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

GOE 187 (SCHTTE-LANZ 2U10) AIRFOIL (goe187-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: GOE 187 (SCHTTE-LANZ 2U10) AIRFOIL (goe187-il)
Reynolds number: 1,000,000
Max Cl/Cd: 112.24 at α=4.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe187-il-1000000-n5.txt
Download as CSV file: xf-goe187-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 187 (SCHTTE-LANZ 2U10) AIRFOIL              
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.250  -0.3318   0.11005   0.10849  -0.0151   1.0000   0.0033
  -9.000  -0.3252   0.10683   0.10529  -0.0166   1.0000   0.0033
  -8.750  -0.3188   0.10364   0.10211  -0.0181   1.0000   0.0033
  -6.750  -0.2509   0.07754   0.07522  -0.0363   0.7388   0.0035
  -6.500  -0.2353   0.07424   0.07185  -0.0402   0.7239   0.0037
  -6.250  -0.2173   0.07072   0.06825  -0.0445   0.7106   0.0038
  -6.000  -0.1972   0.06705   0.06451  -0.0491   0.6983   0.0039
  -5.750  -0.1752   0.06327   0.06066  -0.0538   0.6871   0.0040
  -5.500  -0.1515   0.05949   0.05680  -0.0583   0.6779   0.0042
  -5.250  -0.1261   0.05570   0.05293  -0.0626   0.6700   0.0043
  -5.000  -0.0993   0.05195   0.04909  -0.0666   0.6636   0.0045
  -4.750  -0.0708   0.04824   0.04528  -0.0703   0.6566   0.0048
  -4.500  -0.0390   0.04456   0.04146  -0.0737   0.6501   0.0051
  -4.250  -0.0084   0.04103   0.03778  -0.0764   0.6426   0.0052
  -4.000   0.0216   0.03766   0.03425  -0.0784   0.6358   0.0052
  -3.750   0.0508   0.03450   0.03092  -0.0799   0.6283   0.0052
  -3.250   0.1054   0.02845   0.02452  -0.0823   0.6150   0.0045
  -3.000   0.1357   0.02550   0.02133  -0.0828   0.6084   0.0042
  -2.750   0.1658   0.02293   0.01852  -0.0831   0.6011   0.0044
  -2.500   0.1959   0.02059   0.01591  -0.0831   0.5932   0.0046
  -2.250   0.2264   0.01843   0.01348  -0.0827   0.5836   0.0051
  -2.000   0.2556   0.01647   0.01123  -0.0827   0.5740   0.0048
  -1.750   0.2852   0.01457   0.00903  -0.0824   0.5631   0.0047
  -1.500   0.3150   0.01246   0.00656  -0.0819   0.5515   0.0046
  -1.250   0.3438   0.01001   0.00371  -0.0813   0.5373   0.0050
  -1.000   0.3714   0.00935   0.00282  -0.0812   0.5134   0.0055
  -0.750   0.3986   0.00926   0.00254  -0.0811   0.4757   0.0061
  -0.500   0.4260   0.00938   0.00252  -0.0811   0.4442   0.0072
  -0.250   0.4538   0.00919   0.00217  -0.0811   0.4253   0.0080
   0.000   0.4816   0.00900   0.00184  -0.0812   0.4082   0.0093
   0.250   0.5095   0.00903   0.00181  -0.0812   0.3930   0.0114
   0.500   0.5375   0.00907   0.00178  -0.0813   0.3809   0.0133
   0.750   0.5657   0.00899   0.00162  -0.0814   0.3717   0.0188
   1.000   0.5936   0.00904   0.00161  -0.0814   0.3618   0.0238
   1.250   0.6217   0.00907   0.00159  -0.0815   0.3535   0.0275
   1.500   0.6496   0.00911   0.00159  -0.0816   0.3451   0.0341
   2.000   0.7033   0.00726   0.00182  -0.0821   0.3293   1.0000
   2.250   0.7310   0.00738   0.00187  -0.0821   0.3220   1.0000
   2.500   0.7588   0.00750   0.00193  -0.0822   0.3164   1.0000
   2.750   0.7865   0.00761   0.00200  -0.0822   0.3100   1.0000
   3.000   0.8140   0.00775   0.00208  -0.0823   0.3034   1.0000
   3.250   0.8417   0.00786   0.00217  -0.0824   0.2976   1.0000
   3.500   0.8692   0.00800   0.00226  -0.0824   0.2909   1.0000
   3.750   0.8966   0.00813   0.00236  -0.0825   0.2840   1.0000
   4.000   0.9237   0.00831   0.00249  -0.0825   0.2733   1.0000
   4.250   0.9507   0.00850   0.00262  -0.0825   0.2615   1.0000
   4.500   0.9776   0.00871   0.00277  -0.0825   0.2479   1.0000
   4.750   1.0037   0.00901   0.00297  -0.0824   0.2249   1.0000
   5.000   1.0270   0.00968   0.00337  -0.0820   0.1739   1.0000
   5.250   1.0523   0.01006   0.00366  -0.0818   0.1553   1.0000
   5.500   1.0781   0.01037   0.00393  -0.0816   0.1418   1.0000
   5.750   1.1028   0.01081   0.00425  -0.0814   0.1187   1.0000
   6.000   1.1269   0.01132   0.00463  -0.0810   0.0936   1.0000
   6.250   1.1520   0.01168   0.00495  -0.0808   0.0858   1.0000
   6.500   1.1775   0.01197   0.00525  -0.0806   0.0795   1.0000
   6.750   1.2024   0.01232   0.00559  -0.0803   0.0734   1.0000
   7.000   1.2275   0.01262   0.00590  -0.0801   0.0682   1.0000
   7.250   1.2458   0.01374   0.00672  -0.0790   0.0194   1.0000
   7.500   1.2678   0.01438   0.00737  -0.0783   0.0067   1.0000
   7.750   1.2911   0.01484   0.00787  -0.0777   0.0047   1.0000
   8.000   1.3139   0.01533   0.00840  -0.0772   0.0036   1.0000
   8.250   1.3366   0.01580   0.00893  -0.0766   0.0029   1.0000
   8.500   1.3578   0.01640   0.00958  -0.0758   0.0025   1.0000
   8.750   1.3790   0.01698   0.01023  -0.0749   0.0022   1.0000
   9.000   1.3995   0.01759   0.01091  -0.0740   0.0020   1.0000
   9.250   1.4191   0.01824   0.01166  -0.0730   0.0018   1.0000
   9.500   1.4378   0.01893   0.01242  -0.0719   0.0016   1.0000
   9.750   1.4534   0.01986   0.01344  -0.0703   0.0014   1.0000
  10.000   1.4705   0.02057   0.01423  -0.0690   0.0013   1.0000
  10.250   1.4849   0.02144   0.01519  -0.0672   0.0012   1.0000
  10.500   1.4967   0.02238   0.01623  -0.0651   0.0011   1.0000
  10.750   1.5041   0.02336   0.01732  -0.0623   0.0010   1.0000
  11.000   1.5105   0.02445   0.01849  -0.0595   0.0010   1.0000
  11.250   1.5163   0.02566   0.01981  -0.0570   0.0009   1.0000
  11.500   1.5212   0.02705   0.02130  -0.0547   0.0009   1.0000
  11.750   1.5255   0.02863   0.02298  -0.0528   0.0008   1.0000
  12.000   1.5276   0.03052   0.02499  -0.0510   0.0008   1.0000
  12.250   1.5271   0.03284   0.02742  -0.0495   0.0007   1.0000
  12.500   1.5223   0.03580   0.03052  -0.0483   0.0007   1.0000
  12.750   1.5118   0.03965   0.03454  -0.0475   0.0007   1.0000
  13.000   1.5093   0.04280   0.03785  -0.0473   0.0007   1.0000
  13.250   1.5049   0.04632   0.04150  -0.0472   0.0007   1.0000
  13.500   1.4989   0.05010   0.04542  -0.0473   0.0006   1.0000
  13.750   1.4908   0.05428   0.04974  -0.0477   0.0006   1.0000
  14.000   1.4806   0.05883   0.05443  -0.0483   0.0006   1.0000
  14.250   1.4700   0.06357   0.05931  -0.0490   0.0006   1.0000
  14.500   1.4592   0.06857   0.06443  -0.0501   0.0006   1.0000
  14.750   1.4470   0.07390   0.06990  -0.0514   0.0006   1.0000
  15.000   1.4349   0.07941   0.07555  -0.0529   0.0006   1.0000
  15.250   1.4226   0.08510   0.08137  -0.0545   0.0006   1.0000
  15.500   1.4102   0.09085   0.08724  -0.0562   0.0006   1.0000
  15.750   1.3970   0.09685   0.09338  -0.0582   0.0005   1.0000
  16.000   1.3846   0.10281   0.09946  -0.0601   0.0005   1.0000
<< Back to GOE 187 (SCHTTE-LANZ 2U10) AIRFOIL (goe187-il)

Polar data table (+)

Polar graphs


<< Back to GOE 187 (SCHTTE-LANZ 2U10) AIRFOIL (goe187-il)