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

GOE 602 AIRFOIL (goe602-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 602 AIRFOIL (goe602-il)
Reynolds number: 100,000
Max Cl/Cd: 57.55 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe602-il-100000-n5.txt
Download as CSV file: xf-goe602-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 602 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.3834   0.09193   0.08719  -0.0400   1.0000   0.0285
  -8.500  -0.3912   0.08887   0.08422  -0.0398   1.0000   0.0286
  -8.250  -0.4014   0.08606   0.08149  -0.0390   1.0000   0.0285
  -8.000  -0.4174   0.08338   0.07893  -0.0378   1.0000   0.0287
  -7.750  -0.4321   0.08102   0.07665  -0.0362   1.0000   0.0285
  -7.500  -0.4434   0.07768   0.07339  -0.0363   1.0000   0.0283
  -7.250  -0.4525   0.07398   0.06973  -0.0371   1.0000   0.0282
  -7.000  -0.4596   0.06973   0.06549  -0.0383   1.0000   0.0280
  -6.750  -0.4432   0.06132   0.05689  -0.0473   0.9944   0.0281
  -6.500  -0.4239   0.05351   0.04876  -0.0546   0.9879   0.0281
  -6.250  -0.4039   0.04612   0.04086  -0.0599   0.9815   0.0283
  -6.000  -0.3782   0.03966   0.03362  -0.0641   0.9764   0.0292
  -5.750  -0.3564   0.03588   0.02939  -0.0658   0.9703   0.0315
  -5.500  -0.3264   0.03342   0.02661  -0.0680   0.9662   0.0337
  -5.250  -0.3002   0.03048   0.02313  -0.0687   0.9605   0.0356
  -5.000  -0.2691   0.02823   0.02026  -0.0699   0.9558   0.0395
  -4.750  -0.2350   0.02584   0.01733  -0.0715   0.9526   0.0425
  -4.500  -0.2096   0.02465   0.01604  -0.0715   0.9457   0.0463
  -4.250  -0.1761   0.02345   0.01452  -0.0727   0.9413   0.0507
  -4.000  -0.1412   0.02231   0.01307  -0.0740   0.9374   0.0537
  -3.750  -0.1146   0.02133   0.01211  -0.0741   0.9303   0.0584
  -3.500  -0.0796   0.02044   0.01113  -0.0755   0.9261   0.0619
  -3.250  -0.0501   0.01974   0.01029  -0.0757   0.9196   0.0652
  -3.000  -0.0179   0.01905   0.00953  -0.0766   0.9137   0.0695
  -2.750   0.0202   0.01844   0.00886  -0.0786   0.9100   0.0789
  -2.500   0.0459   0.01798   0.00839  -0.0782   0.9010   0.0940
  -2.250   0.0827   0.01719   0.00777  -0.0800   0.8966   0.1363
  -2.000   0.1090   0.01641   0.00749  -0.0800   0.8882   0.2562
  -1.750   0.1439   0.01576   0.00719  -0.0815   0.8828   0.3808
  -1.500   0.1678   0.01503   0.00715  -0.0806   0.8742   0.5576
  -1.250   0.1969   0.01418   0.00712  -0.0793   0.8689   0.8054
  -1.000   0.2742   0.01382   0.00674  -0.0889   0.8677   1.0000
  -0.750   0.2997   0.01380   0.00654  -0.0883   0.8568   1.0000
  -0.500   0.3284   0.01375   0.00633  -0.0884   0.8471   1.0000
  -0.250   0.3617   0.01363   0.00606  -0.0893   0.8391   1.0000
   0.250   0.4145   0.01361   0.00581  -0.0884   0.8157   1.0000
   0.500   0.4431   0.01357   0.00568  -0.0884   0.8044   1.0000
   0.750   0.4728   0.01353   0.00554  -0.0885   0.7931   1.0000
   1.000   0.5032   0.01348   0.00540  -0.0888   0.7816   1.0000
   1.250   0.5307   0.01349   0.00535  -0.0885   0.7681   1.0000
   1.500   0.5587   0.01351   0.00531  -0.0884   0.7543   1.0000
   1.750   0.5867   0.01354   0.00527  -0.0882   0.7401   1.0000
   2.000   0.6147   0.01359   0.00526  -0.0880   0.7253   1.0000
   2.250   0.6424   0.01367   0.00528  -0.0877   0.7100   1.0000
   2.500   0.6697   0.01376   0.00532  -0.0874   0.6943   1.0000
   2.750   0.6954   0.01390   0.00541  -0.0868   0.6773   1.0000
   3.000   0.7210   0.01406   0.00554  -0.0862   0.6600   1.0000
   3.250   0.7464   0.01423   0.00568  -0.0856   0.6428   1.0000
   3.500   0.7715   0.01442   0.00582  -0.0849   0.6247   1.0000
   3.750   0.7963   0.01463   0.00600  -0.0842   0.6062   1.0000
   4.000   0.8199   0.01487   0.00623  -0.0832   0.5870   1.0000
   4.250   0.8435   0.01512   0.00646  -0.0823   0.5680   1.0000
   4.500   0.8672   0.01539   0.00673  -0.0814   0.5501   1.0000
   4.750   0.8904   0.01569   0.00703  -0.0805   0.5320   1.0000
   5.000   0.9133   0.01599   0.00736  -0.0795   0.5145   1.0000
   5.250   0.9358   0.01632   0.00771  -0.0785   0.4965   1.0000
   5.500   0.9582   0.01666   0.00811  -0.0774   0.4789   1.0000
   5.750   0.9800   0.01703   0.00849  -0.0763   0.4606   1.0000
   6.000   1.0006   0.01741   0.00893  -0.0749   0.4394   1.0000
   6.250   1.0180   0.01785   0.00931  -0.0730   0.4082   1.0000
   6.500   1.0319   0.01841   0.00970  -0.0705   0.3639   1.0000
   6.750   1.0447   0.01911   0.01021  -0.0680   0.3137   1.0000
   7.000   1.0581   0.01989   0.01080  -0.0658   0.2662   1.0000
   7.250   1.0691   0.02092   0.01154  -0.0633   0.2117   1.0000
   7.500   1.0759   0.02237   0.01252  -0.0604   0.1364   1.0000
   7.750   1.0803   0.02411   0.01379  -0.0572   0.0798   1.0000
   8.000   1.0839   0.02579   0.01511  -0.0538   0.0379   1.0000
   8.250   1.0926   0.02700   0.01641  -0.0509   0.0331   1.0000
   8.500   1.1005   0.02827   0.01782  -0.0481   0.0302   1.0000
   8.750   1.1036   0.02987   0.01956  -0.0449   0.0274   1.0000
   9.000   1.1101   0.03122   0.02113  -0.0422   0.0258   1.0000
   9.250   1.1155   0.03266   0.02277  -0.0395   0.0247   1.0000
   9.500   1.1195   0.03424   0.02453  -0.0370   0.0239   1.0000
   9.750   1.1224   0.03598   0.02644  -0.0345   0.0231   1.0000
  10.000   1.1251   0.03783   0.02843  -0.0322   0.0222   1.0000
  10.250   1.1280   0.03979   0.03052  -0.0302   0.0211   1.0000
  10.500   1.1309   0.04186   0.03270  -0.0284   0.0200   1.0000
  10.750   1.1340   0.04410   0.03503  -0.0267   0.0190   1.0000
  11.000   1.1398   0.04645   0.03745  -0.0252   0.0184   1.0000
  11.250   1.1513   0.04892   0.03999  -0.0239   0.0178   1.0000
  11.500   1.1717   0.05195   0.04315  -0.0230   0.0173   1.0000
  11.750   1.1819   0.05454   0.04592  -0.0219   0.0170   1.0000
  12.000   1.1864   0.05714   0.04885  -0.0207   0.0168   1.0000
  12.250   1.1873   0.05991   0.05188  -0.0196   0.0166   1.0000
  12.500   1.1850   0.06289   0.05514  -0.0187   0.0162   1.0000
  12.750   1.1804   0.06618   0.05870  -0.0181   0.0159   1.0000
  13.000   1.1742   0.06982   0.06260  -0.0179   0.0156   1.0000
  13.250   1.1657   0.07382   0.06686  -0.0181   0.0153   1.0000
  13.500   1.1558   0.07819   0.07148  -0.0188   0.0152   1.0000
  13.750   1.1437   0.08304   0.07658  -0.0200   0.0151   1.0000
  14.000   1.1306   0.08830   0.08208  -0.0218   0.0151   1.0000
  14.250   1.1160   0.09406   0.08806  -0.0243   0.0150   1.0000
  14.500   1.1001   0.10041   0.09464  -0.0274   0.0150   1.0000
  14.750   1.0836   0.10731   0.10174  -0.0311   0.0151   1.0000
  15.000   1.0671   0.11471   0.10933  -0.0355   0.0151   1.0000
  15.250   1.0502   0.12271   0.11751  -0.0404   0.0153   1.0000
  15.500   1.0328   0.13142   0.12636  -0.0460   0.0155   1.0000
<< Back to GOE 602 AIRFOIL (goe602-il)

Polar data table (+)

Polar graphs


<< Back to GOE 602 AIRFOIL (goe602-il)