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=50,000 Ncrit=5


Details Polar file
Airfoil: GOE 535 AIRFOIL (goe535-il)
Reynolds number: 50,000
Max Cl/Cd: 29.44 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe535-il-50000-n5.txt
Download as CSV file: xf-goe535-il-50000-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 =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750  -0.2392   0.11896   0.11183  -0.0372   1.0000   0.1065
  -9.500  -0.2355   0.11707   0.11000  -0.0350   1.0000   0.1045
  -9.250  -0.2422   0.11538   0.10842  -0.0332   1.0000   0.1037
  -9.000  -0.2442   0.11304   0.10616  -0.0333   0.9981   0.1029
  -8.750  -0.2257   0.10838   0.10148  -0.0390   0.9891   0.1024
  -8.500  -0.2083   0.10362   0.09670  -0.0448   0.9800   0.1011
  -8.250  -0.1964   0.09865   0.09173  -0.0504   0.9687   0.0993
  -8.000  -0.1881   0.09364   0.08670  -0.0559   0.9563   0.0984
  -7.750  -0.1755   0.08949   0.08254  -0.0604   0.9439   0.0991
  -7.500  -0.1631   0.08515   0.07820  -0.0653   0.9323   0.1001
  -7.250  -0.1584   0.08072   0.07376  -0.0696   0.9178   0.1010
  -7.000  -0.1567   0.07526   0.06829  -0.0750   0.9021   0.1012
  -6.750  -0.1589   0.06768   0.06061  -0.0835   0.8866   0.1014
  -6.500  -0.1657   0.05582   0.04830  -0.0997   0.8711   0.1029
  -6.250  -0.1482   0.04674   0.03824  -0.1125   0.8600   0.1060
  -6.000  -0.1252   0.04498   0.03645  -0.1131   0.8473   0.1081
  -5.750  -0.0895   0.04266   0.03394  -0.1162   0.8395   0.1118
  -5.500  -0.0664   0.04005   0.03087  -0.1181   0.8265   0.1155
  -5.250  -0.0296   0.03777   0.02827  -0.1209   0.8187   0.1194
  -5.000  -0.0056   0.03668   0.02711  -0.1207   0.8057   0.1231
  -4.750   0.0322   0.03481   0.02482  -0.1230   0.7979   0.1282
  -4.500   0.0563   0.03380   0.02370  -0.1227   0.7849   0.1321
  -4.250   0.0934   0.03259   0.02237  -0.1240   0.7769   0.1373
  -4.000   0.1181   0.03175   0.02127  -0.1236   0.7636   0.1423
  -3.750   0.1541   0.03073   0.02023  -0.1245   0.7555   0.1479
  -3.500   0.1774   0.03021   0.01962  -0.1237   0.7422   0.1532
  -3.250   0.2136   0.02930   0.01855  -0.1246   0.7340   0.1604
  -3.000   0.2358   0.02895   0.01820  -0.1235   0.7206   0.1665
  -2.750   0.2718   0.02820   0.01729  -0.1244   0.7124   0.1758
  -2.500   0.2931   0.02796   0.01709  -0.1233   0.6992   0.1848
  -2.250   0.3275   0.02730   0.01638  -0.1239   0.6909   0.1984
  -2.000   0.3498   0.02711   0.01626  -0.1231   0.6780   0.2149
  -1.750   0.3842   0.02653   0.01580  -0.1239   0.6698   0.2488
  -1.500   0.4048   0.02665   0.01636  -0.1225   0.6577   0.2972
  -1.250   0.4350   0.02683   0.01656  -0.1220   0.6492   0.3672
  -1.000   0.4567   0.02733   0.01700  -0.1206   0.6381   0.4088
  -0.750   0.4834   0.02767   0.01728  -0.1195   0.6294   0.4392
  -0.500   0.5067   0.02811   0.01765  -0.1182   0.6201   0.4638
  -0.250   0.5294   0.02855   0.01807  -0.1167   0.6109   0.4909
   0.000   0.5565   0.02887   0.01832  -0.1154   0.6040   0.5242
   0.250   0.5719   0.02942   0.01895  -0.1130   0.5938   0.5482
   0.500   0.5991   0.02951   0.01894  -0.1122   0.5866   0.5696
   0.750   0.6229   0.02976   0.01910  -0.1114   0.5786   0.5842
   1.000   0.6469   0.03003   0.01926  -0.1108   0.5701   0.5977
   1.250   0.6782   0.02994   0.01904  -0.1109   0.5640   0.6087
   1.500   0.6986   0.03044   0.01949  -0.1100   0.5554   0.6180
   1.750   0.7245   0.03063   0.01961  -0.1096   0.5481   0.6266
   2.000   0.7590   0.03061   0.01939  -0.1104   0.5426   0.6362
   2.250   0.7747   0.03127   0.02012  -0.1090   0.5339   0.6434
   2.500   0.8021   0.03155   0.02031  -0.1089   0.5272   0.6529
   2.750   0.8363   0.03149   0.02013  -0.1096   0.5221   0.6624
   3.000   0.8497   0.03238   0.02112  -0.1080   0.5135   0.6710
   3.250   0.8760   0.03268   0.02139  -0.1077   0.5070   0.6806
   3.500   0.9104   0.03266   0.02127  -0.1084   0.5021   0.6934
   3.750   0.9208   0.03367   0.02246  -0.1064   0.4938   0.7042
   4.000   0.9453   0.03404   0.02286  -0.1060   0.4872   0.7189
   4.250   0.9791   0.03396   0.02276  -0.1065   0.4824   0.7380
   4.500   0.9879   0.03504   0.02409  -0.1044   0.4745   0.7613
   4.750   1.0100   0.03535   0.02462  -0.1037   0.4677   0.8306
   5.000   1.0415   0.03538   0.02448  -0.1039   0.4627   1.0000
   5.250   1.0486   0.03689   0.02604  -0.1019   0.4550   1.0000
   5.500   1.0661   0.03783   0.02694  -0.1008   0.4480   1.0000
   5.750   1.1018   0.03789   0.02681  -0.1016   0.4429   1.0000
   6.000   1.1040   0.03958   0.02857  -0.0990   0.4355   1.0000
   6.250   1.1132   0.04087   0.02987  -0.0971   0.4284   1.0000
   6.500   1.1486   0.04086   0.02970  -0.0976   0.4233   1.0000
   6.750   1.1462   0.04272   0.03163  -0.0945   0.4164   1.0000
   7.000   1.1375   0.04485   0.03382  -0.0909   0.4088   1.0000
   7.250   1.1729   0.04475   0.03362  -0.0913   0.4040   1.0000
   7.500   1.1702   0.04676   0.03567  -0.0885   0.3977   1.0000
   7.750   1.1331   0.05122   0.04030  -0.0842   0.3885   1.0000
   8.000   1.1712   0.05062   0.03960  -0.0842   0.3845   1.0000
   8.250   1.0841   0.06067   0.04991  -0.0807   0.3705   1.0000
   8.500   1.1103   0.06074   0.04994  -0.0799   0.3665   1.0000
   8.750   1.1540   0.05912   0.04825  -0.0793   0.3641   1.0000
   9.250   1.0745   0.07414   0.06349  -0.0791   0.3423   1.0000
   9.750   1.0613   0.08281   0.07226  -0.0797   0.3271   1.0000
  10.750   1.0023   0.10676   0.09642  -0.0846   0.2947   1.0000
  11.000   1.0252   0.10708   0.09674  -0.0837   0.2925   1.0000
  11.500   0.9925   0.12027   0.11005  -0.0877   0.2786   1.0000
  11.750   1.0117   0.12123   0.11102  -0.0871   0.2762   1.0000
<< Back to GOE 535 AIRFOIL (goe535-il)

Polar data table (+)

Polar graphs


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