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

GOE 400 AIRFOIL (goe400-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 400 AIRFOIL (goe400-il)
Reynolds number: 100,000
Max Cl/Cd: 59.08 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe400-il-100000.txt
Download as CSV file: xf-goe400-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 400 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4100   0.11661   0.11167   0.0022   1.0000   0.0654
  -8.750  -0.4105   0.11610   0.11123  -0.0021   1.0000   0.0664
  -8.500  -0.4131   0.11584   0.11105  -0.0071   1.0000   0.0668
  -8.250  -0.3950   0.10776   0.10295  -0.0040   1.0000   0.0679
  -8.000  -0.3827   0.10342   0.09863  -0.0032   1.0000   0.0700
  -7.750  -0.3742   0.10042   0.09564  -0.0041   1.0000   0.0728
  -7.500  -0.3682   0.09805   0.09332  -0.0064   1.0000   0.0761
  -7.250  -0.3610   0.09808   0.09341  -0.0171   1.0000   0.0781
  -7.000  -0.3493   0.09406   0.08943  -0.0213   1.0000   0.0789
  -6.750  -0.3409   0.08863   0.08404  -0.0159   1.0000   0.0803
  -6.500  -0.3293   0.08516   0.08059  -0.0159   1.0000   0.0825
  -6.250  -0.3162   0.08210   0.07756  -0.0182   1.0000   0.0855
  -6.000  -0.2894   0.08076   0.07618  -0.0308   1.0000   0.0903
  -5.750  -0.2756   0.07650   0.07195  -0.0336   1.0000   0.0916
  -5.500  -0.2700   0.07256   0.06809  -0.0296   1.0000   0.0934
  -5.250  -0.2586   0.06966   0.06523  -0.0296   1.0000   0.0960
  -5.000  -0.2455   0.06710   0.06270  -0.0314   1.0000   0.0995
  -4.750  -0.2130   0.06623   0.06160  -0.0433   1.0000   0.1041
  -4.500  -0.2198   0.06327   0.05880  -0.0391   1.0000   0.1047
  -4.250  -0.2275   0.06142   0.05705  -0.0355   1.0000   0.1054
  -4.000  -0.1941   0.05743   0.05305  -0.0392   0.9927   0.1101
  -3.750  -0.1290   0.05296   0.04827  -0.0527   0.9834   0.1188
  -3.500  -0.0897   0.04913   0.04441  -0.0572   0.9734   0.1233
  -3.250  -0.0299   0.04540   0.04035  -0.0676   0.9636   0.1328
  -3.000   0.0260   0.04234   0.03702  -0.0756   0.9560   0.1464
  -2.750   0.0606   0.03906   0.03379  -0.0782   0.9445   0.1530
  -2.500   0.1051   0.03636   0.03088  -0.0832   0.9327   0.1652
  -2.250   0.1459   0.03403   0.02839  -0.0868   0.9208   0.1799
  -2.000   0.1823   0.03204   0.02625  -0.0891   0.9082   0.1976
  -1.750   0.2138   0.03029   0.02438  -0.0903   0.8950   0.2243
  -1.500   0.2691   0.02605   0.01895  -0.0941   0.8821   0.1292
  -1.250   0.3025   0.02352   0.01581  -0.0942   0.8686   0.1159
  -1.000   0.3317   0.02201   0.01390  -0.0938   0.8549   0.1163
  -0.750   0.3584   0.02083   0.01257  -0.0932   0.8414   0.1211
  -0.500   0.3856   0.02005   0.01150  -0.0924   0.8282   0.1326
  -0.250   0.4125   0.01927   0.01049  -0.0915   0.8155   0.1456
   0.000   0.4381   0.01846   0.00967  -0.0906   0.8033   0.1667
   0.250   0.4637   0.01808   0.00921  -0.0897   0.7904   0.1930
   0.500   0.4892   0.01792   0.00903  -0.0889   0.7772   0.2171
   0.750   0.5151   0.01776   0.00882  -0.0882   0.7644   0.2338
   1.000   0.5410   0.01751   0.00859  -0.0874   0.7520   0.2418
   1.250   0.5666   0.01738   0.00842  -0.0865   0.7403   0.2509
   1.500   0.5929   0.01727   0.00826  -0.0855   0.7290   0.2587
   1.750   0.6191   0.01724   0.00824  -0.0848   0.7160   0.2666
   2.000   0.6451   0.01726   0.00829  -0.0841   0.7034   0.2781
   2.250   0.6711   0.01725   0.00836  -0.0834   0.6912   0.2968
   2.500   0.6965   0.01687   0.00846  -0.0828   0.6797   0.4220
   2.750   0.7255   0.01612   0.00835  -0.0820   0.6684   1.0000
   3.000   0.7515   0.01642   0.00855  -0.0813   0.6552   1.0000
   3.250   0.7773   0.01673   0.00878  -0.0806   0.6422   1.0000
   3.500   0.8031   0.01703   0.00901  -0.0799   0.6294   1.0000
   3.750   0.8287   0.01728   0.00921  -0.0791   0.6162   1.0000
   4.000   0.8542   0.01748   0.00938  -0.0782   0.6024   1.0000
   4.250   0.8798   0.01767   0.00951  -0.0772   0.5882   1.0000
   4.500   0.9053   0.01784   0.00966  -0.0764   0.5740   1.0000
   4.750   0.9309   0.01803   0.00985  -0.0755   0.5600   1.0000
   5.000   0.9565   0.01822   0.01003  -0.0747   0.5460   1.0000
   5.250   0.9818   0.01840   0.01021  -0.0738   0.5311   1.0000
   5.500   1.0071   0.01859   0.01043  -0.0730   0.5159   1.0000
   5.750   1.0322   0.01879   0.01070  -0.0722   0.5001   1.0000
   6.000   1.0570   0.01895   0.01091  -0.0713   0.4826   1.0000
   6.250   1.0817   0.01914   0.01115  -0.0704   0.4655   1.0000
   6.500   1.1065   0.01940   0.01148  -0.0696   0.4492   1.0000
   6.750   1.1311   0.01968   0.01183  -0.0687   0.4322   1.0000
   7.000   1.1555   0.01997   0.01214  -0.0678   0.4144   1.0000
   7.250   1.1781   0.02023   0.01254  -0.0667   0.3900   1.0000
   7.500   1.1989   0.02033   0.01261  -0.0653   0.3572   1.0000
   7.750   1.2182   0.02062   0.01292  -0.0640   0.3157   1.0000
   8.000   1.2367   0.02127   0.01349  -0.0627   0.2709   1.0000
   8.250   1.2531   0.02233   0.01438  -0.0614   0.2185   1.0000
   8.500   1.2649   0.02410   0.01578  -0.0598   0.1594   1.0000
   8.750   1.2675   0.02714   0.01826  -0.0573   0.1032   1.0000
   9.000   1.2715   0.02995   0.02088  -0.0546   0.0813   1.0000
   9.250   1.2807   0.03217   0.02305  -0.0525   0.0706   1.0000
   9.500   1.2928   0.03413   0.02500  -0.0507   0.0639   1.0000
   9.750   1.3091   0.03633   0.02725  -0.0490   0.0596   1.0000
  10.000   1.3258   0.03822   0.02927  -0.0475   0.0555   1.0000
  10.250   1.3424   0.04045   0.03139  -0.0464   0.0518   1.0000
  10.500   1.3594   0.04293   0.03413  -0.0450   0.0495   1.0000
  10.750   1.3760   0.04577   0.03727  -0.0437   0.0481   1.0000
  11.000   1.3889   0.04893   0.04078  -0.0422   0.0471   1.0000
  11.250   1.3965   0.05235   0.04456  -0.0404   0.0465   1.0000
  11.500   1.3984   0.05582   0.04839  -0.0385   0.0460   1.0000
  11.750   1.3942   0.05913   0.05202  -0.0363   0.0455   1.0000
  12.000   1.3859   0.06244   0.05560  -0.0342   0.0451   1.0000
  12.250   1.3744   0.06602   0.05944  -0.0326   0.0447   1.0000
  12.500   1.3587   0.07012   0.06382  -0.0319   0.0446   1.0000
  12.750   1.3363   0.07522   0.06923  -0.0323   0.0449   1.0000
  13.000   1.3099   0.08137   0.07569  -0.0341   0.0455   1.0000
  13.250   1.2808   0.08853   0.08313  -0.0372   0.0464   1.0000
  13.500   1.2512   0.09649   0.09133  -0.0414   0.0473   1.0000
  13.750   1.2221   0.10516   0.10018  -0.0465   0.0483   1.0000
  14.000   1.1946   0.11439   0.10955  -0.0521   0.0493   1.0000
  14.250   1.1700   0.12389   0.11911  -0.0578   0.0502   1.0000
  14.500   1.1510   0.13302   0.12828  -0.0628   0.0509   1.0000
<< Back to GOE 400 AIRFOIL (goe400-il)

Polar data table (+)

Polar graphs


<< Back to GOE 400 AIRFOIL (goe400-il)