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

GOE 446 AIRFOIL (goe446-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: GOE 446 AIRFOIL (goe446-il)
Reynolds number: 500,000
Max Cl/Cd: 121.96 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-goe446-il-500000.txt
Download as CSV file: xf-goe446-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: GOE 446 AIRFOIL                                 
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.0327   0.08631   0.08371  -0.1028   0.9234   0.0369
  -8.750  -0.0472   0.08069   0.07806  -0.1100   0.9041   0.0390
  -8.500  -0.0430   0.07722   0.07455  -0.1100   0.8927   0.0393
  -8.250  -0.0324   0.07521   0.07250  -0.1096   0.8808   0.0396
  -8.000  -0.0209   0.07330   0.07054  -0.1096   0.8693   0.0400
  -7.750  -0.0103   0.07124   0.06843  -0.1101   0.8582   0.0404
  -7.250   0.0089   0.06640   0.06348  -0.1139   0.8370   0.0420
  -7.000   0.0189   0.05706   0.05400  -0.1308   0.8263   0.0456
  -6.750   0.0350   0.05540   0.05230  -0.1311   0.8189   0.0460
  -6.500   0.0281   0.02586   0.02142  -0.1560   0.8109   0.0393
  -6.250   0.0487   0.02288   0.01818  -0.1566   0.8044   0.0384
  -5.750   0.0946   0.01796   0.01238  -0.1573   0.7922   0.0382
  -5.500   0.1204   0.01665   0.01076  -0.1574   0.7863   0.0382
  -5.250   0.1472   0.01570   0.00955  -0.1574   0.7811   0.0384
  -5.000   0.1741   0.01492   0.00860  -0.1574   0.7754   0.0386
  -4.750   0.2014   0.01431   0.00783  -0.1574   0.7699   0.0388
  -4.500   0.2293   0.01383   0.00718  -0.1574   0.7646   0.0390
  -4.250   0.2561   0.01302   0.00628  -0.1573   0.7591   0.0395
  -4.000   0.2835   0.01243   0.00562  -0.1574   0.7544   0.0401
  -3.750   0.3115   0.01207   0.00521  -0.1575   0.7504   0.0410
  -3.500   0.3396   0.01179   0.00491  -0.1576   0.7463   0.0421
  -3.250   0.3674   0.01150   0.00459  -0.1576   0.7417   0.0430
  -3.000   0.3955   0.01125   0.00429  -0.1576   0.7373   0.0439
  -2.750   0.4240   0.01107   0.00403  -0.1577   0.7331   0.0449
  -2.500   0.4520   0.01078   0.00375  -0.1578   0.7290   0.0466
  -2.250   0.4800   0.01060   0.00358  -0.1578   0.7247   0.0488
  -2.000   0.5083   0.01046   0.00342  -0.1579   0.7205   0.0517
  -1.750   0.5370   0.01034   0.00329  -0.1580   0.7165   0.0591
  -1.500   0.5651   0.01024   0.00331  -0.1581   0.7121   0.0811
  -1.250   0.5931   0.01025   0.00337  -0.1581   0.7073   0.0962
  -1.000   0.6214   0.01032   0.00339  -0.1582   0.7027   0.1037
  -0.750   0.6499   0.01029   0.00335  -0.1584   0.6983   0.1107
  -0.500   0.6775   0.01028   0.00336  -0.1583   0.6936   0.1166
  -0.250   0.7053   0.01023   0.00332  -0.1583   0.6887   0.1225
   0.000   0.7335   0.01021   0.00328  -0.1584   0.6841   0.1291
   0.250   0.7613   0.01021   0.00328  -0.1585   0.6793   0.1346
   0.500   0.7887   0.01012   0.00324  -0.1584   0.6737   0.1424
   0.750   0.8164   0.01011   0.00321  -0.1584   0.6684   0.1515
   1.000   0.8439   0.01005   0.00321  -0.1584   0.6628   0.1660
   1.250   0.8710   0.00995   0.00321  -0.1583   0.6564   0.1909
   1.500   0.8981   0.00957   0.00327  -0.1585   0.6506   0.3724
   1.750   0.9195   0.00834   0.00346  -0.1571   0.6445   1.0000
   2.000   0.9462   0.00842   0.00347  -0.1569   0.6375   1.0000
   2.250   0.9729   0.00852   0.00350  -0.1566   0.6308   1.0000
   2.500   0.9991   0.00860   0.00355  -0.1563   0.6226   1.0000
   2.750   1.0252   0.00872   0.00360  -0.1560   0.6147   1.0000
   3.000   1.0508   0.00882   0.00367  -0.1556   0.6050   1.0000
   3.250   1.0763   0.00895   0.00375  -0.1552   0.5956   1.0000
   3.500   1.1012   0.00910   0.00383  -0.1546   0.5855   1.0000
   3.750   1.1260   0.00924   0.00395  -0.1541   0.5743   1.0000
   4.000   1.1501   0.00943   0.00408  -0.1534   0.5623   1.0000
   4.250   1.1733   0.00964   0.00422  -0.1526   0.5494   1.0000
   4.500   1.1957   0.00989   0.00438  -0.1516   0.5359   1.0000
   4.750   1.2186   0.01013   0.00457  -0.1508   0.5231   1.0000
   5.000   1.2414   0.01037   0.00477  -0.1499   0.5117   1.0000
   5.250   1.2636   0.01065   0.00499  -0.1490   0.5013   1.0000
   5.500   1.2851   0.01094   0.00523  -0.1479   0.4903   1.0000
   5.750   1.3064   0.01122   0.00547  -0.1468   0.4786   1.0000
   6.000   1.3267   0.01153   0.00573  -0.1455   0.4678   1.0000
   6.250   1.3470   0.01183   0.00599  -0.1442   0.4594   1.0000
   6.500   1.3674   0.01211   0.00627  -0.1429   0.4528   1.0000
   6.750   1.3878   0.01239   0.00655  -0.1417   0.4455   1.0000
   7.000   1.4066   0.01274   0.00687  -0.1402   0.4375   1.0000
   7.250   1.4255   0.01307   0.00720  -0.1387   0.4281   1.0000
   7.500   1.4442   0.01343   0.00756  -0.1373   0.4202   1.0000
   7.750   1.4634   0.01377   0.00791  -0.1359   0.4121   1.0000
   8.000   1.4814   0.01417   0.00831  -0.1344   0.4038   1.0000
   8.250   1.4991   0.01457   0.00872  -0.1329   0.3945   1.0000
   8.500   1.5172   0.01498   0.00915  -0.1314   0.3859   1.0000
   8.750   1.5326   0.01551   0.00965  -0.1296   0.3753   1.0000
   9.000   1.5499   0.01596   0.01014  -0.1281   0.3634   1.0000
   9.250   1.5649   0.01654   0.01071  -0.1263   0.3498   1.0000
   9.500   1.5767   0.01729   0.01141  -0.1241   0.3309   1.0000
   9.750   1.5827   0.01838   0.01237  -0.1211   0.3012   1.0000
  10.000   1.5806   0.02002   0.01379  -0.1173   0.2627   1.0000
  10.250   1.5753   0.02199   0.01553  -0.1132   0.2255   1.0000
  10.500   1.5723   0.02394   0.01732  -0.1097   0.1961   1.0000
  10.750   1.5719   0.02581   0.01907  -0.1068   0.1728   1.0000
  11.000   1.5734   0.02764   0.02081  -0.1043   0.1538   1.0000
  11.250   1.5756   0.02947   0.02258  -0.1019   0.1385   1.0000
  11.500   1.5780   0.03134   0.02441  -0.0998   0.1250   1.0000
  11.750   1.5785   0.03346   0.02645  -0.0976   0.1106   1.0000
  12.000   1.5783   0.03570   0.02863  -0.0955   0.0952   1.0000
  12.250   1.5721   0.03855   0.03133  -0.0932   0.0696   1.0000
  12.500   1.5519   0.04282   0.03538  -0.0901   0.0384   1.0000
  12.750   1.5480   0.04568   0.03827  -0.0883   0.0330   1.0000
  13.000   1.5465   0.04842   0.04106  -0.0868   0.0303   1.0000
  13.250   1.5495   0.05075   0.04349  -0.0856   0.0289   1.0000
  13.500   1.5500   0.05339   0.04621  -0.0845   0.0277   1.0000
  13.750   1.5478   0.05637   0.04927  -0.0833   0.0266   1.0000
  14.000   1.5421   0.05982   0.05280  -0.0821   0.0256   1.0000
  14.250   1.5443   0.06246   0.05554  -0.0814   0.0250   1.0000
  14.500   1.5447   0.06534   0.05851  -0.0807   0.0243   1.0000
  14.750   1.5433   0.06847   0.06173  -0.0800   0.0236   1.0000
  15.000   1.5407   0.07180   0.06514  -0.0794   0.0230   1.0000
  15.250   1.5361   0.07546   0.06888  -0.0790   0.0224   1.0000
  15.500   1.5279   0.07964   0.07316  -0.0786   0.0219   1.0000
  15.750   1.5167   0.08427   0.07788  -0.0784   0.0215   1.0000
  16.000   1.5163   0.08754   0.08124  -0.0783   0.0212   1.0000
<< Back to GOE 446 AIRFOIL (goe446-il)

Polar data table (+)

Polar graphs


<< Back to GOE 446 AIRFOIL (goe446-il)