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

USA 46 AIRFOIL (usa46-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: USA 46 AIRFOIL (usa46-il)
Reynolds number: 500,000
Max Cl/Cd: 75.06 at α=2.25°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa46-il-500000-n5.txt
Download as CSV file: xf-usa46-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 46 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.500 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.5701   0.08426   0.08207  -0.0129   1.0000   0.0047
  -9.250  -0.5984   0.07471   0.07260  -0.0179   1.0000   0.0046
  -9.000  -0.6240   0.06670   0.06465  -0.0231   1.0000   0.0045
  -8.500  -0.6607   0.02898   0.02591  -0.0571   1.0000   0.0041
  -8.250  -0.6450   0.02489   0.02133  -0.0582   1.0000   0.0042
  -8.000  -0.6260   0.02209   0.01812  -0.0585   1.0000   0.0044
  -7.750  -0.6049   0.02002   0.01572  -0.0583   1.0000   0.0046
  -7.500  -0.5828   0.01840   0.01380  -0.0580   1.0000   0.0048
  -7.250  -0.5600   0.01709   0.01224  -0.0576   1.0000   0.0050
  -7.000  -0.5377   0.01562   0.01058  -0.0572   1.0000   0.0055
  -6.750  -0.5140   0.01487   0.00973  -0.0567   1.0000   0.0060
  -6.500  -0.4901   0.01407   0.00879  -0.0562   1.0000   0.0065
  -6.250  -0.4662   0.01322   0.00774  -0.0555   1.0000   0.0070
  -6.000  -0.4421   0.01252   0.00689  -0.0549   1.0000   0.0075
  -5.750  -0.4180   0.01178   0.00600  -0.0542   1.0000   0.0081
  -5.500  -0.3902   0.01106   0.00515  -0.0543   0.9990   0.0097
  -5.250  -0.3557   0.01057   0.00459  -0.0559   0.9957   0.0119
  -5.000  -0.3228   0.01005   0.00397  -0.0570   0.9921   0.0166
  -4.750  -0.2904   0.00963   0.00358  -0.0581   0.9884   0.0275
  -4.500  -0.2584   0.00937   0.00332  -0.0590   0.9843   0.0368
  -4.250  -0.2266   0.00916   0.00308  -0.0599   0.9792   0.0423
  -4.000  -0.1948   0.00890   0.00282  -0.0607   0.9736   0.0493
  -3.750  -0.1621   0.00870   0.00257  -0.0618   0.9677   0.0550
  -3.500  -0.1299   0.00845   0.00230  -0.0627   0.9608   0.0609
  -3.250  -0.0962   0.00823   0.00208  -0.0639   0.9536   0.0691
  -3.000  -0.0625   0.00799   0.00184  -0.0651   0.9435   0.0816
  -2.750  -0.0267   0.00775   0.00166  -0.0668   0.9293   0.1020
  -2.500   0.0075   0.00761   0.00152  -0.0680   0.9105   0.1228
  -2.250   0.0372   0.00755   0.00140  -0.0683   0.8898   0.1358
  -2.000   0.0652   0.00751   0.00129  -0.0681   0.8695   0.1451
  -1.750   0.0922   0.00750   0.00120  -0.0677   0.8480   0.1529
  -1.500   0.1189   0.00747   0.00112  -0.0672   0.8265   0.1615
  -1.250   0.1454   0.00746   0.00105  -0.0668   0.8035   0.1720
  -1.000   0.1716   0.00745   0.00098  -0.0663   0.7780   0.1854
  -0.750   0.1977   0.00744   0.00093  -0.0658   0.7501   0.2019
  -0.500   0.2236   0.00739   0.00088  -0.0653   0.7210   0.2337
  -0.250   0.2495   0.00727   0.00086  -0.0649   0.6941   0.2969
   0.000   0.2752   0.00708   0.00086  -0.0646   0.6694   0.3948
   0.250   0.3008   0.00690   0.00088  -0.0642   0.6450   0.4962
   0.500   0.3261   0.00676   0.00093  -0.0636   0.6198   0.5888
   0.750   0.3507   0.00668   0.00103  -0.0628   0.5925   0.6773
   1.000   0.3757   0.00672   0.00112  -0.0621   0.5661   0.7380
   1.250   0.4013   0.00678   0.00119  -0.0615   0.5440   0.7736
   1.500   0.4268   0.00680   0.00125  -0.0609   0.5270   0.8073
   1.750   0.4518   0.00674   0.00131  -0.0600   0.5107   0.8461
   2.000   0.4904   0.00664   0.00137  -0.0623   0.4878   1.0000
   2.250   0.5164   0.00688   0.00147  -0.0619   0.4473   1.0000
   2.500   0.5415   0.00726   0.00159  -0.0614   0.3837   1.0000
   2.750   0.5667   0.00765   0.00176  -0.0609   0.3346   1.0000
   3.000   0.5921   0.00804   0.00196  -0.0605   0.2931   1.0000
   3.250   0.6180   0.00835   0.00216  -0.0602   0.2607   1.0000
   3.500   0.6436   0.00872   0.00236  -0.0598   0.2179   1.0000
   3.750   0.6667   0.00950   0.00270  -0.0592   0.1236   1.0000
   4.000   0.6897   0.01035   0.00319  -0.0586   0.0469   1.0000
   4.250   0.7145   0.01090   0.00362  -0.0580   0.0181   1.0000
   4.500   0.7402   0.01133   0.00409  -0.0575   0.0123   1.0000
   4.750   0.7659   0.01173   0.00458  -0.0570   0.0109   1.0000
   5.000   0.7914   0.01218   0.00511  -0.0565   0.0098   1.0000
   5.250   0.8166   0.01263   0.00561  -0.0560   0.0084   1.0000
   5.500   0.8405   0.01332   0.00638  -0.0553   0.0073   1.0000
   5.750   0.8637   0.01417   0.00733  -0.0545   0.0068   1.0000
   6.000   0.8874   0.01487   0.00812  -0.0537   0.0064   1.0000
   6.250   0.9103   0.01572   0.00908  -0.0528   0.0060   1.0000
   6.500   0.9327   0.01669   0.01015  -0.0518   0.0057   1.0000
   6.750   0.9555   0.01754   0.01108  -0.0510   0.0052   1.0000
   7.000   0.9789   0.01816   0.01173  -0.0505   0.0048   1.0000
   7.250   0.9991   0.01949   0.01315  -0.0494   0.0044   1.0000
   7.500   1.0201   0.02083   0.01470  -0.0483   0.0042   1.0000
   7.750   1.0410   0.02223   0.01630  -0.0472   0.0040   1.0000
   8.000   1.0604   0.02407   0.01838  -0.0459   0.0038   1.0000
   8.250   1.0783   0.02632   0.02094  -0.0444   0.0036   1.0000
   8.500   1.0938   0.02897   0.02392  -0.0427   0.0035   1.0000
   8.750   1.1061   0.03211   0.02744  -0.0408   0.0034   1.0000
   9.000   1.1143   0.03568   0.03141  -0.0387   0.0033   1.0000
   9.250   1.1168   0.03984   0.03598  -0.0363   0.0033   1.0000
   9.500   1.1127   0.04445   0.04098  -0.0336   0.0033   1.0000
   9.750   1.1024   0.04907   0.04592  -0.0310   0.0033   1.0000
  10.000   1.0862   0.05283   0.04991  -0.0279   0.0034   1.0000
  10.250   1.0646   0.05689   0.05417  -0.0260   0.0034   1.0000
  10.500   1.0451   0.06142   0.05887  -0.0265   0.0034   1.0000
  10.750   1.0245   0.06726   0.06486  -0.0295   0.0035   1.0000
  11.000   1.0057   0.07438   0.07212  -0.0349   0.0035   1.0000
  11.250   0.9849   0.08392   0.08179  -0.0425   0.0035   1.0000
  11.500   0.9634   0.09630   0.09428  -0.0515   0.0036   1.0000
<< Back to USA 46 AIRFOIL (usa46-il)

Polar data table (+)

Polar graphs


<< Back to USA 46 AIRFOIL (usa46-il)