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

BOEING 103 AIRFOIL (boe103-il) Xfoil prediction polar at RE=500,000 Ncrit=9


Details Polar file
Airfoil: BOEING 103 AIRFOIL (boe103-il)
Reynolds number: 500,000
Max Cl/Cd: 94.57 at α=3.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-boe103-il-500000.txt
Download as CSV file: xf-boe103-il-500000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: BOEING 103 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 =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -11.750  -0.2866   0.11058   0.10831  -0.0414   1.0000   0.0303
 -11.500  -0.3957   0.10895   0.10649  -0.0416   1.0000   0.0298
 -11.250  -0.3886   0.10705   0.10460  -0.0411   1.0000   0.0301
 -11.000  -0.3829   0.10487   0.10244  -0.0410   1.0000   0.0305
 -10.750  -0.3793   0.10230   0.09988  -0.0412   1.0000   0.0310
 -10.500  -0.3779   0.09934   0.09694  -0.0416   1.0000   0.0316
 -10.250  -0.3796   0.09586   0.09350  -0.0424   1.0000   0.0324
  -9.250  -0.5685   0.03950   0.03586  -0.0786   0.9838   0.0296
  -9.000  -0.5570   0.03471   0.03065  -0.0798   0.9773   0.0297
  -8.750  -0.5373   0.03022   0.02579  -0.0815   0.9737   0.0300
  -8.500  -0.5094   0.02737   0.02270  -0.0834   0.9717   0.0304
  -8.250  -0.4825   0.02625   0.02161  -0.0840   0.9669   0.0311
  -8.000  -0.4543   0.02440   0.01955  -0.0850   0.9631   0.0316
  -7.750  -0.4225   0.02288   0.01783  -0.0865   0.9606   0.0324
  -7.500  -0.3903   0.02137   0.01608  -0.0879   0.9584   0.0333
  -7.250  -0.3655   0.02006   0.01455  -0.0875   0.9518   0.0337
  -7.000  -0.3353   0.01887   0.01316  -0.0881   0.9474   0.0342
  -6.750  -0.3028   0.01801   0.01212  -0.0891   0.9440   0.0347
  -6.500  -0.2776   0.01688   0.01083  -0.0887   0.9364   0.0354
  -6.250  -0.2484   0.01548   0.00936  -0.0891   0.9311   0.0361
  -6.000  -0.2204   0.01464   0.00849  -0.0891   0.9244   0.0366
  -5.750  -0.1926   0.01396   0.00777  -0.0891   0.9171   0.0373
  -5.500  -0.1645   0.01338   0.00713  -0.0890   0.9099   0.0379
  -5.250  -0.1377   0.01289   0.00660  -0.0887   0.9015   0.0388
  -5.000  -0.1104   0.01246   0.00611  -0.0884   0.8938   0.0397
  -4.750  -0.0841   0.01204   0.00564  -0.0880   0.8851   0.0404
  -4.500  -0.0575   0.01167   0.00521  -0.0875   0.8766   0.0409
  -4.250  -0.0309   0.01136   0.00483  -0.0871   0.8679   0.0414
  -4.000  -0.0057   0.01087   0.00432  -0.0865   0.8589   0.0424
  -3.750   0.0207   0.01051   0.00391  -0.0861   0.8503   0.0439
  -3.500   0.0467   0.01026   0.00364  -0.0856   0.8406   0.0456
  -3.250   0.0739   0.01006   0.00337  -0.0852   0.8315   0.0474
  -3.000   0.1002   0.00988   0.00314  -0.0847   0.8201   0.0490
  -2.750   0.1264   0.00966   0.00290  -0.0841   0.8085   0.0534
  -2.500   0.1528   0.00944   0.00270  -0.0837   0.7974   0.0670
  -2.250   0.1780   0.00907   0.00257  -0.0831   0.7853   0.1344
  -2.000   0.2038   0.00885   0.00247  -0.0826   0.7726   0.1767
  -1.750   0.2297   0.00863   0.00236  -0.0822   0.7608   0.2215
  -1.500   0.2543   0.00822   0.00226  -0.0817   0.7503   0.3251
  -1.250   0.2775   0.00771   0.00222  -0.0809   0.7388   0.4784
  -1.000   0.3015   0.00741   0.00221  -0.0801   0.7277   0.5767
  -0.750   0.3258   0.00721   0.00222  -0.0791   0.7174   0.6610
  -0.500   0.3498   0.00704   0.00223  -0.0781   0.7063   0.7328
  -0.250   0.3729   0.00687   0.00229  -0.0767   0.6954   0.8072
   0.000   0.3968   0.00682   0.00235  -0.0753   0.6851   0.8700
   0.250   0.4238   0.00686   0.00244  -0.0746   0.6745   0.9231
   0.500   0.4603   0.00698   0.00252  -0.0759   0.6624   0.9597
   0.750   0.5004   0.00711   0.00257  -0.0783   0.6491   0.9773
   1.000   0.5437   0.00724   0.00259  -0.0815   0.6365   0.9861
   1.250   0.5869   0.00735   0.00262  -0.0848   0.6235   0.9940
   1.500   0.6327   0.00745   0.00262  -0.0887   0.6077   1.0000
   1.750   0.6553   0.00753   0.00264  -0.0877   0.5937   1.0000
   2.000   0.6779   0.00763   0.00267  -0.0866   0.5798   1.0000
   2.250   0.7004   0.00774   0.00271  -0.0855   0.5655   1.0000
   2.500   0.7227   0.00786   0.00276  -0.0844   0.5507   1.0000
   2.750   0.7447   0.00799   0.00282  -0.0831   0.5336   1.0000
   3.000   0.7662   0.00816   0.00289  -0.0818   0.5142   1.0000
   3.500   0.8086   0.00855   0.00308  -0.0791   0.4705   1.0000
   3.750   0.8294   0.00878   0.00321  -0.0777   0.4474   1.0000
   4.000   0.8504   0.00904   0.00335  -0.0763   0.4229   1.0000
   4.250   0.8708   0.00935   0.00353  -0.0749   0.3978   1.0000
   4.500   0.8919   0.00966   0.00372  -0.0736   0.3737   1.0000
   4.750   0.9130   0.01000   0.00394  -0.0724   0.3523   1.0000
   5.000   0.9347   0.01033   0.00416  -0.0713   0.3343   1.0000
   5.250   0.9570   0.01063   0.00439  -0.0703   0.3186   1.0000
   5.500   0.9798   0.01092   0.00462  -0.0694   0.3042   1.0000
   5.750   1.0028   0.01120   0.00485  -0.0686   0.2899   1.0000
   6.000   1.0257   0.01150   0.00509  -0.0677   0.2751   1.0000
   6.250   1.0482   0.01181   0.00534  -0.0669   0.2592   1.0000
   6.500   1.0698   0.01218   0.00562  -0.0659   0.2408   1.0000
   6.750   1.0910   0.01259   0.00593  -0.0648   0.2179   1.0000
   7.000   1.1101   0.01312   0.00631  -0.0635   0.1900   1.0000
   7.250   1.1279   0.01374   0.00677  -0.0619   0.1633   1.0000
   7.500   1.1459   0.01434   0.00724  -0.0604   0.1429   1.0000
   7.750   1.1646   0.01487   0.00770  -0.0590   0.1305   1.0000
   8.000   1.1834   0.01537   0.00816  -0.0576   0.1228   1.0000
   8.250   1.2020   0.01580   0.00859  -0.0562   0.1176   1.0000
   8.500   1.2187   0.01633   0.00908  -0.0544   0.1127   1.0000
   8.750   1.2371   0.01676   0.00955  -0.0530   0.1094   1.0000
   9.000   1.2557   0.01719   0.01000  -0.0517   0.1063   1.0000
   9.250   1.2731   0.01769   0.01051  -0.0502   0.1034   1.0000
   9.500   1.2888   0.01829   0.01111  -0.0485   0.1004   1.0000
   9.750   1.3046   0.01889   0.01174  -0.0469   0.0977   1.0000
  10.000   1.3233   0.01933   0.01224  -0.0457   0.0957   1.0000
  10.250   1.3406   0.01985   0.01281  -0.0444   0.0936   1.0000
  10.500   1.3567   0.02046   0.01345  -0.0430   0.0914   1.0000
  10.750   1.3709   0.02119   0.01420  -0.0415   0.0891   1.0000
  11.000   1.3820   0.02213   0.01517  -0.0396   0.0864   1.0000
  11.250   1.4003   0.02264   0.01575  -0.0387   0.0846   1.0000
  11.500   1.4175   0.02322   0.01640  -0.0377   0.0824   1.0000
  11.750   1.4328   0.02395   0.01717  -0.0366   0.0801   1.0000
  12.000   1.4461   0.02482   0.01808  -0.0353   0.0777   1.0000
  12.250   1.4558   0.02598   0.01925  -0.0337   0.0750   1.0000
  12.500   1.4731   0.02662   0.01997  -0.0330   0.0730   1.0000
  12.750   1.4890   0.02736   0.02079  -0.0322   0.0701   1.0000
  13.000   1.5020   0.02835   0.02178  -0.0312   0.0666   1.0000
  13.250   1.5140   0.02943   0.02290  -0.0301   0.0635   1.0000
  13.500   1.5275   0.03041   0.02394  -0.0293   0.0592   1.0000
  13.750   1.5377   0.03169   0.02521  -0.0283   0.0539   1.0000
  14.000   1.5440   0.03333   0.02680  -0.0272   0.0457   1.0000
  14.250   1.5469   0.03531   0.02876  -0.0260   0.0403   1.0000
  14.500   1.5494   0.03740   0.03088  -0.0250   0.0365   1.0000
  14.750   1.5510   0.03963   0.03316  -0.0241   0.0333   1.0000
  15.000   1.5514   0.04206   0.03564  -0.0233   0.0303   1.0000
  15.250   1.5531   0.04447   0.03812  -0.0227   0.0272   1.0000
  15.750   1.5415   0.05112   0.04482  -0.0219   0.0161   1.0000
  16.000   1.5341   0.05485   0.04862  -0.0219   0.0150   1.0000
  16.250   1.5288   0.05845   0.05232  -0.0220   0.0142   1.0000
  16.500   1.5235   0.06215   0.05614  -0.0223   0.0139   1.0000
  16.750   1.5167   0.06616   0.06029  -0.0229   0.0136   1.0000
  17.000   1.5085   0.07047   0.06473  -0.0237   0.0133   1.0000
  17.250   1.4983   0.07514   0.06955  -0.0247   0.0131   1.0000
  17.500   1.4864   0.08020   0.07475  -0.0260   0.0129   1.0000
  17.750   1.4729   0.08564   0.08033  -0.0276   0.0128   1.0000
  18.000   1.4582   0.09137   0.08620  -0.0294   0.0127   1.0000
  18.250   1.4426   0.09738   0.09236  -0.0315   0.0126   1.0000
  18.500   1.4258   0.10373   0.09885  -0.0338   0.0125   1.0000
  18.750   1.4083   0.11032   0.10559  -0.0365   0.0124   1.0000
  19.000   1.3904   0.11707   0.11247  -0.0393   0.0123   1.0000
  19.250   1.3729   0.12388   0.11941  -0.0423   0.0123   1.0000
<< Back to BOEING 103 AIRFOIL (boe103-il)

Polar data table (+)

Polar graphs


<< Back to BOEING 103 AIRFOIL (boe103-il)