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

GOE 367 AIRFOIL (goe367-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: GOE 367 AIRFOIL (goe367-il)
Reynolds number: 100,000
Max Cl/Cd: 41.19 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe367-il-100000.txt
Download as CSV file: xf-goe367-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 367 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.2807   0.10645   0.10170  -0.0546   0.9721   0.1415
  -8.500  -0.2778   0.10269   0.09793  -0.0605   0.9669   0.1474
  -8.250  -0.3545   0.09993   0.09519  -0.0691   0.9501   0.1500
  -8.000  -0.3277   0.09499   0.09031  -0.0681   0.9465   0.1518
  -7.750  -0.2770   0.09181   0.08710  -0.0675   0.9449   0.1550
  -7.500  -0.2641   0.08843   0.08371  -0.0704   0.9394   0.1592
  -7.250  -0.2854   0.08576   0.08103  -0.0710   0.9279   0.1634
  -7.000  -0.2992   0.07978   0.07489  -0.0782   0.9203   0.1694
  -6.750  -0.2993   0.07817   0.07334  -0.0741   0.9103   0.1710
  -6.500  -0.2722   0.07537   0.07052  -0.0757   0.9060   0.1769
  -6.250  -0.3089   0.07270   0.06752  -0.0754   0.8927   0.1865
  -6.000  -0.2723   0.06937   0.06435  -0.0761   0.8894   0.1906
  -5.500  -0.2851   0.05604   0.04921  -0.0772   0.8715   0.1342
  -5.250  -0.2556   0.04870   0.04136  -0.0794   0.8687   0.1186
  -5.000  -0.2546   0.04718   0.03965  -0.0757   0.8578   0.1184
  -4.750  -0.2220   0.04438   0.03641  -0.0771   0.8533   0.1186
  -4.500  -0.1807   0.04158   0.03316  -0.0798   0.8506   0.1188
  -4.250  -0.1782   0.04079   0.03212  -0.0758   0.8387   0.1192
  -4.000  -0.1381   0.03866   0.02992  -0.0784   0.8350   0.1227
  -3.750  -0.1205   0.03809   0.02927  -0.0769   0.8261   0.1257
  -3.500  -0.0881   0.03685   0.02776  -0.0776   0.8198   0.1293
  -3.250  -0.0424   0.03544   0.02590  -0.0801   0.8165   0.1346
  -3.000  -0.0302   0.03492   0.02544  -0.0779   0.8060   0.1386
  -2.750   0.0099   0.03389   0.02429  -0.0797   0.8011   0.1463
  -2.500   0.0594   0.03238   0.02272  -0.0830   0.7986   0.1570
  -2.250   0.0691   0.03247   0.02273  -0.0800   0.7868   0.1646
  -2.000   0.1135   0.03111   0.02143  -0.0824   0.7829   0.1810
  -1.750   0.1605   0.02964   0.02000  -0.0849   0.7805   0.2046
  -1.500   0.1707   0.02970   0.02018  -0.0821   0.7683   0.2196
  -1.250   0.2120   0.02843   0.01910  -0.0837   0.7646   0.2470
  -1.000   0.2557   0.02714   0.01799  -0.0855   0.7621   0.2831
  -0.750   0.4914   0.02385   0.01688  -0.1201   0.7663   1.0000
  -0.500   0.5256   0.02323   0.01602  -0.1206   0.7616   1.0000
  -0.250   0.5344   0.02369   0.01640  -0.1176   0.7493   1.0000
   0.000   0.5679   0.02312   0.01564  -0.1179   0.7443   1.0000
   0.250   0.5778   0.02357   0.01603  -0.1150   0.7325   1.0000
   0.500   0.6103   0.02306   0.01536  -0.1152   0.7271   1.0000
   0.750   0.6217   0.02347   0.01572  -0.1125   0.7159   1.0000
   1.000   0.6530   0.02304   0.01514  -0.1126   0.7098   1.0000
   1.250   0.6661   0.02341   0.01546  -0.1101   0.6995   1.0000
   1.500   0.6950   0.02312   0.01506  -0.1099   0.6928   1.0000
   1.750   0.7111   0.02339   0.01527  -0.1078   0.6834   1.0000
   2.000   0.7371   0.02322   0.01501  -0.1071   0.6757   1.0000
   2.250   0.7570   0.02333   0.01507  -0.1056   0.6672   1.0000
   2.500   0.7795   0.02331   0.01497  -0.1044   0.6587   1.0000
   2.750   0.8034   0.02330   0.01489  -0.1035   0.6511   1.0000
   3.000   0.8210   0.02350   0.01505  -0.1016   0.6419   1.0000
   3.250   0.8536   0.02324   0.01467  -0.1021   0.6361   1.0000
   3.500   0.8610   0.02382   0.01529  -0.0987   0.6256   1.0000
   3.750   0.8982   0.02345   0.01478  -0.0999   0.6203   1.0000
   4.000   0.9005   0.02422   0.01562  -0.0956   0.6096   1.0000
   4.250   0.9361   0.02391   0.01519  -0.0966   0.6039   1.0000
   4.500   0.9402   0.02466   0.01600  -0.0927   0.5941   1.0000
   4.750   0.9717   0.02453   0.01578  -0.0931   0.5879   1.0000
   5.000   0.9818   0.02519   0.01646  -0.0902   0.5799   1.0000
   5.250   1.0028   0.02548   0.01673  -0.0890   0.5730   1.0000
   5.500   1.0423   0.02533   0.01646  -0.0908   0.5685   1.0000
   5.750   1.0328   0.02662   0.01790  -0.0849   0.5596   1.0000
   6.000   1.0616   0.02679   0.01802  -0.0851   0.5544   1.0000
   6.250   1.0829   0.02728   0.01851  -0.0841   0.5488   1.0000
   6.500   1.0842   0.02831   0.01963  -0.0800   0.5412   1.0000
   6.750   1.1232   0.02807   0.01930  -0.0816   0.5358   1.0000
   7.000   1.1255   0.02915   0.02048  -0.0778   0.5287   1.0000
   7.250   1.1438   0.02960   0.02095  -0.0764   0.5219   1.0000
   7.500   1.1948   0.02901   0.02021  -0.0797   0.5167   1.0000
   7.750   1.1749   0.03089   0.02232  -0.0729   0.5093   1.0000
   8.000   1.2004   0.03118   0.02263  -0.0726   0.5036   1.0000
   8.250   1.2502   0.03077   0.02211  -0.0758   0.4992   1.0000
   8.500   1.2216   0.03305   0.02464  -0.0680   0.4921   1.0000
   8.750   1.2488   0.03319   0.02480  -0.0679   0.4862   1.0000
   9.000   1.3014   0.03260   0.02413  -0.0715   0.4816   1.0000
   9.250   1.2623   0.03547   0.02726  -0.0626   0.4754   1.0000
   9.500   1.2686   0.03656   0.02844  -0.0599   0.4702   1.0000
   9.750   1.3294   0.03556   0.02742  -0.0644   0.4659   1.0000
  10.000   1.2904   0.03851   0.03055  -0.0559   0.4600   1.0000
  10.250   1.2651   0.04126   0.03343  -0.0502   0.4533   1.0000
  10.500   1.3762   0.03728   0.02936  -0.0596   0.4483   1.0000
  10.750   1.0760   0.06350   0.05589  -0.0374   0.4311   1.0000
  11.000   1.2016   0.05203   0.04450  -0.0393   0.4330   1.0000
  11.250   1.4184   0.03900   0.03133  -0.0543   0.4306   1.0000
  11.500   1.1074   0.06791   0.06052  -0.0348   0.4133   1.0000
  11.750   1.2111   0.05807   0.05078  -0.0351   0.4143   1.0000
  12.000   1.4232   0.04237   0.03504  -0.0455   0.4118   1.0000
  12.250   0.9419   0.09846   0.09121  -0.0350   0.3682   1.0000
  12.500   0.9865   0.09540   0.08825  -0.0337   0.3652   1.0000
  12.750   1.0407   0.09073   0.08370  -0.0321   0.3637   1.0000
  13.000   1.1015   0.08474   0.07783  -0.0303   0.3631   1.0000
  13.250   1.1647   0.07791   0.07112  -0.0283   0.3622   1.0000
  13.500   1.2490   0.06868   0.06204  -0.0268   0.3607   1.0000
  13.750   1.4747   0.04555   0.03868  -0.0304   0.3451   1.0000
  14.000   1.4254   0.05125   0.04464  -0.0263   0.3373   1.0000
  14.250   1.4301   0.05202   0.04546  -0.0247   0.3243   1.0000
  14.500   1.4275   0.05358   0.04706  -0.0229   0.3102   1.0000
  14.750   1.4133   0.05659   0.05014  -0.0213   0.2951   1.0000
  15.000   1.3887   0.06119   0.05483  -0.0200   0.2789   1.0000
  15.250   1.3583   0.06707   0.06081  -0.0195   0.2606   1.0000
  15.500   1.3444   0.07108   0.06470  -0.0191   0.2359   1.0000
  15.750   1.3255   0.07602   0.06949  -0.0190   0.2118   1.0000
  16.000   1.3177   0.07938   0.07249  -0.0185   0.1913   1.0000
  16.250   1.3120   0.08253   0.07538  -0.0179   0.1755   1.0000
  16.500   1.3133   0.08472   0.07734  -0.0170   0.1621   1.0000
  16.750   1.3202   0.08621   0.07861  -0.0159   0.1502   1.0000
  17.000   1.3165   0.08965   0.08213  -0.0159   0.1412   1.0000
  17.250   1.3185   0.09220   0.08466  -0.0157   0.1327   1.0000
  17.500   1.3278   0.09356   0.08583  -0.0150   0.1243   1.0000
  17.750   1.3210   0.09772   0.09022  -0.0157   0.1182   1.0000
  18.000   1.3308   0.09908   0.09141  -0.0152   0.1115   1.0000
  18.250   1.3282   0.10263   0.09514  -0.0157   0.1063   1.0000
  18.500   1.3451   0.10290   0.09519  -0.0147   0.1002   1.0000
<< Back to GOE 367 AIRFOIL (goe367-il)

Polar data table (+)

Polar graphs


<< Back to GOE 367 AIRFOIL (goe367-il)