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

GOE 654 AIRFOIL (goe654-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 654 AIRFOIL (goe654-il)
Reynolds number: 100,000
Max Cl/Cd: 41.01 at α=6.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe654-il-100000.txt
Download as CSV file: xf-goe654-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 654 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.2297   0.10859   0.10424  -0.0362   1.0000   0.1090
  -9.000  -0.2612   0.10938   0.10523  -0.0332   1.0000   0.1103
  -8.750  -0.2939   0.11031   0.10633  -0.0317   0.9982   0.1110
  -8.500  -0.2635   0.10401   0.10002  -0.0357   0.9926   0.1137
  -8.250  -0.2258   0.09965   0.09562  -0.0385   0.9862   0.1191
  -8.000  -0.2252   0.09613   0.09213  -0.0539   0.9703   0.1264
  -7.750  -0.1946   0.09037   0.08636  -0.0558   0.9622   0.1291
  -7.500  -0.1539   0.08617   0.08212  -0.0588   0.9570   0.1346
  -7.250  -0.1649   0.08085   0.07677  -0.0778   0.9319   0.1434
  -7.000  -0.1156   0.07719   0.07313  -0.0730   0.9307   0.1471
  -6.750  -0.1079   0.07157   0.06737  -0.0895   0.9124   0.1598
  -6.500  -0.0588   0.06852   0.06436  -0.0858   0.9091   0.1654
  -6.250  -0.0466   0.06408   0.05983  -0.0941   0.8923   0.1781
  -6.000  -0.0196   0.06166   0.05738  -0.0940   0.8787   0.1864
  -5.750  -0.0083   0.05828   0.05391  -0.0972   0.8626   0.1971
  -5.500   0.0000   0.05550   0.05097  -0.1001   0.8466   0.2119
  -5.250   0.0218   0.05354   0.04901  -0.0983   0.8337   0.2187
  -5.000   0.0351   0.05113   0.04650  -0.0989   0.8213   0.2332
  -4.750   0.0494   0.04903   0.04430  -0.0988   0.8096   0.2496
  -4.500   0.0606   0.03837   0.03236  -0.1080   0.7973   0.1480
  -4.250   0.0839   0.03263   0.02544  -0.1077   0.7889   0.1060
  -4.000   0.1051   0.03038   0.02268  -0.1064   0.7775   0.1013
  -3.750   0.1295   0.02938   0.02099  -0.1048   0.7662   0.0969
  -3.500   0.1576   0.02751   0.01883  -0.1044   0.7574   0.0960
  -3.250   0.1798   0.02640   0.01752  -0.1031   0.7458   0.0956
  -3.000   0.2075   0.02533   0.01619  -0.1025   0.7376   0.0957
  -2.750   0.2324   0.02461   0.01528  -0.1015   0.7279   0.0965
  -2.500   0.2592   0.02360   0.01420  -0.1010   0.7201   0.0994
  -2.250   0.2838   0.02299   0.01361  -0.1001   0.7109   0.1031
  -2.000   0.3108   0.02241   0.01296  -0.0995   0.7035   0.1065
  -1.750   0.3347   0.02201   0.01252  -0.0983   0.6945   0.1104
  -1.500   0.3604   0.02139   0.01193  -0.0975   0.6875   0.1171
  -1.250   0.3821   0.02111   0.01169  -0.0962   0.6783   0.1286
  -1.000   0.4090   0.02034   0.01113  -0.0955   0.6718   0.1771
  -0.750   0.4174   0.01846   0.01140  -0.0915   0.6629   0.6873
  -0.500   0.5258   0.01779   0.01082  -0.1053   0.6522   1.0000
  -0.250   0.5465   0.01799   0.01081  -0.1039   0.6433   1.0000
   0.000   0.5688   0.01820   0.01079  -0.1028   0.6351   1.0000
   0.250   0.5912   0.01839   0.01078  -0.1016   0.6260   1.0000
   0.500   0.6129   0.01864   0.01087  -0.1004   0.6171   1.0000
   0.750   0.6373   0.01880   0.01082  -0.0995   0.6084   1.0000
   1.000   0.6582   0.01912   0.01103  -0.0982   0.5989   1.0000
   1.250   0.6846   0.01927   0.01096  -0.0976   0.5908   1.0000
   1.500   0.7045   0.01964   0.01127  -0.0961   0.5805   1.0000
   1.750   0.7333   0.01977   0.01114  -0.0959   0.5728   1.0000
   2.000   0.7517   0.02019   0.01154  -0.0942   0.5617   1.0000
   2.250   0.7781   0.02046   0.01163  -0.0937   0.5538   1.0000
   2.500   0.7989   0.02087   0.01199  -0.0924   0.5438   1.0000
   2.750   0.8234   0.02125   0.01224  -0.0917   0.5356   1.0000
   3.000   0.8458   0.02163   0.01256  -0.0907   0.5263   1.0000
   3.250   0.8694   0.02205   0.01289  -0.0899   0.5183   1.0000
   3.500   0.8917   0.02247   0.01325  -0.0889   0.5097   1.0000
   3.750   0.9172   0.02290   0.01357  -0.0884   0.5030   1.0000
   4.000   0.9360   0.02347   0.01418  -0.0870   0.4945   1.0000
   4.250   0.9658   0.02373   0.01425  -0.0871   0.4888   1.0000
   4.500   0.9802   0.02448   0.01514  -0.0852   0.4803   1.0000
   4.750   1.0067   0.02479   0.01535  -0.0849   0.4743   1.0000
   5.000   1.0275   0.02544   0.01601  -0.0838   0.4682   1.0000
   5.250   1.0464   0.02606   0.01669  -0.0826   0.4614   1.0000
   5.500   1.0760   0.02626   0.01676  -0.0827   0.4564   1.0000
   5.750   1.0898   0.02717   0.01781  -0.0808   0.4499   1.0000
   6.000   1.1114   0.02769   0.01834  -0.0799   0.4440   1.0000
   6.250   1.1426   0.02786   0.01839  -0.0803   0.4397   1.0000
   6.500   1.1514   0.02899   0.01972  -0.0778   0.4332   1.0000
   6.750   1.1736   0.02947   0.02020  -0.0770   0.4278   1.0000
   7.000   1.2065   0.02956   0.02018  -0.0776   0.4237   1.0000
   7.250   1.2108   0.03090   0.02175  -0.0746   0.4173   1.0000
   7.500   1.2328   0.03135   0.02221  -0.0738   0.4119   1.0000
   7.750   1.2685   0.03129   0.02203  -0.0748   0.4079   1.0000
   8.000   1.2670   0.03288   0.02389  -0.0711   0.4015   1.0000
   8.250   1.2893   0.03329   0.02431  -0.0704   0.3962   1.0000
   8.500   1.3290   0.03307   0.02393  -0.0719   0.3920   1.0000
   8.750   1.3192   0.03497   0.02615  -0.0672   0.3856   1.0000
   9.000   1.3424   0.03533   0.02654  -0.0666   0.3804   1.0000
   9.250   1.3797   0.03524   0.02634  -0.0678   0.3758   1.0000
   9.500   1.3648   0.03737   0.02877  -0.0628   0.3699   1.0000
   9.750   1.3887   0.03772   0.02913  -0.0623   0.3647   1.0000
  10.000   1.4263   0.03772   0.02906  -0.0635   0.3600   1.0000
  10.250   1.3986   0.04035   0.03200  -0.0572   0.3544   1.0000
  10.500   1.4201   0.04084   0.03253  -0.0564   0.3492   1.0000
  10.750   1.4681   0.04061   0.03221  -0.0589   0.3445   1.0000
  11.000   1.4041   0.04464   0.03658  -0.0489   0.3400   1.0000
  11.250   1.3882   0.04705   0.03912  -0.0450   0.3350   1.0000
  11.500   1.2291   0.06244   0.05476  -0.0383   0.3273   1.0000
  12.000   1.4702   0.04658   0.03868  -0.0440   0.3161   1.0000
  12.250   1.2754   0.06540   0.05788  -0.0362   0.3130   1.0000
  12.750   1.5131   0.04777   0.03993  -0.0400   0.2964   1.0000
  13.000   0.8577   0.14343   0.13600  -0.0619   0.2799   1.0000
  13.250   0.8670   0.14748   0.14012  -0.0629   0.2811   1.0000
<< Back to GOE 654 AIRFOIL (goe654-il)

Polar data table (+)

Polar graphs


<< Back to GOE 654 AIRFOIL (goe654-il)