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

GOE 481A AIRFOIL (goe481a-il) Xfoil prediction polar at RE=200,000 Ncrit=5


Details Polar file
Airfoil: GOE 481A AIRFOIL (goe481a-il)
Reynolds number: 200,000
Max Cl/Cd: 59.03 at α=6°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe481a-il-200000-n5.txt
Download as CSV file: xf-goe481a-il-200000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 481A AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.1182   0.09337   0.08904  -0.0858   0.9463   0.0543
  -8.500  -0.1259   0.08944   0.08513  -0.0866   0.9404   0.0547
  -8.250  -0.1247   0.08558   0.08127  -0.0886   0.9358   0.0546
  -7.750  -0.3469   0.04212   0.03696  -0.1084   0.9040   0.0562
  -7.500  -0.3558   0.03685   0.03121  -0.1070   0.8970   0.0567
  -7.250  -0.3486   0.03434   0.02843  -0.1053   0.8904   0.0572
  -7.000  -0.3289   0.03199   0.02580  -0.1053   0.8871   0.0578
  -6.750  -0.3049   0.02986   0.02337  -0.1057   0.8849   0.0584
  -6.500  -0.3016   0.02875   0.02206  -0.1016   0.8748   0.0588
  -6.250  -0.2783   0.02725   0.02029  -0.1012   0.8710   0.0596
  -6.000  -0.2514   0.02576   0.01851  -0.1014   0.8684   0.0606
  -5.750  -0.2400   0.02495   0.01747  -0.0983   0.8593   0.0612
  -5.500  -0.2157   0.02396   0.01633  -0.0976   0.8542   0.0619
  -5.250  -0.1862   0.02307   0.01539  -0.0977   0.8511   0.0624
  -5.000  -0.1694   0.02258   0.01488  -0.0955   0.8423   0.0630
  -4.750  -0.1448   0.02190   0.01414  -0.0946   0.8358   0.0636
  -4.500  -0.1128   0.02108   0.01324  -0.0951   0.8319   0.0644
  -4.250  -0.0975   0.02067   0.01276  -0.0925   0.8201   0.0651
  -4.000  -0.0657   0.01991   0.01190  -0.0929   0.8146   0.0660
  -3.750  -0.0444   0.01944   0.01133  -0.0913   0.8032   0.0670
  -3.500  -0.0099   0.01870   0.01047  -0.0922   0.7960   0.0683
  -3.250   0.0164   0.01821   0.00999  -0.0916   0.7839   0.0693
  -3.000   0.0544   0.01756   0.00930  -0.0932   0.7740   0.0708
  -2.750   0.0914   0.01697   0.00862  -0.0946   0.7605   0.0724
  -2.500   0.1315   0.01638   0.00790  -0.0966   0.7447   0.0740
  -2.250   0.1724   0.01585   0.00728  -0.0988   0.7270   0.0757
  -2.000   0.2112   0.01548   0.00681  -0.1008   0.7073   0.0779
  -1.750   0.2449   0.01524   0.00643  -0.1016   0.6881   0.0807
  -1.500   0.2740   0.01509   0.00618  -0.1016   0.6702   0.0835
  -1.250   0.3007   0.01501   0.00602  -0.1012   0.6542   0.0869
  -1.000   0.3265   0.01495   0.00586  -0.1005   0.6399   0.0906
  -0.750   0.3492   0.01495   0.00585  -0.0993   0.6266   0.0948
  -0.500   0.3727   0.01497   0.00582  -0.0981   0.6147   0.1005
  -0.250   0.3955   0.01504   0.00583  -0.0969   0.6027   0.1083
   0.000   0.4180   0.01514   0.00589  -0.0956   0.5921   0.1154
   0.250   0.4410   0.01525   0.00595  -0.0944   0.5824   0.1226
   0.500   0.4635   0.01535   0.00597  -0.0931   0.5733   0.1299
   0.750   0.4854   0.01545   0.00603  -0.0917   0.5643   0.1351
   1.000   0.5079   0.01558   0.00606  -0.0904   0.5556   0.1415
   1.250   0.5288   0.01562   0.00613  -0.0889   0.5470   0.1462
   1.500   0.5510   0.01574   0.00616  -0.0876   0.5390   0.1516
   1.750   0.5719   0.01579   0.00620  -0.0861   0.5314   0.1559
   2.000   0.5929   0.01583   0.00626  -0.0846   0.5239   0.1606
   2.250   0.6148   0.01595   0.00632  -0.0833   0.5170   0.1666
   2.500   0.6345   0.01598   0.00640  -0.0815   0.5091   0.1730
   2.750   0.6538   0.01607   0.00647  -0.0797   0.5005   0.1792
   3.000   0.6729   0.01618   0.00655  -0.0779   0.4917   0.1841
   3.250   0.6913   0.01624   0.00662  -0.0759   0.4830   0.1879
   3.500   0.7104   0.01634   0.00671  -0.0741   0.4751   0.1917
   3.750   0.7296   0.01643   0.00682  -0.0723   0.4674   0.1960
   4.000   0.7492   0.01657   0.00693  -0.0706   0.4607   0.2009
   4.250   0.7684   0.01668   0.00707  -0.0688   0.4535   0.2070
   4.500   0.7865   0.01681   0.00722  -0.0669   0.4449   0.2148
   4.750   0.8052   0.01696   0.00738  -0.0651   0.4381   0.2284
   5.000   0.8246   0.01703   0.00759  -0.0635   0.4306   0.2639
   5.500   1.0031   0.01701   0.00901  -0.0909   0.4056   1.0000
   5.750   1.0204   0.01732   0.00924  -0.0889   0.3979   1.0000
   6.000   1.0390   0.01760   0.00950  -0.0872   0.3903   1.0000
   6.250   1.0562   0.01793   0.00975  -0.0852   0.3835   1.0000
   6.500   1.0750   0.01823   0.01004  -0.0836   0.3773   1.0000
   6.750   1.0932   0.01855   0.01034  -0.0818   0.3707   1.0000
   7.000   1.1102   0.01893   0.01065  -0.0799   0.3647   1.0000
   7.250   1.1288   0.01926   0.01098  -0.0784   0.3588   1.0000
   7.500   1.1466   0.01962   0.01134  -0.0767   0.3526   1.0000
   7.750   1.1628   0.02004   0.01172  -0.0747   0.3466   1.0000
   8.000   1.1804   0.02043   0.01211  -0.0731   0.3410   1.0000
   8.250   1.1976   0.02084   0.01253  -0.0713   0.3347   1.0000
   8.500   1.2124   0.02133   0.01298  -0.0693   0.3283   1.0000
   8.750   1.2283   0.02179   0.01345  -0.0675   0.3217   1.0000
   9.000   1.2426   0.02231   0.01396  -0.0654   0.3136   1.0000
   9.250   1.2552   0.02291   0.01453  -0.0632   0.3064   1.0000
   9.500   1.2699   0.02344   0.01509  -0.0613   0.2983   1.0000
   9.750   1.2803   0.02416   0.01575  -0.0589   0.2900   1.0000
  10.000   1.2953   0.02473   0.01637  -0.0572   0.2828   1.0000
  10.250   1.3061   0.02548   0.01711  -0.0550   0.2743   1.0000
  10.500   1.3186   0.02619   0.01784  -0.0530   0.2667   1.0000
  10.750   1.3289   0.02702   0.01867  -0.0509   0.2583   1.0000
  11.000   1.3395   0.02787   0.01953  -0.0489   0.2502   1.0000
  11.250   1.3486   0.02882   0.02048  -0.0468   0.2418   1.0000
  11.500   1.3572   0.02984   0.02151  -0.0447   0.2333   1.0000
  11.750   1.3641   0.03100   0.02266  -0.0426   0.2250   1.0000
  12.000   1.3710   0.03221   0.02387  -0.0405   0.2163   1.0000
  12.250   1.3764   0.03356   0.02522  -0.0385   0.2094   1.0000
  12.500   1.3831   0.03489   0.02657  -0.0367   0.2032   1.0000
  12.750   1.3881   0.03637   0.02806  -0.0349   0.1972   1.0000
  13.000   1.3912   0.03806   0.02975  -0.0330   0.1923   1.0000
  13.250   1.3977   0.03955   0.03129  -0.0316   0.1874   1.0000
  13.500   1.4023   0.04123   0.03300  -0.0301   0.1832   1.0000
  13.750   1.4052   0.04311   0.03491  -0.0287   0.1796   1.0000
  14.000   1.4088   0.04498   0.03682  -0.0274   0.1763   1.0000
  14.250   1.4149   0.04670   0.03861  -0.0264   0.1729   1.0000
  14.500   1.4181   0.04872   0.04068  -0.0253   0.1692   1.0000
  14.750   1.4200   0.05090   0.04289  -0.0243   0.1661   1.0000
  15.000   1.4207   0.05322   0.04523  -0.0234   0.1632   1.0000
  15.250   1.4259   0.05519   0.04730  -0.0227   0.1601   1.0000
  15.500   1.4294   0.05736   0.04956  -0.0221   0.1569   1.0000
  15.750   1.4311   0.05974   0.05200  -0.0215   0.1538   1.0000
  16.000   1.4313   0.06230   0.05459  -0.0210   0.1509   1.0000
  16.250   1.4314   0.06488   0.05720  -0.0206   0.1481   1.0000
  16.500   1.4348   0.06723   0.05970  -0.0203   0.1451   1.0000
  16.750   1.4356   0.06988   0.06244  -0.0201   0.1416   1.0000
  17.000   1.4345   0.07276   0.06538  -0.0200   0.1383   1.0000
  17.250   1.4313   0.07589   0.06852  -0.0200   0.1354   1.0000
  17.500   1.4329   0.07857   0.07135  -0.0201   0.1320   1.0000
  17.750   1.4312   0.08169   0.07457  -0.0203   0.1281   1.0000
  18.000   1.4260   0.08525   0.07818  -0.0207   0.1242   1.0000
  18.250   1.4243   0.08842   0.08143  -0.0210   0.1209   1.0000
  18.500   1.4211   0.09185   0.08496  -0.0215   0.1168   1.0000
  18.750   1.4149   0.09568   0.08883  -0.0222   0.1132   1.0000
  19.000   1.4112   0.09917   0.09239  -0.0229   0.1102   1.0000
<< Back to GOE 481A AIRFOIL (goe481a-il)

Polar data table (+)

Polar graphs


<< Back to GOE 481A AIRFOIL (goe481a-il)