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

DELFT DU84-132V3 AIRFOIL (MEASURED) (du84132v-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: DELFT DU84-132V3 AIRFOIL (MEASURED) (du84132v-il)
Reynolds number: 100,000
Max Cl/Cd: 53.62 at α=7.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-du84132v-il-100000.txt
Download as CSV file: xf-du84132v-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DELFT DU84-132V3 AIRFOIL (MEASURED)             
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -9.000  -0.3719   0.10064   0.09577  -0.0359   1.0000   0.1533
  -8.750  -0.3811   0.09861   0.09382  -0.0346   1.0000   0.1576
  -8.500  -0.4638   0.09780   0.09327  -0.0359   1.0000   0.1624
  -8.250  -0.4478   0.09383   0.08932  -0.0334   1.0000   0.1645
  -8.000  -0.4273   0.09165   0.08714  -0.0302   1.0000   0.1670
  -7.750  -0.4298   0.08986   0.08541  -0.0276   1.0000   0.1703
  -7.500  -0.4503   0.08813   0.08377  -0.0254   1.0000   0.1743
  -7.250  -0.5717   0.08400   0.07973  -0.0314   1.0000   0.1800
  -7.000  -0.5394   0.08173   0.07760  -0.0258   1.0000   0.1818
  -6.750  -0.5277   0.08037   0.07630  -0.0216   1.0000   0.1844
  -6.500  -0.5339   0.07806   0.07402  -0.0204   1.0000   0.1888
  -6.250  -0.5627   0.07324   0.06915  -0.0238   1.0000   0.1988
  -6.000  -0.5473   0.04407   0.03740  -0.0475   0.9932   0.0865
  -5.750  -0.5155   0.03996   0.03291  -0.0503   0.9879   0.0816
  -5.500  -0.4751   0.03697   0.02877  -0.0534   0.9825   0.0764
  -5.250  -0.4410   0.03490   0.02636  -0.0555   0.9767   0.0764
  -5.000  -0.4057   0.03287   0.02405  -0.0576   0.9710   0.0768
  -4.750  -0.3669   0.03133   0.02227  -0.0599   0.9662   0.0770
  -4.500  -0.3365   0.03000   0.02080  -0.0607   0.9589   0.0773
  -4.250  -0.2979   0.02878   0.01954  -0.0628   0.9539   0.0785
  -4.000  -0.2678   0.02788   0.01865  -0.0635   0.9467   0.0805
  -3.750  -0.2324   0.02711   0.01788  -0.0650   0.9404   0.0837
  -3.500  -0.1969   0.02651   0.01721  -0.0664   0.9343   0.0883
  -3.250  -0.1670   0.02561   0.01650  -0.0673   0.9264   0.0976
  -3.000  -0.1244   0.02469   0.01583  -0.0703   0.9217   0.1302
  -2.750  -0.1058   0.02256   0.01536  -0.0699   0.9128   0.4054
  -2.500  -0.0825   0.02239   0.01635  -0.0674   0.9067   0.6157
  -2.250  -0.0663   0.02322   0.01724  -0.0640   0.8960   0.7037
  -2.000  -0.0360   0.02425   0.01818  -0.0621   0.8896   0.7608
  -1.750  -0.0253   0.02476   0.01863  -0.0578   0.8781   0.7907
  -1.500   0.0053   0.02484   0.01861  -0.0565   0.8728   0.8196
  -1.250   0.0127   0.02490   0.01861  -0.0520   0.8610   0.8443
  -1.000   0.0283   0.02467   0.01833  -0.0481   0.8530   0.8688
  -0.750   0.0468   0.02442   0.01800  -0.0454   0.8441   0.8903
  -0.500   0.0653   0.02421   0.01772  -0.0428   0.8354   0.9108
  -0.250   0.0974   0.02396   0.01739  -0.0426   0.8286   0.9349
   0.000   0.1905   0.02371   0.01700  -0.0531   0.8271   0.9592
   0.250   0.3377   0.02278   0.01586  -0.0736   0.8280   0.9740
   0.500   0.4235   0.02181   0.01478  -0.0838   0.8257   0.9808
   0.750   0.4700   0.02163   0.01456  -0.0880   0.8158   0.9881
   1.000   0.5306   0.02093   0.01380  -0.0940   0.8114   0.9927
   1.250   0.5741   0.02069   0.01354  -0.0975   0.8017   0.9990
   1.500   0.6190   0.02008   0.01289  -0.1004   0.7954   1.0000
   1.750   0.6305   0.02015   0.01295  -0.0977   0.7831   1.0000
   2.000   0.6510   0.02001   0.01281  -0.0964   0.7727   1.0000
   2.250   0.6921   0.01938   0.01213  -0.0984   0.7656   1.0000
   2.500   0.7103   0.01932   0.01206  -0.0966   0.7541   1.0000
   2.750   0.7278   0.01929   0.01204  -0.0947   0.7422   1.0000
   3.000   0.7456   0.01925   0.01200  -0.0928   0.7301   1.0000
   3.250   0.7679   0.01910   0.01184  -0.0916   0.7183   1.0000
   3.500   0.7932   0.01890   0.01163  -0.0909   0.7064   1.0000
   3.750   0.8127   0.01883   0.01155  -0.0891   0.6931   1.0000
   4.000   0.8284   0.01879   0.01150  -0.0867   0.6785   1.0000
   4.250   0.8487   0.01865   0.01134  -0.0850   0.6636   1.0000
   4.500   0.8656   0.01861   0.01128  -0.0828   0.6476   1.0000
   4.750   0.8832   0.01857   0.01122  -0.0807   0.6309   1.0000
   5.000   0.9067   0.01848   0.01106  -0.0795   0.6135   1.0000
   5.250   0.9303   0.01848   0.01099  -0.0785   0.5948   1.0000
   5.500   0.9490   0.01864   0.01113  -0.0768   0.5750   1.0000
   5.750   0.9726   0.01884   0.01129  -0.0761   0.5562   1.0000
   6.000   0.9964   0.01913   0.01155  -0.0755   0.5378   1.0000
   6.250   1.0208   0.01946   0.01186  -0.0750   0.5202   1.0000
   6.500   1.0453   0.01980   0.01216  -0.0745   0.5030   1.0000
   6.750   1.0687   0.02017   0.01250  -0.0739   0.4856   1.0000
   7.000   1.0899   0.02051   0.01284  -0.0729   0.4677   1.0000
   7.250   1.1110   0.02080   0.01315  -0.0719   0.4508   1.0000
   7.500   1.1336   0.02114   0.01345  -0.0712   0.4353   1.0000
   7.750   1.1547   0.02156   0.01385  -0.0703   0.4188   1.0000
   8.000   1.1753   0.02203   0.01428  -0.0693   0.4025   1.0000
   8.250   1.1968   0.02259   0.01478  -0.0685   0.3864   1.0000
   8.500   1.2167   0.02324   0.01541  -0.0675   0.3697   1.0000
   8.750   1.2338   0.02390   0.01610  -0.0660   0.3528   1.0000
   9.000   1.2475   0.02453   0.01681  -0.0640   0.3353   1.0000
   9.250   1.2575   0.02517   0.01750  -0.0614   0.3166   1.0000
   9.500   1.2642   0.02589   0.01822  -0.0583   0.2967   1.0000
   9.750   1.2691   0.02677   0.01907  -0.0551   0.2756   1.0000
  10.000   1.2751   0.02783   0.01998  -0.0522   0.2550   1.0000
  10.250   1.2789   0.02905   0.02121  -0.0492   0.2338   1.0000
  10.500   1.2833   0.03024   0.02239  -0.0465   0.2160   1.0000
  10.750   1.2890   0.03147   0.02358  -0.0441   0.2012   1.0000
  11.000   1.2954   0.03281   0.02486  -0.0421   0.1882   1.0000
  11.250   1.3019   0.03425   0.02624  -0.0401   0.1763   1.0000
  11.500   1.3086   0.03576   0.02766  -0.0384   0.1653   1.0000
  11.750   1.3146   0.03735   0.02934  -0.0367   0.1551   1.0000
  12.000   1.3241   0.03910   0.03104  -0.0354   0.1453   1.0000
  12.250   1.3380   0.04096   0.03279  -0.0345   0.1353   1.0000
  12.500   1.3490   0.04281   0.03463  -0.0335   0.1263   1.0000
  12.750   1.3545   0.04489   0.03690  -0.0320   0.1189   1.0000
  13.000   1.3719   0.04664   0.03849  -0.0316   0.1113   1.0000
  13.250   1.3725   0.04892   0.04110  -0.0299   0.1062   1.0000
  13.500   1.3977   0.05074   0.04269  -0.0300   0.0988   1.0000
  13.750   1.3929   0.05343   0.04575  -0.0282   0.0951   1.0000
  14.000   1.3973   0.05587   0.04832  -0.0272   0.0906   1.0000
  14.250   1.4222   0.05866   0.05105  -0.0274   0.0851   1.0000
  14.500   1.4131   0.06194   0.05468  -0.0259   0.0832   1.0000
  14.750   1.4052   0.06546   0.05851  -0.0249   0.0812   1.0000
  15.000   1.3988   0.06903   0.06234  -0.0242   0.0792   1.0000
  15.250   1.3918   0.07268   0.06618  -0.0239   0.0773   1.0000
  15.750   1.3930   0.08072   0.07440  -0.0240   0.0729   1.0000
  16.000   1.3684   0.08599   0.07999  -0.0246   0.0726   1.0000
  16.250   1.3416   0.09199   0.08629  -0.0261   0.0724   1.0000
  16.500   1.3131   0.09878   0.09336  -0.0285   0.0724   1.0000
  16.750   1.2824   0.10651   0.10135  -0.0320   0.0725   1.0000
  17.000   1.2491   0.11540   0.11047  -0.0366   0.0728   1.0000
  17.250   1.2143   0.12548   0.12073  -0.0424   0.0732   1.0000
<< Back to DELFT DU84-132V3 AIRFOIL (MEASURED) (du84132v-il)

Polar data table (+)

Polar graphs


<< Back to DELFT DU84-132V3 AIRFOIL (MEASURED) (du84132v-il)