RBCC Research Articles

Mixing enhancement assisted by dual plate cavity in supersonic flow

Fast and efficient mixing of high-speed shear flow formed by fuel and oxidizer is of great importance for the improvement of rocket-based combined cycle engine performance. Nevertheless, the existence of compressibility effects of high-speed flow significantly inhibits the growth process of a mixing layer. Moreover, a finite-length combustor of the engine calls for more effective enhanced-mixing strategies to complete mixing in a shorter streamwise distance. To this end, in present paper, the strategy called dual plate cavity (DPC) is proposed to promote mixing. Three cases including the benchmark, front-DPC and back-DPC cases are selected to perform the comparative study. By means of high-order direct numerical simulations, the structure evolution characteristics and turbulence intensity distributions are researched. The index of velocity thickness is utilized to assess the mixing layer growth. The results indicate that with the introduction of DPC, the mixing process is dramatically promoted. The penetration behavior of newly found T-shaped structures into the upper main stream can engulf more fluid into the mixing region. Specifically, in the back-DPC case, the coexistence of both large-scale and small-scale structures in the far flow field can improve the turbulence intensity. The spatial correlation analysis results show that with the influence of DPC, the structure sizes are much larger than that of the benchmark case in the same streamwise position. Meanwhile, the contour line equal to 0.5 possesses property of distortion for the back-DPC case. The drastic pulsation of a mixing layer edge can obviously promote the mixing process. Through exploration of the enhanced-mixing mechanisms, this work indicates that the proposed DPC strategy is a good candidate for efficient mixing, and in the future, more detailed work including three-dimensional simulations concerning the strategy optimization is suggested to be performed.

Computational analysis of the scramjet mode of the RBCC inlet using micro vortex generators

Abstract A computational study of a three ramped dual duct Rocket Based Combined Cycle (RBCC) engine inlet at scramjet mode using different types of MVG (micro vortex generator) arrays were conducted. The definite geometry of engine inlet was operated at hypersonic speeds of Mach 5 and 7 to study the effect of the arrays of delta ramp (DR), rectangular vane (RRV) and ramp vane (RV) on the pressure recovery, exit Mach number, the mass flow rate and Shock wave Boundary Layer interaction (SWBLI). The study was performed considering same heights for all the configurations of MVG array and were positioned at the point of shock impingement on the ramp which caused the separation bubble. The computational analysis was done using k-omega model in Fluent Workbench of Ansys.

Experimental investigation for temperature and emissivity by flame emission spectrum in a cavity of rocket based combined cycle combustor chamber

Flame temperature and spectral emissivity were the important parameters characterizing the sufficient degree of fuel combustion and the particle radiative characteristics in the Rocket Based Combined Cycle (RBCC) combustor. To investigate the combustion characteristics of the complex supersonic flame in the RBCC combustor, a new radiation thermometry combined with Levenberg-Marquardt (LM) algorithm and the least squares method was proposed to measure the temperature, emissivity and spectral radiative properties based on the flame emission spectrum. In-situ measurements of the flame temperature, emissivity and spectral radiative properties were carried out in the RBCC direct-connected test bench with laser-induced plasma combustion enhancement (LIPCE) and without LIPCE. The flame average temperatures at fuel global equivalence ratio (α) of 1.0b and 0.6 with LIPCE were 4.51% and 2.08% higher than those without LIPCE. The flame combustion oscillation of kerosene tended to be stable in the recirculation zone of cavity with the thermal and chemical effects of laser induced plasma. The differences of flame temperature at α = 1.0b and 0.6 were 503 K and 523 K with LIPCE, which were 20.07% and 42.64% lower than those without LIPCE. The flame emissivity with methane assisted ignition was 80.46% lower than that without methane assisted ignition, due to the carbon-hydrogen ratio of kerosene was higher than that of methane. The spectral emissivities at 600 nm with LIPCE were 1.25%, 22.2%, and 4.22% lower than those without LIPCE at α = 1.0a (with methane assisted ignition), 1.0b (without methane assisted ignition) and 0.6. The effect of concentration in the emissivity was removed by normalization to analyze the flame radiative properties in the RBCC combustor chamber. The maximum differences of flame normalized emissivity were 50.91% without LIPCE and 27.53% with LIPCE. The flame radiative properties were stabilized under the thermal and chemical effects of laser induced plasma at α = 0.6.

Numerical study on throttling characteristics of RBCC engines in ejector mode

Investigating the throttling characteristics of Rocket-Based Combined Cycle (RBCC) engines in ejector mode is crucial for adapting to varied flight mission demands. This study employs validated RANS numerical methods to investigate throttling effects on an experimental kerosene-fueled RBCC engine under two representative flight conditions: ground take-off (Mach 0) and supersonic flight (Mach 2.0). Results indicate that, at Mach 0, throttling reduces the entrained secondary stream and increases mixing efficiency, but exacerbates mixing losses, which cannot be offset by the thrust produced by the secondary stream due to its weak working capacity. Therefore, specific impulse performance deteriorates during throttling. Moreover, the optimum mixture ratio remains at 2.4, consistent with conventional rockets. Conversely, at Mach 2.0, throttling mitigates the primary stream’s squeezing effect to augment the entrained secondary stream. Simultaneously, throttling enhances mixing efficiency to improve the utilization of residual fuel. Consequently, the optimum mixture ratio shifts from 2.4 to below 1.4, which contributes to harnessing the work capacity of the secondary stream generated by ram effects, and significantly improves specific impulse performance. This study emphasizes the distinct throttling characteristics at different Mach numbers and provides pivotal insights for developing more efficient RBCC engines in ejector mode.

Numerical Investigation on the Effect of Fuel-Rich Degree in the RBCC Engine under the Ejector Mode

The ejector mode of the Rocket-Based Combined-Cycle (RBCC) engine is characterized by high fuel consumption. This study is aimed at investigating the influence of the rocket fuel-rich degree on the RBCC engine’s performance under the ejector mode combined with simultaneous mixing and combustion (SMC). Numerical simulations were conducted for various rocket mixing ratios (Φ=1.6~3.2) under subsonic (Maf=0.9) and supersonic (Maf=1.8) flight conditions. It was observed that a high fuel-rich degree in the rocket plume negatively impacts the eject performance under all conditions. However, it improves the overall performance (Isp) at high flight Mach numbers (Maf). For supersonic conditions, increasing the fuel-rich degree promotes greater fuel participation in combustion, thereby enhancing RBCC engine performance. Nevertheless, the subsonic-supersonic mixing layer exhibits low evolution, resulting in a decrease in reaction efficiency from 29.2% to 12.0% as the Φ decreases from 3.2 to 1.6. Consequently, there is an inefficient utilization of fuel. To optimize RBCC engine performance, the rocket fuel-rich degree can be appropriately increased. However, this increase should be limited to prevent fuel wastage arising from low reaction efficiency. Under subsonic conditions (Maf=0.9), the low kinetic energy of captured air leads to the occurrence of “negative thrust surface” and “wall impact” phenomena, which hinder the efficient and stable operation of the RBCC engine. Consequently, adjusting the fuel-rich degree alone cannot promote specific impulse (Isp), and a low fuel-rich degree is considered an ideal strategy when combined with adjustable nozzle technology.

Analysis of the combustion modes in a rocket-based combined cycle combustor using unsupervised machine learning methodology

Combustion mode analysis is essential to a rocket-based combined cycle (RBCC) combustor because it may experience multiple combustion modes during the operation. In this study, a method based on an autoencoder and a K-means algorithm was proposed for combustion mode analysis. Flame chemiluminescence images and schlieren images of three combustion modes observed in an RBCC combustor were used to evaluate this method. Two autoencoders that followed the same encoder–decoder architecture were developed separately to generate the latent space representations of flame chemiluminescence images and schlieren images. In the latent space, the centroids and boundaries of different combustion modes were determined using the K-means algorithm. Each autoencoder was trained using 750 images and tested using another 3000 images. The method achieved an accuracy up to 99% on both flame chemiluminescence images and schlieren images. The images generated by the decoder suggested that the autoencoder captured the important features (e.g., primary reaction zone and shock wave) of the reacting flow field. The autoencoder developed for flame chemiluminescence images also successfully detected the combustion mode transition during an ignition process, which suggested that it had the potential to monitor the combustion mode in a real time manner. However, the autoencoder failed on monitoring combustion mode transition when it came to the schlieren images because the optical access of the training data was not exactly the same. Therefore, it is essential to ensure that the optical accesses of different combustion modes are exactly the same when schlieren images are employed for combustion mode analysis.

Effects of excess oxidizer coefficient on RBCC engine performance in ejector mode: A theoretical investigation

This study investigates the effects of the primary rocket’s excess oxidizer coefficient (EOC) on the performance of a kerosene-fueled rocket-based combined cycle (RBCC) engine operating in ejector mode using the diffuser and afterburning (DAB) cycle. Employing complete mixing and chemical equilibrium assumptions, a quasi-one-dimensional model is developed to simulate combustion processes in the primary rocket chamber, mixer, and afterburner. Experimental data and full three-dimensional simulation validate the model’s accuracy. Results indicate that increasing EOC mitigates thermal choking effects at the mixer exit, decreasing mixing and Rayleigh losses, enhancing the entrainment ratio, and providing more oxygen for secondary combustion. However, the gas mixture’s compression ratio and total temperature decrease, rendering secondary combustion more susceptible to thermal choking in a constant cross-sectional afterburner. As the flight Mach number increases from 0.8 to 2.5, the optimum EOC for maximum specific impulse shifts from 0.725 to 2.375. While the optimum EOC for maximum thrust remains consistently around 2.0. The findings underscore the necessity of a higher EOC to fully unlock the DAB cycle’s thrust augmentation potential.

Experimental and numerical study on RBCC engines performance in simultaneous and combustion cycle

The simultaneous mixing and combustion (SMC) cycle is a simple yet effective approach to augment the specific impulse of rocket-based combined cycle (RBCC) engines in ejector mode. However, the thrust performance of RBCC engines in the SMC cycle has not been adequately explored. This study investigates the effects of the flight Mach number, mixer shapes, and mixture ratio (MR) on the thrust performance of an axisymmetric kerosene-fueled RBCC engine. Ground free-suction experiments were conducted, employing a supersonic air ejector to simulate varying flight backpressures. Wall pressure distributions and axial thrusts were obtained. To eliminate undesirable interactions between MRs and backpressures inherent in experiments, numerical simulations were performed. The experimental results show that as flight Mach number increases, specific impulse improves rapidly, and specific impulse augmentation is achieved in the transonic regime. Simulations conducted at H4.0km/Mach0.88 with the expansion mixer reveal that as MR decreases, both mixing efficiency and combustion intensity increase. The balance between specific impulse gain from the secondary stream and loss from the primary stream is achieved at MR = 2.2, with the highest specific impulse being achieved at MR = 2.0. These insights provide valuable guidance for advancing the development of more efficient RBCC engines using the SMC cycle.

A simplified chemical model for RBCC engines operating in ejector mode

The ejector mode in rocket-based combined-cycle (RBCC) engines involves complex mixing and combustion processes between the primary and secondary streams and second fuel, which poses a challenge to numerical simulation. This paper introduces an efficient model to predict RBCC engine performance in ejector mode. Given the slow growth rate of compressible mixing layers with strong compressibility effects and high chemical reactivity of immediate species within the primary stream, the combustion intensity is primarily governed by the mixing process. The model converts residual chemical energy within the primary stream into suitable mass flow rates of H2 and CO and simplifies the intricate chemical kinetics into single-step models. The model was tested on hydrogen- and kerosene-fueled RBCC engine experiments involving diffusion and afterburning (DAB) and simultaneous mixing and combustion cycles. Simulation results align well with experimental data regarding wall pressure distributions, entrainment ratio, and axial thrust. The model suggests intense combustion occurs in both the mixer and afterburner, even in the DAB cycle, which notably reduces the entrainment ratio. Additionally, early combustion within the mixing process incurs extra Rayleigh loss and impairs pressurization performance. The secondary fuel exhibits a jet-wake stabilized combustion mode under a high-temperature and low Mach number environment. In conclusion, the proposed model provides an efficient way of predicting RBCC engine performance in the ejector mode.

Rapid ascent trajectory optimization of rocket-based combined cycle hypersonic vehicle

The ascent trajectory design of a rocket-based combined cycle (RBCC) hypersonic vehicle has many typical characteristics, including the complex working mode of the RBCC power system, strong coupling between thrust and flight state, strong nonlinear model and many complex constraints etc. This paper proposes a rapid trajectory optimization method based on sequential convex optimization to optimize the complex ascent trajectory with the RBCC power system. Firstly, a mathematical model for optimizing the ascent trajectory of the RBCC hypersonic vehicle is established to describe the angle of attack control system with and without the second-order lag. Then, the optimization model is made convex and discretized based on the convex optimization theory, and a trajectory optimization strategy is improved. Finally, taking the maximum terminal mechanical energy as optimization objective, the ascent trajectory optimization is simulated in cases with and without the second-order lag in the angle of attack control system. The simulation results show that the proposed mathematical model and the ascent trajectory optimization method are rapid and reliable and that the optimization results meet the characteristics of the RBCC power system and provide reference for the application of a RBCC power system and the design of an angle of attack control system.

Drag reduction characteristics of recirculation flow at rocket base in an RBCC engine under ramjet/scramjet mode

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.

The Influence of External Flow Field on the Flow Separation of Overexpanded Single-Expansion Ramp Nozzle

Flow separation and transitions of separation patterns are common phenomena of nozzles working with a wide Mach range. The maximum thrust method is applied to design the single-expansion ramp nozzle (SERN) for specific operating conditions. The nozzle is used to numerically simulate the transition processes of separation patterns under the linear change in the external flow Mach number and the actual trajectory take-off condition of a rocket-based combined cycle (RBCC), to investigate the mechanism through which the external flow field influences the separation pattern transition during acceleration. The computational fluid dynamics (CFD) method is briefly introduced, followed by experimental validation. Then, the design procedure of SERN is described in detail. The simulation results indicate that as the external Mach number increases, the flow field in the nozzle undergoes transitions from RSS (ramp) to FSS, and finally exhibits a no-flow separation pattern. The rate at which the external Mach number varies has little effect on the transition principle of the nozzle flow separation patterns, but it has a significant effect on the critical Mach number of the transition points. The external flow field of the nozzle has an airflow accumulation effect during acceleration, which can delay the transition of the flow separation pattern.

Combustion organization strategies for a variable geometry RBCC combustor under low total temperature conditions

A rocket-based combined cycle (RBCC) engine experiences a low Mach number phase during flight operations. Through combustor geometry adjustment technology, the engine can combust more efficiently under low-temperature inflow conditions during this phase, thereby improving the engine efficiency. In this paper, we investigate the fuel blending and combustion processes in rocket ramjet and pure ramjet modes under Ma 2.5 low-temperature inflow conditions, analyze the fuel combustion and flow and engine performance changes under different operating modes, and propose a new method for centralized heat release combustion organization under low-temperature inflow conditions. The method is based on the RBCC engine combustor with a geometry throat and relies on both direct-connect experiments and numerical calculations. The results show that (1) when the primary rocket operates, its jet reacts with the inflow in the shear layer and ignites the kerosene fuel injected by the pylons, forming a high-temperature zone downstream of the fuel pylons. When the primary rocket stops working, this flame is stabilized in the low-speed reflux region at the exit of the rocket by injecting a small equivalent ratio of fuel in the isolator to mix with the air. This approach can be a good substitute for the rocket jet to stabilize the flame. (2) When the primary rocket operates, the secondary fuel is mainly concentrated in the area from the fuel pylons to the geometric throat. After the rocket is turned off, the interval of heat release from combustion is also concentrated in the combustor from the fuel pylons to the geometric throat. (3) After the primary rocket stops working, the performance of the combustor improves significantly.

Spatiotemporal flow evolution in a rocket-based combined-cycle inlet during ejector-to-ramjet mode transition

The spatiotemporal distribution characteristics and flow stability of a rocket-based combined-cycle (RBCC) inlet during the ejector-to-ramjet mode transition are investigated numerically. The operational pressure of the embedded rocket is adjusted to three different levels, and the time-sequences of the rocket and back pressure regulation are varied. The pressure in feature sections is monitored to reveal the coupling relationship and stability of the internal flowfield. The inlet is more adaptable to severe disturbances under the “throttle-maintaining” regulation and is susceptible under the “direct-shutdown” regulation. The severe fluctuation period is relatively short within “medium throttle-maintaining,” while is lengthy within the “high throttle-maintaining.” The severe fluctuation under the direct-shutdown develops with the propagation of the regulation and decays with its establishment. The ultimate flowfields driven by different time-sequences reach unanimity with the same adjustable parameters of embedded rocket and back pressure; however, the dynamic evolutions show distinct characteristics. During the mode transition, pressure “valleys” are formed in any selected sections with the rocket regulations, and “peaks” are developed in many sections due to the propagation of back pressure or the instability of the rocket jet. For the medium throttle-maintaining regulation, the effect of time-sequence on the flowfield is relatively weak. For the high throttle-maintaining regulation, the pressure disturbance rises abruptly under the rocket priority regulation, with a most severe amplitude of 100.7%. For the direct-shutdown regulation, the maximum pressure disturbance of 125% is observed within the rocket priority regulation, and the minimum disturbance occurs within the back pressure priority regulation.

Thermodynamic performance modeling, optimization and numerical simulation of RBCC ejector mode

Ejector mode is a unique and critical phase of wide-range rocket-based combined cycle (RBCC) engine. In this paper, a quasi-one-dimensional thermodynamic performance modeling method, with more detailed model treatments for the inlet-diffuser system, primary/secondary flow interaction, and pressure feedback matching, was developed for operating characteristics studies and multi-objective optimization analysis of the ejector mode of an actual RBCC engine. A series of three-dimensional simulations of separate inlet and full flowpath was completed to validate the modeling study and provide further insight into the operating characteristics. The primary/secondary equilibrium pressure ratio functions a significant effect on ejector mode performance, a higher performance augmentation can be obtained by lower rocket pressure ratio, larger mixing section area ratio, smaller throttling throat and higher equivalence ratio, within an appropriate range. The positive performance augmentation can be realized at low flight Mach conditions, the coordination and trade-off relationships between specific impulse, performance augmentation ratio and thrust-to-area ratio during ejector mode are present by the Pareto-front from MOP analysis. It is further verified by CFD simulation that, the operating back-pressure at the exit of inlet-diffuser system functions a decisive influence on the airbreathing characteristics, the pressure feedback and matching should be well-controlled for secondary flowrate and performance augmentation. The thermodynamic modeling analysis results are basically consistent with those of numerical simulation, to validate the rationality and effectiveness of the modeling method.

Thrust performance of the rocket-based combined-cycle engine under ejector mode

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.

Numerical study on the dynamic process of ramjet/scramjet mode transition in the integrated full flow path for RBCC engine

As the inflow velocity increases, rocket-based combined cycle(RBCC) engine typically require a ramjet/scramjet mode transition within the flight range Ma∞ = 5.5 to 6.0. However, during the mode transition process, the reduction of the fuel mass flow rate usually causes the combustion intensity in the engine to drop sharply in a very short time, which will further lead to severe thrust fluctuations. So some technical methods need to be applied in this process to realize smooth mode transition. In this paper, different throttling strategies (kerosene mass flow rate dominant and rocket mass flow rate dominant) were designed and employed to study the dynamic mode transition process in the integrated RBCC engine flow path through three-dimensional numerical simulation for flight Ma∞ = 5.5 and Ma∞ = 6.0, respectively. The results show that when RBCC engine operates under the condition of large mass flow rate of fuel, the wall flow separation zone caused by oblique shock wave and combustion cannot be effectively improved by the bleed block and the inlet is still in the limit working state. Under flight Ma∞ = 5.5 condition, with the reduction of fuel equivalence ratio, RBCC engine successfully transited from ramjet mode to scramjet mode, and the flow separation zone near the bleed block disappeared. At this time, the high temperature gas of the rocket has no obvious effect on the thrust gain of the engine. By adjusting fuel mass flow rate in a small range and rocket mass flow rate in a large range, stable mode transition can be achieved, with relatively small time consumption and minimum thrust fluctuation during mode transition. With the increase of flight altitude, under flight Ma∞ = 6.0, the incoming flow speed increases, but the air capture ability of the inlet decreases, which is not conducive to the mixing of fuel and air combustion. When Ma∞ = 5.5, the mode transition by reducing the fuel equivalence ratio will cause the RBCC engine to produce huge thrust fluctuations, and the range of the high-pressure zone in the combustor will be greatly reduced. In order to ensure the stable operation of the RBCC engine, the rocket needs to operate at a larger mass flow rate.

Investigation of the influence mechanism of geometric throats in variable-geometry rocket-based combined-cycle combustors

To explore the reaction mechanism a geometric throat and its influence on the performance of the combustor, this paper studies the working process of a rocket-based combined-cycle (RBCC) variable geometry combustor with geometric throats under Ma4 inflow conditions by both numerical simulation and ground direct-connect testing. The main research contents and conclusions are as follows: (1) The geometric throat can accurately control the choking position. The fuel in the variable geometry combustor can complete the concentrated heat release and bring higher combustor pressure. The thrust of the geometric throat combustor is 8.7% higher than that of the thermal throat combustor. Therefore, the geometric throat combustor provides a better combustor performance than the thermal throat combustor. (2) When the throat area of the Ma4 configuration is adopted, the thrust and specific impulse in the combustor are at their the highest, reaching 6.0% and 0.4% higher than those of the Ma5 and Ma3 throat configuration combustors, respectively. This shows that a certain amount of combustor heat release needs to match the corresponding geometric throat area to ensure the optimal performance of the combustor. (3) In the Ma4 combustor configuration, after closing the primary rocket, the thrust in the combustor decreases by approximately 10.8%, but its specific impulse performance greatly improved by 39.3%. Therefore, the geometric throat can be used in the RBCC mode transition process in combination with the secondary fuel injection strategy to avoid the thrust trap in the mode transition process.

Numerical study on the dynamic process of ramjet/scramjet mode transition in the integrated RBCC full flow path through multistage fuel injection strategies

The rocket-based combined-cycle (RBCC) engine usually completes the transition process from ramjet to scramjet mode under the Ma∞ = 5.5–6.0. The dual-mode (ramjet/scramjet) transient process is an important technology for engine operation. CFD modeling based on Reynolds average Navier-Stokes was for studying the influence of fuel injection position, high-temperature rocket jet and fuel throttling time on the combustion flow structure and transient characteristics of RBCC engine during the mode transiton . The results showed that whether RBCC engine could successfully achieve stable mode transition was mainly influenced by flame stability in the combustor. When the rocket shuts down, it was more suitable to choose the fuel injection arrangement which distributed downstream of the combustor to realize relatively smooth mode transition. The analysis shows that in the mode transition process, using the synergistic combustion-supporting effect of the cavity and the strut is beneficial to improve the fuel combustion efficiency and reduce the thrust fluctuation. Although the high temperature rocket jet with low mass flow rate cannot maintain the fuel combustion in the isolator at high Mach number, it can expand the high temperature combustion area width in the central flow of RBCC engine and increase the secondary fuel combustion efficiency. Therefore, keeping the rocket running at low mass flow rate can reduce the thrust drop caused by insufficient fuel heat release. When adopting a mode transition strategy that throttles the fuel linearly over time, the engine can go through the thrust dip more smoothly, but this correspondingly increases the time required for the internal flow field to stabilize.

Study on SABRE and rocket-based combined cycle engine

Since the 20th century, countries have been conducting researches on near-spacecraft and have proposed a series of representative space vehicle concepts. This paper reviews the brief history of the development of reusable near-spacecraft and focuses on two combined cycle propulsion modes, SABRE and RBCC, with the goal of improving the performance of future near-spacecraft propulsion systems. An improving idea of combining the RBCC with the SABRE is proposed, and its possibility is discussed by analyzing the specific impulse curve. Two combining configurations of them and design principles are obtained by citing the literature and studying similar configurations. The two designs are also compared and the more optimized one is suggested, so as to offer some inspiration for future studies.