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

LISSAMAN 7769 AIRFOIL (l7769-il) Xfoil prediction polar at RE=500,000 Ncrit=5


Details Polar file
Airfoil: LISSAMAN 7769 AIRFOIL (l7769-il)
Reynolds number: 500,000
Max Cl/Cd: 111.05 at α=8°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-l7769-il-500000-n5.txt
Download as CSV file: xf-l7769-il-500000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: LISSAMAN 7769 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 =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.4546   0.11028   0.10623   0.0145   0.4942   0.0211
 -10.000  -0.4495   0.10685   0.10280   0.0128   0.4927   0.0214
  -8.500  -0.6320   0.03222   0.02727  -0.0348   0.4943   0.0249
  -8.250  -0.6150   0.02881   0.02350  -0.0354   0.4924   0.0252
  -8.000  -0.5949   0.02641   0.02078  -0.0354   0.4902   0.0254
  -7.750  -0.5727   0.02459   0.01868  -0.0353   0.4880   0.0257
  -7.500  -0.5494   0.02309   0.01694  -0.0351   0.4857   0.0259
  -7.250  -0.5251   0.02181   0.01543  -0.0348   0.4836   0.0262
  -7.000  -0.5002   0.02071   0.01410  -0.0346   0.4814   0.0264
  -6.750  -0.4748   0.01973   0.01292  -0.0343   0.4792   0.0267
  -6.500  -0.4488   0.01884   0.01187  -0.0340   0.4770   0.0269
  -6.250  -0.4225   0.01805   0.01092  -0.0337   0.4748   0.0272
  -6.000  -0.3958   0.01736   0.01008  -0.0334   0.4726   0.0275
  -5.750  -0.3688   0.01679   0.00938  -0.0331   0.4705   0.0278
  -5.500  -0.3416   0.01630   0.00876  -0.0328   0.4684   0.0282
  -5.250  -0.3143   0.01581   0.00814  -0.0325   0.4664   0.0283
  -5.000  -0.2872   0.01527   0.00748  -0.0322   0.4646   0.0286
  -4.750  -0.2602   0.01461   0.00674  -0.0319   0.4628   0.0289
  -4.500  -0.2328   0.01412   0.00621  -0.0316   0.4610   0.0293
  -4.250  -0.2052   0.01371   0.00576  -0.0314   0.4590   0.0297
  -4.000  -0.1774   0.01336   0.00538  -0.0311   0.4568   0.0300
  -3.750  -0.1496   0.01305   0.00502  -0.0309   0.4546   0.0304
  -3.500  -0.1217   0.01276   0.00469  -0.0306   0.4527   0.0309
  -3.250  -0.0938   0.01250   0.00439  -0.0304   0.4509   0.0314
  -3.000  -0.0658   0.01226   0.00410  -0.0301   0.4492   0.0319
  -2.750  -0.0378   0.01206   0.00385  -0.0299   0.4474   0.0324
  -2.500  -0.0096   0.01187   0.00363  -0.0297   0.4459   0.0330
  -2.250   0.0187   0.01170   0.00344  -0.0295   0.4443   0.0337
  -2.000   0.0468   0.01146   0.00321  -0.0293   0.4426   0.0349
  -1.750   0.0750   0.01129   0.00303  -0.0291   0.4408   0.0360
  -1.500   0.1033   0.01114   0.00288  -0.0289   0.4390   0.0372
  -1.250   0.1316   0.01102   0.00273  -0.0287   0.4371   0.0387
  -1.000   0.1599   0.01090   0.00260  -0.0286   0.4353   0.0411
  -0.750   0.1881   0.01077   0.00249  -0.0284   0.4336   0.0463
  -0.500   0.2162   0.01064   0.00242  -0.0282   0.4320   0.0626
  -0.250   0.2445   0.01057   0.00240  -0.0281   0.4306   0.0773
   0.000   0.2729   0.01051   0.00239  -0.0280   0.4292   0.0879
   0.250   0.3014   0.01047   0.00240  -0.0279   0.4276   0.0976
   0.500   0.3299   0.01044   0.00241  -0.0278   0.4260   0.1062
   0.750   0.3584   0.01043   0.00242  -0.0276   0.4245   0.1141
   1.000   0.3869   0.01042   0.00244  -0.0275   0.4229   0.1220
   1.250   0.4154   0.01042   0.00245  -0.0274   0.4212   0.1287
   1.500   0.4438   0.01041   0.00246  -0.0273   0.4195   0.1371
   1.750   0.4722   0.01041   0.00248  -0.0273   0.4179   0.1456
   2.000   0.5005   0.01043   0.00250  -0.0272   0.4164   0.1536
   2.250   0.5289   0.01045   0.00255  -0.0271   0.4151   0.1609
   2.500   0.5573   0.01045   0.00260  -0.0270   0.4138   0.1694
   2.750   0.5857   0.01045   0.00266  -0.0270   0.4124   0.1804
   3.000   0.6140   0.01042   0.00273  -0.0269   0.4106   0.2015
   3.500   0.6781   0.00861   0.00290  -0.0287   0.4058   1.0000
   3.750   0.7059   0.00869   0.00294  -0.0285   0.4033   1.0000
   4.000   0.7336   0.00879   0.00299  -0.0283   0.4009   1.0000
   4.250   0.7615   0.00888   0.00308  -0.0282   0.3986   1.0000
   4.500   0.7894   0.00895   0.00318  -0.0280   0.3955   1.0000
   4.750   0.8174   0.00902   0.00326  -0.0279   0.3923   1.0000
   5.000   0.8452   0.00912   0.00335  -0.0278   0.3899   1.0000
   5.250   0.8729   0.00922   0.00344  -0.0276   0.3872   1.0000
   5.500   0.9007   0.00931   0.00353  -0.0275   0.3832   1.0000
   5.750   0.9287   0.00937   0.00364  -0.0274   0.3779   1.0000
   6.000   0.9564   0.00947   0.00374  -0.0273   0.3736   1.0000
   6.250   0.9839   0.00960   0.00386  -0.0272   0.3705   1.0000
   6.500   1.0118   0.00971   0.00403  -0.0271   0.3666   1.0000
   6.750   1.0394   0.00982   0.00417  -0.0270   0.3617   1.0000
   7.000   1.0669   0.00995   0.00431  -0.0269   0.3571   1.0000
   7.250   1.0944   0.01009   0.00450  -0.0268   0.3525   1.0000
   7.500   1.1219   0.01023   0.00468  -0.0268   0.3465   1.0000
   7.750   1.1490   0.01040   0.00486  -0.0267   0.3398   1.0000
   8.000   1.1760   0.01059   0.00506  -0.0266   0.3279   1.0000
   8.250   1.2022   0.01088   0.00532  -0.0265   0.3075   1.0000
   8.500   1.2267   0.01143   0.00574  -0.0263   0.2771   1.0000
   8.750   1.2496   0.01223   0.00639  -0.0261   0.2461   1.0000
   9.000   1.2719   0.01309   0.00714  -0.0259   0.2204   1.0000
   9.250   1.2948   0.01383   0.00783  -0.0257   0.2022   1.0000
   9.500   1.3174   0.01457   0.00852  -0.0255   0.1860   1.0000
   9.750   1.3384   0.01548   0.00934  -0.0253   0.1665   1.0000
  10.000   1.3581   0.01649   0.01027  -0.0250   0.1443   1.0000
  10.250   1.3762   0.01762   0.01131  -0.0247   0.1258   1.0000
  10.500   1.3947   0.01863   0.01229  -0.0244   0.1134   1.0000
  10.750   1.4122   0.01968   0.01334  -0.0241   0.1027   1.0000
  11.000   1.4270   0.02095   0.01460  -0.0238   0.0918   1.0000
  11.250   1.4370   0.02265   0.01630  -0.0236   0.0796   1.0000
  11.500   1.4365   0.02625   0.01993  -0.0256   0.0671   1.0000
  11.750   1.4197   0.03030   0.02403  -0.0258   0.0622   1.0000
  12.000   1.4049   0.03409   0.02782  -0.0258   0.0554   1.0000
  12.250   1.3914   0.03777   0.03150  -0.0257   0.0479   1.0000
  12.500   1.3807   0.04124   0.03497  -0.0257   0.0412   1.0000
  12.750   1.3697   0.04480   0.03853  -0.0258   0.0346   1.0000
  13.000   1.3618   0.04812   0.04185  -0.0260   0.0295   1.0000
  13.250   1.3563   0.05126   0.04501  -0.0262   0.0258   1.0000
  13.500   1.3529   0.05434   0.04812  -0.0265   0.0234   1.0000
  13.750   1.3525   0.05722   0.05105  -0.0269   0.0213   1.0000
  14.000   1.3505   0.06038   0.05424  -0.0274   0.0195   1.0000
  14.250   1.3519   0.06321   0.05713  -0.0279   0.0185   1.0000
  14.500   1.3528   0.06616   0.06013  -0.0284   0.0176   1.0000
  14.750   1.3531   0.06924   0.06327  -0.0291   0.0169   1.0000
  15.000   1.3524   0.07252   0.06660  -0.0298   0.0161   1.0000
  15.250   1.3519   0.07582   0.06997  -0.0306   0.0155   1.0000
  15.500   1.3533   0.07892   0.07315  -0.0314   0.0151   1.0000
  15.750   1.3532   0.08231   0.07662  -0.0324   0.0147   1.0000
  16.000   1.3532   0.08568   0.08006  -0.0333   0.0142   1.0000
  16.250   1.3518   0.08937   0.08382  -0.0345   0.0139   1.0000
  16.500   1.3500   0.09315   0.08767  -0.0357   0.0135   1.0000
  16.750   1.3471   0.09721   0.09180  -0.0371   0.0132   1.0000
  17.000   1.3429   0.10149   0.09616  -0.0387   0.0128   1.0000
<< Back to LISSAMAN 7769 AIRFOIL (l7769-il)

Polar data table (+)

Polar graphs


<< Back to LISSAMAN 7769 AIRFOIL (l7769-il)