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GOE 392 AIRFOIL (goe392-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 392 AIRFOIL (goe392-il)
Reynolds number: 1,000,000
Max Cl/Cd: 125.76 at α=2.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe392-il-1000000.txt
Download as CSV file: xf-goe392-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 392 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2509   0.09301   0.09148  -0.0546   0.9933   0.0067
  -8.500  -0.2385   0.08837   0.08685  -0.0588   0.9906   0.0068
  -8.250  -0.2239   0.08418   0.08266  -0.0633   0.9879   0.0068
  -8.000  -0.2084   0.07918   0.07767  -0.0691   0.9852   0.0068
  -7.750  -0.1923   0.07442   0.07291  -0.0749   0.9821   0.0068
  -7.500  -0.1929   0.06750   0.06603  -0.0813   0.9729   0.0072
  -7.250  -0.1765   0.06303   0.06155  -0.0877   0.9656   0.0074
  -7.000  -0.1535   0.05844   0.05693  -0.0955   0.9582   0.0076
  -6.750  -0.1261   0.05380   0.05223  -0.1040   0.9472   0.0082
  -6.500  -0.0858   0.04689   0.04520  -0.1172   0.9356   0.0088
  -6.250  -0.0475   0.03932   0.03738  -0.1287   0.9150   0.0094
  -6.000  -0.0222   0.03438   0.03210  -0.1329   0.8839   0.0101
  -5.750   0.0002   0.03187   0.02927  -0.1331   0.8573   0.0110
  -5.500   0.0195   0.02988   0.02699  -0.1326   0.8382   0.0113
  -4.500   0.0817   0.01404   0.00933  -0.1274   0.7915   0.0081
  -4.250   0.1053   0.01274   0.00777  -0.1265   0.7827   0.0086
  -4.000   0.1305   0.01205   0.00694  -0.1260   0.7741   0.0098
  -3.750   0.1554   0.01121   0.00592  -0.1253   0.7659   0.0105
  -3.500   0.1808   0.01072   0.00531  -0.1248   0.7571   0.0111
  -3.250   0.2056   0.01004   0.00450  -0.1242   0.7476   0.0118
  -3.000   0.2271   0.00873   0.00303  -0.1229   0.7374   0.0131
  -2.750   0.2526   0.00857   0.00281  -0.1224   0.7264   0.0153
  -2.500   0.2780   0.00830   0.00246  -0.1220   0.7161   0.0170
  -2.250   0.3047   0.00827   0.00238  -0.1217   0.7069   0.0188
  -2.000   0.3292   0.00779   0.00177  -0.1211   0.6979   0.0238
  -1.750   0.3552   0.00767   0.00158  -0.1208   0.6882   0.0283
  -1.500   0.3817   0.00762   0.00149  -0.1205   0.6787   0.0304
  -1.250   0.4075   0.00748   0.00123  -0.1201   0.6691   0.0372
  -0.750   0.4549   0.00660   0.00108  -0.1190   0.6482   0.3379
  -0.500   0.4808   0.00657   0.00110  -0.1188   0.6378   0.3792
   0.000   0.5327   0.00660   0.00110  -0.1182   0.6169   0.4186
   0.250   0.5584   0.00660   0.00111  -0.1179   0.6059   0.4435
   0.500   0.5836   0.00652   0.00113  -0.1175   0.5953   0.4904
   0.750   0.6056   0.00619   0.00123  -0.1165   0.5857   0.6664
   1.000   0.6217   0.00571   0.00134  -0.1139   0.5768   0.8880
   1.250   0.6886   0.00572   0.00142  -0.1230   0.5656   1.0000
   1.500   0.7138   0.00581   0.00147  -0.1226   0.5571   1.0000
   1.750   0.7386   0.00593   0.00153  -0.1221   0.5486   1.0000
   2.000   0.7629   0.00607   0.00160  -0.1214   0.5329   1.0000
   2.250   0.7860   0.00626   0.00168  -0.1206   0.5103   1.0000
   2.500   0.8099   0.00644   0.00177  -0.1199   0.4905   1.0000
   2.750   0.8331   0.00665   0.00188  -0.1191   0.4671   1.0000
   3.000   0.8528   0.00708   0.00204  -0.1176   0.4133   1.0000
   3.250   0.8608   0.00832   0.00257  -0.1141   0.2793   1.0000
   3.500   0.8558   0.01058   0.00373  -0.1082   0.0207   1.0000
   3.750   0.8796   0.01082   0.00402  -0.1074   0.0160   1.0000
   4.000   0.9027   0.01109   0.00436  -0.1066   0.0145   1.0000
   4.250   0.9253   0.01139   0.00469  -0.1057   0.0129   1.0000
   4.500   0.9471   0.01175   0.00509  -0.1046   0.0118   1.0000
   4.750   0.9680   0.01216   0.00557  -0.1033   0.0106   1.0000
   5.000   0.9863   0.01275   0.00622  -0.1015   0.0096   1.0000
   5.250   1.0011   0.01353   0.00709  -0.0990   0.0091   1.0000
   5.500   1.0088   0.01464   0.00831  -0.0953   0.0086   1.0000
   5.750   1.0059   0.01616   0.00992  -0.0895   0.0082   1.0000
   6.000   1.0299   0.01619   0.00998  -0.0889   0.0077   1.0000
   6.250   1.0422   0.01681   0.01065  -0.0861   0.0073   1.0000
   6.500   1.0535   0.01754   0.01144  -0.0831   0.0068   1.0000
   6.750   1.0626   0.01850   0.01245  -0.0798   0.0064   1.0000
   7.000   1.0720   0.01983   0.01383  -0.0766   0.0062   1.0000
   7.250   1.1603   0.02795   0.02181  -0.0880   0.0081   1.0000
   7.500   1.1862   0.03003   0.02401  -0.0876   0.0081   1.0000
   7.750   1.2083   0.03195   0.02606  -0.0866   0.0080   1.0000
   8.000   1.2265   0.03356   0.02783  -0.0849   0.0080   1.0000
   8.250   1.2376   0.03280   0.02723  -0.0818   0.0075   1.0000
   8.500   1.2529   0.03265   0.02719  -0.0795   0.0065   1.0000
   8.750   1.2704   0.03404   0.02870  -0.0779   0.0058   1.0000
   9.000   1.2863   0.03581   0.03060  -0.0763   0.0055   1.0000
   9.250   1.2991   0.03763   0.03255  -0.0744   0.0052   1.0000
   9.500   1.3108   0.03970   0.03474  -0.0725   0.0050   1.0000
   9.750   1.3226   0.04201   0.03714  -0.0709   0.0048   1.0000
  10.500   1.3008   0.05675   0.05272  -0.0586   0.0043   1.0000
  10.750   1.2860   0.05896   0.05511  -0.0531   0.0043   1.0000
  11.000   1.2728   0.06124   0.05754  -0.0486   0.0043   1.0000
  11.250   1.2563   0.06380   0.06027  -0.0444   0.0043   1.0000
  11.500   1.2407   0.06655   0.06317  -0.0410   0.0043   1.0000
  11.750   1.2222   0.06967   0.06646  -0.0382   0.0043   1.0000
  12.000   1.2052   0.07301   0.06993  -0.0363   0.0043   1.0000
  12.250   1.1871   0.07671   0.07378  -0.0350   0.0043   1.0000
  12.500   1.1701   0.08068   0.07787  -0.0345   0.0043   1.0000
  12.750   1.1510   0.08512   0.08245  -0.0346   0.0043   1.0000
  13.000   1.1315   0.08994   0.08741  -0.0353   0.0043   1.0000
  13.250   1.1144   0.09494   0.09253  -0.0367   0.0043   1.0000
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