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Jet-Swirl Injector Spray Characteristics in Combustion Waste of a Liquid Propellant Rocket Thrust Chamber
Published in Dzaraini Kamarun, Ramlah Mohd. Tajuddin, Bulan Abdullah, Engineering and Technical Development for a Sustainable Environment, 2017
Zulkifli Abdul Ghaffar, Salmiah Kasolang, Ahmad Hussein Abdul Hamid
A liquid propellant rocket involves the combustion process in generating the thrust for the purpose of propelling the rocket. However, this process causes the emission of heat and wastes which is harmful to the environment and health. An injector discharging fine droplets size facilitates the combustion process and capable to increase the combustion efficiency. Besides the droplets size, other important spray characteristics of an injector is the spray angle. A larger spray angle increases the exposure of the droplets to the surrounding air or gas, which improves the rates of heat and mass transfer. Jet-swirl injector is one of the injector types capable of producing both fine droplets size and wide spray angle. In real applications, the fineness of the droplets size and the wideness of the spray angle are predominantly dependent on the operating conditions of the system and the injector geometrical design. In this chapter, a fundamental investigation into the characterization of the jet-swirl spray behavior has been discussed. The effect of liquid injection pressure, swirl chamber diameter and flow regimes on the characteristics of spray angle are presented.
Rocket Engine with Continuously Rotating Liquid-Film Detonation
Published in Combustion Science and Technology, 2020
S. M. Frolov, I. O. Shamshin, V. S. Aksenov, P. A. Gusev, V. A. Zelensky, E. V. Evstratov, M. I. Alymov
Currently, there are several promising trends of development of liquid-propellant rocket engines (LREs) in space propulsion technology. One of such trends is to replace continuous deflagrative (subsonic) combustion by continuous detonative (supersonic) combustion of the propellant mixture in the LRE combustor. The transition to continuous detonative combustion is advisable because the thermodynamic cycle efficiency of the detonative combustion engine is higher than that of the conventional deflagrative combustion engine (Zel’dovich, 1940), theoretically, by 13–15% (Chvanov et al., 2012). Moreover, in the detonation liquid-propellant rocket engine (DLRE), the combustor and the nozzle are more compact, and the detonative combustion is characterized by low emissions of hazardous pollutants. The energy efficiency of using continuous detonative combustion in a DLRE was experimentally proven in our previous studies with gaseous hydrogen–gaseous oxygen DLRE (Frolov et al., 2014; Frolov et al., 2015a). The specific impulse was shown to increase up to 7–8% when deflagrative combustion was replaced by detonative combustion in the same engine at the same supply pressures of mixture components.
Combustion Dynamics Simulation of a 30-Injector Rocket Engine
Published in Combustion Science and Technology, 2022
Juntao Xiong, Feng Liu, William A. Sirignano
Combustion instability in liquid-propellant rocket engines is a phenomenon involving coupling mechanisms between acoustic waves and flames that affect combustion dynamics and combustion instability (Poinsot and Veynante 2011). The acoustic oscillations occur at natural resonant frequencies for the injection-combustor-nozzle configuration and have amplitudes that rise above the level of combustor noise related to vortex shedding. For the combustion instability, flames engage in the resonant interaction and provide oscillatory burning rates that reinforce the acoustic oscillations. Combustion instability has been studied for many decades and is considered in all development programs for new rocket designs because it is an undesired and harmful phenomenon; the resulting large-amplitude pressure oscillation can modify the thrust vector and damage the engine through increased heat transfer to the walls and injector face. Harrje and Reardon (1972) edited an excellent compilation of works during the 1950s and 1960s. Oefelein and Yang (1993) presented a discussion of instability problems of the F-1 rocket-motor. During the early period, the physics of the oscillation were identified by Crocco and Cheng (1953, 1956)) for the longitudinal-mode linear instability. They developed the two-parameter (n, τ) coupling between combustion and acoustics for the stability analysis while Sirignano and Crocco (1964) and Mitchell, Crocco, and Sirignano (1969) performed nonlinear analyses for spontaneous and triggered longitudinal mode instabilities with shock-waves formation. Later, Zinn (1968) extended the theory for a transverse mode analysis. Two general types of acoustical combustion instability exist according to Culick (2006): driven instability and self-excited instability. Driven instability is more common in solid-propellant rocket engines while self-excited linear and nonlinear instability are common in liquid-propellant rocket engines.