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

GOE 550 AIRFOIL (goe550-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 550 AIRFOIL (goe550-il)
Reynolds number: 50,000
Max Cl/Cd: 33.72 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe550-il-50000-n5.txt
Download as CSV file: xf-goe550-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 550 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3737   0.10182   0.09406  -0.0477   1.0000   0.1022
  -9.750  -0.3843   0.09763   0.08995  -0.0486   1.0000   0.1027
  -9.500  -0.3980   0.09332   0.08574  -0.0494   1.0000   0.1031
  -9.250  -0.3927   0.09185   0.08434  -0.0477   1.0000   0.1046
  -9.000  -0.3903   0.09039   0.08293  -0.0461   1.0000   0.1066
  -8.750  -0.3979   0.08810   0.08073  -0.0449   1.0000   0.1083
  -8.500  -0.4130   0.08552   0.07827  -0.0435   1.0000   0.1096
  -8.250  -0.4358   0.08292   0.07581  -0.0415   1.0000   0.1106
  -8.000  -0.4615   0.07986   0.07289  -0.0401   1.0000   0.1112
  -7.750  -0.4865   0.07564   0.06877  -0.0405   1.0000   0.1120
  -7.500  -0.5123   0.06962   0.06280  -0.0429   1.0000   0.1131
  -7.250  -0.5459   0.05116   0.04373  -0.0601   0.9950   0.1178
  -7.000  -0.5112   0.05264   0.04531  -0.0598   0.9885   0.1215
  -6.750  -0.4883   0.04656   0.03873  -0.0672   0.9804   0.1288
  -6.500  -0.4595   0.04352   0.03538  -0.0713   0.9731   0.1349
  -6.250  -0.4308   0.04154   0.03311  -0.0740   0.9651   0.1426
  -6.000  -0.3973   0.03998   0.03132  -0.0769   0.9586   0.1500
  -5.750  -0.3701   0.03746   0.02830  -0.0796   0.9500   0.1602
  -5.500  -0.3361   0.03777   0.02867  -0.0807   0.9431   0.1670
  -5.250  -0.3075   0.03641   0.02699  -0.0824   0.9346   0.1770
  -5.000  -0.2732   0.03537   0.02563  -0.0847   0.9280   0.1883
  -4.750  -0.2455   0.03536   0.02565  -0.0849   0.9191   0.1960
  -4.250  -0.1837   0.03401   0.02387  -0.0874   0.9038   0.2172
  -4.000  -0.1466   0.03331   0.02299  -0.0896   0.8982   0.2270
  -3.750  -0.1211   0.03253   0.02188  -0.0900   0.8887   0.2380
  -3.500  -0.0851   0.03215   0.02151  -0.0915   0.8828   0.2468
  -3.250  -0.0588   0.03151   0.02060  -0.0918   0.8736   0.2563
  -3.000  -0.0233   0.03104   0.02004  -0.0933   0.8675   0.2663
  -2.750   0.0041   0.03067   0.01955  -0.0936   0.8588   0.2759
  -2.500   0.0392   0.03016   0.01878  -0.0951   0.8521   0.2874
  -2.250   0.0695   0.02984   0.01849  -0.0956   0.8446   0.2959
  -2.000   0.1012   0.02949   0.01800  -0.0964   0.8367   0.3067
  -1.750   0.1402   0.02907   0.01746  -0.0983   0.8316   0.3192
  -1.500   0.1620   0.02897   0.01740  -0.0974   0.8211   0.3275
  -1.250   0.2011   0.02855   0.01688  -0.0992   0.8160   0.3397
  -1.000   0.2229   0.02852   0.01684  -0.0982   0.8051   0.3485
  -0.750   0.2619   0.02813   0.01640  -0.0998   0.7996   0.3594
  -0.500   0.2855   0.02815   0.01634  -0.0992   0.7887   0.3686
  -0.250   0.3245   0.02775   0.01594  -0.1007   0.7828   0.3785
   0.000   0.3480   0.02780   0.01595  -0.0999   0.7716   0.3880
   0.250   0.3878   0.02737   0.01552  -0.1015   0.7655   0.4006
   0.500   0.4101   0.02742   0.01561  -0.1004   0.7534   0.4105
   0.750   0.4431   0.02717   0.01537  -0.1009   0.7449   0.4239
   1.000   0.4719   0.02702   0.01529  -0.1007   0.7348   0.4400
   1.250   0.4979   0.02695   0.01531  -0.1002   0.7243   0.4580
   1.500   0.5320   0.02658   0.01504  -0.1007   0.7163   0.4833
   1.750   0.5539   0.02657   0.01523  -0.0996   0.7049   0.5133
   2.000   0.5880   0.02588   0.01493  -0.0998   0.6981   0.5767
   2.250   0.6132   0.02520   0.01515  -0.0988   0.6861   1.0000
   2.500   0.6407   0.02543   0.01519  -0.0985   0.6761   1.0000
   2.750   0.6715   0.02551   0.01513  -0.0986   0.6670   1.0000
   3.000   0.6930   0.02591   0.01544  -0.0975   0.6554   1.0000
   3.250   0.7249   0.02594   0.01537  -0.0976   0.6469   1.0000
   3.500   0.7472   0.02632   0.01571  -0.0966   0.6356   1.0000
   3.750   0.7702   0.02669   0.01604  -0.0956   0.6247   1.0000
   4.000   0.8029   0.02668   0.01596  -0.0959   0.6163   1.0000
   4.250   0.8212   0.02724   0.01654  -0.0943   0.6043   1.0000
   4.500   0.8464   0.02756   0.01684  -0.0936   0.5943   1.0000
   4.750   0.8746   0.02775   0.01701  -0.0933   0.5849   1.0000
   5.000   0.8937   0.02833   0.01764  -0.0919   0.5736   1.0000
   5.250   0.9246   0.02847   0.01775  -0.0919   0.5652   1.0000
   5.500   0.9436   0.02909   0.01843  -0.0905   0.5541   1.0000
   5.750   0.9657   0.02962   0.01901  -0.0895   0.5443   1.0000
   6.000   0.9933   0.02993   0.01935  -0.0892   0.5356   1.0000
   6.250   1.0099   0.03074   0.02024  -0.0876   0.5252   1.0000
   6.500   1.0418   0.03090   0.02043  -0.0878   0.5176   1.0000
   6.750   1.0535   0.03196   0.02161  -0.0856   0.5070   1.0000
   7.000   1.0847   0.03220   0.02188  -0.0857   0.4998   1.0000
   7.250   1.0956   0.03326   0.02308  -0.0835   0.4891   1.0000
   7.500   1.1164   0.03382   0.02373  -0.0823   0.4795   1.0000
   7.750   1.1414   0.03405   0.02402  -0.0814   0.4691   1.0000
   8.000   1.1508   0.03496   0.02504  -0.0788   0.4571   1.0000
   8.250   1.1673   0.03542   0.02557  -0.0769   0.4451   1.0000
   8.500   1.1890   0.03555   0.02576  -0.0754   0.4330   1.0000
   8.750   1.2022   0.03615   0.02646  -0.0732   0.4211   1.0000
   9.000   1.2074   0.03718   0.02761  -0.0702   0.4102   1.0000
   9.250   1.2254   0.03764   0.02816  -0.0685   0.4000   1.0000
   9.500   1.2377   0.03834   0.02897  -0.0663   0.3890   1.0000
   9.750   1.2405   0.03959   0.03037  -0.0633   0.3780   1.0000
  10.000   1.2515   0.04032   0.03120  -0.0611   0.3661   1.0000
  10.250   1.2613   0.04098   0.03191  -0.0587   0.3525   1.0000
  10.500   1.2664   0.04187   0.03284  -0.0560   0.3377   1.0000
  10.750   1.2679   0.04306   0.03407  -0.0533   0.3224   1.0000
  11.000   1.2649   0.04471   0.03580  -0.0507   0.3069   1.0000
  11.250   1.2604   0.04665   0.03783  -0.0484   0.2909   1.0000
  11.500   1.2553   0.04883   0.04007  -0.0464   0.2744   1.0000
  11.750   1.2481   0.05146   0.04280  -0.0449   0.2576   1.0000
  12.000   1.2408   0.05434   0.04577  -0.0436   0.2402   1.0000
  12.250   1.2345   0.05731   0.04878  -0.0427   0.2226   1.0000
  12.500   1.2294   0.06028   0.05176  -0.0419   0.2059   1.0000
  12.750   1.2246   0.06336   0.05479  -0.0414   0.1903   1.0000
  13.000   1.2198   0.06658   0.05796  -0.0410   0.1763   1.0000
  13.250   1.2150   0.06993   0.06123  -0.0407   0.1640   1.0000
  13.500   1.2107   0.07333   0.06453  -0.0406   0.1529   1.0000
  13.750   1.2072   0.07675   0.06788  -0.0406   0.1429   1.0000
  14.000   1.2031   0.08053   0.07172  -0.0408   0.1334   1.0000
  14.250   1.2017   0.08384   0.07494  -0.0409   0.1250   1.0000
  14.500   1.1987   0.08762   0.07880  -0.0413   0.1168   1.0000
  14.750   1.1976   0.09115   0.08234  -0.0416   0.1097   1.0000
  15.000   1.1951   0.09507   0.08637  -0.0423   0.1030   1.0000
  15.250   1.1958   0.09840   0.08968  -0.0427   0.0970   1.0000
  15.500   1.1895   0.10328   0.09481  -0.0442   0.0918   1.0000
  15.750   1.1929   0.10612   0.09760  -0.0446   0.0868   1.0000
  16.000   1.1858   0.11138   0.10310  -0.0465   0.0830   1.0000
  16.250   1.1781   0.11681   0.10875  -0.0487   0.0794   1.0000
  16.500   1.1805   0.12006   0.11202  -0.0496   0.0758   1.0000
  16.750   1.1750   0.12521   0.11731  -0.0518   0.0732   1.0000
  17.000   1.1554   0.13375   0.12615  -0.0563   0.0715   1.0000
  17.250   1.1294   0.14433   0.13699  -0.0625   0.0702   1.0000
  17.500   1.0869   0.16043   0.15327  -0.0723   0.0697   1.0000
<< Back to GOE 550 AIRFOIL (goe550-il)

Polar data table (+)

Polar graphs


<< Back to GOE 550 AIRFOIL (goe550-il)