RBCC Research Articles

Investigation on geometric throat matching characteristics of variable geometry rocket-based combined cycle combustor

One of the most effective ways to realize multi-mode and high-performance matching operation of rocket-based combined cycle (RBCC) engine is to adopt a variable geometry combustor. In this study, direct-connect experiments and three-dimensional numerical simulations were used to investigate the matching characteristics between the isolator and the geometric throat of the variable geometry rocket-based combined cycle combustor under Ma 6 condition. The influences of equivalence ratio and geometric throat height variations on the combustor matching characteristics were obtained. The study results show that: (1) The pressure in the combustor and the outlet pressure of the isolator increase gradually with the increase of equivalence ratio, which causes the pre-combustion shock train position in the isolator to move forward gradually. Meanwhile, the combined injection of secondary fuel with the isolator and the pylon is beneficial to the improvement of combustor pressure, thus further improving the engine performance. (2) Different geometric throat heights have a substantial impact on the isolator and combustor matching when the equivalent ratio of the combustor is fixed. Under the conditions of Ma 6 operating at an altitude of 24 km inflow, the starting position of the pre-combustion shock train in the isolator is precisely at the isolator entrance, and the corresponding geometric throat height is 2.10H. The isolator and the combustor can complete matching operation at this point which the geometric throat height is the minimal throat height for Equivalent ratio = 1.0. (3) The reduction in geometric throat height can result in an increase in combustor pressure but a loss in combustor heat release and combustion efficiency of fuel combustion, which are reduced by 21.1% and 7.2% respectively. (4) The heat release and combustion efficiency in the combustor will decrease with a decrease in equivalent ratio at the minimal throat height condition where the isolator and the combustor can complete matching operation. This happens when the equivalent ratio and the geometric throat height change synergistically.

Nitrogen Jet Array for Ignition Enhancement in a Rocket-Based Combined Cycle Combustor

In a rocket-based combined cycle combustor operating at scramjet mode, the size of the cavity recirculation flow might be reduced when the cavity is subjected to an expansion fan, which is detrimental to ignition. To address this challenge, a method based on a nitrogen jet array is proposed and evaluated in a combustor with a Mach 2.92 inflow. Nitrogen is supplied through injectors in the cavity. In the case without nitrogen injection, the cavity recirculation flow is so small that the spark plug cannot ignite the combustor. In stark contrast, the interaction between the supersonic inflow and the nitrogen jets on the cavity ramp significantly elevates the pressure in the cavity, and it enlarges the cavity recirculation flow. During the ignition process, the enlarged cavity recirculation flow is essential to holding the weak self-sustained flame. The heat, hot products, and radicals generated by the weak self-sustained flame elevate the pressure in the cavity and further enlarge the cavity recirculation flow, which paves the way for a successful ignition. Once a successful ignition is achieved, turning off the nitrogen jet array does not lead to flame extinction. Compared to the mass flow rate of the inflow, the nitrogen consumption for ignition enhancement is minor.

Effect of the heat release on ejector-to-ramjet dynamic mode transition process in the rocket-based combined cycle engine

Effect of the heat release on ejector-to-ramjet dynamic mode transition process in the rocket-based combined cycle engine

Experimental detection of two-dimensional temperature distribution in Rocket-Based Combined Cycle combustion chamber using multispectral imaging processing

The extremely harsh environment and highly unsteady and transient nature during supersonic combustion bring great challenges for quantitative diagnostics. In the present work, for the first time, we report measurements of two-dimensional distributions of flame temperatures in a Rocket-Based Combined Cycle (RBCC) combustion chamber using a multispectral imaging technique with both spatial and spectral resolutions. The relation between the multispectral radiation intensity images and the flame 2-D temperature distributions was established by combing the Lvevenberg-Marquardt algorithm and the Hottel and Broughton emissivity models. The experimental results revealed that both laser-induced plasma (LIP) and plasma jet (PJ) torches had positive effects on combustion stabilization, and LIP had better effect on combustion enhancement. In the ignition stage of a fuel/air mixture with an equivalence ratio of 1.0, the average temperature of a flame initiated by LIP combustion enhancement was 100 K higher than that by PJ torch combustion enhancement. A recirculation zone was formed inside a cavity, which improved the combustion stability. The injected fuel and inflow were most fully mixed on the slope at the rear of the cavity, where 70 % and 65 % of the peak flame temperature points of LIP and JP were located, respectively. The combustion in the RBCC chamber showed an oscillation phenomenon that peak flame temperature and the flame structure oscillated continuously around the central axis with a horizontal extent of 162 mm. Due to the insufficient fuel blending, the oscillation amplitude was minimal when the equivalence ratio was 1.0. Overall, the results of flame temperature distributions demonstrated that the proposed multispectral thermometry was suitable for measuring temperatures of supersonic combustion flames.

A Mixing Process Influenced by Wall Jet-Induced Shock Waves in Supersonic Flow

With the development of hypersonic air-breathing propulsion systems, such as the supersonic combustion ramjet (Scramjet) and rocket-based combined cycle (RBCC) engines, the mixing process of supersonic airstream with fuel in the engine combustor has been drawing more and more attention. Due to the compressibility effects, the mixing process in a supersonic condition is significantly inhibited. In the present paper, the novel strategy of wall-jet induced shock waves (WJISW) is put forward to realize mixing enhancement. The interaction process between WJISW and the supersonic mixing layer is researched and the enhanced-mixing mechanism is revealed, employing large eddy simulation (LES) methods. The fine vortex structures of the flow field are well captured and presented, utilizing the numerical schlieren technique. Detailed visualization results indicate that WJISW in a low frequency condition can result in the ‘region action mode’ (RAM) never reported before. The drastic dynamic behaviors including growth, deformation, and distortion in the interaction region can undoubtedly promote the mixing of upper and lower streams. The Reynolds stress distributions along the streamwise x-direction suggest that more intense fluctuations can be achieved with a low frequency WJISW. Moreover, a sharp increase in mixing layer thickness can be realized in the interaction region. The dynamic mode decomposition (DMD) analysis results show that the mixing layer evolution process is dominated by the mode induced by WJISW, which leads to the coexistence of both large- and small-scale structures in the flow field. The entrainment process corresponding to large-scale vortices and the nibbling process corresponding to small-scale vortices can obviously promote mixing enhancement. It is suggested that the present proposed strategy is a good candidate for enhanced-mixing with application to Scramjet and RBCC combustors.

Numerical study on the lobed nozzles for enhancing the mixing and combustion performance of rocket-based combined cycle engine

Numerical study on the lobed nozzles for enhancing the mixing and combustion performance of rocket-based combined cycle engine

Ejector mode performance improvement of rocket-based combined cycle engine designed for dual-ramjet mode

Rocket-based combined cycle engine integrating multi-mode operation accomplishes an excellent performance under various flight speed conditions. Generally researchers selected dual-ramjet mode as the cruise and design point with adjusting heat addition method to improve the performance under other modes. To analyze the difficulties during the multi-mode design, a unilateral expansive flowpath through central rocket ejection was investigated to analyze the effects of the mixing ratio of oxygen and fuel and the mass flow rate of the primary rocket plume. The results indicate that the mixing ratio increase can significantly strengthen the entrainment amount. There is an optimal operating rate to obtain the highest entrainment ratio due to Fabri choking limit induced by a confined path width. However, the excessively divergent afterbody based on dual-ramjet results in the under-expansion issue of the exhaust gas and negative thrust of the nozzle. The multistage ejection method can effectively overcome this shortage by compressing the exhaust gas. But with the flight speed increasing, the existence of upstream bow shock induces high inlet resistance. Adding extra duct can effectively solve the problem by decreasing inlet convergent ratio and sucking more secondary flow for combustion.

Effect of area ratio and heat release method on the Ma∞ = 2-6 operation performance of rocket-based combined cycle engine

Ballistic trajectories with rapidly changing and variable ambient conditions impose higher requirements on integral multi-component rocket-based combined cycle (RBCC) engines. A central strut-based RBCC engine was numerically simulated by solving Reynolds-averaged Navier-Stokes equations together with a k-ω SST turbulence model under various flight conditions, Ma∞=2 to Ma∞=6, which cover ejector, ramjet and scramjet modes. Ground direct-connect tests were carried out to analyze heat release performance in the combustor and the pressure balance relationship between various components. Results indicate that scramjet combustion should be conducted in the foremost combustor where geometry expansion is slightest. As the oncoming velocity decreases, the combustor expansion ratio should be increased and the location of heat addition moves to downstream larger area region. Distributed fuel injection with 20%∼30% upstream is beneficial for increasing the initial temperature of the mixing gas and thus the combustion efficiency under Ma∞3-6 conditions. At lower flight Mach numbers (Ma∞<3), different dual-ramjet engine, a section with larger area ratio, over 2, should be added behind the ramjet combustor. Since off-design inlet operation causes a decrease on the intake mass flow rate, concentrated heat addition should be adopt to ensure effective thermal choking.

Start Limits and Positive Control of RBCC Inlet

AbstractA rocket-based combined-cycle (RBCC) inlet is always required to start at a relatively low Mach number in consideration of the fuel consumption in the ejector mode and the comprehensive eng...

Experimental and Numerical Investigations on the Ignition of RBCC Gas-Oxygen/Kerosene Rocket Ejector

The rocket ejector refers to a core component of a rocket-based combined cycle (RBCC) engine. The ignition is of critical significance for rocket ejection. Reliable and stable ignition crucially determines the normal operation of the engine. In this paper, a thrust chamber with coaxial swirl injector for the RBCC rocket ejection was developed and tested. Gas oxygen (GOX) and kerosene acted as propellants. As revealed from the test results, the process of ignition pressurizing comprised four phases. The oxygen prefilling time before ignition slightly impacted the ignition time, whereas it affected the peak pressure of ignition. In a confined range, the peak pressure decreased as the prefilling time was extended. The ignition was simulated by building a numerical model, and the results well complied with the experimentally achieved results. The numerical model is capable of specifically indicating the position of the kernel of fire and the process of flame propagation. The simulation results reveal that the propellant could form a combustible condition within 4 ms. The kernel was 6 mm away from the injector, located at the oxygen and kerosene mixing interface and approaching the upper wall. The above results reflected the vital role of the central recirculation zone formed by the prefilled oxygen. The ignition energy was transported near the injector under the convection effect, which ignited the stoichiometric mixture, and the entire ignition could reach a stable state within 20 ms. The numerical model which was developed in this paper can help clarify the combustion mechanism.

Numerical study on the induced-ramp mixed and combustion characteristics of rocket-based combined cycle engine

In the study, the supersonic flow mechanism and mixed characteristics are discussed by changing the ramp angle and momentum flux ratio of the rocket-based combined cycle engines. The three-dimensional Reynolds-averaged Navier-Stokes equations, the shear stress transport k-ω turbulence model, and the Eddy-dissipation concept model are adopted to analyze the flow field structure. The results reveal that the oblique shock wave causes the rocket gas and parallel shear layer to slope downward, thus promoting more efficient combustion. The oblique shock wave interacts with the bottom wall boundary layer and forms a low-velocity region. The increase of the ramp angle is beneficial for jet to enter the combustion chamber. The ramp angle increases from 20° to 40°, and the combustion efficiency increases by about 49%. Furthermore, the increase of the momentum flux ratio is conducive to induce vortex structure, and the best penetration depth of jet fuel is obtained when the momentum flux ratio is 0.72.

Vortex evolution and flame propagation driven by oblique shock wave in supersonic reactive mixing layer

Vortex evolution and flame propagation of supersonic reactive mixing layer with the interaction of single/double oblique shock waves are researched by employing direct numerical simulation methods. The vortex dynamics are found to be modulated by oblique shock waves, resulting in mixing enhancement of vorticity values. Notably, wider range of vortex scale distribution induced by double oblique shock waves leads to intense entrainment and nibbling process of large and small scale structures, which can remarkably increase the mixing level of hydrogen and airstreams. The analysis of vorticity transport equation indicates that compressing effects induced by shock waves are responsible for the increase of vortex strength. By tracking the vortex braid dynamical evolution process, it is revealed that local auto-ignition induced by double shock waves can accelerate the consumption of hydrogen and finally lead to high efficiency combustion. Since complex background waves are inherently present in confined space of Rocket-based Combined Cycle (RBCC) engine, the newly observed mixing enhancement mechanism for double shocked-reactive mixing layer can provide a guidance for the future design of mixing strategy using the inherent wave structures in RBCC combustor chambers.

Ignition dynamics and combustion mode transitions in a rocket-based combined cycle combustor operating in the ramjet/scramjet mode

Cavity ignition was investigated in a rocket-based combined cycle combustor. The Mach number, total pressure, and total temperature of the inflow were 2.92, 2.60 MPa, and 1650 K, respectively. A backward-facing step was employed to replace the rocket engine that was shut down. The expansion fan and the reattachment shock wave resulting from the backward-facing step distorted the mainstream. The combustor was 40 mm in height (H) at the entrance. Three typical cases with d/H=2.0, 4.5, and 7.5 were reported in this study, where d was the distance between the backward-facing step and the cavity. Schlieren images and flame chemiluminescence images were captured to reveal the ignition processes. The ignition process in the case with d/H=2.0 was similar to the conventional cavity ignition because only the rear of the cavity was subjected to the expansion fan. Two combustion modes, namely, the local combustion mode and the cavity stabilized scramjet mode, were witnessed in this case. The first combustion mode was characterized by a small reaction zone in the fore part of the cavity. The primary reaction zone of the second mode was in the region near the cavity ramp. The rapid flame spreading in the cavity shear layer led to the transition between these two combustion modes. In the case with d/H=4.5, the reattachment shock wave enlarged the cavity recirculation flow. The expansion fan decreased the static pressure upstream of the cavity, which made the combustor more sensitive to backpressure. Although the first two combustion modes of this case were similar to those of the case with d/H=2.0, a new combustion mode (unsteady jet-wake stabilized scramjet mode) was observed in this case. The mode transition between the cavity stabilized scramjet mode and the unsteady jet-wake stabilized scramjet mode was attributed to the backpressure resulting from the heat release. The backpressure gradually pushed the reaction zone upstream. In the case with d/H=7.5, the cavity was not directly exposed to the expansion fan or the reattachment shock wave. However, the ignition process in this case was more complex. The combustor successively experienced the aforementioned three combustion modes and eventually settled into the ramjet mode. The mode transition between the last two modes was the result of thermal choking. The flame propagated upstream rapidly when the combustor was chocked.

A numerical study on flow structure and combustion mechanism of supersonic mixed inflow with transverse jet

The flow mixing mechanism between sonic fuel jet and supersonic mixed inflow is one of the key technologies in rocket-based combined cycle engines. In this study, the flow mechanism of different supersonic airflow conditions on the flow field is discussed. The variable parameters are the Mach number and total pressure of the airflow. The three-dimensional Reynolds-averaged Navier-Stokes equations, two equations of the shear stress transport k-ω turbulence model, and the Eddy-dissipation concept reaction model are adopted to evaluate the flow field structure. The results reveal that the shear layer between the airflow and the rocket gases gradually slopes downward with the increase of Mach number. The bow shock induced by the jet develops to the upper wall and cuts off the shear layer vortex. Additionally, as the Mach number of the airflow decreases, the mixing and combustion efficiencies gradually increase. The Mach number has a slight influence on the penetration depth. Furthermore, when the total pressure of the airflow increases, the penetration depth and mixing efficiency sharply increase.

A survey on the conceptual design of hypersonic aircraft powered by RBCC engine

Rocket-based combined-cycle (RBCC) powered vehicles have been widely recognized as the most promising aircraft solution that could dramatically reduce the cost of space transportation. Researchers and scientists worldwide have conducted considerable overall design researches to cope with the challenges in RBCC development including mode transition, thermal protection and thrust enhancement. According to the way to orbit and the configuration characteristics, the hypersonic aircraft powered by RBCC engine are classified as four categories: single-stage-to-orbit (SSTO) two-dimensional configuration, SSTO axisymmetric configuration, two-stage-to-orbit (TSTO) two-dimensional configuration, and TSTO axisymmetric configuration. This paper systematically presents the development of the conceptual design of RBCC-powered vehicles. Both the structural and operating key parameters like the weight distribution, the RBCC propulsion performance and take-off mode, et al. are introduced in detail. On this basis, a comparative analysis of the advantages and disadvantages of the orbit model, the configuration selection and takeoff modes are conducted. In addition, the application prospect and technology development direction for hypersonic aircraft are also discussed. At the same time, the lessons that can be drew from previous hypersonic vehicle concept design are explored.

Secondary fuel jet strategies on mixing enhancement performance of rocket-based combined cycle engine

The mixed augmentation performance between secondary fuel jet and uneven supersonic inflow is a significant evaluation index of the rocket-based combined cycle engine. In this study, the effects of the secondary fuel jet on the flow field structure and mixing enhancement performance are discussed by changing both the secondary fuel jet angle and spacing of multiple rows of jet holes. The three-dimensional coupled implicit Reynolds averaged Navier-Stokes equations, SST k-ω turbulence model, and eddy dissipation concept (EDC) reaction model were adopted to evaluate the flow field structure with a sonic hydrogen jet. It is indicated that the 135° fuel jet has the largest penetration depth value, followed by the 90° fuel jet, whereas the 45° fuel jet has the smallest penetration depth value. However, the 45° fuel jet possesses the best mixing performance, it easily induces a pair of downstream counter-rotating vortex, and the oblique torque is further generated to aggravate the tearing ability to the jet. The three jet holes induce a stronger shock wave and form a larger barrier under the 90° jet. Additionally, when the spacing between the multiple rows of jet holes reduces, penetration depth is easily created. Further, when the spacing of the three rows of jet holes is larger, the fluctuation of the mixing coefficient increases.

Design and Numerical Investigations on a Dual-Duct Variable Geometry RBCC Inlet

Abstract A widely applicable and variable geometry 2-D rocket based combined cycle (RBCC) inlet characterized by the dual-duct design is conceptually put forward. The inlet operates as dual-duct status in the low Mach range (0~4), and transits to single-flowpath status in the following high Mach range (4~7). It accomplishes operational status transition through an 8.0-degree ramp rotation and a 4.0-degree cowl rotation at Mach 4. Through numerical simulations on typical flight Mach numbers, the observed starting Mach number is 2.2, which provides a sufficient operational window for a smooth ejector-to-ramjet mode transition. The RBCC inlet achieves comprehensive high mass capture coefficients in the overall wide flight range, especially in the low speed regimes. Suitable Mach numbers satisfying various combustion requirements in different modes together with high total pressure recovery coefficients are also obtained since the physical throat areas, compression angles, and the corresponding contraction ratios can be adjusted by a large margin through very limited rotations. The variable geometry scheme is not only feasible for practical realizations, but is also simple to arrange the dynamic sealing issues in a low-temperature environment in the RBCC engine.

Cavity flameholding subjected to an expansion fan in a rocket-based combined cycle combustor

The effect of expansion fan on cavity flameholding was investigated in a rocket-based combined cycle combustor fueled by ethylene. The inflow of the combustor was at a condition where the Mach number, stagnation temperature, and total pressure were 2.92, 1650 K, and 2.6 MPa, respectively. The global equivalence ratio was 0.25. A backward-facing step (BFS) was used to simulate a rocket that was shut down and create an expansion fan. The dimensionless distance between the BFS and the cavity (d/H) varied from 1.5 to 7.5, where d = 60–300 mm was the actual distance and H = 40 mm was the height of the combustor entrance. The result indicated that the cavity shear layer reattached at the cavity floor in the combustor with 2.5 ≤d/H≤ 3.5 because of the pressure gradient resulting from the expansion fan. The expansion fan also significantly decreased the static pressure in the cavity. As a consequence, the combustor with 2.5 ≤d/H≤ 3.5 failed in ignition. In the combustors with d/H = 4.0 and 6.0, the cavity flows were open-type. However, the ignition attempt failed because the expansion fan and the reattachment shock wave might worsen the local equivalence ratio in the cavity. Successful ignition and flame stabilization were achieved in the combustors with d/H = 1.5, 2.0, 4.5, and 7.5. Three combustion modes and three kinds of combustion oscillation were identified. In the combustor with d/H = 1.5, only a small part of the cavity was exposed to the expansion fan. The combustor operated in the cavity stabilized scramjet mode because the flame front was anchored at the cavity leading edge. The cavity shear layer impingement onto the cavity ramp lead to the combustion oscillation with a frequency ranging from 300 to 400 Hz. The combustion mode and combustion oscillation in the combustor with d/H = 2.0 were the same as their counterparts in the combustor with d/H = 1.5. Unsteady jet-wake stabilized scramjet mode was witnessed in the combustor with d/H = 4.5. In this combustor, the fuel jet and the front of the cavity were exposed to the expansion fan. The reattachment shock wave lifted the cavity shear layer. The combustion oscillation induced by shock-flame interaction showed a dominant frequency at 107.1 Hz. The expansion fan and the reattachment shock wave were upstream of the cavity in the combustor with d/H = 7.5. Both of them made the boundary layer more sensitive to adverse pressure gradient. The flame front was upstream of the cavity because the heat release resulting from vigorous chemical reaction choked the combustor and induced large-scale boundary layer separation. The combustor operated in ramjet mode. The unsteady shock train and thermal throat induced the combustion oscillation with a frequency of approximately 800 Hz.

Artificial neural network based chemical mechanisms for computationally efficient modeling of hydrogen/carbon monoxide/kerosene combustion

To effectively simulate the combustion of hydrogen/hydrocarbon-fueled supersonic engines, such as scramjet and rocket-based combined cycle (RBCC) engines, a detailed mechanism for chemistry is usually required but computationally prohibitive. In order to accelerate chemistry calculation, an artificial neural network (ANN) based methodology was introduced in this study. This methodology consists of two different layers: self-organizing map (SOM) and back-propagation neural network (BPNN). The SOM is for clustering the dataset into subsets to reduce the nonlinearity, while the BPNN is for regression for each subset. Compared with previous studies, the chemical reaction mechanism involved in this study is more complex, therefore, the particle swarm optimization (PSO) method is employed for accelerating training process in this study. Then we were committed to constructing an ANN-based mechanism of hydrogen and kerosene for supersonic turbulent combustion and verifying it in a practical RBCC combustion chamber. The training data was generated by RANS simulations of the RBCC combustion chamber, and then fed into the SOM-BPNN with six different topologies (three different SOM topologies and two different BPNN topologies). Through LES simulation of the Rocket-Based Combined Cycle (RBCC) combustor, the six ANN-based mechanisms were verified. By comparing the predicted results of six cases with those of the conventional ODE solver, it is found that if the topology is properly designed, high-precision results in terms of ignition, quenching and mass fraction prediction can be achieved. As for efficiency, 8~20 times speedup of the chemical system integration was achieved, which will greatly improve the computational efficiency of combustion simulation of hydrogen/carbon monoxide/kerosene mixture.

Research and development on inlets for rocket based combined cycle engines

Rocket based combined cycle (RBCC) engine is regarded as one of the most efficient propulsion systems for reusable space transportation vehicles and hypersonic cruise vehicles. It can perform excellently over a very broad range of Mach numbers and flight altitudes through multi-mode operations by effectively integrating high thrust-to-weight ratio of the rocket engine and high specific impulse of the airbreathing engine. The RBCC inlet, commonly operating in an extremely wide flight speed range covering static, subsonic, transonic, supersonic, and hypersonic, plays a vital role in the RBCC propulsion system and flight vehicles. This paper presents a comprehensive and systematical review of the research progress on the inlets for RBCC engines in different countries and different historical periods. A number of key technical and scientific issues specially concerning RBCC inlets are sorted out, including the wide range RBCC inlet design and optimization, inlet/embedded rocket/combustor matching and control, dynamic operation, etc. Several typical RBCC inlet applications and the corresponding research are discussed, including the Strutjet inlet, A5 inlet, GTX inlet, and the Astrox's inlet in the U.S., the RBCC inlet in JAXA, and a sidewall compression RBCC inlet in China. Finally, experience gained and some recommendations are summarized for future research of the RBCC inlet.