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

GOE 685 AIRFOIL (goe685-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 685 AIRFOIL (goe685-il)
Reynolds number: 1,000,000
Max Cl/Cd: 119.21 at α=5.25°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe685-il-1000000.txt
Download as CSV file: xf-goe685-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 685 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -17.500  -0.7724   0.11251   0.11014  -0.0446   1.0000   0.0226
 -17.250  -0.7971   0.10407   0.10157  -0.0496   1.0000   0.0228
 -17.000  -0.8157   0.09698   0.09437  -0.0539   1.0000   0.0229
 -16.750  -0.8390   0.08982   0.08710  -0.0581   1.0000   0.0232
 -16.500  -0.8482   0.08499   0.08222  -0.0608   1.0000   0.0234
 -16.250  -0.8525   0.08097   0.07817  -0.0631   1.0000   0.0237
 -16.000  -0.8588   0.07663   0.07378  -0.0656   1.0000   0.0238
 -15.750  -0.8651   0.07243   0.06954  -0.0680   1.0000   0.0240
 -15.500  -0.8713   0.06828   0.06534  -0.0705   1.0000   0.0241
 -15.250  -0.8809   0.06390   0.06091  -0.0731   1.0000   0.0243
 -15.000  -0.8913   0.05969   0.05665  -0.0756   1.0000   0.0244
 -14.750  -0.9023   0.05540   0.05232  -0.0783   1.0000   0.0246
 -14.500  -0.9169   0.05089   0.04774  -0.0810   1.0000   0.0248
 -14.250  -0.9352   0.04619   0.04299  -0.0836   1.0000   0.0249
 -14.000  -0.9656   0.04093   0.03766  -0.0859   1.0000   0.0249
 -13.750  -1.0321   0.03357   0.03019  -0.0871   1.0000   0.0245
 -13.500  -1.0081   0.02954   0.02598  -0.0939   0.9984   0.0249
 -13.250  -0.9807   0.02746   0.02378  -0.0978   0.9959   0.0254
 -13.000  -0.9525   0.02579   0.02200  -0.1008   0.9936   0.0258
 -12.750  -0.9268   0.02449   0.02059  -0.1026   0.9906   0.0262
 -12.500  -0.8987   0.02344   0.01945  -0.1043   0.9879   0.0265
 -12.250  -0.8679   0.02267   0.01858  -0.1060   0.9859   0.0268
 -12.000  -0.8430   0.02064   0.01644  -0.1082   0.9835   0.0275
 -11.750  -0.8128   0.01955   0.01529  -0.1101   0.9822   0.0280
 -11.500  -0.7858   0.01873   0.01444  -0.1109   0.9789   0.0285
 -11.250  -0.7562   0.01799   0.01364  -0.1120   0.9761   0.0289
 -11.000  -0.7258   0.01730   0.01289  -0.1132   0.9735   0.0294
 -10.750  -0.6946   0.01670   0.01223  -0.1144   0.9712   0.0301
 -10.500  -0.6625   0.01609   0.01155  -0.1158   0.9689   0.0306
 -10.250  -0.6369   0.01560   0.01099  -0.1156   0.9636   0.0309
 -10.000  -0.6096   0.01513   0.01046  -0.1158   0.9588   0.0312
  -9.750  -0.5833   0.01424   0.00948  -0.1161   0.9542   0.0318
  -9.500  -0.5609   0.01359   0.00878  -0.1154   0.9479   0.0326
  -9.250  -0.5358   0.01312   0.00826  -0.1150   0.9422   0.0332
  -9.000  -0.5094   0.01271   0.00779  -0.1148   0.9366   0.0339
  -8.750  -0.4854   0.01233   0.00736  -0.1141   0.9294   0.0344
  -8.500  -0.4595   0.01198   0.00693  -0.1137   0.9224   0.0351
  -8.250  -0.4342   0.01165   0.00655  -0.1131   0.9151   0.0356
  -8.000  -0.4083   0.01136   0.00618  -0.1126   0.9075   0.0361
  -7.750  -0.3831   0.01094   0.00568  -0.1120   0.8999   0.0370
  -7.500  -0.3578   0.01054   0.00524  -0.1115   0.8912   0.0383
  -7.250  -0.3315   0.01025   0.00489  -0.1111   0.8835   0.0394
  -7.000  -0.3050   0.00998   0.00458  -0.1107   0.8748   0.0406
  -6.750  -0.2781   0.00977   0.00429  -0.1103   0.8670   0.0417
  -6.500  -0.2517   0.00945   0.00393  -0.1099   0.8578   0.0439
  -6.250  -0.2251   0.00919   0.00364  -0.1095   0.8484   0.0466
  -6.000  -0.1986   0.00892   0.00334  -0.1091   0.8369   0.0515
  -5.750  -0.1724   0.00858   0.00301  -0.1086   0.8242   0.0640
  -5.500  -0.1471   0.00811   0.00265  -0.1082   0.8109   0.0944
  -5.250  -0.1210   0.00785   0.00245  -0.1077   0.7968   0.1200
  -5.000  -0.0942   0.00776   0.00233  -0.1073   0.7819   0.1326
  -4.750  -0.0674   0.00769   0.00222  -0.1069   0.7669   0.1421
  -4.500  -0.0401   0.00765   0.00213  -0.1066   0.7530   0.1482
  -4.250  -0.0128   0.00759   0.00205  -0.1063   0.7410   0.1549
  -4.000   0.0147   0.00759   0.00198  -0.1059   0.7302   0.1595
  -3.750   0.0420   0.00753   0.00189  -0.1056   0.7195   0.1658
  -3.500   0.0696   0.00750   0.00183  -0.1054   0.7094   0.1714
  -3.000   0.1244   0.00745   0.00172  -0.1048   0.6869   0.1832
  -2.750   0.1518   0.00744   0.00167  -0.1045   0.6759   0.1889
  -2.500   0.1791   0.00743   0.00162  -0.1041   0.6645   0.1960
  -2.250   0.2068   0.00740   0.00160  -0.1039   0.6533   0.2037
  -2.000   0.2341   0.00739   0.00157  -0.1036   0.6424   0.2123
  -1.750   0.2615   0.00740   0.00155  -0.1033   0.6303   0.2198
  -1.500   0.2891   0.00739   0.00153  -0.1031   0.6187   0.2267
  -1.250   0.3165   0.00741   0.00152  -0.1027   0.6063   0.2342
  -1.000   0.3437   0.00746   0.00151  -0.1024   0.5928   0.2405
  -0.750   0.3706   0.00749   0.00152  -0.1020   0.5777   0.2478
  -0.500   0.3975   0.00757   0.00154  -0.1016   0.5602   0.2536
  -0.250   0.4241   0.00764   0.00155  -0.1012   0.5410   0.2595
   0.000   0.4508   0.00773   0.00158  -0.1008   0.5236   0.2656
   0.250   0.4775   0.00783   0.00162  -0.1004   0.5071   0.2706
   0.500   0.5041   0.00793   0.00166  -0.1000   0.4914   0.2756
   1.000   0.5576   0.00811   0.00177  -0.0993   0.4647   0.2868
   1.250   0.5845   0.00821   0.00182  -0.0989   0.4541   0.2910
   1.750   0.6381   0.00835   0.00195  -0.0982   0.4359   0.3019
   2.000   0.6648   0.00846   0.00202  -0.0978   0.4277   0.3061
   2.250   0.6917   0.00852   0.00208  -0.0975   0.4200   0.3114
   2.500   0.7182   0.00861   0.00216  -0.0971   0.4115   0.3174
   2.750   0.7451   0.00869   0.00224  -0.0968   0.4047   0.3229
   3.000   0.7719   0.00876   0.00231  -0.0964   0.3981   0.3290
   3.250   0.7981   0.00887   0.00241  -0.0960   0.3914   0.3355
   3.500   0.8253   0.00892   0.00249  -0.0957   0.3858   0.3413
   3.750   0.8515   0.00900   0.00259  -0.0953   0.3791   0.3501
   4.250   0.9046   0.00914   0.00279  -0.0946   0.3681   0.3722
   4.500   0.9298   0.00926   0.00291  -0.0940   0.3584   0.3875
   4.750   0.9559   0.00932   0.00302  -0.0936   0.3486   0.4089
   5.000   0.9808   0.00938   0.00316  -0.0930   0.3376   0.4495
   5.250   1.0192   0.00855   0.00349  -0.0956   0.3244   1.0000
   5.500   1.0432   0.00878   0.00363  -0.0948   0.3089   1.0000
   5.750   1.0665   0.00905   0.00381  -0.0939   0.2889   1.0000
   6.000   1.0879   0.00944   0.00405  -0.0927   0.2631   1.0000
   6.250   1.1060   0.01004   0.00441  -0.0909   0.2206   1.0000
   6.500   1.1233   0.01068   0.00484  -0.0891   0.1862   1.0000
   6.750   1.1443   0.01107   0.00515  -0.0879   0.1733   1.0000
   7.000   1.1665   0.01137   0.00542  -0.0868   0.1664   1.0000
   7.250   1.1880   0.01170   0.00571  -0.0856   0.1588   1.0000
   7.500   1.2110   0.01192   0.00594  -0.0847   0.1547   1.0000
   7.750   1.2322   0.01224   0.00624  -0.0835   0.1482   1.0000
   8.000   1.2537   0.01253   0.00651  -0.0824   0.1410   1.0000
   8.250   1.2732   0.01291   0.00682  -0.0809   0.1284   1.0000
   8.500   1.2779   0.01393   0.00753  -0.0770   0.0793   1.0000
   8.750   1.2905   0.01447   0.00805  -0.0744   0.0700   1.0000
   9.000   1.3032   0.01501   0.00857  -0.0718   0.0630   1.0000
   9.250   1.3175   0.01550   0.00904  -0.0695   0.0564   1.0000
   9.500   1.3315   0.01602   0.00954  -0.0673   0.0499   1.0000
   9.750   1.3451   0.01657   0.01007  -0.0650   0.0451   1.0000
  10.000   1.3580   0.01717   0.01065  -0.0628   0.0412   1.0000
  10.250   1.3721   0.01773   0.01122  -0.0608   0.0390   1.0000
  10.500   1.3857   0.01833   0.01184  -0.0589   0.0372   1.0000
  10.750   1.3976   0.01904   0.01256  -0.0567   0.0355   1.0000
  11.000   1.4087   0.01982   0.01337  -0.0546   0.0341   1.0000
  11.250   1.4222   0.02049   0.01409  -0.0529   0.0333   1.0000
  11.500   1.4346   0.02125   0.01488  -0.0511   0.0323   1.0000
  11.750   1.4454   0.02214   0.01579  -0.0493   0.0312   1.0000
  12.000   1.4539   0.02321   0.01690  -0.0473   0.0303   1.0000
  12.250   1.4591   0.02454   0.01829  -0.0452   0.0292   1.0000
  12.500   1.4703   0.02554   0.01934  -0.0438   0.0288   1.0000
  12.750   1.4807   0.02662   0.02048  -0.0424   0.0282   1.0000
  13.000   1.4899   0.02784   0.02174  -0.0410   0.0274   1.0000
  13.250   1.4979   0.02921   0.02316  -0.0397   0.0268   1.0000
  13.500   1.5031   0.03087   0.02487  -0.0384   0.0262   1.0000
  13.750   1.5055   0.03285   0.02690  -0.0372   0.0255   1.0000
  14.000   1.5021   0.03548   0.02961  -0.0359   0.0249   1.0000
  14.250   1.5101   0.03711   0.03132  -0.0352   0.0246   1.0000
  14.500   1.5152   0.03909   0.03336  -0.0346   0.0242   1.0000
  14.750   1.5198   0.04118   0.03553  -0.0340   0.0238   1.0000
  15.000   1.5234   0.04343   0.03784  -0.0335   0.0233   1.0000
  15.250   1.5269   0.04573   0.04020  -0.0332   0.0228   1.0000
  15.500   1.5272   0.04847   0.04301  -0.0330   0.0224   1.0000
  15.750   1.5256   0.05150   0.04610  -0.0329   0.0220   1.0000
  16.000   1.5213   0.05492   0.04960  -0.0329   0.0216   1.0000
  16.250   1.5100   0.05933   0.05410  -0.0333   0.0211   1.0000
  16.500   1.5061   0.06294   0.05779  -0.0337   0.0209   1.0000
  16.750   1.5066   0.06604   0.06098  -0.0341   0.0207   1.0000
  17.000   1.5046   0.06951   0.06454  -0.0347   0.0204   1.0000
  17.250   1.5029   0.07300   0.06811  -0.0353   0.0201   1.0000
  17.500   1.5002   0.07671   0.07189  -0.0361   0.0197   1.0000
  17.750   1.4974   0.08045   0.07571  -0.0369   0.0194   1.0000
  18.000   1.4940   0.08431   0.07964  -0.0379   0.0191   1.0000
  18.250   1.4904   0.08828   0.08369  -0.0390   0.0188   1.0000
  18.500   1.4853   0.09255   0.08801  -0.0403   0.0185   1.0000
  18.750   1.4786   0.09710   0.09263  -0.0417   0.0182   1.0000
  19.000   1.4699   0.10199   0.09759  -0.0434   0.0179   1.0000
<< Back to GOE 685 AIRFOIL (goe685-il)

Polar data table (+)

Polar graphs


<< Back to GOE 685 AIRFOIL (goe685-il)