Mixture ratio distribution - Its impact on rocket thrust chamber performance.

1967 ◽  
Vol 4 (6) ◽  
pp. 786-789 ◽  
Author(s):  
J. L. PlEPER ◽  
L. E. DEAN ◽  
R. S. VALENTINE
2021 ◽  
pp. 1-11
Author(s):  
Thomas Govaert ◽  
Wolfgang Armbruster ◽  
Justin S. Hardi ◽  
Dmitry Suslov ◽  
Michael Oschwald ◽  
...  

2019 ◽  
Vol 35 (3) ◽  
pp. 632-644 ◽  
Author(s):  
Wolfgang Armbruster ◽  
Justin S. Hardi ◽  
Dmitry Suslov ◽  
Michael Oschwald

1962 ◽  
Vol 84 (1) ◽  
pp. 19-28 ◽  
Author(s):  
William E. Welsh ◽  
Arvel B. Witte

Experimental data are presented showing heat-flux distributions measured calorimetrically with several liquid-propellant rocket thrust-chamber configurations. Thrust levels of the experimental chambers were from 300 to 5000 lb. Enzian-type and axial-stream showerhead propellant injectors were utilized with hydrazine (N2H4) and nitrogen tetroxide (N2O4) propellants. Nozzle-contraction-area ratios of 8 to 1, 4 to 1, and 1.64 to 1 were tested, each having a 5-in. inlet diameter. Characteristic chamber lengths ranged from 16.95 to 62.8 in. The comparison between the experimental heat flux and the analytical heat flux using the method of Bartz [1] was found to be closest in the nozzle-expansion region. The experimental heat-flux measurements ranged between 80 per cent above and 45 per cent below the analytical estimates at the nozzle throat, however. These differences were dependent upon thrust-chamber configuration, injector type, and chamber pressure, and apparently resulted from nonideal combustion and flow characteristics. It is concluded that a priori determination of heat-flux distribution along the thrust-chamber length was possible only to a first approximation for the conditions of these tests.


2014 ◽  
Vol 592-594 ◽  
pp. 1692-1696
Author(s):  
A. Anoop ◽  
M.P. Assiz ◽  
M. John Manu

Initiation of chemical reactions liberating thermal energy in non hypergolic propellants require ignition systems which are reliable as well as robust in nature. This necessitates the igniter and its subsystems be on par or exceed every quality criteria required for the main stage. The ignition systems which provide pilot action and flame holding in supersonic environment adds new challenges in its structural and thermal stability. The paper aims to conduct a numerical analysis of the Augmented Spark Torch igniter required for an air breathing rocket (SCRAMJET) engine with the emphasis on the optimization of the mixing length ratio required to achieve a uniform mixture ratio distribution in the zone of the igniter combustion chamber. Nomenclature


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