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GOE 11K AIRFOIL (goe11k-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 11K AIRFOIL (goe11k-il)
Reynolds number: 50,000
Max Cl/Cd: 31.58 at α=5.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe11k-il-50000-n5.txt
Download as CSV file: xf-goe11k-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 11K 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.4716   0.10596   0.09897  -0.0315   1.0000   0.0578
  -8.500  -0.4738   0.10278   0.09584  -0.0309   1.0000   0.0559
  -7.500  -0.5214   0.08856   0.08200  -0.0343   1.0000   0.0477
  -7.250  -0.5241   0.08485   0.07830  -0.0342   1.0000   0.0476
  -7.000  -0.5256   0.08112   0.07457  -0.0340   1.0000   0.0475
  -6.750  -0.5267   0.07698   0.07038  -0.0343   1.0000   0.0475
  -6.500  -0.5260   0.07274   0.06606  -0.0345   1.0000   0.0475
  -6.250  -0.5231   0.06867   0.06189  -0.0344   1.0000   0.0474
  -6.000  -0.5185   0.06439   0.05743  -0.0343   1.0000   0.0473
  -5.750  -0.5121   0.05989   0.05271  -0.0342   1.0000   0.0473
  -5.500  -0.5040   0.05579   0.04839  -0.0339   1.0000   0.0485
  -5.250  -0.4936   0.05376   0.04632  -0.0326   1.0000   0.0512
  -5.000  -0.4811   0.05018   0.04244  -0.0321   1.0000   0.0532
  -4.750  -0.4666   0.04617   0.03798  -0.0316   1.0000   0.0546
  -4.500  -0.4500   0.04225   0.03344  -0.0309   1.0000   0.0572
  -4.250  -0.4319   0.03878   0.02929  -0.0300   1.0000   0.0610
  -4.000  -0.4133   0.03669   0.02689  -0.0289   1.0000   0.0656
  -3.750  -0.3926   0.03416   0.02371  -0.0277   1.0000   0.0722
  -3.500  -0.3717   0.03245   0.02156  -0.0265   1.0000   0.0801
  -3.250  -0.3508   0.03096   0.01984  -0.0255   1.0000   0.0877
  -3.000  -0.3292   0.02965   0.01820  -0.0244   1.0000   0.0981
  -2.750  -0.3064   0.02855   0.01670  -0.0233   1.0000   0.1073
  -2.500  -0.2840   0.02770   0.01571  -0.0225   1.0000   0.1176
  -2.250  -0.2588   0.02681   0.01444  -0.0218   1.0000   0.1236
  -2.000  -0.2343   0.02608   0.01360  -0.0211   1.0000   0.1295
  -1.750  -0.2093   0.02564   0.01294  -0.0207   1.0000   0.1400
  -1.500  -0.1850   0.02524   0.01240  -0.0202   1.0000   0.1510
  -1.250  -0.1607   0.02481   0.01187  -0.0196   1.0000   0.1594
  -1.000  -0.1368   0.02445   0.01145  -0.0191   1.0000   0.1726
  -0.750  -0.1126   0.02408   0.01112  -0.0187   1.0000   0.1958
  -0.500  -0.0508   0.02139   0.01114  -0.0260   1.0000   1.0000
  -0.250  -0.0298   0.02163   0.01100  -0.0249   1.0000   1.0000
   0.000  -0.0090   0.02189   0.01097  -0.0239   1.0000   1.0000
   0.250   0.0118   0.02216   0.01099  -0.0230   1.0000   1.0000
   0.500   0.0325   0.02245   0.01109  -0.0220   1.0000   1.0000
   0.750   0.0530   0.02277   0.01123  -0.0211   1.0000   1.0000
   1.000   0.0774   0.02316   0.01148  -0.0211   0.9984   1.0000
   1.250   0.1075   0.02372   0.01193  -0.0222   0.9948   1.0000
   1.500   0.1367   0.02420   0.01233  -0.0232   0.9905   1.0000
   1.750   0.1675   0.02483   0.01289  -0.0245   0.9866   1.0000
   2.000   0.1977   0.02531   0.01335  -0.0257   0.9807   1.0000
   2.250   0.2333   0.02587   0.01392  -0.0279   0.9724   1.0000
   2.500   0.2734   0.02639   0.01448  -0.0309   0.9608   1.0000
   2.750   0.3131   0.02682   0.01499  -0.0336   0.9489   1.0000
   3.000   0.3531   0.02734   0.01561  -0.0365   0.9402   1.0000
   3.250   0.3808   0.02763   0.01602  -0.0369   0.9302   1.0000
   3.500   0.4142   0.02804   0.01659  -0.0385   0.9214   1.0000
   3.750   0.4496   0.02841   0.01716  -0.0403   0.9119   1.0000
   4.000   0.4800   0.02868   0.01767  -0.0410   0.9001   1.0000
   4.250   0.5134   0.02889   0.01814  -0.0422   0.8877   1.0000
   4.500   0.5638   0.02818   0.01781  -0.0451   0.8608   1.0000
   4.750   0.6248   0.02493   0.01495  -0.0459   0.7885   1.0000
   5.000   0.6438   0.02416   0.01446  -0.0425   0.7392   1.0000
   5.250   0.6639   0.02357   0.01411  -0.0393   0.6701   1.0000
   5.500   0.7357   0.02330   0.01212  -0.0426   0.3207   1.0000
   5.750   0.7419   0.02515   0.01319  -0.0392   0.2161   1.0000
   6.000   0.7581   0.02669   0.01441  -0.0374   0.1726   1.0000
   6.250   0.7805   0.02809   0.01576  -0.0366   0.1484   1.0000
   6.500   0.8066   0.02955   0.01717  -0.0365   0.1229   1.0000
   6.750   0.8379   0.03118   0.01886  -0.0374   0.0966   1.0000
   7.000   0.8757   0.03327   0.02109  -0.0390   0.0756   1.0000
   7.250   0.9103   0.03560   0.02349  -0.0403   0.0629   1.0000
   7.500   0.9431   0.03822   0.02655  -0.0408   0.0540   1.0000
   7.750   0.9699   0.04117   0.02985  -0.0405   0.0491   1.0000
   8.000   0.9895   0.04409   0.03331  -0.0389   0.0449   1.0000
   8.250   1.0037   0.04651   0.03601  -0.0370   0.0418   1.0000
   8.500   1.0161   0.04969   0.03935  -0.0351   0.0402   1.0000
   8.750   1.0207   0.05323   0.04352  -0.0315   0.0393   1.0000
   9.000   1.0190   0.05680   0.04765  -0.0273   0.0382   1.0000
   9.250   1.0142   0.06031   0.05159  -0.0233   0.0375   1.0000
   9.500   1.0056   0.06377   0.05545  -0.0192   0.0369   1.0000
   9.750   0.9932   0.06711   0.05908  -0.0150   0.0365   1.0000
  10.000   0.9773   0.07042   0.06262  -0.0109   0.0364   1.0000
  10.250   0.9594   0.07394   0.06634  -0.0074   0.0364   1.0000
  10.500   0.9406   0.07772   0.07026  -0.0049   0.0365   1.0000
  10.750   0.9203   0.08198   0.07467  -0.0035   0.0368   1.0000
  11.000   0.8988   0.08688   0.07969  -0.0036   0.0370   1.0000
  11.250   0.8786   0.09240   0.08529  -0.0050   0.0374   1.0000
  11.500   0.8616   0.09845   0.09138  -0.0077   0.0380   1.0000
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