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GOE 55 AIRFOIL (goe55-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 55 AIRFOIL (goe55-il)
Reynolds number: 100,000
Max Cl/Cd: 43.3 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe55-il-100000-n5.txt
Download as CSV file: xf-goe55-il-100000-n5.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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.250  -0.6092   0.10560   0.10055   0.0330   1.0000   0.0615
  -8.000  -0.6048   0.10207   0.09705   0.0312   1.0000   0.0630
  -7.750  -0.6016   0.09849   0.09351   0.0285   1.0000   0.0647
  -7.500  -0.5958   0.09363   0.08863   0.0145   1.0000   0.0678
  -7.000  -0.5747   0.08457   0.07959   0.0113   1.0000   0.0697
  -6.750  -0.5619   0.08110   0.07612   0.0110   1.0000   0.0710
  -6.250  -0.5181   0.06646   0.06108  -0.0067   1.0000   0.0563
  -6.000  -0.4975   0.06171   0.05621  -0.0101   1.0000   0.0560
  -5.750  -0.4755   0.05715   0.05152  -0.0132   1.0000   0.0557
  -5.500  -0.4530   0.05300   0.04724  -0.0156   1.0000   0.0544
  -5.250  -0.4268   0.04832   0.04234  -0.0189   1.0000   0.0530
  -5.000  -0.3981   0.04351   0.03721  -0.0221   1.0000   0.0520
  -4.750  -0.3687   0.03949   0.03285  -0.0247   1.0000   0.0529
  -4.500  -0.3384   0.03571   0.02868  -0.0268   1.0000   0.0538
  -4.250  -0.3081   0.03226   0.02483  -0.0283   1.0000   0.0535
  -4.000  -0.2777   0.02931   0.02148  -0.0295   1.0000   0.0535
  -3.750  -0.2474   0.02680   0.01856  -0.0303   1.0000   0.0538
  -3.500  -0.2172   0.02471   0.01608  -0.0309   1.0000   0.0549
  -3.250  -0.1869   0.02305   0.01398  -0.0312   1.0000   0.0567
  -3.000  -0.1578   0.02150   0.01223  -0.0314   1.0000   0.0576
  -2.750  -0.1291   0.02024   0.01087  -0.0315   1.0000   0.0584
  -2.500  -0.1004   0.01920   0.00977  -0.0315   1.0000   0.0595
  -2.250  -0.0719   0.01828   0.00881  -0.0314   1.0000   0.0609
  -2.000  -0.0434   0.01744   0.00795  -0.0313   1.0000   0.0627
  -1.750  -0.0151   0.01677   0.00727  -0.0311   1.0000   0.0660
  -1.500   0.0132   0.01610   0.00664  -0.0309   1.0000   0.0693
  -1.250   0.0520   0.01563   0.00616  -0.0327   0.8377   0.0729
  -1.000   0.0734   0.01594   0.00577  -0.0300   0.6411   0.0766
  -0.750   0.0989   0.01599   0.00536  -0.0290   0.5431   0.0816
  -0.500   0.1262   0.01598   0.00511  -0.0287   0.4977   0.0912
  -0.250   0.1542   0.01585   0.00484  -0.0285   0.4711   0.1053
   0.000   0.1823   0.01559   0.00458  -0.0283   0.4531   0.1275
   0.250   0.2106   0.01530   0.00449  -0.0283   0.4398   0.1852
   0.500   0.2306   0.01297   0.00433  -0.0261   0.4305   1.0000
   0.750   0.2591   0.01322   0.00432  -0.0259   0.4209   1.0000
   1.000   0.2876   0.01347   0.00437  -0.0257   0.4123   1.0000
   1.250   0.3159   0.01373   0.00447  -0.0255   0.4044   1.0000
   1.500   0.3443   0.01399   0.00460  -0.0254   0.3957   1.0000
   1.750   0.3724   0.01427   0.00471  -0.0252   0.3839   1.0000
   2.000   0.4005   0.01452   0.00483  -0.0250   0.3708   1.0000
   2.250   0.4287   0.01475   0.00501  -0.0249   0.3595   1.0000
   2.500   0.4568   0.01503   0.00520  -0.0247   0.3512   1.0000
   2.750   0.4851   0.01528   0.00544  -0.0246   0.3423   1.0000
   3.000   0.5132   0.01557   0.00570  -0.0244   0.3346   1.0000
   3.250   0.5413   0.01582   0.00595  -0.0243   0.3242   1.0000
   3.500   0.5694   0.01602   0.00617  -0.0241   0.3111   1.0000
   3.750   0.5973   0.01620   0.00636  -0.0240   0.2961   1.0000
   4.000   0.6253   0.01634   0.00653  -0.0238   0.2794   1.0000
   4.250   0.6532   0.01647   0.00669  -0.0237   0.2611   1.0000
   4.500   0.6813   0.01663   0.00690  -0.0236   0.2416   1.0000
   4.750   0.7092   0.01683   0.00713  -0.0234   0.2180   1.0000
   5.000   0.7367   0.01713   0.00734  -0.0233   0.1900   1.0000
   5.250   0.7638   0.01764   0.00771  -0.0233   0.1621   1.0000
   5.500   0.7908   0.01829   0.00828  -0.0233   0.1392   1.0000
   5.750   0.8175   0.01906   0.00899  -0.0232   0.1156   1.0000
   6.000   0.8442   0.01990   0.00982  -0.0232   0.0915   1.0000
   6.250   0.8705   0.02082   0.01068  -0.0232   0.0754   1.0000
   6.500   0.8968   0.02169   0.01160  -0.0230   0.0679   1.0000
   6.750   0.9223   0.02268   0.01258  -0.0229   0.0636   1.0000
   7.000   0.9478   0.02360   0.01366  -0.0227   0.0609   1.0000
   7.250   0.9728   0.02458   0.01477  -0.0224   0.0586   1.0000
   7.500   0.9972   0.02564   0.01592  -0.0221   0.0568   1.0000
   7.750   1.0209   0.02681   0.01714  -0.0218   0.0550   1.0000
   8.000   1.0439   0.02810   0.01853  -0.0214   0.0533   1.0000
   8.250   1.0672   0.02937   0.01999  -0.0209   0.0517   1.0000
   8.500   1.0898   0.03079   0.02159  -0.0204   0.0505   1.0000
   8.750   1.1117   0.03236   0.02333  -0.0199   0.0495   1.0000
   9.000   1.1330   0.03406   0.02524  -0.0194   0.0486   1.0000
   9.250   1.1534   0.03588   0.02724  -0.0189   0.0477   1.0000
   9.500   1.1727   0.03775   0.02926  -0.0184   0.0466   1.0000
   9.750   1.1907   0.03992   0.03151  -0.0180   0.0454   1.0000
  10.000   1.2060   0.04255   0.03444  -0.0175   0.0445   1.0000
  10.250   1.2186   0.04546   0.03780  -0.0170   0.0439   1.0000
  10.500   1.2269   0.04883   0.04161  -0.0166   0.0434   1.0000
  10.750   1.2295   0.05267   0.04590  -0.0165   0.0428   1.0000
  11.000   1.2250   0.05701   0.05065  -0.0167   0.0423   1.0000
  11.250   1.2111   0.06198   0.05598  -0.0177   0.0419   1.0000
  11.500   1.1843   0.06846   0.06276  -0.0213   0.0421   1.0000
  11.750   1.1455   0.08096   0.07554  -0.0336   0.0428   1.0000
  12.000   1.0909   0.09814   0.09287  -0.0474   0.0437   1.0000
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