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To suppress the high frequency combustion instability of rocket engines, quasi-direct current (quasi-DC) discharge plasma is considered in combination with special arrangements in a rocket engine combustor. After establishing an internal reacting flow field model of a rocket combustor and a phenomenological plasma model, the influence of quasi-DC discharge plasma on high frequency combustion instability is numerically investigated. The results show that the characteristics of the high frequency combustion instability in the combustor are captured very well through the simulation. Based on the DES turbulence model, a positive feedback process is clearly observed for the self-excited combustion instability. The mean temperature and pressure of the combustor are not sensitive to the occurrence of quasi-DC discharge plasma. The pressure histories of the combustor and their characteristic oscillating frequencies are not affected by the plasma when only one actuator is in operation. However, the oscillating amplitudes of the pressure are reduced within a certain time when several actuators are in operation. The heat release becomes more diffusive and more eddies form around the injector faceplate with an increasing number of plasma filaments. Mainly through the reduction of local heat release, the plasma is capable of affecting the high frequency combustion instability in the combustor. To better control high frequency combustion instability, the optimal arrangement is a double-plasma filaments with symmetrical actuators.

This study numerically investigated the thermoacoustic combustion instability characteristics of a scaled rocket combustor based on a hybrid of the Reynolds-averaged Navier–Stokes and large–eddy simulation method. The turbulence–combustion interactions were treated using flamelet generated manifold approach. An unstable case was simulated with detailed reaction mechanisms (GRI-Mech 3.0). The obtained results agree well with experiment data from Purdue University, in terms of pressure oscillations frequency and power spectral density spectrum. The combustion instability mode was identified to be coupled with the first longitudinal acoustic mode of the combustion chamber by dynamic model decomposition method. According to Rayleigh index analysis, the unstable driving source was found to be located near the combustor step, which was further confirmed by time-averaged flow fields. Detailed three-dimensional vortex ring shedding evolutions at the combustor step were tracked with fine time resolution. Results indicate that the combustion instability arises from periodic vortex ring shedding at the combustor step and interacting with the chamber wall. The unburnt reactants were rolled up by the shedding vortex ring, which would not break up until impact with the chamber wall. Therefore, the mixing performance was significantly enhanced, leading to sudden heat release. Consequently, the thermal energy is added to the acoustic field, and the first longitudinal mode is thus reinforced, giving rise to large amplitude axial velocity oscillations which prompt the generation of the new vortex ring. The results of the present investigation will support the design and development of high-performance rocket engines.

There is a lack of understanding of the spray characteristics of gas-centered swirl coaxial (GCSC) injectors during self-pulsation occurs. Therefore, the self-pulsation of a GCSC injector was investigated experimentally in this study. Experiments were conducted at atmospheric pressure with filtered water and dried air supplied through a propellant feed system. A back-lighting high-speed photography technique was used to capture unsteady spray features. A laser-based particle size analyzer (LPSA) was used to measure the size of the droplets in the spray. The effects of recess and gas-liquid ratio on spray self-pulsation were analyzed. It was found that the recess of the injector strongly determines the spray pattern. When spray self-pulsation occurs without recess, both the center and periphery of the spray oscillate. With an increase in the mass flow rate of the gas, the boundary between the center and the periphery of the spray becomes more noticeable. Meanwhile, small droplets in the spray center oscillate, with the periphery of the spray being characterized by a periodic “shoulder.” Under the same operating conditions but with a small recess (2 mm), the spray adheres to the injector faceplate. With a larger recess (7 mm), when spray self-pulsation occurs, the spray periodically forms “shoulder” and “neck,” similar to the behavior of self-pulsation in a liquid-centered coaxial injector. Therefore, it can be concluded that spray self-pulsation enhances atomization at the center of the spray to a certain extent. However, atomization becomes worse in the periphery with an oscillating spray.

The numerical calculation of three-dimensional unsteady combustion for the combustion chamber of LOX/kerosene high pressure staged combustion rocket engine was carried out. By changing the offset ratio of oxygen mass flow rate in the edge area of the injector face, computational studies were conducted to investigate the effects of non-uniform distribution of oxidizer flow on combustion instability for a liquid-propellant rocket engine. The calculation results show that the offset ratio of oxygen mass flow rate changes the distribution of heat release in the combustion chamber. Within a certain range of offset ratio, the non-uniform distribution degree of oxidizer flow enhances the coupling between the pressure and heat release. As a result, it leads to an increase in the pressure oscillation amplitude in the combustion chamber. However, if the offset ratio is too large, the oxygen-fuel ratio will be too small in some regions, which will reduce coupling between the pressure and heat release and increase the damping of combustion instability.

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.

Gelled hypergolic propellants offer potential safety and performance improvements over conventional liquid and solid propellants. Understanding the combustion process of single gelled droplets is the first basic step to predict their behavior in the future combustion chambers. Experimental studies were performed to investigate the burning behavior of gelled unsymmetrical dimethylhydrazine (UDMH) droplets under oxidant convective conditions. The effects of oxidant convective conditions including velocity and temperature on the burning behavior were analyzed. The burning process was broken down into four stages: heating and swelling period, initial combustion period, violent combustion period, and stable and extinguished combustion period. Sometimes the droplet inside seemed to be porous or botryoidal. The microexplosion period lasted for a long time sometimes exceeding about 70% of the burning lifetime, and the phenomenon of gas jet combustion due to the burst steam from microexplosions was founded. The conversion of burning flame from a layered and enveloped flame structure to an escaped flame structure with increase in the convective velocity was observed. When the enveloped flame appeared, it was more helpful to the burning process, and the intensity of microexplosions and the burning rate constant were found to increase with the convective temperature. When the escaped flame appeared, it was disadvantageous to the burning process, and the intensity of microexplosions decreases with rise in convective velocity. Compared with the escaped flame, the intensity of microexplosions of the enveloped flame was lower, but the frequency was higher.

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.

The evolutions of dielectric barrier discharge (DBD) plasma induced starting vortex and the transformation of the DBD plasma induced flow field at different operating pressures are obtained by particle image velocimetry. The results show that the starting vortex will diffuse to the embedded electrode side and then disappear, and the diffusivity will decrease as the time goes on at high operating pressure. But the starting vortex will not change in location, and it will not disappear at low operating pressure. As the operating pressure decreases, the starting time of plasma induced flow field decreases, the ordinate orientation of the induced flowfield increases, and the landscape orientation of the induced flowfield decreases. The form of the induced streamline will come through L→U→V as the pressure decreases. There is no induced vortex in the form of L. There are two induced vortexes in the form of U which lay in the middle and on the right hand of the form U respectively, and there is one induced vortex in the form of V which lays in the middle of form V.

The organic gel propellants have the advantages of both solid and liquid propellant, but the catching fire and burning process of organic gel droplet appears obvious vapor ejection and flame disturbance. This unsteady combustion phe- nomenon of organic gel droplet is different from the quasi-steady combustion of conventional liquid droplet and produces great influence on the spray combustion flow field of gel liquid rocket engine and its optimization design. Based on the un- steady evaporation characteristics inside UDMH (unsymmetrical dimethylhydrazine) organic gel droplet and the micro molecule structure of organic gel propellant, three kinds of liquid phase evaporation model, the unsteady multicomponent evaporation model, the interface tracking multiphase flow evaporation model and the discrete element multiphase flow evaporation model, are proposed. The unsteady multicomponent evaporation model regards the gel droplet as a binary multi- conponent droplet, in which the gellant inside the gel droplet is been equivalent to a species with low volatility and high boiling temperature. In order to depict the shape and location variety of vapor bubble inside the gel droplet more precisely, the interface tracking multiphase flow evaporation model regards the gel droplet and the vapor bubble inside it as a single fluid, and uses the Level Set function tracking the interface location of vapor bubble. Based on the micro molecule structure of organic gel propellant, the discrete element multiphase flow evaporation model for the gel droplet regards the gellant and vapor bubble as the aggregation of micro-particles with different properties respectively, and uses the Lagrange method tracking the movement and aggregation of each kind micro-particle. The multicomponent evaporation model has been im- plemented in our foregoing work, so a summary of the model is presented. And next the main difficulties and the technique preparation about the last two multiphase flow models have been presented and analyzed in detail. In view of mixing and combustion process of gas phase surrounding the UDMH organic gel droplet, we propose that the finite rate chemical reac- tion model and the flamelet model should be used for the stagnation ambient and convective ambient respectively. The diffi- culties and numerical methods for handling the stiff chemical source term in finite rate chemical reaction model have been presented. And next the consistent flamelet model which takes consideration of species diffusion coefficient and non-unity Lewis number is presented. It would be more precise for the temperature and species interpolation of the gel droplet laminar non-premixed combustion flow field. Keywords organic gel; droplet; evaporation model; finite rate chemistry model; flamelet model; stiff source term

The experiments to generate vortices continuously using sub-microsecond pulsed surface dielectric barrier discharge (SDBD) actuator are conducted by particle image velocimetry (PIV). The double-frequencies actuation mode is presented which includes repetitive pulse frequency and vortex frequency. It is found that the empty zone of PIV particles appears in the place where the particles are quite few even nil during the experiments. When discharges occur the primary empty zone is produced by the micro explosion due to the released heat of plasma, and when discharges end the secondary empty zone appears. The induced flow is farther apart form wall and the influence of wall friction should be suppressed due to primary empty zone. When the secondary empty zone on the left side of actuator exposed electrode is blown away completely, the next actuation can start. In order to control the flow more effectively, the pulse voltage with higher repetitive frequency should be applied. The pulse number during one vortex time should be more than 10. As the pulse number increases, the maximal velocity of induced flow increases but the momentum transfer efficiency decreases. The mechanisms releasing heat and body force can be triggered by using the sub-microsecond pulse SDBD actuator.

Gel propellant has the advantage of controllable flux as liquid propellant and long-term reservation as solid propellant, however, the evaporation and combustion problem of gel spray droplet bores with the gel propellant development and combustor design all the time, and hampers gel propellant practical engineering applications. In this paper, the gel single droplet combustion experiment system is designed and constructed, and then the evaporation and combustion mechanism is explored deeply based on the experimental phenomena of organic gel unsymmetrical dimethylhydrazine (UDMH) single droplet burning in nitrogen tetroxide. The organic gel spray droplet multi-component evaporation model is developed for three different evaporation phases of the gel layer, i.e. the forming, expanding and bursting of the gel layer based on the single droplet evaporation characteristics in experiment, and then the gel single droplet vaporization in high temperature gas phase is numerically simulated and compared with the result of conventional liquid droplet using the elementary model parameters and physic property parameters. The result shows that the gel content on the droplet surface increases slowly at the beginning of evaporation, however it would increase rapidly to a mass fraction of 95% and form the gel layer in a limited time after exceeding a critical evaporation time, which results in surface mass flux dropping to 0 and surface temperature reaching the UDMH boil point rapidly. After the gel layer forming, the droplet radius and surface UDMH vapor mass fraction exhibit oscillation as the swelling-bursting phenomena in experiment. The gel droplet surface temperature holds above the boil point and the mass flux of gel droplet inner boiling evaporation is stronger than the conventional liquid droplet surface steady evaporation which makes the life time of gel droplet much shorter.

Based on the physical processes of dielectric barrier discharge (DBD) and quasi-direct-current (quasi-DC) discharge, the plasma aerodynamic actuation mechanism is analyzed, then the numerical model of plasma aerodynamic actuation is founded, finally the DBD and the quasi-DC discharge plasma flow control processes are simulated in the cases of low velocity and high velocity. The results show that the aerodynamic actuation mechanism of DBD plasma is that the discharge changes three kinds of forces in continuum fluent medium, these being shear stress caused by Newton friction, body force caused by electro hydrodynamic and impulsive active force caused by pressure change, and the main aerodynamic actuation mechanism of DBD plasma is body force caused by electro hydrodynamic. The effect of body force is stronger in near space than in the sea level, plasma induced flow velocity increases in near space. The main aerodynamic actuation mechanism of quasi-DC discharge plasma in supersonic air flow is the thermal mechanism of heat plasma, the exploding wires diathermanous model found in this paper is good for the simulation of the process of surface quasi-DC discharge plasma incident shock. The effect of quasi-DC discharge plasma on the supersonic flow field is conform with the effect of protuberance with a bevel to the supersonic flow field, so it can be used to control the shock wave in supersonic aircraft.