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GOE 562 AIRFOIL (goe562-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 562 AIRFOIL (goe562-il)
Reynolds number: 50,000
Max Cl/Cd: 25.84 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe562-il-50000-n5.txt
Download as CSV file: xf-goe562-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 562 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.2245   0.13192   0.12580  -0.0418   1.0000   0.0894
 -10.250  -0.2360   0.13123   0.12529  -0.0422   1.0000   0.0901
 -10.000  -0.2502   0.13071   0.12495  -0.0420   1.0000   0.0904
  -9.750  -0.2395   0.12568   0.12001  -0.0394   1.0000   0.0920
  -9.250  -0.1999   0.11710   0.11145  -0.0472   0.9818   0.1034
  -9.000  -0.2042   0.11549   0.10992  -0.0552   0.9568   0.1055
  -8.750  -0.1691   0.10868   0.10308  -0.0540   0.9436   0.1115
  -8.500  -0.1596   0.10555   0.09996  -0.0586   0.9245   0.1183
  -8.250  -0.1685   0.10386   0.09831  -0.0679   0.9054   0.1208
  -8.000  -0.1238   0.09680   0.09116  -0.0668   0.8990   0.1269
  -7.750  -0.1175   0.09379   0.08812  -0.0737   0.8839   0.1347
  -7.500  -0.0903   0.08877   0.08304  -0.0768   0.8730   0.1416
  -7.250  -0.0934   0.08649   0.08064  -0.0864   0.8552   0.1515
  -7.000  -0.0554   0.08122   0.07533  -0.0854   0.8449   0.1608
  -6.750  -0.0484   0.07789   0.07192  -0.0903   0.8298   0.1700
  -6.500  -0.0276   0.07445   0.06836  -0.0922   0.8172   0.1785
  -6.250  -0.0160   0.07122   0.06503  -0.0946   0.8044   0.1880
  -6.000  -0.0147   0.06454   0.05764  -0.1040   0.7913   0.1056
  -5.500   0.0120   0.05752   0.05039  -0.1035   0.7688   0.0940
  -5.250   0.0294   0.05413   0.04665  -0.1041   0.7599   0.0862
  -5.000   0.0384   0.05159   0.04388  -0.1027   0.7491   0.0833
  -4.750   0.0518   0.04876   0.04066  -0.1019   0.7405   0.0804
  -4.500   0.0656   0.04653   0.03810  -0.1006   0.7315   0.0804
  -4.250   0.0801   0.04464   0.03591  -0.0991   0.7231   0.0814
  -4.000   0.0971   0.04265   0.03353  -0.0978   0.7152   0.0814
  -3.750   0.1127   0.04083   0.03129  -0.0959   0.7073   0.0802
  -3.500   0.1317   0.03913   0.02915  -0.0944   0.6999   0.0794
  -3.250   0.1520   0.03764   0.02724  -0.0930   0.6930   0.0791
  -3.000   0.1709   0.03652   0.02585  -0.0915   0.6853   0.0804
  -2.750   0.1990   0.03527   0.02422  -0.0915   0.6798   0.0832
  -2.500   0.2145   0.03457   0.02326  -0.0892   0.6716   0.0843
  -2.250   0.2447   0.03346   0.02176  -0.0893   0.6660   0.0848
  -2.000   0.2666   0.03282   0.02087  -0.0881   0.6590   0.0854
  -1.750   0.2950   0.03211   0.01988  -0.0880   0.6526   0.0865
  -1.500   0.3344   0.03130   0.01869  -0.0898   0.6478   0.0895
  -1.250   0.3580   0.03107   0.01842  -0.0894   0.6401   0.0931
  -1.000   0.3984   0.03052   0.01768  -0.0918   0.6346   0.0974
  -0.750   0.4280   0.03028   0.01727  -0.0921   0.6290   0.1007
  -0.500   0.4489   0.03032   0.01719  -0.0909   0.6224   0.1042
  -0.250   0.4805   0.03000   0.01677  -0.0916   0.6175   0.1108
   0.000   0.4998   0.03019   0.01692  -0.0902   0.6116   0.1213
   0.250   0.5201   0.03029   0.01705  -0.0891   0.6056   0.1367
   0.750   0.6859   0.02841   0.01690  -0.1121   0.5935   1.0000
   1.000   0.7044   0.02892   0.01723  -0.1106   0.5881   1.0000
   1.250   0.7322   0.02918   0.01722  -0.1105   0.5839   1.0000
   1.500   0.7370   0.03014   0.01816  -0.1068   0.5772   1.0000
   1.750   0.7550   0.03069   0.01857  -0.1052   0.5718   1.0000
   2.000   0.7837   0.03091   0.01858  -0.1053   0.5678   1.0000
   2.250   0.7833   0.03210   0.01979  -0.1009   0.5606   1.0000
   2.500   0.8028   0.03257   0.02015  -0.0996   0.5551   1.0000
   2.750   0.8309   0.03275   0.02015  -0.0995   0.5505   1.0000
   3.000   0.8265   0.03402   0.02148  -0.0946   0.5422   1.0000
   3.250   0.8576   0.03393   0.02123  -0.0948   0.5369   1.0000
   3.500   0.8552   0.03512   0.02244  -0.0902   0.5286   1.0000
   3.750   0.8799   0.03520   0.02241  -0.0895   0.5224   1.0000
   4.000   0.8873   0.03602   0.02321  -0.0863   0.5154   1.0000
   4.250   0.8978   0.03671   0.02388  -0.0837   0.5084   1.0000
   4.500   0.9368   0.03625   0.02328  -0.0850   0.5041   1.0000
   4.750   0.9125   0.03845   0.02560  -0.0777   0.4948   1.0000
   5.000   0.9431   0.03832   0.02538  -0.0778   0.4902   1.0000
   5.250   0.9220   0.04036   0.02752  -0.0712   0.4820   1.0000
   5.500   0.9369   0.04090   0.02802  -0.0693   0.4763   1.0000
   5.750   0.9797   0.04030   0.02735  -0.0710   0.4729   1.0000
   6.000   0.9291   0.04396   0.03114  -0.0617   0.4620   1.0000
   6.250   0.9650   0.04345   0.03058  -0.0622   0.4585   1.0000
   6.500   0.9199   0.04785   0.03509  -0.0555   0.4467   1.0000
   6.750   0.9520   0.04740   0.03463  -0.0553   0.4434   1.0000
   7.250   0.9443   0.05174   0.03905  -0.0505   0.4277   1.0000
   7.750   0.9400   0.05649   0.04387  -0.0470   0.4118   1.0000
   8.250   0.9355   0.06189   0.04937  -0.0444   0.3960   1.0000
   8.750   0.8750   0.07504   0.06263  -0.0431   0.3713   1.0000
   9.000   0.8885   0.07642   0.06404  -0.0425   0.3679   1.0000
   9.250   0.9091   0.07685   0.06449  -0.0417   0.3658   1.0000
   9.750   0.8870   0.08622   0.07396  -0.0413   0.3520   1.0000
  10.250   0.8611   0.09642   0.08425  -0.0417   0.3400   1.0000
  10.500   0.8724   0.09837   0.08626  -0.0414   0.3378   1.0000
  10.750   0.8854   0.10015   0.08812  -0.0411   0.3364   1.0000
  11.000   0.8485   0.10901   0.09703  -0.0426   0.3308   1.0000
  11.250   0.8450   0.11317   0.10126  -0.0431   0.3282   1.0000
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