.03683   0.03446  -0.0263   1.0000   0.0033
  -4.250  -0.3800   0.03259   0.03004  -0.0265   1.0000   0.0034
  -4.000  -0.3600   0.02957   0.02684  -0.0261   1.0000   0.0036
  -3.750  -0.3379   0.02648   0.02354  -0.0253   1.0000   0.0039
  -3.500  -0.3148   0.02347   0.02029  -0.0242   1.0000   0.0043
  -3.250  -0.2894   0.02068   0.01722  -0.0224   1.0000   0.0049
  -2.000  -0.1697   0.01065   0.00580  -0.0165   1.0000   0.0109
  -1.500  -0.1182   0.00773   0.00254  -0.0141   1.0000   0.0042
  -1.250  -0.0932   0.00696   0.00164  -0.0132   1.0000   0.0047
  -1.000  -0.0690   0.00660   0.00121  -0.0124   1.0000   0.0075
  -0.750  -0.0400   0.00512   0.00090  -0.0133   0.9989   0.4119
  -0.500  -0.0001   0.00349   0.00096  -0.0163   1.0000   1.0000
   0.000   0.0789   0.00354   0.00091  -0.0219   0.9952   1.0000
   0.250   0.1190   0.00350   0.00086  -0.0248   0.9910   1.0000
   0.500   0.1672   0.00339   0.00078  -0.0295   0.9834   1.0000
   0.750   0.2079   0.00330   0.00072  -0.0324   0.9740   1.0000
   1.000   0.2710   0.00349   0.00050  -0.0400   0.8012   1.0000
   1.250   0.2731   0.00576   0.00075  -0.0350   0.2326   1.0000
   1.500   0.2925   0.00701   0.00130  -0.0334   0.0066   1.0000
   1.750   0.3164   0.00758   0.00195  -0.0323   0.0042   1.0000
   5.750   0.6629   0.05108   0.04895  -0.0119   0.0021   1.0000
   6.000   0.6794   0.05563   0.05365  -0.0122   0.0020   1.0000
   6.250   0.6942   0.06023   0.05837  -0.0129   0.0020   1.0000
   6.500   0.7085   0.06509   0.06334  -0.0143   0.0020   1.0000
   6.750   0.7207   0.07004   0.06838  -0.0164   0.0020   1.0000
   7.000   0.7304   0.07501   0.07340  -0.0191   0.0020   1.0000
   7.250   0.7375   0.07976   0.07819  -0.0221   0.0019   1.0000
   7.500   0.7409   0.08427   0.08274  -0.0252   0.0020   1.0000
   7.750   0.7412   0.08855   0.08702  -0.0285   0.0019   1.0000
   8.000   0.7365   0.09256   0.09101  -0.0313   0.0019   1.0000
   8.250   0.7362   0.09663   0.09506  -0.0337   0.0019   1.0000
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