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

GOE 626 AIRFOIL (goe626-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: GOE 626 AIRFOIL (goe626-il)
Reynolds number: 200,000
Max Cl/Cd: 71.23 at α=7.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe626-il-200000.txt
Download as CSV file: xf-goe626-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 626 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
  ------ -------- --------- --------- -------- -------- --------
 -13.500  -0.6644   0.07045   0.06576  -0.0806   1.0000   0.0553
 -13.250  -0.7150   0.06285   0.05795  -0.0841   1.0000   0.0548
 -13.000  -0.7448   0.05845   0.05345  -0.0847   1.0000   0.0546
 -12.750  -0.7704   0.05506   0.04998  -0.0843   1.0000   0.0546
 -12.500  -0.7960   0.05218   0.04703  -0.0830   1.0000   0.0545
 -12.250  -0.8246   0.04959   0.04439  -0.0811   1.0000   0.0544
 -12.000  -0.8602   0.04709   0.04183  -0.0784   1.0000   0.0542
 -11.750  -0.8749   0.04425   0.03890  -0.0787   1.0000   0.0544
 -11.500  -0.8785   0.04155   0.03610  -0.0798   1.0000   0.0547
 -11.250  -0.8734   0.03920   0.03362  -0.0810   0.9997   0.0550
 -11.000  -0.8435   0.03705   0.03134  -0.0853   0.9957   0.0558
 -10.750  -0.8119   0.03514   0.02927  -0.0894   0.9918   0.0569
 -10.500  -0.7843   0.03342   0.02739  -0.0922   0.9871   0.0579
 -10.250  -0.7533   0.03202   0.02589  -0.0947   0.9831   0.0589
 -10.000  -0.7189   0.03096   0.02490  -0.0974   0.9800   0.0601
  -9.750  -0.6935   0.02982   0.02375  -0.0985   0.9740   0.0613
  -9.500  -0.6589   0.02869   0.02257  -0.1013   0.9697   0.0629
  -9.250  -0.6196   0.02761   0.02139  -0.1048   0.9668   0.0648
  -9.000  -0.5967   0.02655   0.02034  -0.1052   0.9599   0.0665
  -8.750  -0.5605   0.02548   0.01925  -0.1082   0.9555   0.0696
  -8.500  -0.5192   0.02441   0.01813  -0.1119   0.9526   0.0740
  -8.250  -0.4941   0.02339   0.01711  -0.1125   0.9451   0.0787
  -8.000  -0.4577   0.02210   0.01581  -0.1154   0.9405   0.0885
  -7.750  -0.4165   0.02036   0.01417  -0.1197   0.9375   0.1144
  -7.500  -0.3893   0.01932   0.01324  -0.1206   0.9300   0.1435
  -7.250  -0.3513   0.01877   0.01277  -0.1228   0.9248   0.1719
  -7.000  -0.3061   0.01865   0.01270  -0.1258   0.9215   0.1945
  -6.750  -0.2709   0.01879   0.01282  -0.1267   0.9148   0.2093
  -6.500  -0.2322   0.01886   0.01279  -0.1282   0.9085   0.2217
  -6.250  -0.1852   0.01912   0.01305  -0.1309   0.9053   0.2306
  -6.000  -0.1358   0.01924   0.01313  -0.1342   0.9033   0.2393
  -5.750  -0.1068   0.01959   0.01350  -0.1336   0.8941   0.2451
  -5.500  -0.0641   0.01948   0.01326  -0.1359   0.8895   0.2536
  -5.250  -0.0184   0.01990   0.01381  -0.1381   0.8861   0.2588
  -5.000   0.0130   0.01990   0.01373  -0.1382   0.8770   0.2655
  -4.750   0.0509   0.01963   0.01337  -0.1396   0.8701   0.2713
  -4.500   0.0902   0.01988   0.01372  -0.1407   0.8634   0.2760
  -4.250   0.1203   0.01991   0.01370  -0.1405   0.8518   0.2826
  -4.000   0.1524   0.01956   0.01322  -0.1410   0.8411   0.2889
  -3.750   0.1862   0.01960   0.01332  -0.1413   0.8302   0.2928
  -3.500   0.2139   0.01959   0.01329  -0.1407   0.8157   0.2978
  -3.250   0.2443   0.01928   0.01268  -0.1411   0.8020   0.3053
  -3.000   0.2743   0.01900   0.01244  -0.1410   0.7886   0.3092
  -2.750   0.3020   0.01895   0.01238  -0.1405   0.7733   0.3129
  -2.500   0.3287   0.01873   0.01205  -0.1400   0.7574   0.3173
  -2.250   0.3565   0.01840   0.01148  -0.1400   0.7419   0.3221
  -2.000   0.3840   0.01792   0.01081  -0.1399   0.7269   0.3249
  -1.750   0.4113   0.01761   0.01042  -0.1396   0.7127   0.3274
  -1.500   0.4385   0.01749   0.01022  -0.1392   0.6988   0.3303
  -1.250   0.4648   0.01736   0.01000  -0.1387   0.6848   0.3333
  -1.000   0.4919   0.01720   0.00969  -0.1385   0.6718   0.3361
  -0.750   0.5203   0.01709   0.00936  -0.1384   0.6597   0.3390
  -0.500   0.5466   0.01708   0.00918  -0.1380   0.6469   0.3421
  -0.250   0.5735   0.01686   0.00886  -0.1378   0.6358   0.3445
   0.000   0.6003   0.01679   0.00871  -0.1374   0.6248   0.3467
   0.250   0.6264   0.01677   0.00865  -0.1370   0.6138   0.3490
   0.500   0.6540   0.01682   0.00859  -0.1368   0.6036   0.3517
   0.750   0.6791   0.01683   0.00858  -0.1361   0.5927   0.3548
   1.000   0.7073   0.01691   0.00848  -0.1360   0.5831   0.3576
   1.250   0.7317   0.01689   0.00844  -0.1352   0.5723   0.3598
   1.750   0.7832   0.01682   0.00827  -0.1341   0.5529   0.3642
   2.000   0.8102   0.01687   0.00824  -0.1338   0.5444   0.3671
   2.250   0.8339   0.01688   0.00830  -0.1329   0.5347   0.3702
   2.500   0.8604   0.01694   0.00829  -0.1324   0.5267   0.3733
   2.750   0.8854   0.01700   0.00835  -0.1317   0.5183   0.3761
   3.000   0.9110   0.01707   0.00836  -0.1312   0.5104   0.3788
   3.250   0.9373   0.01718   0.00841  -0.1307   0.5031   0.3814
   3.500   0.9613   0.01715   0.00845  -0.1299   0.4948   0.3844
   3.750   0.9879   0.01727   0.00851  -0.1295   0.4879   0.3879
   4.000   1.0115   0.01737   0.00869  -0.1286   0.4804   0.3920
   4.250   1.0367   0.01750   0.00880  -0.1280   0.4737   0.3964
   4.500   1.0640   0.01772   0.00894  -0.1278   0.4678   0.4003
   4.750   1.0866   0.01776   0.00912  -0.1267   0.4605   0.4042
   5.000   1.1111   0.01788   0.00923  -0.1259   0.4533   0.4088
   5.250   1.1349   0.01805   0.00943  -0.1251   0.4460   0.4140
   5.500   1.1573   0.01818   0.00958  -0.1240   0.4382   0.4193
   5.750   1.1821   0.01834   0.00974  -0.1233   0.4310   0.4254
   6.000   1.2022   0.01845   0.00998  -0.1218   0.4232   0.4327
   6.250   1.2253   0.01862   0.01014  -0.1208   0.4159   0.4413
   6.500   1.2470   0.01880   0.01042  -0.1197   0.4090   0.4513
   6.750   1.2675   0.01894   0.01067  -0.1183   0.4018   0.4630
   7.250   1.3089   0.01927   0.01124  -0.1156   0.3882   0.5033
   7.500   1.3287   0.01931   0.01151  -0.1142   0.3809   0.5643
   7.750   1.3499   0.01895   0.01178  -0.1128   0.3738   1.0000
   8.000   1.3665   0.01922   0.01207  -0.1107   0.3660   1.0000
   8.250   1.3852   0.01958   0.01231  -0.1091   0.3589   1.0000
   8.500   1.3973   0.01986   0.01269  -0.1063   0.3509   1.0000
   8.750   1.4113   0.02022   0.01299  -0.1038   0.3428   1.0000
   9.000   1.4229   0.02059   0.01342  -0.1011   0.3336   1.0000
   9.250   1.4351   0.02104   0.01382  -0.0985   0.3250   1.0000
   9.500   1.4468   0.02149   0.01435  -0.0960   0.3152   1.0000
   9.750   1.4570   0.02207   0.01486  -0.0933   0.3061   1.0000
  10.000   1.4668   0.02265   0.01551  -0.0907   0.2949   1.0000
  10.250   1.4755   0.02335   0.01621  -0.0881   0.2839   1.0000
  10.500   1.4818   0.02420   0.01701  -0.0853   0.2729   1.0000
  10.750   1.4901   0.02506   0.01790  -0.0829   0.2611   1.0000
  11.000   1.4974   0.02604   0.01888  -0.0805   0.2508   1.0000
  11.250   1.5022   0.02720   0.01999  -0.0780   0.2413   1.0000
  11.500   1.5094   0.02832   0.02115  -0.0760   0.2313   1.0000
  11.750   1.5135   0.02968   0.02248  -0.0737   0.2227   1.0000
  12.000   1.5190   0.03104   0.02386  -0.0718   0.2144   1.0000
  12.250   1.5225   0.03259   0.02539  -0.0698   0.2075   1.0000
  12.500   1.5273   0.03413   0.02698  -0.0681   0.2003   1.0000
  12.750   1.5276   0.03605   0.02887  -0.0662   0.1939   1.0000
  13.000   1.5320   0.03778   0.03068  -0.0648   0.1871   1.0000
  13.250   1.5305   0.04004   0.03291  -0.0633   0.1806   1.0000
  13.500   1.5323   0.04215   0.03510  -0.0621   0.1740   1.0000
  13.750   1.5299   0.04473   0.03768  -0.0610   0.1673   1.0000
  14.000   1.5284   0.04737   0.04039  -0.0601   0.1605   1.0000
  14.250   1.5241   0.05040   0.04345  -0.0595   0.1532   1.0000
  14.500   1.5201   0.05356   0.04668  -0.0590   0.1455   1.0000
  14.750   1.5133   0.05715   0.05027  -0.0588   0.1381   1.0000
  15.000   1.5094   0.06059   0.05381  -0.0588   0.1300   1.0000
  15.250   1.5013   0.06455   0.05774  -0.0590   0.1238   1.0000
  15.500   1.4999   0.06785   0.06114  -0.0591   0.1177   1.0000
  15.750   1.4951   0.07160   0.06491  -0.0595   0.1133   1.0000
  16.000   1.4933   0.07501   0.06833  -0.0598   0.1099   1.0000
  16.250   1.4935   0.07821   0.07162  -0.0602   0.1065   1.0000
  16.500   1.4928   0.08158   0.07502  -0.0606   0.1038   1.0000
  16.750   1.4920   0.08496   0.07838  -0.0612   0.1015   1.0000
  17.000   1.4942   0.08792   0.08137  -0.0615   0.0996   1.0000
  17.250   1.4973   0.09082   0.08437  -0.0620   0.0977   1.0000
  17.500   1.4996   0.09388   0.08750  -0.0626   0.0958   1.0000
<< Back to GOE 626 AIRFOIL (goe626-il)

Polar data table (+)

Polar graphs


<< Back to GOE 626 AIRFOIL (goe626-il)