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

GOE 546 AIRFOIL (goe546-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 546 AIRFOIL (goe546-il)
Reynolds number: 100,000
Max Cl/Cd: 59.41 at α=6.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe546-il-100000.txt
Download as CSV file: xf-goe546-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 546 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
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3718   0.09672   0.09170  -0.0366   1.0000   0.1168
  -8.500  -0.3911   0.09495   0.09006  -0.0377   1.0000   0.1208
  -8.250  -0.4247   0.09413   0.08945  -0.0374   1.0000   0.1218
  -8.000  -0.4610   0.09258   0.08805  -0.0389   1.0000   0.1223
  -7.750  -0.4068   0.08753   0.08289  -0.0319   1.0000   0.1282
  -7.500  -0.4215   0.08611   0.08156  -0.0289   1.0000   0.1308
  -7.250  -0.4424   0.08450   0.08005  -0.0272   1.0000   0.1338
  -7.000  -0.4919   0.08105   0.07658  -0.0378   1.0000   0.1376
  -6.750  -0.4754   0.07820   0.07389  -0.0294   1.0000   0.1400
  -6.500  -0.4683   0.07682   0.07254  -0.0247   1.0000   0.1453
  -6.250  -0.4867   0.07160   0.06722  -0.0331   1.0000   0.1541
  -6.000  -0.4763   0.07071   0.06646  -0.0257   1.0000   0.1585
  -5.750  -0.4792   0.06652   0.06219  -0.0300   1.0000   0.1705
  -5.500  -0.4719   0.06506   0.06079  -0.0258   1.0000   0.1759
  -5.250  -0.4674   0.06169   0.05736  -0.0274   1.0000   0.1880
  -5.000  -0.4581   0.05927   0.05460  -0.0327   1.0000   0.2144
  -4.750  -0.4511   0.05651   0.05207  -0.0279   1.0000   0.2173
  -4.500  -0.3918   0.03738   0.03036  -0.0458   1.0000   0.1082
  -4.250  -0.3654   0.03409   0.02697  -0.0465   0.9985   0.1026
  -4.000  -0.3244   0.03096   0.02318  -0.0496   0.9937   0.0988
  -3.750  -0.2851   0.02899   0.02074  -0.0522   0.9876   0.0994
  -3.500  -0.2448   0.02732   0.01861  -0.0546   0.9820   0.0993
  -3.250  -0.2065   0.02622   0.01720  -0.0567   0.9752   0.1022
  -3.000  -0.1676   0.02540   0.01608  -0.0588   0.9685   0.1061
  -2.750  -0.1293   0.02432   0.01490  -0.0608   0.9617   0.1091
  -2.500  -0.0932   0.02352   0.01417  -0.0626   0.9546   0.1155
  -2.250  -0.0546   0.02287   0.01353  -0.0647   0.9474   0.1255
  -2.000  -0.0205   0.02235   0.01306  -0.0660   0.9391   0.1377
  -1.750   0.0190   0.02151   0.01250  -0.0684   0.9323   0.1726
  -1.500   0.0443   0.01986   0.01262  -0.0681   0.9241   0.5481
  -1.250   0.0788   0.01936   0.01269  -0.0684   0.9175   0.7097
  -1.000   0.1455   0.01881   0.01253  -0.0751   0.9136   1.0000
  -0.750   0.1801   0.01898   0.01244  -0.0767   0.9039   1.0000
  -0.500   0.2172   0.01915   0.01239  -0.0787   0.8956   1.0000
  -0.250   0.2563   0.01925   0.01232  -0.0810   0.8877   1.0000
   0.000   0.2894   0.01943   0.01236  -0.0821   0.8786   1.0000
   0.250   0.3314   0.01943   0.01224  -0.0847   0.8718   1.0000
   0.500   0.3621   0.01961   0.01232  -0.0853   0.8622   1.0000
   0.750   0.4071   0.01950   0.01213  -0.0883   0.8562   1.0000
   1.000   0.4361   0.01968   0.01226  -0.0886   0.8463   1.0000
   1.250   0.4794   0.01952   0.01206  -0.0911   0.8402   1.0000
   1.500   0.5054   0.01973   0.01224  -0.0907   0.8295   1.0000
   1.750   0.5489   0.01948   0.01197  -0.0930   0.8238   1.0000
   2.000   0.5729   0.01971   0.01219  -0.0922   0.8123   1.0000
   2.250   0.6183   0.01928   0.01177  -0.0945   0.8069   1.0000
   2.500   0.6438   0.01927   0.01177  -0.0936   0.7939   1.0000
   2.750   0.6740   0.01901   0.01153  -0.0931   0.7811   1.0000
   3.000   0.7058   0.01869   0.01122  -0.0929   0.7687   1.0000
   3.250   0.7433   0.01827   0.01082  -0.0936   0.7595   1.0000
   3.500   0.7705   0.01824   0.01084  -0.0930   0.7473   1.0000
   3.750   0.7966   0.01826   0.01090  -0.0921   0.7343   1.0000
   4.000   0.8241   0.01825   0.01095  -0.0915   0.7217   1.0000
   4.250   0.8530   0.01819   0.01095  -0.0911   0.7092   1.0000
   4.500   0.8819   0.01802   0.01082  -0.0904   0.6940   1.0000
   4.750   0.9103   0.01776   0.01057  -0.0895   0.6757   1.0000
   5.000   0.9413   0.01735   0.01010  -0.0887   0.6542   1.0000
   5.250   0.9628   0.01726   0.01000  -0.0866   0.6275   1.0000
   5.500   0.9850   0.01724   0.00994  -0.0847   0.6004   1.0000
   5.750   1.0062   0.01729   0.00996  -0.0827   0.5730   1.0000
   6.000   1.0252   0.01740   0.01008  -0.0805   0.5435   1.0000
   6.250   1.0425   0.01757   0.01025  -0.0781   0.5112   1.0000
   6.500   1.0586   0.01782   0.01046  -0.0755   0.4747   1.0000
   6.750   1.0723   0.01822   0.01074  -0.0725   0.4309   1.0000
   7.000   1.0819   0.01888   0.01119  -0.0691   0.3728   1.0000
   7.250   1.0804   0.02019   0.01191  -0.0640   0.2730   1.0000
   7.500   1.0691   0.02264   0.01340  -0.0582   0.1585   1.0000
   7.750   1.0682   0.02451   0.01483  -0.0538   0.1255   1.0000
   8.000   1.0721   0.02604   0.01620  -0.0502   0.1125   1.0000
   8.250   1.0814   0.02735   0.01750  -0.0475   0.1036   1.0000
   8.500   1.0937   0.02886   0.01895  -0.0452   0.0978   1.0000
   8.750   1.1127   0.03025   0.02038  -0.0439   0.0926   1.0000
   9.000   1.1448   0.03242   0.02232  -0.0449   0.0869   1.0000
   9.250   1.1684   0.03378   0.02389  -0.0442   0.0829   1.0000
   9.500   1.2007   0.03572   0.02596  -0.0450   0.0795   1.0000
   9.750   1.2362   0.03829   0.02844  -0.0466   0.0755   1.0000
  10.000   1.2608   0.04079   0.03121  -0.0463   0.0729   1.0000
  10.250   1.2799   0.04315   0.03393  -0.0450   0.0715   1.0000
  10.500   1.2940   0.04557   0.03669  -0.0433   0.0697   1.0000
  10.750   1.3058   0.04797   0.03940  -0.0414   0.0678   1.0000
  11.000   1.3183   0.05060   0.04224  -0.0398   0.0661   1.0000
  11.250   1.3257   0.05377   0.04573  -0.0377   0.0655   1.0000
  11.500   1.3293   0.05704   0.04930  -0.0353   0.0650   1.0000
  11.750   1.3301   0.06042   0.05293  -0.0330   0.0643   1.0000
  12.000   1.3275   0.06427   0.05699  -0.0306   0.0636   1.0000
  12.250   1.3176   0.06836   0.06133  -0.0277   0.0632   1.0000
  12.500   1.3013   0.07158   0.06481  -0.0242   0.0632   1.0000
  12.750   1.2791   0.07441   0.06793  -0.0208   0.0635   1.0000
  13.000   1.0716   0.07980   0.07448  -0.0137   0.0702   1.0000
  13.250   1.0346   0.08693   0.08182  -0.0155   0.0710   1.0000
<< Back to GOE 546 AIRFOIL (goe546-il)

Polar data table (+)

Polar graphs


<< Back to GOE 546 AIRFOIL (goe546-il)