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

GOE 615 AIRFOIL (goe615-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 615 AIRFOIL (goe615-il)
Reynolds number: 500,000
Max Cl/Cd: 93.51 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe615-il-500000-n5.txt
Download as CSV file: xf-goe615-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 615 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3717   0.02653   0.02228  -0.1340   0.8596   0.0275
  -9.750  -0.3679   0.02389   0.01919  -0.1322   0.8489   0.0279
  -9.500  -0.3536   0.02238   0.01738  -0.1310   0.8402   0.0283
  -9.250  -0.3337   0.02165   0.01657  -0.1300   0.8318   0.0286
  -9.000  -0.3124   0.02104   0.01585  -0.1292   0.8240   0.0288
  -8.750  -0.2912   0.02038   0.01509  -0.1283   0.8161   0.0290
  -8.500  -0.2697   0.01972   0.01428  -0.1275   0.8079   0.0293
  -8.250  -0.2479   0.01905   0.01348  -0.1266   0.8000   0.0296
  -8.000  -0.2259   0.01836   0.01265  -0.1257   0.7916   0.0299
  -7.750  -0.2033   0.01772   0.01186  -0.1249   0.7840   0.0303
  -7.500  -0.1805   0.01706   0.01105  -0.1240   0.7759   0.0307
  -7.250  -0.1573   0.01642   0.01026  -0.1232   0.7686   0.0311
  -7.000  -0.1335   0.01588   0.00958  -0.1224   0.7604   0.0317
  -6.750  -0.1094   0.01540   0.00893  -0.1217   0.7527   0.0321
  -6.500  -0.0849   0.01493   0.00834  -0.1210   0.7444   0.0324
  -6.250  -0.0612   0.01438   0.00769  -0.1202   0.7362   0.0328
  -6.000  -0.0364   0.01397   0.00723  -0.1195   0.7274   0.0332
  -5.750  -0.0117   0.01366   0.00685  -0.1188   0.7186   0.0335
  -5.500   0.0135   0.01336   0.00650  -0.1182   0.7088   0.0340
  -5.250   0.0384   0.01309   0.00616  -0.1174   0.6988   0.0344
  -5.000   0.0634   0.01283   0.00582  -0.1167   0.6877   0.0349
  -4.750   0.0884   0.01257   0.00549  -0.1160   0.6768   0.0355
  -4.500   0.1133   0.01235   0.00517  -0.1153   0.6661   0.0360
  -4.250   0.1385   0.01215   0.00490  -0.1146   0.6544   0.0367
  -4.000   0.1637   0.01198   0.00464  -0.1139   0.6432   0.0373
  -3.750   0.1880   0.01174   0.00434  -0.1130   0.6320   0.0380
  -3.500   0.2128   0.01156   0.00412  -0.1123   0.6199   0.0386
  -3.250   0.2375   0.01144   0.00394  -0.1115   0.6066   0.0393
  -3.000   0.2620   0.01135   0.00377  -0.1107   0.5934   0.0401
  -2.750   0.2865   0.01127   0.00361  -0.1098   0.5815   0.0411
  -2.500   0.3118   0.01118   0.00346  -0.1092   0.5719   0.0421
  -2.250   0.3370   0.01113   0.00334  -0.1085   0.5638   0.0430
  -2.000   0.3624   0.01100   0.00319  -0.1079   0.5561   0.0444
  -1.500   0.4136   0.01089   0.00302  -0.1067   0.5433   0.0475
  -1.250   0.4392   0.01085   0.00294  -0.1061   0.5364   0.0491
  -1.000   0.4643   0.01081   0.00285  -0.1054   0.5303   0.0510
  -0.750   0.4905   0.01074   0.00279  -0.1049   0.5247   0.0535
  -0.500   0.5163   0.01072   0.00274  -0.1044   0.5192   0.0563
  -0.250   0.5416   0.01071   0.00271  -0.1038   0.5144   0.0605
   0.250   0.5933   0.01056   0.00268  -0.1028   0.5055   0.0942
   0.500   0.6175   0.01034   0.00271  -0.1021   0.5009   0.1808
   1.000   0.6673   0.01019   0.00282  -0.1008   0.4914   0.2784
   1.250   0.6920   0.01009   0.00288  -0.1002   0.4861   0.3389
   1.500   0.7153   0.00996   0.00296  -0.0993   0.4808   0.4243
   1.750   0.7369   0.00971   0.00306  -0.0981   0.4756   0.5542
   2.000   0.7548   0.00932   0.00316  -0.0960   0.4692   0.7216
   2.500   0.8687   0.00933   0.00351  -0.1084   0.4369   1.0000
   2.750   0.8893   0.00951   0.00358  -0.1069   0.4191   1.0000
   3.000   0.9082   0.00975   0.00369  -0.1051   0.3971   1.0000
   3.250   0.9258   0.01005   0.00384  -0.1031   0.3724   1.0000
   3.500   0.9428   0.01039   0.00403  -0.1010   0.3506   1.0000
   4.000   0.9784   0.01104   0.00446  -0.0971   0.3187   1.0000
   4.250   0.9974   0.01132   0.00467  -0.0954   0.3081   1.0000
   4.500   1.0156   0.01162   0.00489  -0.0936   0.2980   1.0000
   4.750   1.0347   0.01186   0.00509  -0.0919   0.2891   1.0000
   5.000   1.0511   0.01215   0.00532  -0.0897   0.2802   1.0000
   5.250   1.0693   0.01239   0.00553  -0.0879   0.2721   1.0000
   5.750   1.1050   0.01293   0.00600  -0.0842   0.2584   1.0000
   6.000   1.1229   0.01322   0.00626  -0.0825   0.2515   1.0000
   6.250   1.1404   0.01354   0.00654  -0.0807   0.2436   1.0000
   6.500   1.1575   0.01388   0.00684  -0.0789   0.2350   1.0000
   6.750   1.1750   0.01422   0.00715  -0.0772   0.2279   1.0000
   7.000   1.1919   0.01459   0.00749  -0.0754   0.2197   1.0000
   7.250   1.2088   0.01497   0.00785  -0.0737   0.2132   1.0000
   7.500   1.2256   0.01537   0.00823  -0.0721   0.2056   1.0000
   7.750   1.2418   0.01581   0.00864  -0.0704   0.1993   1.0000
   8.000   1.2579   0.01627   0.00908  -0.0687   0.1914   1.0000
   8.250   1.2726   0.01681   0.00959  -0.0669   0.1830   1.0000
   8.500   1.2866   0.01740   0.01015  -0.0651   0.1740   1.0000
   8.750   1.2996   0.01807   0.01076  -0.0632   0.1629   1.0000
   9.000   1.3127   0.01876   0.01143  -0.0615   0.1543   1.0000
   9.250   1.3232   0.01962   0.01223  -0.0595   0.1419   1.0000
   9.500   1.3302   0.02072   0.01323  -0.0572   0.1239   1.0000
   9.750   1.3200   0.02296   0.01519  -0.0533   0.0804   1.0000
  10.000   1.3225   0.02452   0.01667  -0.0510   0.0660   1.0000
  10.250   1.3332   0.02558   0.01775  -0.0496   0.0633   1.0000
  10.500   1.3435   0.02670   0.01890  -0.0482   0.0611   1.0000
  10.750   1.3534   0.02789   0.02012  -0.0469   0.0588   1.0000
  11.000   1.3640   0.02906   0.02132  -0.0457   0.0571   1.0000
  11.250   1.3747   0.03024   0.02256  -0.0447   0.0560   1.0000
  11.500   1.3845   0.03153   0.02390  -0.0436   0.0550   1.0000
  11.750   1.3937   0.03289   0.02531  -0.0426   0.0540   1.0000
  12.000   1.4018   0.03440   0.02687  -0.0416   0.0531   1.0000
  12.250   1.4096   0.03596   0.02848  -0.0407   0.0524   1.0000
  12.500   1.4154   0.03774   0.03031  -0.0397   0.0512   1.0000
  12.750   1.4209   0.03961   0.03224  -0.0389   0.0504   1.0000
  13.000   1.4260   0.04157   0.03427  -0.0382   0.0497   1.0000
  13.250   1.4311   0.04356   0.03632  -0.0376   0.0492   1.0000
  13.500   1.4387   0.04535   0.03819  -0.0371   0.0487   1.0000
  13.750   1.4457   0.04724   0.04015  -0.0367   0.0483   1.0000
  14.000   1.4517   0.04928   0.04226  -0.0363   0.0476   1.0000
  14.250   1.4572   0.05140   0.04446  -0.0360   0.0469   1.0000
  14.500   1.4616   0.05366   0.04679  -0.0358   0.0462   1.0000
  14.750   1.4644   0.05614   0.04934  -0.0356   0.0451   1.0000
  15.000   1.4664   0.05878   0.05204  -0.0355   0.0445   1.0000
  15.250   1.4669   0.06162   0.05496  -0.0355   0.0436   1.0000
  15.500   1.4657   0.06468   0.05808  -0.0355   0.0426   1.0000
  15.750   1.4665   0.06753   0.06101  -0.0356   0.0419   1.0000
  16.000   1.4703   0.07004   0.06360  -0.0357   0.0408   1.0000
  16.250   1.4717   0.07287   0.06651  -0.0359   0.0397   1.0000
  16.500   1.4717   0.07591   0.06960  -0.0362   0.0375   1.0000
  16.750   1.4689   0.07934   0.07309  -0.0366   0.0366   1.0000
  17.000   1.4681   0.08254   0.07637  -0.0370   0.0340   1.0000
  17.250   1.4578   0.08707   0.08086  -0.0378   0.0200   1.0000
  17.500   1.4357   0.09339   0.08720  -0.0391   0.0157   1.0000
  17.750   1.4210   0.09881   0.09272  -0.0404   0.0147   1.0000
<< Back to GOE 615 AIRFOIL (goe615-il)

Polar data table (+)

Polar graphs


<< Back to GOE 615 AIRFOIL (goe615-il)