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

N-22 (n22-il) Xfoil prediction polar at RE=100,000 Ncrit=9


Details Polar file
Airfoil: N-22 (n22-il)
Reynolds number: 100,000
Max Cl/Cd: 51.81 at α=5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n22-il-100000.txt
Download as CSV file: xf-n22-il-100000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: N-22                                            
  
 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
  ------ -------- --------- --------- -------- -------- --------
 -10.250  -0.3285   0.09633   0.09143  -0.0339   0.9999   0.1364
 -10.000  -0.3554   0.09537   0.09062  -0.0331   0.9999   0.1397
  -9.750  -0.3956   0.09529   0.09073  -0.0302   0.9999   0.1406
  -9.500  -0.4384   0.09431   0.08989  -0.0317   0.9999   0.1413
  -9.250  -0.4412   0.09067   0.08632  -0.0287   0.9999   0.1427
  -9.000  -0.4226   0.08847   0.08414  -0.0224   0.9999   0.1452
  -8.750  -0.4284   0.08682   0.08254  -0.0196   0.9999   0.1475
  -8.500  -0.4393   0.08501   0.08079  -0.0181   0.9999   0.1503
  -8.000  -0.4607   0.07763   0.07340  -0.0261   0.9981   0.1596
  -7.750  -0.4283   0.07266   0.06806  -0.0398   0.9891   0.1743
  -7.500  -0.4031   0.06881   0.06440  -0.0376   0.9851   0.1766
  -7.250  -0.3739   0.06627   0.06185  -0.0388   0.9792   0.1843
  -7.000  -0.3438   0.06180   0.05725  -0.0453   0.9718   0.1960
  -6.750  -0.2958   0.04018   0.03342  -0.0660   0.9658   0.1097
  -6.500  -0.2595   0.03685   0.02960  -0.0690   0.9592   0.1097
  -6.250  -0.2155   0.03375   0.02598  -0.0730   0.9551   0.1094
  -6.000  -0.1842   0.03169   0.02345  -0.0741   0.9465   0.1101
  -5.750  -0.1409   0.02970   0.02118  -0.0775   0.9413   0.1139
  -5.500  -0.1073   0.02870   0.02009  -0.0788   0.9330   0.1183
  -5.250  -0.0643   0.02740   0.01845  -0.0815   0.9267   0.1234
  -5.000  -0.0259   0.02628   0.01723  -0.0834   0.9195   0.1304
  -4.750   0.0136   0.02553   0.01630  -0.0854   0.9115   0.1409
  -4.500   0.0646   0.02439   0.01517  -0.0894   0.9075   0.1570
  -4.250   0.0929   0.02373   0.01458  -0.0894   0.8963   0.1745
  -4.000   0.1431   0.02252   0.01362  -0.0932   0.8920   0.2111
  -3.750   0.1712   0.02189   0.01322  -0.0931   0.8808   0.2506
  -3.500   0.2154   0.02076   0.01247  -0.0955   0.8757   0.3223
  -3.250   0.2404   0.01990   0.01227  -0.0948   0.8643   0.4414
  -3.000   0.3018   0.01788   0.01181  -0.0981   0.8611   1.0001
  -2.750   0.3294   0.01788   0.01154  -0.0978   0.8494   1.0001
  -2.500   0.3691   0.01750   0.01092  -0.0991   0.8429   1.0001
  -2.250   0.3934   0.01754   0.01079  -0.0982   0.8300   1.0001
  -2.000   0.4208   0.01751   0.01061  -0.0976   0.8187   1.0001
  -1.750   0.4554   0.01719   0.01011  -0.0980   0.8107   1.0001
  -1.500   0.4799   0.01724   0.01004  -0.0970   0.7975   1.0001
  -1.250   0.5065   0.01723   0.00991  -0.0962   0.7852   1.0001
  -1.000   0.5361   0.01709   0.00964  -0.0959   0.7746   1.0001
  -0.750   0.5645   0.01700   0.00943  -0.0953   0.7628   1.0001
  -0.500   0.5900   0.01706   0.00939  -0.0945   0.7489   1.0001
  -0.250   0.6163   0.01710   0.00934  -0.0937   0.7352   1.0001
   0.000   0.6432   0.01712   0.00926  -0.0930   0.7218   1.0001
   0.250   0.6709   0.01714   0.00916  -0.0925   0.7087   1.0001
   0.500   0.6993   0.01713   0.00903  -0.0920   0.6957   1.0001
   0.750   0.7250   0.01726   0.00907  -0.0913   0.6804   1.0001
   1.000   0.7506   0.01741   0.00916  -0.0906   0.6651   1.0001
   1.250   0.7764   0.01759   0.00927  -0.0900   0.6500   1.0001
   1.500   0.8022   0.01778   0.00939  -0.0894   0.6350   1.0001
   1.750   0.8281   0.01800   0.00953  -0.0888   0.6201   1.0001
   2.000   0.8540   0.01825   0.00972  -0.0883   0.6059   1.0001
   2.250   0.8802   0.01851   0.00991  -0.0879   0.5923   1.0001
   2.500   0.9070   0.01876   0.01007  -0.0876   0.5795   1.0001
   2.750   0.9333   0.01905   0.01030  -0.0872   0.5667   1.0001
   3.000   0.9578   0.01943   0.01068  -0.0867   0.5531   1.0001
   3.250   0.9826   0.01978   0.01099  -0.0861   0.5396   1.0001
   3.500   1.0076   0.02009   0.01127  -0.0856   0.5263   1.0001
   3.750   1.0330   0.02039   0.01150  -0.0851   0.5132   1.0001
   4.000   1.0596   0.02068   0.01168  -0.0848   0.5011   1.0001
   4.250   1.0826   0.02108   0.01214  -0.0841   0.4878   1.0001
   4.500   1.1059   0.02152   0.01261  -0.0835   0.4755   1.0001
   4.750   1.1305   0.02195   0.01302  -0.0830   0.4640   1.0001
   5.000   1.1570   0.02233   0.01331  -0.0828   0.4530   1.0001
   5.250   1.1783   0.02282   0.01391  -0.0819   0.4406   1.0001
   5.500   1.2008   0.02333   0.01447  -0.0812   0.4288   1.0001
   5.750   1.2249   0.02380   0.01493  -0.0807   0.4173   1.0001
   6.000   1.2492   0.02425   0.01533  -0.0803   0.4054   1.0001
   6.250   1.2690   0.02479   0.01599  -0.0792   0.3923   1.0001
   6.500   1.2893   0.02533   0.01659  -0.0782   0.3790   1.0001
   6.750   1.3097   0.02588   0.01719  -0.0772   0.3654   1.0001
   7.000   1.3298   0.02645   0.01777  -0.0761   0.3514   1.0001
   7.250   1.3494   0.02708   0.01839  -0.0750   0.3372   1.0001
   7.500   1.3682   0.02774   0.01906  -0.0738   0.3226   1.0001
   7.750   1.3862   0.02845   0.01975  -0.0726   0.3079   1.0001
   8.000   1.4037   0.02920   0.02047  -0.0712   0.2937   1.0001
   8.250   1.4173   0.03005   0.02143  -0.0694   0.2801   1.0001
   8.500   1.4314   0.03104   0.02254  -0.0677   0.2681   1.0001
   8.750   1.4480   0.03203   0.02357  -0.0664   0.2577   1.0001
   9.000   1.4676   0.03280   0.02422  -0.0655   0.2476   1.0001
   9.250   1.4738   0.03375   0.02540  -0.0628   0.2375   1.0001
   9.500   1.4831   0.03458   0.02627  -0.0606   0.2279   1.0001
   9.750   1.4937   0.03520   0.02681  -0.0586   0.2182   1.0001
  10.000   1.4908   0.03627   0.02810  -0.0548   0.2099   1.0001
  10.250   1.4968   0.03717   0.02896  -0.0524   0.2016   1.0001
  10.500   1.4959   0.03841   0.03036  -0.0495   0.1937   1.0001
  10.750   1.4998   0.03967   0.03164  -0.0474   0.1862   1.0001
  11.000   1.5004   0.04114   0.03325  -0.0452   0.1790   1.0001
  11.250   1.5042   0.04266   0.03482  -0.0435   0.1722   1.0001
  11.500   1.5021   0.04449   0.03683  -0.0415   0.1654   1.0001
  11.750   1.5026   0.04632   0.03869  -0.0400   0.1583   1.0001
  12.000   1.4961   0.04867   0.04122  -0.0384   0.1512   1.0001
  12.250   1.4911   0.05109   0.04372  -0.0371   0.1438   1.0001
  12.500   1.4812   0.05404   0.04680  -0.0361   0.1360   1.0001
  12.750   1.4687   0.05756   0.05048  -0.0355   0.1278   1.0001
  13.000   1.4564   0.06115   0.05404  -0.0351   0.1187   1.0001
  13.250   1.4374   0.06594   0.05903  -0.0353   0.1098   1.0001
  13.500   1.4189   0.07105   0.06423  -0.0359   0.1005   1.0001
  13.750   1.4037   0.07596   0.06911  -0.0366   0.0923   1.0001
  14.000   1.3890   0.08117   0.07440  -0.0378   0.0853   1.0001
  14.250   1.3769   0.08616   0.07946  -0.0387   0.0794   1.0001
  14.500   1.3679   0.09081   0.08414  -0.0399   0.0745   1.0001
  14.750   1.3617   0.09505   0.08843  -0.0407   0.0705   1.0001
<< Back to N-22 (n22-il)

Polar data table (+)

Polar graphs


<< Back to N-22 (n22-il)