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=100,000 Ncrit=5


Details Polar file
Airfoil: WORTMANN FX 082-512 AIRFOIL (fx082512-il)
Reynolds number: 100,000
Max Cl/Cd: 47.66 at α=3.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-fx082512-il-100000-n5.txt
Download as CSV file: xf-fx082512-il-100000-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 =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -12.250  -0.6350   0.08894   0.08361  -0.0569   1.0000   0.0311
 -12.000  -0.6628   0.08117   0.07576  -0.0606   1.0000   0.0309
 -11.750  -0.6883   0.07476   0.06925  -0.0634   1.0000   0.0308
 -11.500  -0.7118   0.06922   0.06360  -0.0654   1.0000   0.0308
 -11.250  -0.7319   0.06465   0.05890  -0.0663   1.0000   0.0309
 -11.000  -0.7508   0.06062   0.05475  -0.0666   1.0000   0.0311
 -10.500  -0.7790   0.05521   0.04927  -0.0646   1.0000   0.0318
 -10.250  -0.7939   0.05299   0.04704  -0.0627   1.0000   0.0321
 -10.000  -0.8010   0.05055   0.04457  -0.0627   1.0000   0.0327
  -9.750  -0.8022   0.04786   0.04181  -0.0638   1.0000   0.0336
  -9.500  -0.7980   0.04498   0.03879  -0.0657   1.0000   0.0352
  -9.250  -0.7891   0.04201   0.03556  -0.0676   1.0000   0.0380
  -9.000  -0.7706   0.03965   0.03311  -0.0722   1.0000   0.0404
  -8.750  -0.7522   0.03742   0.03063  -0.0744   1.0000   0.0445
  -8.500  -0.7300   0.03564   0.02864  -0.0772   1.0000   0.0490
  -8.250  -0.7054   0.03409   0.02685  -0.0801   1.0000   0.0545
  -8.000  -0.6811   0.03257   0.02511  -0.0820   1.0000   0.0607
  -7.750  -0.6549   0.03127   0.02356  -0.0843   1.0000   0.0677
  -7.500  -0.6251   0.03021   0.02237  -0.0862   0.9986   0.0739
  -7.250  -0.5917   0.02922   0.02117  -0.0891   0.9966   0.0814
  -7.000  -0.5594   0.02845   0.02027  -0.0913   0.9948   0.0886
  -6.750  -0.5326   0.02803   0.01997  -0.0912   0.9924   0.0926
  -6.500  -0.5017   0.02762   0.01945  -0.0924   0.9898   0.0968
  -6.250  -0.4700   0.02723   0.01911  -0.0939   0.9875   0.1011
  -6.000  -0.4348   0.02679   0.01854  -0.0963   0.9855   0.1045
  -5.750  -0.3987   0.02595   0.01762  -0.0996   0.9833   0.1081
  -5.500  -0.3610   0.02502   0.01655  -0.1033   0.9810   0.1133
  -5.250  -0.3203   0.02389   0.01527  -0.1080   0.9793   0.1200
  -5.000  -0.2784   0.02271   0.01397  -0.1129   0.9778   0.1310
  -4.750  -0.2399   0.02247   0.01460  -0.1164   0.9758   0.2582
  -4.500  -0.2046   0.02252   0.01436  -0.1183   0.9731   0.2875
  -4.250  -0.1744   0.02272   0.01434  -0.1190   0.9690   0.3018
  -4.000  -0.1427   0.02312   0.01463  -0.1198   0.9653   0.3121
  -3.750  -0.1078   0.02338   0.01471  -0.1214   0.9625   0.3215
  -3.500  -0.0795   0.02370   0.01497  -0.1215   0.9583   0.3290
  -3.250  -0.0489   0.02382   0.01497  -0.1223   0.9540   0.3357
  -3.000  -0.0174   0.02406   0.01518  -0.1230   0.9505   0.3406
  -2.750   0.0198   0.02399   0.01492  -0.1254   0.9481   0.3473
  -2.500   0.0466   0.02414   0.01509  -0.1252   0.9432   0.3514
  -2.250   0.0749   0.02430   0.01527  -0.1254   0.9384   0.3565
  -2.000   0.1107   0.02423   0.01505  -0.1275   0.9352   0.3631
  -1.750   0.1448   0.02428   0.01515  -0.1287   0.9325   0.3672
  -1.500   0.1706   0.02437   0.01528  -0.1285   0.9270   0.3713
  -1.250   0.2013   0.02433   0.01520  -0.1294   0.9220   0.3761
  -1.000   0.2380   0.02416   0.01489  -0.1317   0.9188   0.3804
  -0.750   0.2737   0.02400   0.01478  -0.1333   0.9160   0.3823
  -0.500   0.3001   0.02394   0.01476  -0.1334   0.9101   0.3839
  -0.250   0.3305   0.02386   0.01472  -0.1341   0.9048   0.3860
   0.000   0.3657   0.02372   0.01461  -0.1357   0.9013   0.3885
   0.250   0.4051   0.02348   0.01440  -0.1380   0.8985   0.3913
   0.500   0.4310   0.02333   0.01426  -0.1378   0.8882   0.3936
   0.750   0.4792   0.02259   0.01352  -0.1412   0.8823   0.3963
   1.000   0.5097   0.02215   0.01315  -0.1413   0.8699   0.3981
   1.250   0.5460   0.02161   0.01272  -0.1424   0.8613   0.4001
   1.500   0.5800   0.02116   0.01240  -0.1431   0.8523   0.4023
   1.750   0.6120   0.02066   0.01201  -0.1432   0.8393   0.4052
   2.000   0.6481   0.01988   0.01132  -0.1438   0.8213   0.4085
   2.250   0.6839   0.01920   0.01068  -0.1443   0.7987   0.4119
   2.500   0.7177   0.01872   0.01028  -0.1447   0.7736   0.4147
   2.750   0.7566   0.01819   0.00984  -0.1459   0.7416   0.4172
   3.000   0.8105   0.01760   0.00913  -0.1498   0.6737   0.4207
   3.250   0.8551   0.01794   0.00884  -0.1522   0.5391   0.4243
   3.500   0.8693   0.01927   0.00926  -0.1495   0.3818   0.4271
   3.750   0.8803   0.02100   0.01001  -0.1468   0.2081   0.4300
   4.000   0.8964   0.02232   0.01079  -0.1450   0.1305   0.4326
   4.500   0.9393   0.02355   0.01214  -0.1427   0.1157   0.4395
   4.750   0.9603   0.02418   0.01285  -0.1414   0.1106   0.4433
   5.000   0.9792   0.02495   0.01365  -0.1399   0.1062   0.4468
   5.250   0.9985   0.02560   0.01447  -0.1383   0.1026   0.4499
   5.500   1.0157   0.02650   0.01544  -0.1365   0.0984   0.4536
   5.750   1.0348   0.02733   0.01636  -0.1350   0.0924   0.4580
   6.000   1.0546   0.02822   0.01730  -0.1337   0.0860   0.4631
   6.250   1.0749   0.02910   0.01831  -0.1325   0.0791   0.4679
   6.500   1.0961   0.03008   0.01935  -0.1315   0.0726   0.4739
   6.750   1.1175   0.03095   0.02040  -0.1304   0.0653   0.4805
   7.000   1.1388   0.03189   0.02157  -0.1293   0.0580   0.4869
   7.250   1.1587   0.03277   0.02255  -0.1281   0.0523   0.4947
   7.500   1.1778   0.03351   0.02347  -0.1266   0.0471   0.5028
   7.750   1.1981   0.03438   0.02455  -0.1254   0.0420   0.5137
   8.000   1.2216   0.03569   0.02607  -0.1247   0.0382   0.5275
   8.250   1.2392   0.03655   0.02708  -0.1232   0.0356   0.5450
   8.500   1.2649   0.03824   0.02911  -0.1229   0.0322   0.5717
   8.750   1.2840   0.03946   0.03063  -0.1216   0.0301   0.6146
   9.250   1.3087   0.04214   0.03415  -0.1162   0.0279   1.0000
   9.500   1.3341   0.04615   0.03872  -0.1160   0.0261   1.0000
   9.750   1.3459   0.04973   0.04277  -0.1138   0.0245   1.0000
  10.000   1.3500   0.05321   0.04666  -0.1108   0.0235   1.0000
  10.250   1.3477   0.05697   0.05082  -0.1071   0.0230   1.0000
  10.500   1.3398   0.06082   0.05505  -0.1032   0.0226   1.0000
  10.750   1.3273   0.06487   0.05946  -0.0992   0.0223   1.0000
  11.000   1.3110   0.06918   0.06410  -0.0956   0.0222   1.0000
  11.250   1.2914   0.07389   0.06914  -0.0926   0.0221   1.0000
  11.500   1.2689   0.07906   0.07462  -0.0905   0.0221   1.0000
  11.750   1.2431   0.08496   0.08080  -0.0896   0.0221   1.0000
  12.000   1.2140   0.09181   0.08793  -0.0902   0.0223   1.0000
  12.250   1.1828   0.09981   0.09617  -0.0928   0.0225   1.0000
  12.500   1.1480   0.10983   0.10641  -0.0983   0.0228   1.0000
  12.750   1.1098   0.12344   0.12021  -0.1077   0.0235   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)