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Glenn Martin 3 (glennmartin3-il) Xfoil prediction polar at RE=50,000 Ncrit=9


Details Polar file
Airfoil: Glenn Martin 3 (glennmartin3-il)
Reynolds number: 50,000
Max Cl/Cd: 32.45 at α=4.5°
Description: Mach=0 Ncrit=9
Source: Xfoil prediction
Download polar: xf-glennmartin3-il-50000.txt
Download as CSV file: xf-glennmartin3-il-50000.csv

  
       XFOIL         Version 6.96
  
 Calculated polar for: Glenn Martin 3                                  
  
 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
  ------ -------- --------- --------- -------- -------- --------
  -4.000  -0.2513   0.07270   0.06738   0.0346   1.0000   0.7917
  -3.750  -0.3042   0.07102   0.06589   0.0417   1.0000   0.7562
  -3.500  -0.3721   0.06901   0.06408   0.0498   1.0000   0.7179
  -3.250  -0.4460   0.06744   0.06273   0.0610   1.0000   0.7099
  -3.000  -0.5879   0.05931   0.05233   0.0418   1.0000   0.2582
  -2.750  -0.5696   0.05377   0.04582   0.0438   1.0000   0.1841
  -2.500  -0.5616   0.05123   0.04310   0.0472   1.0000   0.1774
  -2.250  -0.5528   0.04920   0.04045   0.0512   1.0000   0.1656
  -2.000  -0.5422   0.04730   0.03830   0.0542   1.0000   0.1610
  -1.750  -0.5311   0.04570   0.03631   0.0573   1.0000   0.1581
  -1.500  -0.5179   0.04439   0.03464   0.0598   1.0000   0.1571
  -1.250  -0.5034   0.04335   0.03336   0.0619   1.0000   0.1591
  -1.000  -0.4859   0.04233   0.03204   0.0633   1.0000   0.1592
  -0.750  -0.4652   0.04146   0.03090   0.0639   1.0000   0.1592
  -0.500  -0.4434   0.04090   0.03012   0.0638   1.0000   0.1623
  -0.250   0.0620   0.04286   0.03406  -0.0287   0.9639   1.0000
   0.000   0.1239   0.04329   0.03409  -0.0361   0.9353   1.0000
   0.250   0.1852   0.04337   0.03388  -0.0426   0.9077   1.0000
   0.500   0.2430   0.04299   0.03327  -0.0479   0.8793   1.0000
   0.750   0.2966   0.04236   0.03245  -0.0520   0.8516   1.0000
   1.000   0.3604   0.04129   0.03121  -0.0573   0.8292   1.0000
   1.250   0.4022   0.04043   0.03025  -0.0588   0.8018   1.0000
   1.500   0.4594   0.03888   0.02859  -0.0622   0.7802   1.0000
   1.750   0.5185   0.03689   0.02651  -0.0655   0.7607   1.0000
   2.000   0.5556   0.03559   0.02515  -0.0654   0.7338   1.0000
   2.250   0.6195   0.03307   0.02254  -0.0689   0.7126   1.0000
   2.500   0.6766   0.03104   0.02040  -0.0717   0.6842   1.0000
   2.750   0.7528   0.02891   0.01799  -0.0782   0.6484   1.0000
   3.000   0.8267   0.02812   0.01675  -0.0856   0.6024   1.0000
   3.250   0.8865   0.02851   0.01674  -0.0914   0.5591   1.0000
   3.500   0.9278   0.02931   0.01727  -0.0941   0.5282   1.0000
   3.750   0.9712   0.03013   0.01779  -0.0973   0.5021   1.0000
   4.000   1.0025   0.03109   0.01863  -0.0983   0.4823   1.0000
   4.250   1.0370   0.03211   0.01953  -0.1000   0.4660   1.0000
   4.500   1.0765   0.03317   0.02045  -0.1028   0.4521   1.0000
   4.750   1.0890   0.03432   0.02170  -0.1001   0.4409   1.0000
   5.000   1.1151   0.03551   0.02286  -0.1003   0.4309   1.0000
   5.250   1.1330   0.03665   0.02406  -0.0988   0.4220   1.0000
   5.500   1.1553   0.03800   0.02541  -0.0983   0.4143   1.0000
   5.750   1.1609   0.03942   0.02704  -0.0946   0.4082   1.0000
   6.000   1.1851   0.04076   0.02838  -0.0945   0.4018   1.0000
   6.250   1.1951   0.04237   0.03011  -0.0918   0.3971   1.0000
   6.500   1.1868   0.04419   0.03217  -0.0858   0.3930   1.0000
   6.750   1.1898   0.04587   0.03401  -0.0819   0.3880   1.0000
   7.000   1.2167   0.04727   0.03539  -0.0824   0.3830   1.0000
   7.250   1.2116   0.04941   0.03770  -0.0774   0.3798   1.0000
   7.500   1.1849   0.05201   0.04056  -0.0689   0.3776   1.0000
   7.750   1.1533   0.05486   0.04362  -0.0601   0.3761   1.0000
   8.000   1.1051   0.05823   0.04717  -0.0495   0.3753   1.0000
   8.250   1.0177   0.06457   0.05370  -0.0364   0.3757   1.0000
   8.500   0.8770   0.08012   0.06935  -0.0280   0.3801   1.0000
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