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GOE 522 AIRFOIL (goe522-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 522 AIRFOIL (goe522-il)
Reynolds number: 50,000
Max Cl/Cd: 9.82 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe522-il-50000-n5.txt
Download as CSV file: xf-goe522-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 522 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
  ------ -------- --------- --------- -------- -------- --------
 -11.250  -0.2453   0.11356   0.10600  -0.0732   0.9611   0.0891
 -11.000  -0.2651   0.10514   0.09757  -0.0775   0.9539   0.0912
 -10.750  -0.3107   0.09211   0.08451  -0.0846   0.9489   0.0944
 -10.500  -0.3386   0.08370   0.07605  -0.0893   0.9415   0.0977
 -10.250  -0.4289   0.06825   0.06016  -0.0978   0.9317   0.1004
 -10.000  -0.4326   0.06494   0.05683  -0.0997   0.9226   0.1032
  -9.750  -0.4485   0.06071   0.05243  -0.1019   0.9127   0.1064
  -9.500  -0.4794   0.05693   0.04841  -0.1013   0.9011   0.1086
  -9.250  -0.4820   0.05497   0.04641  -0.1010   0.8922   0.1125
  -9.000  -0.4940   0.05299   0.04429  -0.0988   0.8814   0.1166
  -8.750  -0.4879   0.05127   0.04248  -0.0985   0.8729   0.1224
  -8.500  -0.4906   0.04940   0.04035  -0.0968   0.8630   0.1286
  -8.250  -0.4785   0.04843   0.03939  -0.0959   0.8540   0.1353
  -8.000  -0.4559   0.04698   0.03777  -0.0970   0.8482   0.1441
  -7.750  -0.4591   0.04611   0.03677  -0.0936   0.8356   0.1500
  -7.500  -0.4318   0.04516   0.03579  -0.0946   0.8297   0.1600
  -7.250  -0.4302   0.04429   0.03471  -0.0918   0.8176   0.1666
  -7.000  -0.4021   0.04366   0.03414  -0.0922   0.8110   0.1741
  -6.750  -0.3906   0.04285   0.03313  -0.0906   0.8010   0.1810
  -6.500  -0.3673   0.04246   0.03282  -0.0900   0.7926   0.1879
  -6.250  -0.3339   0.04148   0.03168  -0.0913   0.7878   0.1972
  -6.000  -0.3288   0.04147   0.03171  -0.0881   0.7744   0.2022
  -5.750  -0.2974   0.04058   0.03067  -0.0890   0.7689   0.2117
  -5.500  -0.2866   0.04065   0.03078  -0.0866   0.7571   0.2178
  -5.250  -0.2590   0.03997   0.02997  -0.0867   0.7502   0.2280
  -5.000  -0.2225   0.03937   0.02933  -0.0878   0.7462   0.2393
  -4.750  -0.2192   0.03955   0.02947  -0.0846   0.7321   0.2461
  -4.500  -0.1861   0.03893   0.02876  -0.0852   0.7273   0.2580
  -4.250  -0.1755   0.03906   0.02889  -0.0828   0.7158   0.2656
  -4.000  -0.1488   0.03855   0.02824  -0.0826   0.7091   0.2776
  -3.750  -0.1132   0.03785   0.02754  -0.0833   0.7052   0.2892
  -3.500  -0.1088   0.03819   0.02772  -0.0805   0.6925   0.2994
  -3.250  -0.0793   0.03780   0.02738  -0.0803   0.6871   0.3118
  -3.000  -0.0434   0.03709   0.02658  -0.0811   0.6836   0.3284
  -2.750  -0.0431   0.03790   0.02738  -0.0776   0.6706   0.3382
  -2.500  -0.0129   0.03749   0.02694  -0.0776   0.6658   0.3539
  -2.250   0.0233   0.03688   0.02620  -0.0783   0.6626   0.3720
  -2.000   0.0197   0.03807   0.02745  -0.0745   0.6500   0.3803
  -1.750   0.0496   0.03778   0.02706  -0.0745   0.6454   0.3963
  -1.500   0.0858   0.03726   0.02643  -0.0752   0.6423   0.4131
  -1.250   0.0814   0.03863   0.02785  -0.0715   0.6303   0.4206
  -1.000   0.1086   0.03851   0.02765  -0.0712   0.6256   0.4352
  -0.750   0.1423   0.03810   0.02721  -0.0714   0.6225   0.4491
  -0.500   0.1383   0.03962   0.02870  -0.0681   0.6120   0.4580
  -0.250   0.1603   0.03980   0.02887  -0.0672   0.6066   0.4706
   0.000   0.1926   0.03950   0.02847  -0.0673   0.6033   0.4890
   0.250   0.2289   0.03902   0.02796  -0.0677   0.6009   0.5078
   0.500   0.2080   0.04166   0.03066  -0.0632   0.5881   0.5144
   0.750   0.2362   0.04158   0.03057  -0.0628   0.5844   0.5302
   1.000   0.2717   0.04110   0.03004  -0.0633   0.5819   0.5455
   1.500   0.2785   0.04408   0.03302  -0.0592   0.5657   0.5665
   1.750   0.3105   0.04383   0.03273  -0.0594   0.5628   0.5828
   2.000   0.3466   0.04330   0.03218  -0.0598   0.5607   0.6006
   2.500   0.3438   0.04719   0.03619  -0.0554   0.5434   0.6281
   2.750   0.3774   0.04672   0.03574  -0.0555   0.5409   0.6511
   3.000   0.4147   0.04599   0.03504  -0.0557   0.5391   0.6794
   3.500   0.4057   0.05042   0.03977  -0.0511   0.5208   0.7269
   3.750   0.4184   0.05143   0.04096  -0.0498   0.5147   0.7681
   4.250   0.4691   0.05315   0.04294  -0.0523   0.4998   1.0000
   4.500   0.5118   0.05211   0.04165  -0.0532   0.4975   1.0000
   5.000   0.5156   0.05636   0.04572  -0.0509   0.4782   1.0000
   5.250   0.5327   0.05729   0.04652  -0.0503   0.4720   1.0000
   5.750   0.5559   0.06004   0.04909  -0.0488   0.4565   1.0000
   6.000   0.5822   0.06012   0.04903  -0.0485   0.4525   1.0000
   7.000   0.5955   0.06912   0.05790  -0.0452   0.4168   1.0000
   7.250   0.6277   0.06854   0.05721  -0.0449   0.4143   1.0000
   7.750   0.6287   0.07395   0.06259  -0.0436   0.3962   1.0000
   8.250   0.6319   0.07932   0.06793  -0.0427   0.3789   1.0000
   8.500   0.6611   0.07899   0.06752  -0.0423   0.3766   1.0000
   9.000   0.6567   0.08571   0.07423  -0.0419   0.3600   1.0000
   9.500   0.6485   0.09333   0.08187  -0.0421   0.3448   1.0000
   9.750   0.6718   0.09379   0.08228  -0.0418   0.3424   1.0000
  10.250   0.6573   0.10257   0.09111  -0.0427   0.3282   1.0000
  10.500   0.6765   0.10369   0.09219  -0.0425   0.3260   1.0000
  11.000   0.6546   0.11416   0.10273  -0.0444   0.3152   1.0000
  11.250   0.6658   0.11638   0.10495  -0.0447   0.3122   1.0000
  11.500   0.6827   0.11780   0.10635  -0.0447   0.3100   1.0000
  11.750   0.7027   0.11887   0.10739  -0.0446   0.3084   1.0000
  12.000   0.6732   0.12670   0.11531  -0.0468   0.3027   1.0000
  12.250   0.6732   0.13057   0.11921  -0.0477   0.3001   1.0000
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