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

GOE 628 AIRFOIL (goe628-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 628 AIRFOIL (goe628-il)
Reynolds number: 200,000
Max Cl/Cd: 64.78 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe628-il-200000.txt
Download as CSV file: xf-goe628-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 628 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.500  -0.4135   0.06263   0.05899  -0.0902   0.9762   0.0841
 -10.250  -0.4572   0.04602   0.04187  -0.1127   0.9543   0.0821
 -10.000  -0.4587   0.03933   0.03457  -0.1189   0.9413   0.0822
  -9.750  -0.4359   0.03509   0.02976  -0.1236   0.9356   0.0830
  -9.500  -0.4174   0.03262   0.02688  -0.1244   0.9246   0.0834
  -9.250  -0.3877   0.03027   0.02414  -0.1266   0.9187   0.0839
  -9.000  -0.3638   0.02816   0.02192  -0.1270   0.9089   0.0849
  -8.750  -0.3325   0.02664   0.02030  -0.1282   0.9018   0.0859
  -8.500  -0.3066   0.02540   0.01890  -0.1282   0.8922   0.0866
  -8.250  -0.2783   0.02420   0.01755  -0.1285   0.8834   0.0874
  -8.000  -0.2534   0.02322   0.01640  -0.1281   0.8734   0.0882
  -7.750  -0.2259   0.02230   0.01534  -0.1280   0.8644   0.0894
  -7.500  -0.2014   0.02154   0.01443  -0.1273   0.8542   0.0906
  -7.250  -0.1743   0.02077   0.01348  -0.1270   0.8453   0.0917
  -7.000  -0.1496   0.02013   0.01270  -0.1262   0.8354   0.0926
  -6.750  -0.1228   0.01941   0.01185  -0.1258   0.8267   0.0935
  -6.500  -0.0982   0.01866   0.01111  -0.1251   0.8174   0.0949
  -6.250  -0.0721   0.01809   0.01053  -0.1246   0.8086   0.0965
  -6.000  -0.0466   0.01766   0.01005  -0.1240   0.8000   0.0986
  -5.750  -0.0209   0.01725   0.00957  -0.1233   0.7910   0.1010
  -5.500   0.0057   0.01691   0.00910  -0.1228   0.7829   0.1034
  -5.250   0.0295   0.01639   0.00863  -0.1219   0.7733   0.1064
  -5.000   0.0555   0.01603   0.00822  -0.1212   0.7642   0.1102
  -4.750   0.0804   0.01571   0.00783  -0.1203   0.7536   0.1152
  -4.500   0.1054   0.01532   0.00744  -0.1196   0.7446   0.1237
  -4.250   0.1298   0.01491   0.00709  -0.1187   0.7354   0.1410
  -4.000   0.1558   0.01450   0.00689  -0.1183   0.7281   0.1921
  -3.750   0.1806   0.01451   0.00697  -0.1175   0.7191   0.2292
  -3.500   0.2090   0.01463   0.00697  -0.1172   0.7117   0.2504
  -3.250   0.2346   0.01475   0.00708  -0.1165   0.7033   0.2647
  -3.000   0.2615   0.01482   0.00709  -0.1160   0.6952   0.2770
  -2.750   0.2890   0.01498   0.00711  -0.1155   0.6878   0.2898
  -2.500   0.3143   0.01505   0.00724  -0.1148   0.6792   0.3008
  -2.250   0.3426   0.01512   0.00717  -0.1145   0.6723   0.3114
  -2.000   0.3674   0.01517   0.00726  -0.1136   0.6636   0.3204
  -1.750   0.3943   0.01518   0.00718  -0.1131   0.6556   0.3304
  -1.500   0.4205   0.01523   0.00722  -0.1125   0.6477   0.3394
  -1.250   0.4464   0.01523   0.00716  -0.1118   0.6391   0.3490
  -1.000   0.4740   0.01526   0.00716  -0.1114   0.6320   0.3586
  -0.750   0.4983   0.01527   0.00715  -0.1105   0.6225   0.3688
  -0.500   0.5257   0.01528   0.00712  -0.1101   0.6148   0.3794
  -0.250   0.5496   0.01526   0.00711  -0.1091   0.6049   0.3895
   0.000   0.5764   0.01524   0.00704  -0.1086   0.5966   0.3997
   0.250   0.6006   0.01521   0.00701  -0.1077   0.5868   0.4087
   0.500   0.6272   0.01516   0.00691  -0.1072   0.5782   0.4171
   0.750   0.6512   0.01516   0.00688  -0.1062   0.5676   0.4260
   1.000   0.6777   0.01511   0.00678  -0.1057   0.5585   0.4349
   1.250   0.7006   0.01511   0.00678  -0.1046   0.5468   0.4446
   1.500   0.7255   0.01508   0.00674  -0.1038   0.5367   0.4538
   1.750   0.7493   0.01509   0.00673  -0.1029   0.5256   0.4649
   2.000   0.7728   0.01511   0.00678  -0.1019   0.5147   0.4765
   2.250   0.7972   0.01515   0.00676  -0.1011   0.5044   0.4890
   2.500   0.8193   0.01520   0.00686  -0.0999   0.4926   0.5034
   2.750   0.8427   0.01527   0.00692  -0.0989   0.4822   0.5210
   3.000   0.8645   0.01529   0.00702  -0.0977   0.4714   0.5425
   3.250   0.8863   0.01532   0.00717  -0.0965   0.4617   0.5754
   3.500   0.9069   0.01512   0.00731  -0.0950   0.4527   0.6775
   3.750   0.9659   0.01506   0.00762  -0.1013   0.4409   1.0000
   4.000   0.9896   0.01538   0.00776  -0.1005   0.4331   1.0000
   4.250   1.0114   0.01568   0.00802  -0.0994   0.4249   1.0000
   4.500   1.0341   0.01598   0.00822  -0.0984   0.4177   1.0000
   4.750   1.0579   0.01636   0.00849  -0.0977   0.4113   1.0000
   5.000   1.0795   0.01667   0.00879  -0.0965   0.4048   1.0000
   5.250   1.1032   0.01703   0.00905  -0.0958   0.3992   1.0000
   5.500   1.1272   0.01744   0.00939  -0.0952   0.3937   1.0000
   5.750   1.1480   0.01777   0.00974  -0.0939   0.3880   1.0000
   6.000   1.1707   0.01812   0.01003  -0.0930   0.3828   1.0000
   6.250   1.1979   0.01862   0.01040  -0.0931   0.3782   1.0000
   6.500   1.2172   0.01898   0.01084  -0.0916   0.3737   1.0000
   6.750   1.2387   0.01936   0.01123  -0.0906   0.3692   1.0000
   7.000   1.2620   0.01976   0.01159  -0.0899   0.3650   1.0000
   7.250   1.2906   0.02033   0.01203  -0.0903   0.3607   1.0000
   7.500   1.3072   0.02071   0.01253  -0.0884   0.3568   1.0000
   7.750   1.3273   0.02114   0.01300  -0.0872   0.3528   1.0000
   8.000   1.3497   0.02158   0.01343  -0.0864   0.3491   1.0000
   8.250   1.3759   0.02208   0.01386  -0.0864   0.3455   1.0000
   8.500   1.3987   0.02267   0.01447  -0.0858   0.3418   1.0000
   8.750   1.4137   0.02314   0.01505  -0.0838   0.3382   1.0000
   9.000   1.4323   0.02362   0.01558  -0.0825   0.3347   1.0000
   9.250   1.4539   0.02408   0.01604  -0.0816   0.3311   1.0000
   9.500   1.4822   0.02465   0.01652  -0.0820   0.3274   1.0000
   9.750   1.4983   0.02527   0.01723  -0.0804   0.3240   1.0000
  10.000   1.5057   0.02576   0.01784  -0.0772   0.3203   1.0000
  10.250   1.5185   0.02623   0.01837  -0.0750   0.3165   1.0000
  10.500   1.5386   0.02668   0.01880  -0.0740   0.3130   1.0000
  10.750   1.5749   0.02741   0.01941  -0.0759   0.3089   1.0000
  11.000   1.5737   0.02801   0.02020  -0.0715   0.3062   1.0000
  11.250   1.5787   0.02867   0.02100  -0.0684   0.3030   1.0000
  11.500   1.5899   0.02929   0.02168  -0.0663   0.2995   1.0000
  11.750   1.6082   0.02982   0.02223  -0.0652   0.2963   1.0000
  12.000   1.6403   0.03039   0.02271  -0.0662   0.2928   1.0000
  12.250   1.6418   0.03130   0.02378  -0.0630   0.2898   1.0000
  12.500   1.6401   0.03226   0.02491  -0.0595   0.2867   1.0000
  12.750   1.6456   0.03311   0.02586  -0.0571   0.2832   1.0000
  13.000   1.6601   0.03375   0.02653  -0.0558   0.2799   1.0000
  13.250   1.6893   0.03415   0.02685  -0.0563   0.2764   1.0000
  13.500   1.6883   0.03542   0.02827  -0.0535   0.2733   1.0000
  13.750   1.6787   0.03693   0.02997  -0.0500   0.2700   1.0000
  14.000   1.6797   0.03814   0.03129  -0.0478   0.2664   1.0000
  14.250   1.6930   0.03883   0.03200  -0.0467   0.2629   1.0000
  14.500   1.7229   0.03907   0.03212  -0.0471   0.2589   1.0000
  14.750   1.6990   0.04156   0.03489  -0.0434   0.2558   1.0000
  15.000   1.6878   0.04369   0.03718  -0.0411   0.2518   1.0000
  15.250   1.6931   0.04489   0.03842  -0.0399   0.2479   1.0000
  15.500   1.7216   0.04476   0.03816  -0.0399   0.2437   1.0000
  15.750   1.6969   0.04819   0.04185  -0.0376   0.2403   1.0000
  16.000   1.6807   0.05129   0.04514  -0.0362   0.2362   1.0000
  16.250   1.6845   0.05287   0.04675  -0.0354   0.2321   1.0000
  16.500   1.7121   0.05259   0.04633  -0.0352   0.2279   1.0000
  16.750   1.6758   0.05794   0.05200  -0.0341   0.2241   1.0000
  17.000   1.6613   0.06164   0.05583  -0.0338   0.2196   1.0000
  17.250   1.6727   0.06265   0.05681  -0.0335   0.2152   1.0000
  17.500   1.6650   0.06593   0.06018  -0.0334   0.2109   1.0000
  17.750   1.6367   0.07175   0.06621  -0.0340   0.2062   1.0000
  18.000   1.6413   0.07372   0.06818  -0.0341   0.2014   1.0000
  18.250   1.6403   0.07650   0.07097  -0.0343   0.1969   1.0000
  18.500   1.6056   0.08387   0.07857  -0.0360   0.1920   1.0000
  18.750   1.6095   0.08615   0.08085  -0.0364   0.1874   1.0000
  19.000   1.6077   0.08924   0.08396  -0.0370   0.1828   1.0000
<< Back to GOE 628 AIRFOIL (goe628-il)

Polar data table (+)

Polar graphs


<< Back to GOE 628 AIRFOIL (goe628-il)