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

GOE 573 AIRFOIL (goe573-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 573 AIRFOIL (goe573-il)
Reynolds number: 500,000
Max Cl/Cd: 100.04 at α=7.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe573-il-500000-n5.txt
Download as CSV file: xf-goe573-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 573 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
  ------ -------- --------- --------- -------- -------- --------
 -10.500   0.0248   0.08797   0.08401  -0.1142   0.6673   0.0122
 -10.250   0.0283   0.08489   0.08088  -0.1154   0.6547   0.0125
 -10.000   0.0181   0.07906   0.07505  -0.1177   0.6487   0.0137
  -9.750   0.0216   0.07634   0.07230  -0.1188   0.6387   0.0139
  -9.500   0.0282   0.07430   0.07022  -0.1196   0.6290   0.0141
  -9.250   0.0374   0.07277   0.06866  -0.1201   0.6204   0.0145
  -9.000   0.0384   0.06989   0.06576  -0.1211   0.6130   0.0146
  -8.750   0.0396   0.06724   0.06311  -0.1220   0.6065   0.0150
  -8.500   0.0343   0.06395   0.05983  -0.1233   0.6002   0.0152
  -8.000   0.0099   0.05737   0.05326  -0.1235   0.5897   0.0160
  -7.750  -0.0209   0.04935   0.04516  -0.1247   0.5873   0.0173
  -7.500  -0.0159   0.04698   0.04271  -0.1241   0.5809   0.0175
  -7.250  -0.0069   0.04507   0.04072  -0.1234   0.5747   0.0178
  -6.750  -0.0434   0.02918   0.02402  -0.1157   0.5687   0.0212
  -6.500  -0.0280   0.02836   0.02307  -0.1140   0.5618   0.0214
  -6.250  -0.0102   0.02764   0.02224  -0.1127   0.5546   0.0217
  -6.000   0.0077   0.02701   0.02148  -0.1112   0.5469   0.0221
  -5.750   0.0231   0.02565   0.01995  -0.1094   0.5403   0.0227
  -5.500   0.0358   0.02365   0.01768  -0.1068   0.5338   0.0236
  -5.250   0.0424   0.02023   0.01368  -0.1029   0.5291   0.0253
  -5.000   0.0597   0.01885   0.01195  -0.1008   0.5230   0.0259
  -4.750   0.0776   0.01769   0.01055  -0.0990   0.5166   0.0266
  -4.500   0.0995   0.01719   0.00995  -0.0979   0.5107   0.0270
  -4.250   0.1217   0.01678   0.00943  -0.0969   0.5043   0.0275
  -4.000   0.1438   0.01625   0.00876  -0.0957   0.4986   0.0279
  -3.750   0.1669   0.01575   0.00813  -0.0947   0.4930   0.0284
  -3.500   0.1899   0.01527   0.00751  -0.0937   0.4873   0.0288
  -3.250   0.2131   0.01484   0.00695  -0.0928   0.4822   0.0294
  -3.000   0.2371   0.01447   0.00648  -0.0919   0.4765   0.0301
  -2.750   0.2608   0.01418   0.00609  -0.0910   0.4715   0.0308
  -2.500   0.2845   0.01385   0.00566  -0.0902   0.4671   0.0312
  -2.250   0.3089   0.01358   0.00531  -0.0894   0.4632   0.0316
  -2.000   0.3328   0.01332   0.00499  -0.0886   0.4591   0.0319
  -1.750   0.3563   0.01312   0.00471  -0.0877   0.4553   0.0321
  -1.500   0.3789   0.01278   0.00432  -0.0866   0.4520   0.0327
  -1.250   0.4021   0.01250   0.00403  -0.0856   0.4490   0.0334
  -1.000   0.4254   0.01234   0.00384  -0.0847   0.4458   0.0341
  -0.750   0.4487   0.01224   0.00373  -0.0838   0.4427   0.0350
  -0.500   0.4717   0.01216   0.00362  -0.0828   0.4399   0.0358
   0.000   0.5178   0.01200   0.00340  -0.0808   0.4346   0.0375
   0.250   0.5413   0.01193   0.00332  -0.0800   0.4318   0.0385
   0.500   0.5647   0.01189   0.00326  -0.0790   0.4291   0.0394
   0.750   0.5878   0.01186   0.00321  -0.0781   0.4268   0.0405
   1.000   0.6104   0.01183   0.00316  -0.0770   0.4245   0.0428
   1.250   0.6331   0.01184   0.00315  -0.0760   0.4223   0.0454
   1.500   0.6563   0.01186   0.00316  -0.0751   0.4204   0.0495
   1.750   0.6796   0.01183   0.00318  -0.0743   0.4187   0.0589
   2.000   0.7019   0.01175   0.00324  -0.0732   0.4168   0.1071
   2.250   0.7221   0.01158   0.00331  -0.0718   0.4147   0.2061
   2.500   0.7131   0.01032   0.00343  -0.0643   0.4133   0.7294
   2.750   1.0007   0.01117   0.00453  -0.1221   0.4071   1.0000
   3.000   1.0222   0.01132   0.00464  -0.1210   0.4055   1.0000
   3.250   1.0445   0.01140   0.00474  -0.1200   0.4038   1.0000
   3.500   1.0666   0.01150   0.00484  -0.1189   0.4018   1.0000
   3.750   1.0882   0.01160   0.00495  -0.1178   0.3992   1.0000
   4.000   1.1092   0.01172   0.00506  -0.1166   0.3967   1.0000
   4.250   1.1296   0.01185   0.00517  -0.1153   0.3941   1.0000
   4.500   1.1490   0.01202   0.00530  -0.1137   0.3910   1.0000
   4.750   1.1695   0.01215   0.00545  -0.1124   0.3889   1.0000
   5.000   1.1903   0.01225   0.00558  -0.1112   0.3863   1.0000
   5.250   1.2107   0.01237   0.00572  -0.1099   0.3844   1.0000
   5.500   1.2306   0.01251   0.00589  -0.1085   0.3826   1.0000
   5.750   1.2500   0.01265   0.00605  -0.1070   0.3808   1.0000
   6.000   1.2687   0.01280   0.00621  -0.1053   0.3789   1.0000
   6.250   1.2869   0.01297   0.00640  -0.1036   0.3771   1.0000
   6.500   1.3041   0.01315   0.00658  -0.1017   0.3754   1.0000
   6.750   1.3209   0.01329   0.00677  -0.0997   0.3738   1.0000
   7.000   1.3373   0.01340   0.00694  -0.0976   0.3717   1.0000
   7.250   1.3506   0.01350   0.00708  -0.0949   0.3667   1.0000
   7.500   1.3634   0.01366   0.00725  -0.0921   0.3629   1.0000
   7.750   1.3758   0.01384   0.00744  -0.0893   0.3586   1.0000
   8.000   1.3911   0.01397   0.00763  -0.0871   0.3537   1.0000
   8.250   1.4041   0.01417   0.00786  -0.0845   0.3487   1.0000
   8.500   1.4167   0.01440   0.00810  -0.0818   0.3434   1.0000
   8.750   1.4308   0.01461   0.00835  -0.0795   0.3363   1.0000
   9.000   1.4418   0.01491   0.00866  -0.0767   0.3296   1.0000
   9.250   1.4530   0.01524   0.00900  -0.0740   0.3188   1.0000
   9.500   1.4593   0.01574   0.00947  -0.0705   0.3031   1.0000
   9.750   1.4623   0.01643   0.01009  -0.0667   0.2841   1.0000
  10.000   1.4602   0.01741   0.01097  -0.0623   0.2630   1.0000
  10.250   1.4537   0.01871   0.01216  -0.0576   0.2407   1.0000
  10.500   1.4446   0.02033   0.01367  -0.0530   0.2189   1.0000
  10.750   1.4370   0.02207   0.01535  -0.0491   0.2006   1.0000
  11.000   1.4199   0.02463   0.01779  -0.0448   0.1715   1.0000
  11.250   1.3928   0.02829   0.02128  -0.0405   0.1340   1.0000
  11.500   1.3633   0.03267   0.02555  -0.0371   0.1033   1.0000
  11.750   1.3325   0.03765   0.03046  -0.0345   0.0723   1.0000
  12.000   1.3127   0.04191   0.03470  -0.0329   0.0533   1.0000
  12.250   1.2829   0.04740   0.04015  -0.0313   0.0282   1.0000
  12.500   1.2725   0.05105   0.04384  -0.0305   0.0213   1.0000
  12.750   1.2664   0.05435   0.04719  -0.0299   0.0183   1.0000
  13.000   1.2642   0.05726   0.05018  -0.0294   0.0174   1.0000
  13.250   1.2592   0.06054   0.05354  -0.0290   0.0160   1.0000
  13.500   1.2546   0.06387   0.05695  -0.0287   0.0150   1.0000
  13.750   1.2518   0.06702   0.06018  -0.0285   0.0142   1.0000
  14.000   1.2493   0.07017   0.06342  -0.0283   0.0138   1.0000
  14.250   1.2476   0.07325   0.06658  -0.0282   0.0131   1.0000
  14.500   1.2445   0.07655   0.06995  -0.0282   0.0125   1.0000
  14.750   1.2405   0.08002   0.07350  -0.0282   0.0120   1.0000
  15.000   1.2363   0.08352   0.07706  -0.0283   0.0115   1.0000
  15.250   1.2301   0.08734   0.08096  -0.0285   0.0110   1.0000
  15.500   1.2278   0.09068   0.08439  -0.0287   0.0108   1.0000
<< Back to GOE 573 AIRFOIL (goe573-il)

Polar data table (+)

Polar graphs


<< Back to GOE 573 AIRFOIL (goe573-il)