The organization of detonation using hydrocarbon fuel imposes stricter requirements on the initiation condition compared to hydrogen fuel. To ensure complete heat release within a combustion chamber of limited size, a well-designed combustion chamber is essential. To address the demand for rapid energy release from hydrocarbon fuels in high-speed vehicle engines, This study proposes a detonation initiation scheme based on a V-shape induced surface. This approach employs a careful design of the induced surface to utilize shock wave interactions and shock wave combinations, compressing the incoming fuel and facilitating successful initiation and heat release within the combustion chamber. Numerical simulations demonstrate that, while neither the wedge surface nor the passivated wedge surface can achieve detonation, the V-shape inducted surface effectively initiates detonation and completes fuel combustion in the limited space. The main wave system in the flow field exhibits an inverted V-shaped structure, with detonation occurring near the symmetry plane of the combustion chamber. As the wave system propagates toward the sidewalls, the combustion mode transitions to a shock-induced combustion. The wave intensity is stronger in the symmetry plane and weaker near the sidewalls, which helps mitigate the risk of instability caused by shock-boundary layer interactions. Moreover, by adjusting the radius ratio of the V-shape leading edge, the interaction mode of the wave system within the leading-edge region can be modified, enabling control of the flow characteristics of the entire flow field.

Research on detonation has traditionally focused on forward solutions, with limited attention to inverse design methods, which has significantly hindered the development of detonation engines. In this paper, the Method of characteristics for Curved-Detonation (MOCD) is proposed to enable the inverse design of detonation waves. MOCD is based on the Method of Curved-shock Characteristics (MOCC) and integrates higher-order aerodynamic parameters from Curved Detonation Equations (CDE), allowing the calculation of the wedge angle given specific wave angle. The effectiveness of MOCD is validated using both oblique and curved detonation waves with single-step and detailed chemical reactions. Various applications demonstrate the ability to meet the inverse design requirements of detonation engines. For example, inverse design for given wave angles can optimize engine thrust and prevent Mach reflections. Additionally, inverse design schemes tailored to incoming flow conditions, such as varying Mach numbers and equivalence ratios, enhance the feasibility of detonation engines. Applying the method to given post-wave aerodynamic parameters enables more precise engine design, which is crucial for improving propulsion performance and effective thermal protection. In summary, the advantages of MOCD include not only performing a fast solution of the detonation flow field, but also allowing the inverse design of the detonation wave.

Numerical simulations and theoretical models are developed in this paper for the Detonation-Wave/Boundary-Layer Interactions (DWBLIs) under reflections. Transient flow fields demonstrate the highly non-stationarity of the DWBLIs when Mach Reflection (MR) occur, and subsequent analyses show that the subsonic region introduced by the boundary layer exacerbates the instability. Further quantitative analyses show that viscosity has little effect on propulsive performance and the separation wave can be considered as an oblique detonation wave. Influence parameters to DWBLIs such as combustion chamber height, incoming Mach number, equivalence ratio, and inlet channel length are categorized and studied. Besides simulations, theoretical analytical modeling is established for Regular Reflection (RR) and MR of DWBLIs. Multiple formulas for the separation zone length are obtained according to the mass conservation under different transformation type between inviscid and viscid reflections. Comparison with the numerical simulations verifies the validity of the model and it can be further generalized to the curved DWBLIs. The developed model makes the theoretical solution process of DWBLIs possible and provides the key foundation for further analysis and solution.

The performance of inlets is critical to combined cycle engines. Aiming at the overall demand of Xiamen Turbine Ejector Ramjet (XTER), the design concept and design elements of the inward-turning Turbine Based Combined Cycle (TBCC) inlet for XTER are reviewed in detail. On this basis, flow structures and characteristics of the inlet are analyzed. Results show that either design elements or design constraints of the XTER inlet are coupled with each other, and the mass modulation mechanism is the core element, which is, however, difficult to design and makes the mutual restriction among design elements more complicated. Flow structures in the XTER inlet vary significantly with the increase of Mach number from 0 to 6. Nevertheless, the total mass flow rate remaines above 0.75, which meets the mass flow demand. The mass distribution mechanism also helps the flow mass varies smoothly during mode transitions. During the turbine-to-ejector transition at Mach 2.5, the total-pressure recovery coefficients of the ejector and scramjet increase steadily. Once the transition ends, the total-pressure recovery coefficient is close to or above 0.85. As to the ejector-to-scramjet transition at Mach 3-4.5, the mass flow rate of the inlet increases from 0.81 to 0.90. Moreover, the ejector tunnel maintains a high total-pressure recovery coefficient in the first 62.5% process, indicating that this tunnel still performs properly in the first half of mode transition process. Taken together, the XTER inlet it is capable of continuous normal operation in a wide speed range since its aerodynamic characteristics meet the demand of power system, i.e., the mass flow rate varies moderately in the full speed range and the tunnel performance transits smoothly in the mode transition process.

Oblique wedge initiation is the predominant initiation mode for oblique detonation waves. While increasing the angle of the oblique wedge accelerates detonation, it also raises the overdrive degree of the detonation wave and the resistance within the combustion chamber. To simultaneously ensure rapid detonation initiation and combustion performance, we present the S-shaped wall initiation method, making full use of the flexibility of the curved shock wave system. Concave wall compression facilitates detonation initiation, while convex wall expansion mitigates the overdrive degree of the detonation wave. Given that the expansion wave affecting the initiation zone could lead to detonation wave extinguishment, precise adjustment of the turning point for the S-shaped wall initiation method becomes crucial. Thus, employing the method of curved-shock characteristics, we present a calculation approach to determining the initiation position of curved detonation, to establish the turning point of the S-shaped curved wall in a rational manner. Findings reveal that the optimized S-shaped wall initiation scheme leads to a 16.5% increase in average thrust potential gain and an 8.3% enhancement in the average total pressure recovery coefficient.

The selection of an appropriate basic detonation wave flow field is crucial for improving the performance and geometric design of standing detonation vehicles. This paper employs a detailed chemical reaction model and solves the unsteady axisymmetric Euler equation to study the characteristics of the Axisymmetric Inward Turning Curved Detonation Wave (AIT-CDW) flow field and the parameters affecting the stability of the wave system structure of AIT-CDW flow field. The numerical results demonstrate a radial compression effect in the AIT-CDW flow field. This effect causes the detonation wave to have a shorter initiation length than oblique detonation wave flow field and the detonation wave angle to gradually increase with the flow direction post-detonation. The AIT-CDW flow field is confined space, making it prone to normal detonation waves when the detonation wave reflects from the wall. This phenomenon is detrimental to the stability of the wave system structure in the flow field. It has been observed that increasing the center body radius and decreasing the fuel equivalent ratio can effectively reduce the height of the normal detonation wave or even eliminate it. Additionally, a well-designed generatrix shape of the center body can enhance airflow, reduce choked flow, and promote the stability of the wave structure in the flow field.

The hypersonic three-dimensional inward-turning inlet has gradually become the preferred design for air-breathing vehicles due to its advantages of compact structure, high compression efficiency, high flow capture capability, high total pressure recovery coefficient, and ease of integration. Since the introduction of the inward-turning inlet concept, it has attracted extensive attention from scholars and research institutions worldwide. Firstly, focusing on the aerodynamic design and optimization of the inward-turning inlet, the recent advancements in shock wave analytic theory, basic flowfield construction, and inward-turning inlet design are summarized. Subsequently, the research on improving the performance of the inward-turning inlet under off-design conditions is introduced from the perspectives of low Mach number starting and anti-backpressure characteristics. The current application of the inward-turning inlet in TBCC combined propulsion systems is then discussed. Finally, based on the analysis and summary of the current research status at home and abroad, four key future research spots of the three-dimensional inward-turning inlet are pointed out.

The Mach reflection in axisymmetric flows, which is compressed by an internal cone, is investigated by using a combination of numerical simulations and theoretical analyses at a free-stream Mach number of 6. The effects of the leading-edge angle and cone length on the convergence behaviour of incident shocks and the flow downstream of Mach disk are examined. The results show that the leading-edge angle determines the point of von Neumann strength of the convergent incident shock, and also the location of the Mach disk. In addition to the shock convergence, the position of Mach disk is influenced by the downstream flow as well. According to whether the sonic throat of the stream tube downstream of the Mach disk is affected by the expansion waves generated from the trailing edge of the wall or not, two types of flow patterns downstream of the Mach disk can be classified depending on the leading-edge angle and the cone length. When the sonic throat is independent of the expansion waves, the pressure upstream of the sonic throat is balanced by the post-shock pressure of the reflected shock, and the position of Mach disk will not shift even if the cone length changes. On the other hand, when the sonic throat depends on the expansion waves, the stream tube upstream of the sonic throat needs to match the change of the pressure resulting from expansion waves. The position of Mach disk can be theoretically predicted based on the matching relations between the flow downstream of reflected shocks and the flow downstream of Mach disk.