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

GOE 428 AIRFOIL (goe428-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 428 AIRFOIL (goe428-il)
Reynolds number: 100,000
Max Cl/Cd: 58.41 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe428-il-100000.txt
Download as CSV file: xf-goe428-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 428 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
  ------ -------- --------- --------- -------- -------- --------
  -7.750  -0.3059   0.10018   0.09570  -0.0311   1.0000   0.0616
  -7.500  -0.3167   0.09945   0.09510  -0.0312   1.0000   0.0618
  -7.250  -0.3237   0.09848   0.09424  -0.0333   1.0000   0.0621
  -7.000  -0.3286   0.09719   0.09302  -0.0353   1.0000   0.0623
  -6.750  -0.3197   0.09006   0.08594  -0.0262   1.0000   0.0643
  -6.500  -0.3219   0.08790   0.08385  -0.0233   1.0000   0.0662
  -6.250  -0.3291   0.08633   0.08237  -0.0213   1.0000   0.0677
  -6.000  -0.3367   0.08486   0.08097  -0.0199   1.0000   0.0698
  -5.750  -0.3427   0.08333   0.07950  -0.0199   1.0000   0.0723
  -5.500  -0.3408   0.08227   0.07844  -0.0259   1.0000   0.0749
  -5.250  -0.3318   0.07984   0.07594  -0.0323   1.0000   0.0760
  -5.000  -0.3339   0.07572   0.07195  -0.0253   0.9983   0.0781
  -4.750  -0.2729   0.07152   0.06746  -0.0424   0.9891   0.0888
  -4.500  -0.2503   0.06591   0.06192  -0.0440   0.9837   0.0915
  -4.250  -0.2182   0.06222   0.05816  -0.0482   0.9753   0.0988
  -4.000  -0.1734   0.05740   0.05318  -0.0571   0.9681   0.1069
  -3.750  -0.1269   0.05310   0.04863  -0.0655   0.9592   0.1192
  -3.500  -0.0883   0.04940   0.04478  -0.0706   0.9503   0.1339
  -3.250  -0.0416   0.04541   0.04063  -0.0765   0.9431   0.1498
  -3.000  -0.0016   0.04202   0.03705  -0.0808   0.9324   0.1661
  -2.750   0.0427   0.03866   0.03344  -0.0857   0.9235   0.1906
  -2.500   0.0830   0.03591   0.03059  -0.0890   0.9161   0.2340
  -2.250   0.1609   0.02772   0.02019  -0.0957   0.9113   0.0888
  -2.000   0.2019   0.02520   0.01740  -0.0982   0.9035   0.0927
  -1.750   0.2453   0.02328   0.01508  -0.1004   0.8965   0.0961
  -1.500   0.2878   0.02180   0.01315  -0.1022   0.8889   0.1034
  -1.250   0.3241   0.02061   0.01190  -0.1033   0.8795   0.1129
  -1.000   0.3673   0.01932   0.01051  -0.1053   0.8728   0.1276
  -0.750   0.3988   0.01839   0.00964  -0.1054   0.8616   0.1564
  -0.500   0.4313   0.01682   0.00857  -0.1058   0.8514   0.2630
  -0.250   0.4853   0.01448   0.00780  -0.1099   0.8444   1.0000
   0.000   0.5149   0.01447   0.00750  -0.1095   0.8313   1.0000
   0.250   0.5437   0.01447   0.00727  -0.1091   0.8177   1.0000
   0.500   0.5717   0.01450   0.00709  -0.1085   0.8036   1.0000
   0.750   0.5990   0.01455   0.00697  -0.1077   0.7890   1.0000
   1.000   0.6257   0.01462   0.00688  -0.1069   0.7741   1.0000
   1.250   0.6521   0.01470   0.00680  -0.1061   0.7588   1.0000
   1.500   0.6783   0.01477   0.00673  -0.1051   0.7429   1.0000
   1.750   0.7023   0.01490   0.00676  -0.1040   0.7245   1.0000
   2.000   0.7275   0.01500   0.00674  -0.1029   0.7065   1.0000
   2.250   0.7534   0.01510   0.00673  -0.1020   0.6891   1.0000
   2.500   0.7797   0.01524   0.00673  -0.1012   0.6721   1.0000
   2.750   0.8038   0.01549   0.00690  -0.1002   0.6527   1.0000
   3.000   0.8289   0.01575   0.00705  -0.0994   0.6347   1.0000
   3.250   0.8543   0.01604   0.00726  -0.0987   0.6179   1.0000
   3.500   0.8797   0.01638   0.00751  -0.0980   0.6023   1.0000
   3.750   0.9049   0.01677   0.00783  -0.0974   0.5880   1.0000
   4.000   0.9297   0.01718   0.00823  -0.0968   0.5746   1.0000
   4.250   0.9545   0.01760   0.00868  -0.0962   0.5618   1.0000
   4.500   0.9792   0.01801   0.00910  -0.0955   0.5497   1.0000
   4.750   1.0042   0.01842   0.00952  -0.0950   0.5387   1.0000
   5.000   1.0306   0.01881   0.00987  -0.0946   0.5292   1.0000
   5.250   1.0540   0.01927   0.01044  -0.0938   0.5182   1.0000
   5.500   1.0781   0.01973   0.01103  -0.0932   0.5081   1.0000
   5.750   1.1037   0.02014   0.01147  -0.0927   0.4991   1.0000
   6.000   1.1277   0.02057   0.01199  -0.0920   0.4889   1.0000
   6.250   1.1511   0.02107   0.01265  -0.0912   0.4791   1.0000
   6.500   1.1746   0.02140   0.01309  -0.0903   0.4669   1.0000
   6.750   1.1965   0.02119   0.01273  -0.0886   0.4442   1.0000
   7.000   1.2139   0.02125   0.01290  -0.0865   0.4217   1.0000
   7.250   1.2339   0.02139   0.01305  -0.0849   0.4024   1.0000
   7.500   1.2490   0.02148   0.01329  -0.0825   0.3766   1.0000
   7.750   1.2634   0.02163   0.01356  -0.0799   0.3490   1.0000
   8.000   1.2727   0.02189   0.01385  -0.0766   0.3047   1.0000
   8.250   1.2730   0.02309   0.01451  -0.0724   0.1850   1.0000
   8.500   1.2606   0.02642   0.01685  -0.0675   0.0842   1.0000
   8.750   1.2603   0.02858   0.01895  -0.0638   0.0651   1.0000
   9.000   1.2624   0.03016   0.02064  -0.0602   0.0579   1.0000
   9.250   1.2605   0.03197   0.02249  -0.0566   0.0528   1.0000
   9.500   1.2607   0.03368   0.02434  -0.0535   0.0493   1.0000
   9.750   1.2591   0.03563   0.02642  -0.0506   0.0466   1.0000
  10.000   1.2561   0.03782   0.02876  -0.0480   0.0453   1.0000
  10.250   1.2529   0.04025   0.03126  -0.0458   0.0441   1.0000
  10.500   1.2523   0.04277   0.03379  -0.0436   0.0430   1.0000
  10.750   1.2655   0.04478   0.03579  -0.0416   0.0421   1.0000
  11.000   1.2913   0.04660   0.03770  -0.0401   0.0412   1.0000
  11.250   1.3160   0.04880   0.04008  -0.0391   0.0394   1.0000
  11.500   1.3426   0.05157   0.04307  -0.0384   0.0382   1.0000
  11.750   1.3677   0.05524   0.04709  -0.0377   0.0386   1.0000
  12.000   1.3769   0.05916   0.05143  -0.0362   0.0393   1.0000
  12.250   1.3780   0.06319   0.05581  -0.0346   0.0401   1.0000
  12.500   1.3746   0.06738   0.06032  -0.0331   0.0408   1.0000
  12.750   1.3690   0.07200   0.06521  -0.0319   0.0416   1.0000
  13.000   1.3762   0.07814   0.07158  -0.0316   0.0428   1.0000
  13.250   1.3587   0.08058   0.07432  -0.0302   0.0434   1.0000
  13.500   1.3338   0.08473   0.07880  -0.0301   0.0442   1.0000
  13.750   1.3052   0.09002   0.08439  -0.0314   0.0446   1.0000
  14.000   1.2758   0.09627   0.09093  -0.0340   0.0450   1.0000
  14.250   1.2478   0.10317   0.09807  -0.0375   0.0453   1.0000
  14.500   1.2194   0.11091   0.10602  -0.0422   0.0455   1.0000
  14.750   1.1907   0.11961   0.11486  -0.0480   0.0458   1.0000
  15.000   1.1605   0.12978   0.12518  -0.0553   0.0460   1.0000
  15.250   1.1261   0.14254   0.13805  -0.0646   0.0465   1.0000
<< Back to GOE 428 AIRFOIL (goe428-il)

Polar data table (+)

Polar graphs


<< Back to GOE 428 AIRFOIL (goe428-il)