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GOE 415 AIRFOIL (goe415-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 415 AIRFOIL (goe415-il)
Reynolds number: 50,000
Max Cl/Cd: 33.98 at α=9.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe415-il-50000.txt
Download as CSV file: xf-goe415-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 415 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.4203   0.10228   0.09612   0.0001   1.0000   0.2055
  -7.500  -0.4270   0.10091   0.09487  -0.0020   1.0000   0.2159
  -7.250  -0.4209   0.09763   0.09162  -0.0019   1.0000   0.2297
  -7.000  -0.4165   0.09464   0.08872  -0.0027   1.0000   0.2440
  -6.750  -0.4139   0.09196   0.08611  -0.0044   1.0000   0.2586
  -6.500  -0.3943   0.08694   0.08112   0.0000   1.0000   0.2772
  -6.250  -0.3863   0.08381   0.07807   0.0006   1.0000   0.2956
  -6.000  -0.3902   0.08210   0.07647  -0.0025   1.0000   0.3155
  -5.750  -0.3729   0.07766   0.07208   0.0017   1.0000   0.3361
  -5.500  -0.3701   0.07503   0.06955   0.0018   1.0000   0.3605
  -5.000  -0.3553   0.06911   0.06380   0.0066   1.0000   0.4180
  -4.750  -0.3448   0.06596   0.06072   0.0105   1.0000   0.4507
  -4.500  -0.3402   0.06326   0.05810   0.0144   1.0000   0.4901
  -3.500  -0.1912   0.03867   0.03113  -0.0435   1.0000   0.1600
  -3.250  -0.1641   0.03520   0.02721  -0.0439   1.0000   0.1448
  -3.000  -0.1379   0.03216   0.02377  -0.0440   1.0000   0.1350
  -2.750  -0.1107   0.02984   0.02097  -0.0437   1.0000   0.1267
  -2.500  -0.0833   0.02778   0.01839  -0.0433   1.0000   0.1207
  -2.250  -0.0568   0.02648   0.01657  -0.0426   1.0000   0.1179
  -2.000  -0.0325   0.02492   0.01498  -0.0423   1.0000   0.1218
  -1.750  -0.0091   0.02400   0.01389  -0.0418   1.0000   0.1281
  -1.500   0.0145   0.02334   0.01303  -0.0413   1.0000   0.1319
  -1.250   0.0390   0.02274   0.01242  -0.0416   1.0000   0.1399
  -1.000   0.0609   0.02255   0.01223  -0.0421   1.0000   0.1581
  -0.750   0.1140   0.02041   0.01182  -0.0480   0.9904   0.4497
  -0.500   0.1840   0.01951   0.01122  -0.0546   0.9603   1.0000
  -0.250   0.2561   0.01988   0.01102  -0.0631   0.9310   1.0000
   0.000   0.3206   0.02002   0.01077  -0.0696   0.9029   1.0000
   0.250   0.3697   0.02011   0.01060  -0.0728   0.8744   1.0000
   0.500   0.4079   0.02021   0.01049  -0.0737   0.8461   1.0000
   0.750   0.4397   0.02033   0.01041  -0.0733   0.8188   1.0000
   1.000   0.4686   0.02048   0.01038  -0.0724   0.7928   1.0000
   1.250   0.4961   0.02064   0.01037  -0.0712   0.7686   1.0000
   1.500   0.5232   0.02082   0.01037  -0.0698   0.7462   1.0000
   1.750   0.5480   0.02124   0.01067  -0.0685   0.7226   1.0000
   2.000   0.5737   0.02161   0.01087  -0.0672   0.7021   1.0000
   2.250   0.5979   0.02223   0.01140  -0.0662   0.6803   1.0000
   2.500   0.6231   0.02277   0.01181  -0.0650   0.6617   1.0000
   2.750   0.6475   0.02349   0.01250  -0.0642   0.6429   1.0000
   3.000   0.6719   0.02425   0.01323  -0.0635   0.6252   1.0000
   3.250   0.6965   0.02502   0.01396  -0.0628   0.6090   1.0000
   3.500   0.7211   0.02585   0.01477  -0.0621   0.5939   1.0000
   3.750   0.7455   0.02677   0.01576  -0.0616   0.5798   1.0000
   4.000   0.7694   0.02776   0.01679  -0.0611   0.5661   1.0000
   4.250   0.7930   0.02884   0.01794  -0.0606   0.5530   1.0000
   4.500   0.8164   0.03001   0.01920  -0.0602   0.5409   1.0000
   4.750   0.8399   0.03115   0.02041  -0.0597   0.5291   1.0000
   5.000   0.8641   0.03215   0.02152  -0.0589   0.5177   1.0000
   5.250   0.8869   0.03327   0.02275  -0.0582   0.5050   1.0000
   5.500   0.9092   0.03425   0.02383  -0.0573   0.4907   1.0000
   5.750   0.9315   0.03519   0.02488  -0.0562   0.4763   1.0000
   6.000   0.9532   0.03624   0.02612  -0.0553   0.4622   1.0000
   6.250   0.9739   0.03750   0.02757  -0.0544   0.4486   1.0000
   6.500   0.9937   0.03895   0.02924  -0.0536   0.4352   1.0000
   6.750   1.0129   0.04045   0.03095  -0.0528   0.4217   1.0000
   7.000   1.0311   0.04213   0.03287  -0.0519   0.4083   1.0000
   7.250   1.0463   0.04442   0.03543  -0.0513   0.3959   1.0000
   7.500   1.0597   0.04698   0.03832  -0.0507   0.3840   1.0000
   7.750   1.0738   0.04939   0.04098  -0.0499   0.3727   1.0000
   8.000   1.0914   0.05081   0.04261  -0.0486   0.3595   1.0000
   8.250   1.1359   0.04490   0.03658  -0.0443   0.3296   1.0000
   8.500   1.1750   0.04092   0.03246  -0.0417   0.3073   1.0000
   8.750   1.1919   0.04073   0.03262  -0.0398   0.2882   1.0000
   9.000   1.2141   0.03800   0.02997  -0.0372   0.2605   1.0000
   9.250   1.2262   0.03609   0.02817  -0.0342   0.2250   1.0000
   9.500   1.2274   0.03663   0.02889  -0.0315   0.1731   1.0000
   9.750   1.2202   0.03929   0.03107  -0.0290   0.1265   1.0000
  10.000   1.2107   0.04268   0.03416  -0.0267   0.1070   1.0000
  10.250   1.2022   0.04588   0.03733  -0.0247   0.0956   1.0000
  10.500   1.1957   0.04922   0.04071  -0.0233   0.0881   1.0000
  10.750   1.1905   0.05262   0.04414  -0.0226   0.0827   1.0000
  11.000   1.1884   0.05600   0.04759  -0.0218   0.0787   1.0000
  11.250   1.1894   0.05930   0.05115  -0.0210   0.0751   1.0000
  11.500   1.1922   0.06244   0.05437  -0.0201   0.0720   1.0000
  11.750   1.2037   0.06493   0.05680  -0.0179   0.0691   1.0000
  12.000   1.2055   0.06891   0.06115  -0.0174   0.0680   1.0000
  12.250   1.2004   0.07365   0.06620  -0.0179   0.0673   1.0000
  12.500   1.1898   0.07911   0.07195  -0.0195   0.0670   1.0000
  12.750   1.1748   0.08534   0.07843  -0.0220   0.0671   1.0000
  13.000   1.1566   0.09234   0.08566  -0.0254   0.0676   1.0000
  13.250   1.1369   0.09995   0.09344  -0.0294   0.0682   1.0000
  13.500   1.1170   0.10798   0.10160  -0.0337   0.0689   1.0000
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