In order to improve the performance of the rocket-based combined-cycle engine, the influence of the rocket exit area under the ejector mode was studied. Through numerical simulation research, the mass flow of captured air was mainly affected by the entrainment performance under the condition of low flight Mach number. The larger the rocket exit area, the better the entrainment performance. However, with the increase of flight Mach number, the kinetic energy of the entrainment air increases, and there is flow choking in the isolator. The mass flow of captured air was mainly limited by the geometric size of the isolator, independent of the rocket exit area. Under subsonic conditions, the smaller the rocket exit area is, the lower the specific impulse of the engine is, and when the exit dimensionless area is 3.15, the rocket plume will expand and impact the wall, which can cause a sudden reduction in performance. Under supersonic conditions, the smaller the rocket exit area, the higher pressure in the combustor, and the better performance of the rocket-based combined cycle engine.

An experimental study is conducted to investigate the sensitive factors affecting ethylene combustion heat release in a scramjet combustion under supersonic flow conditions with of Mach number 2.52, total pressure 1.34 MPa and total temperature 1 650 K. The effects of key configuration parameters such as the isolator length, injection distance, cavity depth, and throat size on the ethylene combustion heat release are systematically analyzed and compared. The results show that the sensitivity of ethylene combustion heat release to configuration parameters is closely related to the combustion mode. When the combustion is in the completely scramjet mode, the combustion heat release is insensitive to changes in configuration parameters, while the opposite is true when the combustion is in the dual-mode scramjet operation (scramjet mode). Later, as the equivalence ratio increases, the combustion gradually transitions to the dual-mode ramjet operation (ramjet mode), and the sensitivity of the heat release to changes in configuration parameters gradually decreases. Overall, the order of the influence of different configuration parameters on combustion heat release from high to low is as follows: throat size > isolator length > injection distance > cavity depth.

A numerical study on the combustion flow field of an ethylene-fueled circular-section scramjet engine under Mach number 6 flight condition is conducted based on the RANS method. A pressure-related flamelet/progress variable model is used in the calculations, and the numerical results obtained are in good agreement with the experiments. Accordingly, the flame structure and combustion characteristics were analyzed for the high and low equivalence ratio cases. Calculations show that the high equivalence ratio case has a high level of heat release, forming a thermal congestion in the flow field, and that the premixed combustion and diffusion combustion together dominate the overall heat release. In the low equivalence ratio case, the center of the flow channel is still supersonic, and the diffusion flame dominates the overall heat release. In both cases, the jet leeward zone and the axisymmetric cavity are important reaction zones, where the former has a high heat release intensity, while the latter plays an important role in enhancing mixing and improving combustion efficiency.

To reduce the drag generated by the recirculation flow at the rocket base in a Rocket-Based Combined Cycle (RBCC) engine operating in the ramjet/scramjet mode, a novel annular rocket RBCC engine based on a central plug cone was proposed. The performance loss mechanism caused by the recirculation flow at the rocket base and the influence of the plug cone configuration on the thrust performance were studied. Results indicated that the recirculation flow at the rocket base extended through the entire combustor, which creates an extensive range of the “low-kinetic-energy zone” at the center and leads to an engine thrust loss. The plug cone serving as a surface structure had a restrictive effect on the internal flow of the engine, making it smoothly transit at the position of the large separation zone. The model RBCC engine could achieve a maximum thrust augmentation of 37.6% with a long plug cone that was twice diameter of the inner isolator. However, a shorter plug cone that was half diameter of the inner isolator proved less effective at reducing the recirculation flow for a supersonic flow and induced an undesirable flow fraction that diminished the thrust performance. Furthermore, the effectiveness of the plug cone increased with the flight Mach number, indicating that it could further broaden the operating speed range of the scramjet mode.

It has been found that the static pressure distribution along the axial direction of liquid kerosene is lower than that of the gaseous kerosene under the same flow condition and overall equivalent ratio from previous studies. To further investigate this phenomenon, a compressible two-phase parallel simulation method is utilized to analyze the mixing and combustion characteristics of gaseous and liquid kerosene jets in a cavity-based supersonic combustor. The numerical results are consistent with the experiments and demonstrate that gaseous injection leads to a cavity shear layer that dives deeper into the cavity, forming two recirculation zones in the front and rear of the cavity. In contrast, the cavity shear layer is closer to the mainstream during liquid injection, and only a large recirculation zone is formed in the rear of the cavity. As a result of the cavity shear layer and the recirculating flow, the fuel vapor of gaseous injection accumulates in the front of the cavity, while for the liquid injection, the fuel vapor disperses in the cavity, cavity shear layer, and the region above, and the rear of the cavity has a higher fuel vapor concentration than the front. This unique fuel distribution causes the combustion area to be concentrated in the cavity during the gaseous injection but dispersed inside and downstream of the cavity during the liquid injection. As a result, forming a thermal throat under the same conditions is more challenging during liquid injection, and the generated static pressure distribution is lower than that during the gaseous injection.

Developing supersonic combustion models with efficiency, accuracy and practicality is important foundation to deeply understand the complex combustion processes in scramjet engines. Characterized by efficiency and intuition, the flamelet-like models are widely used models in computational combustion methods. However, the supersonic combustion flow field has the nature of strong compressibility, multiple modality, and multiple scales, which poses a great challenge to the traditional flamelet-like models with fixed boundary conditions, and then the complex chemical reaction mechanisms that may face will impose additional computational burden. In this context, this paper reviews the flamelet-like models used in scramjet engines, and summarizes prominent issues in the application practice, including modeling partially premixed combustion, defining progress variable, solving temperature efficiently, evaluating assumed Probability Density Function (PDF) models, and treating mixture fraction variance. Furthermore, possible prospects and directions of improvements are proposed and highlighted for the flamelet-like models. To fully describe the physicochemical scenario and address the raised challenges, these improvements are dedicated to dealing with the compressibility, temperature rise, time-scales, species of interest, multi-inlet combustion, the progress variable definition, and the higher Mach number flight condition.

In this study, a hybrid Reynolds-Averaged Navier-Stokes/Large Eddy Simulation (RANS/LES) approach coupled with an improved Flamelet/Progress Variable model is used to numerically investigate the combustion flow field of an ethylene-fueled circular-section model scramjet at high Mach numbers. The computational results show highly consistent combustion structure and heat release characteristics with those from the experiments. On this basis, the analysis results indicate that the mixing process of the flow field is dominated by the large-scale flow structures, and the shock train at the center of the channel is the direct cause of promoting vorticity generation, fuel mixing, and downstream reaction zone wrinkling. In this case, the cavity does not directly participate in the heat release process, but rather plays a role in inducing shock waves and promoting mixing and combustion. Overall, combustion exhibits spatiotemporal multi-scale characteristics, and flames gradually develop towards a near equilibrium state as they propagate downstream. The scramjet mode and diffusion combustion dominate the entire reaction zone, and most of the fast-chemistry combustion is in the corrugated flamelets regime.

The asymmetric separation has a crucial effect on the performance of the scramjet. In this study, the asymmetric separation and combustion in both rectangular and circular scramjets are investigated numerically, and the effect of injection scheme is analyzed. The characteristics of the flow field are analyzed based on sufficient code verification. In the rectangular scramjet, the separation tends to occur in the corner due to the corner boundary-layer effect. The separation is asymmetric and only two corners have serious separation. The fuel penetration depth in the separation zone increases and the combustion is intensified. When the injection scheme is uniform, both the combustion and separation become weak. In the circular scramjet, the separation and combustion are basically axisymmetric in the scramjet with one-row injection scheme. The asymmetric combustion becomes obvious in cases with multi-row injection scheme. When the injection orifices distribute intensively on the top and bottom sides, the strongest and weakest separations occur near these two sides respectively. When the distribution of orifices becomes uniform, the direction of separation cannot be predicted. For multi-row cases, the nonuniform injection scheme could result in violent combustion and asymmetric flow structures compared with the uniform injection scheme.