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

USA-35B AIRFOIL (usa35b-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: USA-35B AIRFOIL (usa35b-il)
Reynolds number: 100,000
Max Cl/Cd: 56.73 at α=5.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa35b-il-100000-n5.txt
Download as CSV file: xf-usa35b-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA-35B 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
  ------ -------- --------- --------- -------- -------- --------
 -10.000  -0.3516   0.09586   0.09083  -0.0403   1.0000   0.0803
  -9.750  -0.3655   0.09363   0.08871  -0.0395   1.0000   0.0804
  -9.500  -0.3821   0.08551   0.08063  -0.0401   1.0000   0.0605
  -9.250  -0.3844   0.08368   0.07887  -0.0377   1.0000   0.0609
  -9.000  -0.3723   0.08104   0.07625  -0.0386   0.9963   0.0617
  -8.500  -0.3372   0.06105   0.05583  -0.0645   0.9745   0.0520
  -8.250  -0.3137   0.05627   0.05089  -0.0696   0.9664   0.0514
  -8.000  -0.2875   0.05059   0.04495  -0.0756   0.9590   0.0509
  -7.750  -0.2643   0.04519   0.03919  -0.0800   0.9493   0.0505
  -7.500  -0.2362   0.03953   0.03292  -0.0846   0.9421   0.0512
  -7.250  -0.2119   0.03546   0.02826  -0.0866   0.9315   0.0518
  -7.000  -0.1824   0.03243   0.02475  -0.0884   0.9235   0.0519
  -6.750  -0.1538   0.02996   0.02182  -0.0894   0.9140   0.0522
  -6.500  -0.1244   0.02800   0.01954  -0.0902   0.9048   0.0526
  -6.250  -0.0936   0.02646   0.01780  -0.0911   0.8957   0.0532
  -6.000  -0.0652   0.02523   0.01639  -0.0914   0.8845   0.0540
  -5.750  -0.0337   0.02419   0.01519  -0.0921   0.8753   0.0556
  -5.500  -0.0053   0.02320   0.01401  -0.0921   0.8632   0.0574
  -5.250   0.0235   0.02218   0.01275  -0.0920   0.8513   0.0589
  -5.000   0.0534   0.02120   0.01154  -0.0920   0.8403   0.0600
  -4.750   0.0818   0.02034   0.01064  -0.0919   0.8279   0.0613
  -4.500   0.1092   0.01969   0.00996  -0.0916   0.8146   0.0629
  -4.250   0.1372   0.01911   0.00932  -0.0914   0.8020   0.0653
  -4.000   0.1656   0.01860   0.00869  -0.0911   0.7900   0.0691
  -3.750   0.1928   0.01809   0.00820  -0.0908   0.7771   0.0730
  -3.500   0.2199   0.01767   0.00774  -0.0904   0.7641   0.0778
  -3.250   0.2472   0.01723   0.00726  -0.0900   0.7522   0.0839
  -3.000   0.2750   0.01685   0.00683  -0.0897   0.7413   0.0952
  -2.750   0.3017   0.01647   0.00652  -0.0894   0.7289   0.1150
  -2.500   0.3288   0.01609   0.00629  -0.0891   0.7176   0.1547
  -2.250   0.3558   0.01561   0.00609  -0.0889   0.7067   0.2481
  -2.000   0.3790   0.01466   0.00611  -0.0881   0.6947   0.5139
  -1.750   0.4067   0.01363   0.00612  -0.0863   0.6832   0.9080
  -1.500   0.4455   0.01364   0.00589  -0.0883   0.6722   1.0000
  -1.250   0.4721   0.01378   0.00583  -0.0878   0.6617   1.0000
  -1.000   0.4989   0.01394   0.00580  -0.0875   0.6524   1.0000
  -0.750   0.5258   0.01411   0.00579  -0.0871   0.6435   1.0000
  -0.500   0.5525   0.01429   0.00584  -0.0867   0.6345   1.0000
  -0.250   0.5795   0.01447   0.00587  -0.0864   0.6260   1.0000
   0.000   0.6061   0.01467   0.00597  -0.0861   0.6174   1.0000
   0.250   0.6330   0.01487   0.00605  -0.0857   0.6090   1.0000
   0.500   0.6596   0.01509   0.00619  -0.0854   0.6006   1.0000
   0.750   0.6863   0.01530   0.00631  -0.0850   0.5921   1.0000
   1.000   0.7128   0.01553   0.00648  -0.0847   0.5837   1.0000
   1.250   0.7394   0.01575   0.00664  -0.0843   0.5752   1.0000
   1.500   0.7657   0.01599   0.00684  -0.0839   0.5666   1.0000
   1.750   0.7920   0.01622   0.00702  -0.0835   0.5578   1.0000
   2.000   0.8181   0.01647   0.00725  -0.0831   0.5489   1.0000
   2.250   0.8442   0.01670   0.00745  -0.0826   0.5399   1.0000
   2.500   0.8698   0.01695   0.00771  -0.0822   0.5302   1.0000
   2.750   0.8958   0.01718   0.00787  -0.0816   0.5206   1.0000
   3.000   0.9202   0.01740   0.00812  -0.0809   0.5071   1.0000
   3.250   0.9444   0.01760   0.00831  -0.0801   0.4927   1.0000
   3.500   0.9686   0.01782   0.00851  -0.0794   0.4785   1.0000
   3.750   0.9930   0.01807   0.00875  -0.0786   0.4657   1.0000
   4.000   1.0173   0.01833   0.00898  -0.0779   0.4536   1.0000
   4.250   1.0411   0.01862   0.00931  -0.0772   0.4402   1.0000
   4.500   1.0647   0.01892   0.00965  -0.0764   0.4270   1.0000
   4.750   1.0881   0.01925   0.00999  -0.0756   0.4140   1.0000
   5.000   1.1110   0.01960   0.01034  -0.0747   0.4010   1.0000
   5.250   1.1335   0.01998   0.01072  -0.0738   0.3882   1.0000
   5.500   1.1557   0.02040   0.01114  -0.0728   0.3754   1.0000
   5.750   1.1775   0.02085   0.01162  -0.0719   0.3628   1.0000
   6.000   1.1989   0.02134   0.01213  -0.0708   0.3508   1.0000
   6.250   1.2196   0.02186   0.01265  -0.0697   0.3391   1.0000
   6.500   1.2395   0.02243   0.01322  -0.0685   0.3268   1.0000
   6.750   1.2584   0.02302   0.01384  -0.0672   0.3128   1.0000
   7.000   1.2758   0.02367   0.01451  -0.0657   0.2976   1.0000
   7.250   1.2920   0.02435   0.01519  -0.0641   0.2820   1.0000
   7.500   1.3074   0.02507   0.01592  -0.0624   0.2670   1.0000
   7.750   1.3221   0.02582   0.01669  -0.0606   0.2530   1.0000
   8.000   1.3357   0.02661   0.01752  -0.0588   0.2402   1.0000
   8.250   1.3469   0.02745   0.01837  -0.0566   0.2286   1.0000
   8.500   1.3577   0.02833   0.01930  -0.0544   0.2178   1.0000
   8.750   1.3680   0.02929   0.02032  -0.0523   0.2074   1.0000
   9.000   1.3760   0.03039   0.02146  -0.0502   0.1970   1.0000
   9.250   1.3826   0.03165   0.02272  -0.0480   0.1868   1.0000
   9.500   1.3904   0.03292   0.02408  -0.0462   0.1773   1.0000
   9.750   1.3955   0.03442   0.02562  -0.0444   0.1683   1.0000
  10.000   1.3997   0.03607   0.02732  -0.0427   0.1593   1.0000
  10.250   1.4042   0.03780   0.02917  -0.0412   0.1502   1.0000
  10.500   1.4053   0.03988   0.03129  -0.0398   0.1416   1.0000
  10.750   1.4067   0.04208   0.03359  -0.0386   0.1316   1.0000
  11.000   1.4067   0.04453   0.03614  -0.0377   0.1212   1.0000
  11.250   1.4041   0.04736   0.03904  -0.0370   0.1109   1.0000
  11.500   1.3995   0.05054   0.04228  -0.0366   0.1010   1.0000
  11.750   1.3942   0.05396   0.04575  -0.0364   0.0914   1.0000
  12.000   1.3885   0.05753   0.04939  -0.0364   0.0831   1.0000
  12.250   1.3806   0.06147   0.05340  -0.0366   0.0775   1.0000
  12.500   1.3717   0.06565   0.05763  -0.0370   0.0731   1.0000
  12.750   1.3618   0.07009   0.06211  -0.0377   0.0700   1.0000
  13.000   1.3535   0.07443   0.06654  -0.0384   0.0672   1.0000
  13.250   1.3439   0.07905   0.07125  -0.0393   0.0649   1.0000
  13.500   1.3335   0.08389   0.07614  -0.0404   0.0631   1.0000
  13.750   1.3250   0.08850   0.08082  -0.0414   0.0613   1.0000
  14.000   1.3191   0.09281   0.08526  -0.0424   0.0594   1.0000
<< Back to USA-35B AIRFOIL (usa35b-il)

Polar data table (+)

Polar graphs


<< Back to USA-35B AIRFOIL (usa35b-il)