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USA 49 AIRFOIL (usa49-il) Xfoil prediction polar at RE=50,000 Ncrit=5


Details Polar file
Airfoil: USA 49 AIRFOIL (usa49-il)
Reynolds number: 50,000
Max Cl/Cd: 33.62 at α=5°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-usa49-il-50000-n5.txt
Download as CSV file: xf-usa49-il-50000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: USA 49 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.4381   0.09759   0.09112  -0.0113   1.0000   0.0524
  -8.750  -0.4355   0.09336   0.08692  -0.0116   1.0000   0.0497
  -8.500  -0.5454   0.10286   0.09613  -0.0020   1.0000   0.0624
  -8.250  -0.5452   0.09753   0.09084  -0.0060   1.0000   0.0512
  -8.000  -0.5430   0.09316   0.08652  -0.0080   1.0000   0.0484
  -7.750  -0.5439   0.08814   0.08158  -0.0124   1.0000   0.0458
  -7.500  -0.5418   0.08140   0.07478  -0.0208   1.0000   0.0416
  -7.250  -0.5356   0.07723   0.07062  -0.0218   1.0000   0.0405
  -7.000  -0.5292   0.07267   0.06603  -0.0239   1.0000   0.0394
  -6.750  -0.5217   0.06790   0.06117  -0.0262   1.0000   0.0383
  -6.500  -0.5127   0.06298   0.05612  -0.0283   1.0000   0.0373
  -6.250  -0.5020   0.05808   0.05101  -0.0301   1.0000   0.0365
  -6.000  -0.4893   0.05342   0.04606  -0.0314   1.0000   0.0363
  -5.750  -0.4746   0.04911   0.04141  -0.0321   1.0000   0.0366
  -5.500  -0.4579   0.04511   0.03702  -0.0323   1.0000   0.0372
  -5.250  -0.4394   0.04131   0.03277  -0.0322   1.0000   0.0377
  -5.000  -0.4192   0.03772   0.02862  -0.0317   1.0000   0.0378
  -4.750  -0.3974   0.03442   0.02479  -0.0310   1.0000   0.0379
  -4.500  -0.3742   0.03152   0.02137  -0.0301   1.0000   0.0383
  -4.250  -0.3500   0.02900   0.01838  -0.0291   1.0000   0.0393
  -4.000  -0.3250   0.02705   0.01592  -0.0280   1.0000   0.0421
  -3.750  -0.3019   0.02520   0.01403  -0.0272   1.0000   0.0466
  -3.500  -0.2772   0.02366   0.01226  -0.0260   1.0000   0.0505
  -3.250  -0.2522   0.02227   0.01059  -0.0249   1.0000   0.0557
  -3.000  -0.2278   0.02119   0.00946  -0.0241   1.0000   0.0707
  -2.750  -0.2035   0.01990   0.00828  -0.0233   1.0000   0.1047
  -2.500  -0.1822   0.01786   0.00740  -0.0230   1.0000   0.2977
  -2.250  -0.1397   0.01544   0.00715  -0.0224   1.0000   0.9502
  -2.000  -0.0985   0.01531   0.00644  -0.0252   1.0000   1.0000
  -1.750  -0.0803   0.01527   0.00605  -0.0237   1.0000   1.0000
  -1.500  -0.0614   0.01525   0.00570  -0.0223   1.0000   1.0000
  -1.250  -0.0420   0.01527   0.00545  -0.0210   1.0000   1.0000
  -1.000  -0.0222   0.01531   0.00526  -0.0199   1.0000   1.0000
  -0.750  -0.0020   0.01539   0.00513  -0.0188   1.0000   1.0000
  -0.500   0.0185   0.01549   0.00504  -0.0178   1.0000   1.0000
  -0.250   0.0391   0.01562   0.00503  -0.0170   1.0000   1.0000
   0.000   0.0599   0.01579   0.00508  -0.0162   1.0000   1.0000
   0.250   0.0806   0.01600   0.00519  -0.0155   1.0000   1.0000
   0.500   0.1012   0.01626   0.00535  -0.0149   1.0000   1.0000
   0.750   0.1430   0.01657   0.00560  -0.0185   0.9880   1.0000
   1.000   0.1906   0.01684   0.00585  -0.0231   0.9706   1.0000
   1.250   0.2399   0.01705   0.00608  -0.0277   0.9528   1.0000
   1.500   0.2855   0.01719   0.00628  -0.0314   0.9317   1.0000
   1.750   0.3283   0.01728   0.00646  -0.0343   0.9096   1.0000
   2.000   0.3664   0.01733   0.00660  -0.0360   0.8853   1.0000
   2.500   0.4329   0.01740   0.00690  -0.0371   0.8314   1.0000
   2.750   0.4634   0.01745   0.00704  -0.0369   0.8026   1.0000
   3.000   0.4920   0.01750   0.00719  -0.0361   0.7708   1.0000
   3.250   0.5181   0.01755   0.00724  -0.0346   0.7307   1.0000
   3.500   0.5422   0.01767   0.00729  -0.0327   0.6857   1.0000
   3.750   0.5651   0.01790   0.00745  -0.0307   0.6385   1.0000
   4.000   0.5875   0.01823   0.00773  -0.0289   0.5920   1.0000
   4.250   0.6100   0.01862   0.00806  -0.0274   0.5475   1.0000
   4.500   0.6317   0.01908   0.00844  -0.0258   0.4985   1.0000
   4.750   0.6542   0.01954   0.00895  -0.0246   0.4536   1.0000
   5.000   0.6757   0.02010   0.00948  -0.0232   0.3912   1.0000
   5.250   0.6954   0.02097   0.01002  -0.0219   0.3057   1.0000
   5.500   0.7147   0.02227   0.01087  -0.0209   0.2043   1.0000
   5.750   0.7351   0.02377   0.01203  -0.0202   0.1455   1.0000
   6.000   0.7569   0.02506   0.01333  -0.0194   0.1253   1.0000
   6.250   0.7784   0.02638   0.01478  -0.0185   0.1150   1.0000
   6.500   0.8002   0.02771   0.01633  -0.0175   0.1075   1.0000
   6.750   0.8214   0.02924   0.01802  -0.0164   0.0983   1.0000
   7.000   0.8418   0.03108   0.01995  -0.0153   0.0867   1.0000
   7.250   0.8614   0.03331   0.02230  -0.0143   0.0723   1.0000
   7.500   0.8816   0.03614   0.02527  -0.0133   0.0607   1.0000
   7.750   0.9020   0.03907   0.02847  -0.0123   0.0522   1.0000
   8.000   0.9198   0.04230   0.03222  -0.0110   0.0461   1.0000
   8.250   0.9361   0.04543   0.03585  -0.0099   0.0417   1.0000
   8.500   0.9501   0.04922   0.03981  -0.0091   0.0392   1.0000
   8.750   0.9565   0.05302   0.04436  -0.0073   0.0373   1.0000
   9.000   0.9590   0.05705   0.04898  -0.0059   0.0354   1.0000
   9.250   0.9580   0.06111   0.05347  -0.0048   0.0341   1.0000
   9.500   0.9530   0.06529   0.05799  -0.0039   0.0333   1.0000
   9.750   0.9428   0.06967   0.06267  -0.0034   0.0330   1.0000
  10.000   0.9261   0.07410   0.06731  -0.0032   0.0332   1.0000
  10.250   0.9066   0.07934   0.07273  -0.0050   0.0336   1.0000
  10.500   0.8868   0.08582   0.07933  -0.0092   0.0343   1.0000
  10.750   0.8689   0.09348   0.08704  -0.0149   0.0353   1.0000
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