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

GOE 510 AIRFOIL (goe510-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: GOE 510 AIRFOIL (goe510-il)
Reynolds number: 1,000,000
Max Cl/Cd: 106.86 at α=6.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe510-il-1000000.txt
Download as CSV file: xf-goe510-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 510 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.0442   0.08892   0.08706  -0.1031   0.9719   0.0323
  -9.750  -0.0296   0.08565   0.08378  -0.1065   0.9710   0.0329
  -9.000  -0.0327   0.06822   0.06635  -0.1218   0.9602   0.0358
  -8.750  -0.0170   0.06491   0.06304  -0.1261   0.9579   0.0359
  -8.500  -0.2108   0.02177   0.01822  -0.1422   0.9230   0.0346
  -8.250  -0.2005   0.02013   0.01631  -0.1396   0.9141   0.0349
  -8.000  -0.1783   0.01928   0.01526  -0.1389   0.9074   0.0352
  -7.750  -0.1691   0.01712   0.01277  -0.1362   0.8975   0.0357
  -7.500  -0.1495   0.01619   0.01172  -0.1349   0.8886   0.0361
  -7.250  -0.1237   0.01566   0.01110  -0.1347   0.8803   0.0364
  -7.000  -0.1014   0.01533   0.01071  -0.1336   0.8700   0.0367
  -6.750  -0.0760   0.01511   0.01041  -0.1330   0.8592   0.0372
  -6.500  -0.0524   0.01477   0.00998  -0.1322   0.8479   0.0376
  -6.250  -0.0307   0.01435   0.00947  -0.1309   0.8379   0.0380
  -6.000  -0.0079   0.01391   0.00890  -0.1299   0.8281   0.0385
  -5.750   0.0135   0.01342   0.00829  -0.1285   0.8178   0.0389
  -5.500   0.0358   0.01305   0.00781  -0.1272   0.8079   0.0394
  -5.000   0.0798   0.01247   0.00700  -0.1245   0.7855   0.0403
  -4.750   0.1020   0.01232   0.00676  -0.1232   0.7741   0.0406
  -4.500   0.1220   0.01156   0.00586  -0.1216   0.7624   0.0412
  -4.250   0.1431   0.01116   0.00541  -0.1201   0.7497   0.0418
  -4.000   0.1643   0.01099   0.00519  -0.1186   0.7361   0.0424
  -3.750   0.1849   0.01088   0.00503  -0.1169   0.7214   0.0430
  -3.500   0.2048   0.01078   0.00484  -0.1151   0.7054   0.0436
  -3.250   0.2245   0.01063   0.00461  -0.1133   0.6902   0.0441
  -3.000   0.2442   0.01050   0.00440  -0.1114   0.6759   0.0447
  -2.750   0.2643   0.01040   0.00422  -0.1096   0.6621   0.0454
  -2.500   0.2850   0.01030   0.00406  -0.1080   0.6490   0.0460
  -2.250   0.3057   0.01024   0.00393  -0.1063   0.6371   0.0464
  -2.000   0.3253   0.01023   0.00384  -0.1044   0.6254   0.0468
  -1.750   0.3435   0.00983   0.00340  -0.1023   0.6147   0.0478
  -1.500   0.3617   0.00968   0.00322  -0.1001   0.6041   0.0488
  -1.250   0.3799   0.00963   0.00313  -0.0980   0.5922   0.0497
  -1.000   0.3999   0.00956   0.00302  -0.0962   0.5799   0.0506
  -0.750   0.4194   0.00952   0.00294  -0.0943   0.5673   0.0516
  -0.500   0.4380   0.00950   0.00286  -0.0922   0.5523   0.0525
  -0.250   0.4561   0.00951   0.00280  -0.0901   0.5356   0.0531
   0.000   0.4739   0.00958   0.00278  -0.0879   0.5157   0.0538
   0.250   0.4909   0.00961   0.00271  -0.0855   0.4955   0.0545
   0.500   0.5080   0.00961   0.00263  -0.0832   0.4777   0.0562
   0.750   0.5268   0.00968   0.00263  -0.0813   0.4631   0.0578
   1.000   0.5460   0.00976   0.00266  -0.0794   0.4511   0.0594
   1.250   0.5655   0.00984   0.00269  -0.0777   0.4409   0.0608
   1.500   0.5863   0.00990   0.00271  -0.0762   0.4335   0.0621
   1.750   0.6066   0.00994   0.00272  -0.0745   0.4265   0.0646
   2.000   0.6265   0.01001   0.00277  -0.0729   0.4201   0.0683
   2.250   0.6485   0.01003   0.00280  -0.0716   0.4155   0.0725
   2.500   0.6654   0.00989   0.00288  -0.0694   0.4103   0.1560
   2.750   0.6827   0.00991   0.00304  -0.0673   0.4044   0.2197
   3.000   0.7046   0.00995   0.00314  -0.0661   0.4008   0.2447
   3.250   0.7265   0.01001   0.00324  -0.0649   0.3968   0.2623
   3.500   0.7476   0.01012   0.00335  -0.0635   0.3923   0.2744
   3.750   0.7674   0.01026   0.00348  -0.0620   0.3875   0.2880
   4.000   0.7899   0.01032   0.00358  -0.0609   0.3840   0.2990
   4.250   0.8120   0.01040   0.00368  -0.0598   0.3801   0.3089
   4.500   0.8326   0.01050   0.00380  -0.0584   0.3763   0.3222
   4.750   0.8519   0.01063   0.00394  -0.0568   0.3720   0.3371
   5.000   0.8721   0.01073   0.00409  -0.0554   0.3684   0.3577
   5.250   0.8892   0.01057   0.00422  -0.0533   0.3654   0.4780
   5.500   1.0317   0.01014   0.00496  -0.0791   0.3557   0.9872
   5.750   1.0704   0.01043   0.00522  -0.0816   0.3512   0.9938
   6.000   1.1131   0.01062   0.00541  -0.0851   0.3468   0.9973
   6.250   1.1533   0.01087   0.00561  -0.0882   0.3398   0.9993
   6.500   1.1816   0.01106   0.00578  -0.0886   0.3337   1.0000
   6.750   1.1958   0.01119   0.00590  -0.0860   0.3278   1.0000
   7.000   1.2076   0.01139   0.00607  -0.0830   0.3223   1.0000
   7.250   1.2220   0.01156   0.00623  -0.0805   0.3163   1.0000
   7.500   1.2357   0.01179   0.00641  -0.0779   0.3074   1.0000
   7.750   1.2503   0.01202   0.00662  -0.0755   0.2991   1.0000
   8.000   1.2637   0.01232   0.00687  -0.0730   0.2901   1.0000
   8.250   1.2798   0.01256   0.00710  -0.0710   0.2836   1.0000
   8.500   1.2936   0.01293   0.00741  -0.0686   0.2735   1.0000
   8.750   1.3095   0.01323   0.00770  -0.0667   0.2654   1.0000
   9.000   1.3232   0.01365   0.00807  -0.0644   0.2558   1.0000
   9.250   1.3376   0.01408   0.00844  -0.0624   0.2447   1.0000
   9.500   1.3522   0.01452   0.00885  -0.0604   0.2350   1.0000
   9.750   1.3650   0.01506   0.00933  -0.0582   0.2239   1.0000
  10.000   1.3769   0.01567   0.00987  -0.0559   0.2113   1.0000
  10.250   1.3870   0.01638   0.01051  -0.0534   0.1960   1.0000
  10.500   1.3918   0.01740   0.01137  -0.0503   0.1715   1.0000
  10.750   1.3894   0.01883   0.01259  -0.0464   0.1404   1.0000
  11.000   1.3784   0.02083   0.01434  -0.0417   0.1044   1.0000
  11.250   1.3639   0.02318   0.01647  -0.0369   0.0690   1.0000
  11.500   1.3430   0.02617   0.01929  -0.0320   0.0266   1.0000
  11.750   1.3470   0.02766   0.02081  -0.0299   0.0210   1.0000
  12.000   1.3545   0.02897   0.02215  -0.0282   0.0194   1.0000
  12.250   1.3634   0.03021   0.02344  -0.0269   0.0186   1.0000
  12.500   1.3702   0.03167   0.02495  -0.0254   0.0177   1.0000
  12.750   1.3754   0.03329   0.02662  -0.0240   0.0169   1.0000
  13.000   1.3792   0.03510   0.02850  -0.0226   0.0163   1.0000
  13.250   1.3855   0.03676   0.03021  -0.0215   0.0159   1.0000
  13.500   1.3910   0.03855   0.03206  -0.0205   0.0155   1.0000
  13.750   1.3961   0.04042   0.03399  -0.0196   0.0151   1.0000
  14.000   1.3988   0.04257   0.03621  -0.0187   0.0146   1.0000
  14.250   1.4007   0.04487   0.03857  -0.0179   0.0143   1.0000
  14.500   1.4002   0.04749   0.04127  -0.0172   0.0140   1.0000
  14.750   1.3995   0.05021   0.04407  -0.0166   0.0138   1.0000
  15.000   1.3889   0.05409   0.04806  -0.0160   0.0134   1.0000
  15.250   1.3891   0.05685   0.05090  -0.0158   0.0132   1.0000
  15.500   1.3874   0.05989   0.05401  -0.0156   0.0131   1.0000
  15.750   1.3848   0.06311   0.05731  -0.0156   0.0129   1.0000
  16.000   1.3803   0.06660   0.06090  -0.0156   0.0128   1.0000
  16.250   1.3746   0.07033   0.06471  -0.0158   0.0126   1.0000
  16.500   1.3689   0.07412   0.06860  -0.0161   0.0124   1.0000
  16.750   1.3615   0.07822   0.07278  -0.0166   0.0123   1.0000
  17.000   1.3549   0.08223   0.07688  -0.0171   0.0121   1.0000
  17.250   1.3477   0.08639   0.08112  -0.0178   0.0118   1.0000
  17.500   1.3393   0.09076   0.08557  -0.0186   0.0117   1.0000
  17.750   1.3293   0.09541   0.09032  -0.0195   0.0117   1.0000
  18.000   1.3204   0.09989   0.09487  -0.0204   0.0114   1.0000
  18.250   1.3092   0.10479   0.09986  -0.0215   0.0114   1.0000
  18.500   1.2987   0.10963   0.10479  -0.0227   0.0112   1.0000
  18.750   1.2860   0.11488   0.11013  -0.0242   0.0111   1.0000
  19.000   1.2687   0.12088   0.11623  -0.0259   0.0109   1.0000
<< Back to GOE 510 AIRFOIL (goe510-il)

Polar data table (+)

Polar graphs


<< Back to GOE 510 AIRFOIL (goe510-il)