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

DAE-51 AIRFOIL (dae51-il) Xfoil prediction polar at RE=1,000,000 Ncrit=9


Details Polar file
Airfoil: DAE-51 AIRFOIL (dae51-il)
Reynolds number: 1,000,000
Max Cl/Cd: 144.36 at α=3.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-dae51-il-1000000.txt
Download as CSV file: xf-dae51-il-1000000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: DAE-51 AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     1.000 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.2599   0.09814   0.09661  -0.0394   0.9753   0.0116
 -10.000  -0.2613   0.09312   0.09158  -0.0404   0.9620   0.0119
  -9.000  -0.3580   0.09210   0.09054  -0.0355   0.9802   0.0119
  -8.750  -0.3489   0.08914   0.08755  -0.0373   0.9590   0.0120
  -8.500  -0.3441   0.08662   0.08498  -0.0377   0.9408   0.0122
  -8.250  -0.3391   0.08389   0.08221  -0.0387   0.9266   0.0124
  -8.000  -0.3338   0.08087   0.07915  -0.0403   0.9144   0.0126
  -7.750  -0.3281   0.07754   0.07579  -0.0426   0.9033   0.0129
  -7.500  -0.3210   0.07385   0.07207  -0.0462   0.8936   0.0134
  -7.250  -0.3088   0.06908   0.06725  -0.0523   0.8848   0.0140
  -7.000  -0.2855   0.06063   0.05875  -0.0656   0.8762   0.0152
  -5.750  -0.1782   0.01518   0.01068  -0.0993   0.8448   0.0124
  -5.500  -0.1510   0.01414   0.00944  -0.0994   0.8396   0.0128
  -5.250  -0.1231   0.01328   0.00844  -0.0996   0.8344   0.0132
  -5.000  -0.0952   0.01240   0.00740  -0.0997   0.8289   0.0135
  -4.750  -0.0673   0.01165   0.00650  -0.0998   0.8238   0.0139
  -4.500  -0.0389   0.01094   0.00568  -0.0999   0.8186   0.0142
  -4.250  -0.0106   0.01034   0.00498  -0.1000   0.8132   0.0145
  -4.000   0.0176   0.00987   0.00440  -0.1001   0.8080   0.0149
  -3.750   0.0464   0.00962   0.00410  -0.1002   0.8025   0.0153
  -3.500   0.0747   0.00877   0.00315  -0.1004   0.7971   0.0166
  -3.250   0.1033   0.00849   0.00283  -0.1006   0.7918   0.0176
  -3.000   0.1322   0.00825   0.00255  -0.1007   0.7859   0.0190
  -2.750   0.1609   0.00806   0.00230  -0.1008   0.7802   0.0201
  -2.500   0.1899   0.00773   0.00194  -0.1010   0.7745   0.0232
  -2.250   0.2187   0.00761   0.00178  -0.1012   0.7684   0.0258
  -2.000   0.2475   0.00741   0.00157  -0.1013   0.7626   0.0330
  -1.750   0.2765   0.00719   0.00140  -0.1016   0.7560   0.0505
  -1.500   0.3052   0.00698   0.00129  -0.1018   0.7496   0.0904
  -1.250   0.3341   0.00669   0.00122  -0.1021   0.7427   0.1603
  -1.000   0.3627   0.00643   0.00116  -0.1024   0.7358   0.2470
  -0.750   0.3915   0.00614   0.00114  -0.1028   0.7285   0.3486
  -0.500   0.4199   0.00586   0.00114  -0.1031   0.7209   0.4609
  -0.250   0.4484   0.00559   0.00116  -0.1034   0.7129   0.5731
   0.000   0.4763   0.00537   0.00119  -0.1034   0.7047   0.6834
   0.250   0.5033   0.00512   0.00123  -0.1032   0.6958   0.7910
   0.500   0.5273   0.00471   0.00119  -0.1019   0.6869   1.0000
   0.750   0.5559   0.00478   0.00119  -0.1020   0.6770   1.0000
   1.000   0.5846   0.00485   0.00120  -0.1022   0.6665   1.0000
   1.250   0.6132   0.00493   0.00122  -0.1023   0.6558   1.0000
   1.500   0.6416   0.00502   0.00124  -0.1024   0.6444   1.0000
   1.750   0.6699   0.00512   0.00128  -0.1025   0.6321   1.0000
   2.000   0.6983   0.00522   0.00133  -0.1026   0.6193   1.0000
   2.250   0.7265   0.00533   0.00138  -0.1027   0.6060   1.0000
   2.500   0.7546   0.00545   0.00144  -0.1027   0.5925   1.0000
   2.750   0.7826   0.00558   0.00151  -0.1028   0.5785   1.0000
   3.000   0.8106   0.00571   0.00160  -0.1029   0.5642   1.0000
   3.250   0.8384   0.00586   0.00169  -0.1029   0.5495   1.0000
   3.500   0.8661   0.00602   0.00179  -0.1029   0.5341   1.0000
   3.750   0.8936   0.00619   0.00191  -0.1029   0.5177   1.0000
   4.000   0.9209   0.00639   0.00203  -0.1029   0.4998   1.0000
   4.250   0.9482   0.00658   0.00217  -0.1029   0.4810   1.0000
   4.500   0.9750   0.00682   0.00232  -0.1028   0.4581   1.0000
   4.750   1.0013   0.00711   0.00251  -0.1026   0.4295   1.0000
   5.000   1.0277   0.00739   0.00269  -0.1025   0.4056   1.0000
   5.250   1.0533   0.00775   0.00291  -0.1023   0.3736   1.0000
   5.500   1.0779   0.00823   0.00320  -0.1019   0.3318   1.0000
   5.750   1.1027   0.00866   0.00349  -0.1016   0.2985   1.0000
   6.000   1.1271   0.00914   0.00380  -0.1012   0.2651   1.0000
   6.250   1.1510   0.00966   0.00415  -0.1007   0.2291   1.0000
   6.500   1.1745   0.01020   0.00453  -0.1003   0.1963   1.0000
   6.750   1.1977   0.01076   0.00493  -0.0997   0.1642   1.0000
   7.000   1.2200   0.01140   0.00539  -0.0991   0.1306   1.0000
   7.250   1.2418   0.01207   0.00588  -0.0983   0.0999   1.0000
   7.500   1.2634   0.01274   0.00640  -0.0975   0.0740   1.0000
   7.750   1.2849   0.01338   0.00692  -0.0967   0.0544   1.0000
   8.000   1.3068   0.01397   0.00744  -0.0960   0.0410   1.0000
   8.250   1.3284   0.01455   0.00798  -0.0952   0.0316   1.0000
   8.500   1.3496   0.01516   0.00856  -0.0943   0.0252   1.0000
   8.750   1.3712   0.01567   0.00907  -0.0935   0.0221   1.0000
   9.000   1.3914   0.01632   0.00974  -0.0924   0.0192   1.0000
   9.250   1.4128   0.01680   0.01026  -0.0915   0.0178   1.0000
   9.500   1.4323   0.01741   0.01089  -0.0904   0.0163   1.0000
   9.750   1.4488   0.01826   0.01181  -0.0888   0.0147   1.0000
  10.000   1.4685   0.01877   0.01236  -0.0878   0.0139   1.0000
  10.250   1.4864   0.01937   0.01300  -0.0865   0.0130   1.0000
  10.500   1.5006   0.02008   0.01376  -0.0845   0.0122   1.0000
  10.750   1.5078   0.02122   0.01497  -0.0816   0.0113   1.0000
  11.000   1.5180   0.02219   0.01603  -0.0793   0.0109   1.0000
  11.250   1.5311   0.02299   0.01691  -0.0776   0.0106   1.0000
  11.500   1.5427   0.02393   0.01792  -0.0757   0.0102   1.0000
  11.750   1.5534   0.02498   0.01903  -0.0739   0.0098   1.0000
  12.000   1.5634   0.02611   0.02023  -0.0722   0.0094   1.0000
  12.250   1.5726   0.02734   0.02153  -0.0705   0.0091   1.0000
  12.500   1.5795   0.02880   0.02306  -0.0687   0.0087   1.0000
  12.750   1.5794   0.03092   0.02527  -0.0665   0.0084   1.0000
  13.000   1.5723   0.03375   0.02826  -0.0641   0.0080   1.0000
  13.250   1.5816   0.03519   0.02977  -0.0630   0.0079   1.0000
  13.500   1.5876   0.03698   0.03167  -0.0618   0.0078   1.0000
  13.750   1.5915   0.03899   0.03378  -0.0606   0.0076   1.0000
  14.000   1.5941   0.04121   0.03610  -0.0595   0.0075   1.0000
  14.250   1.5956   0.04362   0.03861  -0.0585   0.0073   1.0000
  14.500   1.5959   0.04619   0.04129  -0.0577   0.0072   1.0000
  14.750   1.5957   0.04891   0.04411  -0.0570   0.0070   1.0000
  15.000   1.5939   0.05190   0.04721  -0.0565   0.0069   1.0000
  15.250   1.5915   0.05507   0.05049  -0.0561   0.0068   1.0000
  15.500   1.5889   0.05840   0.05392  -0.0561   0.0066   1.0000
  15.750   1.5850   0.06203   0.05766  -0.0562   0.0065   1.0000
  16.000   1.5806   0.06585   0.06158  -0.0566   0.0064   1.0000
  16.250   1.5745   0.06999   0.06583  -0.0571   0.0063   1.0000
  16.500   1.5670   0.07448   0.07043  -0.0579   0.0063   1.0000
  16.750   1.5580   0.07932   0.07540  -0.0590   0.0062   1.0000
  17.000   1.5475   0.08447   0.08066  -0.0603   0.0061   1.0000
  17.250   1.5355   0.08997   0.08629  -0.0618   0.0061   1.0000
  17.500   1.5226   0.09577   0.09221  -0.0636   0.0060   1.0000
  17.750   1.5083   0.10193   0.09850  -0.0657   0.0060   1.0000
  18.000   1.4928   0.10836   0.10505  -0.0681   0.0059   1.0000
  18.250   1.4772   0.11497   0.11180  -0.0707   0.0059   1.0000
  18.500   1.4619   0.12174   0.11870  -0.0736   0.0058   1.0000
  18.750   1.4448   0.12895   0.12605  -0.0769   0.0058   1.0000
  19.000   1.4298   0.13595   0.13318  -0.0803   0.0058   1.0000
<< Back to DAE-51 AIRFOIL (dae51-il)

Polar data table (+)

Polar graphs


<< Back to DAE-51 AIRFOIL (dae51-il)