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

GOE 406 AIRFOIL (goe406-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 406 AIRFOIL (goe406-il)
Reynolds number: 500,000
Max Cl/Cd: 94.84 at α=3°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe406-il-500000.txt
Download as CSV file: xf-goe406-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 406 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.1497   0.09034   0.08800  -0.0745   0.9756   0.0266
  -8.750  -0.1320   0.08558   0.08324  -0.0819   0.9717   0.0267
  -8.500  -0.1201   0.08194   0.07961  -0.0825   0.9642   0.0269
  -8.250  -0.0986   0.07871   0.07638  -0.0846   0.9594   0.0271
  -8.000  -0.0852   0.07596   0.07363  -0.0859   0.9475   0.0273
  -7.750  -0.0687   0.07278   0.07044  -0.0893   0.9320   0.0278
  -7.500  -0.0430   0.06861   0.06622  -0.0968   0.9117   0.0282
  -7.250   0.0018   0.06311   0.06059  -0.1098   0.8885   0.0295
  -6.750   0.0532   0.05133   0.04814  -0.1302   0.8288   0.0312
  -6.500   0.0646   0.04942   0.04612  -0.1295   0.8146   0.0315
  -6.250   0.0750   0.04768   0.04428  -0.1286   0.8024   0.0317
  -6.000   0.0858   0.04605   0.04258  -0.1276   0.7930   0.0321
  -5.750   0.0969   0.04428   0.04071  -0.1266   0.7836   0.0325
  -5.500   0.1071   0.04235   0.03870  -0.1254   0.7751   0.0332
  -5.250   0.1186   0.04026   0.03648  -0.1242   0.7675   0.0341
  -5.000   0.1243   0.03572   0.03147  -0.1214   0.7608   0.0363
  -4.750   0.1363   0.03410   0.02984  -0.1197   0.7535   0.0367
  -4.500   0.1515   0.03275   0.02838  -0.1183   0.7467   0.0370
  -4.250   0.1658   0.03154   0.02714  -0.1165   0.7401   0.0375
  -4.000   0.1811   0.03022   0.02570  -0.1147   0.7331   0.0383
  -3.750   0.1969   0.02875   0.02408  -0.1127   0.7267   0.0396
  -3.500   0.2073   0.02602   0.02094  -0.1086   0.7204   0.0426
  -3.250   0.2251   0.02497   0.01983  -0.1071   0.7145   0.0431
  -3.000   0.2437   0.02413   0.01894  -0.1057   0.7092   0.0440
  -2.750   0.2619   0.02319   0.01791  -0.1038   0.7031   0.0458
  -2.500   0.2774   0.02154   0.01589  -0.1007   0.6976   0.0497
  -2.250   0.2974   0.02069   0.01500  -0.0993   0.6923   0.0506
  -2.000   0.3171   0.01995   0.01421  -0.0977   0.6865   0.0523
  -1.750   0.3359   0.01903   0.01295  -0.0952   0.6812   0.0577
  -1.500   0.3571   0.01818   0.01209  -0.0941   0.6757   0.0590
  -1.250   0.3780   0.01758   0.01146  -0.0926   0.6694   0.0616
  -0.750   0.1566   0.02088   0.01668  -0.0460   0.6593   0.0488
  -0.500   0.4450   0.01518   0.00835  -0.0870   0.6507   0.0563
  -0.250   0.4679   0.01428   0.00748  -0.0863   0.6446   0.0622
   0.000   0.4895   0.01395   0.00715  -0.0851   0.6373   0.0700
   0.250   0.5256   0.01154   0.00425  -0.0850   0.6307   0.0409
   0.500   0.5488   0.01123   0.00390  -0.0838   0.6227   0.0411
   0.750   0.5712   0.01102   0.00365  -0.0825   0.6139   0.0416
   1.000   0.5934   0.01085   0.00346  -0.0811   0.6052   0.0419
   1.250   0.6139   0.01059   0.00318  -0.0794   0.5956   0.0428
   1.500   0.6351   0.01045   0.00305  -0.0779   0.5852   0.0443
   1.750   0.6557   0.01040   0.00297  -0.0762   0.5737   0.0463
   2.000   0.6753   0.01037   0.00289  -0.0743   0.5602   0.0480
   2.250   0.6942   0.01037   0.00283  -0.0723   0.5440   0.0492
   2.500   0.7120   0.01039   0.00281  -0.0700   0.5238   0.0542
   2.750   0.7277   0.01048   0.00284  -0.0673   0.4994   0.0667
   3.000   0.9768   0.01030   0.00414  -0.1188   0.4089   1.0000
   3.250   0.9956   0.01061   0.00432  -0.1169   0.3925   1.0000
   3.500   1.0144   0.01092   0.00452  -0.1151   0.3800   1.0000
   3.750   1.0348   0.01115   0.00470  -0.1137   0.3703   1.0000
   4.000   1.0541   0.01143   0.00491  -0.1120   0.3625   1.0000
   4.250   1.0748   0.01165   0.00510  -0.1106   0.3556   1.0000
   4.500   1.0937   0.01194   0.00532  -0.1088   0.3493   1.0000
   4.750   1.1141   0.01216   0.00553  -0.1074   0.3441   1.0000
   5.000   1.1344   0.01238   0.00574  -0.1059   0.3392   1.0000
   5.250   1.1533   0.01265   0.00598  -0.1042   0.3348   1.0000
   5.500   1.1712   0.01298   0.00626  -0.1023   0.3306   1.0000
   5.750   1.1918   0.01316   0.00648  -0.1009   0.3275   1.0000
   6.000   1.2115   0.01337   0.00671  -0.0993   0.3242   1.0000
   6.250   1.2298   0.01362   0.00696  -0.0975   0.3208   1.0000
   6.500   1.2443   0.01395   0.00724  -0.0950   0.3159   1.0000
   6.750   1.2611   0.01419   0.00751  -0.0928   0.3127   1.0000
   7.000   1.2781   0.01434   0.00770  -0.0908   0.3083   1.0000
   7.250   1.2944   0.01456   0.00794  -0.0886   0.3044   1.0000
   7.500   1.3093   0.01488   0.00823  -0.0862   0.3004   1.0000
   7.750   1.3254   0.01520   0.00856  -0.0841   0.2968   1.0000
   8.000   1.3431   0.01539   0.00882  -0.0823   0.2937   1.0000
   8.250   1.3598   0.01560   0.00908  -0.0803   0.2896   1.0000
   8.500   1.3750   0.01591   0.00938  -0.0781   0.2854   1.0000
   8.750   1.3896   0.01633   0.00979  -0.0759   0.2814   1.0000
   9.000   1.4073   0.01651   0.01007  -0.0742   0.2779   1.0000
   9.250   1.4240   0.01677   0.01038  -0.0724   0.2735   1.0000
   9.500   1.4387   0.01712   0.01075  -0.0703   0.2694   1.0000
   9.750   1.4527   0.01755   0.01119  -0.0681   0.2651   1.0000
  10.000   1.4702   0.01779   0.01153  -0.0666   0.2605   1.0000
  10.250   1.4841   0.01819   0.01195  -0.0645   0.2541   1.0000
  10.500   1.4982   0.01861   0.01240  -0.0625   0.2473   1.0000
  10.750   1.5116   0.01907   0.01289  -0.0604   0.2387   1.0000
  11.000   1.5246   0.01959   0.01343  -0.0584   0.2282   1.0000
  11.250   1.5335   0.02033   0.01413  -0.0559   0.2113   1.0000
  11.500   1.5321   0.02171   0.01531  -0.0523   0.1770   1.0000
  11.750   1.5124   0.02439   0.01762  -0.0469   0.1246   1.0000
  12.000   1.4973   0.02708   0.02013  -0.0427   0.0928   1.0000
  12.250   1.4800   0.03019   0.02302  -0.0388   0.0589   1.0000
  12.500   1.4746   0.03255   0.02535  -0.0363   0.0489   1.0000
  13.000   1.4735   0.03675   0.02967  -0.0327   0.0418   1.0000
  13.250   1.4731   0.03895   0.03193  -0.0312   0.0397   1.0000
  13.500   1.4676   0.04172   0.03476  -0.0296   0.0377   1.0000
  13.750   1.4665   0.04415   0.03729  -0.0283   0.0366   1.0000
  14.000   1.4662   0.04657   0.03981  -0.0273   0.0353   1.0000
  14.250   1.4633   0.04935   0.04269  -0.0264   0.0342   1.0000
  14.500   1.4590   0.05238   0.04580  -0.0256   0.0334   1.0000
  14.750   1.4479   0.05628   0.04978  -0.0249   0.0322   1.0000
  15.000   1.4367   0.06039   0.05400  -0.0245   0.0316   1.0000
  15.250   1.4294   0.06417   0.05789  -0.0243   0.0310   1.0000
  15.500   1.4251   0.06764   0.06147  -0.0243   0.0303   1.0000
  15.750   1.4174   0.07167   0.06561  -0.0244   0.0295   1.0000
  16.000   1.4094   0.07583   0.06987  -0.0247   0.0290   1.0000
  16.250   1.4006   0.08018   0.07433  -0.0252   0.0285   1.0000
  16.500   1.3915   0.08463   0.07887  -0.0258   0.0280   1.0000
  16.750   1.3808   0.08940   0.08373  -0.0265   0.0277   1.0000
  17.000   1.3693   0.09433   0.08874  -0.0274   0.0271   1.0000
  17.250   1.3567   0.09940   0.09386  -0.0283   0.0266   1.0000
<< Back to GOE 406 AIRFOIL (goe406-il)

Polar data table (+)

Polar graphs


<< Back to GOE 406 AIRFOIL (goe406-il)