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GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 55 AIRFOIL (goe55-il)
Reynolds number: 100,000
Max Cl/Cd: 42.35 at α=5.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe55-il-100000.txt
Download as CSV file: xf-goe55-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 55 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.500  -0.6211   0.11228   0.10724   0.0386   1.0000   0.0842
  -8.250  -0.6202   0.10922   0.10421   0.0354   1.0000   0.0872
  -8.000  -0.6275   0.10713   0.10219   0.0249   1.0000   0.0889
  -7.750  -0.6215   0.10214   0.09720   0.0160   1.0000   0.0898
  -7.500  -0.6100   0.09728   0.09238   0.0247   1.0000   0.0913
  -7.250  -0.5985   0.09359   0.08871   0.0261   1.0000   0.0935
  -7.000  -0.5877   0.08987   0.08500   0.0239   1.0000   0.0964
  -6.750  -0.5658   0.08555   0.08038   0.0012   1.0000   0.1031
  -6.500  -0.5580   0.08008   0.07509   0.0062   1.0000   0.1046
  -6.250  -0.5464   0.07682   0.07188   0.0097   1.0000   0.1073
  -6.000  -0.5292   0.07314   0.06818   0.0072   1.0000   0.1121
  -5.750  -0.5021   0.06761   0.06241  -0.0039   1.0000   0.1190
  -5.500  -0.4877   0.06428   0.05916  -0.0020   1.0000   0.1217
  -5.250  -0.4551   0.06011   0.05462  -0.0110   1.0000   0.1332
  -5.000  -0.4408   0.05661   0.05128  -0.0087   1.0000   0.1362
  -4.750  -0.4127   0.05309   0.04755  -0.0130   1.0000   0.1494
  -4.500  -0.3930   0.05031   0.04485  -0.0123   1.0000   0.1562
  -4.250  -0.3665   0.04704   0.04146  -0.0147   1.0000   0.1682
  -4.000  -0.3328   0.04543   0.03947  -0.0186   1.0000   0.1931
  -3.750  -0.3136   0.04149   0.03573  -0.0179   1.0000   0.1989
  -3.500  -0.2721   0.02399   0.01866  -0.0182   1.0000   0.2269
  -3.250  -0.2520   0.02183   0.01660  -0.0174   1.0000   0.2480
  -3.000  -0.2303   0.01974   0.01454  -0.0171   1.0000   0.2802
  -2.750  -0.1574   0.02671   0.01862  -0.0285   1.0000   0.1186
  -2.500  -0.1272   0.02431   0.01611  -0.0288   1.0000   0.1117
  -2.250  -0.0961   0.02249   0.01402  -0.0290   1.0000   0.1104
  -2.000  -0.0654   0.02094   0.01224  -0.0291   1.0000   0.1100
  -1.750  -0.0350   0.01953   0.01062  -0.0290   1.0000   0.1091
  -1.500  -0.0053   0.01837   0.00935  -0.0288   1.0000   0.1103
  -1.250   0.0240   0.01740   0.00834  -0.0286   1.0000   0.1137
  -1.000   0.0530   0.01642   0.00741  -0.0284   1.0000   0.1201
  -0.750   0.0818   0.01553   0.00672  -0.0282   1.0000   0.1271
  -0.500   0.1113   0.01460   0.00610  -0.0283   1.0000   0.1366
  -0.250   0.1452   0.01491   0.00573  -0.0276   0.6712   0.1602
   0.000   0.1712   0.01472   0.00535  -0.0266   0.6013   0.2091
   0.250   0.1922   0.01224   0.00482  -0.0240   0.5717   1.0000
   0.500   0.2205   0.01264   0.00480  -0.0236   0.5487   1.0000
   0.750   0.2487   0.01304   0.00490  -0.0234   0.5314   1.0000
   1.000   0.2773   0.01343   0.00506  -0.0232   0.5174   1.0000
   1.250   0.3058   0.01383   0.00528  -0.0231   0.5043   1.0000
   1.500   0.3339   0.01422   0.00550  -0.0229   0.4889   1.0000
   1.750   0.3620   0.01461   0.00573  -0.0227   0.4731   1.0000
   2.000   0.3900   0.01502   0.00599  -0.0224   0.4591   1.0000
   2.250   0.4180   0.01546   0.00631  -0.0222   0.4470   1.0000
   2.500   0.4465   0.01584   0.00668  -0.0222   0.4347   1.0000
   2.750   0.4748   0.01628   0.00710  -0.0221   0.4227   1.0000
   3.000   0.5027   0.01675   0.00753  -0.0219   0.4106   1.0000
   3.250   0.5305   0.01724   0.00795  -0.0216   0.3976   1.0000
   3.500   0.5581   0.01767   0.00835  -0.0214   0.3819   1.0000
   3.750   0.5856   0.01799   0.00871  -0.0211   0.3627   1.0000
   4.000   0.6127   0.01824   0.00895  -0.0208   0.3425   1.0000
   4.250   0.6398   0.01847   0.00913  -0.0204   0.3232   1.0000
   4.500   0.6669   0.01869   0.00930  -0.0200   0.3048   1.0000
   4.750   0.6943   0.01878   0.00951  -0.0197   0.2838   1.0000
   5.000   0.7216   0.01859   0.00936  -0.0194   0.2608   1.0000
   5.250   0.7492   0.01841   0.00928  -0.0191   0.2349   1.0000
   5.500   0.7767   0.01834   0.00917  -0.0188   0.1998   1.0000
   5.750   0.8021   0.01971   0.01007  -0.0187   0.1416   1.0000
   6.000   0.8277   0.02129   0.01151  -0.0185   0.1162   1.0000
   6.250   0.8537   0.02258   0.01279  -0.0181   0.1046   1.0000
   6.500   0.8793   0.02408   0.01426  -0.0177   0.0983   1.0000
   6.750   0.9057   0.02550   0.01583  -0.0172   0.0936   1.0000
   7.000   0.9314   0.02702   0.01738  -0.0168   0.0901   1.0000
   7.250   0.9564   0.02917   0.01951  -0.0165   0.0873   1.0000
   7.500   0.9817   0.03094   0.02170  -0.0161   0.0846   1.0000
   7.750   1.0058   0.03307   0.02419  -0.0157   0.0822   1.0000
   8.000   1.0286   0.03570   0.02721  -0.0154   0.0810   1.0000
   8.250   1.0493   0.03883   0.03079  -0.0152   0.0802   1.0000
   8.500   1.0671   0.04247   0.03494  -0.0150   0.0795   1.0000
   8.750   1.0841   0.04579   0.03858  -0.0149   0.0783   1.0000
   9.000   1.1000   0.04914   0.04212  -0.0148   0.0769   1.0000
   9.250   1.0819   0.05961   0.05382  -0.0177   0.0792   1.0000
   9.500   1.0253   0.07659   0.07149  -0.0284   0.0841   1.0000
   9.750   0.9890   0.09133   0.08628  -0.0408   0.0889   1.0000
  10.000   0.9980   0.09324   0.08819  -0.0390   0.0881   1.0000
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