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Thermoacoustic instability (TAI) presents a critical challenge for lean-burning combustors and rocket engines. The early detection of instability is crucial, and to address this, a data-driven prediction framework has been established for TAI in a sub-scale rocket combustor with variable chamber length. Nonlinear combustion features are generated from time series of dynamic pressure using recurrence matrices. Deep learning models are then utilized to train these features and predict the proximity of impending TAI. The performance of the proposed method is investigated through cross-validations of 12 groups of hot-fire test datasets. Remarkably, the prediction performances are in good agreement with measured experimental data, with most instabilities being predicted dozens of milliseconds in advance. This capability paves the way for the early implementation of active control systems in full-scale combustors in the future. The prediction performances are also compared with state-of-the-art TAI prediction methods. [ABSTRACT FROM AUTHOR]

A methane/oxygen mixture is considered to be an appropriate propellant for many future rocket engines due to its practicality and low cost. To better understand the combustion instability in methane/oxygen-fed rocket engines, the spontaneous transverse combustion instability in a rectangular multi-element combustor (RMC) was analyzed both experimentally and numerically. Severe combustion instabilities occurred in the RMC during repeatable hot-fire tests. The physical mechanisms were systematically investigated through numerical simulations based on the stress-blended eddy simulation and flamelet-generated manifolds method with detailed chemical mechanisms (GRI Mech 3.0). The numerical results for the frequency spectrum and spatial modes agree well with the theoretical analysis and experimental data. The driven regions of the combustion instability were identified on both sides of the combustion chamber through a Rayleigh index analysis. The longitudinal pressure oscillations in the oxidizer post were found to be coupled with the transverse pressure waves in the combustion chamber and led to periodic oscillations of the mass flow rate of propellant. Moreover, the mixing was highly enhanced when the pressure wave interacted with walls of the combustion chamber. Therefore, a sudden release of heat occurred. The pressure oscillations were enhanced by pulsated heat release. A closed-loop system with positive feedback associated with periodic oscillations mass flow rate of the propellant, and sudden heat release, was believed to account for the present combustion instability. [ABSTRACT FROM AUTHOR]

In order to study the effect of different equivalence ratios on plasma assisted detonation initiation under a rich burn condition, a loosely coupled method was adopted to simulate the detonation formation by alternating current dielectric barrier discharge in a hydrogen–oxygen gas mixture with different equivalence ratios. The spatial and temporal evolution of discharge products were analyzed first. Then, the species distribution, Mach number, thrust wall pressure, and whole history of detonation formation were examined in detail. Results showed that the shapes of the temporal and spatial distribution of discharge products do not change when equivalence ratio varies in one discharge cycle. However, due to the variation of the percentage of fuel and oxidizer, and its impact on the discharge elementary reactions, the number density of every key active particle declines while the decline amplitude decreases when equivalence ratio rises. Although the magnitudes of species concentration are not altered by the plasma, the reacted region extends towards the downstream flow more quickly. The influence of equivalence ratio on plasma assisted detonation initiation becomes remarkable until late in the subsonic stage of the flowfield evolution process in the combustor under a rich burn condition. Furthermore, a larger equivalence ratio leads to a better accelerating effect of the plasma. Through analyzing the dynamic process of thrust wall pressure, it is also found that the DDT process is expedited more notably under a larger equivalence ratio, yet the magnitude of pressure is independent of the plasma. The provision of active particles for the combustion reaction by the plasma and its dilution effect on the fuel are the two intrinsic reasons for the reduction of DDT time and distance and the increase of this reduction degree, when equivalence ratio rises. [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]

Rocket engines are prone to suffer from transverse combustion instabilities, manifesting as periodic oscillations in pressure and heat release rate. Self-excited transverse instabilities in a rectangle multi-injector rocket combustor powered by n-C 12 H 26 and O 2 , are investigated experimentally and numerically in this paper. Transverse combustion instabilities are captured during repeatable experiments, while dynamic characteristics in the combustion chamber and driving mechanisms of instabilities are analyzed via numerical simulation based on stress-blended eddy simulation. Dynamic mode decomposition analysis reveals coupling behaviors between the transverse mode in the combustion chamber and longitudinal mode in the oxidizer post, resulting in propellant mass flow rate oscillations. Substantial evidence demonstrates that the intensities of heat release rate near the end walls are higher than that detected around the chamber center due to its intense coupling to the local first transverse (1W) acoustic mode. Flow field animations indicate strong deflection of oxygen posts and liquid particles by a transverse moving gas mixture leading to flame interactions and collision between adjacent share injectors. The propellant injected into the chamber does not burn immediately and is transported to the combustor end wall. Propellant mixing performance is significantly enhanced during interactions between acoustic pressure and end wall, causing a sudden heat release near 1W pressure anti-node, thereby strengthening pressure oscillations. Finally, a combustion instability driving mechanism associated with propellant mass flow rate oscillations, and interactions between unsteady acoustic pressure and heat release rate is proposed. The insights obtained from this work can aid comprehension of transverse combustion instabilities encountered by rocket engines. • Both liquid fuel and gaseous oxygen are highly deflected by transverse moving gas mixture when transverse combustion instability appearing. • Large flame interactions and collision structures appear inside the combustion chamber. • Combustion instability is driven by pulsated heat release rate near the end walls and periodic supply of propellant mass flow rate. [ABSTRACT FROM AUTHOR]

The diffusion process and spatial distribution of a liquid jet in supersonic crossflow (Mach number of 2) in a model combustor with an expanded section were investigated on the basis of experiments and simulations. Experiments were performed using high-speed photography, high-speed shadowgraph and laser-based particle size analyzer to investigate the spray structures, flow fields and droplet diameter, respectively. The Eulerian-Lagrangian method coupled with the Kelvin-Helmholtz/Rayleigh-Taylor breakup model were used in the simulations. Moreover, delayed detached eddy simulation numerical method based on the two-equation shear stress transport turbulence model was applied. The spray distributions and flow fields predicted in the simulation, including the spray penetration, cross-sectional distribution, droplet size in the far-field, the bow shock, large-scale vortices and recirculation zones were in good agreement with experiments. The evolution of spray in supersonic crossflow could be divided into three stages, aerodynamic induced liquid column fracturing, expansion wave promoted spray accelerating, and compression wave disturbed spray blending. Overall, the channel configuration (with expended section) changed the direction and strength of the supersonic airflow. The force of the crossflow determined the spray diffusion. • The diffusion process and spatial distribution of LJISC in a combustor with an expanded section were investigated. • The simulation results such as spray structures, airflow structures were comprehensively verified experimentally. • Typical structures of LJISC in a rectangular channel with an expaned section could be divided into three regions. [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]