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

GOE 423 AIRFOIL (goe423-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: GOE 423 AIRFOIL (goe423-il)
Reynolds number: 100,000
Max Cl/Cd: 40.21 at α=8°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-goe423-il-100000-n5.txt
Download as CSV file: xf-goe423-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 423 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.2348   0.07618   0.07049  -0.1060   0.9277   0.0698
  -9.750  -0.2716   0.06467   0.05887  -0.1159   0.9116   0.0696
  -9.500  -0.3202   0.05670   0.05057  -0.1179   0.8931   0.0696
  -9.250  -0.3505   0.05077   0.04417  -0.1170   0.8774   0.0699
  -9.000  -0.3456   0.04807   0.04129  -0.1162   0.8653   0.0704
  -8.750  -0.3314   0.04573   0.03878  -0.1162   0.8556   0.0709
  -8.500  -0.3247   0.04332   0.03613  -0.1148   0.8417   0.0713
  -8.250  -0.3140   0.04101   0.03354  -0.1138   0.8292   0.0718
  -8.000  -0.2985   0.03876   0.03098  -0.1132   0.8185   0.0724
  -7.750  -0.2839   0.03703   0.02899  -0.1121   0.8050   0.0733
  -7.500  -0.2679   0.03529   0.02696  -0.1110   0.7921   0.0745
  -7.250  -0.2497   0.03347   0.02475  -0.1102   0.7807   0.0757
  -7.000  -0.2319   0.03180   0.02270  -0.1091   0.7677   0.0767
  -6.750  -0.2124   0.03030   0.02083  -0.1081   0.7552   0.0775
  -6.500  -0.1899   0.02894   0.01906  -0.1074   0.7442   0.0783
  -6.250  -0.1667   0.02790   0.01786  -0.1069   0.7327   0.0792
  -6.000  -0.1422   0.02715   0.01701  -0.1066   0.7225   0.0805
  -5.750  -0.1169   0.02643   0.01614  -0.1063   0.7134   0.0821
  -5.500  -0.0921   0.02572   0.01527  -0.1059   0.7050   0.0837
  -5.250  -0.0664   0.02497   0.01432  -0.1056   0.6976   0.0853
  -5.000  -0.0400   0.02426   0.01340  -0.1054   0.6912   0.0867
  -4.750  -0.0144   0.02366   0.01263  -0.1050   0.6845   0.0882
  -4.500   0.0119   0.02311   0.01207  -0.1049   0.6789   0.0900
  -4.250   0.0381   0.02270   0.01161  -0.1048   0.6737   0.0925
  -4.000   0.0634   0.02233   0.01117  -0.1044   0.6680   0.0952
  -3.750   0.0897   0.02193   0.01066  -0.1041   0.6630   0.0980
  -3.500   0.1166   0.02153   0.01020  -0.1040   0.6589   0.1005
  -3.250   0.1414   0.02126   0.00994  -0.1036   0.6543   0.1038
  -3.000   0.1666   0.02103   0.00968  -0.1032   0.6500   0.1085
  -2.750   0.1920   0.02079   0.00944  -0.1029   0.6462   0.1139
  -2.500   0.2184   0.02059   0.00920  -0.1027   0.6429   0.1211
  -2.250   0.2454   0.02036   0.00895  -0.1026   0.6401   0.1298
  -2.000   0.2695   0.02019   0.00884  -0.1020   0.6367   0.1426
  -1.750   0.2936   0.02002   0.00874  -0.1015   0.6332   0.1609
  -1.500   0.3183   0.01984   0.00865  -0.1011   0.6298   0.1855
  -1.250   0.3437   0.01964   0.00860  -0.1009   0.6267   0.2220
  -1.000   0.3697   0.01940   0.00855  -0.1008   0.6238   0.2831
  -0.750   0.3955   0.01898   0.00853  -0.1007   0.6214   0.3873
  -0.500   0.4154   0.01849   0.00878  -0.0992   0.6188   0.5881
  -0.250   0.4525   0.01818   0.00931  -0.0999   0.6160   0.8197
   0.250   0.5636   0.01878   0.00983  -0.1103   0.6106   0.9635
   0.500   0.6207   0.01901   0.00992  -0.1165   0.6081   0.9886
   0.750   0.6673   0.01917   0.00995  -0.1207   0.6057   1.0000
   1.000   0.6900   0.01937   0.01003  -0.1200   0.6035   1.0000
   1.250   0.7127   0.01961   0.01017  -0.1193   0.6014   1.0000
   1.500   0.7292   0.01995   0.01052  -0.1176   0.5988   1.0000
   1.750   0.7466   0.02030   0.01087  -0.1160   0.5963   1.0000
   2.000   0.7648   0.02066   0.01122  -0.1146   0.5939   1.0000
   2.250   0.7838   0.02101   0.01156  -0.1132   0.5915   1.0000
   2.500   0.8039   0.02134   0.01185  -0.1120   0.5891   1.0000
   2.750   0.8258   0.02164   0.01212  -0.1111   0.5868   1.0000
   3.000   0.8492   0.02195   0.01237  -0.1105   0.5848   1.0000
   3.250   0.8736   0.02228   0.01265  -0.1101   0.5832   1.0000
   3.500   0.8907   0.02277   0.01318  -0.1084   0.5810   1.0000
   3.750   0.9012   0.02339   0.01387  -0.1057   0.5779   1.0000
   4.000   0.9146   0.02395   0.01449  -0.1034   0.5748   1.0000
   4.250   0.9310   0.02448   0.01506  -0.1017   0.5721   1.0000
   4.500   0.9504   0.02493   0.01552  -0.1004   0.5695   1.0000
   4.750   0.9731   0.02529   0.01588  -0.0997   0.5671   1.0000
   5.000   0.9996   0.02557   0.01614  -0.0996   0.5649   1.0000
   5.250   1.0196   0.02602   0.01661  -0.0984   0.5619   1.0000
   5.500   1.0174   0.02701   0.01774  -0.0938   0.5564   1.0000
   5.750   1.0320   0.02753   0.01831  -0.0917   0.5521   1.0000
   6.000   1.0560   0.02778   0.01858  -0.0912   0.5488   1.0000
   6.250   1.0861   0.02787   0.01867  -0.0916   0.5461   1.0000
   6.500   1.1042   0.02837   0.01921  -0.0902   0.5427   1.0000
   6.750   1.0895   0.02987   0.02089  -0.0840   0.5364   1.0000
   7.000   1.1049   0.03036   0.02144  -0.0822   0.5321   1.0000
   7.250   1.1426   0.02991   0.02097  -0.0835   0.5283   1.0000
   7.500   1.1498   0.03048   0.02161  -0.0804   0.5225   1.0000
   7.750   1.1329   0.03193   0.02318  -0.0740   0.5154   1.0000
   8.000   1.1976   0.02978   0.02092  -0.0784   0.5088   1.0000
   8.250   1.1603   0.03223   0.02355  -0.0696   0.5014   1.0000
   8.500   1.1826   0.03189   0.02324  -0.0685   0.4941   1.0000
   8.750   1.1944   0.03225   0.02366  -0.0663   0.4875   1.0000
   9.000   1.1734   0.03461   0.02615  -0.0611   0.4789   1.0000
   9.250   1.2205   0.03286   0.02437  -0.0624   0.4734   1.0000
   9.500   1.1604   0.03839   0.03012  -0.0549   0.4618   1.0000
   9.750   1.1368   0.04220   0.03403  -0.0516   0.4508   1.0000
  10.000   1.1549   0.04236   0.03424  -0.0506   0.4431   1.0000
  11.000   1.1306   0.05381   0.04602  -0.0447   0.3947   1.0000
  11.250   1.1306   0.05619   0.04846  -0.0437   0.3803   1.0000
  11.500   1.1382   0.05768   0.04999  -0.0428   0.3661   1.0000
  11.750   1.1539   0.05814   0.05044  -0.0419   0.3507   1.0000
  12.000   1.1764   0.05766   0.04985  -0.0409   0.3312   1.0000
  12.250   1.1887   0.05850   0.05061  -0.0398   0.3122   1.0000
  12.500   1.1956   0.05988   0.05183  -0.0386   0.2879   1.0000
  12.750   1.1946   0.06229   0.05409  -0.0374   0.2623   1.0000
  13.000   1.1892   0.06530   0.05695  -0.0364   0.2350   1.0000
  13.250   1.1812   0.06876   0.06025  -0.0355   0.2043   1.0000
  13.500   1.1683   0.07287   0.06410  -0.0347   0.1611   1.0000
  13.750   1.1478   0.07796   0.06881  -0.0341   0.1222   1.0000
  14.000   1.1369   0.08215   0.07283  -0.0337   0.1009   1.0000
  14.250   1.1310   0.08586   0.07646  -0.0334   0.0855   1.0000
  14.500   1.1275   0.08935   0.07992  -0.0333   0.0764   1.0000
  14.750   1.1254   0.09272   0.08330  -0.0333   0.0714   1.0000
<< Back to GOE 423 AIRFOIL (goe423-il)

Polar data table (+)

Polar graphs


<< Back to GOE 423 AIRFOIL (goe423-il)