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NACA M16 AIRFOIL (m16-il) Xfoil prediction polar at RE=100,000 Ncrit=5


Details Polar file
Airfoil: NACA M16 AIRFOIL (m16-il)
Reynolds number: 100,000
Max Cl/Cd: 53.65 at α=5.75°
Description: Mach=0 Ncrit=5
Source: Xfoil prediction
Download polar: xf-m16-il-100000-n5.txt
Download as CSV file: xf-m16-il-100000-n5.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: NACA M16 AIRFOIL                                
  
 1 1 Reynolds number fixed          Mach number fixed         
  
 xtrf =   1.000 (top)        1.000 (bottom)  
 Mach =   0.000     Re =     0.100 e 6     Ncrit =   5.000
  
   alpha    CL        CD       CDp       CM     Top_Xtr  Bot_Xtr
  ------ -------- --------- --------- -------- -------- --------
  -7.250  -0.4502   0.09546   0.09133  -0.0002   1.0000   0.0290
  -7.000  -0.4407   0.09096   0.08686  -0.0004   1.0000   0.0297
  -6.750  -0.4294   0.08696   0.08284  -0.0023   1.0000   0.0303
  -6.500  -0.4162   0.08308   0.07896  -0.0051   1.0000   0.0311
  -6.250  -0.4010   0.07921   0.07509  -0.0083   1.0000   0.0320
  -6.000  -0.3836   0.07532   0.07119  -0.0119   1.0000   0.0332
  -5.750  -0.3631   0.07155   0.06739  -0.0159   1.0000   0.0360
  -5.500  -0.3139   0.06912   0.06460  -0.0274   0.9566   0.0401
  -5.250  -0.2974   0.06293   0.05837  -0.0295   0.9235   0.0411
  -5.000  -0.2848   0.05911   0.05450  -0.0290   0.8972   0.0430
  -4.750  -0.2660   0.05596   0.05121  -0.0297   0.8752   0.0454
  -4.250  -0.2165   0.04956   0.04425  -0.0331   0.8417   0.0563
  -4.000  -0.1967   0.04684   0.04137  -0.0330   0.8274   0.0609
  -3.500  -0.1435   0.04092   0.03483  -0.0349   0.8023   0.0716
  -3.250  -0.1085   0.03647   0.02974  -0.0354   0.7921   0.0523
  -3.000  -0.0828   0.03352   0.02646  -0.0354   0.7820   0.0511
  -2.750  -0.0535   0.03092   0.02330  -0.0355   0.7719   0.0551
  -2.250   0.0026   0.02694   0.01844  -0.0351   0.7528   0.0531
  -2.000   0.0301   0.02508   0.01622  -0.0349   0.7437   0.0528
  -1.750   0.0583   0.02354   0.01429  -0.0347   0.7344   0.0530
  -1.500   0.0860   0.02232   0.01265  -0.0343   0.7264   0.0534
  -1.250   0.1140   0.02103   0.01110  -0.0343   0.7172   0.0548
  -1.000   0.1406   0.02035   0.01036  -0.0343   0.7088   0.0595
  -0.750   0.1682   0.01963   0.00943  -0.0340   0.7006   0.0614
  -0.500   0.1962   0.01884   0.00844  -0.0338   0.6922   0.0608
  -0.250   0.2235   0.01819   0.00761  -0.0334   0.6849   0.0604
   0.000   0.2513   0.01763   0.00697  -0.0333   0.6762   0.0602
   0.250   0.2781   0.01717   0.00640  -0.0329   0.6691   0.0603
   0.500   0.3053   0.01677   0.00597  -0.0327   0.6606   0.0607
   0.750   0.3323   0.01649   0.00559  -0.0323   0.6532   0.0614
   1.000   0.3597   0.01629   0.00531  -0.0322   0.6451   0.0627
   1.250   0.3871   0.01617   0.00509  -0.0320   0.6375   0.0648
   1.750   0.4419   0.01599   0.00485  -0.0317   0.6219   0.0783
   2.000   0.4621   0.01440   0.00486  -0.0306   0.6149   0.6808
   2.250   0.5072   0.01407   0.00490  -0.0337   0.6058   1.0000
   2.500   0.5335   0.01423   0.00493  -0.0333   0.5988   1.0000
   2.750   0.5605   0.01440   0.00509  -0.0333   0.5900   1.0000
   3.000   0.5869   0.01457   0.00522  -0.0329   0.5830   1.0000
   3.250   0.6138   0.01476   0.00544  -0.0329   0.5741   1.0000
   3.500   0.6405   0.01495   0.00564  -0.0327   0.5665   1.0000
   3.750   0.6672   0.01514   0.00590  -0.0325   0.5582   1.0000
   4.000   0.6940   0.01536   0.00619  -0.0324   0.5498   1.0000
   4.250   0.7206   0.01556   0.00644  -0.0322   0.5422   1.0000
   4.500   0.7475   0.01580   0.00682  -0.0322   0.5330   1.0000
   4.750   0.7741   0.01602   0.00716  -0.0319   0.5254   1.0000
   5.000   0.8009   0.01627   0.00759  -0.0318   0.5162   1.0000
   5.250   0.8266   0.01621   0.00762  -0.0312   0.4924   1.0000
   5.500   0.8518   0.01616   0.00764  -0.0305   0.4592   1.0000
   5.750   0.8755   0.01632   0.00763  -0.0298   0.3906   1.0000
   6.000   0.8868   0.01946   0.00893  -0.0294   0.0920   1.0000
   6.250   0.9046   0.02166   0.01073  -0.0289   0.0343   1.0000
   6.500   0.9262   0.02288   0.01218  -0.0283   0.0299   1.0000
   6.750   0.9464   0.02418   0.01370  -0.0277   0.0261   1.0000
   7.000   0.9620   0.02600   0.01575  -0.0269   0.0230   1.0000
   7.250   0.9786   0.02749   0.01744  -0.0258   0.0221   1.0000
   7.500   0.9937   0.02911   0.01920  -0.0244   0.0213   1.0000
   7.750   1.0093   0.03080   0.02098  -0.0229   0.0204   1.0000
   8.000   1.0268   0.03248   0.02276  -0.0215   0.0189   1.0000
   8.250   1.0453   0.03427   0.02466  -0.0203   0.0173   1.0000
   8.500   1.0656   0.03641   0.02693  -0.0191   0.0167   1.0000
   8.750   1.0865   0.03883   0.02954  -0.0180   0.0163   1.0000
   9.000   1.1060   0.04154   0.03258  -0.0169   0.0161   1.0000
   9.250   1.1228   0.04449   0.03586  -0.0157   0.0160   1.0000
   9.500   1.1360   0.04766   0.03940  -0.0144   0.0160   1.0000
   9.750   1.1448   0.05086   0.04295  -0.0131   0.0157   1.0000
  10.000   1.1493   0.05408   0.04648  -0.0117   0.0153   1.0000
  10.250   1.1495   0.05743   0.05010  -0.0104   0.0148   1.0000
  10.500   1.1451   0.06095   0.05388  -0.0090   0.0145   1.0000
  10.750   1.1352   0.06435   0.05752  -0.0074   0.0143   1.0000
  11.000   1.1234   0.06802   0.06143  -0.0067   0.0142   1.0000
  11.250   1.1104   0.07204   0.06567  -0.0068   0.0141   1.0000
  11.500   1.0968   0.07634   0.07019  -0.0076   0.0141   1.0000
  11.750   1.0827   0.08098   0.07504  -0.0091   0.0142   1.0000
  12.000   1.0676   0.08608   0.08035  -0.0113   0.0143   1.0000
  12.250   1.0517   0.09172   0.08619  -0.0142   0.0144   1.0000
  12.500   1.0348   0.09806   0.09273  -0.0179   0.0146   1.0000
  12.750   1.0160   0.10547   0.10032  -0.0225   0.0150   1.0000
  13.000   0.9854   0.11714   0.11223  -0.0307   0.0162   1.0000
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