Optimization of mixture ratio distribution in liquid propellant rocket thrust chamber

1992 ◽  
Vol 8 (3) ◽  
pp. 605-608 ◽  
Author(s):  
K. Ramamurthi ◽  
A. Jayashree
Keyword(s):  
Liquid Propellant ◽  
Ratio Distribution ◽  
Mixture Ratio ◽  
Thrust Chamber ◽  
Rocket Thrust ◽  
Propellant Rocket

A Comparison of Analytical and Experimental Local Heat Fluxes in Liquid-Propellant Rocket Thrust Chambers

1962 ◽  
Vol 84 (1) ◽  
pp. 19-28 ◽  
Author(s):  
William E. Welsh ◽  
Arvel B. Witte
Keyword(s):  
Heat Flux ◽  
Flow Characteristics ◽  
Heat Fluxes ◽  
Chamber Pressure ◽  
Liquid Propellant ◽  
Nitrogen Tetroxide ◽  
Thrust Chamber ◽  
Experimental Heat ◽  
Rocket Thrust ◽  
Propellant Rocket

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.

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
Keyword(s):  
Ratio Distribution ◽  
Mixture Ratio ◽  
Thrust Chamber ◽  
Rocket Thrust

scholarly journals Thermal Analysis on Regenerative Cooled Liquid Propellant Rocket Thrust Chamber Using ANSYS

2021 ◽  
Vol 1059 (1) ◽  
pp. 012068
Author(s):  
Pon. Maheskumar ◽  
R. Girimurugan ◽  
G. Sivaraman ◽  
S. Purushothaman ◽  
M. Vairavel
Keyword(s):  
Thermal Analysis ◽  
Liquid Propellant ◽  
Thrust Chamber ◽  
Rocket Thrust ◽  
Propellant Rocket

scholarly journals Investigation of a Radial Turbines Compatibility for Small Rocket Engine

2020 ◽  
Vol 197 ◽  
pp. 11009
Author(s):  
Angelo Leto
Keyword(s):  
Rocket Engine ◽  
Pressure Ratio ◽  
Liquid Propellant ◽  
Axial Turbine ◽  
Thrust Chamber ◽  
Radial Turbine ◽  
Liquid Propellant Rocket Engine ◽  
Radial Turbines ◽  
High Thrust ◽  
Propellant Rocket

In the radial turbine preliminary design for an expander rocket engine, a comparison was made with axial turbine used in Pratt & Whitney RL10 engine. One of the primary requirements of a liquid propellant rocket engine is the generation of a high thrust, which depends on both the mass flow rate of the propellant and the pressure in the thrust chamber. In expander-cycle engines, which are the subject of the present study, the liquid propellant is first compressed using centrifugal turbo-pumps, then it is used to cool the combustion chamber and the nozzle and, once vaporized, it flows through the turbines used to drive the turbo-pumps. The aim was to demonstrate the greater efficiency of the radial turbine with a reduction of the pressure ratio with respect to the axial turbine.

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