The gas–liquid mixing process of a liquid jet in supersonic crossflow with a gas–liquid momentum ratio of 7.67 in the near-wall region is investigated numerically. The two-phase flow large eddy simulation is based on the Eulerian–Lagrangian approach and considers the droplet–wall interaction. The results indicate the penetration depth and the lateral extension width, which are in good agreement with the experimental data. The $\Omega$ shape, especially the spray foot structure of spray in the cross-sectional plane, is captured well. The transport process of spray toward the wall and the formation of spray foot are systematically studied. Under the influence of the upper CVP (counter-rotating vortex pair), partial droplets in the center region of the spray are transported to the near-wall region and move toward both sides when encountering the wall CVP. Under the current gas–liquid momentum ratio, droplets collide with the wall mainly in the central region at the bottom, which will produce splashed droplets. Affected by the horseshoe vortex, the instantaneous distribution of droplets on both sides near the wall shows stripes shape. The spray foot structure forms the shape that is narrow on the top and wide on the bottom and is mainly formed by splashed droplets. Some splashed droplets in the low-speed boundary layer constitute the lower half of the spray foot; meanwhile, some splashed droplets enter mainstream and constitute the upper half of the spray foot. Moreover, the spray is mainly distributed in the core region, and the spray concentration is very sparse in the spray foot region.

Impinging-jet injectors, liquid–liquid coaxial swirl injectors, and liquid–liquid pintle injectors are representative liquid–liquid injectors in liquid rocket engines (LRE). For these liquid–liquid injectors, the atomization processes all involve the liquid impingement, including jet–jet, sheet–sheet, and jets/sheet–sheet impingement, respectively. After impingement, a liquid sheet forms and fragments. Based on these similarities, reviewing published literature on the spray characteristics of these three liquid–liquid injectors in LRE is necessary and will facilitate the investigation of spray characteristics of liquid–liquid pintle injectors to meet the progress of variable-thrust LRE. This review covers the following aspects of these injectors: basic spray morphology, liquid sheet characteristics and disintegration mechanisms, and atomization characteristics. For impinging-jet injectors, rim instability and impact wave play crucial roles in spray morphology and disintegration. Jet Weber number is of great importance for liquid sheet breakup length and mean droplet diameter. In the case of liquid–liquid coaxial swirl injectors, the overall spray morphology is similar to that of pressure swirl injectors, but it may feature two separate liquid sheets. The recess length strongly influences spray morphology, spray angle, breakup length, and Sauter mean diameter. Liquid–liquid pintle injectors can be simplified to injection element, in which the spray morphology resembles a cloak-like shape. In a complete pintle injector, the spray forms a conical liquid sheet. Momentum ratio proves to be the most significant parameter for predicting spray angle. Although the review indicates substantial progress has been made in understanding spray characteristics of liquid–liquid injectors, there remain several shortcomings that require further research, particularly for pintle injectors, which can be learned from the other two injectors.

The mixing and combustion characteristics of liquid kerosene spray in a cavity-based supersonic combustor under a Mach 2.52 inflow are experimentally and numerically investigated. The combustion mode transition of kerosene spray includes two aspects: 1. As the injection pressure increases, the combustion mode shows a transition of weak-intensive-transitional(weak to intensive)-weak-blowoff; 2. in the transitional combustion mode, there is a transition from weak to intensive combustion over time. In the weak combustion mode, the combustion chamber pressure increases slightly, and the flame is inside the cavity. In the intensive combustion mode, the pressure rises sharply from the isolation section and the flame is a typical cavity stabilized combustion mode. The weak combustion at low/high injection pressure is due to rich/lean fuel inside the cavity. In the transitional combustion mode, the flame on the downstream wall of the cavity gradually expands and joins with the flame inside the cavity, evolving from weak combustion to intensive combustion. The transition is accomplished in a very short time, on the order of milliseconds. This study deepens the understanding of the combustion mode of scramjet engines and provides a reference for the design of fuel injection schemes for scramjet engines.

The Rocket-Based Combined-Cycle (RBCC) engine integrates the preponderance of ramjet and rocket engines. It can perform excellently at a lower Mach number than a ramjet while consuming less fuel than the rocket. The higher specific impulse under lower flight Mach conditions guarantees the competitiveness and application prospect of an RBCC engine for reusable space transportation and hypersonic cruise vehicles. With a wide range of Mach numbers, the flow choking between primary and secondary streams in the inner flow passage of the engine becomes complicated. The flow choking not only affects the mass flow ratio between the air stream and rocket plume but also determines the thrust performance of the engine. However, the relationship between flow-choking states and thrust performance has not been revealed. This investigation aimed to provide a better design reference for the RBCC trajectory design and a basis for the RBCC engine geometry design so that the thrust performance under different flow choking states was studied. The findings indicate that the engine is more favorable in the supersonic regime status than others. At the premise of the lower total pressure air stream condition, the thrust of the RBCC engine is attributed to the rocket pressurization effect. On the contrary, its thrust is ascribed to the air stream when the air stream total pressure is higher. Besides, the thrust augmentation of the RBCC engine is essentially due to the larger entrainment ratio. There is a rocket mass flow rate range in different mixer diameters, named region B in this paper, in which the air mass flow rate can be boosted with the rocket mass flow rate and the specific impulse ratio is higher than the rocket specific impulse. Significantly, the RBCC design state should be maintained in region B to meet the operational requirements of engines under the precondition of an insufficient flight Mach number.

Flame stabilization characteristics in the supersonic combustor with a circular cross-section strut were experimentally investigated. The Mach number, stagnation pressure, and stagnation temperature of the inflow were 2.52, 1.60 MPa, and 1486 K, respectively. Compared to the wedge-shaped recirculation flow at the base of the rectangular cross-section strut, the conical recirculation flow at the base of the circular cross-section strut possessed a more favorable geometry characterized by a longer length, a larger volume, and a small surface. When fueled by hydrogen, the combustor equipped both with a circular cross-section strut and a rectangular cross-section strut could achieve self-sustaining combustion. In the tests with equivalence ratios of 0.08 and 0.12, the time-averaged flame chemiluminescence intensity in the combustor with a circular cross-section strut was higher than its counterpart in the combustor with a rectangular cross-section strut by 80% and 53%, respectively. Nevertheless, the standard deviations of the flame chemiluminescence intensities showed an opposite trend. Therefore, the reaction zone downstream of the circular cross-section strut was more vigorous and stable. When both kerosene and hydrogen were supplied to the combustor, these two struts failed to stabilize the flame. Although an unsteady hydrogen–kerosene flame was witnessed in the combustor, it extinguished within 2 ms because the recirculation flow at the base of the strut was too small. Shock wave generators were employed to enhance the flame stabilization ability of the strut with a circular cross section. The experimental results suggested that shock wave generators significantly enlarged the recirculation flow and created a region with high temperature and high pressure. The self-sustaining hydrogen–kerosene flame was achieved in the combustor equipped with shock wave generators.