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N-12 (n12-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: N-12 (n12-il)
Reynolds number: 50,000
Max Cl/Cd: 28.91 at α=4.75°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-n12-il-50000.txt
Download as CSV file: xf-n12-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: N-12                                            
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.050 e 6     Ncrit =   9.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -9.500  -0.6140   0.09376   0.08635  -0.0132   1.0000   0.1884
  -9.250  -0.6665   0.08121   0.07395  -0.0240   1.0000   0.1583
  -9.000  -0.7332   0.07390   0.06657  -0.0252   1.0000   0.1482
  -8.750  -0.7359   0.06900   0.06162  -0.0245   1.0000   0.1432
  -8.500  -0.7871   0.06315   0.05516  -0.0213   1.0000   0.1347
  -8.250  -0.7789   0.05876   0.05068  -0.0200   1.0000   0.1326
  -8.000  -0.7783   0.05470   0.04636  -0.0178   1.0000   0.1302
  -7.750  -0.7782   0.05079   0.04208  -0.0152   1.0000   0.1281
  -7.500  -0.7745   0.04724   0.03812  -0.0125   1.0000   0.1271
  -7.250  -0.7665   0.04426   0.03475  -0.0099   1.0000   0.1284
  -7.000  -0.7558   0.04149   0.03156  -0.0075   1.0000   0.1302
  -6.750  -0.7420   0.03880   0.02844  -0.0051   1.0000   0.1317
  -6.500  -0.7248   0.03630   0.02549  -0.0031   1.0000   0.1330
  -6.250  -0.7053   0.03411   0.02279  -0.0012   1.0000   0.1351
  -6.000  -0.6810   0.03187   0.02064  -0.0006   1.0000   0.1410
  -5.750  -0.6581   0.03020   0.01869   0.0008   1.0000   0.1490
  -5.500  -0.6306   0.02828   0.01687   0.0013   1.0000   0.1584
  -5.250  -0.6022   0.02650   0.01514   0.0018   1.0000   0.1724
  -5.000  -0.5768   0.02463   0.01347   0.0028   1.0000   0.2008
  -4.750  -0.5653   0.02215   0.01204   0.0057   1.0000   0.2941
  -4.500  -0.5636   0.02054   0.01140   0.0110   1.0000   0.4502
  -4.250  -0.5552   0.01979   0.01133   0.0161   1.0000   0.5634
  -4.000  -0.5408   0.01965   0.01147   0.0206   1.0000   0.6499
  -3.750  -0.5190   0.01979   0.01169   0.0240   1.0000   0.7177
  -3.500  -0.4861   0.02057   0.01246   0.0265   1.0000   0.7847
  -3.250  -0.4282   0.02209   0.01368   0.0253   1.0000   0.8462
  -3.000  -0.2859   0.02453   0.01533   0.0097   1.0000   0.9084
  -2.750  -0.1488   0.02459   0.01476  -0.0097   1.0000   0.9645
  -2.500  -0.0589   0.02311   0.01294  -0.0240   1.0000   0.9979
  -2.250  -0.0466   0.02226   0.01202  -0.0237   1.0000   1.0000
  -2.000  -0.0394   0.02161   0.01132  -0.0221   1.0000   1.0000
  -1.750  -0.0316   0.02108   0.01076  -0.0202   1.0000   1.0000
  -1.500  -0.0236   0.02066   0.01032  -0.0180   1.0000   1.0000
  -1.250  -0.0159   0.02032   0.00997  -0.0157   1.0000   1.0000
  -1.000  -0.0089   0.02006   0.00970  -0.0132   1.0000   1.0000
  -0.750  -0.0029   0.01986   0.00951  -0.0103   1.0000   1.0000
  -0.500   0.0016   0.01974   0.00940  -0.0072   1.0000   1.0000
  -0.250   0.0045   0.01968   0.00935  -0.0038   1.0000   1.0000
   0.000   0.0061   0.01969   0.00936  -0.0002   1.0000   1.0000
   0.250   0.0071   0.01976   0.00943   0.0035   1.0000   1.0000
   0.500   0.0084   0.01989   0.00956   0.0070   1.0000   1.0000
   0.750   0.0107   0.02009   0.00974   0.0104   1.0000   1.0000
   1.000   0.0141   0.02035   0.01000   0.0134   1.0000   1.0000
   1.250   0.0186   0.02069   0.01033   0.0161   1.0000   1.0000
   1.500   0.0240   0.02111   0.01074   0.0185   1.0000   1.0000
   1.750   0.0300   0.02162   0.01125   0.0207   1.0000   1.0000
   2.000   0.0365   0.02220   0.01186   0.0225   1.0000   1.0000
   2.250   0.1125   0.02330   0.01318   0.0112   0.9751   1.0000
   2.500   0.1999   0.02395   0.01414  -0.0011   0.9425   1.0000
   2.750   0.2896   0.02404   0.01461  -0.0128   0.9065   1.0000
   3.000   0.4051   0.02316   0.01431  -0.0272   0.8641   1.0000
   3.250   0.4940   0.02181   0.01337  -0.0349   0.8099   1.0000
   3.500   0.5334   0.02140   0.01307  -0.0345   0.7562   1.0000
   3.750   0.5622   0.02127   0.01297  -0.0325   0.7066   1.0000
   4.000   0.5854   0.02128   0.01291  -0.0296   0.6587   1.0000
   4.250   0.6027   0.02125   0.01269  -0.0256   0.6063   1.0000
   4.500   0.6132   0.02128   0.01245  -0.0206   0.5471   1.0000
   4.750   0.6227   0.02154   0.01251  -0.0161   0.4886   1.0000
   5.000   0.6304   0.02200   0.01272  -0.0114   0.4202   1.0000
   5.250   0.6315   0.02344   0.01338  -0.0058   0.3058   1.0000
   5.500   0.6422   0.02555   0.01476  -0.0022   0.2203   1.0000
   5.750   0.6608   0.02692   0.01595  -0.0002   0.1872   1.0000
   6.000   0.6839   0.02838   0.01733   0.0012   0.1700   1.0000
   6.250   0.7087   0.03007   0.01907   0.0023   0.1589   1.0000
   6.500   0.7314   0.03181   0.02082   0.0035   0.1492   1.0000
   6.750   0.7537   0.03376   0.02290   0.0047   0.1421   1.0000
   7.000   0.7757   0.03632   0.02560   0.0059   0.1384   1.0000
   7.250   0.7905   0.03894   0.02874   0.0082   0.1358   1.0000
   7.500   0.8023   0.04160   0.03185   0.0107   0.1324   1.0000
   7.750   0.8130   0.04445   0.03503   0.0130   0.1299   1.0000
   8.000   0.8173   0.04801   0.03906   0.0159   0.1303   1.0000
   8.250   0.8156   0.05204   0.04357   0.0189   0.1319   1.0000
   8.500   0.8096   0.05627   0.04819   0.0218   0.1338   1.0000
   8.750   0.8012   0.06064   0.05285   0.0242   0.1356   1.0000
   9.000   0.7919   0.06503   0.05743   0.0263   0.1370   1.0000
   9.250   0.7874   0.06939   0.06189   0.0277   0.1385   1.0000
   9.500   0.7150   0.07712   0.06984   0.0285   0.1493   1.0000
   9.750   0.7156   0.08191   0.07463   0.0287   0.1522   1.0000
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