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

GOE 535 AIRFOIL (goe535-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 535 AIRFOIL (goe535-il)
Reynolds number: 100,000
Max Cl/Cd: 50.88 at α=6.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe535-il-100000-n5.txt
Download as CSV file: xf-goe535-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 535 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -13.500  -0.2335   0.14937   0.14355  -0.0386   1.0000   0.0777
 -13.250  -0.2345   0.14682   0.14104  -0.0396   1.0000   0.0782
 -13.000  -0.2351   0.14418   0.13843  -0.0405   1.0000   0.0784
 -12.750  -0.2335   0.14152   0.13582  -0.0410   1.0000   0.0784
 -12.500  -0.2329   0.13879   0.13313  -0.0414   1.0000   0.0785
 -12.250  -0.2312   0.13609   0.13047  -0.0416   1.0000   0.0785
 -11.750  -0.2303   0.13075   0.12522  -0.0417   1.0000   0.0785
 -11.250  -0.2325   0.12201   0.11653  -0.0418   1.0000   0.0645
 -11.000  -0.2281   0.11998   0.11455  -0.0406   1.0000   0.0639
 -10.750  -0.2285   0.11772   0.11236  -0.0398   1.0000   0.0633
 -10.500  -0.2317   0.11544   0.11014  -0.0389   1.0000   0.0626
 -10.250  -0.2273   0.11187   0.10660  -0.0408   0.9979   0.0621
 -10.000  -0.2116   0.10680   0.10152  -0.0463   0.9918   0.0623
  -9.750  -0.1985   0.10176   0.09647  -0.0516   0.9846   0.0629
  -9.500  -0.1882   0.09626   0.09094  -0.0574   0.9772   0.0636
  -9.250  -0.1786   0.09078   0.08545  -0.0628   0.9684   0.0637
  -9.000  -0.1731   0.08511   0.07978  -0.0680   0.9580   0.0638
  -8.750  -0.1772   0.07717   0.07181  -0.0756   0.9474   0.0642
  -8.500  -0.1794   0.06991   0.06454  -0.0827   0.9353   0.0653
  -8.250  -0.2572   0.04613   0.04017  -0.1123   0.9036   0.0663
  -8.000  -0.2425   0.04117   0.03480  -0.1192   0.8899   0.0679
  -7.750  -0.2208   0.03639   0.02940  -0.1253   0.8791   0.0697
  -7.500  -0.2015   0.03319   0.02564  -0.1277   0.8651   0.0716
  -7.250  -0.1680   0.03209   0.02452  -0.1293   0.8566   0.0733
  -7.000  -0.1420   0.03078   0.02304  -0.1299   0.8434   0.0754
  -6.750  -0.1133   0.02886   0.02061  -0.1315   0.8321   0.0783
  -6.500  -0.0829   0.02772   0.01940  -0.1323   0.8207   0.0801
  -6.250  -0.0545   0.02684   0.01842  -0.1326   0.8080   0.0825
  -6.000  -0.0220   0.02571   0.01696  -0.1337   0.7973   0.0858
  -5.750   0.0044   0.02484   0.01599  -0.1336   0.7834   0.0882
  -5.500   0.0342   0.02416   0.01523  -0.1339   0.7715   0.0911
  -5.250   0.0631   0.02344   0.01429  -0.1340   0.7587   0.0945
  -5.000   0.0905   0.02276   0.01351  -0.1338   0.7458   0.0974
  -4.750   0.1201   0.02221   0.01290  -0.1339   0.7343   0.1006
  -4.500   0.1462   0.02174   0.01231  -0.1334   0.7211   0.1040
  -4.250   0.1745   0.02123   0.01168  -0.1332   0.7096   0.1074
  -4.000   0.2003   0.02083   0.01130  -0.1326   0.6968   0.1107
  -3.750   0.2273   0.02048   0.01085  -0.1322   0.6853   0.1144
  -3.500   0.2538   0.02016   0.01044  -0.1316   0.6735   0.1182
  -3.250   0.2797   0.01985   0.01015  -0.1311   0.6621   0.1222
  -3.000   0.3062   0.01961   0.00983  -0.1305   0.6508   0.1267
  -2.750   0.3319   0.01939   0.00957  -0.1299   0.6395   0.1315
  -2.500   0.3584   0.01917   0.00931  -0.1294   0.6288   0.1375
  -2.250   0.3843   0.01901   0.00910  -0.1288   0.6181   0.1442
  -2.000   0.4108   0.01880   0.00887  -0.1284   0.6080   0.1536
  -1.750   0.4368   0.01860   0.00869  -0.1279   0.5978   0.1680
  -1.500   0.4630   0.01835   0.00848  -0.1275   0.5878   0.1942
  -1.250   0.4889   0.01803   0.00842  -0.1272   0.5786   0.2554
  -1.000   0.5143   0.01806   0.00872  -0.1265   0.5691   0.3413
  -0.750   0.5414   0.01826   0.00882  -0.1259   0.5610   0.3817
  -0.500   0.5671   0.01847   0.00898  -0.1252   0.5515   0.4075
  -0.250   0.5936   0.01869   0.00910  -0.1245   0.5438   0.4269
   0.000   0.6190   0.01892   0.00932  -0.1237   0.5356   0.4428
   0.250   0.6445   0.01918   0.00953  -0.1229   0.5277   0.4607
   0.500   0.6710   0.01946   0.00968  -0.1222   0.5212   0.4791
   0.750   0.6952   0.01974   0.01000  -0.1212   0.5132   0.4957
   1.000   0.7200   0.02000   0.01025  -0.1202   0.5064   0.5108
   1.250   0.7462   0.02024   0.01039  -0.1195   0.5005   0.5248
   1.500   0.7705   0.02045   0.01061  -0.1186   0.4933   0.5352
   1.750   0.7965   0.02061   0.01067  -0.1181   0.4869   0.5421
   2.000   0.8234   0.02075   0.01071  -0.1177   0.4816   0.5474
   2.250   0.8477   0.02094   0.01092  -0.1170   0.4750   0.5535
   2.500   0.8731   0.02114   0.01106  -0.1164   0.4688   0.5595
   2.750   0.8996   0.02130   0.01114  -0.1160   0.4634   0.5646
   3.000   0.9253   0.02152   0.01135  -0.1155   0.4577   0.5710
   3.250   0.9504   0.02178   0.01159  -0.1150   0.4515   0.5779
   3.500   0.9766   0.02198   0.01175  -0.1146   0.4457   0.5836
   3.750   1.0046   0.02221   0.01187  -0.1145   0.4407   0.5900
   4.000   1.0279   0.02251   0.01222  -0.1137   0.4343   0.5973
   4.250   1.0528   0.02273   0.01247  -0.1131   0.4281   0.6043
   4.500   1.0794   0.02297   0.01262  -0.1128   0.4228   0.6131
   4.750   1.1034   0.02326   0.01296  -0.1122   0.4168   0.6215
   5.000   1.1265   0.02354   0.01331  -0.1114   0.4105   0.6308
   5.250   1.1514   0.02378   0.01355  -0.1108   0.4049   0.6405
   5.500   1.1771   0.02406   0.01381  -0.1105   0.3999   0.6521
   5.750   1.1979   0.02439   0.01431  -0.1094   0.3937   0.6656
   6.000   1.2207   0.02465   0.01467  -0.1085   0.3879   0.6844
   6.500   1.2669   0.02490   0.01535  -0.1070   0.3775   1.0000
   6.750   1.2865   0.02539   0.01587  -0.1058   0.3716   1.0000
   7.000   1.3079   0.02582   0.01625  -0.1048   0.3665   1.0000
   7.250   1.3312   0.02625   0.01658  -0.1041   0.3621   1.0000
   7.500   1.3459   0.02685   0.01730  -0.1022   0.3561   1.0000
   7.750   1.3632   0.02736   0.01783  -0.1006   0.3505   1.0000
   8.000   1.3832   0.02779   0.01818  -0.0994   0.3458   1.0000
   8.250   1.3972   0.02839   0.01884  -0.0974   0.3405   1.0000
   8.500   1.4087   0.02904   0.01955  -0.0950   0.3348   1.0000
   8.750   1.4235   0.02959   0.02011  -0.0931   0.3300   1.0000
   9.000   1.4426   0.03008   0.02052  -0.0919   0.3259   1.0000
   9.250   1.4486   0.03100   0.02158  -0.0891   0.3203   1.0000
   9.500   1.4590   0.03176   0.02241  -0.0869   0.3151   1.0000
   9.750   1.4735   0.03239   0.02301  -0.0852   0.3106   1.0000
  10.000   1.4831   0.03329   0.02396  -0.0831   0.3058   1.0000
  10.250   1.4885   0.03441   0.02519  -0.0808   0.3005   1.0000
  10.500   1.4986   0.03534   0.02615  -0.0789   0.2959   1.0000
  10.750   1.5148   0.03601   0.02677  -0.0777   0.2921   1.0000
  11.000   1.5160   0.03756   0.02848  -0.0754   0.2873   1.0000
  11.250   1.5194   0.03902   0.03004  -0.0734   0.2825   1.0000
  11.500   1.5279   0.04016   0.03119  -0.0719   0.2780   1.0000
  11.750   1.5383   0.04125   0.03226  -0.0706   0.2738   1.0000
  12.000   1.5318   0.04359   0.03480  -0.0686   0.2686   1.0000
  12.250   1.5332   0.04541   0.03668  -0.0671   0.2637   1.0000
  12.500   1.5456   0.04640   0.03760  -0.0661   0.2594   1.0000
  12.750   1.5375   0.04926   0.04063  -0.0648   0.2549   1.0000
  13.000   1.5297   0.05220   0.04371  -0.0638   0.2497   1.0000
  13.250   1.5343   0.05400   0.04552  -0.0630   0.2451   1.0000
  13.500   1.5341   0.05641   0.04797  -0.0623   0.2407   1.0000
  13.750   1.5154   0.06097   0.05274  -0.0620   0.2356   1.0000
  14.000   1.5123   0.06389   0.05571  -0.0618   0.2308   1.0000
  14.250   1.5220   0.06532   0.05708  -0.0613   0.2265   1.0000
  14.500   1.4891   0.07233   0.06436  -0.0623   0.2214   1.0000
  14.750   1.4777   0.07677   0.06890  -0.0629   0.2166   1.0000
  15.250   1.4619   0.08513   0.07743  -0.0642   0.2087   1.0000
  15.500   1.4204   0.09471   0.08724  -0.0672   0.2032   1.0000
  15.750   1.4323   0.09603   0.08858  -0.0671   0.1997   1.0000
  16.000   1.4460   0.09708   0.08960  -0.0668   0.1965   1.0000
  16.250   1.3407   0.11832   0.11123  -0.0757   0.1867   1.0000
  16.500   1.3738   0.11587   0.10875  -0.0739   0.1848   1.0000
<< Back to GOE 535 AIRFOIL (goe535-il)

Polar data table (+)

Polar graphs


<< Back to GOE 535 AIRFOIL (goe535-il)