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

GOE 514 AIRFOIL (goe514-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 514 AIRFOIL (goe514-il)
Reynolds number: 100,000
Max Cl/Cd: 45.3 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe514-il-100000.txt
Download as CSV file: xf-goe514-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 514 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
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.3012   0.11193   0.10796  -0.0466   0.9607   0.1425
  -8.000  -0.3306   0.10903   0.10510  -0.0505   0.9496   0.1430
  -7.750  -0.4379   0.08106   0.07695  -0.0572   0.9428   0.0806
  -7.500  -0.4560   0.07423   0.06998  -0.0569   0.9326   0.0791
  -7.250  -0.4798   0.06219   0.05748  -0.0587   0.9262   0.0780
  -7.000  -0.5105   0.05349   0.04807  -0.0546   0.9151   0.0774
  -6.750  -0.5151   0.04643   0.03970  -0.0520   0.9076   0.0799
  -6.500  -0.4923   0.04469   0.03806  -0.0515   0.8967   0.0836
  -6.250  -0.4724   0.04268   0.03566  -0.0504   0.8875   0.0889
  -6.000  -0.4446   0.04158   0.03450  -0.0504   0.8797   0.0961
  -5.750  -0.4252   0.03990   0.03238  -0.0489   0.8719   0.1042
  -5.500  -0.3970   0.03981   0.03227  -0.0488   0.8637   0.1145
  -5.250  -0.3674   0.04021   0.03287  -0.0489   0.8556   0.1238
  -5.000  -0.3394   0.04012   0.03275  -0.0487   0.8472   0.1349
  -4.750  -0.3129   0.03973   0.03220  -0.0482   0.8396   0.1466
  -4.500  -0.2871   0.03927   0.03146  -0.0475   0.8310   0.1587
  -4.250  -0.2562   0.03957   0.03205  -0.0478   0.8231   0.1676
  -4.000  -0.2291   0.03854   0.03077  -0.0473   0.8148   0.1786
  -3.750  -0.2010   0.03809   0.03020  -0.0468   0.8066   0.1894
  -3.500  -0.1688   0.03769   0.02984  -0.0471   0.7985   0.1993
  -3.250  -0.1431   0.03685   0.02878  -0.0462   0.7896   0.2111
  -3.000  -0.1090   0.03616   0.02793  -0.0465   0.7823   0.2260
  -2.750  -0.0814   0.03599   0.02785  -0.0459   0.7728   0.2361
  -2.500  -0.0445   0.03518   0.02695  -0.0466   0.7660   0.2500
  -2.250  -0.0168   0.03451   0.02610  -0.0459   0.7569   0.2626
  -2.000   0.0213   0.03360   0.02509  -0.0467   0.7501   0.2744
  -1.750   0.0704   0.03211   0.02352  -0.0492   0.7472   0.2860
  -1.500   0.0909   0.03167   0.02283  -0.0473   0.7350   0.2947
  -1.250   0.1411   0.03008   0.02127  -0.0501   0.7318   0.3061
  -1.000   0.1710   0.02929   0.02037  -0.0497   0.7215   0.3162
  -0.750   0.2234   0.02780   0.01871  -0.0530   0.7162   0.3274
  -0.500   0.2954   0.02580   0.01660  -0.0600   0.7127   0.3367
  -0.250   0.3276   0.02516   0.01579  -0.0602   0.6986   0.3443
   0.000   0.3731   0.02423   0.01477  -0.0629   0.6855   0.3520
   0.250   0.4282   0.02340   0.01372  -0.0675   0.6720   0.3630
   0.500   0.4812   0.02266   0.01289  -0.0721   0.6569   0.3754
   0.750   0.5250   0.02218   0.01230  -0.0749   0.6411   0.3891
   1.000   0.5649   0.02184   0.01188  -0.0771   0.6254   0.4051
   1.250   0.5965   0.02168   0.01170  -0.0779   0.6096   0.4216
   1.500   0.6302   0.02159   0.01160  -0.0791   0.5946   0.4397
   1.750   0.6613   0.02153   0.01152  -0.0799   0.5812   0.4587
   2.000   0.6942   0.02135   0.01134  -0.0810   0.5703   0.4831
   2.250   0.7118   0.02109   0.01131  -0.0791   0.5599   0.5243
   2.500   0.9828   0.02214   0.01298  -0.1279   0.5335   1.0000
   2.750   1.0042   0.02247   0.01315  -0.1268   0.5253   1.0000
   3.000   1.0232   0.02288   0.01346  -0.1253   0.5180   1.0000
   3.250   1.0403   0.02329   0.01380  -0.1235   0.5108   1.0000
   3.500   1.0708   0.02364   0.01393  -0.1242   0.5047   1.0000
   3.750   1.0770   0.02417   0.01453  -0.1203   0.4980   1.0000
   4.000   1.0983   0.02456   0.01482  -0.1193   0.4919   1.0000
   4.250   1.1229   0.02501   0.01515  -0.1189   0.4863   1.0000
   4.500   1.1299   0.02559   0.01579  -0.1153   0.4807   1.0000
   4.750   1.1495   0.02607   0.01623  -0.1140   0.4757   1.0000
   5.000   1.1840   0.02653   0.01653  -0.1155   0.4712   1.0000
   5.250   1.1852   0.02724   0.01737  -0.1109   0.4665   1.0000
   5.500   1.1948   0.02786   0.01805  -0.1078   0.4616   1.0000
   5.750   1.2175   0.02837   0.01849  -0.1071   0.4571   1.0000
   6.000   1.2476   0.02899   0.01901  -0.1079   0.4530   1.0000
   6.250   1.2438   0.02985   0.02003  -0.1024   0.4491   1.0000
   6.500   1.2494   0.03064   0.02092  -0.0988   0.4452   1.0000
   6.750   1.2658   0.03133   0.02162  -0.0970   0.4413   1.0000
   7.000   1.2973   0.03195   0.02215  -0.0981   0.4377   1.0000
   7.250   1.3004   0.03294   0.02324  -0.0941   0.4341   1.0000
   7.500   1.2904   0.03403   0.02450  -0.0878   0.4304   1.0000
   7.750   1.2935   0.03498   0.02553  -0.0839   0.4266   1.0000
   8.000   1.3097   0.03580   0.02637  -0.0823   0.4232   1.0000
   8.250   1.3443   0.03649   0.02702  -0.0840   0.4200   1.0000
   8.500   1.3337   0.03791   0.02858  -0.0779   0.4170   1.0000
   8.750   1.2918   0.03964   0.03049  -0.0666   0.4139   1.0000
   9.000   1.2602   0.04147   0.03247  -0.0577   0.4106   1.0000
   9.250   1.2608   0.04279   0.03386  -0.0541   0.4071   1.0000
   9.500   1.2943   0.04333   0.03439  -0.0554   0.4040   1.0000
   9.750   1.3220   0.04433   0.03538  -0.0560   0.4007   1.0000
  10.000   0.7202   0.10319   0.09506  -0.0261   0.3835   1.0000
  10.250   0.7036   0.10934   0.10127  -0.0267   0.3835   1.0000
<< Back to GOE 514 AIRFOIL (goe514-il)

Polar data table (+)

Polar graphs


<< Back to GOE 514 AIRFOIL (goe514-il)