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

B-29 ROOT AIRFOIL (b29root-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: B-29 ROOT AIRFOIL (b29root-il)
Reynolds number: 100,000
Max Cl/Cd: 13.81 at α=4°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-b29root-il-100000.txt
Download as CSV file: xf-b29root-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: B-29 ROOT 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
  ------ -------- --------- --------- -------- -------- --------
 -13.000  -0.3152   0.12064   0.11502  -0.0477   1.0000   0.1929
 -12.750  -0.3086   0.11929   0.11374  -0.0473   1.0000   0.2056
 -12.500  -1.0385   0.05087   0.04332  -0.0427   0.9991   0.1304
 -12.250  -0.9972   0.04866   0.04107  -0.0472   0.9918   0.1333
 -12.000  -0.9731   0.04645   0.03863  -0.0493   0.9819   0.1358
 -11.750  -0.9565   0.04444   0.03632  -0.0502   0.9712   0.1382
 -11.500  -0.9447   0.04273   0.03419  -0.0499   0.9606   0.1407
 -11.250  -0.8902   0.04024   0.03179  -0.0558   0.9556   0.1448
 -11.000  -0.8466   0.03840   0.02984  -0.0600   0.9501   0.1492
 -10.750  -0.8195   0.03692   0.02809  -0.0613   0.9412   0.1533
 -10.500  -0.7605   0.03482   0.02609  -0.0676   0.9370   0.1595
 -10.250  -0.7145   0.03324   0.02440  -0.0718   0.9318   0.1662
 -10.000  -0.6713   0.03177   0.02297  -0.0752   0.9236   0.1732
  -9.750  -0.6321   0.03054   0.02171  -0.0780   0.9159   0.1819
  -9.500  -0.5889   0.02927   0.02057  -0.0813   0.9086   0.1917
  -9.250  -0.5617   0.02834   0.01961  -0.0818   0.8983   0.2021
  -9.000  -0.5359   0.02751   0.01884  -0.0821   0.8890   0.2145
  -8.750  -0.5161   0.02678   0.01826  -0.0813   0.8786   0.2275
  -8.500  -0.5018   0.02616   0.01775  -0.0795   0.8692   0.2428
  -8.250  -0.4942   0.02566   0.01736  -0.0765   0.8586   0.2598
  -8.000  -0.4898   0.02522   0.01703  -0.0728   0.8494   0.2806
  -7.750  -0.4919   0.02494   0.01683  -0.0680   0.8390   0.3035
  -7.500  -0.4912   0.02468   0.01673  -0.0632   0.8315   0.3295
  -7.250  -0.5008   0.02487   0.01703  -0.0568   0.8206   0.3514
  -7.000  -0.5010   0.02484   0.01702  -0.0514   0.8135   0.3779
  -6.750  -0.5065   0.02517   0.01740  -0.0456   0.8038   0.3985
  -6.500  -0.5034   0.02530   0.01751  -0.0409   0.7961   0.4203
  -6.250  -0.4920   0.02532   0.01748  -0.0374   0.7910   0.4422
  -6.000  -0.4906   0.02582   0.01810  -0.0330   0.7807   0.4575
  -5.750  -0.4744   0.02587   0.01816  -0.0304   0.7743   0.4761
  -5.500  -0.4516   0.02586   0.01812  -0.0288   0.7699   0.4949
  -5.250  -0.4577   0.02643   0.01871  -0.0234   0.7594   0.5072
  -5.000  -0.4420   0.02647   0.01870  -0.0209   0.7536   0.5241
  -4.750  -0.4099   0.02640   0.01865  -0.0208   0.7495   0.5412
  -4.500  -0.4090   0.02703   0.01936  -0.0166   0.7395   0.5525
  -4.250  -0.3951   0.02707   0.01935  -0.0139   0.7334   0.5677
  -4.000  -0.3630   0.02701   0.01928  -0.0139   0.7295   0.5838
  -3.750  -0.3519   0.02757   0.01992  -0.0113   0.7220   0.5959
  -3.500  -0.3512   0.02792   0.02024  -0.0068   0.7139   0.6093
  -3.250  -0.3176   0.02784   0.02016  -0.0070   0.7098   0.6263
  -3.000  -0.2781   0.02785   0.02020  -0.0079   0.7067   0.6456
  -2.750  -0.2963   0.02918   0.02167  -0.0013   0.6954   0.6566
  -2.500  -0.2725   0.02938   0.02188   0.0000   0.6905   0.6746
  -2.250  -0.2404   0.02919   0.02163   0.0002   0.6873   0.6919
  -2.000  -0.1790   0.02910   0.02153  -0.0047   0.6849   0.7047
  -1.750  -0.2302   0.03080   0.02337   0.0066   0.6713   0.7165
  -1.500  -0.2000   0.03073   0.02326   0.0068   0.6678   0.7325
  -1.250  -0.1273   0.03069   0.02319   0.0001   0.6659   0.7440
  -1.000  -0.1910   0.03281   0.02546   0.0129   0.6529   0.7586
  -0.750  -0.1657   0.03292   0.02553   0.0137   0.6484   0.7754
  -0.500  -0.0701   0.03296   0.02554   0.0034   0.6467   0.7840
  -0.250  -0.0183   0.03281   0.02530   0.0004   0.6447   0.7979
   0.000  -0.1268   0.03648   0.02920   0.0184   0.6292   0.8181
   0.250  -0.0129   0.03648   0.02910   0.0054   0.6278   0.8226
   0.500   0.0845   0.03628   0.02879  -0.0048   0.6262   0.8274
   0.750   0.1442   0.03592   0.02834  -0.0090   0.6244   0.8383
   1.000   0.2384   0.03580   0.02810  -0.0191   0.6230   0.8402
   1.250   0.1498   0.04076   0.03332  -0.0055   0.6075   0.8612
   1.500   0.2427   0.04016   0.03262  -0.0147   0.6060   0.8643
   1.750   0.2912   0.03954   0.03191  -0.0170   0.6043   0.8765
   2.000   0.3785   0.03887   0.03114  -0.0258   0.6030   0.8792
   2.250   0.2788   0.04540   0.03793  -0.0128   0.5864   0.9020
   2.500   0.3627   0.04430   0.03676  -0.0204   0.5853   0.9079
   2.750   0.4305   0.04324   0.03563  -0.0256   0.5841   0.9157
   3.000   0.5140   0.04202   0.03433  -0.0336   0.5831   0.9198
   3.250   0.3816   0.05059   0.04314  -0.0190   0.5651   0.9489
   3.500   0.4766   0.04869   0.04120  -0.0276   0.5645   0.9517
   3.750   0.5655   0.04687   0.03933  -0.0357   0.5639   0.9560
   4.000   0.6298   0.04559   0.03800  -0.0404   0.5631   0.9641
   4.250   0.4328   0.06184   0.05457  -0.0298   0.5363   0.9925
   4.500   0.4645   0.06366   0.05642  -0.0344   0.5311   1.0000
   4.750   0.4751   0.06364   0.05635  -0.0324   0.5280   1.0000
   5.000   0.5050   0.06268   0.05534  -0.0315   0.5259   1.0000
   5.250   0.5375   0.06179   0.05441  -0.0309   0.5241   1.0000
   5.500   0.3797   0.07092   0.06355  -0.0183   0.5146   1.0000
   5.750   0.3806   0.07174   0.06433  -0.0157   0.5108   1.0000
   6.000   0.4148   0.07094   0.06349  -0.0149   0.5074   1.0000
   6.250   0.4625   0.06969   0.06220  -0.0150   0.5052   1.0000
   6.500   0.3645   0.07652   0.06903  -0.0077   0.5016   1.0000
   6.750   0.3409   0.07932   0.07181  -0.0046   0.4997   1.0000
   7.000   0.3314   0.08184   0.07433  -0.0026   0.5005   1.0000
   7.250   0.3344   0.08418   0.07665  -0.0013   0.5026   1.0000
   7.500   0.3420   0.08642   0.07889  -0.0004   0.5039   1.0000
   7.750   0.3624   0.08852   0.08099  -0.0002   0.5050   1.0000
   8.000   0.3720   0.08741   0.07984   0.0026   0.4876   1.0000
   8.250   0.3726   0.08997   0.08240   0.0037   0.4873   1.0000
   8.500   0.3741   0.09259   0.08503   0.0048   0.4859   1.0000
   8.750   0.3737   0.09174   0.08414   0.0079   0.4652   1.0000
   9.000   0.3645   0.09470   0.08712   0.0091   0.4648   1.0000
   9.250   0.3703   0.09747   0.08991   0.0096   0.4659   1.0000
   9.500   0.3752   0.09783   0.09026   0.0113   0.4529   1.0000
   9.750   0.3740   0.10069   0.09314   0.0119   0.4526   1.0000
  10.000   0.4080   0.09773   0.09013   0.0144   0.4225   1.0000
  10.250   0.4578   0.09613   0.08852   0.0152   0.4150   1.0000
  10.500   0.4344   0.09961   0.09202   0.0165   0.4050   1.0000
  10.750   0.4698   0.09891   0.09134   0.0173   0.3981   1.0000
  11.000   0.4614   0.10149   0.09393   0.0184   0.3881   1.0000
  11.250   0.4867   0.10153   0.09399   0.0192   0.3812   1.0000
  11.500   0.5439   0.09868   0.09115   0.0204   0.3779   1.0000
  11.750   0.5116   0.10337   0.09587   0.0211   0.3640   1.0000
  12.000   0.5613   0.10083   0.09335   0.0224   0.3609   1.0000
  12.250   0.5359   0.10527   0.09782   0.0228   0.3471   1.0000
  12.500   0.5840   0.10241   0.09498   0.0243   0.3440   1.0000
<< Back to B-29 ROOT AIRFOIL (b29root-il)

Polar data table (+)

Polar graphs


<< Back to B-29 ROOT AIRFOIL (b29root-il)