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

USA 40 AIRFOIL (usa40-il) Xfoil prediction polar at RE=1,000,000 Ncrit=5


Details Polar file
Airfoil: USA 40 AIRFOIL (usa40-il)
Reynolds number: 1,000,000
Max Cl/Cd: 87.02 at α=2.5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa40-il-1000000-n5.txt
Download as CSV file: xf-usa40-il-1000000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 40 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -14.750  -0.8709   0.06217   0.05952  -0.0678   1.0000   0.0214
 -14.500  -0.9282   0.04797   0.04506  -0.0797   1.0000   0.0214
 -14.250  -0.9512   0.03931   0.03620  -0.0883   1.0000   0.0214
 -14.000  -0.9653   0.03403   0.03076  -0.0932   1.0000   0.0215
 -13.750  -0.9798   0.03065   0.02726  -0.0948   1.0000   0.0215
 -13.500  -0.9702   0.02845   0.02492  -0.0968   0.9991   0.0216
 -13.250  -0.9427   0.02672   0.02306  -0.1001   0.9964   0.0218
 -13.000  -0.9135   0.02527   0.02149  -0.1030   0.9941   0.0220
 -12.750  -0.8854   0.02397   0.02008  -0.1051   0.9910   0.0221
 -12.500  -0.8571   0.02276   0.01876  -0.1070   0.9874   0.0222
 -12.250  -0.8277   0.02167   0.01757  -0.1089   0.9848   0.0224
 -12.000  -0.7976   0.02070   0.01649  -0.1106   0.9825   0.0225
 -11.750  -0.7715   0.01985   0.01556  -0.1112   0.9772   0.0226
 -11.500  -0.7422   0.01905   0.01467  -0.1124   0.9729   0.0227
 -11.250  -0.7126   0.01832   0.01386  -0.1134   0.9688   0.0228
 -11.000  -0.6850   0.01766   0.01312  -0.1140   0.9625   0.0228
 -10.750  -0.6562   0.01669   0.01206  -0.1150   0.9567   0.0230
 -10.500  -0.6292   0.01592   0.01122  -0.1155   0.9482   0.0233
 -10.250  -0.6009   0.01529   0.01053  -0.1160   0.9399   0.0235
 -10.000  -0.5734   0.01476   0.00994  -0.1162   0.9300   0.0236
  -9.750  -0.5458   0.01429   0.00940  -0.1164   0.9211   0.0238
  -9.500  -0.5185   0.01387   0.00891  -0.1165   0.9109   0.0240
  -9.250  -0.4911   0.01348   0.00845  -0.1165   0.9009   0.0242
  -9.000  -0.4638   0.01312   0.00801  -0.1164   0.8901   0.0244
  -8.750  -0.4364   0.01278   0.00760  -0.1164   0.8781   0.0246
  -8.500  -0.4091   0.01248   0.00722  -0.1163   0.8637   0.0249
  -8.250  -0.3820   0.01221   0.00685  -0.1161   0.8461   0.0252
  -8.000  -0.3549   0.01195   0.00648  -0.1159   0.8274   0.0254
  -7.750  -0.3276   0.01169   0.00611  -0.1157   0.8098   0.0257
  -7.500  -0.3001   0.01144   0.00577  -0.1156   0.7942   0.0259
  -7.250  -0.2725   0.01121   0.00545  -0.1155   0.7779   0.0261
  -7.000  -0.2449   0.01100   0.00514  -0.1154   0.7602   0.0263
  -6.750  -0.2171   0.01082   0.00486  -0.1152   0.7424   0.0265
  -6.500  -0.1893   0.01065   0.00460  -0.1151   0.7240   0.0266
  -6.250  -0.1616   0.01042   0.00427  -0.1150   0.7055   0.0270
  -5.750  -0.1057   0.01004   0.00374  -0.1149   0.6722   0.0278
  -5.500  -0.0775   0.00990   0.00353  -0.1149   0.6571   0.0283
  -5.250  -0.0492   0.00978   0.00335  -0.1148   0.6435   0.0287
  -5.000  -0.0209   0.00967   0.00317  -0.1148   0.6317   0.0292
  -4.750   0.0077   0.00955   0.00301  -0.1148   0.6201   0.0297
  -4.500   0.0362   0.00946   0.00286  -0.1148   0.6088   0.0302
  -4.250   0.0647   0.00938   0.00272  -0.1147   0.5971   0.0307
  -4.000   0.0934   0.00927   0.00257  -0.1148   0.5875   0.0315
  -3.750   0.1219   0.00918   0.00244  -0.1147   0.5760   0.0327
  -3.500   0.1505   0.00910   0.00233  -0.1147   0.5630   0.0338
  -3.250   0.1791   0.00905   0.00223  -0.1147   0.5500   0.0350
  -3.000   0.2076   0.00901   0.00213  -0.1147   0.5369   0.0368
  -2.750   0.2359   0.00897   0.00206  -0.1146   0.5210   0.0397
  -2.500   0.2642   0.00895   0.00199  -0.1146   0.5027   0.0446
  -2.250   0.2923   0.00895   0.00194  -0.1145   0.4836   0.0523
  -2.000   0.3205   0.00896   0.00191  -0.1145   0.4672   0.0605
  -1.750   0.3485   0.00899   0.00190  -0.1144   0.4485   0.0681
  -1.500   0.3763   0.00905   0.00189  -0.1143   0.4265   0.0753
  -1.250   0.4043   0.00912   0.00189  -0.1142   0.4084   0.0822
  -1.000   0.4325   0.00915   0.00188  -0.1142   0.3957   0.0912
  -0.750   0.4611   0.00908   0.00186  -0.1142   0.3864   0.1142
  -0.500   0.4894   0.00900   0.00188  -0.1143   0.3768   0.1622
  -0.250   0.5179   0.00899   0.00191  -0.1143   0.3673   0.1891
   0.000   0.5462   0.00901   0.00195  -0.1143   0.3581   0.2096
   0.250   0.5743   0.00905   0.00199  -0.1143   0.3459   0.2301
   0.500   0.6027   0.00904   0.00203  -0.1144   0.3364   0.2638
   0.750   0.6307   0.00901   0.00210  -0.1144   0.3255   0.3210
   1.000   0.6589   0.00901   0.00217  -0.1145   0.3142   0.3713
   1.250   0.6868   0.00906   0.00225  -0.1145   0.3031   0.4076
   1.500   0.7147   0.00909   0.00234  -0.1144   0.2924   0.4549
   1.750   0.7427   0.00910   0.00244  -0.1145   0.2820   0.5068
   2.000   0.7703   0.00916   0.00256  -0.1144   0.2706   0.5616
   2.250   0.7974   0.00927   0.00272  -0.1142   0.2558   0.6147
   2.500   0.8241   0.00947   0.00289  -0.1140   0.2370   0.6458
   2.750   0.8500   0.00978   0.00310  -0.1136   0.2130   0.6673
   3.000   0.8752   0.01016   0.00335  -0.1131   0.1847   0.6847
   3.250   0.8980   0.01082   0.00377  -0.1123   0.1378   0.7019
   3.500   0.9196   0.01162   0.00431  -0.1113   0.0833   0.7147
   3.750   0.9460   0.01185   0.00453  -0.1110   0.0780   0.7265
   4.000   0.9728   0.01204   0.00472  -0.1108   0.0743   0.7376
   4.250   0.9990   0.01229   0.00494  -0.1104   0.0693   0.7472
   4.500   1.0249   0.01257   0.00519  -0.1100   0.0608   0.7545
   4.750   1.0483   0.01311   0.00562  -0.1093   0.0308   0.7646
   5.000   1.0741   0.01333   0.00587  -0.1089   0.0293   0.7782
   5.250   1.0999   0.01356   0.00613  -0.1085   0.0285   0.7907
   5.500   1.1255   0.01380   0.00639  -0.1081   0.0278   0.8032
   5.750   1.1505   0.01404   0.00668  -0.1076   0.0271   0.8179
   6.000   1.1752   0.01429   0.00699  -0.1070   0.0265   0.8359
   6.250   1.1996   0.01450   0.00727  -0.1063   0.0262   0.8580
   6.500   1.2217   0.01454   0.00752  -0.1052   0.0259   0.9805
   6.750   1.2463   0.01486   0.00784  -0.1047   0.0255   1.0000
   7.000   1.2706   0.01519   0.00817  -0.1041   0.0251   1.0000
   7.250   1.2945   0.01555   0.00853  -0.1035   0.0247   1.0000
   7.500   1.3179   0.01592   0.00891  -0.1028   0.0243   1.0000
   7.750   1.3408   0.01632   0.00932  -0.1021   0.0240   1.0000
   8.000   1.3629   0.01676   0.00976  -0.1012   0.0237   1.0000
   8.250   1.3841   0.01724   0.01027  -0.1003   0.0233   1.0000
   8.500   1.4048   0.01773   0.01077  -0.0992   0.0231   1.0000
   8.750   1.4259   0.01816   0.01122  -0.0982   0.0229   1.0000
   9.000   1.4464   0.01860   0.01167  -0.0972   0.0226   1.0000
   9.250   1.4647   0.01905   0.01216  -0.0957   0.0223   1.0000
   9.500   1.4818   0.01956   0.01269  -0.0941   0.0221   1.0000
   9.750   1.4980   0.02012   0.01328  -0.0924   0.0218   1.0000
  10.000   1.5134   0.02072   0.01391  -0.0907   0.0216   1.0000
  10.250   1.5281   0.02137   0.01459  -0.0889   0.0214   1.0000
  10.500   1.5420   0.02208   0.01534  -0.0872   0.0211   1.0000
  10.750   1.5554   0.02284   0.01614  -0.0855   0.0209   1.0000
  11.000   1.5681   0.02368   0.01701  -0.0838   0.0207   1.0000
  11.250   1.5794   0.02465   0.01803  -0.0821   0.0204   1.0000
  11.500   1.5887   0.02581   0.01924  -0.0803   0.0202   1.0000
  11.750   1.5956   0.02720   0.02069  -0.0785   0.0200   1.0000
  12.000   1.6060   0.02839   0.02194  -0.0771   0.0199   1.0000
  12.250   1.6155   0.02971   0.02332  -0.0758   0.0197   1.0000
  12.500   1.6242   0.03115   0.02482  -0.0747   0.0196   1.0000
  12.750   1.6323   0.03271   0.02644  -0.0736   0.0194   1.0000
  13.000   1.6393   0.03444   0.02823  -0.0726   0.0191   1.0000
  13.250   1.6451   0.03634   0.03020  -0.0718   0.0189   1.0000
  13.500   1.6497   0.03845   0.03238  -0.0710   0.0187   1.0000
  13.750   1.6531   0.04077   0.03478  -0.0705   0.0185   1.0000
  14.000   1.6549   0.04334   0.03742  -0.0700   0.0184   1.0000
  14.250   1.6554   0.04615   0.04031  -0.0697   0.0182   1.0000
  14.500   1.6550   0.04914   0.04338  -0.0696   0.0181   1.0000
  14.750   1.6534   0.05238   0.04670  -0.0696   0.0179   1.0000
  15.000   1.6506   0.05580   0.05022  -0.0697   0.0178   1.0000
  15.250   1.6464   0.05952   0.05402  -0.0700   0.0177   1.0000
  15.500   1.6412   0.06344   0.05803  -0.0705   0.0175   1.0000
  15.750   1.6328   0.06787   0.06256  -0.0711   0.0174   1.0000
  16.000   1.6226   0.07264   0.06742  -0.0719   0.0173   1.0000
  16.250   1.6106   0.07780   0.07269  -0.0730   0.0172   1.0000
  16.500   1.5960   0.08338   0.07837  -0.0742   0.0171   1.0000
<< Back to USA 40 AIRFOIL (usa40-il)

Polar data table (+)

Polar graphs


<< Back to USA 40 AIRFOIL (usa40-il)