The Rocket-Based Combined Cycle (RBCC) engine is a promising propulsion system for hypersonic and space launch applications due to its capability to operate efficiently over a broad range of flight conditions. This study investigates the influence of total temperature and total pressure on flow patterns in the rocket-ejector mode of a RBCC engine using two-dimensional numerical simulations—a simplification that facilitates efficient parametric analysis while inherently omitting three-dimensional effects. The transition between stable and wavy flow patterns under the Diffusion and Afterburning (DAB) combustion mode is analyzed. Higher total temperatures enhance mixing efficiency but can induce wavy flow patterns, leading to potential instability. Conversely, increased total pressures promote stability through Fabri-choking mechanisms while reducing mixing efficiency by limiting entrainment capacity. A significant hysteresis effect is observed, where transition thresholds for stable and wavy states vary based on operational history. Key mechanisms contributing to this effect are discussed in depth, including momentum flux dynamics, Fabri-choking behavior, shock wave reformation, and mass and heat exchange processes. These findings provide critical insights for optimizing RBCC engine performance by balancing flow stability and mixing efficiency under varying conditions. This study’s insights into flow pattern dynamics, particularly the hysteresis effect, are crucial for developing robust control strategies and optimizing RBCC engine designs for hypersonic and space launch applications.

The effects of pilot hot gas injection on flame stabilization in a cavity-based scramjet were investigated using three-dimensional unsteady Reynolds-averaged Navier-Stokes simulations. Ethylene fuel was injected within a dual-solution regime exhibiting two possible flame stabilization modes. Pilot injection assisted both ignition and flame stabilization. Two representative pilot heating levels, low and high enough, were examined to assess their impact on flame behavior and mode transition. Results show that flame stabilization was highly sensitive to pilot heating power: low heating sustained cavity shear-layer stabilization, whereas high enough heating induced a transition to jet-wake stabilization. This transition was primarily driven by reduced ignition delay in the unburned jet-wake region due to elevated temperatures from pilot heating. Additionally, pilot injection enhanced fuel-air mixing through vortex generation. Both effects intensified with increasing pilot heating power. Consequently, high enough heating power facilitated upstream flame propagation and flow separation, ultimately triggering the transition. A theoretical analysis based on the Semenov thermal ignition theory further showed that high enough pilot heating promoted chemical heat release within a control volume in the unburned jet-wake region, which exceeded the enthalpy outflow in the initial cavity shear-layer stabilization mode. This energy imbalance led to a sustained temperature rise, initiating the transition.

The design of scramjets for flight at Mach numbers above 9 is constrained by severe aerothermodynamic effects and chemical nonequilibrium. This work utilizes a validated quasi-one-dimensional model with finite-rate chemistry to conduct a systematic parametric investigation of scramjet performance. The study analyzes the influence of inlet contraction ratio, combustor geometry, and flight conditions (Mach number and dynamic pressure) on specific impulse. Results reveal that performance is governed by a critical tradeoff between maximizing combustion completeness, which is favored by high pressure, and mitigating high-temperature energy losses from nitric oxide (NO) formation and product dissociation. An optimal “Max Isp island” is identified in the combustor geometric design space, the location of which is highly sensitive to the operating point. Increasing flight Mach number forces the optimal geometry toward more aggressive expansion to manage thermal loads, while increasing dynamic pressure shifts it back toward less expansion to capitalize on the suppression of fuel dissociation. These findings demonstrate that an optimal scramjet configuration is not fixed but is a dynamic function of the flight conditions, highlighting the necessity of multipoint optimization for designing hypersonic vehicle propulsion systems.

The structural integrity of thermal protection systems under cyclic supersonic aerothermal loads remains a critical bottleneck for high-performance aerospace vehicles. Here, we elucidate the dynamic surface roughness evolution and its role in accelerating failure mechanisms within 2D-woven oxide/oxide ceramic matrix composites (CMCs). By integrating wind tunnel thermal shock experiments with computational fluid dynamics (CFD) simulations, we establish a probabilistic diagnostic framework based on the Gaussian fitting of roughness height deviations. Our findings identify an apparent transition damage regime around 1373 K; beyond this limit, fiber fracture triggers nonlinear roughness proliferation, which subsequently intensifies intra-cavity vortices. This aerodynamic coupling elevates local wall shear stresses by an order of magnitude compared to pristine surfaces, establishing a self-reinforcing “topography-vortex-damage” feedback loop. The identified asymmetric damage modes and their governing laws provide fundamental physical criteria for the life-cycle prediction and holistic design of long-service thermal protection architectures.

This paper presents a special design concept for a fixed-geometry combustor, which is adaptable to a wide range of requirements for future scramjets. The design is based on the principles of staged combustion and the integration of a geometric throat. In low-speed inflow conditions, the combustor optimizes subsonic combustion performance through the choking effect created by the geometric throat. In contrast, under high-speed inflow conditions, the aerodynamic properties of the open cavity establish a boundary downstream of the combustor. This design mitigates the negative impacts of the geometric throat in supersonic flow, resulting in improved combustion performance under high-speed inflow conditions. This paper examines how the combustor flow channel expansion ratio affects the combustion flow field and performance parameters at flight Mach numbers from 3 to 7, with a kerosene equivalence ratio of 1.0. This paper verifies the feasibility of the new design concept with a specific adaptive throat combustor. To illustrate the advantages of this combustor, a traditional thermal throat combustor is present for comparative analysis of combustion performance and internal flow field parameters at typical flight Mach numbers. The results show that under inflow conditions corresponding to Mach numbers 3 to 5, the combustion performance of the adaptive throat combustor significantly exceeded that of the thermal throat combustor. Under inflow conditions corresponding to Mach numbers 6 to 7, the adaptive throat combustor's performance was slightly lower than that of the thermal throat combustor. Therefore, this design concept holds considerable promise for practical engineering applications. (c) 2025 The Author(s). Published by Elsevier Ltd on behalf of Chinese Society of Aeronautics and Astronautics. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/ licenses/by-nc-nd/4.0/).

This paper focuses on the numerical study on the phenomenon and mechanism of spin-to-orbital angular momentum conversion for the head-on collision between a spinning droplet and a non-spinning droplet of unequal sizes. The droplet deformation process was phenomenologically described, where the gas film suppresses the rotating flow inside the spinning droplet and, in turn, rotates the non-spinning droplet through its shearing effects. The gas film tends to be curved by the size disparity, leading to a more complex gas film flow. The hysteresis mechanism of spin-to-orbital angular momentum conversion was analyzed. The decrease in spin angular momentum and increase in orbital angular momentum have an upper limit as increasing the size ratio, which is attributed to the droplet deformation that is similar to a droplet impacting on a liquid film. In addition, a wider range of size ratio, Weber number, and droplet spin angular speed were discussed, where some numerical findings are significant to the angular momentum conversion coefficient modeling.

Normal shock trains are a flow phenomenon of significance to ramjet engines, but it remains unclear what its structure is decided by and how it evolves with the incoming Mach number. To seek a theoretical explanation, the minimum entropy production principle is generalized to the quasi-steady behavior of normal shock trains in two-dimensional straight channels with uniform incoming flow. Numerical simulations are also performed to validate the model together with the data collected from public literature. The analysis suggests that the flow parameters of a normal shock train depend on the inviscid shock-shock interactions rather than the local boundary-layer separations, though the angles of two incident shocks should still be equal as similar to the case that complies with the free-interaction theory. The shock feet’s positions, meanwhile, are allowed to be coincident or not, free from the entropy restriction. This freedom of position explains why both symmetric and partially asymmetric normal shock trains could be found previously. Further theoretical calculations reveal the inclinations of two incident shocks increase first and then decrease with the incoming Mach number, peaking at 48.570 degrees when the Mach number reaches 1.753. It is also indicated that the Mach number range allowing for a normal shock train is 1.652 to 2.254, giving evidence for past observations.

Wind tunnel experiments were conducted at incoming Mach numbers ([Formula: see text]) of 2.01 and 2.81 to investigate the interactions between shock waves and weak, moderate, and strong streamwise vortices in a concave channel representing the inward-turning inlet/isolator flow of hypersonic engines. Synchronized high-speed schlieren visualizations and high-frequency wall static pressure measurements were collected to evaluate the steady-state and dynamic characteristics of the global and local flowfields. The interference of streamwise vortices bifurcated the oblique shock wave; the stronger the former, the more severe the deformation of the latter with increasing intensity during downstream development. When [Formula: see text], moderate and strong streamwise vortices caused the Mach stem in the shock train leading edge (STLE) to bend upstream; when [Formula: see text], moderate and strong streamwise vortices caused the top- or bottom-wall shock waves in the STLE to bend and bifurcate into multiple shock waves. An increase in streamwise vortex intensity weakened the self-excited oscillations of the shock train and enhanced the pressure fluctuations caused by STLE oscillations. The dominant frequency of the interacting flowfield oscillation was 20–40 Hz, and the dominant wave structure in the shock train oscillated rigidly. These results can inform hypersonic engine inlet development.

The restart ability is a significant concern in the design of hypersonic inlets, but existing theoretical and empirical formulas can hardly accurately predict the restart characteristics. To develop a universal rapid prediction method, this paper explores the similarity law of the restart phenomenon at variable Mach numbers based on systematic numerical simulation. The corrected contraction angle and internal contraction ratio (ICR) that reflect the influence of multiple geometric and aerodynamic parameters are proposed, and an empirical formula for restart boundaries is established and validated by the literature data. It is revealed that the restart boundaries obey an inherent geometric similarity law, and the discrete relation between Mach number and ICR, emphasized by most existing studies, is not essential. The restart-critical wall pressure characteristics are analyzed, and a self-similar pressure criterion is obtained with the incident shock strength considered.

Shock-induced combustion in a direct-connected combustor was experimentally and numerically investigated at Mach 10 flight conditions. This study aimed to reveal the establishment of shock-induced combustion and its transition to partially premixed-dominated combustion in practical combustor flows, highlighting the critical effects of fuel injection and shock wave/boundary-layer interaction on shock-induced combustion stability. The combustor consisted of a constant-area section, a diverging section, and two shock-generating wedges symmetrically mounted on the top and bottom walls of the combustor. Hydrogen fuel was injected through two symmetrical fuel injectors. Each fuel injector had three identical struts equally spaced along the spanwise direction. Oblique shock waves were generated through wedges at the end of the combustor diverging section. Wall pressure and hydroxyl radical (OH*) flame chemiluminescence imaging were used to characterize the transient phenomenon of flame stabilization and upstream propagation. Large-eddy simulations were conducted to analyze the unsteady flame propagating processes in association with experimental results. The nonuniformity of prewave parameters induced by strut injection, coupled with the exacerbation of boundary-layer separation caused by shock-induced combustion, led to combustion instability and engine performance deterioration.

To understand the reflection type of incident shock wave under the influence of the wave structures near the intersection point of shock/shock interaction in confined-space double-wedge. Shock polar analysis is conducted to reveal the displacement of critical conditions due to the interaction of wave structures near the intersection point of Type VI and I. The evolution of the flow field and shock reflection is comprehensively analyzed for various relative positions of the incident shock wave and expansion waves, which controlled by contraction ratio (Cr) of the double-wedge. Computational results show that expansion wave of Type VI has limited impact on the reflection type of oblique shock wave above detachment condition. However, its influence on the shock reflection configuration becomes appreciable within the displacement band of the detachment condition. When the expansion waves cover the incident shock wave, The critical Cr for the RR-MR transition on the opposite wall is determined when expansion waves fully cover the incident shock, and a subsonic region following RR is observed. Additional insight into the trend of Mach stem height along increasing Cr shows linear or nonlinear mode, which is dominated by the related position of expansion fans. It is also shown that the shock wave near the intersection point of Type I contributes subtle effect on the reflection type at the opposite wall.

To better understand combustion unsteadiness in scramjets, a cavity-based circular combustor is experimentally studied at five total pressures, namely 0.76, 1.52, 2.54, 4.56, and 5.06 MPa, with the total temperature kept at 1578 K and the incoming Mach number 2.6 in all cases. The time-averaged pressure curves show that the cavity pressure, overall, rises with the incoming total pressure, and consequently, the shock train is pushed more upstream. But one more pressure peak appears after the total pressure reaches 5.06 MPa, reflecting a shift in the flow pattern. Severer pressure oscillations are also observed as the total pressure goes up, particularly for the part within the isolator, which signifies the behavior of the shock train and tends to increase faster in terms of the oscillatory amplitude. By contrast, the cavity-pressure oscillations weaken gradually till basically unchanged. Though different in amplitude trend, pressure signals in those two places always share a dominating frequency, which decreases constantly. To predict it more precisely, the classic frequency formula is improved based on one-dimensional analysis and experimental data, and the error is thus substantially reduced to 10%.

The prediction of hypersonic inlet unstart has been a difficult aerodynamic problem over the past half century. Different from inviscid unstart theories, this study develops empirical and theoretical prediction methods for viscous flow. By proposing a corrected contraction angle and contraction ratio, a generalized similarity law for unstart boundary is established and validated by sufficient numerical and experimental results, describing the effects of aerodynamic and geometric parameters simultaneously. To explain the physical implications of the generalized unstart similarity law, two-dimensional flow-based theoretical models are established for three unstart modes using oblique shock relations, free interaction theory, and scaling laws for separation length. For the short-cowl and long-cowl modes, the self-similarity of the unstart boundary is attributed to the separation scale relative to the effective throat height and the mass-averaged throat Mach number accompanied by shock reflections, respectively. The transitional unstart is triggered by the decrease of the shock-impingement distance relative to the separation scale, and the self-similarity is broken by the wedge length independence and Reynolds number dependence. Moreover, to facilitate unstart detection in wind tunnel or flight tests, the corrected dimensionless pressure rise is proposed, correlating the critical wall pressure characteristics with the incident shock pressure rise and dimensionless wedge length.

This study employs three-dimensional unsteady numerical simulations to analyze fluid-combustion interactions in a scramjet combustor under freestream oxygen perturbations at Mach 6 flight conditions. The combustor utilizes ethylene fuel injection, with systematic examinations of dynamic responses to abrupt reductions in freestream oxygen mole fractions from baseline 0.21 to 0, 0.05, and 0.10. A characteristic mode shift occurs from jet-wake stabilization to cavity shear-layer stabilization during oxygen depletion, followed by reversion to the original mode upon freestream recovery. This transition correlates strongly with perturbations in combustion intensity and heat release redistribution. Rapid alterations in shock train positioning and flow separation characteristics were observed, directly influencing static pressure distributions along the flowpath. The resultant pressure fluctuations exhibited temporal coherence with thrust variations. Thrust modulation magnitudes demonstrated linear dependence on oxygen depletion severity, while combustion sustainability remained intact during 0.2 ms perturbations. Prolonged perturbations can induce cumulative effects and are capable of altering hypersonic vehicle trajectories through thrust vectoring changes. These results highlight the critical time-dependent coupling between freestream composition transients and scramjet operability, providing quantitative benchmarks for robust combustor design in atmospheric disturbance scenarios.

The influence of thermal/chemical nonequilibrium on a hydrogen-fueled [Formula: see text] 10 scramjet was analyzed by combining the JF-2 4 s hock tunnel test and improved delayed detached eddy simulation modeling. A remarkable change in flame stabilization mode when incorporating the two-temperature nonequilibrium model was observed. The nonequilibrium heating and cooling effects were analyzed for different sections of the scramjet. The nonequilibrium heating effect facilitates the upstream flame propagation by inducing an early ignition and thickening the inlet boundary layer. The nonequilibrium heating or cooling effect is generally weak in the nearly constant-area isolator and combustor, where the flow is mainly influenced by the chemical nonequilibrium with a variation of 10% in reaction rate. The nonequilibrium cooling effect mainly exists in the expanding nozzle, where [Formula: see text], but the energy replenishment from vibrational mode to trans-rotational mode is nearly frozen when [Formula: see text]. Under both conditions, the final mixing is nearly complete, and net thrust has been achieved. When considering the nonequilibrium effects, the final combustion efficiency increases from 87.68 to 99%, together with a 53.94% rise in peak pressure ratio and a 137.04% rise in specific impulse.

This article is based on the supersonic directly connected wind tunnel. Through a specially designed experimental chamber, combined with infrared temperature measurement, high-speed camera, etc., in -situ monitoring of composite materials under airflow at Ma 3.0 with a total temperature of 950 similar to 1473K was carried out. The dimensional analysis method was used to propose dimensionless parameters to characterize the thermal coupling caused by high-speed airflow thermal shock. Research has shown that the thermal coupling effect of supersonic airflow causes uneven temperature inside the material, and the thermal stress caused by temperature gradient changes (including increasing and decreasing processes) is the main reason for material damage. The damage of ceramic matrix composites under thermal shock mainly manifests as a decrease in surface roughness, surface fiber fracture and a decrease in elastic modulus. In addition, the study also found that there are damage thresholds for the thermal shock effect of airflow at different total temperatures, which helps to further understand the thermomechanical damage mechanism and degradation law of composite structure under high -temperature transient conditions.

Suppressing shock-induced flow separation has been a long-standing problem in the design of supersonic vehicles. To reduce the structural and design complexity of control devices, a passive control technique based on micro-serrations is proposed and its controlling effects are preliminarily investigated under test conditions in which the Mach number is 2.5 and the ramp creating an incident shock is 15 deg. Meanwhile, a vorticity-based criterion for assessing separation scales is developed to resolve the inapplicability of the zero skin friction criterion caused by wall unevenness. The simulations demonstrate that the height of the first stair significantly influences the separation length. Generally, the separation length is shorter at higher stairs, but when the height is greater than half of the thickness of the incoming boundary layer, the corresponding separation point moves upstream. A stair with a height of only 0.4 times the thickness of the boundary layer reduces the separation length by 2.69%. Further parametric analysis reveals that while the remaining serrations have limited effects on the flow separation, an optimization of their shape (depth and width) can create more favorable spanwise vortices and offer a modest improvement of the overall controlling performance. Compared to the plate case, a 9.13% reduction in the separation length can be achieved using a slightly serrated design in which the leading stair is 0.1 high and the subsequent serrations are 0.2 deep and 0.05 wide (nondimensionalized, with the thickness of the incoming boundary layer). Meanwhile, the micro-serration structure even brings less drag. Considering the minor modification to the structure, the proposed method has the potential for use in conjunction with other techniques to exert enhanced control on separations.

Inlet unstart is detrimental to hypersonic airbreathing flights. By unstart, we mean the separation-induced collapse of a supersonic duct flow. The unstart characteristics of a two-dimensional contraction duct with an expansion corner are numerically investigated to develop a universal prediction method and to determine the unstart mechanism. The unstart boundaries under varying contraction angles, entrance heights, and Reynolds numbers are obtained by translating the cowl plate. A self-similar empirical equation of unstart boundary that unifies multiple variables is proposed through similarity analysis. To estimate the unstart boundaries of large-angle ducts, the alternative area ratio is presented. Three types of unstarts are identified according to the distribution of the unstart boundaries and the evolution mechanism of separation region: the long-cowl, transitional, and short-cowl unstart. The long-cowl unstart stems from the mass flow limit at the throat, while the other two unstarts stem from the separation region growth. The wall pressure criteria for unstart are correlated with the dimensionless cowl length and the characteristic separation scale in the critical state to facilitate unstart detection. Moreover, the self-similar compression intensity is proposed to characterize the maximum pressure increase induced by the unit effective area contraction in the started state.