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

WORTMANN FX 082-512 AIRFOIL (fx082512-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: WORTMANN FX 082-512 AIRFOIL (fx082512-il)
Reynolds number: 50,000
Max Cl/Cd: 33.45 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx082512-il-50000-n5.txt
Download as CSV file: xf-fx082512-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: WORTMANN FX 082-512 AIRFOIL                     
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.000  -0.5376   0.10218   0.09477  -0.0503   1.0000   0.0687
 -10.750  -0.5620   0.09417   0.08682  -0.0538   1.0000   0.0686
 -10.500  -0.5897   0.08655   0.07924  -0.0571   1.0000   0.0682
 -10.250  -0.6156   0.08018   0.07287  -0.0596   1.0000   0.0677
 -10.000  -0.6433   0.07454   0.06721  -0.0614   1.0000   0.0674
  -9.750  -0.6691   0.07003   0.06268  -0.0619   1.0000   0.0671
  -9.500  -0.6944   0.06573   0.05833  -0.0627   1.0000   0.0672
  -9.250  -0.7108   0.06065   0.05307  -0.0661   1.0000   0.0677
  -9.000  -0.7190   0.05514   0.04718  -0.0717   1.0000   0.0693
  -8.750  -0.7128   0.05209   0.04400  -0.0725   1.0000   0.0732
  -8.500  -0.7038   0.04888   0.04050  -0.0741   1.0000   0.0783
  -8.250  -0.6912   0.04635   0.03778  -0.0742   1.0000   0.0835
  -8.000  -0.6768   0.04397   0.03519  -0.0749   1.0000   0.0908
  -7.750  -0.6609   0.04205   0.03310  -0.0746   1.0000   0.0976
  -7.500  -0.6449   0.04045   0.03144  -0.0740   1.0000   0.1055
  -7.250  -0.6273   0.03898   0.02988  -0.0736   1.0000   0.1128
  -7.000  -0.6117   0.03817   0.02916  -0.0715   1.0000   0.1175
  -6.750  -0.5919   0.03686   0.02779  -0.0717   1.0000   0.1247
  -6.500  -0.5691   0.03555   0.02638  -0.0728   1.0000   0.1320
  -6.250  -0.5456   0.03444   0.02536  -0.0736   1.0000   0.1377
  -6.000  -0.5169   0.03305   0.02398  -0.0762   1.0000   0.1469
  -5.750  -0.4848   0.03163   0.02271  -0.0798   1.0000   0.1609
  -5.500  -0.4627   0.03183   0.02338  -0.0790   1.0000   0.1803
  -5.250  -0.4568   0.03430   0.02629  -0.0715   1.0000   0.2023
  -5.000  -0.4228   0.03383   0.02549  -0.0748   1.0000   0.2655
  -4.500  -0.3599   0.03392   0.02474  -0.0793   1.0000   0.3226
  -4.250  -0.3369   0.03439   0.02500  -0.0787   1.0000   0.3365
  -4.000  -0.3119   0.03461   0.02498  -0.0789   1.0000   0.3494
  -3.750  -0.2967   0.03526   0.02561  -0.0760   1.0000   0.3562
  -3.500  -0.2675   0.03492   0.02497  -0.0778   1.0000   0.3656
  -3.250  -0.2517   0.03535   0.02540  -0.0753   1.0000   0.3715
  -3.000  -0.2225   0.03507   0.02488  -0.0772   1.0000   0.3815
  -2.750  -0.2080   0.03549   0.02532  -0.0744   1.0000   0.3873
  -2.500  -0.1693   0.03518   0.02472  -0.0785   0.9976   0.3973
  -2.250  -0.1447   0.03580   0.02542  -0.0774   0.9930   0.4036
  -2.000  -0.1064   0.03567   0.02507  -0.0810   0.9889   0.4141
  -1.750  -0.0796   0.03605   0.02550  -0.0807   0.9842   0.4201
  -1.500  -0.0450   0.03582   0.02514  -0.0833   0.9795   0.4254
  -1.250  -0.0033   0.03540   0.02452  -0.0878   0.9756   0.4285
  -1.000   0.0348   0.03502   0.02398  -0.0914   0.9712   0.4304
  -0.750   0.0683   0.03483   0.02375  -0.0934   0.9660   0.4318
  -0.500   0.1051   0.03476   0.02364  -0.0961   0.9617   0.4338
  -0.250   0.1353   0.03462   0.02348  -0.0975   0.9555   0.4362
   0.250   0.2094   0.03451   0.02329  -0.1032   0.9455   0.4428
   0.500   0.2437   0.03440   0.02312  -0.1057   0.9389   0.4459
   0.750   0.2806   0.03441   0.02317  -0.1082   0.9337   0.4480
   1.000   0.3124   0.03445   0.02329  -0.1097   0.9270   0.4501
   1.250   0.3454   0.03451   0.02342  -0.1114   0.9200   0.4529
   1.500   0.3857   0.03459   0.02358  -0.1143   0.9151   0.4567
   1.750   0.4142   0.03468   0.02372  -0.1154   0.9057   0.4606
   2.000   0.4546   0.03472   0.02386  -0.1182   0.8996   0.4644
   2.250   0.4834   0.03477   0.02408  -0.1189   0.8888   0.4673
   2.500   0.5242   0.03450   0.02397  -0.1211   0.8772   0.4708
   2.750   0.5684   0.03377   0.02337  -0.1234   0.8590   0.4749
   3.000   0.6170   0.03262   0.02237  -0.1258   0.8378   0.4803
   3.250   0.6607   0.03147   0.02148  -0.1270   0.8188   0.4848
   3.500   0.6983   0.03061   0.02086  -0.1275   0.8005   0.4900
   3.750   0.7282   0.02999   0.02042  -0.1269   0.7778   0.4950
   4.000   0.7657   0.02887   0.01956  -0.1267   0.7534   0.4997
   4.250   0.7930   0.02831   0.01924  -0.1254   0.7227   0.5042
   4.500   0.8249   0.02763   0.01874  -0.1246   0.6835   0.5103
   4.750   0.8643   0.02664   0.01778  -0.1242   0.6081   0.5171
   5.000   0.9021   0.02697   0.01672  -0.1233   0.3738   0.5247
   5.250   0.9083   0.02897   0.01781  -0.1201   0.2353   0.5300
   5.500   0.9202   0.03062   0.01899  -0.1180   0.1738   0.5351
   5.750   0.9374   0.03186   0.02014  -0.1165   0.1557   0.5418
   6.000   0.9552   0.03310   0.02140  -0.1151   0.1440   0.5490
   6.250   0.9734   0.03441   0.02276  -0.1138   0.1345   0.5574
   6.500   0.9949   0.03565   0.02413  -0.1128   0.1254   0.5674
   6.750   1.0220   0.03710   0.02562  -0.1128   0.1141   0.5802
   7.000   1.0599   0.03861   0.02732  -0.1142   0.1013   0.5968
   7.250   1.1088   0.04071   0.02966  -0.1174   0.0872   0.6202
   7.500   1.1479   0.04288   0.03206  -0.1191   0.0760   0.6527
   7.750   1.1785   0.04507   0.03484  -0.1189   0.0685   0.7031
   8.000   1.1904   0.04627   0.03641  -0.1158   0.0636   0.8793
   8.250   1.2175   0.04927   0.03982  -0.1155   0.0588   1.0000
   8.500   1.2398   0.05185   0.04256  -0.1150   0.0548   1.0000
   8.750   1.2548   0.05473   0.04579  -0.1133   0.0514   1.0000
   9.000   1.2674   0.05781   0.04927  -0.1113   0.0488   1.0000
   9.250   1.2808   0.06084   0.05249  -0.1097   0.0473   1.0000
   9.500   1.2847   0.06438   0.05644  -0.1069   0.0460   1.0000
   9.750   1.2793   0.06815   0.06072  -0.1030   0.0448   1.0000
  10.000   1.2716   0.07165   0.06468  -0.0994   0.0437   1.0000
  10.250   1.2618   0.07488   0.06820  -0.0958   0.0428   1.0000
  10.500   1.2520   0.07812   0.07167  -0.0926   0.0420   1.0000
  10.750   1.2418   0.08147   0.07522  -0.0900   0.0414   1.0000
  11.000   1.2327   0.08498   0.07886  -0.0880   0.0409   1.0000
  11.250   1.2112   0.08957   0.08373  -0.0861   0.0408   1.0000
  11.500   1.1892   0.09459   0.08899  -0.0853   0.0409   1.0000
  11.750   1.1448   0.10254   0.09737  -0.0868   0.0418   1.0000
  12.000   1.1061   0.11164   0.10665  -0.0911   0.0430   1.0000
<< Back to WORTMANN FX 082-512 AIRFOIL (fx082512-il)

Polar data table (+)

Polar graphs


<< Back to WORTMANN FX 082-512 AIRFOIL (fx082512-il)