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As a promising technology, plasma-assisted combustion (PAC) has attracted many researchers to explore the effect of PAC on improving the combustion in propulsion devices, such as scramjet, detonation engines, internal engines, and so on. In this chapter, we aim to exhibit the influence of quasi-DC discharge plasma on the operating performance of scramjet combustor and find the internal mechanisms, which may contribute to the development of PAC technology in supersonic combustion. For case one, a plasma filament is generated upstream of fuel jet through quasi-DC discharge in a scramjet combustor; for case two, the plasma is formed across the backward facing step of a flame holding cavity to improve the flame stabilization of the cavity in the scramjet combustor. The two cases are investigated in detail through three-dimensional numerical simulation based on the dominant thermal blocking mechanism. Important parameters including temperature distribution, separation zone, water production, stagnation pressure loss, combustion efficiency, cavity drag, mass exchange rate, and cavity oscillating characteristics are obtained and analyzed. It shows that the quasi-DC discharge plasma does benefit for the improvement of the combustion in a scramjet combustor.

Non-equilibrium plasma has significant chemical kinetic effect and aerodynamic effect, both of which play an important role in the plasma-assisted combustion. In this paper, the numerical study on plasma-assisted combustion in gaseous methane rocket engine is carried out to investigate the impact of kinetic effect on its combustion performance. There are three schemes of plasma excitation via oxygen, methane and both of them in the numerical simulation model. The variations of combustion chamber temperature, pressure, species concentration and special impulse with different dissociation degrees are analysed in three discharge schemes. The results show that non-equilibrium plasma can effectively increase methane burning rate, shorten unburnt area upstream of combustion chamber, and increase the width of combustion zone in shear layer upstream of combustion chamber. And this phenomenon becomes more obvious with the increase of excitation intensity. In addition, more methane and oxygen are added to the combustion process under plasma actuation, and the mole fraction of water is increased in combustion products. The specific impulse of engine is also increased under plasma actuation. At the same degree of dissociation, the combustion-supporting effect of simultaneous discharge of oxygen and methane is significantly greater than that of any component discharge. And in the condition of single component discharge, the combustion-supporting effect of methane is better than that of oxygen.

The combustion instability encountered unexpectedly in hot-firing tests of the YF-960 liquid oxygen/gaseous hydrogen (LOX=GH2) rocket engine was numerically researched. Here, the emphasis was on evaluating the control effects of two kinds of baffle configurations on combustion instability. The integrated physical model and numerical method were presented to solve the turbulent, three-dimensional, two-phase, and full-size LOX=GH2 rocket combustion process. The dominant oscillation frequency characteristics that appeared in the hot-firing tests were reproduced andconfirmedfirst,which validated the computationalfluiddynamics code to a certain extent. Thebaffle effects were investigated by the developedmass flow flux perturbingmodel. Results show that the control effect of the hub-and-six-blade baffle on the one-order tangential-mode oscillation frequency of 5800 Hz is better than that of the hub-and-three-blade baffle. Although the two kinds of baffles have similar effects on the radial-mode instability, the control effect is better fordamping the5800Hzoscillation than the2900Hzoscillation.This researchhasprovided direct design data for engineers, and it is valuable in further understanding the baffle-eliminating mechanisms.

[Abstract] The present study has been focused on measuring the combustion process of gelled propellant in high-pressure oxidant environment. The research object consisted of gel droplet and gel spray due to different extruding speed for the gel injector. Results show that the burning behavior of gel droplet in high-pressure oxidant environment is very different from the burning behavior of gel droplet in room pressure conditions. The burning flame structures of gel droplets in high-pressure conditions evolved very violently. The microexplosion phenomenon was more vigorous usually resulting that the maternal droplets broken into several smaller droplets which dropped off during their burning processes. And while the smaller droplets were dropping off, the long-plume from burning flames were protracted behind them. In a moment when the smaller droplets burned out, their burning flames become more luminous instantaneously. And it is interesting that the dropping route of the smaller droplet was not downward vertically rather than floated due to the synthetic effects of gravity, buoyancy, and microexplosion force, and so on. Additionally, the combustion process of gel spray was accelerated as the pressure in chamber was increased.

[Abstract] The present study is focused on measuring the atomization and combustion behavior of gelled propellant. The main aim is to examine the characteristics of the flame structure, the spray breakup length and the spray angle of gelled droplets at different initial chamber conditions and initial spray parameters. Results show that the gelled liquid ligament burns up quickly after only about 0.0099 sec at a certain condition. The gelled liquid ligament burns up more quickly than a gelled droplet. Additionally, the injected velocity and the width of gelled liquid ligament have an influence on the burning time. The spraying ligament would be shrunk and a droplet separated from the liquid ligament during the burning process.

[Abstract] The active control technology for preventing the combustion instability in a liquid rocket engine has been developed. Aiming at the liquid rocket engine combustion process, the integrated physical model and calculation model was presented. Then the numerical simulation of the combustion instability in the tripropellant liquid rocket engine has been conducted by studying the control equation and physical model of turbulent twophase chemical reaction flow for the combustion process. The results show that combustion instability may be restrained when the proportion of hydrogen is increased, however, it is introduced more easily for smaller average diameter of droplet and too high or too low eject velocity. Finally, According to real time adjusting of propellant flow based on the variety of combustor pressure, combustion instability is controlled actively. The simulation results show that it is feasible that combustion instability of liquid rocket engine can be controlled actively by adjusting propellant flow flux. Nomenclature U = axial velocity respectively V = radial velocity respectively ρ = density T = temperature Sϕ = source terms due to coordinator transform Sd,ϕ = source terms due to two phase flow SC,ϕ = source terms due to chemical reaction P

[Abstract] Based on detailed analyzing of flow field, accurate simulation of radiation of the liquid rocket engine exhaust plume plays an important role for detecting, identifying, tracking in missiles defense systems. The fully coupled Navier-Stokes equations with additional terms describing one step chemical kinetics of CH4 and O2 in chamber have been used for incorporate numerical simulation of flow fields in engine and plume at 5km, 8km,20km, 30km, 40km 50km and 80km. Considering two steps reaction in chamber, plume flow field at 5km is also studied particularly. According to the effect of atmosphere attenuation and transmissivity on radiation, a generated model has been developed to calculate optical radiation parameters of the plumes at 5km. Some key features in expanding of the plume into the ambient environment have been gained, such as the location of the barrel shock impingement, the shock reflection structure, thickness of the shear layer and afterburning processes, as well as flow field temperature, pressure, velocity and mass fraction of products. The simulation results show that the shape and characteristic of exhaust plume are determined by combustion model in chamber and ambient environment. The plume produces powerful radiation over the infrared spectral ranges, and conditions of atmosphere and transmission distance have a great influence on the distribution of spectrum.

[Abstract] The aluminum oxide (Al2O3) liquid or solid particle radiation is significant in the plume radiation of an aluminized solid-propellant rocket motor. This paper has researched the influences of Al2O3 particles size and initial velocity at the burning surface in chamber on the exhaust plume flow field and radiation, as well as the Mach number of the missile. Numerical simulation has been done to learn inner and outer integrative two-phase flow field and flow field parameters distribution of gaseous components and Al2O3 particles in many situations have been obtained, compared and analyzed. Structured grid generated by TTM method has been employed. The two-phase flow field was solved by discrete particle model, where the PSIC (Particle Source in Cell) method was employed. The Rosin-Rammler algorithm was used for particle diameter distribution. Then spectral radiance of two dimension axial symmetry plume has been computed by developed model of infrared radiation of solid rocket exhaust plume, and compared with that in document. The study shows that the radiation of particles plays the main roal in the infrared radiation of exhaust plume. The quicker the missile is, the weaker the gaseous radiation is, however, the stronger the solid radiation is. For the same particle dimension, the quicker the burning surface particles are, the weaker the gaseous radiation is, however, the stronger the solid radiation is, and the more particle dimension is, the bigger the influence of particle velocity on plume radiation is. Examination of the gaseous molecular spectra shows that there are three spectral peaks at 1.9-2.2μm, 2.7-2.8μm and 4.3-4.66μm. In addition, their intensities are related to particle dimension and velocity, as well as solid radiation. It indicates the method of this paper is feasible and the solution is proper.