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

GOE 377 AIRFOIL (goe377-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: GOE 377 AIRFOIL (goe377-il)
Reynolds number: 500,000
Max Cl/Cd: 91.71 at α=2.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe377-il-500000-n5.txt
Download as CSV file: xf-goe377-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 377 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.2746   0.08930   0.08723  -0.0294   1.0000   0.0091
  -8.750  -0.2747   0.08607   0.08401  -0.0294   1.0000   0.0091
  -7.750  -0.3603   0.08918   0.08709  -0.0253   1.0000   0.0091
  -7.250  -0.2618   0.05985   0.05788  -0.0393   0.9856   0.0058
  -6.750  -0.3128   0.06989   0.06778  -0.0442   0.9831   0.0055
  -6.500  -0.2865   0.06333   0.06117  -0.0532   0.9778   0.0057
  -6.250  -0.2617   0.05862   0.05641  -0.0593   0.9703   0.0055
  -6.000  -0.2310   0.05283   0.05051  -0.0668   0.9644   0.0055
  -5.750  -0.2000   0.04560   0.04312  -0.0741   0.9563   0.0057
  -5.500  -0.1714   0.03847   0.03575  -0.0795   0.9482   0.0061
  -5.250  -0.1422   0.03632   0.03349  -0.0819   0.9427   0.0068
  -5.000  -0.1163   0.03060   0.02746  -0.0840   0.9348   0.0072
  -4.750  -0.1075   0.01436   0.00951  -0.0820   0.9240   0.0076
  -4.500  -0.0821   0.01271   0.00747  -0.0815   0.9174   0.0084
  -4.250  -0.0547   0.01167   0.00613  -0.0813   0.9118   0.0090
  -4.000  -0.0293   0.01076   0.00503  -0.0807   0.9048   0.0108
  -3.500   0.0251   0.01012   0.00423  -0.0803   0.8915   0.0147
  -3.250   0.0527   0.00988   0.00394  -0.0801   0.8849   0.0186
  -3.000   0.0799   0.00994   0.00396  -0.0798   0.8771   0.0236
  -2.750   0.1076   0.00996   0.00389  -0.0797   0.8700   0.0260
  -2.500   0.1335   0.00966   0.00352  -0.0792   0.8613   0.0285
  -2.250   0.1600   0.00953   0.00334  -0.0789   0.8517   0.0307
  -2.000   0.1863   0.00942   0.00315  -0.0784   0.8384   0.0329
  -1.750   0.2116   0.00920   0.00281  -0.0777   0.8199   0.0341
  -1.500   0.2368   0.00901   0.00250  -0.0770   0.7990   0.0351
  -1.250   0.2616   0.00878   0.00215  -0.0762   0.7783   0.0353
  -1.000   0.2863   0.00861   0.00185  -0.0754   0.7581   0.0356
  -0.750   0.3112   0.00849   0.00162  -0.0747   0.7368   0.0362
  -0.500   0.3363   0.00843   0.00144  -0.0740   0.7175   0.0370
  -0.250   0.3611   0.00833   0.00125  -0.0732   0.6979   0.0408
   0.000   0.3862   0.00832   0.00116  -0.0726   0.6781   0.0458
   0.250   0.4112   0.00835   0.00111  -0.0719   0.6570   0.0511
   0.500   0.4359   0.00838   0.00108  -0.0712   0.6339   0.0627
   1.000   0.4845   0.00825   0.00113  -0.0698   0.5938   0.1878
   1.250   0.5048   0.00769   0.00121  -0.0687   0.5754   0.4715
   1.500   0.5896   0.00677   0.00138  -0.0818   0.5421   1.0000
   1.750   0.6128   0.00697   0.00144  -0.0808   0.5173   1.0000
   2.000   0.6363   0.00715   0.00152  -0.0800   0.4956   1.0000
   2.250   0.6601   0.00732   0.00160  -0.0791   0.4758   1.0000
   2.500   0.6838   0.00750   0.00169  -0.0783   0.4559   1.0000
   2.750   0.7071   0.00771   0.00179  -0.0774   0.4304   1.0000
   3.000   0.7298   0.00797   0.00193  -0.0764   0.4005   1.0000
   3.250   0.7525   0.00825   0.00208  -0.0754   0.3720   1.0000
   3.500   0.7751   0.00854   0.00224  -0.0744   0.3448   1.0000
   3.750   0.7973   0.00886   0.00245  -0.0734   0.3164   1.0000
   4.000   0.8203   0.00914   0.00264  -0.0725   0.2960   1.0000
   4.250   0.8439   0.00936   0.00283  -0.0717   0.2826   1.0000
   4.500   0.8674   0.00959   0.00303  -0.0709   0.2693   1.0000
   4.750   0.8899   0.00990   0.00327  -0.0700   0.2462   1.0000
   5.000   0.9123   0.01023   0.00351  -0.0691   0.2227   1.0000
   5.250   0.9328   0.01072   0.00381  -0.0678   0.1824   1.0000
   5.500   0.9510   0.01144   0.00422  -0.0663   0.1258   1.0000
   5.750   0.9663   0.01247   0.00486  -0.0643   0.0568   1.0000
   6.000   0.9843   0.01326   0.00548  -0.0626   0.0195   1.0000
   6.250   1.0060   0.01368   0.00595  -0.0615   0.0126   1.0000
   6.500   1.0270   0.01418   0.00652  -0.0603   0.0102   1.0000
   6.750   1.0467   0.01483   0.00730  -0.0588   0.0083   1.0000
   7.000   1.0679   0.01528   0.00782  -0.0577   0.0075   1.0000
   7.250   1.0882   0.01581   0.00844  -0.0564   0.0065   1.0000
   7.500   1.1074   0.01644   0.00915  -0.0550   0.0060   1.0000
   7.750   1.1228   0.01742   0.01023  -0.0530   0.0054   1.0000
   8.000   1.1391   0.01826   0.01118  -0.0511   0.0052   1.0000
   8.250   1.1550   0.01911   0.01214  -0.0491   0.0048   1.0000
   8.500   1.1703   0.01997   0.01310  -0.0471   0.0046   1.0000
   8.750   1.1857   0.02077   0.01398  -0.0453   0.0042   1.0000
   9.000   1.2000   0.02162   0.01491  -0.0433   0.0039   1.0000
   9.250   1.2136   0.02247   0.01582  -0.0413   0.0036   1.0000
   9.500   1.2169   0.02401   0.01747  -0.0376   0.0034   1.0000
   9.750   1.2235   0.02524   0.01886  -0.0344   0.0033   1.0000
  10.000   1.2315   0.02641   0.02018  -0.0316   0.0032   1.0000
  10.250   1.2371   0.02792   0.02184  -0.0286   0.0031   1.0000
  10.500   1.2426   0.02951   0.02360  -0.0258   0.0030   1.0000
  10.750   1.2466   0.03135   0.02564  -0.0230   0.0028   1.0000
  11.000   1.2489   0.03345   0.02793  -0.0204   0.0028   1.0000
  11.250   1.2502   0.03564   0.03033  -0.0179   0.0028   1.0000
  11.500   1.2490   0.03811   0.03302  -0.0155   0.0027   1.0000
  11.750   1.2457   0.04087   0.03599  -0.0133   0.0026   1.0000
  12.000   1.2403   0.04381   0.03915  -0.0114   0.0025   1.0000
  12.250   1.2327   0.04719   0.04275  -0.0100   0.0026   1.0000
  12.500   1.2204   0.05129   0.04710  -0.0091   0.0025   1.0000
  12.750   1.2091   0.05540   0.05142  -0.0090   0.0025   1.0000
  13.000   1.1961   0.06007   0.05629  -0.0096   0.0025   1.0000
  13.250   1.1807   0.06544   0.06187  -0.0111   0.0025   1.0000
  13.500   1.1643   0.07143   0.06804  -0.0134   0.0026   1.0000
  13.750   1.1481   0.07778   0.07458  -0.0164   0.0025   1.0000
  14.000   1.1303   0.08498   0.08195  -0.0202   0.0026   1.0000
  14.250   1.1116   0.09290   0.09003  -0.0247   0.0025   1.0000
  14.500   1.0957   0.10071   0.09798  -0.0292   0.0026   1.0000
  14.750   1.0786   0.10931   0.10671  -0.0343   0.0026   1.0000
  15.000   1.0604   0.11860   0.11612  -0.0397   0.0027   1.0000
<< Back to GOE 377 AIRFOIL (goe377-il)

Polar data table (+)

Polar graphs


<< Back to GOE 377 AIRFOIL (goe377-il)