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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.

This study presented the numerical simulation of the tangential combustion instability in a hypergolic liquid bipropellant rocket thrust chamber, which applied fuel liquid film cooling method and unlike impinging injectors. The liquid spray was modeled using Lagrangian approach, while the gas was regarded as Euler phase. Stress-blended eddy simulation and finite rate/eddy–dissipation model were adopted to simulate the turbulent combustion process. Consistent with the experiment results, this work successfully simulated the transformation of tangential combustion instability from standing mode to spinning mode. The mean pressure, amplitude and frequency of limit cycle oscillation were in good agreement with the experiment. There was a detailed analysis about the flow field, Rayleigh index, and driving mechanism of the combustion instability. It was found that the oscillation began with hot spots of heat release rate due to the interaction between the spray of impinging injectors and cooling fuel jet. More than that, cooling fuel jet also contributed to drive the oscillation. In the standing mode, injectors in the inner and outer rings drive the oscillation together, while the spinning mode is mainly driven by injectors in the outer ring. The pressure wave is subsonic and its Mach number is close to 1. It was shown that the pressure wave contained a complex structure divided into three parts. This led to the in-phase of the pressure along the axial direction and the double-peak characteristic of the downstream pressure signal. Besides, a positive feedback closed-loop system associated with periodic oxidizer/fuel ratio was believed to sustain the combustion instability. The oscillation can be maintained when pressure, heat release and oxidizer/fuel ratio are coupled together. The analysis results indicate that rotating detonation is an implication to tangential combustion instability. [ABSTRACT FROM AUTHOR]

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. • High frequency combustion instability in rocket engine is well captured. • Local heat release is altered by plasma, leading to a weakening in oscillation. • Plasma does not change the phase relationship between pressure and heat release. • Double-plasma actuators scheme is better for combustion instability control. [ABSTRACT FROM AUTHOR]

The quasi-DC discharge plasma across the backward facing step of a flame-holding cavity can be used to improve the flame stabilization of the cavity in a scramjet combustor. The effects of plasma filaments on the main flow structures, cavity drag, mass exchange rate, cavity oscillating characteristics, stagnation pressure loss, and combustion efficiency are numerically investigated based on the dominant thermal blocking mechanism. The results show that the cavity shear layer fluctuates dramatically due to the ‘cutting’ effect of plasma, which leads to obvious changes in the fuel mixing and burning in the combustor, and the pressure distribution near the rear edge of the cavity. Some symmetrical structures in the shape of ‘branch pipes’ are observed, which represent the distribution of product water. The periodical producing of them can be attributed to the movement of the cavity shear layer. Compared with the no plasma case, the cavity drag coefficient overall increases 11.3%, yet the mass exchange rate rises sharply over 10 times. The sound pressure level of the cavity representative points is raised by the plasma, and the first dominant frequency of pressure oscillation at the monitor point on the rear edge is twice the plasma actuation frequency. Although the stagnation pressure loss increases, the plasma shows benefits in energy producing that outweigh the energy consumed by the plasma actuator. [ABSTRACT FROM AUTHOR]

A plasma filament is generated upstream of the fuel jet through quasi-dc discharge in a scramjet combustor. The effects of the plasma filament on fuel mixing and combustion are numerically investigated based on the dominant thermal blocking mechanism. Results have shown that temperature distribution of the whole flowfield is changed prominently due to heat release from the plasma filament. Since the shock waves induced by the transverse fuel jet are partly weakened and move upstream by the plasma filament, the separation zone upstream of the fuel orifice can be controlled to provide a better combustion environment. Compared with the no-plasma case, more water appears in higher space above the bottom wall while less near the bottom wall of combustor. Although the stagnation pressure loss increases a little owing to the plasma filament, the combustion efficiency rises obviously with only a little external energy addition. [ABSTRACT FROM PUBLISHER]

The process of dielectric barrier discharge (DBD) plasma flow-separation control on screw propeller is numerically simulated based on the unsteady compressible Navier-Stokes equations loosely with DBD-plasma body force aerodynamic actuation mechanism. The effect and the mechanism of plasma flow control to enhance the propeller aerodynamic characteristics are discussed. The results show that the propeller aerodynamic characteristics will decrease as the altitude increase without plasma flow control, but it will first increase and then decrease as the altitude increases with plasma flow control. The effect of DBD-plasma flow control on the distribution of pressure is small at ground condition, and the difference of pressure distribution between general propeller and plasma efficiency-enhance propeller will be increased as the altitude increases. The effect of DBD-plasma flow-separation control will increase as the altitude increases. Mainly, the mechanism of the plasma flow control to enhance the propeller aerodynamic is obtained as follows. The flow separation on the propeller blade surface could be effectually controlled by DBD-plasma aerodynamic actuation, thereby increasing the efficiency of the propeller. The pressure of the windward surface of the blade will decrease by the plasma body force, which will increase the thrust of the propeller. [ABSTRACT FROM PUBLISHER]

The detonation wave in a rotating detonation engine is highly adaptable to the incoming flow, making the wave easier to control. In this study, a numerical simulation method is used to analyze the working process and flow field structure of a rotating detonation model engine with dual cavity injection of an H 2 /air mixture by controlling the injection pressure ratio of the dual cavity and the number of detonation wave heads. It is found that the rotating detonation engine offers the possibility to control the thrust vector with two different modes. The first is a one-cycle alternate control mode with a small injection pressure ratio. Here two deflections occur in different directions occur across one detonation wave propagation cycle, but the overall deflection direction is in the low-pressure region. The second is a one-way control mode, with a large injection pressure ratio, and the deflection direction towards the low-pressure region. For the multi wave-mode, it belongs to one-way control mode because of constant deflection direction in the low-pressure area. From the perspective of thrust distribution along the circumference, the one-way control strategy satisfies the ability of a rotating detonation thrust vector control. • RDE theoretically owns the ability of the thrust vector control with two modes. • Analyze the specific impulse and thrust of the rotating detonation engine during the processes of the thrust vector control. • One-cycle alternate and one-way control modes have different processes, but the latter is more suitable for RDE applications. • Numerical simulation with dual cavity injection was adopted to analyze the TVC working process and flow field structure. [ABSTRACT FROM AUTHOR]

Self-excited longitudinal combustion instabilities were analyzed in a sub-scale combustor operating with a propellant combination of O 2 /CH 4. Couplings between acoustics, hydrodynamics, and heat release were investigated numerically based on an experimental setup. The numerical results were in good agreement with the experimental data. Various post-processing methods have been adopted to explain the thermoacoustic behavior and reveal the underlying unstable combustion mechanisms. The classical longitudinal acoustic characteristics of the combustion chamber were identified by the pressure mode shape and dynamic mode decomposition (DMD) analysis. It was found that the incoming streams along the injector were periodically hindered by the traveling pressure wave originating from the chamber, thus, resulting in an oscillatory propellant supply and a pulsating heat release. This, in turn, feeds energy into the acoustic pressure oscillations, and maintains the combustion instability. Meanwhile, the coherent vortex structures in the reacting mixing-shear layer were found to be the principal reason for sustaining combustion instabilities. Besides, a positive feedback closed-loop system associated with periodic vortex shedding and merging is believed to sustain combustion instabilities. The plot of the Rayleigh index further confirmed that the shear layer region drives thermoacoustic instabilities. [ABSTRACT FROM AUTHOR]