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USA 35 A AIRFOIL (usa35a-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: USA 35 A AIRFOIL (usa35a-il)
Reynolds number: 100,000
Max Cl/Cd: 35.6 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-usa35a-il-100000.txt
Download as CSV file: xf-usa35a-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 35 A 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
  ------ -------- --------- --------- -------- -------- --------
  -9.250   0.0272   0.10921   0.10426  -0.0904   0.9375   0.1190
  -9.000   0.0524   0.10525   0.10029  -0.0956   0.9335   0.1233
  -8.750   0.0294   0.10342   0.09849  -0.1028   0.9183   0.1270
  -8.500   0.0780   0.09811   0.09314  -0.1029   0.9158   0.1291
  -8.250   0.0943   0.09549   0.09053  -0.1030   0.9030   0.1325
  -8.000   0.0718   0.09414   0.08919  -0.1083   0.8869   0.1386
  -7.750   0.1028   0.08955   0.08458  -0.1079   0.8806   0.1409
  -7.500   0.1223   0.08705   0.08207  -0.1068   0.8691   0.1445
  -7.250   0.1061   0.08529   0.08029  -0.1108   0.8562   0.1516
  -7.000   0.1039   0.08237   0.07739  -0.1108   0.8422   0.1537
  -6.750   0.1377   0.07923   0.07418  -0.1090   0.8364   0.1569
  -6.500   0.1420   0.07752   0.07249  -0.1083   0.8224   0.1614
  -6.250   0.1135   0.07496   0.06990  -0.1154   0.8085   0.1684
  -6.000   0.1385   0.07261   0.06754  -0.1113   0.8004   0.1709
  -5.750   0.1543   0.07052   0.06540  -0.1107   0.7924   0.1761
  -5.500   0.1442   0.06747   0.06229  -0.1171   0.7802   0.1856
  -5.250   0.1685   0.06533   0.06011  -0.1143   0.7749   0.1894
  -5.000   0.1625   0.06318   0.05785  -0.1198   0.7629   0.2031
  -4.750   0.1850   0.06109   0.05580  -0.1163   0.7568   0.2068
  -4.500   0.1944   0.05873   0.05326  -0.1204   0.7496   0.2230
  -4.250   0.2072   0.05717   0.05180  -0.1173   0.7407   0.2263
  -4.000   0.2243   0.04042   0.03320  -0.1327   0.7359   0.1252
  -3.750   0.2539   0.03766   0.03005  -0.1337   0.7319   0.1242
  -3.500   0.2651   0.03659   0.02876  -0.1320   0.7220   0.1241
  -3.250   0.2923   0.03459   0.02635  -0.1321   0.7162   0.1240
  -3.000   0.3255   0.03268   0.02395  -0.1327   0.7120   0.1251
  -2.750   0.3428   0.03211   0.02314  -0.1313   0.7036   0.1276
  -2.500   0.3688   0.03131   0.02237  -0.1310   0.6967   0.1314
  -2.250   0.4035   0.03006   0.02082  -0.1315   0.6920   0.1355
  -2.000   0.4258   0.02966   0.02023  -0.1306   0.6848   0.1401
  -1.750   0.4484   0.02934   0.01996  -0.1297   0.6775   0.1453
  -1.500   0.4818   0.02860   0.01896  -0.1300   0.6726   0.1522
  -1.250   0.5145   0.02787   0.01826  -0.1305   0.6683   0.1605
  -1.000   0.5279   0.02833   0.01871  -0.1284   0.6597   0.1670
  -0.750   0.5568   0.02781   0.01829  -0.1283   0.6541   0.1776
  -0.500   0.5916   0.02711   0.01757  -0.1289   0.6500   0.1920
  -0.250   0.6085   0.02741   0.01799  -0.1273   0.6429   0.2062
   0.000   0.6308   0.02730   0.01803  -0.1263   0.6362   0.2293
   0.250   0.6627   0.02635   0.01740  -0.1265   0.6315   0.2967
   0.500   0.6842   0.02438   0.01715  -0.1238   0.6281   0.7519
   0.750   0.7110   0.02504   0.01805  -0.1237   0.6187   1.0000
   1.000   0.7400   0.02517   0.01792  -0.1237   0.6131   1.0000
   1.250   0.7771   0.02502   0.01745  -0.1247   0.6088   1.0000
   1.500   0.7869   0.02605   0.01847  -0.1223   0.6006   1.0000
   1.750   0.8122   0.02633   0.01859  -0.1217   0.5939   1.0000
   2.000   0.8495   0.02614   0.01815  -0.1228   0.5893   1.0000
   2.250   0.8637   0.02705   0.01902  -0.1209   0.5818   1.0000
   2.500   0.8839   0.02761   0.01952  -0.1198   0.5746   1.0000
   2.750   0.9196   0.02753   0.01923  -0.1206   0.5698   1.0000
   3.000   0.9414   0.02816   0.01978  -0.1198   0.5635   1.0000
   3.250   0.9537   0.02908   0.02071  -0.1176   0.5551   1.0000
   3.500   0.9894   0.02896   0.02041  -0.1185   0.5499   1.0000
   3.750   1.0194   0.02929   0.02060  -0.1187   0.5446   1.0000
   4.000   1.0203   0.03076   0.02219  -0.1152   0.5356   1.0000
   4.250   1.0525   0.03088   0.02217  -0.1157   0.5304   1.0000
   4.500   1.0942   0.03074   0.02183  -0.1174   0.5266   1.0000
   4.750   1.0797   0.03299   0.02430  -0.1121   0.5170   1.0000
   5.000   1.1071   0.03327   0.02450  -0.1119   0.5114   1.0000
   5.250   1.1495   0.03298   0.02403  -0.1137   0.5074   1.0000
   5.500   1.1374   0.03532   0.02654  -0.1088   0.4996   1.0000
   5.750   1.1446   0.03667   0.02793  -0.1064   0.4932   1.0000
   6.000   1.1822   0.03654   0.02768  -0.1076   0.4892   1.0000
   6.250   1.2321   0.03602   0.02697  -0.1103   0.4861   1.0000
   6.500   1.1231   0.04324   0.03467  -0.0950   0.4735   1.0000
   6.750   1.1788   0.04181   0.03312  -0.0976   0.4708   1.0000
   7.000   1.2533   0.03965   0.03075  -0.1028   0.4680   1.0000
   7.250   1.0532   0.05614   0.04778  -0.0850   0.4494   1.0000
   7.500   1.1500   0.04936   0.04081  -0.0875   0.4502   1.0000
   7.750   1.2684   0.04288   0.03404  -0.0944   0.4495   1.0000
   8.000   1.0465   0.06503   0.05671  -0.0811   0.4277   1.0000
   8.250   1.1084   0.06054   0.05212  -0.0809   0.4276   1.0000
   8.500   1.1768   0.05601   0.04748  -0.0816   0.4275   1.0000
   8.750   1.2646   0.05062   0.04194  -0.0843   0.4275   1.0000
   9.000   0.8620   0.10238   0.09445  -0.0812   0.3798   1.0000
   9.250   0.9007   0.10067   0.09268  -0.0801   0.3775   1.0000
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