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

GOE 15 AIRFOIL (goe15-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 15 AIRFOIL (goe15-il)
Reynolds number: 500,000
Max Cl/Cd: 94.27 at α=8°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe15-il-500000-n5.txt
Download as CSV file: xf-goe15-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 15 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.750   0.0260   0.09337   0.09042  -0.0908   0.8469   0.0197
  -9.500   0.0332   0.09096   0.08785  -0.0918   0.8117   0.0210
  -9.250   0.0382   0.08884   0.08558  -0.0921   0.7814   0.0210
  -9.000   0.0431   0.08678   0.08340  -0.0923   0.7598   0.0211
  -8.750   0.0483   0.08460   0.08115  -0.0927   0.7463   0.0211
  -8.500   0.0557   0.08251   0.07899  -0.0927   0.7352   0.0213
  -8.250   0.0633   0.08059   0.07704  -0.0928   0.7260   0.0214
  -8.000   0.0712   0.07890   0.07531  -0.0929   0.7177   0.0219
  -7.750   0.0785   0.07719   0.07359  -0.0930   0.7103   0.0227
  -7.500   0.0822   0.07528   0.07166  -0.0931   0.7041   0.0239
  -7.250   0.0840   0.07344   0.06983  -0.0931   0.6990   0.0240
  -7.000   0.0909   0.07132   0.06769  -0.0941   0.6935   0.0241
  -6.750   0.0996   0.06906   0.06541  -0.0957   0.6878   0.0241
  -6.500   0.1104   0.06672   0.06303  -0.0975   0.6828   0.0242
  -6.250   0.1232   0.06435   0.06064  -0.0994   0.6767   0.0242
  -6.000   0.1354   0.06230   0.05857  -0.0992   0.6708   0.0244
  -5.750   0.1500   0.06058   0.05683  -0.0999   0.6651   0.0250
  -5.500   0.1662   0.05867   0.05491  -0.1011   0.6589   0.0257
  -5.250   0.1837   0.05656   0.05274  -0.1027   0.6525   0.0269
  -5.000   0.2037   0.05418   0.05030  -0.1050   0.6465   0.0272
  -4.750   0.2246   0.05187   0.04793  -0.1070   0.6389   0.0273
  -4.500   0.2456   0.04966   0.04563  -0.1087   0.6309   0.0273
  -4.000   0.2839   0.04552   0.04137  -0.1101   0.6096   0.0277
  -3.750   0.3040   0.04387   0.03963  -0.1107   0.5958   0.0281
  -3.500   0.3253   0.04223   0.03788  -0.1113   0.5802   0.0286
  -3.250   0.3475   0.04061   0.03613  -0.1118   0.5622   0.0294
  -3.000   0.3726   0.03888   0.03422  -0.1127   0.5418   0.0305
  -2.750   0.3975   0.03720   0.03233  -0.1132   0.5202   0.0306
  -2.500   0.4223   0.03562   0.03053  -0.1135   0.4994   0.0307
  -2.250   0.4472   0.03404   0.02874  -0.1137   0.4818   0.0308
  -2.000   0.4690   0.03246   0.02703  -0.1135   0.4685   0.0308
  -1.750   0.4879   0.03117   0.02568  -0.1130   0.4582   0.0311
  -1.500   0.5106   0.03004   0.02444  -0.1128   0.4497   0.0315
  -1.250   0.5344   0.02893   0.02322  -0.1126   0.4427   0.0319
  -1.000   0.5590   0.02784   0.02202  -0.1124   0.4363   0.0328
  -0.500   0.6128   0.02556   0.01940  -0.1115   0.4252   0.0342
  -0.250   0.6378   0.02438   0.01810  -0.1110   0.4208   0.0342
   0.000   0.6595   0.02337   0.01705  -0.1106   0.4165   0.0333
   0.250   0.6843   0.02229   0.01582  -0.1100   0.4126   0.0325
   0.500   0.7091   0.02136   0.01475  -0.1093   0.4087   0.0325
   0.750   0.7348   0.02047   0.01375  -0.1087   0.4055   0.0328
   1.000   0.7603   0.01965   0.01281  -0.1080   0.4018   0.0330
   1.250   0.7875   0.01846   0.01136  -0.1068   0.3986   0.0343
   1.500   0.8114   0.01796   0.01080  -0.1062   0.3954   0.0346
   1.750   0.8354   0.01758   0.01034  -0.1056   0.3922   0.0349
   2.000   0.8603   0.01714   0.00984  -0.1050   0.3894   0.0354
   2.250   0.8856   0.01666   0.00929  -0.1043   0.3865   0.0359
   2.500   0.9108   0.01620   0.00874  -0.1036   0.3833   0.0366
   2.750   0.9366   0.01537   0.00769  -0.1025   0.3805   0.0382
   3.000   0.9605   0.01518   0.00745  -0.1018   0.3772   0.0386
   3.250   0.9841   0.01504   0.00727  -0.1011   0.3731   0.0390
   3.500   1.0092   0.01479   0.00700  -0.1005   0.3691   0.0396
   3.750   1.0337   0.01456   0.00672  -0.0998   0.3652   0.0406
   4.000   1.0581   0.01416   0.00621  -0.0988   0.3618   0.0422
   4.250   1.0813   0.01400   0.00600  -0.0980   0.3581   0.0428
   4.500   1.1049   0.01396   0.00596  -0.0972   0.3541   0.0433
   4.750   1.1287   0.01387   0.00590  -0.0965   0.3500   0.0441
   5.000   1.1518   0.01383   0.00585  -0.0956   0.3460   0.0453
   5.250   1.1741   0.01380   0.00578  -0.0945   0.3426   0.0467
   5.500   1.1955   0.01380   0.00575  -0.0933   0.3394   0.0474
   5.750   1.2183   0.01378   0.00578  -0.0924   0.3367   0.0481
   6.000   1.2407   0.01381   0.00586  -0.0915   0.3334   0.0489
   6.250   1.2624   0.01388   0.00595  -0.0904   0.3298   0.0502
   6.500   1.2829   0.01399   0.00606  -0.0891   0.3260   0.0518
   6.750   1.3017   0.01411   0.00618  -0.0875   0.3223   0.0525
   7.000   1.3221   0.01416   0.00627  -0.0861   0.3184   0.0533
   7.250   1.3414   0.01427   0.00642  -0.0846   0.3137   0.0544
   7.500   1.3597   0.01445   0.00661  -0.0830   0.3093   0.0555
   7.750   1.3781   0.01465   0.00684  -0.0814   0.3053   0.0580
   8.000   1.3980   0.01483   0.00706  -0.0801   0.3009   0.0592
   8.250   1.4161   0.01504   0.00729  -0.0785   0.2958   0.0607
   8.500   1.4326   0.01534   0.00758  -0.0766   0.2908   0.0627
   8.750   1.4515   0.01557   0.00786  -0.0753   0.2860   0.0653
   9.000   1.4681   0.01588   0.00818  -0.0735   0.2790   0.0695
   9.250   1.4829   0.01627   0.00855  -0.0716   0.2698   0.0736
   9.500   1.4954   0.01676   0.00902  -0.0693   0.2587   0.0857
  10.000   1.5603   0.01716   0.01049  -0.0740   0.2312   1.0000
  10.250   1.5724   0.01780   0.01110  -0.0720   0.2204   1.0000
  10.500   1.5829   0.01856   0.01183  -0.0698   0.2097   1.0000
  10.750   1.5921   0.01943   0.01266  -0.0676   0.1992   1.0000
  11.000   1.6005   0.02039   0.01357  -0.0655   0.1877   1.0000
  11.250   1.6090   0.02138   0.01454  -0.0635   0.1782   1.0000
  11.750   1.6260   0.02350   0.01666  -0.0598   0.1639   1.0000
  12.000   1.6321   0.02479   0.01794  -0.0580   0.1567   1.0000
  12.250   1.6407   0.02596   0.01914  -0.0565   0.1504   1.0000
  12.500   1.6463   0.02741   0.02059  -0.0549   0.1434   1.0000
  12.750   1.6532   0.02881   0.02203  -0.0535   0.1380   1.0000
  13.000   1.6584   0.03042   0.02365  -0.0522   0.1306   1.0000
  13.250   1.6607   0.03232   0.02556  -0.0509   0.1219   1.0000
  13.500   1.6598   0.03458   0.02781  -0.0495   0.1118   1.0000
  13.750   1.6517   0.03758   0.03075  -0.0481   0.0962   1.0000
  14.000   1.6403   0.04102   0.03415  -0.0467   0.0825   1.0000
  14.250   1.6236   0.04517   0.03825  -0.0455   0.0685   1.0000
  14.500   1.6102   0.04917   0.04226  -0.0448   0.0581   1.0000
  14.750   1.5887   0.05431   0.04738  -0.0442   0.0444   1.0000
  15.000   1.5646   0.06004   0.05311  -0.0441   0.0299   1.0000
  15.250   1.5442   0.06562   0.05874  -0.0444   0.0209   1.0000
  15.500   1.5305   0.07050   0.06369  -0.0448   0.0179   1.0000
  15.750   1.5217   0.07483   0.06812  -0.0452   0.0167   1.0000
  16.000   1.5129   0.07923   0.07261  -0.0458   0.0158   1.0000
  16.500   1.4976   0.08779   0.08138  -0.0471   0.0145   1.0000
  16.750   1.4914   0.09192   0.08560  -0.0479   0.0141   1.0000
  17.000   1.4847   0.09615   0.08994  -0.0487   0.0137   1.0000
  17.250   1.4768   0.10060   0.09448  -0.0496   0.0133   1.0000
  17.500   1.4698   0.10495   0.09893  -0.0507   0.0130   1.0000
  17.750   1.4622   0.10947   0.10354  -0.0518   0.0126   1.0000
  18.000   1.4534   0.11425   0.10842  -0.0532   0.0122   1.0000
<< Back to GOE 15 AIRFOIL (goe15-il)

Polar data table (+)

Polar graphs


<< Back to GOE 15 AIRFOIL (goe15-il)