Airfoil Tools
Search 1636 airfoils Google+
You have 0 airfoils loaded.
Your Reynold number range is 50,000 to 1,000,000. (set)

GOE 796 AIRFOIL (goe796-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 796 AIRFOIL (goe796-il)
Reynolds number: 200,000
Max Cl/Cd: 73.25 at α=4.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe796-il-200000-n5.txt
Download as CSV file: xf-goe796-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 796 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.500  -0.4050   0.11368   0.10971  -0.0392   1.0000   0.0269
 -11.250  -0.4048   0.10950   0.10554  -0.0408   1.0000   0.0267
 -10.250  -0.7010   0.04280   0.03846  -0.0739   0.9946   0.0240
 -10.000  -0.6953   0.03719   0.03230  -0.0775   0.9877   0.0241
  -9.750  -0.6762   0.03374   0.02843  -0.0798   0.9834   0.0242
  -9.500  -0.6561   0.03124   0.02560  -0.0809   0.9781   0.0244
  -9.250  -0.6313   0.02912   0.02319  -0.0823   0.9740   0.0246
  -9.000  -0.6027   0.02727   0.02105  -0.0841   0.9712   0.0249
  -8.750  -0.5785   0.02584   0.01940  -0.0844   0.9655   0.0252
  -8.500  -0.5504   0.02455   0.01791  -0.0854   0.9611   0.0257
  -8.250  -0.5199   0.02334   0.01649  -0.0867   0.9580   0.0263
  -8.000  -0.4951   0.02233   0.01530  -0.0867   0.9514   0.0268
  -7.750  -0.4666   0.02133   0.01413  -0.0873   0.9465   0.0272
  -7.500  -0.4358   0.02039   0.01303  -0.0883   0.9429   0.0276
  -7.250  -0.4122   0.01963   0.01217  -0.0877   0.9350   0.0281
  -7.000  -0.3832   0.01888   0.01133  -0.0882   0.9298   0.0288
  -6.750  -0.3562   0.01827   0.01063  -0.0882   0.9234   0.0296
  -6.500  -0.3292   0.01769   0.00997  -0.0882   0.9164   0.0307
  -6.250  -0.3001   0.01711   0.00929  -0.0885   0.9112   0.0323
  -6.000  -0.2748   0.01661   0.00872  -0.0880   0.9025   0.0343
  -5.750  -0.2459   0.01612   0.00818  -0.0882   0.8968   0.0383
  -5.500  -0.2202   0.01570   0.00772  -0.0877   0.8878   0.0435
  -5.250  -0.1915   0.01532   0.00723  -0.0878   0.8816   0.0500
  -5.000  -0.1655   0.01497   0.00687  -0.0874   0.8724   0.0559
  -4.750  -0.1374   0.01465   0.00648  -0.0874   0.8656   0.0617
  -4.500  -0.1106   0.01440   0.00623  -0.0872   0.8566   0.0681
  -4.250  -0.0828   0.01416   0.00595  -0.0871   0.8493   0.0736
  -4.000  -0.0554   0.01398   0.00574  -0.0869   0.8405   0.0806
  -3.750  -0.0276   0.01382   0.00556  -0.0868   0.8327   0.0892
  -3.500   0.0001   0.01367   0.00535  -0.0867   0.8247   0.0974
  -3.250   0.0276   0.01348   0.00513  -0.0866   0.8170   0.1047
  -3.000   0.0553   0.01333   0.00491  -0.0864   0.8091   0.1122
  -2.750   0.0826   0.01313   0.00471  -0.0863   0.8018   0.1212
  -2.500   0.1102   0.01298   0.00452  -0.0862   0.7943   0.1295
  -2.250   0.1375   0.01281   0.00433  -0.0860   0.7866   0.1379
  -2.000   0.1649   0.01266   0.00415  -0.0859   0.7782   0.1462
  -1.750   0.1922   0.01252   0.00400  -0.0857   0.7695   0.1562
  -1.500   0.2195   0.01238   0.00385  -0.0855   0.7610   0.1687
  -1.250   0.2466   0.01225   0.00376  -0.0853   0.7523   0.1865
  -1.000   0.2738   0.01208   0.00365  -0.0852   0.7447   0.2159
  -0.750   0.3004   0.01188   0.00360  -0.0850   0.7361   0.2615
  -0.500   0.3267   0.01162   0.00353  -0.0847   0.7284   0.3315
  -0.250   0.3511   0.01119   0.00352  -0.0842   0.7188   0.4510
   0.000   0.3736   0.01067   0.00352  -0.0830   0.7108   0.6151
   0.250   0.3955   0.01023   0.00362  -0.0812   0.7031   0.7843
   0.500   0.4440   0.01012   0.00370  -0.0847   0.6972   0.9364
   0.750   0.4842   0.01019   0.00373  -0.0873   0.6899   0.9764
   1.000   0.5264   0.01027   0.00372  -0.0904   0.6828   0.9964
   1.250   0.5560   0.01034   0.00374  -0.0908   0.6750   1.0000
   1.500   0.5812   0.01043   0.00376  -0.0902   0.6675   1.0000
   1.750   0.6064   0.01052   0.00381  -0.0896   0.6598   1.0000
   2.000   0.6315   0.01062   0.00384  -0.0890   0.6517   1.0000
   2.250   0.6564   0.01071   0.00390  -0.0884   0.6421   1.0000
   2.500   0.6812   0.01081   0.00395  -0.0877   0.6318   1.0000
   2.750   0.7059   0.01092   0.00400  -0.0870   0.6201   1.0000
   3.000   0.7305   0.01103   0.00409  -0.0863   0.6079   1.0000
   3.250   0.7552   0.01115   0.00418  -0.0856   0.5954   1.0000
   3.500   0.7799   0.01128   0.00427  -0.0849   0.5822   1.0000
   3.750   0.8042   0.01143   0.00437  -0.0842   0.5673   1.0000
   4.000   0.8284   0.01159   0.00449  -0.0834   0.5507   1.0000
   4.250   0.8522   0.01178   0.00462  -0.0826   0.5323   1.0000
   4.500   0.8753   0.01200   0.00476  -0.0816   0.5115   1.0000
   4.750   0.8980   0.01226   0.00494  -0.0807   0.4883   1.0000
   5.250   0.9422   0.01288   0.00538  -0.0785   0.4447   1.0000
   5.500   0.9639   0.01324   0.00564  -0.0774   0.4249   1.0000
   5.750   0.9848   0.01364   0.00595  -0.0762   0.4053   1.0000
   6.000   1.0062   0.01403   0.00628  -0.0751   0.3893   1.0000
   6.250   1.0278   0.01440   0.00661  -0.0741   0.3753   1.0000
   6.500   1.0487   0.01482   0.00698  -0.0730   0.3601   1.0000
   6.750   1.0686   0.01528   0.00738  -0.0717   0.3432   1.0000
   7.000   1.0883   0.01574   0.00778  -0.0704   0.3270   1.0000
   7.250   1.1090   0.01615   0.00819  -0.0692   0.3134   1.0000
   7.500   1.1294   0.01655   0.00860  -0.0681   0.3004   1.0000
   7.750   1.1494   0.01697   0.00905  -0.0669   0.2871   1.0000
   8.000   1.1688   0.01741   0.00950  -0.0656   0.2726   1.0000
   8.250   1.1870   0.01788   0.00997  -0.0641   0.2546   1.0000
   8.500   1.2027   0.01845   0.01048  -0.0622   0.2317   1.0000
   8.750   1.2142   0.01915   0.01108  -0.0597   0.2067   1.0000
   9.000   1.2245   0.01995   0.01177  -0.0571   0.1872   1.0000
   9.250   1.2352   0.02079   0.01255  -0.0547   0.1732   1.0000
   9.500   1.2462   0.02167   0.01339  -0.0525   0.1619   1.0000
   9.750   1.2578   0.02255   0.01426  -0.0504   0.1514   1.0000
  10.000   1.2711   0.02336   0.01512  -0.0487   0.1411   1.0000
  10.250   1.2832   0.02425   0.01603  -0.0469   0.1308   1.0000
  10.500   1.2950   0.02520   0.01697  -0.0451   0.1198   1.0000
  10.750   1.3070   0.02616   0.01794  -0.0435   0.1083   1.0000
  11.000   1.3178   0.02723   0.01901  -0.0418   0.0974   1.0000
  11.250   1.3275   0.02840   0.02019  -0.0402   0.0885   1.0000
  11.500   1.3360   0.02970   0.02150  -0.0385   0.0812   1.0000
  11.750   1.3455   0.03097   0.02282  -0.0370   0.0750   1.0000
  12.000   1.3529   0.03244   0.02432  -0.0355   0.0696   1.0000
  12.250   1.3617   0.03384   0.02581  -0.0342   0.0641   1.0000
  12.500   1.3675   0.03553   0.02754  -0.0328   0.0588   1.0000
  12.750   1.3749   0.03714   0.02923  -0.0316   0.0531   1.0000
  13.000   1.3796   0.03903   0.03119  -0.0304   0.0480   1.0000
  13.250   1.3843   0.04097   0.03319  -0.0293   0.0425   1.0000
  13.500   1.3874   0.04311   0.03540  -0.0283   0.0379   1.0000
  13.750   1.3886   0.04551   0.03784  -0.0274   0.0339   1.0000
  14.000   1.3897   0.04797   0.04039  -0.0266   0.0308   1.0000
  14.250   1.3885   0.05074   0.04322  -0.0259   0.0282   1.0000
  14.500   1.3868   0.05368   0.04626  -0.0254   0.0262   1.0000
  14.750   1.3850   0.05672   0.04940  -0.0251   0.0245   1.0000
  15.000   1.3810   0.06011   0.05288  -0.0249   0.0231   1.0000
  15.250   1.3746   0.06392   0.05680  -0.0250   0.0220   1.0000
  15.500   1.3710   0.06751   0.06054  -0.0253   0.0211   1.0000
  15.750   1.3657   0.07141   0.06457  -0.0257   0.0202   1.0000
  16.000   1.3593   0.07559   0.06888  -0.0264   0.0194   1.0000
  16.250   1.3515   0.08011   0.07353  -0.0274   0.0188   1.0000
  16.500   1.3419   0.08502   0.07857  -0.0286   0.0183   1.0000
  16.750   1.3305   0.09035   0.08401  -0.0302   0.0179   1.0000
<< Back to GOE 796 AIRFOIL (goe796-il)

Polar data table (+)

Polar graphs


<< Back to GOE 796 AIRFOIL (goe796-il)