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

NLR-1T AIRFOIL (nlr1t-il) Xfoil prediction polar at RE=200,000 Ncrit=9


Details Polar file
Airfoil: NLR-1T AIRFOIL (nlr1t-il)
Reynolds number: 200,000
Max Cl/Cd: 53.41 at α=2.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-nlr1t-il-200000.txt
Download as CSV file: xf-nlr1t-il-200000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NLR-1T AIRFOIL                                  
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.200 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -8.750  -0.5471   0.08498   0.08152  -0.0344   1.0000   0.0417
  -8.500  -0.5563   0.08130   0.07775  -0.0359   1.0000   0.0419
  -8.250  -0.5638   0.07816   0.07450  -0.0357   1.0000   0.0420
  -8.000  -0.5744   0.07572   0.07196  -0.0332   1.0000   0.0421
  -7.750  -0.5893   0.06931   0.06561  -0.0308   1.0000   0.0428
  -7.500  -0.5902   0.06619   0.06255  -0.0282   1.0000   0.0436
  -7.250  -0.5925   0.06392   0.06030  -0.0252   1.0000   0.0444
  -7.000  -0.5947   0.06164   0.05800  -0.0223   1.0000   0.0454
  -6.750  -0.5959   0.05921   0.05549  -0.0197   1.0000   0.0466
  -6.500  -0.5955   0.05657   0.05274  -0.0172   1.0000   0.0483
  -6.250  -0.5927   0.05383   0.04980  -0.0147   1.0000   0.0513
  -6.000  -0.5961   0.05184   0.04706  -0.0104   1.0000   0.0550
  -5.750  -0.5866   0.04721   0.04266  -0.0097   1.0000   0.0570
  -5.500  -0.5760   0.04500   0.04045  -0.0079   1.0000   0.0595
  -5.250  -0.5644   0.04284   0.03808  -0.0058   1.0000   0.0637
  -5.000  -0.5553   0.04006   0.03480  -0.0031   1.0000   0.0697
  -4.750  -0.5285   0.02565   0.02086  -0.0017   1.0000   0.0721
  -4.500  -0.5142   0.02375   0.01879   0.0001   1.0000   0.0775
  -4.250  -0.5092   0.03336   0.02767   0.0016   1.0000   0.0858
  -4.000  -0.4923   0.03197   0.02594   0.0034   1.0000   0.0980
  -3.750  -0.4726   0.02986   0.02388   0.0043   1.0000   0.1040
  -3.500  -0.4353   0.02503   0.01785   0.0078   1.0000   0.0520
  -3.250  -0.4075   0.02216   0.01450   0.0089   1.0000   0.0453
  -3.000  -0.3814   0.02122   0.01322   0.0102   1.0000   0.0430
  -2.750  -0.3560   0.02029   0.01217   0.0110   1.0000   0.0417
  -2.500  -0.3296   0.01880   0.01057   0.0116   1.0000   0.0411
  -2.250  -0.3041   0.01781   0.00950   0.0122   1.0000   0.0412
  -2.000  -0.2793   0.01706   0.00870   0.0130   1.0000   0.0418
  -1.750  -0.2548   0.01609   0.00775   0.0136   1.0000   0.0433
  -1.500  -0.2309   0.01549   0.00721   0.0142   1.0000   0.0482
  -1.250  -0.1960   0.01515   0.00683   0.0127   0.9974   0.0539
  -1.000  -0.0848   0.01205   0.00711  -0.0036   1.0000   1.0000
  -0.750  -0.0631   0.01215   0.00706  -0.0027   0.9992   1.0000
  -0.500   0.0184   0.01230   0.00702  -0.0141   0.9872   1.0000
  -0.250   0.1403   0.01194   0.00651  -0.0332   0.9702   1.0000
   0.000   0.1850   0.01181   0.00633  -0.0368   0.9620   1.0000
   0.250   0.2468   0.01153   0.00602  -0.0438   0.9569   1.0000
   0.500   0.2853   0.01138   0.00587  -0.0459   0.9486   1.0000
   0.750   0.3330   0.01107   0.00556  -0.0496   0.9412   1.0000
   1.000   0.3618   0.01092   0.00541  -0.0494   0.9295   1.0000
   1.250   0.3890   0.01074   0.00525  -0.0486   0.9170   1.0000
   1.500   0.4131   0.01055   0.00507  -0.0471   0.9027   1.0000
   1.750   0.4349   0.01033   0.00486  -0.0450   0.8861   1.0000
   2.000   0.4551   0.01011   0.00464  -0.0425   0.8661   1.0000
   2.250   0.4754   0.00986   0.00439  -0.0401   0.8409   1.0000
   2.500   0.4955   0.00965   0.00419  -0.0378   0.8002   1.0000
   2.750   0.5122   0.00959   0.00341  -0.0337   0.6242   1.0000
   3.000   0.5263   0.01089   0.00365  -0.0310   0.4290   1.0000
   3.250   0.5444   0.01186   0.00401  -0.0295   0.3263   1.0000
   3.500   0.5656   0.01248   0.00434  -0.0285   0.2864   1.0000
   3.750   0.5882   0.01295   0.00467  -0.0276   0.2654   1.0000
   4.000   0.6110   0.01338   0.00503  -0.0267   0.2511   1.0000
   4.250   0.6339   0.01378   0.00536  -0.0259   0.2387   1.0000
   4.500   0.6568   0.01417   0.00569  -0.0250   0.2269   1.0000
   4.750   0.6794   0.01460   0.00602  -0.0242   0.2159   1.0000
   5.000   0.7029   0.01486   0.00637  -0.0234   0.2064   1.0000
   5.250   0.7258   0.01526   0.00675  -0.0226   0.1976   1.0000
   5.500   0.7484   0.01562   0.00709  -0.0217   0.1885   1.0000
   5.750   0.7714   0.01593   0.00749  -0.0209   0.1797   1.0000
   6.000   0.7937   0.01635   0.00791  -0.0200   0.1706   1.0000
   6.250   0.8158   0.01665   0.00825  -0.0190   0.1595   1.0000
   6.500   0.8375   0.01692   0.00857  -0.0180   0.1454   1.0000
   6.750   0.8591   0.01718   0.00891  -0.0169   0.1255   1.0000
   7.000   0.8801   0.01759   0.00928  -0.0157   0.1042   1.0000
   7.250   0.9000   0.01825   0.00987  -0.0143   0.0910   1.0000
   7.500   0.9203   0.01896   0.01064  -0.0130   0.0834   1.0000
   7.750   0.9385   0.02001   0.01162  -0.0114   0.0780   1.0000
   8.000   0.9591   0.02077   0.01253  -0.0101   0.0741   1.0000
   8.250   0.9789   0.02158   0.01342  -0.0087   0.0703   1.0000
   8.500   0.9980   0.02282   0.01465  -0.0074   0.0673   1.0000
   8.750   1.0184   0.02425   0.01621  -0.0063   0.0653   1.0000
   9.000   1.0386   0.02541   0.01757  -0.0050   0.0637   1.0000
   9.250   1.0581   0.02668   0.01905  -0.0036   0.0618   1.0000
   9.500   1.0760   0.02768   0.02021  -0.0021   0.0590   1.0000
   9.750   1.0929   0.02916   0.02168  -0.0010   0.0554   1.0000
  10.000   1.1063   0.03030   0.02312   0.0011   0.0524   1.0000
  10.250   1.1204   0.03083   0.02381   0.0030   0.0486   1.0000
  10.500   1.1347   0.03175   0.02472   0.0046   0.0453   1.0000
  10.750   1.1425   0.03405   0.02730   0.0070   0.0426   1.0000
  11.000   1.1515   0.03503   0.02858   0.0096   0.0395   1.0000
  11.250   1.1605   0.03589   0.02953   0.0119   0.0360   1.0000
  11.500   1.1578   0.03902   0.03287   0.0151   0.0330   1.0000
  11.750   1.1555   0.04037   0.03455   0.0190   0.0302   1.0000
  12.000   1.1549   0.04139   0.03567   0.0223   0.0279   1.0000
  12.250   1.1564   0.04278   0.03700   0.0247   0.0261   1.0000
  12.500   1.1399   0.04644   0.04093   0.0280   0.0251   1.0000
  12.750   1.1286   0.04895   0.04376   0.0304   0.0243   1.0000
  13.000   1.1141   0.05237   0.04745   0.0319   0.0237   1.0000
  13.250   1.0967   0.05656   0.05190   0.0323   0.0233   1.0000
  13.500   1.0766   0.06170   0.05729   0.0312   0.0232   1.0000
  13.750   1.0540   0.06819   0.06402   0.0280   0.0233   1.0000
  14.000   1.0285   0.07645   0.07248   0.0225   0.0237   1.0000
  14.250   1.0006   0.08636   0.08255   0.0157   0.0244   1.0000
  14.500   0.9700   0.09708   0.09337   0.0092   0.0252   1.0000
<< Back to NLR-1T AIRFOIL (nlr1t-il)

Polar data table (+)

Polar graphs


<< Back to NLR-1T AIRFOIL (nlr1t-il)