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GOE 391 AIRFOIL (goe391-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: GOE 391 AIRFOIL (goe391-il)
Reynolds number: 50,000
Max Cl/Cd: 39.62 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe391-il-50000.txt
Download as CSV file: xf-goe391-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 391 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.000  -0.4482   0.10238   0.09546  -0.0187   1.0000   0.1294
  -7.750  -0.4601   0.10166   0.09489  -0.0208   1.0000   0.1316
  -7.500  -0.4482   0.09645   0.08972  -0.0181   1.0000   0.1381
  -7.250  -0.4512   0.09444   0.08779  -0.0198   1.0000   0.1440
  -7.000  -0.4527   0.09182   0.08526  -0.0218   1.0000   0.1474
  -6.750  -0.4461   0.08826   0.08174  -0.0202   1.0000   0.1571
  -6.500  -0.4438   0.08500   0.07856  -0.0206   1.0000   0.1635
  -6.250  -0.4430   0.08333   0.07689  -0.0239   1.0000   0.1735
  -6.000  -0.4358   0.07909   0.07273  -0.0207   1.0000   0.1846
  -5.750  -0.4311   0.07587   0.06956  -0.0201   1.0000   0.1967
  -5.500  -0.4262   0.07295   0.06667  -0.0200   1.0000   0.2120
  -5.250  -0.4216   0.07024   0.06392  -0.0212   1.0000   0.2307
  -5.000  -0.0732   0.04279   0.03551  -0.0247   1.0000   1.0000
  -4.750  -0.0638   0.04081   0.03357  -0.0252   1.0000   1.0000
  -4.500  -0.0546   0.03888   0.03169  -0.0256   1.0000   1.0000
  -4.250  -0.0455   0.03706   0.02993  -0.0260   1.0000   1.0000
  -4.000  -0.0872   0.03836   0.03147  -0.0138   1.0000   0.9734
  -3.750  -0.1339   0.03965   0.03301  -0.0015   1.0000   0.9413
  -3.500  -0.1743   0.03989   0.03349   0.0082   1.0000   0.9040
  -3.250  -0.2200   0.04014   0.03400   0.0182   1.0000   0.8733
  -3.000  -0.2717   0.04030   0.03444   0.0288   1.0000   0.8442
  -2.750  -0.3281   0.04015   0.03456   0.0393   1.0000   0.8100
  -2.250  -0.1816   0.03343   0.02509  -0.0320   1.0000   0.2676
  -2.000  -0.1381   0.03122   0.02190  -0.0339   1.0000   0.1986
  -1.750  -0.1057   0.02920   0.01924  -0.0336   1.0000   0.1697
  -1.500  -0.0766   0.02751   0.01704  -0.0329   1.0000   0.1556
  -1.250  -0.0495   0.02603   0.01520  -0.0322   1.0000   0.1546
  -1.000  -0.0221   0.02481   0.01358  -0.0316   1.0000   0.1563
  -0.750   0.0042   0.02369   0.01219  -0.0306   1.0000   0.1548
  -0.500   0.0332   0.02279   0.01099  -0.0300   1.0000   0.1566
  -0.250   0.0612   0.02198   0.01002  -0.0296   1.0000   0.1637
   0.000   0.0862   0.02131   0.00935  -0.0289   1.0000   0.1799
   0.250   0.1105   0.02075   0.00878  -0.0281   1.0000   0.1949
   0.750   0.1698   0.01796   0.00777  -0.0282   1.0000   1.0000
   1.000   0.1907   0.01828   0.00764  -0.0271   1.0000   1.0000
   1.250   0.2107   0.01863   0.00776  -0.0261   1.0000   1.0000
   1.500   0.2304   0.01902   0.00799  -0.0253   1.0000   1.0000
   1.750   0.2496   0.01945   0.00831  -0.0245   1.0000   1.0000
   2.000   0.2684   0.01993   0.00872  -0.0237   1.0000   1.0000
   2.250   0.2868   0.02047   0.00922  -0.0230   1.0000   1.0000
   2.500   0.3046   0.02106   0.00980  -0.0223   1.0000   1.0000
   2.750   0.3219   0.02172   0.01047  -0.0218   1.0000   1.0000
   3.000   0.3387   0.02244   0.01124  -0.0212   1.0000   1.0000
   3.250   0.3549   0.02325   0.01213  -0.0208   1.0000   1.0000
   3.500   0.3706   0.02414   0.01310  -0.0204   1.0000   1.0000
   3.750   0.3986   0.02536   0.01445  -0.0228   0.9940   1.0000
   4.000   0.4662   0.02699   0.01639  -0.0324   0.9649   1.0000
   4.250   0.5459   0.02796   0.01789  -0.0426   0.9262   1.0000
   4.500   0.6297   0.02754   0.01809  -0.0512   0.8781   1.0000
   4.750   0.7417   0.02316   0.01480  -0.0572   0.8018   1.0000
   5.000   0.7888   0.01991   0.01186  -0.0519   0.6806   1.0000
   5.250   0.7939   0.02153   0.01123  -0.0428   0.3287   1.0000
   5.500   0.8041   0.02504   0.01331  -0.0397   0.1905   1.0000
   5.750   0.8358   0.02752   0.01551  -0.0395   0.1598   1.0000
   6.000   0.8709   0.03010   0.01814  -0.0398   0.1422   1.0000
   6.250   0.8995   0.03248   0.02070  -0.0392   0.1277   1.0000
   6.500   0.9283   0.03536   0.02396  -0.0384   0.1214   1.0000
   6.750   0.9514   0.03843   0.02738  -0.0372   0.1143   1.0000
   7.000   0.9711   0.04143   0.03087  -0.0355   0.1084   1.0000
   7.250   0.9885   0.04534   0.03539  -0.0334   0.1085   1.0000
   7.500   1.0005   0.04975   0.04046  -0.0311   0.1106   1.0000
   7.750   1.0089   0.05434   0.04567  -0.0288   0.1130   1.0000
   8.000   1.0146   0.05893   0.05070  -0.0268   0.1143   1.0000
   8.250   1.0179   0.06367   0.05581  -0.0250   0.1153   1.0000
   8.500   1.0161   0.06861   0.06118  -0.0232   0.1192   1.0000
   8.750   1.0018   0.07478   0.06774  -0.0222   0.1281   1.0000
   9.000   0.9674   0.08247   0.07576  -0.0234   0.1503   1.0000
   9.250   0.8511   0.07972   0.07339  -0.0167   0.1482   1.0000
   9.500   0.8163   0.08733   0.08097  -0.0205   0.1532   1.0000
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